SECgGEORGEGe

JUNE 25, 1968

4s

Unel asDufkS OPSTUTS

SATURN V LAUNCH VEHICLE

FLIGHT EVALUATION REPORT-AS-502

APOLLO 6 MISSION )

PREPARED BYSATURN V FLIGHT EVALUATION WORKING GROUP

SMSF Form 774 Rev Octoher 1967)

56 Cepee

GEORGE C. MARSHALL SPACE FLIGHT CENTER

MPR-SAT~FE-68-3

SATURN V LAUNCH VEHICLE

FLIGHT EVALUATION REPORT - AS-502

APOLLO 6 MISSION

PREPARED BY

‘SATURN ¥ FLIGHT EVALUATION WORKING GROUP

AS-502 Laurch Vehicle

MPR-SAT-FE-68-3

SATURHK Y LAUNCH VEHICLE FLIGHT EVALUATION REPORT - AS-S02APOLLO 6 MISSION

BY

Saturn Flight Evaluation Working GroupGeorge C. Marshal] Space Flight Center

ABSTRACT

Saturn V AS-502 (Apollo 6 Mission) was launched at 07:00:01 Eastern Stan-dard Time on April 4, 1968, from Kennedy Space Center, Complex 39, Pad A.The vehicle lifted off on schedule on a launch azimuth of 90 degrees eastof north and rolled to a flight azimuth of 72 degrees east of north.

The actual trajectory parameters of the AS-502 were close to nominal untilthe premature shutdown of two engines in the S-II stage. After thisoccurred, the trajectory deviated significantly from the nominal through-out the remainder of the mission.

Nine of the sixteen primary objectives of this mission were completelyaccomplished, six partially accomplished, and one (S-IVB restart) was notaccomplished. One of tre two secondary objectives was completelyaccomplished, and one partially accomplished.

Any questions or comments pertaining to tae information contained in thisreport are invited and should be directed to:

Director, George C. Marshall Space Fligit CenterHuntsville, Alabama 36812Attention: Chairman, Saturn Flight Evaluation

Working Group, R-AERO-F (Phone 876-4575)

Section

TABLE OF CONTENTS

TABLE OF CONTENTSLIST OF ILLUSTRATIONSLIST OF TABLESACKNOWLEDGEMENTABBREVIATIONSMISSION PLANFLIGHT TEST SUMMARY

INTRODUCTION1 Purpose1.2 Scope

EVENT TIMES21 Summary of Events2.2 Secuence of Events

LAUNCH OPERATIONS

31 Summary3.2 Prelaunch Milestones3.3 Countdown Events34 Propellant Loading3.4.1 RP-1 Loading34.2 LOX Loadinga LH2 Loading

3.4.4 Auxiliary Propulsion System Prone] lantLoading

Page

iiixiii

xxviixxxtd

xxxiiiwo

xxvty

Section

TABLE OF CONTENTS (CONTINUED)

Mow

TRAJECTORY4a4.24.2.1

2

ape

oe

S-IC Stage Propellant LoadS-II Stage Propellant Load5-IVB Stage Propellant LoadS-II Insulation Purge and -eakDetectionGround Support fquipment

SummaryTracking Data UtilizationTracking During the Ascent Phase ofFlightTracking During Orbital FlightTrajectory EvaluationAscent TrajectoryOrbital Trajectory

S-1€ PROPULSION5.1

5.25.36.4

5.55.65.6.15.6.2

5.75.85.9

SumnaryS-IC Igniticn Transient PerformanceS-IC Main Stage PerformanceS-IC Engine Shutdown Transient PerformanceS-IC Stage Proneltant ManagementS-IC Pressurization SystemS-IC Fuel Pressurization SystemS-IC LOX Pressurization SystemS-IC Pneumatic Control Pressure SystemS-IC Purge SystemS-IC Camera Zjection and Purge System

S-11 PROPULSION6.)

6.2

SuemaryS-1L Chitldown and Bui tdup ~ransientPerformance

iv

Page62636

6-1

6-2

Section6.3

ooNome

TABLE OF CONTENTS (CONTINUED)

S-II Main Stage PerformanceS-IE Stage Shutdown Transient PerformanceS-I Propellant Management$-IL Pressurization Systems$-I1 Fuel Pressurization SystemS-I1 LOX Pressurization SystemS-IL Pneumatic Control Pressure SystemCamera Ejection SystemHelium Injection System

S-1¥B PROPULSION

77.2

7.3

24

22

7.8

7.3

7.107.19.17.10.2aU

712

SumnaryS-I¥B ChilIdown and Buildup TransientPerformance for First BurnS-IVB Main Stage Performance for FirstBurnS-1VB Shutdown Transient Performance forFirst BurnS-IVB Coast Phase ConditioningS-I¥B Shilldown and Attempted Restartfor Second BurnS-1¥3 Main Stage Performance far SecondBurn5-1¥8 Shutdown Transient Performance forSecond Burn5-1¥8 Stage Prooellant Utilization$-1V8 Pressurization System5-18 LHe Tank Pressurization SystemS-1VB LOX Pressurtzation SystenS-1¥B Pnaumatic Control System5-1¥8 Auxiliary Propulsion System

Page6-96-196-226-266-296-306-346-356-35

7-2

7-4

7-11yl

7-28

7-287-287-347-347-397-507-61

Section8

98

TABLE OF CONTENTS (CONTINUED)

HYDRAULIC SYSTEMS8.1

8.28.38.48.5

8.6

STRUCTURES

133 SECOND98.19A.298.2.19A.2.2

90.2.39A.2.4

98.2.598.39A.d9A.S

SummaryS-IC Hydraulic systemS-IL Hydraulic SystemS-I¥B Hydraulic System (First Burn)S-IV8 Hydraulic System (Coast Phase)S-IVB Hydraulic System (Second Burn)

SummaryTotal Vehicle Structures EvaluationLongitudinal LoadsBending MomentsVehicle Dynamic CharacteristicsS-I€ Fin DynamicsVibration EvatuationS-IC Stage and Engine EvaluationS-II Stage and Engine EvaluationS-IVB Stage anc Engine EvaluationS-IVB Stage Forward Skirt UynamicsInstrument Lnit Evaluation

TRANSIENTSummaryInstrument UnitMechanical Versus Electrical DisturbancePressure, Flowvate, and TemperatureMeasurements,Radio Frequency (RF) MeasurementsST-12aM-3 Stabilized ?latform Subsystemand Control Subsystem AnalysisStructures and DynanicsS-IVB StageS-II StageRF Systems

vi

Page

8-1

8-289B128-12

9A-19A-1DA-2

9A-109A-13

98-139A-159A-159a-229A-22

Section

VW

W

TABLE OF CONTENTS (CONTINUE)

9A.6 Smergency Detection Systen (EDS}

GULDANCE AND NAVIGATION10.1 Summary10.1.3 Flight Progran19.1.2 Instrument Unit Components19, Suidance and Navigation System Description10.2.1 Instrument Unit System Description

Flight Program DescriptionGuidance InteTtigence ErrorsNavigation and Guidance SchemeEvaluationInertial 2Jatform and NavigationParameter ComparisonsFlight Program zvaluationOrbital GuidanceOrbital RoutinesEvent Sequencing

1122

10.2.234

5 Guidance System Component Evaluation10.5.1 LYDC Performance10.5.2 VDA Performance10.5. Ladder Outputs10.5.4 Telemetry Outputs10.5.5 Discrete Outputs19.5.8 Switch Selector Functions10.5.7 ST-124M-3 Inertial Platform Performance

CONTROL SYSTEMWa Summary2 Zontrol System Description11.3 S-1C Control System Evaluation11.3.1 Liftoff ClearancesW132 S-IC Flight Dynamics1.4 S-I1 Controt System Evaluation11.4.7 Attitude Contro? Dynamics and StabilityVW.4.2 Liquid Propellant Dynamics and Their

Effects on Flight Control

vit

Page9A-22

10-110-110-130-210-210-219-7

19-9

19-913-1019-1719-2919-2019-2919-2019-2010-2113-2119-2119-2110-21

Ve)

11-21-3n-511-811-2511-26

11-33

Section

12

13

TABLE OF CONTENTS (CONTINUED)

i S-IVB Control System Evaluation11.5.1 S-1¥B Control System Zvaluation Before

and During First durnV1.5.2

—

Contro? System Evaluation DuringParking Orbit

11.5.3 Control System Evatuation During RestartAttempt

11.5.4 Control System Evaluation After RestartAttempt,

11.6 Instrument Unit Control ComponentEvaluation

11.6.1 Contro1-EDS Rate Gyros/Concro? SignaProcessor Analysis

11.6.2 Flight Contrat Computer Analysis

SEPARATION12.4 Summary12.2 S-IC/S-11 Separation Evalustior12.2.1 S-IC Retro Notor Performance12.2, S-I] Ullage Hotor Performance12.2, S-IC/S-H] Separation Dynamic:12.3 S-II Second Plane Separation Evaluation12.4 $-11/S-1¥B Separation Evaluation12.4.1 S-IT Retro Motor Performance12.4.2 S-1¥8 Ul Tage Motors12.4.3 S-II/S-IVB Separation Dynamics12.5 S-IVB/IU/Spacecraft Separation Evaluation

ELECTRICAL NETWORKS13.113.213.3

13.413.5

SummaryS-IC Stage Electrical SystemS-II Stage Electrical System$-IVB Stage Electrical SystemInstrument Unit Electrica? System

vit

Page11-36

11-36

11-45

11-84

11-54

11-68

11-5811-58

12-112-312-312-312-312-712-1412-1412-1812-1512-21

13-113-2

13-513-913-15

Section14

aE:

16

7

TABLE OF CONTENTS (CONTINUED)

RANGE SAFETY AND COMMAND SYSTEMS.1,

.214,

4,

1

3

SummaryRange Safety Command SystemsCommand and Communications System

EMERGENCY DETECTION SYSTEM15.18.15,18.3.18.3.16.3.15.3,is.

SummarySystem DescriptionSystem EvaluationGeneral PerformancePropulsion System SensorsFlight Dynamics and Control SensorsEDS Sequential Events

Interface Considerations

VEHICLE PRESSURE AND ACOUSTIC ENVIRONMENT16,16.

16.16.16.

12

BERBWwNNN

SummarySurface Pressures and ConpartmentVenting$-1C StageS-II Stage5-198 StageBase PressuresS-IC Base PressuresS-I] Base Pressures

Acoustic EnvironmentExternal AcousticsInternal Acoustics

VEHICLE THERMAL ENYE RONMENT17,

Wa1

2

17.2.7

SutiraryS-IC Base Heating and SeparationEnvironment§-1C/S-11 Separation

ix

Page

141Ww14-2

15-1

1-116-115-115-315-3515-5

16-1

16-136-116-516-916-1216-1216-1216-1916-1916-20

TABLE OF CONTENTS (CONTINUED)

Section3 S-II Base lleat Shield and Separation

Environment74 $-11/S-1VB Separation Environnent12.5 Vehicle Aeroheating Thermal Environment17.5. S-IC Stage Aeroheating EnvironmentW738. S-II Stage Aeroheating Environment17.5.3 $-1¥B Stage Aeroheating Environment17.5.4 Instrument Unit Aeroheating Environment17.6 Vehicle Orbital Heating Environment17.6.1 S-IVB Orbital Heating17.6.2 IU Orbital Heating

18 ENVIRONMENTAL CONTROL SYSTEM8.1 Summary18.2 $-I¢ Environmental Contro18.3 S-II Environmental Contro18.4 $-IVB Environmental contro?18.4.1 Ascent Powered Flight Phase18.4.2 Parking Orbit hase18.5 IU Environmental contro18.5.1 Thermal Conditioning System18.8.2 Gas Bearing Supply System

q9 SYSTEMS

SummaryVehicle Heasurements EvaluationS-IC Stage Measurement. Analysis

12.12.2 S-II Stage Measurement Analysis2.3 S-L¥8 Stage Measurement Analysis

19.2.4 Instrument Unit Measurenent Analysis19.3 Airborne Telerietry Systems19.3.1 S-IC Stage Telemetry System19.3.2 S-II Stage Telemetry System19.3.3 S-1¥B Stage Telemetry System19.3.4 Instrument Unit Telemetry System

Page

WN17-2217-2317-2317-3317-3617-3917-4217-4217-46

18-118-118-718-1418-1118-1218-1318-1518-21

19-119-119-219-419-619-619-619-1419-1619-1619-17

Section

20

2

22

23

TABLE OF CONTENTS (CONTINUED)

Airborne Tape RecordersS-IC Stage Recorder5-11 Stage RecordersS-1VB Stage RecorderInstrument Unit RecorderRF Systems EvaTuationTelemetry Systems 2F PropagationEvaluation

19.5.2 Tracking Systems RF PropagationEvaluation

19.5.3 Command Systems RF Evatuation19.5.4 Televiston Propagation Evaluation19.6 Optica? Instrumentation19.6.1 Onboard Cameras19.6.2 Ground Engineering Cameras

VEHICLE AERODYNAMIC CHARACTERISTICS20.1 Summary20.2 Yehicle Axtal Force Characteristic20,3 Vehicle Static Stability20.4 Fin Pressure Loading

MASS CHARACTERIST CS21.1 Sumary21.2 Mass Evaluation

MISSION OBJECTIVES ACCOMPLISHMENT

FAILURES, ANOMALIES AND DEVIATIONS23.1 Sumary23.2 System Failures and Anomalies23.3 System Deviations

x

Page19-1719-1719-1719-2019-2019-21

19-21

19-2419-2819-3119-3119-3119-32

20-120-120-420-4

21-1

21-1

22-1

23-122-723-1

Section

24

AppendixA

TABLE OF CONTENTS (CONTINUED)

SPACECRAFT SLMMARY24.1 Summary24.2 Spacecraft Performance Evaluation

ATMOSPHERE,Ad SummaryAz General Atmospheric Conditions at

Launch TimeAS Surface Observations at Launca TimeAa Upper Air MeasurementsAa] Wind SpeedA4.2 Wind DirectionAa3 Pitch Wind ComponentAad Yaw Hind ComponentALAS Component Wind ShearsAS Thermodynamic DataA517 Temperature45.2 PressureA.5.3 Density4.5.4 Optical Index of Refraction46 Comparison of Selected Atmospheric Data

for all Saturn Launches

AS-502 VEHICLE DESCRIPTIONB.1 Summary8.2 S-IC StageB21 S-EC Configuration8.3 S-IE Stage8.3.1 S-IL ConfigurationBa S-1¥8 Stage

Beat S-I¥B ConfigurationB.5 Iu

8.5.1 IU Canfiguration8.6 Spacecraft8.6.1 Spacecraft Configuration

Page

24-124-1

Figure

4-142

5-105-115-12

6-16-26-36-4

6-56-6

LIST OF ILLUSTRATIONS

Ascent Trajectory Position ComparisonAscent Trajectory Earth-Fixed VelocityComparisonAscent Trajectory Space-Fixed VelocityComparisonAscent Trajectory Acceleration ComparisonDyanmic Pressure and Mach Number VersusRange TimeAS-902 Acceleration Due to Yenting

AS-502 Ground TrackS-IC Start Box RequirementsS-IC Engine Buildup Transient:S-IC Steady State OperationS-IC Engine Shuzdown vransient Performance

S-IC Fuel Lilage Pressure Uuring Countdown

S-[C Fuel Ullage Pressure During BoostS-£C Holium Bottle Pressure for FuelPressuri zationS-1C LOX Ullage PressureS-IC Pneumatic Control Regu’ator dutletPressureS-IC Control Sphere PressureS+IC Camera Ejection SystenS-IC Predicted Camera Ejection and PurgeSphere PressureS-IT LH2 and LOX Recircu.ation Systems

S-I1 Thrust Chamber TemperatureS-II Engine Start Tank PerformanceSéIL Start Box RequirementsS-II LHe and LOX Recirculation Systen PerformanceS-LI Engine Theust Bui Tdup

xiii

Page

44

45

5-19

6-46-56-7

6-86-10

Figure

LIST GF ILLUSTRATIONS (CONTINUED)

S-II Steady State OperationS-II d-2 Engine CiagramS-II Injector after 28 Second Test at High ASYMixture RatioS-II Engine No. 2 Perfarmance ShiftS-I1 Engine No. 2 Cutoff Performance

2S-II Engine No. Gas Generator Performance atShutdownS-I Engine Ke. 3 Performance at ShutdownS-I1 Engine Shutdown TransientS-II Stage Thrust DecayS-I1 Stage PU Valve PositionS-II LOX Probe/Tank MismatchS-11 LHe Probe/Tank MismatchS-I1 Mass at Ignition and CutoffS-1 Fuel Tank Ullage PressureS-IL Fuel Pump Inlet ConditionsS-I1 LO¥ Tank Ullage PressureSeI LOX Pump Inlet ConditionsS-II Recirculation Valves Receiver and RegulatorOutlet Pressures$-I1 Camera Ejection PressuresS-IVB Start Box and Run Requirements - First BurnS-IVE Thrust Chamber Temperature - First BurnS-IVB Start Tank PerformanceS-IVB Buildup Transient - First BurnS-IVB Steady State Performance - First BurnS-IVB Shutdown Transient Performance - First BurnS-1VB CVS Performance - Coast PhaseS-IVB UTlage Conditions - Coast Phase

xiv

Page

6-136-16

6-166-176-18

6-206-216-236-236-246-276-276-286-296-316-226-33

6-346-367-37-57-6

7-87127-13

7-14

7-157-16

77

7-187-19

7-207-217-227-23

7-267-25

7-267-27

7-28

7-267-307-31

LIST GF ILLUSTRATIONS (CONTINUED)

S-IVB Engine Start Tank Temperature andPressure - Coast Phase§-1VB Start Box and Run Requirerents - RestartAttempt,S-1VB Thrust Chamber Chilldewn Performance - Restart AttemptS-IVB Start Tank Pressure - Restart AttemptS-IVB Ergine Contrcl belium Sphere Pressure -Restart Attempt.S-IVB Fuel Turbine Inlet lemperature - RestartAttemptS-IVB GG Chamber Pressure - Restart éttemptS-IVB Engine Environmental Charges Luring J-2Engine anoralyS-IVB Environmental Copcitiors During J-2Performance ShiftS-IVB Summary of Environment EffectsS-IYB Pumps Performance During Restart Attemp:§-1VB J-2 Engine &S1 Schematic

§-1¥B J-2 Engine Injector Schematic

S-IVB PU Mass Sensor Volumetric Nonlinearity

S-IVB Ignition and Cutoff Best Estimate Mass* -First BurnS-IVB PU Valve Positions - First Burn

S-IVB LOX Coarse and Fine Mass DataS-1VB Servo BridgeFailure Modes of S-I¥B PU ProbesS-I¥B LHy Ullage Pressure - First Burn endOroitS-1VB LH Ullage Pressure - Restart AttemptS-IVB Fue? Pump Iniet Conditions - First BurnS-IVB Fuel Pump Inlet Conditions - Restart Attempt

xv

7-187-19

7-20

7-217-22

7-377-387-407-41

Figure

7-32

7-337-347-357-367-377-387-397-40

7-417-427-437-487-45

8-1Ba2

8-3

LIST OF ILLUSTRATIONS (CONTINUED)

S-I¥B LOX Tank Ullage Pressure - First Burnand OrbitS-1VB LOX Tank U1lage Pressure - Restart AttemptS-1VB LOX Pump Inlet Conditions - First BurnS-I¥B LOX Pump Inlet Cond?tiuns - Restart AttemptS-IVB Cold Helium Supply DecayS-1¥B Cold Heliun Pressurization SchematicS-IVB Pneumatic Controi Performance - First BurnS-IVB Pneumatic Control Performance - Coast PhaseS-IVB Pneumatic Contro: Perfaruance - RestartAttempt.S-IVB APS Propellant Predicitons - Module Ho. 1S-IVB APS Propellant Predictions - Module No. 2S-1VB Helium Bottle Pressure - Nodule No. 1S-IVB Helium Bottle Pressure - Nadule No, ?S-1VB Chamber Pressure, APS Engine Ho. 2 -Module Ho. 2S-IC Hydraulic System PerformanceS-II Hydraulic System PerformanceS-II Engines Ko, 2 and 3 Hydraulic SystemTemperatureS-II Engine No, 2 Actuator ForcesS-II Engine No, 2 Actuator Commands and PositionsS-I1 Engines No, ? and 3 Hydraulic Reservoir LevelsS-IVB Hydraulic System Performance - First BurrS-IVB Hydraulic System Performance - Coast PkaseS-1V8 Hydraulc System Pressure During AttenptecRestart

S-1VB Hydraul‘e System Performance During AttemptedRestartd-2 Engine Hydraulic Component Locations

xvi

8-13

a48-15

Figure

9-29-3gee

9-5

LIST OF ILLUSTRATIONS (CONTINUED)

Longitudinal Structural Dynamic Response Dueto Thrust Buildup and Release

Slow Release Rod Loads During ReleaseLongitudirat Acceleration Time Kistory

Longitudinal Loads at Maximum Bencing Momentand Inboard Engine Cutoff

Longitudine] Acceleration During S-I¢ ThrustCutoff (0ECO)Longitudinal Loads Subsequent to $-II Engines ut

Lateral (Pitch) Structural Dynamic ResponseDuring Thrust Buildup and ReleaseHaxinun and $-1! Engines Out BendingMoment

First Longitudinal Modal Frequencies and Accel-erations During S-IC Powered Flight

Longitudinal Oscillation Trends, 110 to 140Secands

S-IC Maximum Individual Engine and ComposChamber Pressure Oscillations

F-1 Engine Chamber Pressure and StructuraAcceleration Response at S-IC Gimbal PlaneDuring Time of High Longitudinal Oscillations

congitudinal Loads at Time of MaximumLongitudinal OscillationFirst Longitudinal Mode Shaves During S-ICPowered Flight

Lateral Modal Frequencies and Acceleration:During S-IC Powered FlightPitch and Yaw Mode Shapes During 5-15 PoweredFlightS-IC Fin Vibration Response and Bending andTorsional Modal FrequenciesS-IC Vibration and Strain Measurements atFin, Fin and Fairing, F-1 Engine, and ThrustStructure

xvii

Page

9-3$49-5

9-6

97

9-8

9-9

9-20

9-21

9-22

9-23

F*gure

9-19

9-209-219-229-23

9-269-27

9-289-299-30

9-31

9-329-339-349A-1

9A-2

9A~39A-4SA-5.

9A-694-7

LIST QF ILLUSTRATIONS (CONTINUED)

S-IC Vibration and Strain Keasurerents atIntertank, LOX Tank, and Forward SkirtS-IC Stage Structure Vibration EnvelopeS-IC Stage Engine Vibration Envelope5-1 Stage Components Vibration EnvelopeS-TE Stage Structure Vibration Envelopes atForward Skirt Stringers and Interstage FramesSII Stage Structure Vibration Envelopes at.Aft Skirt Stringers and Interstage FrancsS-Ii Stage Structure Vibration Envelope atThrust Cone Longerons, Engine 1 Bear, andEngine 1 Gimbal PadS-II Stage Component Vibration EnvelopesS-1VB Acoustics, Vibration and Dynamic StrainMeasurementsS-IVB Stage Vibration EnvelopesS-IVB Forward Skirt Dynamic StrainPressure Differential Across S-1VBForward Skirt PanelsTine Slices of Dynamic Strain Output ShowingWave FormsS-IVB Forward Skirt RMS Strain AmplitudesInstrument Jnit Vibration EnvelopesAS-502 Versus AS-501 Inertial Gimbal VibrationsFirst Instrament Unit Measurements A‘fectedby the 133 Second DisturbanceLocation of Earliest Events That Were Notedat 133 SecondsComposite Electrical EffectsSublimator Inlet Water Differential PressureS-IVB Strain Gage LocationsAxial Strain, Aft Skirt Station 282Axial Strain, Forward Skirt Station 3145, SideGages

xviii

Page

9-249 259-269-27

9-39

9-31

9-429-439-449 46

9A-3

9A-498-119A-129R-1E

9A-17

9A-18

Figure

98-8SA-9

9A-10

9A-11W-1

10-219-3

10-4

10-5

10-610-7

10-810-910-10

W1W-2y-3Vee1-811-6711-81-911-10

LIST QF ILLUSTRATIONS (CONTINUED)

Axial Strain, Forward Skirt Station 3145, Top Gages

S-1VB Forward Skirt Strains Bar Chart ShowingStrain Changes at 133 Seconds Through ECO

"Quick Look” Assessment, ALOTS 70 MM File,133 to 135 Seconds

Simplified EDS Power and Auto-Abort Interface

Navigation, Guidance, and Control System BlockDiagramPlatform Gimbal ConfigurationVelocity from Postflight Trajectory MinusST-124H-3 Platform Measured VelocityPredicted and Measured Accelerations in ParkingOrbit

Composite Pitch Steering Angle (Not Rate Limited)Referenced to Local Horizontal

Actual and Nominal Velocities-to-be-Gained

Commanded Attitude Angles (Rate Limited) DuringBoost PhaseAttitude Timeline - Liftoff to Spacecraft Separation

Accelerometer Pickup Signals During Liftoff

Envelope of Maximum Deviations of the Gyro andAccelerometer Servo Amplifier Outputs

Engines, Actuators and Nozzle Arrangement

Surface Wind Speeds

Liftoff Vertical Motion and Scft Release Forces

Hotddown Post Clearances (Positicn I)

Liftoff Trajectories of Fin Tip A

S-IC Center Engine Trajectories and Plume Angles

Pitch Plane Dynamics During $-1C Burn

Yaw Plane Dynamics During S-IC Burn

Roil Plane Dynamics During $-IC Burn

Free Streari Angle-of Attack During S-1¢ Burn

Page

9A-19

9A-20

9-23

DA-24

10-310-8

10-8

10-12

10-1510-16

10-1810-1910-22

10-241-4V6W-71-9

W1-10WeW-12W13811-14

11-16

Figure

WiWe.

1-13

W-14W185

11-16

WeV-181-1911-2011-21

W1-22

11-2311-24

11-25

11-26W12711-28

11-2911-30

11-3111-3211-3311-34

LIST OF ILLUSTRATIONS (CONT:

VED}

Pitch and Yaw Plane Wind Velocity Durire S-IC BurnQscillograns of IU Control Sensors curing 5-ICTransientOscillograms of Ergire Fesition and S-IC SensorsDuring S-IC TransientPredominate Slosh Frequencies During S-IC BurnS-IC Propellant Slosh Amplitudes at the HallDuring S-IC BurnS-IT and $-IVB Propellant Slosh Anplitudes OuringS-IC BurnS-1C Engine Deflection Response to Propellant SloshPitch Plane Dynamics During S-II BuraYaw Plane Dynamies Quring S-I7 3urnRoll Plane Dynamics During S-IT BurnPitch and Yaw Steering Misalignrent CorrectionDuring S-IT BuenS-HI Engine No. 2 Pitch and Yaw Actuator Responseto the External Compression Load at 319 Secondsand to the Consequent Cormand at 320 SecondsS-T LHe and LOX Slosh Anp1‘tudes During S-11 BurnS-IVB Lig Slosh Amplitude at the Probe DuringS-I1 BurnPredomnant Slosh Frequencies During S-I1 BunS-IT Engine Deflection Response to Propellant SleshS-INB Pitch Actuator Excitattor During $-IC BurnPitch Plane Dynamics During S-IVE rirst BurnYaw Plane Oyramies During S-IVE First BurnRol] Plane Bynamics During S-IVB First BurnS-IVE LOX Slosh Frequency and Height at the ProbeS-IVB LHe Slosh Frequency and HeightPitch Contro? Dynamics Following S-IVB First BurnYaw Control Dynamics Following $-1VB First Buri

x

Page

W-17

11-18

W-19V2)

11-22

11-23I-2411-26N-2711-28

V.3r

Figure

11-3511-36

V1-37

11-38

11-39

11-40

11-41

VW-42

11-4311-44

12-1

12-2

12-3

12-4

12-5

12-6

12-7

12-8

12-912-10

LIST OF ILLUSTRATIONS (CONTINUED)

Roll Control Dynamics Following S-IVB First BurnPitch Control Dynamics During 20 Degree PitchDown Maneuver

Pitch Control Dynamics During Pitch and RollManeuversYaw Control Dynamics During Pitch and Roll ManeuversRoll Control Dynamics During Pitch and Rol]Maneuvers.S-1VB Pitch Attitude Errors and Rates DuringAttempted Second BurnS-IVB Yaw Attitude Errors and Rates DuringAttempted Second BurnS-IVB Rol] Attitude Errors and Rates DuringAttempted Second BurnPitch, Yaw, and Roll Dynamics at Loss of ContreS-IVB/IU Tumble Rate HistoryS-IC Retro Motor ThrustS-IC/S-II Relative VeTocity and Separation DistanceDuring First Plane SeparetionS-IC/S-IT Clearance Distance and LongitudinalAcceleration During First Plane SeparationS-IC Pitch and Yaw Angular Dynamics FollowingS-IC/S-HI SeparationS-11 Angular Dispersions During $-IC/S-II FirstPlane SeparationInterstage/S-II Relative Velocity and SeparationDistance During Second Plane SeparationS-II Angular Dispersions During Second PlaneSeparationLateral Clearance Distance and Interstage BodyRates During Second Plane SeparationS-11 Retro Motor ThrustS-IVB Ullage Motor Thrust

xxi

W-57

11-5911-60

12-5

12-6

12-8

12-9

12-10

12-11

12-12

12-132-1712-17

Figure

12111a-1212-132-14

12-1512-1613-1

13-2

3-3

13-4

13-513-613-7

13-813-9

13-10

13-11

13-13

1314

13-15

13-16

LIST OF ILLUSTRATIONS (CONTINUED)

S-I1/S-I¥B Relative Velocity and Separation DistanceS-II/S-I¥B Longitudinal Accelerat‘onsS-II Angular Dispersions During $-I1/S+IVB SeparationS-II and $-IVB Angular Dispersions During S-I1/5-IVBSeparationS-11/S-1¥B Separation ( earanceRelative Motion During Spacecraft SeparationS-IC Power Generation and Distribution SystemsBlock DiagramS-IC Stage Yoltage (on Bus 1010} and Current{at the Battery)

S-IC Stage Voltage (on Bus 120} and Current(at the Battery)S-IT Power Generation and Distribution System:Block DiagramS-I1 Stage Main DC Bus Voltage anc CurrentS-II Stage Instrumentation Bus Voltace ard CurrentS-II Stage Recirculation DC Bus Yoltage and CurrentS-II Stage Ignition DC VoltageS-IVB Power Generation and Listritution SystensBlock DiagramS-IVB Stage Forward Battery Na. 1 Voltage.Current, and TemperatureS-IVB Stage Forward Battery No. 2 Voltage,Currert, and Temperature

S-IVB Stage Aft Battery No. 1 Yoltage, Current,and TemperatureS-IVE Stage Aft Battery No. 2 Voltage, Current,and TemperatureInstrument Unit Power Generation and DistributionSystems Block DiagrarIU Battery 6D10 YoTtage, Current, and TenperatureIU Battery 6D20 Voltage, Current, and Temperature

xxtT

Page

12-1812-1912-20

12-2012-2212-22

13-2

13-3

13-5713-713-813-8

13-9

13-11

13-12

13-13

13-18

13-1613-1713-17

Figure

1713-1813-19

15-1158-216-116-216-316-416-516-616-716-816-916-1016-1116-1216-13

16-1415-1515-16

16-1716-1816-19

16-2016-21

LIST OF ILLUSTRATIONS (CONTINUED)

Page

IU Battery 6030 Yoltage, Current, and Temperature 33-78

iU Hattery 6D40 Voltage, Current, and Temperature 13-18

Currents and Voltages Associated with 133.3 SecondTransient 13-20EDS Functional Diagram 15-2Q-Ball aP Versus Flight Time 15-5S-IC Engine Fairing Compartment Pressure Oifferential 16-2S-IC Engine Fairing Pressure Loading 16-3S-IC Compartment Pressure Differential 16-6S-IC Compartment Pressure Loading 16-7S-IT Compartment Pressure Loading 16-8S-II Lg Sidewall Insulation Differential Pressure 16-10S-IVB Forward Compartment Differential Pressure 16-11S-1VB AFt Skirt and Interstage Differential Pressure 16-11S-IVB Aft Skirt and Interstage Pressure Loading 16-12S-IC Base Pressure Differentials 16-13S-IC Base Heat Shield Pressure Loading 16-14S-IT Heat Shield Aft face Pressures 16-16S-II Base Heat Shield Forward Tace and Thrust ConePressures 16-165-1 Base Heat Shield and Thrust Cone Pressures 16-17S-II ase Heat Shield Pressures 16-18Vehicle External Overall Sound Pressure Level atLiftoff 16-19Vehicle External Sound Pressure Spectral Densities 16-21

Yehicle External Overall Fluctuating Pressure Level 16-23

Vehicle External Fluctuating Pressure SpectralDensities 16-25

S-IC Internal Acoustic Environment 16-2:

S-I1 Internal Acoustics 16-28

xxiti

Figure

Figure

16-2216-237-117-217-317-417-5

17-6

W-778

17-917-10

WMV2

17-13

17-1417-15

17-16VW?17-18

17-1917-20

17-2117-2217-23

LIST OF ILLUSTRATIONS (CONTINUED)

S-1VB Forward Skirt Acoustic LevelsS-IVB Aft Skirt Acoustic LevelsS-IC Base Heat Shield Trermal EnvironmentF-1 Engiae Thermal EnvironmentS-IC Base of =in D Total Heating RateSIC deat Shield Therwel EnvironmentS-IC Thermal Environment Effect of M-31 Losson Heat Shield Temperatures at OECOThermal Environment, Temperature Under Insulationon Inboard Side of Engine Wo. 7S-IC/S-II Separation Thermal EnvironmentS-II Heat Shield Aft Face Heating Rates andSurface TemperatureS-I] Thrust Cone Total Heating RatesS-I! Thrust Cone Tota? and Radiation HeatingRatesS-II Heat Shield Forward Face TemperatureEngine Compartment Ges TemperatureS-I1 Heat Shield Aft Face and Thrust ConeSurface TemperaturesS-IC Aerodynamic HeatingS-1¢ Engine Fairing Aerodynamic HeatingS-IC Fin Aerodynamic HeatingS-IC Body Aerodynamic HeatingForward Location of Separated FlowS-IC Thermal EnvironmentS-IC Thermal Environments Aerodynamic HeatingIndicator (AHI)S-IC Thermal Environment Fin Skin “TemperaturesS-II LH2 Feedline Fatring Total Heating RatesS-IE Ullage Motor Fairing and Aft Skirt TotalHeating Rates

xxiv

Page

Page

16-2916-2917-317-37-417-5

VW?

LIST CF ILLUSTRATIONS {CONTINUED}

5-II Structural TenperaturesS-II Forward Skirt Skin and Insulation Temperature:5-1VB Aeroheating EnvironmentS-IVB Protuberance Aercheating Environrent$-IV8 LH2 Heating During BoostIU Inner Skin Temperatures for AscentS-1VB APS Fairing TemperatureS-IVB APS Propellant Control Module TemperatureIU Inner Skin Temperature in Earth Orbit$-IC Environnental Control Systers ForwardCompartment Canister Conditioning SystemS-IC Forward Skirt Canister TemperaturesS-IC Environnental Control Syste and CompartmentTemperaturesContainer Equiprent Mount TemperatureS-I] Engine Compartment Conditioning SystemTrarsducer LecationsLOX and L¥2 Chilldown Inverter TemperaturesIv Environmental Control Syste Schematic DiagramSchematic of Theral Control Using WaterCoolant ValveTemperature Control Parameter:Sublimator Water Inlet PerformanceIU AS-802 Sublimator Startup Characteristic:TCS Glig Supply Pressure and TemperatureComponent TemperaturesGES Pressure Regulation PerformanceGBS GNz Supply Pressure and Tenperature(T4 and GET Teleretry CoverageGBI Telemetry CoverageVEF Telemetry Caverage SummaryADWSA/GLOTRAC Coverage Sumrary

xxv

Page

17-3647-3717-3817-4017-4217-4317-4517-4617-47

18-318-4

18-518-10

18-1118-1318-14

18-1618-1718-1818-1918-21

18-2218-2318-2419-2219-2319-2519-29

Figure

19-520-1

20-2

20-320-4ai

21-2

21-3

A-T

A-2AdAada5

AG

AT

BelB-2B-3BedB-5B-6

LIST OF ILLUSTRATIONS (CONTINUED)

C-Band Radar Coverage SummaryAverage Base Differential PressureForebody Axial Force Coefficient, Besed on WindTunne’ TestsVehicle Total Aerodynamic Axial ForceS-I€ Fin Pressure DifferentialTotal Vehicle Nass, Center of Gravity, and MassNoment of Inertia During $-IC Stage Powered FlightTotal vehicle Mass, Center of Gravity, and MassMoment of Inertia During $-II Stage Powered PlightTotal Vehicle Mass, Center of Gravity, and MassMoment of Irertia Ouring $-iVB Stage Powered FlightAS~502 Launch Time Scalar Wind SpeedAS-502 Launch Tire Wind DirectionAS-502 Launch Time Pitch Wind Speed Component {lix)AS-502 Launch Time Yaw Wind Speed Component (bz)AS-502 Launch Time Pitch (Sx) and Yaw {$,) CommonentWind ShearsRelative Deviation of AS-502 Temperature and DensityFrom PAFB (63) Reference AtmosnhereRelative Deviation of Pressure and AdsoluteDeviation of the Index of Refraction from the PAFE(63) Reference Atmosphere, AS-502

Saturn ¥ Apollo Flight ConfigurationS-I Stage ConfigurattonS-II Stage ConfigurationS-IVB Stage ConfigurationInstrument tinit ConfigurationApollo Space Vehicle

xxv

Page

19-30

20-2

26-320-320-5

21-18

21-19

21-20AnsAGATA-8

A-13B-28-3B-6

B-10B13b-16

Table

2-12-22-32-43413-23-334

4-1

a24-34-44-54-6,

5-15425-3

5-4

5-55-65-7

6-26-3

ft

LIST OF TABLES

Time Base Sumrary

Significant Event Times SummarySequence of Switch Selector EventsUnscheduled Switch Selector Events

AS-502 MilestonesS-IC Stage Propellant Mass at Ignition CommandS-11 Stage Propellant Mass at S+IC Ignition CommandS-IVB Stage Propellant Mass at 5-IC Ignition CommandGround/Vehicle Interface Events

Summary of AS-502 Orbital C-Band Tracking StationsS-11 Engine Ho. 2 Premature Cutoff ConditionsComparison of Significant Trajectory Everts

Comparison of Cutoff EventsComparison of Separation Events

Stage Impact Location

Parking Orbit Insertion Conditions

S-IC Stage Engine Startup Event TimesS-IC Engine Performance Deviations

S+IC Velocity and Time Deviation Analysis at OECO(Simulation Versus Predicted)Comparison of S-IC Stage Flight Reconstruction Datawith Trajectory Simulation ResultsSIC Cutoff ImpulseS-IC Stage Propellant Nass tlistoryS-IC Residuals at Guthoard Engine Cutoff$-I1 Engine Start Sequence EventsS-II Engine Performance Deviations (ESC +60 SecordsS-11 Flight Reconstruction Comparisor. withimulation Trajectory Match Results

S-IE Engine No. 2 Performance Shift anc Cutoff

xxvit

Page

2-12-3

2-19

3-2

3-7

3-73-83-94-3

4-9

41aan4-12

4-144-14545-7

5-8

Table

6-56-66-77-17-2

7-3

747-57-6cd7-8797-10a1

9-29-39-49A-T9A-210-110-210-310-4Wel1-2

ne

LIST OF TABLES (CONTIRJEO)

S-IT Cutoff TnsulseS-II Stage 2ropellant Mass IlistoryS-II delium Mass5-1V3 Engine Start Sequence Events - First BurnS-IVB Steady State Performance + First Burn(ESC +60 Second Time Slice)Comparison of S-IVB Stage Flight ReconstructionData - First BurnS-EVB Simulation Burn Tire DeviationsS-IYB Cutof* Impulse - First BuriS-IVB Engine Start Sequence - Second Burn AttemptS-IVB Stage J-2 Engine Failure to RestertS-IVB Stage Prope’ lant Nass HistoryS-IVB Pneumatic Heliur Dottie Mass$-IVB APS Propellant ConsumptionS-INB Hydraulic System PressuresMaximum Modal Accelerations at 5.5 Hertz for110 to 140 Seconds Range TimeS-IC Stage Vibretfon SummaryS-Ii Vibrations, Vehicle StructureS-IVB Vibration SummaryList of IU Measurements with 133 Second Resnonse133 Second Transient Survey, IU StageInertial Platfcrm Velocity ComparisonsGuidarce Comparisons (Navigation System)Rate Limited Steering Command TimesStart and Stop Guidance CommandsSunrary of Liftoff ClearancesAS-502 Misalignnent SummaryMaximum Control Parameters During S-IC Burn

xxviii

Page

6-226-266-35a7

Table

W41-6V€W7Ve-112-212-312413-1

13-213-313-44-115-115-215-316-1

Vel17-218-1

18-2

18-3

18-4

18-5

IST GF TABLES (COMTTNLEB)

S-1C Eynamie End Condi tions*Waximur Control Parameters During S-I1 BuriVaxirur Control Parameters During $-IVB BurnAPS Irpulse RequirementsSeparation Event TimesSeIC Retro Motor PerformanceS-11 Retro Motor Performance

S-IVB Ul lage Motor PerformanceS-IC Stage Electrical System Battery PerformanceDuring FlightS-HI Stage Battery ConsumptionS-IYB Stage Battery ConsumptionInstrument Unit Battery ConsumptionCCS Command History, AS~502Performance Sureary of Thrust OX Pressure SwitcheDiscrete EDS EventsSwitch Selector EDS EventsS-II Acoustic Noise Levels Comparison of AS-501and AS-502 DataRetro Motor Plure HeatingAPS Orbital TemperaturesS-IC Environirental Control System CanisterTerperaturesS-IC Environmental Control System Aft CompartitentTemperatures

S-II Forward Thermal Control System AS-801 andAS-502 Prelaunch and Flight DataS-I1 Aft Thermal Control System AS-SO1 and AS-502Prelaunch and Flight Data$-II Engine Compartment Tenporature DataComparison of AS~501 and AS-802 Flights

xix

Page

11-2011-2911-3911-4312-212-412-1412-16

13-413-6

13-1513-1914-315-415-6

16-2817-2217-44

18-2

13-4

19-8

18-9

18-10

Table

18-618-719-119-219-319-419-519-619-7

19-82-121-221-321-4

21-5

21-921-1022-123-1

23-2Al

A-2A3Aa

LIST OF TABLES (CONTINUED)

Forward and Aft Skirt Component Temperature*IU TCS Flowrates and Coolant Pressure VataVehicle Measurements SummaryMeasurements Waived/Serubbed Prior to LaunchMeasurement Malfunctions During 71ightMeasurements with Imroner RangeQuestionable HeasurementsLaunch Vehicle Telemetry LinksTaze Recorder SummaryAS-502 Onboard Tracking SystemsTotal Vehicle Mass, S-IC Burn Phase, KilograniTotal Yehicle Mass, S-IC Burn Phase, Pounds MassTotal Vehicle Mass, S-II Burn Phase, K*1ogramsTotal Yehicle Mass, S+I1 Burn Phase, Pounds NassTotal Vehicle Mass, S-I¥B First Burn Phase,KiTogransTotal Vehicle Mass, S-IVB First Burn Phese,Pounds MassTotal Vehicie Mass, S-IVB Second Burn Phase,Ki7ogramsTotal Vehicle Mass, $-IVB Second Burn Phase,Pounds MassFlight Sequence Mess SummaryMass Characteristics ComparisenMission Objectives Accoup]ishment SummarySummary of Failures and AnomaliesSummary of DeviationsSurface Observations at AS-602 Launch TimeSystems Used to Measure Upper Air Wind Data, AS-502Moximum Wind Speed in High Bynamic Pressure RegionExtreme Win Shear ir High Dynaric Pressure Region

Xxx

Page

18-1218-2019-219-319-7

19-1119-1319-16“9-8

19-2721-321-421-621-6

21-102-1121-1422-223-223-5AZAWBAg

A-11

Table

AS

B-18-28-3B-4

LIST OF TABLES (CONTINUED)

Selected Atrospheric Observations for SixteenSaturn Vehicle Launches at Kennedy Space Center,Florida

S-IC Significant Configuration ChangesS-IL Significant Configuration Changes

S-IVB Significant Configuration ChangesIU Significant Configuration Changes

xx

Page

A-148-5B-9

B-128-14

ACKNOWLEDGEMENT

chis report is publisned by the Seturn Flicht Evaluation Horking Group--cofposed of representatives of Marsiall Space Flight Center, John F. KennedySpace Center, and MSFC's prime contracters--and in cooperation with theManned Spacecraft Center. Significant contributions to the evaluationhave been made by:

George C, Marshall Space Flight CenterResearch and Levelopnent Operations

Aero~Astrodynamics LaboratoryAstrionics LaboratoryComputation LaboratoryPropulsion and Vehicle Engineering Laboratory

Industrial Operations

John F. Kennedy Space CenterManned Spacecraft Center

The 3oeing Company

Douglas Aircraft Company

International Business Machines CorporationNorth American Rockwell/Rocketdyne DivisionNorth American RockweTl/Space Division

xaxtt

AEDC

ANT

APS

Asc

AST

BDA

AL

ccs.

coor

G

CIF

ABBREVIATIONS

Arnold EngineeringDeve lopment Center

Anzigua

Auxiliary propulsion system

Ascension Island

Augmented spark igniter

Bermuda

Goldstone, California

Command and communicationssystem

Countdown demonstration test

Center of grav‘ty

Central instrumentationfacility

Command module

CutorF

Carnarvon

Command and servicemodule

Cape telemetry 4

Continuous vent system

Grand Canary Island

Digital events evaluator

wot fi

EBu

ECO

ECP

ECs

Eps

EMR

Esc

Fett

FM/ FI

GBI

BFCY

6G

GLOTRAC

GRR

GSE.

GSFC

6TL

ey

HAW

HREY

Exploding bridge wire

Engire cutoff

Engineering changeproposat

Environment control system

Emergency detection system

Engine mixture ratio

Engine start command

Flight combustion monitor

Frequency modu)ation/frequency modulation

Grand Bahama Island

GOX flow control valve

Gas generator

Global tracking

Guidance reference release

Ground support equipment

Goddard Space FlightCenter

Grand Turk Island

Guaynas

Hawaii

Helium flow control valve

TECO

16M

TRIG

Iu

Les

LET

Let

LTA

Lut

w

LDA

LypC

MAB

MCC-H

MEY

MILA

ML

Muy

Mov

HR

MSC

MSFC

NPSP

Inboard engine cutoff

Iterative guidance mode

ter range instrumentationgroup

Instrument Unit

Launch escape system

Launch escape tower

Lunar module

Lunar module test article

Launch umbitical tower

Launch vehicle

Launch vehicle data adapter

Launch vehicle digitalcomputer

Materials analysis branch

Mission contra? center-Houston

Main fuel valve

Merritt Island Launch Area

Mobile Tauncher

Main LOX valve

Main oxidizer valve

Mixture ratio

Manned Spacecraft Center

Marshall Space FlightCenter

Net positive suctionpressure

Xxxiv

OAT

opoP

ECO

PAH/FM/ FM

Por

PCM/ FM

Psp

PTCS.

Pu

PUS

RACS,

RED

RET

RMR.

RMS,

RP-1

RPM

Overall test.

Offset frequency copper

Outboard engine cutoff

Pulse amplitudemodutation/frequencynodulatton/ frequencymodulation

Pulse code modulation

Pulse code modulation/Frequency modulation

Programmed mixture ratio

Programmed mixture ratioshift

Power spectral density

Propellant tankingcomputer system

Propellant utilization

Potable water system

Remote analog calibrationsystem

Redstone (ship)

Transition Reynoldsnumber

Radio frequency

Ground hydraulic fluid(kerosene)

Reference mixture ratio

Root mean square

S-IC stage fuel (kerosene)Revolutions per minute

RPS.

R/S

RSS.

SA

SAR

se

SLA

SM

suc

38

SS/FM

sTes¥

sTDY

sv

TAN

TEL4

THR

TSM

Te

Revolutions per second YAB

Range safetyveo

Root sun square

Swing arm VSWR

Start arm retract

Spacecraft

Spacecraft LM adapter

Service module

Steering misalignmentcorrection

Service propulsion system

Secure range safety commandsystem

Switch selector

Single sideband/ frequencymodulation

Start tank controlsolengid valve

Start tank discharge valve

Space vehicle

Tananarive

Cape telemetry 4

Triple modular redundant

Tail service mast

Torust vector contro?

RXY

Vehicle assembly buildingat KSC

Voltage controlledoscillators

Voltage standing waveratio

MISSION PLAY

45-802 (Apolto 6 Mission) was the second flight vehicle of the Saturn ¥Apollo flight test program. The basic purpose of che flight was to demonstrate the compatibility and performance of the launch vehicle and theApollo Command end Service Modules (CSH) for a future manned flight{he AS-502 vehicle was launched fron Launch Complex 39, Pad & at KennedySpace Center (KSC) on a launch azimuth of 90 degrees, then rolled to aFlight azimuth of 72 degrees. The S-I¢ and S-I1 stage powered flightEines were to be approximately 147 and 368 secords, respectively. FirstS-.VB stage powered Flight time was to be approximately 138 seconds ter.minating with insertion into a 185.2 kilometer (100 n mi} circularbarking orbit of the S-IVB ard spacecraft. Vehicle weight at parkingorbit insertion was to te approximately 127,000 kitograms (280,000 lem)Aeproximately, 18 seconds after first buen cutoff, the S-I¥B was to alignwith the Toca? horizontal, then the vehicte was to roli 180 degrees toobtain a position 111 down configuration, tn this configuration thewenicle would be subjected to a 20 degree pitch-down anda 2) degreepitch-up maneuver, then roll 180 degrees again to obtain the origiaaposition I down configuration. These maneuvers were to produce informatioon the SoI¥B restart bottle repressurization and propellant slosh excl caetion while qualifying these maneuyers for manned Flight. On menned Flightthese maneuvers may be used to orient tne astronauts for landnark tracking

Chilidown and reignition sequencing were to begin between Hawaii andCalifornia during the second revolution and continue across the continentalUnited States, After the restart preparation was initiated, an orientationmaneuver was to be performed to yield the high apogee elliptical orbitafter second burn. The $-1¥8 was to be reign'ted over KSC, near the endOf the second revolution for translunar injection boost, and aimed atPaget simulating che moon. This was in order to preclude hitting themoon while verifying launch vehicle guidance techniqueThe S-IVB second burn, which was to be approximately 316 seconds, was toterminate with the injection of the S-IVB/IU/CSM into an elliptical orbitwith an apogee radius of approximately 528,024 kiloneters (285.110 nai).Foliowing S-1¥B second cutoff, the vehicle was to coast on a simulatedlunar trajectory for about 3 minutes before separation. A pitch rotetionof approximately 188 degrees was to be accomplished during this coast toposition the CSM for retrograde burn. After this rotation, the SpacecraftLunar Module Adapter (SLA) panels were to open to free the CSM. FollowingCSM separation, the S-IVB was to be oriented to give satisfactory attitude

xvi

for ground communications. Four minutes and 40 seconds after $-I¥B secondcutoff, the Service Module Service Propulsion System (SPS) was to beg!a 254-second retrograde burn to retard the CSM onto an earth intersectingellipse naving an apogee altitude of approximately 22,204 kilometer:(11,989 1 mf)

About two hours after separation, the S-IVB was to be aligned by groundcormand with the Tocal horizontal to test the ground comand capabilitynear the limit of S-1¥B stage active lifetime. The S-IVB was to reenterthe atmosphere over the Pacific Ocean on the return leg of the high apogeeellipse

After SPS first cutoff, the CSM was to coast fn the 22,204 kilometer(11,989 1 mi) apogee orbit for about 6 hours, oriented to cold soak theheat shield, approximating lunar return thermal conditions. During de-scent portion of the orbit, the second SPS burn was to accelerate theCSM to the approximate lunar return velocity of 11,125 m/s (36,500 ft/s)with an inertial flight path entry angle of -6.$ degrees. The servicemodule was to be jettisoned before reentry. The command modute was toenter the atmosphere approximately 4 minutes after SPS cutoff and splashdown near Hawai? at approximately 9 hours and 50 minutes after liftoff.

xxxvii

FLIGHT TEST SUMMARY

The second Saturn ¥ Apollo space vehicle, AS-502 (Apollo 6 Mission’, walaunched at Kennedy Space Center (KSC), Florida on April 4, 1968 at07:00:01 Eastern Standard Time (ES7} from Launch Complex 39, Pad A. Thiswas the second mission to use a Saturn ¥ launch vehicle with an un-manned block 1 Command Service Modute (CSM020), and a Lunar Module TestArticle (LTA-2R). Nine of the sixteen primary objectives were completelyaccomplished, six partially accomplished, and one (S-IVB restert) wasnat accomplished. One of the two secondary objectives was completelyaccomplished, and one partially accomplished

The launch countdown was completed without any unscheduled countdown Folds.Ground systems performance was highly satisfactory, The relatively fewproblems encountered in countdown were overcome suck that vehicle Taurcrreadiness was not compromised.

The vehicle was launched on an azimuth of 90 degrees east of north andafter 11.1 seconds of vertical flight, (which incluced a small yaumaneuver for tower clearance) AS-502 becan to roll into a flight azi-muth of 72 degrees east of north. Actual trajectory parameters af theAS-502 were close to nominal until the premature shutdcun of two engineson the S-I1 stage. After this premature skutdoun, the trajectory deviatedsignificantly from the nominal throughout the remainder of the missionSpace-fixed velocity et S-IC Outbearc Ercire Cutoff (CECO) was 7.28 m/:(23.89 ft/s) greater than nominal. At S-II Engine Cutoff (ECO) thespace-fixed velocity was 102.36 m/s (336.82 ft/s) less than nominal andthe altitude was 6.39 kilometers (2.45 n mi) higher than nominal, AtS-IVB velocity cutoff conmard the space-fixed velocity was 48.94 m/s(160.8€ ft/s) oreater than nominal. The altitude at S-IVB velocity cut-off comard was 6.7% kilometers (C.42 n mi) Tower than nominal and thesurface range was 448.46 kilometers (269.15 n i) Yonger than nominal.Parking orbit insertion conditions deviated considerably from nominabecause of anomalies that occurred during the powered portion of flightThe space-fixed velocity at insertion was 48.16 m/s (158.00 ft/sgreater than nowinal and the flight path angle (elevation of space-fixedvelocity vector from local horizontal) was 0.378 degree less than nominal.These conditions produced an orbit which was quite elliptical with aneccentricity 0.0138 greater than nominal. The resulting apogee of theparking orbit was 171.54 kilometers (92.63 n mi) higher than nominal,and the perigee was 12.17 kilometers (6.57 nmi) Tess than nominal. ‘TheS-IVB stage failed to reignite. Shortly after the attempted reignition,the spacecraft separated from the launch vehicte on ground command to thespacecraft. The S-IVB stage reentered due to orbital decay on April 25, 1968

wAKVETT

S-1C propulsion systems performed satisfactori’y. In general, atl per-formance fight data fell close to the nominat predictions. Average enginethrust reduced to standard sea level conditions from 35 to 38 secondswas 0.20 percent lover than predicted and average specific impulse was0.10 percent lower than predicted. The veh*cle first long‘ tudinal struc-tural mode frequency coupled with the engine response to the oxidizersuction lines resonant frequency within the 110 to 140 second periodThis resulted ina vehicle longitudinal oscillation termed "POGO". InboardEngine Cutoff {IECO} and Outboard Engines Cutoff {0ECO) occurred 0.11 and0.85 seconds later than predicted, respectively. An intentional fuellevel cutoff of the outboard engines and a LOX level cutoff of the in-board engine were planned and attained, demonstrating the adequacy ofthese cutoff modes. All the subsystems except the camera ejection systemand the contral pressure system performed as expected. The camera ejectionsystem ejected only one of the four film cameras and the contro! pressuresystem sphere pressure decayed unexpectedly after separation

The $-II propulsion systen performed satisfactorily during the first169 seconds af operation following Engine Start Command {ESC}. Enginethrust, at 60 seconds after ESC, was only 0.43 percent below predici tonand specific impulse 0.08 percent above predictions. At 319 seconds asudden performance shift was exhibited on engine No. 2 with thrust de-creasing approximately 33,806 Newtons (7600 Ibf). The engine continuedperformance at the reduced level until 412.3 seconds. By 412,92 second:the dropout of thrust OK switches indicated engine No. 2 cutoff, and at414.18 seconds engine to. 3 also cut off, Postflight evaluation oftelemetered data led to the conclusion that the engine No. 2 AugmentedSpark Igniter (ASI} fuel line failed and ultimately caused failure ofthe engine, Since the flight, testing at Marsail Space “light Center(HSFC) and the engine manufacturer's facility has substantiated thisconclusion. The testing reveals that an oxidizer rich mixture, causedby 2 fuel leak, creates very high temperatures and rapidly erodes trejnjector. Because of this erosion the LOX dome of engine No. 2 eventuallyfailed, opening the LOX high pressure system and causing Sngine Cutoff(ECO).” A modification of the ASE propellant feedlines (both fuel andLOK) and their installation is seing accomplished. Interchanged LOXprevalve control wiring connections between angines No. 2 and 3 solenoidscaused the preanature cutoff of engine No. 3. When engine No. 2 cutoff,the 0X prevalve on engine No. 3 was commanded closed. Individual sheck~out of the srevalve wiring during orevalve timing checks is planned forFuture vehicles. S-IT burn time was 425.31 seconds which is 37.81 sec:onds longer tnan predicted. The extended burn time was caused by thepremature cutoff of engines Yo. 2 and 3. Loss of the twa engines re-duced prozellant consimption approximately 40 perceat and required alonger ourn time to reach propellant-depletion. The S-11 sroaulsiosubsystems met al] performance requirements

Tne S-1VB J-2 engine operated satisfactorily throughout the operationalpnase of first burn. However, a total performance shift of 2.3 percentdecrease in thrust occurred during first burn from 684 to 702 seconds.

XXXIX

The engine continued to operate at tae shifted performance level anéhad a normal shutdown. S-IVB first ourn time was 166.52 seconds whichwas 28.95 seconds longer than aredicted due to the two engines out con-dition on the S-II stage. Stage performance during first burn deviatedfron the predicted at the 60 second time slice by -0.08 percent for thrustand 0.96 sercent for specific impulse. The LOX mass measuring side ofthe Propellant Utilization (PU) system malfunctioned during orbit priorto the attempted restart. The LOX mass measuring system nalfunctiocaused a 100 percent LOX Toad indication prior to and during the restartattempt. The probable cause of the erroneous 100 percent LOX mass indication may have been due to shorting of the inner and outer element:of the LOX Pil probe from metallic debris that could have been in theLOX tank. Also, an intermittent short in the cable shield between thmass probe and the PU electronics assembly may have occurred. Enginerestart conditions were within limits even though main chamber secondignition failed to occur. Results thus far indicate that a leak in theAugmented Spark Igniter (ASI) fuel supply system probably occurred duringfirst burn. Additional engine tests have essentially verified the per-formance shi“t and the restart failure. A modification of the AST pro-pellanz feedlines (both fuel and LOX) and their installation is beingaccomplished. A11 subsystens operationally met all performance an:stage requirements. However, there were two unexpected deviations whickare discussed in Sect‘on 7,

In general, the hydraulic systems performed satisfactorily ir that thevehicle remained stable during all portions of guidarce-centretled peeredflight. No hydraulic system probleme occurrec during §-IC pewered flightS-IT hydraulic systems performed within predicted limits, and operatedsatisfactorily until 280 seconds. At tris time, the S-Ii engine No. 2yaw actuator delta pressure transducer tegan to deviate significantlyfrom expected values. From 319 secords until engine No. 2 cutoff, botthe pitch and yew actuators showed apparent side loads from the engineAfter engine No. 2 cutoff, the yav actuator performance indicates that¥t locked up. The engine No. 3 system performed normally unti? engineShutdown when the system pump stopped operation and the pressures decayed,The engine No. 1 ard engine No. 4 hydraulic systems performed normallythrougrout S-I1 powered flight. The S-IVB hydraulic system performedwithin predictec limits during liftoff, boost, and first burn, Duringengine restart preparation and restart attempt, the system failed toBreduce hydraulic pressure. System temperatures observed during S-IVBfirst turn indicated the existance of a cryogenic fuel leak which ledto the freezing of the hydrauTic fluid and system blockage. During therestart attewpt, measurements indicated that both the main and the auxil-jary hydraulic pumps cavitated during operation and virtually no systempressure was produced,

The AS-502 flight vehicle experienced considerably wore structural activitythan AS-501; however, this activity did not reach sufficient magnitudeto pose a threat to the launch vehicle structual integrity. Thrustbuildup and vehicle release transients, resulted in maximum longitudinaland lateral {pitch plane) dynamic load factors of £0.4 and 40.03 g,

x]

respectively, at the covmand module. The maximum steady-state bendingmoment condition, 9.89 x 105 N-n (7.29 x 10° Ibf-ft), was exserizncedat 66.6 seconds. The ~aximum longitudinal loads were exoerienced at144.72 seconds (1ECO) at a rigid body acceleration of 4.3 3. Altnoughthe 4.8 g IECO condition was greater than the 1.63 g design value, nomainline structural problens were encountered during this phase of FlightThrust oscillation-structural dynanic response coupling (P80) was evidentduring the 110 to 140 second region of S-IC range tine. The longitudinadynamics of the launch vehicle induced lateral accelerations of 5.65 Gpeakat the LTA. Oscillations in the first longitudinal rode during the 119to 140 second time period exceeded that experienced during 5-501 flightby approximately a factor of three. Maximum response occurred in the5.2 to 5.5 hertz bandwidth. Fin bending and torsional nodes comparedwell with analytical predictions. Fin vibrations exceeded the range ofthe accelerometers but the modal frequencies did not coalesce and flutterdid not occur. S-1C, S-1VB and Instrument (I!) vibrations were as ex-pected. S-I] stage vibrations were as expected, except that forwardskirt vibrations exceeded the sine and randor criteria at liftoff. Yoadverse effects were noted, S-IVB forward skirt experienced Timi tedamplitude panel flutter, The stress amplitudes encountered due to flutterwere about three times higher than those of AS-204 but were stil] withina tolerable level.

At approximately 123 seconds abrupt changes of strain, vibration, andacceleration measurements were indicated in the S-IVB, IU, Spacecraft/Lunar Module Adapter (SLA), LTA, and Command and Service Modute (CSM).Photographic coverage, Airborne Light Optical Tracking System (ALOTS)and ground camera film showed pieces separating from the area of theSLA. There were no known structural failures noted on the launch vehicle.

The perforvance of the guidance and navigation system was as predictedfrom liftoff to 412.9 seconds. When engine No. 2 cut off a discretesignal was recognized by the [¥ indicating a single engine failure. How-ever, since only single engine failures were considered in constructingthe flight program, no program action was taken for engine Yo. 3 failure.with the discrete signal and loss in acceleration the program entereda guidance mode where guidance and navigation computations were madebased cn a thrust change for the single engine failure which was 50 per-cent of the total actual change. This node (artificial tau) lasted untithe It sensed an acceleration change due to S-II Programmed Mixture Ratfo(PHR) shift. Guidance computations responded to variations in altitudeand velocity caused by the decrease in thrust during the S-IT burn periodThe cortrol system responded well to the guidance comrands for the re-mainder of tre beost period. [ue to the two-engine-out perturbation,flight path angle and velocity were not optimun at the time guidancecommanded S-IV ECC, This resulted in an overspeed of 48.9 m/s (160 ft/s)A parking orbit whicr was acceptable though off nominal was achievedAM ortital quiaance maneuvers were satisfactorily performed. IU commandswere properly executed for S-IVE restart, tut the engine did not reignite.

xT

When acceleration test conditons were not met, Time Base 7 (T7) was ini-tiated and a cutoff command was issued to the S-IVB stage.

The AS-502 Flight Control Computer (FCC), Thrust Yector Contro? (T¥C),and Auxiliary Propulsion System (APS} satisfied all requirements for at-titude contro} and stability of bending and propellant sTosh modes inboth the boost and orzital coast modes of operation. During 1i*to“f,all vehicte clearance requirements were met. During first stage boostthe control system was required to correct for a steady-state roll at-titude error of aporoximately -0,5 degree. This roll coraue was not ob-served on AS-501, as the attitude error was essentially null after about50 seconds. Control system performance was consistent with events wh'caccurred during S-IT boost. The performance shift of engine No. 2 at319 seconds was evidenced in the TYC as well as in che FCC parametersHowever, this derformance shift caused no control problens and resultedonly in a new steady-state trim condition, The FCC and TYC respondedsatisfactorily to the perturbations caused by the shutdown of enginesNo. 2 and 3. “This shutdown resulted primarity im a large pitch planedisturbance during which the pitch rate built up to a viaximum of 2.8 deg/:{nose-up) and the pitch attitude error reached a maxirum of 13.4 degreesA maximum engine deflection of 5.95 degrees was required to stabilize theattitude excursions. At S-[1/S-I¥B separacion, the gu*dance computerswitched to the S-IVB coast mode for 9.3 second. The 7.4 degrees pitchattitude error caused a full-on APS pitch-engine firing of 9.3 secondduration to correct the attitude, At 0.3 second after separation, theguidance computer switched to the $-IVB burn mode, The pitch attituderror was trimmed out by the TYC after S-1VB stage J-2 engine ignitionControl system performance was nominal for the remainder of $-IV6 firstburn. Orbital attitude controt requirenents required considerably roreAPS activity than anticipated. The APS systen was required to overcome2 80 degrees nose-up “rom Tocal horizontal attitude and a | deg/snose-up angular rate to align the vehicle along the local horizontal.The vehicle as subsequently exercised through a sequence of fourmaneuvers as follows: 180 degrees roll, 20 degrees pitch down, 20 de-grees pitch up, and 180 degrees roll. che pitch and roll maneuver:were planned to produce information on the 5-IVB restart bottle repres-surization and propellant slosh excitation while qualifying these naneu-vers for manned flight. Each naneuver was executed as planned, Noappreciable effect was noted on the restart bottle conditions. LH» sloshingwas not appreciable ducing any of the maneuvers. Significant LOX sloshingexisted at the initiation of each pitch maneuvers however, tre initiaamplitude was not sustained due to hgh damping. An auxiliary Fydravlicpump failure prevented the S-IVG stage J-2 engine from being certered atthe time of second S-IVB CSC. The engine position at ESC was approximately1.5 degrees in pitch and -2.3 degrees in yaw. Appreciable attitude errorsresulted from this engine position during restart attempt; however, vehiclecontrol was maintained by the APS system following the switch from thrustvector to coast made contro’. Subsequent to spacecraft separation theAPS system maintained contro] until APS module I fuel cepletion at approxi-mately 21,953 seconds. Veh‘cle attitude rates began to builc up signifi-cantly following module ITI fuel depletion (22,602 seconds} and continued

xii

to increase as indicated by reduced radar data until a tumble rate of180 deg/s was recorded by the ninth day following launch.

Launch vehicle separation systems performed satisfactorily, Separationand associated sequencing occurred as planned with adequate clearance be-tween stages. Al) ullage and retro motor performance was satisfactoryThe S-IC retro rotor data indicate that some parameters were either abovenormal or possibly above the maximum Vimits but caused no problem. Space-craft separation was initiated by ground comand to the spacecraft duringthe maneuver to separation attitude. Even though there was a possiblerorentary interference between a SLA panel and the CSM at the separationplane, the momentary interference was not detriirental to the separation

In general, launch vehicle electrical systers performed satisfactorily.Battery valtages and currents were satisfactory and battery temperatureremained within acceptable liits. $-IC battery No. 2 experienced asharp current rise and voltage drop after S-IC/5-!] separation whichlasted fer 11 seconds; however, tape recorder performance was not im-paired. A similar anomaly was experienced by Sattery ‘Yo. 1 on the AS-50flight. Disturbances were experienced on the S-It main and instrurentattion batteries during the engine No. 2 and 3 shutdown pertod. A currentsurge was experienced on the IU 6010 battery at the tire of the 133 sec~ond transient.

Tata indicated that the redundant Secure Range Safety Comrand Systems(SRSCS) cn the S-IC, S11, and $-1VB stages were ready to perform theirfunctions properly on cenmand if flight conditions during the launchphase had required vehicle destruct. The system properly safed the $-IVESRSCS an command from KSC, The performance of the corrand and comruni-cations system in the IU was very good.

The space vehicle Emergency Cetection System (EES) was flight tested inthe aLtomatic abort closed-Icep configuration on AS-502. Launch vehiclemeasurenents irdicated that nc CLS limits ware exceeded and the systemfurctioned properly. There were some anomalies indicated in the spacecraft.

The vehicle interral, external and base region pressure environments weregererally ir geod agreement with the predictions and corpared well withthe AS-501 cata. The pressure enviranment was well below the designlevel. The measured acoustic levels were also generally in good agree-rent with the precictions and with AS-5C] data

The vehicle thermal envirorment was generally less severe than that forwhich the vehicle was designed. One exception was the S-IC forward skirtthermal environment whick exceeded design after S-IC/S-IT separation. Lossof H-31 to the level of the crushed core en the S-IC base heat shield wasvisually observed on this flight via the television careras which viewedthe heat shielc. This was a repeat of an AS-5CI anomaly and ne adverseeffects were noted, The effect of the premature shutdown of engines No. 2

xlidi

and 3 on the S-IT heat shield and base region environment was minor, With -the exception of abrupt spikes due to the engine anomalies, the base re-gion thermal data compared favorabiy with AS-501 data

The S-IC canister conditioning system and the aft environmental candi tion-ing system performed satisfactorily during the AS-502 countdown with onlyone canister and one ambient tenperature measurement dropping below theminimum requirement. The S-!£ therral contral and compartment condition-ing systen maintained temperatures within the design limits throughoutthe prelaunch operations. Temperatures monitored on the S-IVB aft skirtcomponents were slightly cooler than on the AS-501 flight but within de~sign limits. Temperatures of a°i components mounted on the forward skirtcold plates were within design “imits at liftoff. The IU EnvirormertaControl System (ECS) performed well throughout the flight. Coolanttemperatures, pressures, and flowrates remained within the predictecranges and design limits for the duration of the flight data. One speci-fication deviation was observed which was expected. At 11,670 secords,the platform gas bearing pressure differential was 9.069 N/emé (0.1 psia}above the 10.7 N/em? {15.5 psid} maximum allowable and remainec therethroughout the remainder of the flight period for which data is available{33,780 seconds).

There were 2758 telenetered measurements active at the start of the AS-502automatic countdown scquence. OF the 2758 measurements, 66 failed 4flight, resulting in an overali system reliability of 97.9 percert. TheAirborne Telenetry Systen operated satisfactorily, includirc preftichtcalibrations and inflight ca‘ibration. Tape recorder performance wasgood; however, due to the extended burn tine of the S-I1 and S-IVB stages,the S-IC/S-I1 separation data playback was not recovered from the S-IS-I¥B, and IU recorders. This was because insufficiert playback timewas progranred to cover the anomalous situation caused by the S-II two-engines-out cond’ t‘on. Performance of the RF systers sas gecd. fnproxi-mately 2 seconds of real time data on all S-IC stage telemetry linkswere lost due to a data dropout at 146.9 seconcs. This condition waalso noted on AS-5U1 and appears to be related to S-IC IFCC. Groundcanera coverage was good as evidenced by 84 percert system efficiency.However, only two of the six onboard film cameras were recovered. Threeof the cameras an the S-IC stage failed to eject and one of the $-1cameras was not recovered due to a weak recovery beacon signal

xliv

SECTION 1

INTRODUCTION

1.1 PURPOSE

This report provides the ‘ational Aeronautics and Space Administration(NASA) Headquarters, and other interastad agencies, with the launch ve-hicle evaluation results of the AS-592 flignt test. The basic objectiveof flight evaluation is to acquire, reduce, analyze, evaluate and *eporton flight test data to the extent required to assure future missian suc-cess and vehicle reliability. To accomplish this objective, actual flightmalfunctions and deviations must ve idetified, their causes accuratelydetermined, and complete information made available so that correctiveaction can be accomplished witsin the established flight schedule.

1.2 SCOPE

This report presents the results of the early engineering flight evaluationof the AS-502 jaunch vehicle, The contents are centered on the performanceevaluation of tie majar Jadsch vehicle systems, with special emphasis onfailures, anomalies, and deviations. Sumnaries of launch operations andSpacecraft performance are included for completeness

The official “SFC position at this time is represented by tnis reportIt will not be followed ay a similar report unless continued analysis ornew information should prove the conclusion presented herein to be significantly incorrect. Final stage evaluation reports will, however, be pub.lished by the stage contractors. Reports covering major subjects andspecial subjects will be puxlished as required

1-1-2

SECTION 2

EVENT TIMES.

2.7 SUNMARY OF EVENTS

Range zero tine, the basic time reference for this report, is 7:00:0Eastern Starcarc Time (EST) (12:00:07 Universal Time [UT]}. This time,by cefinitior, is tased on the nearest second priar to S-If tail plugcisconrect, which eccurred at 7:00:01.74 EST. Range tine is caTculatedas the elapsed time from range zero time and unless otherwise noted 7the time used thraughovt this report

Guidance Reference Release (GRR) occurred at -16.85 seconds and start ofTisie Base 1 (Ty) occurred 17,54 seconds later at 0.69 second. The tine:noted above were established by the Digital Events Evaluator (DEE-6.except for Ti which was determined by the Launch Vehicle Digital Computer(LyDC}), First motion of the vehicle was established by ground camerasas having occurred at 0.38 second.

Range Lime for each time bese used in the flight sequence arogram and thesignal for initiating each time base are presented in Table 2-1.

Table 2-1. Time Base Sunnary

RANGE TIMETIME BASE SEC SIGNAL START

(HRS MEN: SEC)

Ty -16.85 Guidance Reference Re°easeTy 0.69 Ill Unbiltca1 Disconnect

Sensed by LDCTz 144.95 S-IC TECO Sensed LYDC13 148.41 S-1C OLCO Sensed by L¥DCM4 876.33 S-11 £00 Sensed by LVUC75 A730 S-1¥B £60 (Velocity) Sensed

by Lvoc16 11,287.73 Restart Cquation Solution

(3:08:7.73)

ie) 11,630.33 Commanded LCO based on Thrust(3:13:50.33} Criteria not being met

21

Start of T2 was initiated approximately 0.23 seconds after $-IC inboardengine cutoff. Two redundant LOX level cutoff sensors were located onthe S-IC stage 180 degrees apart. The system was designed so that thesensor which first detected the cutoff level would initiate InboardEngine Cutoff (IECO) by means of redundant circuitry and cutoffsolenoids. Only one sensor circuit, however, was wired to provide theLVOC interrupt signal which would initiate Tp. ECO was achieved at144.72 seconds by means of the cutoff circuit which did not save theLYOG interrupt capability. The second circuit subsequently sensedcutoff level and initiated Tz via the LYDC interrupt at 144.95 seconds.The reason why the level sensors detectad cutoff level at differenttimes is not known at this time and the problem is under investigation.Because of the time discrepancy, both times are listed in Table 2-2

Failure to restart the S-1VB engine terminated 1g early and started T7at 11,630.33 seconds. The flight sequence program norrally commandsengine cutoff and initiates T7 based on velocity attained. On AS-502,however, engine cutoff was commanded and T7 initiated because thrustcriteria were not being met due to the S-IVB stage restart failure.

A summary of significant events for AS-502 is given in Table 2-2. Theost significant deviations from nominal predicted times occurred inguidance and navigation events because of perturbations to the Guidanceand Navigation System occasioned by the premature shutdown of tuo S-1stage engines. R more datailed discussion of these problens is containe:in paragraph 10.4.

2.2 SEQUENCE OF EVENTS

Table 2-3 lists the sequence of switch selector events. “erminology inthis table agrees with the terminology in document 40H33622¢ "InterfaceControl Document Definition of Saturn SA-502 Flight Sequence Program."Eignt events, including S-II engine start, were not verified because oftelemetry dropout during S-IC/S-IT staging, although subsequent eventsindicate that these events did in fact occur, Additionally, 21 orbitaevents and 10 events in T7 were not verified because of stationvisibility constraints and loss of data due to fight perturbations.Probable times for all but two of these events were derived from theflight program. Times were also derived for six switch selectorfunctions which vere verified to 40.5 seconds by compressed date, Tourswitch selector events (Fuel Injection Temperature OK Bypess ResetFlight Control Computer Switch Point No. 5, Point Level Sensor Armingand Cutoff S-IVB Velocity) were missed at the end of 1 due to theearly start of T7.

Table 2-4 lists the unscheduTed switch selector events, which aredependent upon vehicle orientation and position and therefore variable,and also ground conmanded events, which have been verifiec fromavailable data.

Table 2-2. Significant Event Times Summary

RARE TIAL TIME FROM BASEactun. [acz-peew actuat acT=PReo

evENT sec) (sre) {sec} (sit)

T. Guidance Refercnce felease “16.86 “o.75] -17. 54] -0.392. Sel¢ Engine Start Secuerce Ccumard 3.77 | 0.02) -9.48: 0.39

3. Range Zero 0.00 | - -

4, ANT Holddosn Arms Released 0.36 ao] 0.33 0.04. First Kotion 0.38 +023) -0.31 146. IU Unbilical Digcornect, Start of

Tine Base 1 (1) 0.69 -0.37 q :

7. Begin Yaw Maneuver 18 20 12 0.28. Erd Yaw aneuver ve aad “9. Begin Pitch and RoI] Maneuver

{Titt and Roll) Wa 0.8} 19-4, -0.21. tnd R011 Harewver na e.7] 30.8 im11. Pach 1 Achfeved 60.5 sla} 59.8] 0.912. Gecurrerce of Max Dynanic

Pressure (Max 9) 75.2 4s) 78.5 “4413, End Pitch Maneuver (Tilt Arrest) 140.9 -2.7/ oe] 24

4, Sc1€ Inboard Engine Cutoff (TECO)(Sclenoid Activation) 144.72 on a -

18. $-IC Inboard Cnaine CutoffSersed by L¥DC, Start of Tirebase 2 (Ta) 144.95 3.34 Te -

16. SIC Outboard Ergine Cutoff (OECO)(Sensed by LIC), Stert of timeBase 3 (13) 148.41 0.88 13 -

VI, S-IC/S-15 Sopavation Command toFire Seoaration Devices andRetro Hotars 149.08 0.82) 0.67] -0.08

18. SeIT Engire Start Command 149.76 0.80] 1.35] -0.0819, SIT Secord Plane Separation

Cannand 179,08 o.eo} 30.65] -0.0820. Launch Escane Tewer (LET)

Jettison Conand v4.77 0.81| 36.36) -0.0821. Initiate Iterative Guidance Mode

(1GH) Phase 1 190.95 1.35 42.54) 0.5422. Initiate Steering Misel1gnnent

Correction (SHC) 71.99 v.76] 63.58] 9.38

Table 2-2. Significant Event ~imes Summary (Continued)

RANGE “IME TIML_ FROMBASEA ner-pneo acruar ACT-PREDEVENT (sec) Gee) a (sec)

23. S-11 Stage Engine Ne. 2 Out 412.92 +

|

266.51 -24, 5-11 Stage ingine Ne. 3 Out 414.18 =

|

265.77 -25. Guidance Sensed Engine Hixture

Retio (ENR) SKIFt, Inftiate 16HPhase 2 490.76

|

78.03 1a2%. Initiate Chi Freeze 817.7 7a 7.0227. $-11 Engine Cutatf (ECO) (Sensed

by LubC), Start of Tine Base 4(ta) 576.33

|

58.64 ty -28, S-II/S-1¥8 Separation Command toFire Separation-Devices and Retra

Hotors sce

|

sese

|

ors] -0.0529. S-IVB Engine Start SequenceCorman s772e

|

sass

|

0.95] 0.0530. Stop Chi Frenze 582.9 55.8 ai “2831, Initiate Tteratve Suidance Made,Phase 3 seize

|

s7.cr7

|

ass

|

-.9732, Pitch Command Hese-up Attitude 644.02 -

|

67.69 -33. Inittate chi Bar Steering N28 25.8

|

136.6 27.234. Initiate Chi Freeze 76.4

|

93.67

|

679.07

|

38.0235, S-IVB Yetocity Cutotr

Cormané (ECO) 77.08

|

e778 |ts-0.26

|

-2,0636. S-1VB Engine Cutoft Sensed byLWOC” Start of Tine Base 5 (TS) 7a7.30

|

a7.eA 15 -37. Coast Period Om mass

|

e779

|

1.25

|

-0.n538. Parking Orbit Insertion 757.08

|

ar.7e

|

9.74

|

-0,0639, Maneuver to Local Horizontal 762.30

|

a7.s4

|

15.00 0.0040, Initiate 189° Roll to Place !Position JF Down 837.30

'

7.84

|

99.00 0.0041. Initiate 20° Pitch Lown Maneuver 3207.30

|

87.81 eas0.00 9.0042, Initiate 20° Pitch Up Maneuver 5427.30

|

67.84 fised.no 0.0043, Initiate 180° Roll to Place

Position I Down 5707.20

|

e748 040.00 0.004a, Inftfate S-1NB Restart Sequence

and Start of Tine Base 6 (Tg) V.2e7.73

|

ate 15 -45. S-IVB Engine Restart Connand V,614.69

|

21.08

|

326.95

|

-0.05

24

Table 2-2. Significant Event Times Summary (Continued)

RANGE THRE TIME FROM ASEacTua, [act-peea acruat

EVENT see se 5eC

a6, §-1VE C0 Command and Start ofTime Base 7:7) 71,630.23 -97.66 y .

47, Coast Period On 11,631.50 -97.79 117} -9.0348, Maneuver to Separation actitude 11,650.33 -97.76 20.00 9,0049. End Cold Soak Attitude and

Spacecraft Separation Cormand™ 11, 666,02 - 35.69 -

50. L¥-LTAYCS¥ Physical Separation® 11,667.82 ~240..27 3.49 142.51

51. Execute Maneuver A* 16,201.0 - 4570.7 -

52, S-1VB LOK Tonk Vent ValvesOpen* 22,023.30 = 0,392.97] -

53, 5-1¥8 Lig Tank Vent VabvesSyen* 22,024.21 = 0,399.8 -

Ground Command

2-5

Table 2-3. Sequence of Switch Selector Events

RANGE TIVE [VINE FROM uaseactu. acTuat] ACT-PREDFunc 70N STARE

|

(See) ($t¢} (sec)1, Start of Tine Base 1 (hy) w 0.63 - -2. buto-Abort Enable Relays. KESET w 5.68 as

|

0.053. Sensor Bias, UW u 6.36 sr] -0.084. Multiple Fugine Cutoft, ENABLE sic} ses

|

13.96

|

0.085. Teleneter Calibrate, oN sac} 25.4

|

ans

|

3.088. Telumetry Cal forator In-FlightCalibrate, DN Mu ase

|

26.or

|

0.037. Telenetor CoTtbrate, OFF sac} aos

|

20.77

|

0.083. Laureh VehicTe Engines £05 cutoff,ENAELE w wo.ss

]

2.97

|

0.089. Telenetty teliarator In-FlightCalibrate, OFF w 32.88

|

5.95

|

0.0510. Fuel Pressurizing YaTve (KECV) ho, 25OPER &'Tape Recorder RECORD sic] sora

|

49.45

|

0.0511, Start pata Recorders sear} 74.76

|

7.05

|

-o.05V2. Cocling, System Electrica? Assembly,

Power, OFF Ww 75.66 74,9713, Telemetry Calibrator In-FlightCal Sbrate, ON w 30.65

|

99.97VA, Telemetry Ca}ibrater In-FlightCalibrate, OFF w 96.65

|

989s

|

no15. Fuel Prascurtaing Valve (HFCV) no. 3.

|

5. .Fuet sc] 95.95

|

95.26 0.0816. Flight Control Computer Switch Poratto. w wos.es

|

08.97

|

0.05Teleneter calibrate, oN sac

|

nises

|

tise

|

oes18, Regutar Calibrate Relays. oW swe] nig.5

|

tia.t6

|

0.0819. Telemeter Calibrate, OFF sic] 16s

|

mer

|

0.0320. Flight Sontrot Computer Switch Poinehon? w 120.84

|

120.15

|

0.0521. Regular Calibrate Relays, OFF seve} teees

|

aaas

|

«0.0522. Start First PRLAVEH Calibration

—|

set] 230.38

|

129.67

|

-0.0328. Fue) Pressurizing Yatve (HFC) Ho. 4] ste

|

aoete

|

saan

|

oan2b. Fast Record, a sa] i383

|

sear

|

0.05

2-6

Table 2-3. Sequence of Switch Selector Events (Continued)

ee eeFUNCTION stage ACTUAL (secafactunt tscen) Neeehe?

ZS. Ston First PANGFIVN Calibratton sat

|

138.38

|

134.65 “0.05

26. Tave Recorder Record, Mk w 135.55

|

134.86 0.04

27, LOK Tank Strobe Lights, OFF sere

|

t35.78

|

135.05 0.05

28. S-I¢ Two Engine Qut &uto-AbortTnhioit, EXMBLE ww wss.o¢

|

138.26 “0.04

29, S-1C Twn Eng’re Out Autocdbort Inhibit] 16 ise. | 135.48 -0.02

40, Excogs Rete {P,¥4R} Auto-Abortinnibit, ENAMLE w 136.34 135.65 0.05

31, Excess Rete {P,Y4R) Aute-Abort innibst

|

1 vies

|

125.86 0.05

32. Two Adjacent Outboard Ergines OutCutorf, ENABLE sc

|

136.76

|

136.07 -0.03

33, Inboord Tagine Cutoff, EAGLE sc

|

136.98

|

136.27 0.03

44. Intcard Engine Cutoff Seckup, EVABLE

«

S-2¢

|

137.74

|

136.45 -0.06

35. Ingourd Ergine Cutoff Intereupt 1520 184,95

38, Start of Tine Base 2 (Tz) 144.95

37. Sil Ordnance, ARM sb] 145.03 0.0828, Separation and Retro EBM Firing

unttsy, ARH s-tc

|

145.20 0.28 -9.05

39. Separation Canera, OW sec

|

145.42 0.47 “0.03

40. Camera Lights, ON sar] vas.st 0.86 -2.08

41, Telemetry Measurement Switch Over sc

|

145.70 0.76 -2.05

42. Outbosrd Engines Cutoff, ERABLE se] 145.91 0.96 “0,04

43. utboard Engines Cutoff-Interrupt 1520 143.4144, Start of Tine Base 3 (Ta) 148.41

46, Camera Motor, OW se] 148.49 6.08 0.02

46, SIT Lez Recirculation Pumps, CFF sett

|

148.66 0.23 0.08

a7, ST Ullage, TaIGBER sr} 148.87 0.a7 “0.08

48, S-IC/S-1I Separation set

|

149.08 0.67 0.03

49, Camera Event Mark sear] 149.17 on 0.03

2-7

Table 2-3. Sequence of Switch Selector Events (Continued)

RAKGE Tier] TIME FROM BASEFUNCTION vase] (GU serPRCD

5a. Switch Caging Cantrol =e S-Ify S-1TEngine Out Indication *A", ENAILEsS-IL Aft’ Interslaye Senarasten Indication "a", EuAacc wu 149,264 ese

|

-9.0651. S-1f Engines Cutoff, RESET 5-11 oss

[

-o.0552. Engines Ready Bypass su 149, 46" res

|

0.0552. Provalvas Lockout. RESET S-lL 149,96" tase

|

-9.05SAL SoEL Engine, START su 149.764 rast

|

0.0558. Camera FyenL Mark st 149,86" nase

|

0,0556. S-11 Engine Out Indication "G*, FYABLEand $-I1 Aft interstoge SeparationInd“cation "8", ENRELE tu 160.06 Tog

|

-0.0587. Engines Ready kypass, RFSET set} 150,26" Tse

|

0.065@. Qefe11 Power, OFF ww 180.76 2.35 -0.0559, SIL fydrau'ic Accumutetors, UMLack set

|

151,38 2.97 0.036D. Chilldown valves, CLOSE sr] 158.76 6.38 -9.0561. S-i1 Start Phase Limiter cutoff, ane

|

$-11

|

188.08 6.67 0.0362. Acevate PU System Sett

|

755.26 6.85 -0.0563, SII Start Phase Liniter Cutoff arm,

|

5-11

|

156.07 7.07 0.03Reser62. Stop Lata Recorders Sf 159.76 11.38 -0.02

Fast Record, OFF S-Ive] 159.96 11.56 -0.0465. Tape Recorder Record, OFF w 160.17 Wa 0.0367. S-LL At Interstage Separation, ov

|

sur

|

179.06 30.85, 0.0558. camera Event hark sD] 479.16 30.75 0.0839. camera Evert Mark se] 189.16 31.75 0.0670. Launch Escape Tower Jettison "A! w 198.77 36.36 -0.0871. Launch Escane lover Jettison "8", ON Ue 194.98 36.57 -0.03

Camera Eject Ha. 1 sat 186.97 37.97 -0.93Camera bgect ho, 2 SIL] 18.97 38.56 -0.04

74. Camera Eject No. 3 $0 137.46 39.08 0.08 — * berived Times — 2-8

Table 2-3. Sequence of Switch Selector Events (Continued)

BINGE TIME [__TIME FROM BASEnner ACTON, mean [ner pacneTTaN srast) (sre) (step (30)

75. flight Contra) Conputar Switch Paintwo. 3 w 29.76 61.36 0.04

Te. Start Second PAM-ESYFY Calibraton [ser 30.26 tet.es -0.0517. Stop Second MAN-FA/FH Calibration ser aas.zr 65.80 “0.04Tk. Fight Contra? Conputer Switch Paint

mw. 8 w aaa.76 161.35 0.0819, Telemetry cat ibratar In-Flight Cali-

rate, OF w 350.26 21a $0.08a0. Twlenetry Calibrator In Might Cait-

rate, OFF w 385.27 206.87 “0.08Ah. Measurerert Sortrot Sui teh "o.

ACTIVATE str} 3et.06 212.66 “0.0882, Start Third 20GFW/FH Calibravion fs-t2 gaa.se z7z.se -0.0285, Stop Third PAN-FA/FH Calibration set] 45.97 277.56 0.08

S-IL LH2 Steo Pressurization st] 46a 319.95 -0.05fequtar C21iarate Relays, Of seve] 47.26 s2a.89 “0.08

8 Teleretry Calstrator In-Flicht Calfrate, Of w arraa 329.07 “0.03

87, Regu’ ar C2liorate Relays, OFF seo] 4az.an 333.07 0.03BE. Telametry Cal‘brator In-FPight

Calibrate, Arr w 4e2.a6 394.05 -0.0588, charge Ullage “gnition, OY s-1ve) 492.08 334.48 -0.029. S-I1/8-1¥8 Ordnance, ARP sr] 4s3.c6 398.68 0.0541, Tape Revorder Revord, OH w 4s3.27 338,87 -0.0392, Fast Record, OW serve] 483.45 335.05 0.0892, Start Dace Recorders st] 483.67 395.26 -0.0844. S-11 LOK Dentetion Senser Cutott,a st] 4a3.28 35.07 -0.03

95. $-11 Li2 Depletion Senser Cutoff,ARM st] asaes 335.68 “0.05

9. Cutoff S-1T J-2 Enginas-Ioterrupt 1520 576.3997. Start of Time Base 4 (T4} 576.99

98. Redundant S-IT Cateff $8 st] 576m 0.08 0.08 2-9

Table 2-3. Sequence of Switch Selector Events (Continued)

fans Tne [Tae tos eeroverro1 sea] (QE FEMA sereo

Sh Start Recorder Tare Sr} aes Poe ts10. eves Close, OFF swe 0.

|

og1D S188 trains Cart, 268 sts ox

|

ear102. trpine Ready dypass seve oa

|

so.s8TW, LOH CHIL don Bump, cet sae ase

|

coon108, Fire Litage pnie-an, o4 sa 06s

|

asVos. SLU/S-18 Separetion sat ars

|

ons106, S.1¥8 Enine start Tntertock Byetss,a sie] gre bas

|

0.ToT. S1¥8 Engine stare, oN sue} sze

|

os

|

as©, FTIghE Contr} Gompster S188 Burrons w

|

sna ft foes10. Fight control Computer S=1¥B Burnroe bn ee w

|

sae foe oo.08Ne. $208 Engine Oat Inbscationfhiate w

|

sas

|

aes

|

oes11, S18 Fagine ave tndieaion *B",euttte wo} sea

|

ar

|

cos12. Fuel Ghitldow Pcp, OFF sete) sree

|

cas

|

os5. LOH Tonk Flight Pressure Systen, ov [serssl 20.08

|

ozs

|

ou14, Fuel criection Tenperacure OF tyoss j5-tv0] 520.0

|

aor

|

a en1S. Satnaive stare, oFF sre) oes

|

ars

|

os16. First fun Relay, om seve] seas

|

srs

|

cos11/.trergeney Playback Enable, ot sre} see.

|

aa fos1B, Fast Record, O47 sun] sean

|

a5

|

os18. pu detiaee, ow sess0

|

3.97

|

-a.09Ten. Charge Jaye uettisony a sen] sae.0s

|

oar

|

00s1, Five Uiage dettison, a swe] sse.c

|

tere

|

aes12, Tye" Injection Tenpereure Uk Bypass.fie serve] st0.28

|

a2.122. lage Charging, neser srel sie

|

t5.08

Table 2-3. Sequence of Switch Selector Events (Continued)

roi swale [eet

|

Ta

134, Start of Time Base 5 {Ts} s-IvB] 747.30

149. Aux. Hydraulic Pump Coast Made, OW ‘$+ 1NB] 751.15 3.85 0.08

Table 2-3. Sequence of Switch Selector Events (Continued)

SeWGE TIME] TINE FROM bsaccu, at hsf s HiWeTHOR STAGE (sec) sec)

145. Aux. Hydraulse Puup Plight Mode,OFF scm] fe1.37 sat -2.05190, S-Teé Engine ou; Indication -#able, SESET w 27.27 ga “2.08151. S18 Cogine Out Insication "6" bn-able, RESET u reras

|

aaas 0.05VP. Letenotry Calibrator In-Flight Salibrate, GH W reg.as

|

12th 0.0813, Regular Calibrate Relays, ov 789.66

|

1.26Vol. Telemetry Calibrator In-Flight ce1i-urate, OFF tw 784.47 -9

Regular calfbrate Relays, OFF Seve) 702.85) 17.35 “2.08SS/FH Transmister, OFF sev] 769.26

|

21.96 “0.84SS/EH Sreua, OFF serve, yeo.ee

|

22.16 -2.08158. Lilz Tank Cont-nuaus Vert ¥aIve Open,fH sive] gus.23

|

50.55 t.08156. Lig Tank Contiouwes Yeat Yuly:Grr s-tve} 08.26Vee. $148 LVage Engine Yo. 1, CFF Serve} 35.28 787.95 9.8TET. SIRE UNage Engine ho, 2, OFF sey] 235.38

|

98,05 0.88182, S-1¥0 UNtage Thrust Present Irosca~tion, 3FF wu 235.96

|

eee 2.04163, Tergency Playback Enable, 08 seve] 335.75] asese 0.8V64, Tamm Kecorder Pleytack Reverse, on fy focas

|

eats “6.05265, Erergency Alayback Enable, OFF serie] gue.as*

|

62 ase 0.98166, Slow Record, 14 903.45

|

ee te “3.08TE? Stow Record, Ch aig.ase

|

12.154 “Vee. Tape Recorder Playoack Reverse, GFF ge.ast

|

173.19 “1.08182. Engine Punp Purge Control alve Enatie, OFF 1349.55170. Siew Recor, a4 2200.78" -0.05V7. Slow Record, FF 2aye.75*

|

16as.ane “0.08172, Recorcer Playback, Ck 2392.95

|

1605.00"

* Derived Tires

2-1

Table 2-3. Sequence of Switch Selector Events (Continued)

eavet_ Tne |_FIF Tan past

romero tat || MacsTa Recorder Maybach, &F wean. 7e |de “ht

74, Slow Record, i 1a13.65¢ “2.0

Vb, Slim Recore, &H 1623.65 “0.08

Vt. Tetenetry Caiigrater in-flight Cals- ‘brates Ly asans 2899.45 “0.05

177. Regwiar Calibrate Relays. ON s-t¥8| a3c.95

|

2593.65 -0.08

138, Telenetry Gol Ibratar Ta-Fl ight Coli-arele, OFF we] ae4s.z7

|

2son.a7 9.03125. fagular Calsbrate Relays, OFF sera] aiee.ye 7588.87 0.08180, ie lemetry Calicrator r-Flicht Cali :

tratr, Of i uaep.7e+ 030.458 “2.68La, Stow Hecord, ah sr 4638.60"TP. Requ' or Uelinrate Relays, 0 see sage as “0.08183. Tevenetry Clibrater [n-Fléght Cali- |

urate, 2° w 3390.75

|

4643.45 0.08

1ae, Regular calibyate Relzys, 1 seve] $351.15.

|

4683.26 “0.0ty, Sow Record, OFF sewn] sarz.9s |4079.65

186. Recorder Playback. 0 ‘sstvs] 5418.95 4671.05 |

Tat. Recordar Playback, CFF sete] stare

|

6027.45 2.0

148. Slaw Recave, £& setun[ resis

|

2027.n6 <n.129. Stow Bacar, OH sewn} rap.t5

|

5037.25 “tus190, Telenetry tuTiurator In-Flight Cali

urate, tu

|

640.75

|

wea ae 0.95191. aegular celitrate Relays, OW sews] oxto.os. 5893.85 “0.08132, Telewetey Sal brator In-Flight Caii-

rate, OFF w

|

exe.76

|

$593.46 2.0193, Regular Gatibrate Releys, OFF 3.86 5808.66 cao194, Slow Record, s-ivu] 7997.754

|

7160.45" -0.9%

19>. Staw Record OTF seve] 7930.27

|

7192.48 “pe

96, Recorder Payback, In sel 7999.97

|

7192.67 0.03197, Recorger Payback, OFF slo] e2to.se

|

7083.26 “0.08+99, Stow Record, GH | setvel eesovrgee

|

reesae*

|

0.08

+ neriveo Tiress+ Verifies te 10.4 by Conpressed Oete

z-13

Table 2-3, Sequence of Switch Selector Events (continued)

RANGE “INT

|

s1Pn TRGBASEFUNCTION stot] (ete

135. Stan Fecerd, OY eeee.ts

|

aaa “Aa700, Regular Calibrate Kelays, 4 8900.75

|

9733.25 1.08201, Telemetry Catibrator In-Fliget Sol t=ps th uw] aseo.ss

|

e2ar.ag “3.08702. legular Calibrate Relays, OFF sts] esee.7s

|

2236.05 3.08203, Teieretey Calibrator to-Flight Cali

brate, OFF w

|

sses.o8

|

geae.e5 £8.08208, Siow Recora, ad 10,410.75"

|

96e3.456 “0.05205. slow Record, FE servi] 10,402.75"

|

yea5.e5¢

|

0.05206. Becncder Playback, CH 0,492.95"

|

9795.65" 2.05207. Reconear Playback, CFF 19,707.79"

|

998645"Slow Record, 04 310,707.95

|

9960.65* 6.08Slow Record, tH 16,717.95"

|

9y20.55*210, Aux. stydewuic Pump Tight Hoce, 04

|

s-tvel 10,822.25¢* |in,aze.gg| -u.0s21. un. Hydrautic Pump Coast Sedo, oF | s-1¥i] 19,822.49" |10,075. 124"212. LOX ChiTIelown Pump, Om seve 10,872.25|rajrea.ase | -2.95213. fuel Chil tdawn Pap, OH s-1¥@] 20,877.26"

|

10,129,954

|

-¢,05Prevalves Close, Of $-t¥8] 10,877.25hro,ra9.95e | 0.35Telemetry Calibrator In-Flight cali-sratey OW Faw |r,260.78 [5,913.5

2b. Regular Calibrate Relays, 0 11,260.95 hia,sia.es 0.08817, Setewtry Calibrator In-Fhight Ca1i-brate, OFF w [11,265.76

|

10,516.48 0.561218. Regular ca iorate Relays, GFF 11,265.96

|

1w,sta.¢6 2.08219, Begrn Restart Preparations ~ rinefase Sackup 1836, seta205. Start of Time Base § (Tg) 11,287.93221, SIP Lage Engire Ww.

1,

of «287,90 ol? 6.03222, S-1VOUNage Engine ho, 2, OH Stee] 11,288.01 ver 0.08223, S-1¥ UMage Thrvst Present

Indica=iory 1

|

rt28e.19 3.46 ~.04

* verived TinesVerified te 20.5 by Carpressed pata

Table 2-3, Sequence of Switch Selector Events (Continued)

fect Tent TiME ROW onsFuey ou ace] Tue acTunl AFpate

22a. U2 Tank Vent Valve boost Clase, ON [5-108] v1,208.00 0.76225. LP Tank Vent Yalve Boost Close, oh s-1vB] 11,268.69] 0.9622. Lil Tank Continuous Vent Yalve Close,

ow Setve] 11,288.89 1.18 0.05207. Ceband Teansponders ho. 1 ard Wo. 2,oa mw rizs9.0y 1.35 0.0228. 1h2 Tonk vent Yolve Boost Close, OFF s-1ve| 11,290.09) 2.75 8.08229. 0X Tank vent Valve Boost Close, OFF Scive] 11,290.69 2.96 1.0823U. Lip Tank onténuous Went Valve Clase,Fe swe] y1z00.91 a. 0.02231, Fuel CriT dons Purp, OW Seve 11,293.09 5.95 “0.08222. LAK GhiTIGewn Pum, oF seve] 11,298.69

|

10.98 0.05233, Prevalves Close, OY S-1¥B] 11,308.69 PR 0.05

234. LOL Tene Repressuization Cantro}Valve pen; Dt Seve] 15,387.62) 99.98 0.05295 “elenetry Calibrator In-Flight Cal¥-brate, OF ww

|

11,436.00

|

14a.27 0.03236. Tetenetry Calbrazer Te-F1 ight Gali- |brate, OFF 11,800.99 183.75 0.05237. LM “ank Repressurization Contra?Wave Open, 08 seve} 11,487.70 199.96 0.08208, SS/EH Group, Ot serv] 11,496.00 20e.26 0.08239, S5/TH Transmitter, ON s-avel r1,ase.2r pon 48 0.02740, Reqular Calibrate Relays, OW s-tve} 11,386.19

|

268.45 0.05241. Requtar Calibrate Relays. GFF seve] 11,569.20

|

272.86 0.08249. PY Yalve lardever Position, Ok Serva) 11,574.69 286.96 0.04243, Frevalves Clase, Ott S-rval 11,603.90 316.16 “0.04244, IVE Restart ert w 1,608.69 316.96 0.08205. S-IVR Engine Cutort, OFF sive} 11,613.29

|

326.56 0.08240. Eng're Ready Bypass seve} 11,613.51

|

325.77 0.03207, Lip Tan Repressurizatior ContrelYalve Deen, OFT s-wvel 11,813.69 225.96 0.05

Table 2-3, Sequence of Switch Selector Fyents (Continued)

ear Tee sen aseFUNeTiON erage

|

ACTURL aca] AL1-RFDSE) asec) Isre} isrcy

Fuel ChiltdownPurp, OFF seve 51,615.99LOX CV dour Purp. DFT Serve 11,514.09

|

326.96 6.08250, i OX Tenk Renressurization Control

valve Open, "FF nest

|

326.77 “03251. S-L4 Fugine Start, DE Heraey 36.35 2.08152. SIH Lagine Cut Indication

ENABLE tu ve1s.49 sze,7823, S-LNM Engine Qut Traication "a",

TeAGL | 11,818.70

|

327.96 9.01258, S208 JFlage Eryn fe, 1, OFF Vie27.69

|

sz9.95 0.05255. Yaye Engine Re. 2, DFF Tels

|

930.08 “0.8256. Seu [lage Thrust Present

Indication, OFF w

|

11,617.93

|

2an.26 0.04257, Tight Cantrol Computer SIR burn

Pode w

|

11,622.29

|

33.55 -0.05,PHB. Flight Cantrat Conputer S-I¥B Burn

Mode Du "b" tu

|

trees

}

29.76 0.789. Fuel Injection cewerature DK bypass

|

S-(¥BF 11,622.69

|

334,¢6 “0.9%260. LOK Tank Tight Fressure system, pk

|

sive] 11,622 90

|

395.17 5.03761, Coast Perind, OFF serum] 11,623 335.36 -0.08287, S-1N@ Engine Start, Orr s-1waf 11,623.30

f

294.06 2.04263. Second Burn Relay, OW serve] 11,625.29 2.95264. PU Activate, CH sore] 11,627.50285. PU Valve Fardover Position, OFF serve] 11,027.68

|

239.55 2.05266. Fuel injection Temperature Ok Ryyass.| S-1¥B] teRESET

207. Fight Control Conputer SwischPaint No. 5 wu a

268. Soin Level Sensor Arming Seve} te268. Cutoff $-1¥8 velacity w ae270, Start ef Time Base 7 (T7) (ECO) s-1¥8] 10,630.93211. Redundant S-1¥8 Cusoff S-18] 11,630.41] ¢.08 6.08

8" Started new time base beforethese events could take olare

Table 2-3. Sequence of Switch Selector Events (Continued)

cin rarrw ome] et Bet |a

mtnsens creas nf tema na nmmt uatowsen nna [taf tannin nmbe seatevan cin] wean oe am[ne socom rpmess ssn or form] numae| nae taarr, tease Period, 04 seve] vosof iar 0.03bone PL Activate, OFF ‘11,631.70 Le 0.92

bn remmaraea fe] mcm nan |2 Ue ng on oeWiehe cma} nana] ose ommi fierrS18ese n sma] sae fommeesseeaay w vane] sar omwo. ha aenap rete ee]te] sae foni emg ceeniee o mer] so fomco. mgmartrante ys. fin] sneeee cae fsve flee unser sanete aw nama] sae ooco. mavens woes or [son team] sar |cm sam renniter om cin] meno] ae neco sme.om cra] wus} see omim. inne vecin fem] sama] sae |2h eto anwe: can] nama esr oncpsessa sanf vis] nse omor unrtmnne,cae em same) nas asmt tsaese, [ee] tte eae |tm. cart wenen cone[ena] sist cae |ote rites soe | in| wae [os

Table 2-3. Sequence of Switch Selector Events (Continued)

TIML FROR PASE

ke7 acral

|

ACTIAL

|

ae) #arpueWACTTON isc) (3e Tsk)257, Swtten PCM to Lew Gaie arccamsail sates ty

|

12ya90.e8*| 1199.95"BH, Sw tch C28 co Low Gain Antenne w 2,890,484] “Zoo.164759. (Hy Tart Conesuuus vent talve Open.oH : 7.ex¢.66*] joer |30,312 Tash Cuntiruods Vent talve tnen,ae seve] 12,232.68] 1202 -2.08201. Telemetry Zaliorator In-Ttignt ¢2"i-trate, fi 12,980. 0.0302. Regular Calébrare Relays, OW 12,9¢0.78"] 13in.ase

y

p

0.419. Ta" @rwery Calibrator In-Fiigat Coli~brates GFF ww

|

32,985.98 ars.r5¢

|

3.05SBE, Reqular CaLibraze Helays, AFF Serve] 12,905.78] 1ais.ase ,-a.aeWE. Switer PCA to high Gain fntenna w

|

17,030.28

|

5239.98 -9.0830D. SuVLeh CCS te High Gein Antenna(t81 Safe) | ¥7,030.48 7 stn0.17 2.02307. Televetry Cubibrator [a-Flight vali=brite, Oh | 17,140.89

|

ssis.28 “6.04303. legu!ae Ca" ibrate lays. aw sive] 17,140.79

|

5510.45 “3.04309, Telenetey Calibruve In-Flight ca°4-hate, SF w

|

17.15.58

|

5538.26 -£.083M. Regan Calibrate Relays, 3F: seve] 17,145.28

|

9515.a5 0.08a ntiruous Vent Vaive Case,

5-16) ho Gata - :ae nh Certinuows Yent Valve Close, j

s-tv3|_ No dara -

+ berived Les

Table 2-4. Unscheduled Switch Selector Events

TAGE Tie teeenw smace ta ese rimacessec) (Sto)Tater Cootant valve, OPW a Tee. Tyeas.ae 7 s181 68

Noe insti oeater Coolant Valve, CLOSEE w 104.08 Tpe06.78 Late eamcttoncoRane Transacederfea. 2, w ase0.e0 tys2eae etCoBane Tranesender Wo. 1, err ce 3500.8" rosaeme.s7 mn, 2 Activec-Bane Transaoncer 1. 18fo. 2, oh w e211 tgss074.21afane Teamanonter 08,1, 06F Tu e218 tyesoma.za nD. 2 setteeUster Coctent Valve, OPE uw 91.18 rpesiaz.aa Lane functionC-Band Transoonder Hl. 1 &Wo. 2, OF uw ses.81 apestta.2tCetand Transponder 90. 2, 0FF IU 125.58 reysi7a.zn wn. 1 setteeater Coolant YaTve, CLOSED w 6191.65 T3#540. 35,}é-tand Transponder 1. 1 4tone, on uw our. Tys5068.91Setard “ransponter md. 1. ttt [av eeu 28 rgesuag.e m0, 2 vectveater Content Valve, CLOSED q 7495.44 tyr7eta.ts We fwetton‘Dond Teorspender Me 16fae, of w 11,069.08 Teete,a2.74Ustad Trarseondar NO. 2, 08F WU rryosa.11 resteazt.er 0, 1 acetveGeert “rarsporder 0. 8ow w 11,650.84 rp20.12

Trarsporder wo, 4, 0rF ww 11,650.54 qpenat 40, 2 acctveG-dand Trarsponder WO. 1 60.2) OW uw 1680.88 weesCotand Teaesonder Wo. 25 off tw 11,688.75 Tess? mt. 1 acuteCotand Trarsnender WO. 18fe 3, Ow w 1699.19 y68.80C-Band Trarspander Wo. 1, ofr tu 11,600.20 triss.a7 uo, 2 activeC-Band Trarsoander NO. 1 &fo. 2, oh n 1,109.58 tetr.onC-Band Transnancer wt. 20Fe rt 3707.40 seca Wo, 1 activeC-Band Transponcer Wi. 1 &ho, ey OF Ww 72.68 rye.28C-Band Transponcer Wa. 1, orf tu lire 68 ty97.36 fw. 2 activeC-Band Tramsnoncer W. 1 8to."of w 17,280.66 tyraaa.aaaeana Trangpancer 10. 2, 08F | 11,260.73, tpveeo.aa Bo. 1 acttee(ater Coolant Yebve, OPH % ree ayesta.4 tive sanctionHyeraulte Funp Coast Nade, off 5-148 14,96? Tat] around cerrandFycraut le Fun Flight Mode, on 5-148 14,961" aysuat around cormandTelenetry Cel torator, 0h w 18,008 reat Ground cormandReguler Calibrator fetays, oN 5-1¥8 s.an4e reser sraund carmandHydraulte Funp Coast Mode, Off s-t¥8 ts,oa7* ryesto7 Ground commandHycraulte Pump Flight Wade, ON Setve 3,037" tata? Ground ermrandcholic Pury Flight Wade, OFF 5-148 15,080 Tres Ground commandTelemetry Calibeator, OFF w 6.173" Treasaa Ground comandRegular Calibrate Relays, orf §-1¥8 e173" Tyrasty round comand “aru Transeatat tives used becausepitse cites not available

Table 2-4, Unscheduled Switch Selector Eveats (Continued)

roy Te eTeam ema Tne ost sentire tse:EeeTovaner F 7 Team Spe

|

tant PadCetin Transponder 89.18reeds oF | verses) resise.teBand tensvorder no. 1.6

|

ou

|

rszaese

|

tyne

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2-29

SECTION 3

LAUNCH OPERATIONS

3.1 SUNHARY

The Saturn AS-502 was the second flight vehicle of the Seturn ¥ ApollaFlight program. The basic purpose of the flight was to demonstrate theconpatibility and performance of the launch vehicle and the Apolto Commandand Service Modules (CSN) for manned flight

The launch countdown for AS-502 was completed without any unscheduled count-down holds and the vehicle was successfully ‘aunched at 07:00:01 EasternStandard vime {ES™) Apr'l 4, 1968.

Ground systems performance was highly satisfactory, The relatively fewproblens encountered in countdown were overcome such that vehicle launchreadiness was not conpromised.

Launch damage to the complex and support equipment was minor. Nodi fications to the ground syszems were effective in reducing the arount of blastdamage below that sustained during AS-501 Taunch

3.2 PRELAUNCH MILESTONES

A chronological summary of events and preparations leading to the launchof AS-502 is contained in Table 3-1.

3.3 COUNTDOWN EVENTS

The launch countdown for AS-502 was picked up at -24 hours at 1:00:00 ESTAoril 3, 1968 and proceeded to -8 hours with no holds. At this point thesgcreduled six hour hold period was initiated. The count was resumed fromB hours at 23:00:00 EST April 3, 1968, and culminated in the successfulJaunca of the vehicle at 07:00:01 EST April 4, 1968.

Only four significant problems developed during the Taunch countdown. Althese problems were resolved vefore the end of the scheduled six hour holdperiod. The items are stated below in chronolagicat order

a. Several LH2 vent bubble caps were found ta save cracks exposing thevehicle vent system directly to the atmosphere

3-1

Table 3-1, AS-502 Milestones

DATE EVENT

February 21, 1967

March 3, 1967

March 20, 1967

March 29, 1967

May 4, 1967

May 19, 1967

May 24, 1967

May 29, 1967

May 29, 1967

May 31, 1967

June 1, 1967

June 8, 1967

dune 13, 1967

June 29, 1967

duly 6, 1967

July 13, 1967

duly 24, 1967

duly 24, 1967

August 8, 1967

S-IVB Stage Arrival

S-IC Stage Arriva?

Instrument Unit Arrival

Erection of Launch Vehicle (L¥) with S-II Spacer

LY Clectrical Interface Mate Test with S-ISpacer

LY Guidance & Control Tests with $-I1 Spacer

S-II Stage Arrival

LY Propetiant Dispersion Test with S-I1 Spacer

LV Power Transfer Test with S-II Spacer

LY Emergency Detection System (EDS) Test withS-II Spacer

LV Flight Sequence and Exploding Bridge Hire(EB) Functional Test with S-II Spacer

LV Sequence Malfunction Test with $-II Spacer

LV Plugs-In Overal? Test (OAT) No. 1 with S-ISpacer

Ce-erection of the L¥ through S-II Spacer

Completed S-11 LH2 Tank Inspection

Erection of LV with S-II Flight Stage

LY Electrical Interface Mate Test

L¥ Switch Selector Functional Test

LY EDS Test

3-2

Table 3-1. AS-502 Milestones (Continued)

DATE EVENT

August 10, 1967

August 11, 1967

August 11, 1967

August 30, 1967

Decerrber 10, 1967

Lecember 11, 1967

Cecember 21, 1967

December 27, 1967

December 29, 1967

January 4/8, 1968

January 16, 1968

January 24, 1968

Janvary 29, 1968

February 2, 1968

February 6, 1968

March 8, 1968

March 22, 1968

March 31, 1968

April 4, 1968

LV Flight Sequence and EBN Functional Test

LV Power Transfer Test

LV Propellant Dispersion Test

LV OAT No, 2, Plugs Out

Erection of Spacecraft (5/C)

Swing Arm Compatibility Test

LY OAT No. 1, Plugs In (Waivered,

LV Combined Guidance and Control System Tests

LV OAT No. 2, Plugs Gut

LV Mission Control Center-Houston {NCC-H)Interface Test Vehicle Assembly Building (VAB)

Space Vehicle (S¥) OAT No. 1, Plugs In

SV GAT Nc, 2, Plugs Cut

SY Swing Arm OAT

Ordrarce Installatior

Transferred Launcher Umbilical Tower (LLT)/Vehicleto Pad

SY Flight Readiness Test (FRT) Completed

RP-1 Load

SV Countdown Demonstration Test {CDDT) Completed

Vehicle Launched on Schedule at 07:00:01 EST

303

b. The S-II stage, BP-1 transmitter (°CM/FM Link} fafled and had to bereplaced.

c. An electronics package in the $-[¢ stage aP-1 loading system requiredreplacement and re-calibration,

d. Tae $-I5 Ground Supsort Zquipment (GSE) stage power supply requiredreplacement.

3.4 PROPELLANT LOADING

3.4.17 RPT Loading

The RP-] system completed al] operations from CDOT through launch satis-factorily. There were no delays or questionable items during loading otherthan one minor anomaly, Ouring adjust level drain in CDDT, a time delayrelay malfunctioned and required replacement.

The level adjust operation in both COD" and the launch countdown ieft theflight mass within 0.01 percent of the intended vaiue, Kennedy SpaceCenter (KSC} mass readout indicated that 610,197.4 kijograms (1,345,255 Ibn}of RP-1 ware onboard the $-2C stage at ignition,

3.4.2 LOX Loading

The LOX system supported CDDT and launch satisfactorily. Minor problemsduring CDOT were corrected without impact on the launch. There were nohardware fa*lures nor propellant leaks during launch countdown. The LOXautomatic loading sequence was initiated at -6 hours 29 minutes with S-IVBslow fi] and terminated at -3 hours 7 minutes. Approximately 3418.2 m3(903,000 gal) of LOX were consumed during CDDT end launck countdown. Atlaunch, about 1703.4 m3 (450,000 gal) were onboard AS-502 and flight LOXmass was within specifications.

Two of the minor anomalies encountered during firal LOX loactng are asfollows:

a. During the S-IC finaT fast fi11 secuence, the speed cf the LOX trans-fer pump was manually edjusted raising the flowrate fram G.524 m3/s(8300 gpm) to a desired level of 0.590 mé/s (935C gpm) in crder toavoid a loeding delay. The need for this manual adjustment had beenanticipated and wes cavered in the released Toading procedure

b. Level shifts were experienced by the S-IC LOX and RP-1 recorders as-sociated with the automatic Propellant Tanking Computer System (PTCS)during LOX loading. The shifts had no effect on the operation of thePICS. The exact cause of the shifts has not yet been established,but investigation will continue,

3-4

3.4.3 LH2 Loading

The LHy system performed satisfactorily throughout CDOT and launch count~down. “The prelaunch fil) sequences were performed with no anomalies ordelays. Preconditioning of the S-IL stage was initiated at -6 hours 25minutes and was centinued through -3 hours successfully cooling down theS-II stage. Automatic loading was’ initiated at -2 hours 58 minutes withtransfer line chilidown and terminated at -1 hour 31 minutes,

Curing CODT approximately 1041.C m3 (275,CCC gal) of LH2 were consumed.This included losses from three S-I1 preconditioning runs, vehicle boi}-off,and drair yelume rot returned tc the storage tank. Launch countdown con-sumed 1748, m3 (462,000 gal) frem the LH2 facility. At launch Liz flightmass was withir specifications.

The followirg anomalies were noted:

a. Water entered the vehicle vent system at the burn pond after the re-plenish sequence was terminated during CDDT and launch countdown. Thisalso occurred during AS-501 loading operations and was attributed tosiphoning action through the standpipes initiated by rapid closing ofthe stage vent. The AS-502 problem may have been caused by insuffictenthelium purge resulting in cold piping, thus allowing the helium tocontract once the purge was terminated. This contraction could re-sult in lowering of the pressure and initiation of the siphoning action

b. After both CDDT and the launch countdown’, several LH2 vent bubble capswere cracked which exposed the vehicle vent system directly to theatmosphere. These cracks are attributed to localized overheatincoupled with rapid cooling by the splash*ng water in the pond

c. The debris vaive in the LH2 fill line was not closed until 12 seconds.This conpronises the integrity of the fil] line by raising the possi-bility that debris may enter the line during launch.

3.4.4 Auxiliary Propulsion System Propellant Loading

There were no problems encountered during propellant loading of the Auxil-

iary Propulsion System {APS}. Propellants consumed during loading wereas Follows:

a. Module 1

(1) Oxidizer System (Nitrogen Tetroxide, N24)

(a) Volume loaded 67,219.7 cm3 (4102 in.3) at 299.8°K (80°F).b) Volume off-Toaded 6096.0 cm3_ (372 in.3) at 297.0°K (75°F).cc) Yolume removed with bubble bleed during burp firing 453.3 cmd

(28 in.3) at 304.5°K (BB.5°F).(4) Yolume removed with bubble bleed during countdown 622.7 cmd

(38 in.3) at 302.6°K (85°F).

3-5

{2} Fuel System (Mono Methy] Hydrazine, MMH.(a) Volume loaded 67,219.7 cm3 (4102 in,3) av 300.6°K (81.5°F)(>) Volume off-loaded 1442.1 om3 (8B in.3} at 301.5°K (83°F).{c) Volume removed with bubble bleed during countdown 622.7° cm3

(38 in.3) at 302.0°% {B4°F)

b. Modute 2

(1) Oxidizer system (Nitrogen Tetroxide, N20¢(a) YoTume loaded 67,219.7 cmd (4102 in.3) at 300.9°K (82°F).fe Volume off-loaded 6096.0 om3 (372 in.3) at 300.9°K (82°F).c} Volume removed with bubble bleed durirg burp firing 344.1 ond(21 in.3) at 305.9% (91°F),({d) Yolume removed with bubble bleed curing countdown 497.6 om3

(30 in.3) at 302.0°K (84°F).

(2) Fuel System (Mono Methyl Hydrazine, MMH){a} Volume loaded 67,219.7 cm3 (41C2 in.3) at 299.3°K (79°F){b) Volume off-loaded 1442.7 cm3 (88 in.3) at 300-1°K (80.5°F).(c) Volume removed with bubble bleed during countdown 458.8 cmd

(28 in.3} at 300.9°K (82°F).

3.4.8 S-IC Stage Propellant Load

Initial propellant loads were obtained from the KSC weight and balance Taganc cempared with the continuous level sensor data. This comparison showedthe LOX load to be 1103 ktlograms (2432 ibn) greater, and the fuel load1259 kilograms (2777 lbm) Tess than the KSC loads. The propulsion perform-ance reconstruction utilizing an RPM match was able to follow the continuouslevel sensor data for both LOX and fuel with an accuracy of +1.27 centt~meters (+0.5 in.), The reconstruction also matched the residuals calcu-lated from level sensor and line pressure data indicating that the pro-pellant loads catculated from the level sensor data are accurate. Thereconstructed LOX load is 0.08 percent aoove <SC indicated values, andthe reconstructed fuel load is 0.20 percent below KSC indicated values.Both are well within the predicted three sigma limits of #0.5 percentTotal propellants onboard at ignition command are shown in Table 3-2.3.4.6 S-II Stage Propellant Load

The perseatage of flight mass onboard the S-II stage just before each tankpressurizing time was indicated by the PTCS to be 99.96 percent for Lliz and100.02 percent for LOX. Table 3-3 presents the S-II stage propellant loadat S-IC ignition command.

3.4.7 S-INB Stage Propellant Load

The PTCS indicated flight mass onboard the $-IVB stage just prior to in-dividuat tank pressurization was 100.08 percent for LH2 and 99.98 percent

3-6

Table 3-2. S-IC Stage Propellant Mass at Ignition Command

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asne cn LIC aersity of 1137.3 kg/m {7.2 Womft2) and Pe dansity of 862.5 Rae1 ewer FS} ane “hasea on 18 cessity of 7196.6 xy(70.97 Tom, BAT dentity af ach. «gin?

95 TheseKSC prope’ Tanz mess reagosts are sa*e as Toacirg table data at “gnitien.

Table 3-3, S-I1 Stage Propellant Mass at S-IC Ignition Command

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for LOX. Table 3-4 lists the S-IVB propellant Toad at S-IC ignition command. Simulation-trajectory match results were used in conjunction withthe listed data to determine the best estimate values. The best estimatePropellant masses are 0.6 percent hicher for LOX and 0.07 percent lowerfor LHg than the predicted values. These deviations werc well within therequired loading accuracy.

3.5 S-I1 INSULATION PURGE AND LEAK DETECTION

The S-II insulation purge and leak detection system performed effectivelyduring prelaunch operations. It was necessary to activate auxiliary backpurge in the sidewall insulation during terminal count; however, detailedinspection by operational television failed to identify any Teak in theexternal irsulation surface. Data recorded during this time indicated anirterconnecticn developed between the sidewall and feedline elbow flowcircuits in terminal count following the LHe #11} sequence. Reevacuationof the conmom bulkhead was accomplished at =1 hour 33 minutes without con-promise to the purge system.

3.6 GROUND SUPPORT EQUIPMENT

Ground systems performance was highly satisfactory, The swing arms, hold-down arms, tail service masts, propellant tanking systems, and all otnerground equipment functioned well in support of AS-502 Taunca, Tasle 3-5gives the start times for some of the pertinent ground/vchicle interfaceavents. There were relatively few anonalies, and launca damage was lightin most areas. Detailed information of ground equipment serformance,problems encountered during Taunch preparations, aad blast damage to thecomplex and equipment is given in #pol?o/Saturn ¥ Sround Systems Evalua-tion Report AS-592, Kennedy Space Center, May 1368.

Table 3-4. S-IVB Stage Propellant Mass at S-IC Ignition Command

| 3-8

Table 3-5. Ground/Vehicle Interface Events

RANGE TIMECVENT HR:MIN:SEC

Arm 4, 5-I1 Aft (Access) (Start Arm Retract (SAR]}

Arm 9, Egress Arm (SAR)

G-Ball Cover (Start Cover Retract)

Arm 1 S-I¢ Irtertank (SAR)

Arm Z, S-IC Forwarc (SAR)

Laurch Initiate Release SystemCommit Arm Liftoff Switches

Folcdowr: Arn Release, Primary (Preumatic)

Liftoff Switches (Position IIT-[¥}, Primary (1 inch) 00:00:00.55

Arm 8, CH/SM Arm {SAR}

Arm 7, U/S-IV8 Forward (SAR) 00:00:00.60

Tail Service Masts (3), Primary (Pneumatic) 00:00:00. 75

Arm 4, 5-I1 Intermediate (SAR) 60:00 :00.83

Arm 5, S-II Forward (SAR) — 00 :00::00.97

Ara 6, S-1¥B Aft (SAR) 00:00:06,92 Two problems associated with the S-II stage oriented pressurization andservicing system were as follows:

a. The system experienced an excessive loss of helium dering COLT andthe launch countdown. Replacomant of two relief valves, suspected assources of leakage, did not reduce the Toss of helium during launchcountdown. Troubleshooting of this system will continue.

b. S-I] engine start tank temperatures, although within required limitsat launch, were colder than expected. Helium used to condition thestart tanks is prechilled by the GSE LHz heet exckancer. An analysisaf the engine servicing system will be performed to isolate thisproblem.

3-9

Overall damage to the launch complex and support equipment was Tess thanoccurred at AS-S0T Taunch. Modifications incorporated to reduce blastdamage belcw that experienced on the previous launch were effective. Someconditions of damage revealed by post launch damage assessment are asfollows:

a, RP-] System, Tha mast cutoff valve in tai? service nast 1-2 openedat liftoff causing activation of the Ansel fire chemical systam inMobile Launcher (ML) Room 48. This resulted in the RP-1 distributorcabinet being filled with fire: extinguishing powder.

b. LOX System. Thirty one cables on Tevel 39 sustaiaed varying degree:of burn damage, The jacket of a section af vacuun-insulated pipingnear level zero was dented,

¢. Environmental Control system (ECS). Launch damage was approximatelythe same as was experienced during the launch of AS-501; however, theECS ducts were more extensively damaged. The horizontal ducts onlevel zero and the first 6 meters (approximately 20 ft) of verticaducting were comaletely destroyed. The second b-meter section ofducting also suffered extensive blast damage and the supporting struc:tare was broken loose and severely warped

dg. Holddown Arms, The holddown arm hoods were warped. Arm 3 hood waswarped extensively, Grouting between the holddown arm bases and theLUT deck prevented recurrence of the AS-501 flame damage to arm in-terior components.

Swing Arms. Damage was somewhat more widespread than for AS-501 launchbut “ewer major components were affected. Lower swing arms, particu-larly arm 1, sustained the greater damage. There were fires in thelower hinge areas of arms 4 and 6 resulting from Fycraulic oi] leakagethrough loosened "B* nuts.

f. S-IC stage oriented mechanical GSE. Storage racks on LUT levels 60, 100and 120 sustained varying amounts of engine exhaust damage. One rackwas completely destroyed but tre others are considered repairable.

SECTION 4

TRAJECTORY ANALYSIS

4.1 SUMMARY

Actual trajectory parameters of the AS-502 were close to naminal until thepremature shutdown of two engines on the S-IT stage. After this orematureshutdown, the trajectory deviated significantly fran tie ‘ominal throughoutthe remainder of the mission

Space-fixed velocity at S-I¢ Launch Vehicle Digital Computer (LYDC) sensedOutboard Engine Cutoff (ECO) was 7.28 m/s (23.89 ft/s) greater thannominal, At S-II LVOC sensed Engine Cutotf (ECO), comared to nominalcutoff, the snace-Fixed velocity was 102.36 m/s (335.82 ft/s) less thannominal and the altitude was 6.39 kilometers (3.45 1 mi) higher thannominal, At S-IVB velocity cutoff comngnd, compared to nominal cutoff,the space-fixed velocity was 48.98 m/s (163.58 Fl/s) greater than nominal;the cause of this overspeed is discussed in Section 10. The altitude atS-IVB velocity cutoff comand was 0.79 kilometers (0.42 9 ni) Tower thannominat and the surface range was 499.95 kilometers (269.15 n ni} longerthan nominal.

Parking orbit insertion conditions deviated considerably from nominal be-cause of anomaties that occurred during the powered portion of FlightThe space-fixed velocily al insertion was 48.16 m/s (158.00 ft/s) greaterthan nominal and the flight path angle (elevation of space-fixed velocityvector from local horizontal) as 9.378 degree less than nominal. Theseconditions produced an orbit which was quite elliptical with an eccentri-city 0.0138 greater than noninal. Tre resilting apogee of the parkingorbit was 171.54 kilometers (92.63 n mi) higher than nominal, and theperigee was 12.17 kilometers (8.57 n mi) less than nominal

The S-IC stage broke up at approximately 397 seconds at an altitude of28.9 kilometers (15.6 n ni) according to photographic coverage. At thistine the actual surface range and altitude as determined from a theoreti-cal free flight sirulation, were within 0.10 kilometers (0.05 nm mi) and1.42 kilometers (0.76 n mi), respectively, of nowinal. The free-flighttrajectory indicates S-II stage impact of 436.82 kilometers (235.86 n mifurther downrange than tae nominal impact point.

The S~IVB stage failed to reignite. Shortly after the attempted reignition,the spacecraft separated from the launch vehicle on ground command to thespacecraft, The $-I¥3 stage reentered due to orbital decay on April 26, 1968

41

4,2 TRACKING DATA UTILIZATION

4.2.1 Tracking Curing the Ascent Phase of Flight

Tracking data were obtained during the period from the time of firstmetion through parking arbit insertion.

Postflight trajectory for the initial portion of flight was estasliszedfrom a Teast squares curve fit of optical tracking data and was mergedwith a best estimate trajectory, The best estitate trajectory utilizedtelenetered guidance velocities as tne generating parameters to fit datafrom GLOTRAC Station I and five different C-Band radar trac<ing stations.These data points were fit througa guidance error model and constrainedto the insertion vector obtained fram the orbital solutinn. Comparisonof the best estimate trajectory with data from alt tae tracking systemsyielded reasonable agreement.

GLOTRAC Segment I date is a best estimate Fit of the data from the variousGLOTRAC sites. GLOTRAC Segment I provided data up to 480 seconds. Com-parisons between these data and the trajectory show maximum differencesof 300 meters (984 ft) in the vertical component, 20 meters (66 ft) inthe crossrange component, and 109 meters (323 ft} in the downrange com-ponent. The vertical coxponent is the least accurately determined bythe GLOTRAC system. These SLOTRAC data were received too late to beconsidered in the establishment of tne trajectory, but are helpful inascertaining the validity of tre trajectory. The GLOIRAC Segment I datawere the only precision tracking data available after 230 seconds, Com-parisons with the GLOTRAC Station I and Offset Frequency Doppler {QDOP)data show deviations which are considerably less than those obtained fromthe GLOTRAC Segrent I data.

4.2.2 Tracking During Orbital Flight

Table 4-1 presents a summary of the C-Band radar stations furnishingdata for use in determining the orbital trajectory. There were alsoconsiderable S-Band tracking data available during the orbital flightwhich were aot used in determining the orbital trajectory due to theabundance of C-Band radar data.

The orbital trajectory was obtained by taking the insertion conditionsand integrating them forward at the desired time intervals. The insertionconditions, as determined by the Orbital Correczion Program (OCP), wereobtained by a differential correction procedure which adjusted theestimated insertion conditions to fit the C-Band radar tracking datain accordance with the wefghts assigned to the data. After all theC-Band radar tracking data were analyzed, sone stations and passes wereeliminated completely from use in the determination of the insertionconditions.

4-2

Taste 4-1. Summary of AS-502 Orbital C-Band Tracking Stations

star ve or mous sev | vere ava] ave] mvs] seve neve nsereion ship car x

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‘Bermudameigia Conary tsTand

Tracked Spacecraft After 5-116/C5H Separation

Final GLOTRAC Segment I data were received for the interval where theattemoted S-IVB stage reignition occurred. These data were received tooJate to be considered in the establishment of the trajectory; however,comparisons of the trajectory with these data help to indicate thevalidity of the trajectory. After the GLOTRAC data becane reliable (about11,590 seconds), the maximum deviations were about 270 meters (886 ft)in the vertical component, 175 meters (574 ft) tn the crossrangecomponent, and 125 meters (410 ft) in the downrange component

4.3 TRAJECTORY EVALUATION

4.3.1 Ascent Trajectory

Actual and nominal altitude, surface range, and cross range for theascent phase are presented in Figure 4-1. The actual and nominal totalearth-fixed velocities, and the elevation angles (elevation of earth-Fixed velocity vector from the local horizontal) of the velocity vectorsare shown in Figure 4-2. Actual and nominal space-fixed velocity andFlight path angie during ascent are shown in Figure 4-3. Comparisonsof total inertial accelerations are shown in Figure 4-4.

The combined burn tine of the S-IC, 5-11, and S-IV3 first burn was87.61 seconds longer than nominal; the 5-IC burned 0.85 seconds longerthan nominal, the S-II burned 57.8% seconds longer than nominal, and theS-IVB burn was 28.95 seconds longer than nominal. The abnormally longS-II bun was the result of the premature cutoff of the two engines,

4-3

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Figure 4-7. Ascent Trajectory Position Conparison

causing the S-IVB_to turn longer in attempting to arrive at the properend conditions. The accuracy of the trajectory at S-IVB cutoff isestimated to be 21.0 1/s (2.3 ft/s) in velocity, and + 500 meters (1,640ft} in altitude.

Mach nunber and dynamic pressure are shown in Figure 4-5, These para-heters were calculated using measured meteorological data to an altitudeof 50.2 kiloneters (27.1 n mi). Above this altitude the measured datawere merged into the U.S, Standard Reference Atmosphere.

Comparisons of the actual S-If engine No. 2 premature cutoff conditions,with their corresponding nominal conditions, are shown in Table 4-2.Actual and nominal values of parameters at significant trajectory eventtimes, cutoff events, and separation events are shown in Tables 4-3,4-4, and 4-5, respectively. The heading angle ig the azimuth of thespace-fixed velocity measured east of north.

UntiT the S-IT engine premature shutdown, the altitude was slightly Towerthan nominal, the surface range was close tp the nominal, and the total‘inertial acceleration was less than nominal.

The theoretical free-flight trajectory simulation data for the discardedS-IC and S-II stages were based on initial conditions obtained from theFinal postflight trajectory at seoaration. The simulation was based

14

%) S000 Teo| \ re

we, 146,4150: 8000: l i : oe ats

1ac eto Sst,70 7000 {|_| _| 2

ELEVATION ANGLE “

634 = son0 /

uf 0) 3 5000 \

= 3 [ EARTH-FIXED VELOCITYBq M1 gS 400

| 30} £3000 \” a

20 2000 7

10 1000 ~

a 0 Vv ¥: pow | 0 100 200 «300 «400 500 600 700800

RANGE TIME, SECONDSFigure 4-2. Ascent Trajectory Earth-Fixed Velocity Comparison

upon the separation impulses for both stages and nominal tumbling dragcoefficients due to lack of tracking data for both stages. Photographiccoverage of the $-IC stage indicated that the stage broke up at 397 secondswithir 0.10 kilometer (0.05 n mi) surface range, and 1.42 kilometers(0,76 n wi) altitude of nominal. Table 4-3 presents the significant param-eters and their deviations fror nominal; including Max Q, maximum accelera~tich, apexes of spent stages, and maximum earth-fixed velocities. A sum-mary of impact times and locations for the S-IC and S-II stages is pre-sertec in Table 4-6. Since there was no tracking or photographic coverageof the discarded S-II stage, its impact was simiated as noted above.

Spacecraft separation was initiated on graund command to the spacecraftafter it was ascertained that the S-IVE stage had failed to reignite. Thetrajectory conditions at spacecraft separation are presented in Table 4-5.

2278000

5 A284 7000: i

Lil PATH ANSLE ||24} 6000 r

g 27 B 8000 | \ SPACE FINED VELOCITY

Z 16 5 4900 \218 \ seta 4= 5 vovcon= 124 & 3000:

Z a4 £ e000 | }

44 1000 ! _

3 PRIS

|_|

“4 yv W ¥. q o wo 200 300 #00 500 sg0C«CtO

RANGE TIME, SECONDSFigure 4-3, Ascent Trajectory Space-Fixed Velocity Comparison

As a result of the S-IVE failure to reignite for second burn, the S-IVBStage renained in orbit after spacecraft separation instead of flying thehigh apogee (lunar distance) orbit planned.

The S-IVE stage reentered on April 26, in the ocean between the eastcoast of Africa and the west coast of India.

4-6

ACCELERATION,

m/s2

40

38

30

28

18

10

WILY,

VY 100

Figure 4-4,

200 «300 400,500RANGE TIME, SECONDS

600 700-800

Ascent Trajectory Acceleration Comparison

4.0: 16

3.5 14 chDYNAH=C PRESSURE -—4+/ \ _ ACTUAL

NOwTNAL

3.0) 12 f

—

2

MACH

NUMBER

S

[4

DYNAMIC

PRESSURE,

H/er®

yf WACH NUNBER—

[|tt —

0 20 40 60 eo 1001204080RANGE TIMZ, S=CONDS

Fégure 4-5. Dynamic Pressure and Mach Number Versus Range Time

4.3.2 Orbital Trajectory

The acceleration due to venting during parking orbit is presented inFigure 4-6. These accelerations were obtained by differentiating theteleretered guidance velocity data and removing accelerometer biasesand the predicted effects of drag.

A family of values for the insertion parameters was obtained dependingupon the combination of data used and the weigats applied to the data.The solutions that were considered reasonable nad a spread of about£500 meters (1,640 ft) in position components and +1,0 m/s (3.3 ft/s) invelocity components. The actual and nominal parking orbit insertionparameters are presented in Table 4-7.The ground track of the first two revolutions in parking orbit for theAS-502 vehicie is given in Figure 4-7.

4-8

Table 4-2. S-II Engine No. 2 Premature Cutoff Conditions

PARAMETER ACTUAL NOMINAL ACT-NOM

Ranae Time, sec 412.9Altitude, ke 177.26 177.80 0.54

(n wi) (95.71) (96.00) (-0.29)Surface Range, kn 933.27 935.43 -3.76

(n ri) (503.93) (905.63) (-1.70)

Space-£*xed Velocity, m/s 5,183.74 5,183.31 -30.17(ft/s) |(16,906.63) |(77,005.61) (-98.98)

Flight Path angle, deg 1.611 1.647 -0.036Heading Angle, deq 78.706 78.607 0.099Cross Range, km 8.87 1.41 -2.54

(n ni) (4.79) (6.16) (-1.37)Cross Range Yelocity, r/s 92.30 82.86 9.48

(ft/s) (302.82) (271.85) (30.97) GUEDAICE COMPUTER (OUT

7 2000 00 008 00) va000 yen20RANGE TIPE, SECANS

a09:30:03 01:00:00 91;30:00 02:00:00 2:30:00 02:00

RINSE TIME, HOLS f#EVTES: SECONDSFigure 4-6. AS-502 Acceleration Due to Venting

4-9

Table 4-3. Comparison of Significant Trajectory Events

eve ecu ovenFine orl Tonge Trey see 7 zap aeFated Preediah ace as? O39 izeLo. a mln?) 3eT earee Tne, sec 3.5PEititage, to hiefate 5 i.eerinZymic Pressure fange Tee, see’ 5.20dmnasie Presture, wees, | Esewis | eee]paride, Ae ioeoe) baalWeiieae Tae arae Te. sec Taetnertins ReteTeration iecrterachin sd 8ase, ansiieh

Str Prange the, seeRecbleratieng mis :ie isre [Range Tine, see

Rca erations "793tres

Farge Tine secpitt suden ko

Ter STC Sage

umes}SeIt Stage Range Time, see 0.61steetuse, tn 50.72net} (197.83)Surface Range, im 1767.65tien iss.ig}Braap oF STE Stage ~ lige Fine, se “Rr ySiettuten ta a aas) aygae | te.78hSuetaee Ringe mete 36ns sinsei ines:Teale Carti-FAnaT Tange’ Tse sc esTeloeltys foie vetaeitye ave zee tersh roce'ést| —s7ten.'S}st lange Tine, see 317.08 33 97vehocttsy ars ees] 2aat.30ete) t20,fbe,¥2)| (2° 130.72)

serve darge Tine, sec 57.04 569.25ehoeity ers wines) reds 51.24tae faaskebdey ces,aa) r1e8.i8)

Table 4-4, Comparison of Cutoff Events

sovua—Y nomtvar ect-wcn accent apeet acr-norPARAMETER SHIC (ECD (SOLENOTO ACTIVATION] S-1¢ ofcO (U¥OC stWsED)

leange Tine, sec wae Tay On van] 147.56 0-85Jaitieuce, er 36.07 55.00 0.47} 5.8 58.33 1

(n'a) (o0.26)| (30-02)

|

(0.26) 4 (2.10) (31.50)

|

(0.605,lsuetace Range, kn 78.38 75.58) -0.29 23.18 81.82 134

(rs) (sore) cancaty] 0.77) cat.do) et) (0.92)Iseace-Fixed Yeloctty, ays zoxc.74| 2636.54] -17,0| 2752.56] 2745.28 228

Urevey

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(e6se.€3)] ¢-58.20}

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(9090.92) (9006.82)

|

(23.693lFiignt path Angle, deg 20.185 | 20,251] -0,196 19.667 19.984 ~0,177]

lneadiang Angie, deg 75.139 75.582] -0.481 76.495

|

2.430cross Range, km 0.12 ese! 0.71 0.65

n'a} (-0.38) (u.52)) 50.38) 0.8)lcross ange Velocity, as =3.05 WaT) 23.28 18.17 §

ts) | (10,98) 6586.39) | (66.34)S-1E £60 {L¥OC SENSED)

Range Tine, see 376.38 317.65] sh.6a var.ca] 659.26] 87.79Iaitieude, kn 195.05 188.70) 6.38 won} astso] 0.7

(ni) (95.38) (ores)| (3.88) |(102.38)) (909.400 |(0.42surface Range, km veio.62 1500.43] 3i0,19 294a.c2)—_zaea.s7 apa. as

(ew) cortées (a0. | certs} se6-11)] 1314-86) (268.45)space-Fised Velecity, a/s 6725.67, 6828.03 ~l0z,36| 7038.85] 77ea.91| aa. ad

tates) (2,068.88) (22,401.67) (395.82) ¢25,727.25)| (25,560.72) (160.56)lFitgnt Patn angle, deg 1.600 o.7ee] cara 0.400) 0.001389Ineadieg Angle, deg 83.388 sico7] 1.78; 90.257].cross Range, ke 30.60 23.56) 7.08 10.52 sear Bas

inn) cvseb2t (12.32)] (2.80) | (8.28) 28.33) 8.85)crass Range Yetocityy fs Vea} 166.77] 18,79 206.81| 256.92 39.28

tet| ewoseint oneted] gorctin eo78.si] (e485) 030.84)

@ Finsy REVOLUT:ONJ) SECOND REVOLUTION

pg

y Tas{ SEPARATIONBg 40 4 $

Zon aKi

;°

s 4

60Yoo 30 ea 20 8 20 A GO AD loo ico 140 160 160 169 Tee 122 100 HD

e$$$Lon 7098 | legFigure 4-7. AS-802 Ground Track

43

Tatle 4-5. Comparison of Separation Events

PARAMETER ACTUAL I NOMINAL ACT-4OM,

S-IC/S-11 SEPARATION COMMAND.

Range Tire, sec 149.08 148.26 0.82Altitude, km 60.08 59.05 1.03

(n mi) (32.44) (31,88) (0.56)Surface Range, kur 84.65 83.51 Va

(n mi} (45.77) (45.08) (0.62)Space-Fixed Velocity, m/s 2,765.14 2,755.03 10.11

(ft/s) (9,071.98) (9,038.87) (33.17)Flight Path Angle, deg 19.530 19.725 | -0.195Heading Anale, deg 74.996 75.491 -0,495Cross Range, km -0.13 0.66 -0.79

(nmi) (-0.07) (0.36) (-2.43}Cross Range Velocity, m/s -3.81 18.26 -22.97

(ft/s} (-12.50) (59.91) (72.41)

feodetic Latitude, deq North 28. 843 28,833 | o.o10Longitude, dea East -79.780 -79,.788 3.008

S-II/S-1VB SEPARATION COMMAND

Range Time, sec 577.08 | 518.49 $8.59Altitude, km 198.25 188.78 6.47

(n mi} (405.43) (107.93) (3.50)Surface Range, kn 1,815.52 1,505.90 309.62

(n wi} (989.33) (813.12) (167.18)Space-Fixed Velocity, m/s 6,728.65 6,834.40 =108.79

(ft/s) (22,075.62) |(22,422.70) (347.08)Flight Path Angle, deq 1.597 0.778 0.819Heading Angle, deg 83.416 81,639 Lvm7Cross Range, ki 30.74 23.70 7.08

{noni} (18,60) (12.80) (3.80)Cross Range Velocity, n/s 172,83 187.33 15.50

(ft/s) ¢ (967.03) (516.17) (50.86)Geodetic Latitude, deg North 32,144 31.747 0.397

Longitude, deq East -62.136 -65.377 | 3.241

Table 4-5. Comparison of Separation Events (Cont

S+1V8/CSM PHYSICAL SEPARATION

PARAMETER ACTUAL KOMIBAL ACT-ROM

Range Time, sec 11,867.82 11,908.09 ~240.27Altitude, ke 196.21 661.31 -465.10

in of) , (105.94) (457.08) (-251.14)Space-Fixed Velocity, m/s 7,846.32 10,574.07 -2,727.75

(ft/s) (25,742.52) (34,691.83) (-8,.949.31Flight Path Angle, deq -0.281 14.360 14.641

Heading Angle, deq 96.240 195.139 -18,699Geodetic Latitude, deq Nurth 7.993 21.612 10.58Longitude, deq Cast 85.117 -45.680 -39.437

Table 4-6. Stage Imnact Location

PARAMETER ACTUAL NOHINAL ACT-NOM

S-IC STAGE IMPACT

Range Time, sec 528.93 507.89 21.13Surface Range, ke 635.38 638.08 =2.68

(n mi) (343.98) (344.52) (-1.44)Cross 2ange, kr 3.79 10.19 6.40

(nmi) (2.08) (5.59) (-3,45)Geodetic Latitude, deq N 30.207 30.157 9.059Lonaitude, dey E -74.314 -74.272 -1.042

S-IT STAGE IMPACT

Range Tire, sec 1,251.24 1,151.52 99.72Surface Ranae, km 4,648.43, 4,211.51 436.82

{nai} (2,809.95) (2,274.03) (235.86Cross Range, kin 152.16 139.43 21.67

(n mi) (82.16) (79.46) (11,79)

Geodetic Latitude, deq North 31.205 31.859 -9,685Longitude, deq East ~ 32.182 = 36,722 4,540

Table 4-7. Parking Orbit Insertion Conditions

PARAMETER ACTUAL NOMINAL

Range Time, sec 757.04 669,26 87,78Space-Fixed Velocity, m/s 7,842.09 } 7,793.93 48.16

(ft/s) (25,728.64) (25,570.64) (788.00)Flight Path Angle, deg -0.377 0.001 -0.378inclination, deg 32.567 32.561 0.006Eceentricity 0.0741 0.9903 0.0138Apogee*, km 360.10 188.56 171.8e

(n mi) (194.44) 1 (901.81) ($2.63)Perigee*, km 173.18 185.32 “12.17

(n mi) (93.493 (100.06)Altitude, kr 190.19 381.51

(n mi) (102.69) (103.41)Period, min 29.84 88.25 1.61Geodetic Lat! tude, deg North 22,730 | 32,683 0.077Longitude, deq East ~49, 388 54.709 5.321 *Based on a spherical earth of radius 6,378,165 km (3.443.934 0 ni).

4-14

SECTION 5

$-1C PROPULSION

5.1 SUNMARY

The S-IC propulsion system consists of the engines, oxidizer system, fuesystem, pneumatic contrel pressure system, and the camera ejection andpurge system. Five F-1 engines provide the thrust to propel the Saturn ¥jaunch vehicle during first stage boost. The F-] engine is a single-start6,770,198 Hlewton (1,322,000 IbF} Fixed-thrust, bipropetlant. rocket. systenusing liquid oxygen as the oxidizer, and RP-1 as fuel, turbopump bearingcoolant, and contral systen fluid, Liquid oxygex is stored in @ cylindri-cal tank having a capacity of 1342 m3 Tay aoe ft2), allowing for a usableoxidizer supply of 1,489,960 kilograms (3,284,000 ibm}. RP=1 {kerosene}fuel is stored in a tank having a capacity of 827 m3 (29,221 ft3}, ai low-ing for a usable fuel supply of 646,823 kilograms {1,426,000 bn}Pressurized Gaseous Nitrogen (Gtp) is used as a source of pneunaticpressure for propellant system yalve actuation and engine purging. Duringflight, GNis used to purge the film camera and televis‘on canera Tensesand to eject the Film cameras

S-IC propulsion systems performed satisfactorily. In general, all perform-ance flight data as determined from the propulsion reconstruction analysisfell close to the nominal predictions. Average engine thrust reduced tostandard sea level conditions from 35 to 38 seconds was 0.20 percent lowerthan predicted. Average reduced specific inpulse was 0.10 percent lowerthan predicted, and reduced propellant consumption rate was 0.07 percentTess than predicted.

The vehicle first longitudinal structural mode frequency coupted withthe engine response to the oxidizer suction lines resonant frequency withinthe 110 to 140 second periad. This resulted in a vehicle longitudinaoscillation termed "POGO"

Inboard Engine Cutoff (IECO) {solenoid activation signal) occurred 0.1seconds later than predicted. Outboard Engines Cutoff (0LCO) occurred0.85 second later than predicted, primarily due to lower than predictedaverage fuel flowrate. An intentional fuel level cutoff of the outboardengines was planned and attained, demonstrating the adequacy of thiscutoff mode, An inboard engine LOX level cutoff was planned and attained,demonstrating the adequacy of this cutoff mode

5-1

Tae usable LOX residual at OECO was 11,873 kilograms (25,725 ibm} of LOXcompared to the usable 2544 kilograms (18,837 Tbmj predicted and theusable fuel residual at OECO was zero, as predicted. The higher thanexpected LOX residual was primarily due to a slightly higher than expectedloading mixture ratio.

All the subsystems except the camera ejection system and the controlpressure system performed as expected. The camera ejection systemejected only one of the four film cameras. It appears that the systempheumatic supply pressure tubing failed during S-IC/S-11 separation.The con2rol pressure system performed satisfactorily during powered flight.After separation, however, the sphere pressure decayed unexpectedly. Thisdecay may be due to a fai‘ure of the pneumatic lines to solencid valvesthat control the LOX vent and relief valves. the planned correction forboth of these problems on S-IC-3 and subsequent stages will be the sub-stitution of stainiess steel tubing for the alumirum tubing that was usedon S-IC+1 and S-IC-2.

5.2 S-IC IGNITION TRANSIENT PERFORMANCE

The fuel pump inlet preignition pressure and temperature were 30.4 W/cm?(44.1 psia) and 274°K (33°F), respectively. These fuel purp inlet condi-tions were within the F-1 engine mode? specification limits (start boxrequirements) as shown in Figure 5-1. The preignition temperature at thefuel pump inlet was corsiderably jewer than the fuel bulk temperature of294°k (70°F). Similarly, fhe LEX pump inlet preignition pressure andtemperature were 56.4 Nyem? (86.4 psia) and 96°(-287°F), respectively,The LOX pump inlet conditions were also within the F-1 engine modelspecification Timits as shawn in Figure 5-1. The fyel and LOK ui lagepressures were 15.1 N/cmé (27.7 psia) and 17.2 N/emé (24.9 psia),respectively, at ignition,

The engine startup sequence was normal, A 1-2-2 start was planned andattained. Engine position starting order was 5, 1-3, 2-4. Two enginesare considereg to Start together if their combustion chamber pressuresreach 69 N/cm® (100 psig) in a 100-millisecond time period. Figure 5-2shows the thrust buildup of each engine indicative of the successfulJ-2-2 start. The major events during engine startup sequence are listedin Table 5-1.

The best estimate of propellants consumed during tie oeriod between ignitionand hotddown arms release were 38,901 kilograms (85,/65 Ibm) as comparedto 38,923 kilograms (85,810 Ibm) by the reconstruction analysis. Theseconsumptions are more than tho predicted consumption of 38,846 kilograms(85,643 Ibm). The more than predicted holddown consumption resulted inbest estimate liftoff propellant loads of 1,398,599 kilograms (3,083,382Tbm) for LOX and 600,604 kilograms (1,324,104 Ibm) for fuel.

5-2

2. Paw LET peso acu

q Maes ts1 “i

hes

e r 4 - ty qei I :

- + 4 4 pon

Terasur [tf fomit ii sO _ a

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25 © a) w a 2vate resoise, tea

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apeELCTED STARING 2mmeen | . 40 g

i if t ¢- ume

ey

& ca P ia We hoLox Pupp TSLET eRESSUR Zen?

Figure 5-1, S-IC Start Box Requirements

: + a Ss‘ engine 4 L iF

|

Feta — f Fusing 5 +F020

ENGINE 2

Table 5-1, S-IC Stage Engine Startup Event Times RANGE TIME, SECONDS

EVENT engine tT evetne 2 [ENGINE 3

|

enaine 4

|

ENGINE ©Start Solenoid Energized

|

-5.949

|

5.979

|

3.739 489

|

5.959MLV 1 Starts Open -5.7e7

|

-§lait -5.587 1381 “5.811MLY 2 Starts Onan 5.793

|

-5'827

|

“81579 85

|

$8997

Thrust Chamber Ignition -2.600 =2.430 -2.760 no -3.020

NEV 1 Starts Open 320659 22.281 -2/617 265 ~2:88§MEV 2 Starts Doen -2.681 -2.289 253 “Bea?Final Thrust OK 693

|

11287 2313 -1.985AT] Engines: Running Vast Launch Commas O15 NOTE: Times taken from data sampled 500 times per second.

Bea

5.3 S-TC MAIN STAGE PERFORMANCE

The F-1 engine has a single bellshaped thrust chamber with an expansioarea ratio of 10:1. The thrust chanber is cooled regeneratively by fuewhich passes through tubes thet form the thrust chamber wall. A doublewa‘ led extension nozzle, utilizing turbopump turbine exhaust gases farinner wall coolant, is used to increase the expansion area ratio from10:1 to :6:1. The propellants are supplied to the thrust chamber by adirect drive turbopunp driven by exhaust gases from a gas generator.

Two analytical techniques were employed in evaluating S-IC stace propulsionsysten performance. The primary method, propulsion reconstructionanalysis, utilized telemetered engine and stage data to compute longitudinal thrust, specific impulse, and stage mass flowrate. Ir the secondmethod, Flight Simulation, a six-degree-of-freedom trajectory simulatiowas utilized to ft propulsion reconstruction analysis results to thetrajectory. Using a differential correction procedure, this simulationdetermined adjustnents to the reconstruction analysis of thrust and messflow histor’es to yield a simulated trajectory which closely matched theobserved postflight trajectory. $-IC stage propulsion performance, asdetermined by reconstruction was completely satisfactory.

Performance parameters compared weli with the nominal predictions over theentire flight as shown ‘n Figure 5-3.

Average engine thrust, reduced to standard sea level conditions, at a 35 to38 second time slice was 0.20 percent lower than predicted, as shown inTable 5-2, Individual eng'ne deviation from predicted thrust rangedfrom 0.86 percent lower (engine No. 2) to 0.33 percent higher (engine No4). Average engine specif'c ‘mpuise was 0.10 percent lower than predicted.Individual engine deviations fron predicted specific impulse ranged from0.30 percent lower (engine No. 2} to 0.04 percent higher (engines No. 7and 4).

Reduced to sea level ambient pressure, the stage average longi tudinathrust for the Flight from propulsion reconstruction was 0.53 percentJower taan predicted, wii ui. ..j. w.-,dge longitudinal specific imputseas reconstructed was 0.17 percent higher than predicted

Flight simulation showed that the stage average specific impulse was 0.68percent greater than predicted. The flight simulation results were usedin an attempt to explain the time and velocity deviations at OECO. Toexplain the velocity deviation, an error analysis was made to determinetne contributing parameters and the magnitude a* the velocity deviat’orcaused oy each of these parameters. Table 5-3 lists the various errorcontrioutars and the cutoff velocity and time deviations assoc‘ated with

eacn.

fof) Ug goog fe og

2HEFEi

3 a4 5

e a z ae

er ee ee 3

2

az e

ig ae

aS a ey@ @sf 3ag 2 ie

be s

ye ee e

S-IC

Steady

State

Operation

Figure

5-3.

Table 5-2, S-I¢ Engine Performance Deviations

PARAHETER evgive preoic~en reconstruction| pevraricn avzRAGEANALYSIS Percent CEvIATION

PERCENT

1 e738 (1518) szgr (sia) 9.26Thrust 2 6752 (1518) 6695 (1808) -0.86

3 3 3 678e (1525) 6766 (1521) -0.26 20.2610? N (10? IF) 4 6699 41504) ane (150 | 9-33

5 6761 (1520) 6730 (183) -0.88

Specific Impuise 1 [2576 (262.7) 2577 (26a a)) 9.082 |2598 (26479) 2590 (264.1)

Nes/kg_ (1bF-5/Tom) 3 [ees (284-2) 2585 (263.7) -0.104 [2580 (263-1) 258) (263.2)5 42690 (264-1) |2588 (263.9)

1 2016 (5765) 2set (5779) 0.23Total Flowrate 2 2598 (5728) deae (S702) -0-45

3 zene (5772) 2616 (5768) -0.97 0.07ka/s (ibys) 4a 2593 (5716) 2600 (5733) 0.30

5 2610 (5755) 2601 (6738) |-0.36

1 2.27 227] 9.0Mixture Ratio 2 2.26 228] 9.00

3 2:24 2.23] 0.45 -0.002Lox/Fuel 4 2.27 2.28] 0.48

5 2.27 227] 0.0

NOTE: Analysis was reduced to standard sea level conditions (standard pumpinlet conditions} at liftoff plus 35 to 38 seconds

Table 5-4 presents a summary of the flight sinulation results, reduced tsea level ambient pressure conditions, on the average values and deviationsOF longitudinal thrust, propellant flowrate, and vehicle ‘ongitudinaspecific impulse.

The vehicle first longitudinal structural node frequency coupted withthe engine response to the oxidizer suction I‘nes resonant frequencywithin the 110 to 140 second period. The S-IC stage engines exper‘encedchamber pressure oscillations building up to a maximum at approximately125 seconds of 5.5 to 6.9 N/emé (8 to 10 psid} peak-to-peak. The "POG"phenomenon is discussed in detail in paragraph 9.2.3.1

5.4 S-IC ENGINE SHUTDOWN TRANSIENT PERFORMANCE

Inboard engine cutoff was initiated by LOX Tevel indication and occurredat 144.72 seconds (solenoid activation signal). Outboard engine cutoffwas initiated by fuel level and occurred at approximately 148.41 seconds(start of Time Base 3 [T3]). This was 0.85 seconds later than thepredicted time of 147,56 seconds, The late OECO was primarily caused bylower than predicted average fuel flowrate.

Table 5-3, S-I€ Velocity and Time Deviation Analysis at OECO(Simulation versus Predicted)

VELOCITY DEVIATION (EARTH-FIXED)

DEV. (AC7-PRED]CONTRIBUTING ERROR FACTORS WY bn/s}

(0.16 Percent) -10.63

Liftoff Weight Increase.........

Total Propellant Flowrate Decrease., (0.68 Percent} -20,82Axial Force Coefficient Difference 6.02Meteorological Data Difference - 8,92Late 1&C0 2.93Late OEco 38.68

Total Contribution 7.56Observed 7,73Difference (Observed - Total Contribution) O17

TIME DEVIATION

DEV. (ACT-PRED)CONTRIBUTING ERROR FACTORS at {eee}

Tnittal Fuel Undertoading......0s.0..s (2369 kg} = 0.60Fuel Flowrate Decrease 1.88Late 1£C0 - 0.08Residual Differences..........ee-ceeeeaee (30 kg) - 0.0)

Total ContributionObservedDifference (Observed - Total Contribution)

Thrust decay of the F-1 engines is shown in Figure 5-4, The decaytransient was normal. The oscillations which occur near the end of“taitoff" are characteristic of the engine shutdown sequence.

The total stage impulse from OECO to separation was indicated by engineanalysis to be greater than predicted, Telemetered guidance data alsoindicated the cutoff impulse was greater than expected, as shown inTable 5-5. These deviations are within the acceptable range consideringthe difference between the actual and predicted vehicle mass.

5-8

Table 5-4. Comparison of S-IC Stage Flight Recanstruction DataWith Trajectory Simulation Results

EDETRIETN SATINj “oemtien guia dewcaconnewer ass erneten neenersuc seutation Fensetnieren

mierape « a waar] mmeana MmeTongitudinal tarust nee TAFAR ROB. T7538. C4538 7,799,187. booee

Yenicte mace at note-] 1g zszee.grec as7e 08.0 2,784,226.9 .down “gem wetease (ey 12135 802.0 8 6,18 TBP. 0.468

srorage ass was razz] weran ‘ 13.78.86Clomés) 23,880.23 Bz 0.68 25,281.48 68

peerage” vesrg 2508.25 . 29y2-09snecific impulse (er 268.54 Oe 266.35

THRTaRSTT PORIEEE foe evel aDTenE PPETEIE

20T ae! Team+ Tan

TRBane 1-——==) TarENGINE 2 - — Mea) 16imme 3 venANGIE & 8.47 3

2 “aLeID ACTEATION *i let

4 #3 PREOLCTSD ENVELOPE oe §

oa

° ae 5

0 02 8 08 08 1.0 12 1

TIME ROM CUTOFF SICHAL, SFCQNDS

Figure 5-4, S-IC Engine Shutdown Transient Performance

5-9

Table 5-5. S-IC Cutoff Inpulse

PERCENT DEVIATIONFLIGHT FROM PREDICTED

PARAMETER PREDICTED ENGINE. GUID. DATA ENGINE. GUID. DATA

Cutoff N-s 8,975 026 10,765 582 9,141,331imputse (Ibf-s)

|

2,017¢686

|

21277;351

|

2057053

|

19-84 13Velocity m/s 30.89 13.14 VAIncrease (ft/s) 35.73 ay 36.65

|

70-66 | 2.87 5.5 S-IC STAGE PROPELLANT MANASEMENT

S-IC stage uses an open oop method for achieving Propellant Utilization(PU). The propellants loaded were 2123 kilograms (4681 Ibm) greater thanpredicted for LOX and 2369 kitograns (5224 Ibm) less than predicted forfuel. This loading resulted in the destred propellant level cutoffsignals. Since the S-1C stage uses an open loop method for achievingpropellant utilization, the usable propellant residual deviations are theresult of propellant loading and perfarmance prediction inaccuraciesA summary of the propellants remaining at major event t'mes is presentedin Table 6-8 and tne residuals are presented in Table 5-7. An inboardengine LOX level cutoff was planned and attained, denonstrating theadequacy of this cutoff mode. An intentional fue? leve: cutoff of theoutboard engines was planned and attained, demonstrating the adequacyof this cutoff mode.

5.6 S-IC PRESSURIZATION SYSTEMS

5.6.1 S-I0 Fuel Pressurization Systen

Tae fuel pressurization system maintains sufficient fuel tank uliagepressure to meet the minimum Net Positive Suction Pressure (NPSP) require-ments of the engine fuel turbopump during engine start and flight. Inaddition, this system helps provide fuel tank structural cepability bykeeping a positive pressure head at ali points instde the tank. Thefuel tank is provected from overpressurization with a pressure reliefsystem design which requires a double failure mode to occur to exceedthe tank design pressure, Before engine ignition the fuel tark ipressurized with heliun from a ground source. During flictt, the tank ipressurized with gaseous helium obtained by using the F-1 engine heatexchanger to heat helium which is supplied from storage bottles locatedin the LOX tank, The helium pressurization system satisfactorilymaintained the required ullage pressure in the fuel tank during flight.The HeTium Flow Controi Valves {HFCV) opened as programmed and the fifthflow contro] valve was not required. In Section Z, Event Time Tables,these valves are designated "Fuel Pressurizing Valves.” The heatexchangers performed as expected

Table 5-6.

LHe

S-IC Stage Propellant Mass History

Woe BL

De Fa waster

feano og

<a

Hai aobey st

Ce rae

n

% ni 1

ss.477122,306Pebeysan2,92east

1

sae!

rae |Ghd

Mastsanes

agains

3,023 382

aie

Tales da ht ivaChapere a3 Tess sanior anterevgurlseti gas (SDN $0 2hey WITT

Table 5-7. S$-IC Residuals at Outboard Engine Cutoff

PROPELLANTS PREDICTED ACTUAL DEVIATION

LOX RESIDUALS *

Usable Mains tage kg 544s 11,673 3129(ton)

|

18,837 25,735 6898Tarust Decay and kg 14,978 14,978 0Unusablew (ibm)

|

33,020 33,020 o

FUEL RESIDUALS

Usable Mainstage kg 0 Q 0{1b} a a 6

Thrust Decay and kg 11,202 11,305 103Unusable (bm)

}

243698 24,924 226

* Does not include GOX pressurization gas.** LOK bias,

Includes 150 kiTograms (330 1bm) in LOX interconnect linesand 14,826 kitograms (32,690 Ibm) in LOX suction ducts.

6-17

The low flow prepressurization system was commanded on at -96.87 secondsand performed satisfactorily, providing ullage pressure as shown in Figure33.

The fuel high flow prepressurization supply valve of the Ground SupportEquipment (GSE) was opened at -1.34 seconds and maintained the ullagepressure within the band, At 0,82 seconds the to. 1 HFCY of the onboardpressurization system was opened. The flow overlap between the onboardand the prepressurization systems seen on AS-501 did not occur for thisflight. HFCY No. T was signalled to open by umbilical disconnect énsteadof launch commit, eliminating the flow overlap between the two systems,HFC¥'s No. 2, 3, and 4 were commanded open by the switch selector withinacceptable times as shown in Table 2-3. These flows held the ullagepressure within the operating band as shown in Figure 5-6, The fifth HFCYwas not required to operate since ullage pressure was maintained abovethe fifth HFCV switch actuation pressure. Helium bottle pressure as shownin Figure 5-7 stayed within expected limits. The heat exchangers performedwithin the expected performance Vimizs,

ze : 7H

2 ToPREPRESS SHETCH SAND] 2

~ 4]é AS-S01 FLIGHT 9 == 4 .3 5 3-602 FLIGHT OATA 75 7 ——T g

= VY START Low FLOM PREPRESS, -96.87 &ar W LOW FLiy REPRESS VALVE CLOSE “ON”, -61,78 22 1 HicH FLOs PREPRESS SUPPLY Val 203 OPEN MOK tes :g ZF caunen comm, c.12 a= az SIC UMBILICAL OISCOMNEET, 0.74 4 eHFCY NO. T OPEN "ORM, 0.82

| : 1610 te

y

& Y 4 1%120 ~180 “80 ~60 =a -20 0 20

RANGE TIME, SECORDSNOTE: ALL TIMES REFER “D AS~SU2 DATA

TIME ASE FOR AS-501 AND AS-502 DATA 1S NOT THE SAME

Figure §-5. S-1¢ Fuel Ullage Pressure During Countdown

Vee wo. 2 o0ry, ste

3 é

z rom BARD| =

g Pepe 3

wla

a: pone

To. + 2SIM. axeo®SON | :

SLAPS. ‘AS-592 an DATAY a2 al HN ~. \4| 22 ium. fee AD WIN] MU sh tw. san &

: st ga ses super garak=—tf | :: a wun2 a0. =f = =

StL:+

° .

ange 11m. SECONDSFigure 5-7, S-IC Helium Bottle Pressure for Fuel Pressurization

5-13

5.6.2 S-IC LOX Pressurization System

The LOX pressurization system provices and maintains sufficient LOX tankullage pressure to meet the minimum NPSP requirements of the LOX turbo-punp during engine start and flight, In addition, the pressure provide:additional LOX tank structural capability by keeping a positive pressurehead at all points inside of the tank, ‘This system also protects the LOXtank from overpressurization. Before engine ignition, the oxidizer tankis pressurized with helium from a ground source. During flight, thetank pressurization is accomplished with Gaseous Oxygen (60x) obtained byusing F-1 engine heat exchangers te convert oxygen from liguid to gasTre LOX pressurization syster performed satisfactorily and al] performancerequirements were met. The ground prepressurization system maintainedullage pressure within acceptable Imits until launch commit. The onboardpressurization system subsequently maintained utlage pressure within theGOX Flow Control Yalve (GFCV) band during the flight. The heat exchanger:performed as expected.

The prepressurization system was initiated by opening of the ground supplyvalve at -66.65 seconds. The ullage pressure increased until it enteredthe switch band zone which resulted in terminating the flow at approxi—mately -58.84 seconds. The ullage pressure increased approximately1,34 N/em@ (1,95 psi) above the prepressurization switch setting to18.78 N/em2 (27.2 psia). This overshoot is similar to that seen on AS-SU1

The LOX tank ullage pressure history is shown in Figure 5-8. Duringflight, the ullage pressure was maintained within required Timits by taeGFCV throughout the flight and followed the anticipated trend. The SFCYreached full open at 120 seconds and remained open until tha end ofFlight. The maximum GOX flowrate during ful] open aosition of the valvewas 25.85 kg/s (57 lbm/s). After IECO, the 30X flow requirements for theremaining four engines increased until SECO.

5.7 S-IC PNCUMATIC CONTROL PRESSURE SYSTIM

The pneumatic control pressure system uses pressurized GN2 as a source ofpneumatic pressure for propellant system valve actuation and enginepurging. GNz is supplied by a groud source to the stage Gp fill systemand to individual ground controlled, stage pneumatic valves duringstage systen test, checkout, static firing, and prelaunch operationslate in the prelaunch operation, the stage GN> system fs charged +oflight storage pressure. The pneumatic control pressure system on theS-IC stage performed satisfactorily during the 148 second flight. Theactual pneumatic control regulator outlet pressure measured 521 N/cmé(758 psia) as shown in Figure 5-9, The contro] pressure system succeededin actuating the orevalves after engine cutoff. “All inscrumented pre-valves indicated closed positions.

5-14

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2 st

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461 4

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|

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=

el

Ss |— yPa z

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ANGE TOME, SECOHDS

Figure 5-8. S-IC LOX Ullage Pressure

The control sphere pressure decreased to 1793 V/cm? (2690 psia) duringprevalve actuation and then leveled off for about 2 seconds as shownin Figure 5-10, At this time, approximately 157 seconds, the spherepressure started decaying rapidly. It is suspected that the pneumaticVines to the solenoid valves that control the LOX tank vent and reliefvalves burst during the severe enviroments imposed by $-IC/S-ITseparation. A thermal analysis of the preumatic lines was conductedassuming twice the design separation environment. ata from AS-591 andAS-502 flights indicate that the separation environment may be of thismagnitude.” The thermal study indicated that tre 6061-T6 aluminum tubingwould reach a temperature of 739°K (870°F) 3 seconds after separation.With a 517 N/cm@ (759 psia) pressure in the lines, the ultimate stressof the lines is exceeded at 544°K (700°F). Flow analysis indicated thata broken line would have a maximum flowrate of 0.087 kg/s (0.147 lom/s},while the average flowrate out of the sphere was 0.086 kg/s (0.19 Ibm/s},The difference between these two values is attributed to fuel tank ventvalve cycling and other system demands.

ECP 441 has been approved and the 6061-T6 aluminum lines to the solenoidvalves that control the LOX tan< vent and relief valves will be replacedwith stainless steel lines for AS-303 and subsequent vehicles. Thiswill eliminate the possiaility of line rupture due to high temperatures.

8.8 S-IC PURSE SYSTEM

The turbopump LOX seal, gas generator actuator housing, and radiationcalorimeter purge systems performed satisfactorily during the 148 secondflight. The LOX dome and Gas Generator (G3) LOX injector purge systemalso met al7 requirements.

5.9 S-IC CAMERA EJECTLON AND PURGE SYSTEM

The camara ejection and purge system utilizes GNz to perform its funct‘onduring flight. The GN ts Supplied by a ground source to the onboardsystem, During Flight Gp from the system's storage sphere is used topurge tne separation viewing cameras lenses, to eject the two separationviewing cameras,and to eject two LOX viewing cameras, A schemat’c of thecamera ejection system is shown in Figure 5-11,

Tne system ejected only one of the four film cameras. “he canere framerate measurement for separation camera No. 1 went to zero at 174,25seconds, indicating ejection of that camera. Frame rates for the remain-ing three cameras did not change, indicating that there was neitherejection nor sufficient motion of the capsules within the ejection tubesto disconnect the electrical plug.

A study of possible system failure modes was conducted to determine themost probable cause of failure. These analyses indicate the mostprobable reason for failure to eject three of the four cameras wasinadequate bottle pressure at ejection command due to failure of thepurge system Hine very near the purge systen soleno‘d valve (seeFigure 3-11).

900

oes L. Trustee 4sauce vaste Palit Sekt asopeg muuesf‘ot

rot

e ws

: + —oe 60

aan, LE AND STF

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: 1 aH becswecl | i iv‘ee rnne, sis

Figure 5-9. S-IC Pneumatic Control Requlator Gutlet Pressure

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1400: bat: ~sl5 Manor“0 7 7 wo oo wo eTRANGE TIME , SECONDS

Figure 5-10. §-IC Control Sphere Pressure

5-17

Pos Lut,

FAO UMHEL FfFLL SOLENOID YALYE5PRESSURE. rPANSOUCER

BLEED ORIFICEWIDE WUGLE LG CAMERA,

PRESSURERELIEForice

FING

~~esecron shia veEGECT OW AND. PURRESip STORAGE SPHERE

‘NUE LINE : SEPARAYION CAMERA 1

TIN A

JECTION

FORWARD COMPARTMENT LOOKING AFTFigure 5-11. S-IC Camera Ejection System

No flight measurement of bottle pressure was made. Ground measurementsshowed bottle pressure to be 1688 N/cm@ (2450 psia) just prior toliftoff. Figure $-12 shows the minimum predicted sphere pressure fornormal operation compared to the predicted pressure assuming failure ofthe purge line.

If the system operates normally, bottle pressure at ejection command isapproximately twice that required for camera ejection as demonstratedby ground ejection tests. However, with a purge line failure, bottlepressure at ejection command would be below the required pressure forejection,

The following improvements to the camera ejection system have beenproposed:

a. The aluminum purge system lines in tre vicinity of the suspectedfat lure will be replaced with stainless steel lines.

b. The lines in this same area will be redesicned and provided with‘mproved supports to reduce the probability of vibration darage.

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1S syoh - - 3z =]

olf ip i wy| ivi i |.fe ts i saNot To

Figure 5-12. S-IC Predicted Camera Ejection and Purge Sphere Pressure

&. Cabling to the thrusters which open the camera doors will haveadditional insulation instalied as protection from S-I] ullage motorheating.

a, Flow balancing orifices in the purge system will be relocated closerLo the purge solenoid vaive to prevent excessive bottle blowdownin the event of purge line breakage downstream of the orifices

e. Orifices will be added to the ejection lines. In the event that oneof the ejection lines fails, suffictent pressure will be maintainedin the second I*ne for ejection of two of the cameras.

219/5-20

SECTION 6

5-11 PROPULSION

6.1 SUMMARY

Tae 5-11 oropulsion system consists of five stngle-start liquidbivropellant J-2 engines, Liquid Oxygen {1.0%} and Liquid Hydrogen (LH2}propellant systems, propellants management, pressurization, pneumaticcontrol pressure, recirculation and camera ejection systems, The f'veengines are functionally independent but are clustered and controlledto form an integrated mainstage propulsion systen for the $-IT stage.

The S-IE propulsion system performed satisfactori*y during the first 169seconds of operation following Engine Start Command (ESC). Engine thrust,as determined by reconstruction analysis and telemetered propulsionmeasurements at 60 seconds after ESC, was only @.43 percent below pre-diction, Total propellant flowrate was 0.53 percent below and specificimpulse 0.08 percent above predéctions at this time slice.

At 319 seconds a sudden performance shift vas exhibited on engine Na. 2with thrust decreas’ng approximately 33,806 Newtons (7600 Ibf). Theengine continved performance al the reduced level until 412.3 seconds.By 4:2.92 seconds the dropout of thrust OK switches indicated engineNo. 2 cutoff, and at.414,18 seconds engine fo. 3 also cut off. Postflightevaluat‘on of telemetered date led to the conclusion that the ergineNo. 2 Augmented Spark Igniter (ASI} fuel line failec ard ultimatelycaused failure of the engine. Since the flight, testing at MarshalSpace Flight Center (H5FC} and the engine manufacturer's facility hassubstantiated this conclusion. The testing reveals that an axidizerrich mixture, caused by a fuel leak, creates very high temperatures andrapidly erodes the injector. Because of tris erasien the LOX dore ofengine No. 2 eventuelly failed, opening the LOX high pressure system andcausing Engine Cutoff (ECQ). A madification of the ASI propellant feed-lines (both fuel and LOX) and their installation is being accomplished.

Interchanged Lt.! prevalve control wiring connections between enginesRo. 2 and 3 celenoids caused the premature cutoff of engine Ko. 3.When engine Ko. 2 cutoff the LOX prevalve on engine No. 3 was commandedclosed. Ar individual checkout of the prevalve wiring during prevalvetiming checks is planned for future vehicles.

6-1

S-II burn time, from engine ignition (Start Tank Discharge Valve [STDVopen) to ECO (start of Time Base 4 [Tg]) was 425.31 seconds which 457.81 seconds longer than predicted. The extended burn tine was causeby the premature cutoff of engines No. 2 and 3, Loss of the two encinesreduced propellants consumption approximately ¢0 percent and requireda longer burn time to reach propel lants-depletionThe propellants management subsystem met all perfcrmance requirenentsHowever , the Propellant Utilization (PU) mixture ratio step, as sensedby the Launch Yehicle Digital Computer (L¥DC), occurred 77.23 secondslater than predicted after ESC because of the two engines out conditionPropellant Toadirg was .48 percent above the predicted. Residual propellants remainirg in tanks at ECO were 3412 kilograms (7523 Ibm) con-pared to the prediction of 3264 kilograms (7195 Ibm), with no 0.5 secondtime delay incorporated, The discrepancy in residuals was caused byJiguid level measurement errors that developed from "tilted" liquidlevel surfaces after engines No. 2 and 3 cutoff,The performances of the LCX and LHg tank pressurization systems weresotisfactory. The premature loss of twa engines supplying GOX throughheat exchargers to the LOX tank caused the ullage pressure to decreasebelow the regulator band late in the flight. LOX pump fnlet NetPositive Suction Pressure (NPSP), however, was more than adequatethroughout the flight.

The engine servicing system operated satisfactorily with the exceptionthat the engine start tanks were chilled more than expected but withinrequired limits. The exact cause of these low temperatures is notknown, However, detailed analyses are being conducted of tre GroundSupport Equipment (GSE) LHz heat exchanger and engine vurge and loadingoperations in order to isolate the problem.Both LHz and LOX recirculation systems performed satisfactorily andmet the required engine pump inlet and/or discharge conditfons at ESCHowever, there were soma deviations of the Liz oump inlet temperature:and the LOX pump discharge temperatures pricr ta ESC. Potentialsysten changes being considered at this time for AS-503 include improve-ments to the LH2 system insulation and increased helium flow from theLOX helium injection system.6,2 S-II €HTLLDOWN AND BUILDUP TRANSIENT PERFORMANCE

During the S-IC boost phase the LOX and Liz recirculation subsystems,shown in Figure 6-1, chill the ducts, turbopunps and ather engineComponents. Prior to engine start the recirculation systens areshut down. This opens the LH prevalves, shuts down the LH, recirculationPumps and stops LOX helium (He) injection, Engine start signal is thenreceived by the engine electrical controller which causes the propellantvalves to open in the proper sequence, The contro’ler also energizespark plugs in the Gas Generator (66) and thrust chamber, ignites the

6-2

RECZROULATION "Hh9 FACES!

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Figure 6-1, S-II LHp and LOX Recirculation Systems

6-3

propellant and also causes Gaseous Hydrogen (GHz) to be released fronthe start tank. Giz provides the initial drive for the turbopumps which —-deliver propeTiant to the G6 and the engine. After propellant ignition,gas generator output accelerates the turbopumps and engine thrust in-creases to main stage operation at which time the spark plugs de-energizeand the engine is in steady-state operation.

The engine servicing and recirculaticn operations were performedsatisfactorily although there were sore deviations. Thrust chambertemperatures, as shawr on Figure 6-2, Tie near the cold edge of theprediction band and are approximately 5.6°K (10°F) colder than forAS-501, Engine servicing procedures and the GSE LHp heat exchanger arebeing evaluated to determine the cause for the low temperatures.

340WLvoc s-11 ESC, 149.76ENG NO. 1J ERG NO. 2 |. t00DENG NO. 3

300 Pesprepreteo UPPER| © ENG NO. aLevel BENG NO. 5

F 50

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THRUST

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TEMPERATURE,

°F

L-200

140

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100 4be - 300

60 WI~0g 360 ~~SCASCSC RD ° 120

RANGE TIVE, SECONDSFigure 6-2, S-II Thrust Chamber Temperature

6-4

Thrust chamber warmup rates during the S-I€ boost ranged from 8.7 to12.5°K/min (15.6 to 2?.4°F/min) and closely paralleled the predictedrate. These warmup rates were nearly the sane as those experfenced onAS-501, confirming the conclusion that AS-501 results should be con-sidered "normal". The AS-501 report recownendation to accommodate thewarmup rate was satisfied and the maximum allowable engine starttemperature was increased fron 161 to 172°K {-170 to -159°F). Thrustchamber temperatures at engine start ranged fron 125 to 138°K (-235 to-210°F), well within the 82 to 172°K (-300 to -150°F} requirement.

Roth temperature and pressure conditions of the J-2 engine start tank:were within the required prelaunch and engine start boxes. The rangeaf data points were near the cold temparature and high pressure sideof the boxes as shown in Figure 6-3. Shilldown temperatures rangedfrom 90 to 105°K (-298 to -271°F}, (lower than predicted), and analysesare being conducted to determine the reason for increased chilling.Start tank pressures, 898 to 815 N/cm® (171 to 1180 psia), were lowerthan for AS-501, as planned. The lower pressures were intended toincrease tank temperatures but the increased chilldown offset thisplanned increase.

STAI TANK “ESPERATURE, *F-306 230 200 “150

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PRESSURIZATION ACTUALS £226 SEC) .750 eae

b

roCr

START TAVK TEMPERATURE,“

Figure 6-3. $-I1 Engine Start Tank Performance

6-5

Prelaunch start tank warmup rates ranged from 16.1 to 19.0 K/cw2/min{23.4 to 27.5 psi/min) and 1.1 to 1.9°</min (2.0 to 3.8 Egmin) Thewarmup rates during the S-IC boost were 10.0 to 14,8 N/cmé/min (14.5 to21,4 psi/min) and 0.9 to 1.6°K/min (1.6 to 2.6°F/min). These rates arehigher than experienced on AS-501 and are the result of very cold initialstart tank temperatures. Engine No, @ start tank indicated a considerablylower warmup rate and it is suspected to have had a small leak. DuringS-II engine operation, the start tanks refilled as designed, However, atapproximately 413.5 seconds engine No, 2 start tank pressure started todecay probably caused by the failure of the start tank refill line.

A change to the start tank servicing system is being considered whichwoutd increase the launch facility start tank vent system Tine sizefrom 2.54 to 3.81 centimeters (1 to 1.5 in.) from the stage umbilicato the main vent line, This would increase start tank temperaturesapproximately 11.1°K (20°F)

All engine helium tank pressures were within the required prelaunch andengine start limits. At -276 seconds, engine No. § helium tank pressuredecreased below tha redline of 2000 N/em@ (2900 psia) to a minirun of1985 N/cm2 (2884 psia), At -240 seconds, the pressure increased to2035 N/cm@ (2950 psia); this was caused by erratic operation of theGSE pressurization regulator,

Helium tank temperature during prelaunch chilldown ranged fror’ 103 to0°K {-274 to -263°T), approximately 7°K (12.6°F) lower than thoseexhibited on AS-501, Engine servicing procedures and the GSE LH2 heatexchanger are being evaluated to determine the cause of these Towtemperatures.

during tae S-IC boost the heliym tank pressure increase due to warmuprates was aooroximately 4 N/cme {3.8 pst) higher than that for AS-501;the rates were 21.9 to 39.2 N/cmé/min (31.8 to 56.7 psi/nin}.

The Liz and LOX recirculation systems performed satisfactor’ly. At S-IESC the predicted engine pump inlet conditions were obtained as shown onFigure 6-4. However, during prelaunch operations difficulties wereexperienced in maintaining engines Ho. 3, 4 and 5 purip discharge temperaturesbelow the launch redlines. As shown in Figure 6-5 the LOX pump d’schargetemperatures of engines No. 3 and 4 and the LH2 punp discharge temperatureon engines No. 3 and 5 were above the prediction bands most of the Limefrom -210 seconds unti? just before ESC at 149.76 seconds. These samedifficulties were experienced during the Countdown Derionsiration Test(CDOT) at which time it was demonstrated thar these temperature:decreased substantially after tank pressurization. Accordingly, theCDDT prelaunch maximum temperature redlfne for engines LOX pumpsdischarge was increased From 96.5 to 97.6°K (-286 to -264°F} and thetime was changed from between ~100 and -50 seconds to @ constraint at-22 seconds, Similarly, the maximum LH7 pump discharge temperature redline,between -52 and -30 seconds, was changed to a constraint at -30 secondsand increased from 21.8 to 22.3°K {-420.5 to -419.5°F] for engine No. 5

6-6

U2PUN

INLET

TEMPERATURE.

°KLOX

PUMP

INLE

TTEMPERATURC,

°K

25

24

a

20

wi

99

7

95

93

a1

a

LH PUMP INLET PRESSURE, psia

28 32 36 40 “41 i t L 1

QENG HO. 1 SATURATION CURV fr -a15ENS NO. 2 URE WL,aENG NO. 3

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FiLii p~lo-5 ..fhPREDICTED ESC aa=f ENVELOPE

L -423

START REQUIREMENTS ENVELOPE f- -425@ 2 2 2 2 2 30 32

Lz PUMP INLET PRESSURE, N/cw?

LOX PUMP INLET PRESSURE, psia3 36 40 4a 48

ete te:SATURATION CURVE L. -2e0i

START ~ -282REQUIREMENTS,ENVELOPE P7284

b -286L -288

FLIGHT DATA ~ ALL [- ~#90‘5 ENGINES 292

} -294popePREDICTED ESCERI ewvevore ~296

[ -298

2 2 oF OOLOX PUMP INLET PRESSURE, N/cm2

Figure 6-4. S-II Start Box Requirements

6-7

LHPUMP

INLE

TTEMPERATURE,

°FLOX

PUMP

INLET

TEMPERATURE,

°F

SATURATION

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DISZAAKGE TEND AAoo

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BESS kee TEMP fs -180 “120 -60 Q 60 120

RANGE TIME, SECONDS

Figure 6-5, $-11 Lz and LOX Recirculation System Performance

6-8

providing that a cooling trend prevails. Engines No. 1 through 4 red~jane limits were maintained at 21.8°K (-420.5°F) maximum but changed toa constraint at -22 seconds. Recaimended changes to the LOX recircuta-tion system are presented in paragraph 6.9. Changes under considerationfor the LH2 recirculation system are:

a. Insulate all bare flanges of the LH2 return manifolds.

b. Provide additional insulation cn J-2 engine carpanents.

c. Retrofit improved evacuatior valves cn the LH2 feed andvecirculatior vacuum jacketed lines.

The thrust buildup profile of each J-2 engine is shown. in Figure 6-€.All engine performance parameters were within the predicted thrustbuilcup envelope. Majer engine start event times are summarized inTable 6-1. The S-I] stage engine start was commanded by the LOC at149.76 seconds and the engine buildup trensient conmenced at 151.02seconds (average) when the STDV opened. S-1I mairstage, average time forengines to reach 90 percent thrust, occurred at 153.08 seconds, 2.06seconds after STDV. The engine thrust levels stabilized between870,000 and 900,000 Newtons (195,500 and 202,200 Tbf,; prior to PU systemactivation at 155.26 seconds.

Table 6-1, S$-II Engine Start Sequence Events

TIME OF EVENT IN RANGE TIME, SECONDS

EVENT ENGINE 1] ENGINE 2] ENGINE 3

|

ENGINE 4) ENGINE 5

Engine Start Conmand

|

149.76

|

149.76

|

149.76

|

149.76

|

149.76

Mainstage Control 150,98

|

150.98

|

150.98

|

150.98

|

150.98Solenoid

Start Tank Discharge

|

151.020

|

151.026

|

151.01)

|

181.011

|

181.020Valve Gpen

90 Percent Thrust 153.57

|

152.93

|

183.03

|

153.02

|

152.86 Main LOX Valve Open

|

153.778

}

153.778

|

183.778

|

153.778

|

153.737

Nainstage OK 153.79

|

183.74

|

183.74

|

153.74

|

153.71 6.3 S-11 MAIN STAGE PERFORMANCE

Each of the five J-2 engines is a nigh performance, 1,000,850 Newtons(225,000 Inf) thrust rated engine, using LOX and Lig at a mixture ratioof 5.5:1 (can vary to as low as 4.5 for the desired propellant utiliza~tion at stage cutoff). It features a tubular wall, bell-shaped thrustchamber (27.511 expansion ratio), and two independently driven, direct-drive turbopumps powered in series by a single gas generator. Afterthe initial start transient the engines operate at a high mixture ratio

6-9

1.2

UvOC S11 ESC, 149.7611 ENG STBV OPEN (AVG), 151.02TT ENG 90 PERCENT THRUST {CALCULATLD AVG), 153.08ENG NO. 1 (32057) L.24

— ENG ND. 2 (92088) |

1-97" ENG No. 3 (32058)= -— Ens wo. 4 (320895~ ENG NO. § (J2041) --20| “,0.8 7 i

THRUST BUILDUP f- 08ENVELOPE 4if

0.2f— f

|

ot

: [J

i.FATa ENGINE ~}

00 1 2 3 4 5 6TIME FROM ESC, SECONDS

t Vu vy 1 i 1 L ‘149) 750) 751 152 153 154 155 156RANGE TIME, SECONDS

Figure 6-6. $-I7 Engine Thrust Buildup

until Engine Mixture Ratio (EMR) shift. At this time the EMR is changedto the Reference Mixture Ratio (RMR) until ECO,

Tho analytical techniques were employed in evaluating $-11 stagePropulsion system performance. The primary method, propulsion recon-struction analysis, utilized telemetered engine and stage date toCompute iongitudinal thrust, specific impulse, and stage mass flowrate.The second method, flight simulation, utilized a six-degree-offreedomtrajectory simulation to adjust propulsion reconstruction analysisresults to fit the observed trasectory, Adjustments to the reconstruc-tion analysts of thrust and mass flow histories were deterrined using a

6-10

differential correction procedure whick yiciced a simulated trajectoryclosely matching the observed postflight trajectory.

Average engine thrust, as cetermineé from reconstruction analysis, at atime slice of ESC +60 secords was 0.43 percent belcw credicted as shaunin Table 6-2, Individual engine deviaticrs from predicted thrust rargefrom 1,36 percent lower (engire Ko. 2) to G.28 percent higher (engineNo. 3), Average engine specific imputse was 6.08 percent greater thanpredicted. Indivicual engine deviatiors from predicted specific impulseranged from 0,02 percent fower (engine No, 4) tc 0.17 percent greater(engine No. 3) than precicted. Predicted averace performance character.istics of the S-II stace engine system are presented in Table 6-3 forthe total flight and for the righ ard reference mixture ratie burn timesTotal flight averages ave considerably below predicticns because cf thepremature cutoff of engines No. 2 and 3. Durire the high mixture ratioburn, however, average thrust and mass loss rate were only ¢.25 and0.33 percent below the predicted, respectively.

Table 6-2. S-II Engine Performance Deviations (ESC + 60 Seconds)

Taner nn TaEier STRUCTION] FROST SEVIAT-Oara oa ects sBUsTy geTrae T tTSTE z

§ rnpacific 1 Tae? aes Tass (a1reste 2b mage ae) eign Hee 3nesite (geet rivalPTs) nM

: 6.2Fowrate T Be 8 (eTaus ines: 2 bela deat

3 fa2'2 fem ati Ha"? ‘eisc0fi bae'3 {e35153

winder T soRecis i PesoLON Fue : fe

s | 51SOE Pralysis -4 ax (50 yrus & seconds L

Flight simulation showed the stage time averaged specific impulse duringthe high mixture ratio step operation to be C.09 percent greater thanpredicted, in good agreemert with the recenstruction analysis. Detailedresults ere presented in Table 6-3.

The S-11 stage propulsion system perfornarce is presented in Figure 6-7.A shift from normal performance occurred at approximately 319 secondsThe performance change is evidenced by a thrust decrease of 33,806 Newtens{7600 bf}. This has been attributed to an ASI fuel line leak on engineNo. 2, At approximately 413 seconds a large step cecrease in staceperformance was evidenced by @ reduction of stage thrust to 3,002,536Newtons (675,000 Tbf} and a change of propellants flowrate from 1213 to730.3 kg/s (2675 to 1610 Ibm/s}. This abrupt change in performanceapproximately 40 percent, was caused by the shutdown of engines No. 2

6-11

6-12

Table

6-3.

S-II

Flight

Reconstruction

Comparison

With

Simulati

onTrajectory

Matc

hRe

sult

s

PARSHE

TERS.

PRFITED

LYSI

SPE

RCEN

TDEW

,FRO

RED.

Beat

sRa

ntTEFPRENCE E

HGR

MCATURE

|wnatRE

iTew

rio.

|

aaGG

eomes)

(3C4

REESENEE

irTata

waaTURe

FLpar

ia.

|

avenue,

Tetera

FLcait

ht

Aver

age

Lang

’tua91

Stag

eTh

rust

Aver

age

Venicle

Yass

Loss

Rete

Average

Stage

Longtt

udinal

‘Speci

ficTmpulse

¥ te

bas

(rbm/s)

Kessks

(bes

e)

5,0375%

32,4

89)

6,219,

923,

1(988

68998.8

(emer

2200

.8

306,50

2,902,799

(679,038;

78.6

eg wa}

276.6

080.570)

|52.

6(eset)

*

(2666.7)

alma

|aro129,673)

1298.7

(495

.3)

923.9)

(419.03

724882

05,22a)

47:8, 5

66|“3.25

Tea

(e25,ans

geg

9.93

13,68

saa]

oar

|

3.25

Vez

VW

PaRa

wETE

RSwars

MURCONT

PERCENT

CEW.=e

PRET

‘

rn

MURT

LRE

RATIO.

5ENGINES]

ry

csTU

RERATIO

(3cwa

uncs)

Tura

rieFLIGeT

misrine

|MeROSC

RATIO

(2ENGINES)

aver

age

Lang

tudi

rat

Stag

eThrust

Aver

age

Yeki

cve

Mass

Lose

kateAverage

Stage

fi 041,282,

153,

333)

22.3

(2678.7)

4158.4

3,003,500

(675,207)

87

2A76

202

car

isee,718}

ecra

£.03

(2200.6)

4140.6

c.06

(422.2)

+Ti

necuration

isme

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at

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+jvarage

prio

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afarproximate'y

433

secones

rang

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lewonbey SaPLLa Et Vir TIES |

S-I]

Steady

State

Operation

Figure

6-7.

and 3 and is discussed in subsequent paragraphs. Stage thrust droppedbelow the predicted vatue for a two engines out condition after enginesHo. 2 and 3 cutoff. Approximately 33 percent of this performance losswas caused by a reduction of axial acceleration that reduced engine LOXinlet pressure but the remaining effect has not been determined et thistime. Analyses by the engine and stage manufacturers are in progress todetermine whether data was in error or if an engine/stage perfornancedeviation occurred. Current information indicates that 21] pressurefor engines No. 4 and 5 dropped excessively due to an instrumentationbus change observed in a 47 ampere surge at the time of engines flo. 2and 3 cutoff. Because of the low consumption rete of pranellants aftercutoff of the two engines the Prograumed Mixture Ratio (PNR) shift vasdelayed 77.23 seconds and was not sensed until 490,76 seconds, Tre tweengine out condition a'so extended burn time 57.81 seconds to 425.3seconds with ECO, init‘ated by LOX depletion, sensed by the LYDC at576,33 seconds.

A chronological list of events that arc believed to Fave led to thefailure of engine No. 2 are discussed briefly in Table €-4. Postflightdata analysis led to the conclusion that the ASL fuel line, shown inFigure 6-8, had cracked at spproximatcly 228 secerds and continued toleak progressively until 319 seconds. Since the flight, testing atMarshall Space Flight Center (NSFC) ard the engine xanufacturer!facility has substantiated this theory. The testing simulated an ASfuel line failure by reducing tre fuel supply, creating an ASI 9.5 LOX/Lg mixture ratio, The high mixture ratio produced abnormally hightemperatures in the main injectcr which caused severe erosion to occur,as evidenced in Figure 6-9. Further indications of the ASI fuel leakare reflected by local decreases of engine compartment, engine instru-mentation packages, anc engine ho. 2 hydraulic actuator return fluidtemperatures (discussed in paracraph 17.2) and evidence of cryogenicimpingement on the No. 2 yaw actuator pressure transducer, as discussedin paragraph 8.3.

The first incicaticr of abnormal behavior was a gradual decrease incramber pressure ef engine No. 2 starting at 260 seconds. The rate ofdecay was approximately 0.07 N/om2/s (0.10 psi/s}. At 318.9 secondthere was a sudden change of engine No. 2 performance. At this time,fuel flowrate started to increase and was followed ty a general encineperformance decrease at 319 seconds as shown in Figure 6-10. Theirerease in fuel flowrate is probably due to thrust charber tube damageresulting from injector debris produced by the high ASI mixture ratioogeraticn preceding this time, This damage resulted in the side loadsshewn by the actuator sP. MSFC testing demonstrated this type of tusebundle darage. Main thrust chamber pressure decreased 15.8 Y/ome(22.9 psi) with proportionate reductions of pump discharge, maininjection and GG injection pressures. During this performance shifthowever, fuel injection and turbine inlet tesperatures abnormallyincreased indicating that a fuel Teak had developed betueen the flow-meter outlet and the engine fuel manifold discharge. Following theperformance shift, engine No. 2 stabilized and continued oneration a

Table 6-4, S-II Engine No, 2 Performance Shift and Cutoff

RANGE TIME, SECONDS EVENT AND/OR PREVAILING CONDITION

225 Cryogenic leak in engine compartment indicatedengine compartment and engine to, 2 yawactuator return fluid temperatures began todecrease (see paragraphs 17. and 8.3)

260 to 318 Chamber pressure decays at rate of 0.07 Nrem2/q(0.10 psi/s}.

280 Cryogenic impingement on pressure trarsducerbelieved ta have occurred, Engine No. 2 yawactuetor ¢P transducer indicatec ar abrormalpressure ramp (see paragraph 8.3)

318.9 An increase in Lz flowrate occurred,

319 Engine No. 2 Lil flowrate ircreased and engineperformance decreasec; a tube bundle leak isbelieved to have beer. the cause,

412.3 Ergire Ho. 2 chamber and fuel injectionpressures began a gradual decrease

412.6 Ergine compartment heating spike occurs {seeparagraph 17.3)

412.7 LOX come failed; the LOX high pressure systemopened and the engine performance decayedrapidly.

42.9 Mainstage OK pressure switches deactivate whenLOX injection pressure decayed below switchsetting and initiated ECO sequence.

the reduced level until 412.3 secords. Inmediately after this tire, a:indicated in Figure 6-11, the LOX pump discharge pressure bagan to decrease,followed by 2 rapid increase of LOX flowrate at 412.6 seconds, indicatingthat the LOX high pressure system had failed. It has been concluded thatthe LOX dome feiled from erosion caused by a hat oxidizer-rich ASI mixtureratio, Failure of the LHz high pressure system was also indicated at412.6 seconds by z large Flowrate increase, Engine failure was definitelyevidenced by the rapid decay of thrust chamber pressure, main LOX andLig injection pressures, and propellant pump discharge pressures at412.7 seconds. Nainstage thrust OK pressure switches opened andinitiated an engine shutdown sequence at 412.92 seconds, These anomalies

6-15

HIYDRAUL IZ. ACTUATOR{T1P.2° PLACES )

ENGINE LH) INLET SUPPLY DUCT

GAS GENERATOR (us)PU VALVE. Lh GG BLEED VALVE

LOK IZGH PRESSURE: Ly HIGH PRESSURE oUCTuctHEAT EXCHANGER TARDINE BYPASS DUCT

OxIDLZER TURBINEBYPASS YALEF (OTE¥)FLIGHT INSTRUMENTATION

PACKAGEELECTRICAL CONTROL FXAUST MANIFOLDPACKAGE

FUEL MANIFOLD

WAIN, FUEL VALVE FUEL ASE LINE

Figure 6-8. $-II J-2 Engine Diagram

ASI INGECTORMAIN INJECTOR

Lox

ERODED AREA

Figure 6-9, S-II

Nyon?

PRESSUR:

TEMPERATURE,

°K

LHp

FLOW,

m/s

LOX

FLOW,

m3/s

‘V FUEL FLOWRATE INCREASE, 318.9VW ENG PERFORM. UECREASE, 319.0

30 TTFUEL PUMP DISCHARGE PRESSURE

+ 1200860 =| =

LOX PUMP DISCHARGE PRESS JRE prt0o &700 Rt + + fae i000 &WAIN [OX INJECTION PRESSURE r yy 100 Ss

T } a600 [MAIN FUEL INJECTION PRESSURE 300 wpee =THRUST CHAM SSURE L goo

sa|— = .> = 240in 200

P- 250MAIN FUEL INGECTION TewPerature_—-t~| OeMo + joo ~ 260 6

=aS - F251 -

+| bez.-50 - i} 79 =?

b7g a-49 Er" 4 's asPex s

-18 [LOX FLOW ae8

a VY 22” 318.4 318.6 318.8 319.0 319.2 319.4 319.6 319.8 320

Figure 6-10.RANGE TIME, SECONDS

S-II Engine No. 2 Performance Shift

6-18

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Figure

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$-11

Engine

Ko.

2Cutoff

Performance

eesd ‘gunssa¥e 13INE

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wws6 201 ‘wor x07

are apparently the effect cf the ASI system failure. As previously stated,testing at NSFC has ropreduced the failure characteristics exhibited byengine Ya. 2 and it Fas been concluded that the ASI fucl lire failure ¢i¢indeec ritiate ergine he, 2 failure. [t has been reconmerded that theASI fuel line and installatior be improved by redesign

Other engire parameters that reflect the cutoff characteristics of engineNe. Z are show ir Figure €-12. Tre GG LOX and fuel inlet pressuresstart to decrease rapidly at 412.7 secerds, This is the sane time thatthe main LOX ard fuel irjecticr pressures similarly decayed.

Ergine Nc. 3 vas shutdown prematurely when its LOX prevalve was in-advertertly closce by the shutdour. cf engire Ne. Z. The sequence ofevents arc changes of engire parameters are showr in Figure 6-13. Thisfigure shows the cccreases in mair LOX injectior and thrust chamberpressures arc LOX flowrate that resulted from the closure of the LOX pre-valve. Approximately 0.60 second later, the engine No. 3 mainstage Okpressure switches actuated normally when the main LOX injection pressurehac srepped to the switch setting band of 262 to 362 N/eme (380 te525 peia

Engine No. 3 cutoff resulted from 2 wiring harness installation error;the control harnesses for engines flo, 2 and 3 LOX prevalve solenoidwere interchanged. (Plug 206W17P7 was miscannected to receptable 206A50701insteed of 206R508J1 end plug 206N17P8 was misconnected to 206850801.)To prevent this from recurring the harnesses for subsequent stages are tobe cither reinspected and/or redesigned. Also, revisions to the enginecheckout program are to provide for individual engine prevalve timingcheckout since the simultaneous checkout of all prevalves currentlyemployed at KSC does not detect an error of the type experienced on AS-502.

6.4 S-1] STAGE SHUTDOWN TRANSIENT PERFORMANCE

The normal S-I] engine shutdown sequence is initiated by 2 LOX Tow level

indication, Five level sensors are located at the bottom of the propellanttank and engine shutdown is initiated when any two probes of the LOX tankdetect a dry condition. This condition initiates the cutoff signe] andcauses the engine propellant valves to close in the proper sequenceresulting in engine thrust decay and the cutoff sequence is complete. &similar system is provided for the LHp tank to provide for a safe shut-down should LH depletion occur first.

S-I1 ECO for AS-502 was initiated by LOX liquid level and ECO was sensed at576,33 seconds. This corresponds to 425.31 seconds of S-II stage burntime. Engine cutoff transients for engines No. 1, 4, and 5 were normabut engines No. 2 and 3 had cut off earlier at 412.92 and 414.18 seconds,respectively. Thrust decay characteristics for engines No. 4 and 5 areshown on Figure 6-14, Thrust decay rates satisfied separation requirements

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$-II

Engine

No.

3Performance

atShutdown

visd ‘3unssdud

wdB 2oL ‘M024 x07

Stage engine thrust decay (including 2 best estimate of engine No. 1) ispresented in Figure 6-15. Thrust dropped from 2,354,744 Nevitons(529,370 Ibf) at ECO to 133,446 Newtons 130,000 lbf) (5 percent thrust)approximately 0.68 seconds Tater,

The shutdown transient performance, based on telemetered perfcrmarce data,was determined for the S-II stage. The calcuTatec total engire cutaffimpulse was 578,088 N-s (129.960 Ibf-s) which corresponds te an equivalentVelocity change of 2.74 m/s (8,99 ft/s). Comparisons of flight and pre-dicted values of cutoff impulse anc velocity change are presented inTable 6-6. Two reasons my be given to account for the large discrepanciesThe most significant reason is that the two engines out condition reducedthe 5-11 stage thrust approximately 4G percent, In addition, thrustverformance just prior to ECC was Tower than normal and is partiallyattributed to loner than normal accelerations. The remaining causes ofthis impulse ciscrepancy have not been identified at this time,Table 6-5. S-II Cutoff Impulse

FLicHT PERCENT DEVIATIONPARAMETER PREDICTED (3_ENGTNES) FROM PREDICTED

(S ENGINES) Ewcrne [curo. vara

|

eweine] cura. naceCutoff Nes 1,161,742

|

578,088

|

423,049

|

-20.2

|

63.5Impulse (Thf-s)

|

°261,171

|

1237960

|

53286Yelocity m/s 5.53 2.74 2.01 -30.5

|

63,7Tnerease (Ft/s) Tele 8199 6.59i 6.5 S-IL PROPELLANT MANAGEMENT

The propellant management system controls Toading and engine mixtureratios (LOX to LH2) te ensure balanced consumption of LOY and LipCapacitance probes mounted in the LOX and Liz tanks monitor the mass ofpropellants. At PU activation (5.5 seconds after J-2 ignition) thesystem senses the LOX to LH2 imbalance and commands the engine to burtat the high rate engine mixture of 5.5:1, When the excess LOX is coneSumed the PU system then commands the engine to burn at a mixture ratioof 4.65:1, striving for simultaneous depletion cf LCX and LH2 formaximum stage performance,The propellant management system satisfactorily performed the functionsOf propellant loading, mass indication, point sensor level indication,Propellant utilizattcn and programed mixture ratio operation.The facility Propellant Tanking Control System (PTCS) functionedsatisfactorily during S-IE Toading and replenishment. The best estimatesOf liquid propellant mass in the tanks at [SC are 69,275 kilograms{152,726 tbm} Lig and 359,033 kilograms (791,532 Ibn} LOX based onFlowmeter integration and the masses. remaining at ECO. These provellant

6-22

THRUST.

10°

WSTAGE

THRU

ST,

105

Nv

12 TOTOENG NO. # fb .28

10 ANG NO. 54b..20

0.3] 5S wd 2. \ ae

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TIME FROM ECO, SECONDS

4576.5 $76.6 576.) 576.8 576.9 577.0 577.1 517.2 577.3RANGE TIME, SECONDS

WLVDC INTERRUPT (S-11 ECO SCRSED)ANIL START OF Ty, 576.33,Figure 6-14. S-IT Engine Shutdown Transient

5L1.0

4a 5oe

3 2bos

2 2fo.4 =

| a bo2 &

0 el 0 a1 0.2 3 0.40.5 0.6 0.7 0.8 0.9TIME FROV ECO, SECONDS

4976.5 576.6 576.7 576.8 576.9 577.0 577.1 577.2 B77,RANGE TIME, SECONDS

‘UV LvDC INTERRUPT (S-IT ECO. SENSED)AND START OF Ty, 576,33

Figure 6-1. S-II Stage Thrust Decay

6-23

quantities were 0.7 percent more than predictec fer LOX and 0.31 percentmore than predicted for LH. Best estimate propellant masses at liftoffwere 0.48 percent above predicted values.

‘The "PU Activate" command was received and the FU valves stepped from thehominal engine start position to the full-clesed pesition at §.5 secondsafter ESC as shown in Figure €-16, This provided a norinal EMR of 5.5for the first phase of S-II Programmed Mixture Ratio (PHR). Upon shut-down of engines No. 2 and 3 (approximately ESC +263 seconds} the PU valvesmoved momentarily off the high EMR stop, returning at ESC +268 secondsafter an excursion of approximately 5 degrees. No change in EMR orperformance resulted since the PU valves are ineffective in this rangeof travel, At ESC +312.3 seconds the PU valves gradually moved to aposition of -24 degrees at ECO, compared to a predicted value of -13,8degrees. Extrapolation of PU error signal data indicates that this stepwould have begun at ESC +268 secends under normal engine operation,which is well within the predicted step tine of 250 450 seconds.

29 y TRCTAL,PREC LEEDJcl

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LESREL

S.

Loe S18 5c, 199.78ACTIVATE 5-F1 PU S¥STEY cOMMpND, 15

PEN

¢ 30 196 Ezest, Srcenas

a

1 L 1 L Zs 1 L200 70 300 3 00 460 Som 550 Bou

RANGE TIME, SECONDSFigure 6-16. S-II Stage PU Valve Position

Other than the temporary excursion that occurred at ESC +263 seconds,the differences between the actual and predicted histories of PU valveposition were caused by the changes in liquid surface angle resultingfrom the early shutdown of engines No, 2 and 3. The initial thrustimbalance at engine failure caused the vehicle to pitch up. This causeda sufficient decrease in the PU error signal to drive the PU valves offthe high EMR stop. The flight contro] system reacted to keep the vehicleon course, the engines gimbaled to keep the thrust vector of the remaining

6-24

three engines passing through the vehicle Center of Gravity (C6). Thimaneuver raised the liquid level at the LOX probe and towered it at theLHg probe, increasing the PU error signal sufficiently to return the PUvalves to the stop and delay the PAR valve step for 44 seconds. Thisdelay was aggravated by having only three engines burning to consume theincrease in LOX error indicated by the PU probes, The tilt of liquidsurface was approximately 3 degrees ‘rom a plane normal to the vehictelongitudinal axis after attitude recovery by the flight control systemdecreasing to approximately 2 degrees at ECO as the remaining propellantswere consumed and the vehicle CG moved forward. The difference in LOXbuen-off rate between the AS-501 and AS-502 flights is indicated bycomparing the rates of change of PU valve angles when PHR occurred foreach flight. The ratio of the valve angle slopes between the two flightsis about 3/5. Since two of the five engines were shutdown prior to PHR,this fs the ratio expected

Analysis of flight data shows that indicated changes in ligu‘d level dueto liquid propellant sloshing were significantly attenuated by the PUsysten, he PU valve response to the indicated change in propellantlevels at 415 seconds (due to shutdown of engines No. 2 and 3} reducedfrom 89 to 17 deg/voit. Later in the flight after PMR step, changes inpropellant leve’s due to sloshing at frequencies of 0.4 hertz {contribu-tion from LHz tank] and 0,6 hertz (contribution from LOX tank) wereobserved in the PU error signal but were not observed at all in the PUvalve response, indicating that slosh dynamics were effectively fii teredfrom commands to the PU valve positioning system, No changes affectingpropellant utilization system dynamic performance are recommended farfuture flights.

The lower than predicted position of the PU valve at ECO was the resultof earlier errors in the measurement of LOX level; the previously mentionedLOX surface tilt caused a larger LOX mass to be measured and to be con-sumed by the engines. At approximately ESC +380 seconds. there wanecessarily a deficiency of actual LOX mass in the tanks relative to LHzmass. The PU system therefore commanded that less LOX be consumed andyeduced the mixture ratio below the 4.65 RMR setting. As ECO was ap-proached, the sensed PU errors became progressively larger based on apercent of propellant remaining, As a result, the PU valves were drivento an unusually Tow EMR position causing the cutoff residuals to beexcessively high. This explanation is supported both by the residualslisted in Table 6-6 and the fact that ECO was initiated by the LOXdepletion sensors.

Figures 6-17 and 6-18 show the PU system nonlinearity es determined bycomparison of mass data from the PU probes and the engine flowmeters withcorrections made for liquid surface tilt. The PU system error at cutoffsignal was 373 kflograms (822 lbm) of LH2 relative to that predicted atthe actual LOX cutoff level. This compares to an allowable LH2 error of3665 kilograns (#1465 Ibm).

6-25

Table 6-8. S-II Stage Propellant Mass History

EtSios

ae So 2a a nee

[ea] RRBs,

Mom sree) 1 ar aie]

(29

em are9, be ke cee

|

tee Ke

eeer | aes dines [ee aee fon aks st

l iSacketodthaOTSanEySaTetneeateSesens ot oy

IEAM” © tater ate tous | LOX liquid level initiated engine shutdown sequence and the LYDC sensedECO at ESC +426.87 secords. LOX remaining in the tank was 1814 kilograms(4000 1bm) versus 1988 kilegrams (4383 Ibm) predicted. LHz remaining inthe tank was 1598 kilograms (3523 Thr) versus 1276 kilograms (2812 1bm)predicted. Cutoff residuals were determined by extrapolation of paintsensor data tc the time af cutoff signal. A normal 0.5 second ECO timedelay (winimize residuals) was not incorporated on AS-502; hence, thepredicted residuals, shown above, do not include this delay.

The test estimate of LOX consumption was 357,219 kilograms (787,532 bm)as compared to a prediction cf 355,218 kilograms (783,114 Ibm). The bestestimate of LH2 consumption was 67,677 kilograms (149,202 bm) comparedte 67,783 kilograms (149,437 lbm) predicted. These correlations weremade from S~IT ignition ta S-IT stage ECO command using data in Table 6-6without a 0.5 second time delay incorporated.

Table 6-6 presents a comparison of propellant masses as measured by theflowreters, point sensors, and PU probes. The best estimate of propellantass is based on the propellant residuals at ECO as indicated by the pointsensors, and on integration of the flowmeter data.

The best estimate of Taunch vehicle tatal mass at S-II ignition and cut-off, as determined from capacitance prohe, point level sensors, flowmeters and the trajectory simulation is 643,856 kilograms (1,419,460 Ibm)and 210,789 kilograms (464,710 Ibm) respectively, as shown in Figure 6-19,

6.6 S-I1 PRESSURIZATION SYSTEMS

The S-1] pressurization system function is to provide the necessarypositive pressure to the J-2 engines propellant pumps and to increase thestructural capability of the tanks. The systen is comprised of tankullage pressure regulators and vent valves, LOX heat exchangers (integral

6-26

LInERB

eee

ab

ror mess, 0} egFigure 6-17. 8-11 LOX Probe/Tank Mismatch

TOTAL LZ 4855, 107 Yena w an va us01 L L Le i

ce

L.-2 io a 16 ey % 7 aTOTAL Ue AS$, 10°bg

Figure 6-18. S-1L LH, Probe/Tank Mismatch

6-27

70

IGNITION

MASS,

103

kg

CLTOFF MASS, 103 Thm

462.5 463.5 484.5 465.5 886.5645.6 1 L L i L. POINT” SENSOR FLIGHT“™ SIMULATIONraze.fest CSTE

695.0 |Po PROBE|BEST ESTIMATE: L yearIGNITION HAS643,856 41469 kg1,419,460 £3239" te _|SM-5F crore Mass210,789 3566 ka L saz.464,710 F248 Tom

J SENSOR

i } rst9,

643.5f 1ate.

FLOW INTEGRAL683.0 |/ VY b 1417

52.5 rue.

542.0 Plate203.5 710.0 210-6 ino 2 72.0CUTOFF MASS, 103 kg

Figure 6-19. S-IT Mass at Ignition and Cutotf

6-28

102

Ton

IGNITION

mAs:

part of J-2 engine), interconnecting tubing and manifolds. Prior toJaunch, the LOX/LH2 tanks are prepressurized by around source Gaseou:Helium (GHe). During powered Flight of the S-II stage, the LOX tank ispressurized by GOX from the LOX heat exchanger. The LH2 tank is pres-surized by GH2 from the thrust chamber hydrogen injector manifold andpressurization is regulated by the LH2 pressure regulator and vent valve.

6.6.1 S-II Fuel Pressurization System

The S-II LH2 tank wllage pressure during S-IC boost was sTiyhtiy “ower,as planned, than for AS-501. Figure 6-20 presents the fuel tank ullagepressure for AS-502 as compared to AS-50T from prepressurization untilS-I1 ECO,

Y VENT VALVE HO. 1 OPEN 560.5 VENT LYE NO, 1 OPEN, 401‘YH Lvoc 5-11 ESC, 149.76 FLVOC INTERRUPT (5-1; ECO SENSED}sof ster nessunzATICN, 466.96 MO START GF Ty 4576.30 Lis

po \sane |* yess aLGaace UM'sZa ,Z oho Ei Kft

3 om saat nea 15-52 2

8 =

Ty Zz vv wy Y pisa

ronne TIME, seconosFigure 6-20, $-II Fuel Tank Ullage Pressure

The Lip tank vent valves for AS-502 were set ta crack between 21.4 and22.8 K/cm2 (31 to 33 psia) as compared to 23.4 to 24.8 N/em2 (34 to36 psia) for AS-501. During prepressurization the LH» tank waspressurized to 22.0 N/cm2 (32 psia) in 34 seconds.

“Hi-press” was utilized as recommended after the AS-501 flight and wasterminated by the cracking of vent valve to, 1. Vent valve reseat oc-curred at an ullage pressure of 21.4 N/cm? (31 psia}, Ullage pressuredecayed approximately 0.7 N/em2 (1 psi) to 20,7 N/cm2 (30 psia) at S-Iengine start which is well above the minimum requirement of 19.9 N/eme{27.5 psia).

6-29

LHg tank pressurization during $-II flight was normal. The regulatorcontrolled the ullage pressure within the requlator band of 19.65 to20.70 Nem? (28.50 to 30 psia) up to the time of step pressurizationThe ullage pressure increased after step pressurization and at 481 seconds,vent valve Ho. 1 cracked, controlling the pressure at 22 N/em2 (32 psia}until S-IT ECO. Yent valve No. 2 did not open during the entire flight

Figure 6-21 presents engine inlet data bands consisting of total pressuretemperature and HPSP. “he band extremes include data from e1l fiveengines during operations; however, after shutdown of engines No. 2 end3, only data from engines No. 1, 4, and § are shown. Engine inlet Liltotal pressure was obtained by adding calculated dynamic pressure toengine inlet LH2 static pressure, The total pressure increase after468.36 seconds was due to the increase in ullage pressure at step pres-surization. Engine ‘niet Lz temperature shows 2 gradual increase whicis the effect of stratification. The stratification was well within ac-ceptable limits. Engine inlet LH2 NPSP, obteined by subtractirg the Lzvapor pressure from engine inlet LHz total pressure, was well above theminimum requirement.

The NPSP increased, as shown on Figure 6-21, after 468.36 seconds as theresult of ullage pressure incrcase at step pressurization. The gradualdecrease in available IPSP, commencing at abaut 5GC seconds, was theresult of warmer LH2 entering the fuel pump,

6.6.2 S-II LOX Pressurizatior System

Figure 6-22 presents the LOX tank ullage pressure, comared to AS-501,from prepressurizaticn until S-II ECO. A pressure decay of only 0.2 N/cm2(0.29 psia) was recerded during S-IC boost as compared to 3.2 Nfeme(4.7 psia) for AS-5C1. The reduction of pressure decay was due to thePretaunch evacuation of the conmon bulkhead. The LOX tank was pressurizedto the pressure switch setting of 26.5 N/om2 (38.5 psia) in 67 seconds.LOX tank "hi-press" was not utilized since ullage pressure decay waspredicted to be negligible with the cormon buTkhead evacuated. The LOtank ullage pressure at S-II engine start was well above the S-I1 enginestart requirement of 22,7 N/emé (33 psia} and above the redline linitsat launch.

LOX tank pressurization system performance through S-I1 boost was adequateto maintain engine HPSP requirements; however, the ullage pressuredecreased below the regulator control band lower linit at abut the timeof EMR shift at anproximately 900 seconds, and was 24.0 K/cine (34,8 asia)at ECO.

The engine heat exchangers supplying the pressurant gas do not aave anyexcess capacity when operating at novinal or low EMR. The regulator at-tempted to keep pace with the demand by fully opening. Slowever, wit!only three engines operating, sufficient pressurant gas wos not availableto maintain the ullage pressure. Although fewer engines were consumingpropellant, the LOX and tank surface areas which caill the ullage gas.

6-30

aTHy

ALLeg

ENLET

ATURE

,MP

TePSP,

Nom?

VY lye $-11 est, 1 49.76VY S-IL ENG NO. 2 GuT, a1z.32Ys tr eso mg. 3 our, 414.18VY Lave TecLeRUPT

(5-1 ECO SCHSED) AD START OF T,, 576.33

}- 26

PuENGINE TLE” Lp — 2;TOTAL PRESSURE RAND SSS pe oy

Ew 8

hen *

” . b-070 47ENGINE INLET LHy Ln ZBTepPeranvee

2

\ Ne2 -\ — 422 SBf-423 #2

var LET UiWPS ER]

lo BSS

ANS ie8 Fig i

10 26 2

+ RINIMUM NPSP. REQUIREMENT4 ft ~ 6; Lb 4“O50 100 180200250 4007 450

TIME FROM ESC,

159 200” 250300380

400460 5008000RANGE TIME, SECONOS

Figure 6-21. $-II Fuel Pump Inlet Conditions

6-31

VY Lunt 5-1 B80, 19.76Pevwe onereer

(SoU Eco S967) sup S=ERT CF Ty

= bs\ Aesowe 190 bes

Nake a3 mu

/ Le4 | es2297 “a8 2 7} 00 30 ry to we

PAIGE “URE, SECIS

LOeU-

aASE

eRESSURL,

HZor

®

LOELLLSGL

FRESSL

RE,si

e

Figure 6-22. S-IT LOX Tank Ullage Pressure

did not change. Moreover, the uTlage shrinkage created a larcer demardof pressurant gases from the remaining engines, resulting in colder heatexchanger outlet temperatures, The data shows thet tovarcs $-II cutoff,the heat exchanger ex't temperatures were at saturation.

Figure 6-23 presents engine inlet data bands consisting of total pressuretemperature and NPSP, The bands include data from all five encires untithe time that engines No, 2 and 3 cut off; thereafter, only cata fromengines No, 1, 4, and § are shawn. Engine inlet LOX total pressure wasdetermined by adding the calculated dynamic pressure anc heac correctionfor the location of the pressure pickup to the pump inlet LOX staticpressure neasurerent. The abrupt decrease in total pressure at approxi-mately 413 seconds is a direct result of acceleration heac loss due tepremature cutoff of engines No. 2 and 3, The gradual total pressureloss which occurred toward the end of S-II boost is the combined resultof the decrease in liquid head and the decrease in ullage pressure. Thegradual increase in liquid temperature toward the enc of $-II beast showsthat the effect of liquid stratification was well within the acceptablelimit.

Engine inlet LOX available NPSP was obtained by subtractire the LOX vaporpressure from engine LOX inlet total pressure. The NPSP was well abovethe minimum NPSP requirement. The abrupt decrease in availatle NPSP at413 seconds is a direct result of the acceleration liauid head last at‘the premature cutoff of engines No. 2 and 3. The gradual decrease ofavailable NPSP toward the end of S-II boost is the combined result of adacrease in liquid head, a decrease in ullace pressure, and an increasein Viquid temperature caused by stratification.

6-32

LET

TEMP

ERAT

URE ,

LOX

IN°K

NPSP,

N/on?

VY Loc S-1I £86, 149.76 YW oLvoc INTERRUPT

YSENG NO. 2 OUT, 412,92 (S-IT ECO SENSED) AND START OF Ty,S-11 ENG nO. 3 QUT, 414.18 576.33

8 [7 EnGine INET TOK) 2TOTAL PRESSURE BANDL gg =

u Zbas 52 af

baoa7 zP38 x

as ~

9 -294fF -295 o

” ENGINE INLET LOX TEMPERATURE po 296“L (Ob -2o7 HE

*° L -208 =

a0) } - 299 7

ENGINE TNLET LOX NPSP BAND [7-34026

zkMINIMUM. NPSP. 1REQUIREMENT &

a106

50 100 150 200 250 «300 «350 «400-450TIME FROM £SG, SECONDS

Qe150 200-250 «300=«380« 400 50—«S0G SO 600

RANGE TINE, SECONDSFigure 6-23. S-I1 LOX Pump Inlet Conditions

6-33

6.7 S-II PNEUMATIC CONTROL PRESSURE SYSTEM

The pneumaric control pressure system provides pressurized relium fer theactuation of propellant system valves during flight.

Performance of the pneumatic control system was satisfactory. Figure6-24 shows receiver pressure and regulator outlet pressure cf the systemfrom before liftoff until S-I1 ECO. The receiver pressure and regulatoroutlet pressure were w'th'n the predicted bends throughcut the flight.The three step-decreases of pressure represer.ted helium demands forvalve actuations, Only the times at which the demancs for prevalveclosures on engines No. 2 and 3 were made ere unusual. Tatle 6-7 showsthe heliun mass used by the syster.

Y Lyoc S-If ESC, 149.76 ‘Y Lyoc INTERRUPT‘Y S-1t ENG NO. 2 OUT, 412.92 {8-11 ECO SENSED) AND STARTWV S-ILENG NO. 30UT,414.18 OF Ty, 576.332500

: 3609waco |__| Auonaoue wren ier ||

[3200eb:sooo ED'CTION BAND

L Leave

1600 ' peesALLORRETE 2% “OT Lover Corr == ATROORNE N : P7000S LINE MT open “+ pre closeo a2 im Oren . + g3 ST] ee #0. 1 ay z3 RECIRC s PREVALYES CLOSED] piso2 aLves

evoseo NY20 L200

ACTLAL REG OUTLET PRESSREGULATOR] PREEICTIONPressurepees ene100 _ as f

4 Inpteares ' ALLOWABLE LIMITS | aoePREDIeTIONQ | ¥ ZI),+100 oO 100 200. 300 400 $20 60

RINGS TIME, SECONDSFigure 6-24. S-I1 Recirculation Valves Receiver and Regulator

Outlet Pressures

6-34

Table 6-7. S-II Helium Kass

PNEUMATIC CONTROL PRESSURE S¥STEVELIUY MASS READINGSLAUNCH etSEQUENCE {as-5C2)

ACTUAL PREDICTED

Launch 1.71 kg 1.65 to 1.76 kgHinus 30 seconds (3.77 Ten} (3.64 to 3.88 IbmS-IT Engine 1.70 kg 1.99 to 1,70 kaStart Comrand (3.76 Tom) —|(3.50 to 3.76 Ibm)S-II Engines No. 2 and 3 1,56 kg HoneCuter (3.44 ibm)S-1] Engine 1.47 kg 1.47 to 1.60 kgCutoff Command {3.25 Tom) (3.25 to 3.52 lbm) 6.8 CAMERA EJECTION SYSTEM

The function of the camera ejection system is to provide GHe, upon command,to eject the camera cannisters. The camera ejection system consists oftwo helium spheres, tubing and valves.The camera ejection system performed satisfactorily. The ejection sequencewas programmed to start at T3 +38.0 seconds and actually started at13 + 37,97 seconds.

Figure 6-25 indicates a gradual pressure decay cf toth subsystems. Thesubsystem located at position III exhibits the greater, but acceptable,rate of decay anc lower stcrage pressure. However, both subsysters hadsufficient pressure to eject the careras. The same pressure dropcharacteristics were exhibited by the systems during ejection, at whichtime the pressures decreased 130 N/om@ (190 psid). A discussion ofcamera recovery operations way be found in paragraph 19.6.

6.9 HELIUM INJECTION SYSTEK

The inflight helium injection system supplements natural convecticn re-circulation in the LOX recirculation lines, This system injects heliuminto the bottom of the return lines to decrease the return line fluiddensity thereby increasing the recirculation driving force.

6-38

In general, performance of the helium injection system was satisfactoryRequirements were met and parameters were in good agreement with predictedvelues. The supply bottTe was pressurized to 2070 Nom2 (3000 psia) priorto liftoff anc by ESC was 787 N/om2 (1140 psia). This usage of heliumresulted in a helium injection flowrate of 1.67 SCMM (59 SCFI),

As discussed in paragraph 6.2, LOX recirculation was satisfactory exceptfor the prelaunch chilldown of engines No. 3 and 4, Because of theproblem experienced in chilling these engines to below the prelaunchred] ine maximum the following changes are being considered for imptemen-tation priay to AS-503 CDDT:

a. fncrease helium injection system total flowrate from 1.13 to 1.70 SCHM(46 to 60 SCFM) to 1,70 to 2,26 SCMM (60 to 80 SCFM) by increasingthe prirary orifice size.

b. During checkout, verify that the helium injection flow is distributedevenly to all engines.

c. Add screens upstream of heliur. injection orifices.

d. Velete the solenoid outlet pressure instrumentation and add newpressure measurement downstream of the primary orifice. Thiseasurenent will be more sensitive to heliun flow distributionwhile still being representative of total flow conditions,

Fso00

L2iso

sca

sm easiTtON T

: +g |_|

Ben |

é Vue 11 fe, 10.8100) (] CAMERA EAECTION, 187.46

vel eel |RaNGe T2ME, SEcoNES

Figure 6-25. S-I1 Camera Ejection Pressures

6-36

S1OR

AGE

QOTTLE

PRESSURE,

s5‘a

SECTION 7

S-1¥B PROPULSION

7.1 SUMMARY

The S-1¥B_stage propulsion system consists of a main and an auxiliarysyste. The main propulsion system consists of a single, bipropellantJ+2 engine, fuel system, oxidizer system, and a propellant managenentsystem. The Auxiliary Propulsion System (APS) is provided to controlthe vehicle attitude during S-IVB operation and position the propellantin the stage during first burn cutoff, parking orbit coast, second burnrestart, and cutoff transicnts. The other systems discussec urder thisection are the Continuous Vent System (CVS) and the pneumatic controsystem,

The J-2 engine operated satisfactorily throughout the operational phaseof first bun, However, a total performance shift of 2.3 percentdecrease in thrust occurred during first burn from 684 to 702 seconds.The engine continued to operate at the shifted performance level andhad a normal shutdown, 5-1V6 first burn time was 166.52 seconds whichwas 28.95 seconds longer than predicted due to the two engines outcondition on the S-II stage. This burn time was computed from StartTank Discharge Valve (STDV) open (occurred 3.24 secones after the LaunchVehicle Digital Computer [LVDC] S-1VB engine start sequence command) tothe LYDC velocity cutoff command,

The stage performance during first burn, as determined from thepropulsion reconstruction analysis, deviated from the predicted at the60 second time slice by -0.08 percent for thrust and 0.06 percent forspecific impulse.

The S-I¥B stage first burn Engine Cutoff (ECO) was initiated by theLVEC at 747.04 seconds. The LOX mass measuring side of the PropellantUtilization (PL) system malfunctioned prior to the attempted restart. TheLOX mass measuring system malfunction caused a 100 percent LOX loadindication prior to and during the restart attempt. The probable causeof the erroneous 100 percent LOX mass indication may have been due toshorting of the inner and outer elements of the LOX PU probe fromMetallic debris that could have been in the LOX tank. Also, an

7-1

interm'ttent short in the cable shield betweer. the mass probe and the PUelectronics assembly may have occurred. Additional tests on cableconnector crimping will be canducted aleng with the evaluation of insulat-‘ing one ar both of the probe elements to prevent shorting by debris.

Engine restart conditions were within limits ever though main chambersecond ignition failed to occur, Results thus far indicate that a leakin the Augmented Spark Igniter (ASI) fuel supply system probably occurred,resulting in a perfomance shift and the failure to achieve restart.Additional engine tests have essentially verified the performance shiftand the restart failure. A modification of the ASI propellant feed] ines{both fuel and LOx) and their instaTlation is being accomplished.

All subsystems operationally met all performance and stage requirements.However, there were tuo unexpected deviations which are summarized asfollows?

#. Two LOX ullage pressure makeup cycles were required during S-IC boost.The ullage pressure decay requiring two makeup cycles is partially dueto ullage cooling, tank geometry response to the yehicle axialacceleration, and possibly vent/relief and/or relief valve intermittentleakage. Investigation is continuing into the LOX ullage pressuredecay.

b. A possible cold helium leak was indicated after first burn ECO. Thecause of the leak is stil] under extensive investigation with ssecialinterest into the cold helium system conoseals and joints, the LOXpressurization module, and the cold helium sprere pressure transducer.The conoseals wil} be changed to 7075 aluminum coated with teflonthroughout the cold helium system on sudsequent vehicles.

7.2 S-IVB CHILLDOWN AND BUILDUP TRANSIENT PERFORMANCE =OR FIRST BURN

LH2 and LOX recirculation systems are required for proper prestartconditions. With the required prestart conditions, a start is obtainedby proper engine sequencing of the engine components with the powerobtained from an engine mounted helium control sphere. A hydrogen starttank is employed to spin the turbopumps during the start transient untilthe Gas Generator (30) systen is ooerating properly to continue enginestart and maintain engine operation. During the start transient andengine operation, tne start tank is refilled for another stars. Theignition system initiates combustion in the thrust chamber and GG. Thesystem includes four spark plugs, four spark exciters, an ignitionchamser, an ASI valve, an ignition detector probe, and the necessaryelectrical harness and plumbing to join the parts into a systemTne propellant recirculation systems performed satisfactorily, meetingstart and run box requirements for fuel and LOX as shown in Figure 7-1.

7-2

Ung POMP THLE ORFSSURE, pss2

m a 2B R Fa a as2 : F

ren storeaLysSeca

1 9al 2 a - ae

5 Fe PT einex |G i stat Box* § 100 ygf é 12 | wegia f -420g? e ra T =

Bo bem - son e

Sa i 2 3

& bee tm sot yjgol — t 42a

_ wa 2 zeLy PUNE (ALT PRESSURE, Neer?

Lik uM aNLLT PRESSURE, psia2B # % 40 a8 6 2

wa i i a ownTien 1M, 0H

ESC SLEONDS a11 1

ca} 2 15 - -} ae5 1 .

x 6 od 2eof _t | z5 ! 1 eae5 tot i 5= ! 5.53) 1 ! F= cap bo —4 53 ( re= 1 ; iEv L - I - 55. TAZ be 1 ~* \ ee 1 296

stam its pny6 — a.

-3002a iaa a aa

0D, PUPP INLET PRESSURE, W/ar2

Figure 7-1, S-I¥8 Start Box and Run Requirements - First Burn

7-3

The thrust chamber at launch was we?l below the maximum allowableredline linit of 167°K (-160°F). At S-IYB first burn Engine Start Commanc(ESC), the temperature was 159°K (-173°F}, which is within the requiremertof 167 228°K (-160 450°F) as shown in Figure 7-2.

The chilldown and loading of the engine Gaseous Hydrogen (Gz) start sphereand pneumatic control sphere prior to liftoff were satisfactory. Figure7-3 shows the start tank performance fron first burn [SC. At firstburn start command the start tank conditians were within the requiredS-IVB region for initial start of 913.56 451,71 N/cm2 and 161 + 16.7°K(1325 475 psia and -170 30°F). The discharge was completed and the refillinitiated by $-1¥B first burn ESC + 3.85 seconds. The refill was satisfac.tory and in good agreement with the acceptance test. The control bottlepressure and temperatures at liftoff were 2061.6 N/em2 (2999 psia) and160.6°K (-171°F), Nominal chilldown system performance levels were observedduring the chilldown operation. LOX system chilldovn, which was continuousfrom before liftoff until just prior to S-IVB first burn ESC, was satisfactory. At ESC the LOX pump inlet temperature was 91.5°K (-295°F),Nominal chiltdown system performance levels were observed during thechil ldoun operation.

The first burn start transient was satisfactory. The thrust. butldup waswithin the Timits set by the engine manufacturer, A faster thrustbuitdup to the 90 percent level as compared to the acceptance testresults was observed on this flight and is shown in Figure 7-4. Thisbuildup was mainly due to the shorter platees tine during main oxidizervalve opening time. A similar thrust duiTdu> was observed on AS-507Table 7-1 shows the major sequence of events during the buildup transient.The PU system provided the proper null setting of the PB valve duringthe start transtent until system activation, The total impulse fromSTDV to STDV #2.5 seconds was 802,233 Hs (180,350 Ibf-s} compared to604,651 N-s (135,931 Ibf-s) during the same interval for the acceptancetest,

7.3 S-IVB MAIN STAGE PERFORMANCE FOR FIRST BURN

The J-2 engine provides the necessary propulsive performance for theS-IVB stage and is a high performance engine utilizing pump-fed propel-Jants. The J-2 engine used on the S-:¥B has restart capabilities. Ataltitude the engine produces a noriinal thrust to 1,000,850 Newtons(225,000 Ib#) at a LOX to LH2 ratio of 5,5:1. The’engine ts capableof operating between 4.5 and 5.5:1 LOX/tHp mixture ratio for thedesired propellant utilization at stage cutoff. The engine.features atubular-watled, betl-shaped thrust chamber and independently driven,direct-drive turbopumps.

Two analytical techniques were enptoyed in evaluating S-IVB stage propulsion system perfomance. The primary method, propulsion reconstruction

74

“K

TARUST

CHANBLR

TEMPERATIR:

zeeINITIATION OF THRUST CHAMBER CHILLDOWN, 600TERMINATION OF THRUST CHAMBER CHTLLOOWN, 0.0

26g}. —— WS-IV ENG START SEQJENCE COMMAND, 577.28 10

240 \ - .

| -s0220

MAXIMUM STUY LAMIT209 7 L -100

MAXIMUM REDLINErap LIMIT AT LALNCH

J 167°(-160°F) L150

160} — sinimun [STDY LIMIT, Lf

L-200tr a or ‘ a _ _

10h L -250

160“1200-1000 -800—~-600.~~-400 20 °“IME FROM FIRST ESC, SECONDS

9-é00 =agg 208 ° 200 400

RANGE TINE, SECONDS

Figure 7-2, S-1VB Thrust Chamber Temperature ~ First Burn

7-5

THRUST

CHAMBER

TEMPERATURE

,

we = ‘eon990} K— || 1ST BURN ESC FJonl - << ec ene foot"

8 roof —4 4 t.4 1000geo ef

e3 #00 F© soo

e

LL 500 =

ent—of - .°

aSTi TANK TEMPCRATIRE,

Figure 7-3. S-I¥B Start Tank Performance

analysis, utilized telenetered engine and stage data to compute Tongi-tudinal thrust, specific impulse, and stage mass flowrate. In the secondmethod, flight simulation, a five-degree-af-freedom trajectory simulationwas utilized to fit propulsion reconstruction analysis results to thetrajectory. Using a differential correction procedure, this simulationdetermined adjustments to the recorstruction analysis of thrust andmass flow histories to yield a simulated trajectory which closelymatched the observed pestflight trajectory.The propulsion reconstruction analysis showed that the stage performanceduring mainstage operation was satisfactory. A comparison of predictedand actual performance of thrust, total flowrate, specific impulse, andmixture ratio versus time is shown in Figure 7-6. Table 7-2 shows theSpecific impulse, flowrates and mixture ratio deviations from thePredicted at the 60 second time slice. This tine slice performance is

7-6

12 T T EAGINE 9-2082]$-1VB 1ST BURN STDY OPEN (IGNITION), 680.52TofS 8=2VB 1ST BuRW 30 PERCENT THRUST, 582.79

b -20208 ” ACTUAL 2% MAXIMUM. 1) “base,20.8 4 7 -z we

. HINIMUB LO

v ACCEPT TEST

0 0.5 1.0) 1.8 2.0 2.5TIME FROM STOV COMMAND, SECONDS

581-0 5815 582.0 582.5 583.0RANGE TIME, SECONDS

Figure 7-4, S-IVB BufTdup Transient ~ First Burn

Table 7-1. S-IVB Engine Start Sequence Events - First Burn

TIME OF EVENT IN RANGE TIME (SECONDS)EVENTPREDICTED ACTUAL

S-1V8 Engine Start SequenceCommand {ESC} 518.69 577.28

Start Tank DischargeValve (STOV) Open 521,69 580.52

Mainstage ControlSolenoid 521.94 580.75

Mainstage OK 524.10 882.03Main LOX Valve Open 524.58 562.94090 Percent Thrust 523.79 582.79

1-7

TPREDSCTED

108

95

TARUST,

904250

4200

sy

FEC

IMPULSE,

Res/kg

4190

SPEC:

4080asp o

500ees PREDICTED: i | basi200

FF

—T— —" INITIAL PERFORMANCE DECAY BEGINS, 684 “soW PERFORMANCE SHIFT ENDS, 702 \175 f-—{—— VW S-1¥6 WELociTy curorr comane, 747.04 fF 00

| F 30TOTAL

FLOWRATE,

KGis

TOTAL

FLOARATE

,Ver 7

s

1505.8

UEL 5.6 MI

XTURE

54 = RATIO,

LOX/?

ENSIN

5.2 9 20 40 500 MOTO 4GSCTeO Tao

TIME FROM STOV + 2.5 SECONUS

600620640 680,660) 700 728740 760RANGE TINE, SECONDS

OTE: THESE DATA DO NOT REFLECT THE TOTAL PERFORMANCE SHIFTAS AFFECTED GY ASI FUEL LEAK.

Figure 7-5. $-IVB Steady Stale Performance - First Burn

7-8

Table 7-2. S-IVB Steady State Performance - First Burn(ESC +60 Second Time Slice)

7 PERCENTPARAWETER PREDICTED 2ECONS™RICT?ON| pressoy OEVIATION

[enon PREDICTOThrust N 1,018,864 | 1,017,771 7a |g og

(ibf) 228,980 228,804} -176 2.Specific Impulse

Nes/kg 4168.5 4167.8 2.3 0.06(ib¥-8/1bm) 424.97 428.0 0124 :

LOX Flourate jkg/s 206 .60 206.33 “0.27 0.13(Tom/s) 495189 455.30 -0.59 *.

Fuel Flowratekg/s 37.73 37.65 -0.08 0.21(b/s) 83.19 3.01 | -0.18 oe

Engine MixtureRatio

UGx/ Fuel 5.4e04 5.488 0.0046 0.08

the standardized performance which is comparable to engine acceptancetests. The 60 second time slice performance agreed with the predictedby -0.08 percent on thrust and 0.06 percent on specific impulse. Theoverall propulsion reconstruction average during first burn was -0.77percent for longitudinal thrust and -0.66 percent in longitudinalspecific impulse compared fo the predicted.

Curing first burn operation the engine showed a 2.3 percent drop inthrust frem 684 to 702 seconds as shown in Figure 7-5. The most probablecause for this performance shift was an ASI fuel line leak. The suspectedAS] fuel line leak is discussed in detail under paragraph 7.6. Enginetests simulating an ASI fuel line failure have been conducted at MarshallSpace Flight Center (NSFC) and the engine contractor's test facilitywhich have results compatible with the observed performance shift, asubsequent stabilization, and then a nanral shutdown.

The flight simulation analysis shewed a decrease of 0.20 percent, comparedto the prediction, in specific impulse. Other corparisons are shown inTable 7-3.

The S-IVE burn time was 28.96 seconds longer than predicted. Table 7-4shows that the primary contributors ta the longer burn tie were devia-tions in the preconditions cf flight. Another large contributor was the

Table 7-3. Comparison of S-IVB Stage FlightReconstruction Data ~ First Burn FLIGHT RECONSTRUCTION "PERCTWT Ey. ¢RaM PRED.

‘1 5 “TSH METUE warzo

|

\ilGe aURTURE RaTIGParancrees

|

nurs. Fi nt aLFUT

ngitedsra? |v agen 7"weniere varus

|

fae 228,65eniete wars

|euss

Loss Rate’ tlksys3Longitucingt

—

Nesrkg 4s

venicleSoecific Iapulse |(3he-$/75mi

PRRAMLTERS

1ST BU FIST BARFLEGHT ANCA! SERIELongitudinal

—|

9 1,003,724 on jjonicle theast

|

{Tee 75938 feenicle Mass

|

kgs 268.01 4Lass Rate Clays 536.09 onLongitus-nat

|

preg 4149.96vericle (esate 423.18 0.20Specific (mpulse

Table 7-4. $-IVB SimTation Burn Tine Deviations

CONTRIBUTOR DEvIATION*| —BLAN TIMEDELTA (SELONGS)

Preconditions of Flight (S-II/ 2.1S-1¥8 Separation Cormand)

Velocity Magnitude /s{Space Fixed) Fes

FlightPath Argle degAltituee Km

nimiOverspeed (ECO} ys 81felS-1¥8 Tarust 4 0bfS-I¥O Mass Flow ka/s -2.0Ion/sS-IVB Initzal Mass kg seeo za13m 1198 —

explained 32.6Unexalained -3.65 |

served ass point trajectory (post‘light) minus final operationaltrajectory (predicted),

overspeed at S-IV8 ECO, The total contributors show a burn time deviationof 32.6 seconds. This is 3.65 seconds more than the actual deviationThe additional 3.65 seconds af burn time nay be accounted for byuncertainties in preconditions of flight and uncertaintics in the thrustaverage obtained fron trajectory reconstruction

7.4 $-1¥B SHUTDOWN TRANSIENT PERFORMANCE FOR FIRST BURN

S-V8 ECO was initiated at 747.04 seconds by a guidance velocity cutoffcormand which was 87,78 seconds later than predicted for first burnThis later cutoff tine wes primarily a result of the two engines out onthe S-Ii stage. The [CO transient was satisfactory and agreed closelywith the acceptance test and predictions. The total cutoff impulse to5 and zero percent of rated thrust was 236,543 N-s (53,177 Ibf=s} and264,118 N-s (59,376 Ibf-s), respectively. Cutoff occurred with the PUvalve in the fully clased position, high Engine Mixture Ratio (EMR)

The Main Oxidizer Valve (MOV) actuator temperature was 156.9°K (-177.5°F.at cutoff. The cutoff impulse was adjusted from these conditions tostandard conditions for comparison with the log book values at null PUvalve position and 233°K (-40°F) MOY actustor temperature. After theseadjustments, the flight values were near the log book values. Thethrust during cutoff is shown in Figure 7-6,

Tetemetered guidance velocity data indicated the cutoff impulsewas greater than expected as presented in Table 7-5. The differencein vehicle mass between the flight and that precicted is the probablecause for the percentage differerce in velocity increase.

7.8 S-IVB COAST PHASE CONDITIONING

The CVS performed satisfactarily, maintaining the fuel tank ullagepressure at an average level of 13.6 N/cnZ (19.75 psia). Nozzle pressuredata, thrust, and acceleration levels for first and second revolutionsin the orbit are presented in Figure 7-7. LHp ullage conditions duringcoast are shown in Figure 7-8.

The continuaus vent regulator was activated at 806.25 seconds. The tankullage pressure dropped from 22.1 to 16.8 N/om@ (32.0 to 24.4 psia) in65 seconds, and then gradually leveled off to 13.6 Nome (19.76 psia)Regulation at this level continued, with the expected operation of themain poppet periodically opening, cycling, and reseating as shown_inFigure 7-7. CVS thrust and acceleration, also shown in Figure 7-7,were based on venting parameters. The acceleration was computed fromthe calculated thrust and vehicle mass without consideration of anyvehicle drag effects. A different approach for obtaining the thrustand acceleration is discussed in paragraph 10.4.1, Continuous ventingwas terminated at 11,288.49 seconds, which was 326.20 seconds before secondburn ESC. The erroneous CVS readings experienced on AS-801 were notobserved on this flight and a further discussion is presented inParagraph 7.10.1.

1.2 rbo. 25

1.0= q VY S-1VB VELOCITY CUTOFF Lg292 oe COMMAND, 747.04 aBo. t 2- = ACCEPT TEST =be © PRECICTED Fo.153 0.6 4 £02 FLIGHT DATA 5= i

=o.t0a4

a2 bo.05

———e0.8 1.0 1s 2.0 2.5

TIME FROM FIRST ECD, SECONDS

a0 os 7B 748.5 749.0 749.5RANGE TIME, SECOKDSFigure 7-6. S-IVB Shutdown Transient Performance - First Burn

Table 7-5. S-1VB Cutaff Impulse - First Burn

PERCENT OLV LATIONFLIGHTPARAMETER oreoccrco | FROM PREDICTEDensIne

|

suto. ava

|

cucise} suto. sataCutoff es 287,177

|

264,118

|

300,669 0Imputse {1bf-s} 64,560 59,376 67 ,593 ~ “7Yelocity m/s 2.26 2.20 2.50 a3]Increase ift/s} ray 7.24 aia

|

3 10.6 Note: The parameters quoted are from velocity cutoff command to zeropercent of rated thrust,

AOE Tass tha seat St 5 yan Swucnany

Coas

tPhase

VSPerforma

S-1V

BFi

gure

7-7.

714

“

TT

FFcoaMaND,

747.04

*HINUQUS

VERT

¥ALV

EJPEH

ON,608.25

11,287.73

aS-1ve

ENG

REST

AGS

cOveuM

na,11,614.69

FPPPe

zy

|i

8

unssxed 307790 ZH

(f=

101PERCENT

SENSOR

8

1percent

sexson

|90

SERCENT

SENSCR

70PERCENT

SensoR

=UW

1CNT

Seto

n99

PERCENT

SENSOR

—1._

ang

BSSe

Zor

“aso

350

Ye CRMMMOAL 30001 zt

1t

i

7

>To

or20003000

8 cot“ADIN2® SHH SAI WLOL

aeeaTe

re

TS)

RANGE

TIME,

SECONDS

Lowe

.Ly

0:0089

a30:00

T000

5T3

000

msETITOD

Tans

RANGE

TIME,

10U95

TAUT

ES38

Figure

7-8.

S-Iv

BUllage

Conditions

-Coast

Phas

e

Psd “unssa4d JovTIA 2H 1 “saan ueaunaL a08TM0 2H way et“HANSA SSHNLOL

Calculations based on estimated temperatures indicate that the mass ventedwas approximately 1270 kilograms (280C Ibm) and that the boiloff mass Kasapproximately 1360 kilograms (3600 Tbm)

Figure 7-9 shows the start tank temperature and pressure during orbitalcoast for AS-501, 8C2, and 203. A 180-degree roll maneuver on the AS-502Flight took place at approximately 837.3 seconds, This orientation ofthe vehicle did not appear to have any effect cn the start bottle, Manyof the S-IVB pressure sphere temperature measurements during orbit arenot considered reliable; however, the temperature reading at second startcomand was very near that for AS-501,

7.6 S-EVB CHILLOQWN AND ATTEMPTED RESTART FOR SECOND BURN

The S-IVE stage provided adequate conditioning of propellants to the J-2engine for the restart attempt. The engine start sphere was rechargedproperly and mairtained sufficient pressure during coast to the restarttime. The engine control sphere gas usage was as predicted during thefirst burn and maintained sufficient system pressure for a proper restart.

Table 7-6 showing the majer events during the start transient, indicatesthat all events cccurred as required and perforrance was as predicteduntil the end of the start bottle blaudown which occurred at approximatelyESC 48.5 seconds. At this time, the wain engine should have ignited withopening of the MCV, and the GG should have bootstrapped to mainstage opera-tion. The GG was ignited and the MOV opened, however, the main enginedid not ignite as evidenced by the lack of an increase in. fuel injectortemperature. Engine operation was terminated by the Instrument Unit (IU)monitor from lack of sufficient thrust at 11,630.33 seconds.

The propellant recirculation systems performed satisfactorily and metstart and run box requirerents for fuel and LOX as shown in Figure 7-10.Second burn fuel lead generally followed the predicted pattern andresulted in satisfactory conditions as indicated by the thrust chambertemperatures and the associated fuel injector temperatures shown inFigure 7-11, The LHp chilldown system performance far second burn wassatisfactory. The LAz pump inlet temperature at second burn ESC was 21.9°K(-420.6°F). “Second burn LOX pump chilldown was also satisfactory. AtS-IVB second burn ESC the LOX pump inlet temperature was 91.7°K (-295°F),

The start tank perfonned satisfactorily during the second burn blowdownas shown in Figure 7-12. The proper energy input to the turbine wasprovided for-a smooth start. Since there was no ignition of the mainchamber for the second start, the start tank did not refill.

Figure 7-13 shows the predicted engine helium control sphere pressurecompared to AS-501 for the restart attempt. After the normal post test

tf * i _ TeTetebotFachate s

I ro

al

-| | 71d

i to

+ ur | i TE

ORTBAE Te9" mek Ta PLACE MOSITION [15 De67,ANE TO" GL Ta PLAGE POSITION |SBtcFigure 7-9. S-IVB Engine Start Tank Tenperature and Pressure - Coast PhaseTable 7-6, S-IVB Engine Start Sequence ~ Second furn Attempt

Command (SC) 11,403.61 11,614.69,

Mainstage OK 17,412.60 Not AttainedMain LOX Valve Open 11,413.58 11,625.20

90 Percent Phrust 11,413,3 Not Attained 7-16

2 TUNP TET PRE

oH, asa

x 8 z a

fnew = From¥ ae Were Atk

: } 2 START 3Ay .z sibs :BiG th oa £Ee ant

Fa 4225

3 x

as 424

" I L |es Bo %SUFT FRESSURE, Ben?

Lo Pam ET PRS. pea2 3 © “ ® 2

10 . 7

heati

foo 8

re\ 2

1 b-296 =

*, iB 4 38 Es

Loe PUM ISE7 PRESSURE, Neen

Figure 7-10. S-IVE Start Box and Run kequirements - Restart Attempt

400| 200

cei 300 ~ *. i = as-501 3°.22 ~~ EDC TEST be23 200 PTs 25ue NR ES -200E400 Sy t~ ae

‘ NS5 400g

} 200 5

g. 30p=pep. Pitt sege pat RL fe geif pertfeZe % 4 i 38ge | aeae I -200 3%arn S-I¥B ENG RESTART COMMAND, 11,614.69 =”S+1VB STDY OPEN, 11,622.92

° | | poyOO eeTIME FROM SECOND ESC, SECONDS.

| RANGE TIME, HOURS :MINUTES: SECONDS.Figure 7-11, S-IVB Thrust Chamber Chilidown Performance -Restart Atterpt

T-second blowdown the pressure was 1310 N/om? (1900 psia} at secondburn ESC, The pressure decayed more than the predicted 550 N/om2(800 psi) during the second burn fuel lead secause of the lower temperatureat the injector. The pressure stabilized at 670 N/cm2 (975 psia) atthe end of the blawdown. This pressure would have been sufficient tocomplete the mission.

During the start of the second burn, the GG experienced a tenneraturespike as indicated by Figure 7-14, ‘The spike measured 1278°K (1840°F) ,the upper limit of the temperature bulb, but an expanded plot of fuelturbine inlet temperature suggests the spike actually reached as highas 1389°K (2040°F). Tne spixe resulted from a high start mixture ratio‘in tne GG, which in turn was caused by the failure of the main chamber

300 ;

£2 1200aoIe

Ves-1ve eng nesTaRr command, 11,614.69VS-LVB STOY OPEN, 11,622.92 F a00SD INATIATION OF START TaNic REFILL, 11,622.79

su

$400

|

—__- con a

W gBao | 32 : aoe

!200

i/ 200roo }

z 7 € 5 vi 12 v4TIME FROM SECORO ESC, SECOROS

Wot. A r Uw nlTiassa SS a ani aE sn Ie

RANGE TEE, HOURS :MINUTES :SECONDSFigure 7-12, S-IVB Start Tank Pressure - Restart Attemst

pressure to rise above 27.57 N/om2 (40 psia). With a low main chamberpressure most of the flow destined for the GG follows the lower pressuredrop path to the main chamber, resulting in a low GG total flowrate.However, because the start load of the oxidizer pump is lower than thatof the fuel pump, the initial oxidizer flow is less affected than thefuel flow. Tnus, the GG chamber pressure, a function of both totalflowrate and mixture ratio, is low as shown in Figure 7-15; whereas, thefuel turbine inlet temperature, a function only of GG mixture ratio, ishigh,

Near engine cutoff command, the GG chamber pressure rose slightly. Thisfact plus the absence of any unexplained vehicle moment during secondburn Suggests that the temperature spike did not burn through the 66combustor wall.

Information supporting the suspected ASI fuel line failure is discussed inthe following paragraphs. Table 7-7 lists the chronological events which

7-19

T8uC T

1409 :

WW E-18 CG RESTART coMfaND, 11,614.65] 29%YS-1VB STOV OPEN, 11,622.92

E1200 ~ { =

1008 — pe

= aon FS 2p= 1009

600 -

400z 4 # 10 12 "aTIME FROM SECOND ESC, SECONDS

2:13:36 3) 40 3:19:42 3:13:44 3:18:46 3:13de RANGE TIME, HOURS:MINUTES :SECONDS

Figure 7-13, S-IVB Engine Control Helium Sphere Pressure -Restart Attempt

Jed to the failure to restart. The first indication of a problemwasengine area chilling of the MOY closing control line at 645 seconds as shownin Figure 7-16. The second indication was the apparent performance shiftsindicated during first burn as discussed in paragraph 7.5. During theperformance shift at approximately 688 seconds, the yaw actuatorexperienced an abrupt increase in its cooling rate that cannot beattributed to radiation Toss. This may have been cue to the impingerentof a cold gas on the actuator at this time and is discussed in detailunder paragraphs 8.4 through §.6.

The third indication of a problem during first burn operation was theapparent flash fire at 696 seconds. The thrust structure temperaturesdecayed from approximately 222 to 216° (-60 to -70°F) by 696 seconds.At this time, the thrust structure temperature No. 1 showed a rise ofapproximately 16.7°K (30°F) as shown in Figure 7-17. The heat flux requiredto obtain the rise rate shown is approximately 1.1 watts/om® (1 Btu/ft2-s)which was approximately eight times that which can be produced from solar

7-20

1408CALCULATED SPIKE. REACHED ——] L200

1 3a0°R (2090°F) AN200 | 47Bs18 ENG RESTART COPAAND, 17,618.69 :veoes-va sTOV ORE, 17,622.92 TN a1200 Bes-ave Ecocommatio AD START OF T7, 11,690.39 UPPER LIPAT OF “gopTRANSDUCER RANGE1100 J _-

# 3400* 006 7

= 1200 22 960. —— é

* suo 1 eeo

400 ~+ 200

Vv 200oe ee

TIME FROM SCCOND ESC, SCCONIS

Bis36 3134) B15ae 13:48 Bibbs3:13:38 3:13:42, 3:13:46 3:13:50, 3:13:54

RANGE TINE, HOURS:MINUTES SECONDS

Figure 7-14, S-I¥B Fuel Turbine Inlet Temperature - Restart Attempt

heating; therefore, a heat source other than the natural environment

was indicated. Figure 7-17 also shows an interstage gas temperatureprobe which showed a temperature rise at the same time. Its responsetimes were smaller and excursions were larger; however, it does confirmthe presence of an additional heat source, Other measurements in thesame area also confirm these findings. Figure 7-16 also shows @ fire in-dication between the LOX punp and the start bottle. At approximately702 seconds these measurements indicated a cooling rate higher than wouldbe expected from radiation to space and the LOX tank, indicating the pos-sibility of a cold gas impinging in the area. After the performanceshift, the engine continued to operate at that level until ECO with somechilling in the area of the fires however, a normal shutdown was stilobtained.

7-21

#00 ZS-1WB ERE RESTART COMMEND, 117674, 65S=1VB STOV OPEN, 11,622.93S18 ECO conmiatn AMD START OF Ty, 12,630.38. 4oo20

L as0

200 300

g250

oso zg3 200 &si i2 A vo 2

nae am

0

[ 50

° hrame-bineerl °

50#0oe 6B eae

TIME FROM SECOND ESC, SECONDS

31338 Tard2 a6 B1s50RANGE TIME, HOURS:HINUTES:SECONDS

Figure 7-1, S-IVB GG Chanber Pressure - Restart Attempt

During the attempt to start the engine for the second burn, no temperaturechange was noted on the thrust structure temperature No. 2. However, thethrust structure temperature No. 1 dropped from 194°K (-110°F) at 11,622.92seconds (the end of fuel lead) to 190°K (-118°F). The engine main LOXPneumatic line surface temperature dropped from 205°K (-90°F) at 11,623seconds to 169°K (-155°F). The GG fuel inlet line wall temperature rose

7-22

Table 7-7. S-IVE Stage J-2 Engine Failure to Restart

Tanne TOM, _ae10 even snsuvsissrrza

|

eraine start trirst curr} orm{L¥9C Commaed}

585

|

start of engine corvartmest

|

rymaenic fuel Teakage ¢ upper engine area,temperature decraese ” * *Sat

|

sritist perterance decay begins

|

Increased Yeacage in unser angie areaGhas chaober prersure Tons) oe

fr

epson arkWE) engine. corpartaert tegenture dncreasing mere rapid’)628

|

taw actuator hed abupt tne Lwoingement of & cold gre on tre ecteatarcrease tn tts cooving’ ote#92

|

Second serfornance decay seqroe

|

Ferformance lose cue to leakage of spproxinatelyie p5* crember pressure loss;

|

E-Y Yon/s propellants.” (possible path LOX and tet}896

|

Engine congartment heaving i=

|

Fotsible LOX and fued teakage from 851 Tinesvicinity of 451 fuel Tine and *Hire indication setyeen totpure and start bett ¢702

|

rertorance stsrt erds with

|

AST chanbor erosion coneleteenera! cool ira77.08

|

engine cutoft Cvelacity eutott)

|

nomat11,622.52

|

tngena farts to ionte Lack ofASI fanttion,(S800 opereg) from 26 to 30°K (413 to -406°F) at 11,618 seconds and then dropped backto 28°k (-410°F), The engine main LOX valve actuator skin temperaturedropped slightly at 11,625 seconds. These locations are shown an adiagram of the engine in Figure 7-18.

Engine reconstruction analysis and engine tests indicated that the initialperformance drop during first burn was probably caused by leakage of theASI fuel line and is substantiated by the chilling experienced near theLz pump as summarized in Figure 7-18. The engine tests simulating anAST fueT line fatlure were compatible with the AS-502 observed performanceshift, a subsequent stabilization and a normal shutdown. One testconducted at an AS! chamber mixture ratio of 7 resulted in injectordamage simulating the performance changes recorded on AS-502 flight.

The propulsion system met all operational requirements during first burn,cutoff transient, and orbital coast. The conditions for restart werenominal, except for a hot start af the GG and lack of main chamberignition as previously discussed. Figure 7-19 is also presented toverify that pump operation and performance was acceptable up to thetime for main chamber ignition. However, the failed ASI system causedthe main chamber not to ignite, resulting in the failure to obtainrestart for second burn.

7-23

AST FUEL cMLET

Pie. Stonl /eR 2uye

LONE 3

810. PUN accEL:REMAINS HORMEL

S-1B FROST CoNhYORAMLICAACCSIMULATOR,

TIME &

ToHE 8.

EIST SiGN OF CHILLOOMN oxcueS8 "Vov CLOSING CONTROL LINEOCCURRING AT 64S. SECRHDS

AST FURL LOH

Figure 7-16. 5-IVB Engine Environmental Changes During J-2 Cngine Anomaly

7-24

3 320)

5

ENO

THRUST

STRUCTURES

TEM

8 &

PIULECLES

7-25

675

695

ar

738

655

250

zan,

8

VELOCITY

CUTOFF

COMMAND,

747.04

Figure

7-17.

LOX VALVE ACT.SKIN TEMP 7

8 UEL66 VALVE INLET LINEPOSITION WALL TEMP:

G6 BOO"STRAPLINE TEMP

cox waure CoePNEUMATIC “CARE

FUEL 1URBMANIF ExTWALL TCMP

TEMPFUEL TUREMAIN E¥D_ PUMP COLL EXTDISCHARGE TEMP MALL TEMP

HYD PUMP. IRLETom AST FUEL LINE

Lox suepLyLINE: FeMp

THRUST STRUCTURES:NO. 2 TEMP <I CAPS NO. 1

ifENGINE

Yad ACT CYL averent

\

THRUST sTRuc-OIL TER | Tee TURES. TENOxID ANTE |GAS. INTERSTAGESUPPLY TEMP

|

AREA TENP PITCH ACT. CrOiL TereFUEL TX PRESS.

CONT MCDULE TENPFigure 7-18, S-IVB Summary of Environment Effects

7-26

SONS Fealcomic 280 STSeT OF Tp, 1,630.32

SIV) E86 HESTAMT uUestRML 11,614.4815 STV CPE, 11 4672.92

2 won Fw Peo 3; sor con rouE22Ls -ee leBey wor saweete 5| ivsop bp e &Po oAA] ueeigen FueepBad 2 “ 1 i 0SBoa fnenhI 4?

ee!

a o a ic 12 a 16. W 20INETra Secon ese, SECTSVe aires aa

Pane TIE, ats areseects

oo fPnte peaa ta 4 us1 | iy Rowe se &

Lily hus Soren 3

Sob pot a2 8

Soa Nj ofeaoa a ee

« 2.05) -{ 28

og 4s Bw i Wore Ww 2TINE FROM SECOND ESC, SECONDS

Biinse_ulaedg HASe Bah SsePRUGE TUNE, 11OuRS:MENLTES: SFCOKDS

Figure 7-19. $-IVB Pumps Performance During Restart Atterpt

7-27

See

The ASI is a small chamber which is center-mounted in the thrust chamberinjector. Its purpose is to create and maintain a small ignition flanefor thrust chamber ignition. The AST system is shown in Figure 7-20.Figure 7-21 sheus the location of the ASI chamber in the engine injector.At the present time a modification af the ASI propellant feed lines tsbeing made as a result of the flight data and the engine tests made tosimulate the failure to restart. A complete report (Rocketdyne fio. Re 7450-2) titles "J-2 Engine AS-502 (Apollo 6) Flight Report - S-IT and S-1VBStages" will be issued covering the flight performance, 5-11 failures(common to 5-1¥8 failure), S-IVB failure, verification testing, and thedesign and modification changes.

7.7 S-IVB MAIN STAGE PERFORMANCE FOR SECOND BURN

(Ne discussion due to failure to achieve restart.)

7.8 $~IVB SHUTDOWN TRANSIENT PERFORMANCE FOR SECOND BURN

(No discussion due to failure to achfeve restart.)

7.9 S-IVB STAGE PROPELLANT UTILIZATION

The privary function of the PU system is to assure simultaneous depletionof propellants by controlling the LOX flowrate to the J-2 engine. Bycontrolling the LOX flowrate to the engine, indicated propeltant loadingerrors and/or deviations from predicted vehicle flight behavior can becorrected and the proper proportion of LOX and LHz in the main tanks canbe maintained. Other functions accomplished by the propellant utiliza-tion system are:

a. Controlling propellant loading

b, Maintaining the propellants at any predetermined level during launchcountdown.

Providing propellant mass indication signals to the telemetry system,

d. Providing a signal signifying the depletion of either propellant,thereby initiating ECO.

Components in this system include ground and onboard electronics, contin-uous capacitance probes, a PU valve, and discrete liquid levat sensors.

The PU system successfully accomplished the requirements associated withpropellant loading and management during burn, The best estimatepropellant mass values at liftoff were 88,060 kilograms (194,140 thm)LOX and 19,254 kilograms (42,448 Ibm) L¥2 as compared to predicted massvalues of 87,655 kiTograns (193,246 Ibm) LOX and 19,268 kflograms(42,479 Ibm) LH2. These values were well within required loading accuracies.

7-28

pasa EARL “ANE NG

a—3:euine

Set bEAS11468 LEU pane

a

Figure 7-20, $-I¥B J-2 Engine ASI Schematic

MOUNTING BOSS ORIFICE PLATEFOR ASI ASSEMBLY

PLATE

MANIFOLD:

Nast CHANBER

Figure 7-21. $-I¥8 J-2 Engine Injector Schematic

7-29

A Gomparison of propellant ass values at critical flight events, asdetermined by various analyses, is presented in Table 7-8, The LectPérimate full load propellant masses were 0.45 percent higher forLOX and 0,07 percent lower for Lip than the predicted values, as shownin fable 3-4 of Launch Operations, Section 3. This deviation mes net]iithin the required Toading accuracy, The iarge dispersion in firesbum cutoff propellart mass values from predicted was a result. of areextended first burn tine needed to achieve orbital insertion consitiensfozJouing the early shutdown of twa engines on the Sell stage, Figurecaineons 2 Sraphical representation of the FU nass Sensor nonl inouri tiesduring S-IVB powered flight.Figure 7-23 is 2 graphical representation of the best estimate, aGatistical weighted average analyses, for ignition and cutott nesses,The third flight stage ignition tiass was 160,690 kilegrane (354,261 1bm)and the cutoff mass was 120,330 kilograms (265,283 Ibn]. The ticepose cement systens used in determining the statistical weighted averagebest estinate masses were:a. PU indicated (corrected).

b. Flow integral,

. PU volumetric,

d. Level sensor.

€. Trajectory reconstruction,Extrapolation of propellant level sensor data to depletion, using thePropellant flowrates to depletion, indicated taat a LOX depletion wouldhave occurred approximately 2 seconds after second burn velocity cutof?,The first burn PU valve positions are illustrated in Figure 7-24. Duringfirst burn the PU valve was positioned at null for stare and renained thereuntil FU activated at first burn ESC +8 seconds. The PU velve was thencommanded to the fully closed (high EMR) position at activation and itrenained there until ESC +171,9 seconds (749.18 seconds).Fyen though the S-IV8 PU system functioned normally during CountdownDemonstration Test (CDOT), preflight operation, and first burn, ananonaly of 100 percent indicated LOX mass was experienced during the secondgreital revolution. The LOX mass bridge experienced disturbaneas ny ninedifferent eccasicns which caused the LOX bridge to slew toned akefull stop. On each occasion the bridge subsequently recovered except forthe last distrubance at 11,091 seconds of flight when the LOX moos bridge

7-30

Table 7-8, S-1¥B Stage Propellant Mass History

Fe HOLST HT deeb] ta NTETRAL |ose

2g tet ybss teiteaeihmai ty |ROE Gs Tae Talo ea

Lig INDLCATEG MASS, 102 Ibeaw x # 4 % 38 “0 2»

| afx = =: soBaa a5 LE {HOLCATED-ACTUAL210 NORHALLZED TD IGUITIONAND CuEOFE MASSES a

° ,ei BAe ao WS wo es eG eS TS Ted ET Toe Lg INCTGHTED MASS, 103 by

Loe tvoresren mass, 194 Thm

” 15 i 2050 ; +200

2-45 180 =i | 3= -+0] :z weg2 as] : -| =Fi I -0 22x0 zWL= INOTEATED-ACTUAL \ 2a6 AORELIZED TO [oNrTION “0fing COTOFF MASSES Tt

oa0 55 488570 75 BO BEDLOX INDLCATED MASS, 194 kg

Figure 7-22. S-1¥B PU Mass Sensor Volumetric Nonlinearity

7-31

tnawe N

EAWITY,

CUTOTT FAS, 105 abe

284 265 266joa t 7 rATTSNUT ROW: 968,701 2476 om

1at cattte:

|

269,283 2400 Ibmwre

POINT LEVEL SENSOR MASSES —— }

we. 358

bes” esrimire ehveLore!BEST ESTIMATE NASSES i

180.4

2 Pl VOLUMETRIC MASSES a Yo a2 Ni 7 35 ets =g ‘ie a5

iW Ri 2Flow sirécent, —| LA

|

z07 +160.4

rrasccToRY160.2 AECONSTRUCTION7

PU troicatea 383{ CORRECTED)10.9 isses

10813 120.0 WF e080. 120.8 TeN.0CUFOFT MASS, 103 kg

* This best estimate 1s determined by a statistical weightedaverage methodFigure 7-23, S-IYB Ignition and Cutoff Best Estimate Nass* ~First Burn

slewed to the full mechanical stop and remained there for the remainder ofthe S-IVB mission. Had second burn restart been attained with the PUsystem in the malfunctioning mode, the engine would have operated in thehigh EMR mode until velocity cutoff. The erroneous 100 percent LOX massbridge indications are shown in Figure 7-25,

At this time there appear to be two possible causes for the PU systemanomaly. These causes are:

7-32

40

S 30

S 205

a SeI¥E ENG START SEQUENCE COMMAND, 577.28= io ACTIVATE S-IVB PU SYSTEM COMMAND, 585.3Z VALVE REACHES TOP STOP, 869.0> PU DEACTIVATES, 749.2

o I

0 100 200TIME FROM FIRST ESC, SECONDS

—575 80C 625 650 675 700 725 «780 775

RANGE TIME, SECONDS

Figure 7-24, $-1¥B PU Valve Positions - First Burn

a. An intermittent cable shield between the ass probe and the PUelectronics assembly as indicated by an “X" in Figure 7-26.

b. Metatlic debris of some type in the LOX tank which caused a shortbetween the inner and outer elements of the LO PU probe. Figure7-27 shows a cutaway of the PU probe and the possible failuremodes.

Debris in the tank during orbital conditions could be distributedanywhere in the tank and possibly lodge between the probe elements.Since the PU system operation was normal during powered flight whilethe LOX mass probe, its essaciated cable and PU electronics assembly,were under the highest vibration levels experienced during flight, thepossibility of an intermittent cable shield appears remote. Therefore,the most probable cause of the PU system anomaly was metallic debrisin the LOX tank shorting the inner and outer element af the LOX probe,thus causing the LOX bridge to slew at a maximum rate to the “over-fill"condition.

7-33

LeSRaRse tes

Woy Wes Se1OGE SLeye 19 108 PIGrPOS-T:ON AND SEWEINED THERE, 71,09

Loe FINE mass.Besa

jegeo eat Styne W107 Tyce 10s ne ana TeeBRE THE, SEooNeSAL 1 \ WY Ly JBOLO GH gosa] OLS saeae sas plage

ANGE TINE, HOURS MERaES SESONnSFigure 7-25. S-IVB LOX Coarse and Fine Mass Data

Additional tests will be conducted to check the crimping of the cableconnector. The desirability of insulating one or both of the probeelements to prevent shorting by debris is also being considered,

7.10 $~1VB PRESSURIZATION SyS’

The functions of the S-IVS pressurization system are to provide thenecessary positive pressure to the propellant pumps while also supplyingan increased structural capability. The system consists of tank ullagepressure regulators and yent valves. LOX and LIZ recirculation systemsprecondition the propulsion feed system prior to starts.

7.10.1 S-IVB LH) Tank Pressurization

Tae fuel pressurization system provides tank pressur“zation by threemethods. Prior to launch, Gaseous Helium (GHe) from a ground sourceis used. After S-IVB engine start, for both first and Second burns,GH2 for LHp tank pressurization is bled from the thrust chamber hydrageninjector manifold. During orbital coast (parking orbit} seven LHz tankrepressurization GHe storage spheres, attached to the thrust structure,

7-34

hiss reg

nea8, emer ADDLST

co suet

TRIS‘quronarunsELC:LOW

iatTrecoaucePRC oMF omeaM

NPT

Ne ’

1-595] sean soxBe stewunToR

15 We © ile !

195 wwe of 'Ly ram ass

—— 4p trim # Lennie:nse: oS :

b_, mmuzariona CIRCJIT AYA om ®he 1a aRCOGE F/8 >)

ierFigure 7-26, S-IVB Servo Bricge

siuune woe EMOrTION aR:WILE 90 Tepani? Teo |oerShnomee rob ruE Stone 10.0 stuse sweetlol 2 5 Hlonreb toe |BerE Mons FA |TLB Sioned v0.8 |FL ceroomue rartste)tik Wurst 3 eer a3 et ean

a Low 2 & ren iL

TER ELEMERT

Skee Figure 7-27. Failure Modes of S-

7-35

I¥B PU Probes

supply repressurization gas to the LHz tank to meet second burn enginestart requirements.

The LH2 pressurization system operationally met ali engine performancerequirements. The LH2 pressurization system indicated acceptableperformance during S-IVB first burn, coast phase, and second burnattempt. The pressure measurement deviations experienced on AS-501,within the CVS and LH2 pressurization system, during orbital coastwere not experienced on this flight. Relocation of the CVS transducersgave acceptable readings. The sequence of events and associated systemperformances are discussed in the following paragrashs. The Liztank prepressurization command was received at -36 seconds. The Liztank pressurized signal was received 21 seconds later when the Liltank ullage pressure reached 23.3 N/om2 (33.8 psia). The ullagepressure continued to increase, reaching 25.0 N/ome {36.2 psial atS-I¥B first burn ESC as shown in Figure 7-28.During first burn an LH2 non-propulsive vent occurred at ESC +7 seconds(584 seconds}. The average pressurization flowrate was approxinately0.28 kg/s (0.62 Ibm/s}, providing a total flow ef 46.4 kilograms (102ibm) during first bum. A slight downward shift in pressurizationflowrate occurred as a result of the engine performance shift at 696seconds.

During the repressurization period the LHp tank was pressurized from14.3 to 23.3 N/cm2 (20.8 to 33.8 psia}. The ullage pressure subsequentlydecayed, reaching 21.7 N/om2 (31.4 psia). At this time, (ESC -16 seconds)a makeup cycle was initiated, increasing the Lllz ullage pressure to 22.6Nyené (32.8 psia) at second burn ESC as shown in Figure 7-29. Approximately20.0 kilograms (43.9 Ibm) of ambient helium were used in the repressuri-zation operation.

After the stage failed to restart, a decision was made to vent the LH2tank in an attenpt to gain additional deta pertaining to stage safing.The LHz tank vent valve was commanded open at 22,024.21 seconds as shownin Figure 7-29,

The LHe pump inlet NPSP was calculated from the pump interface temperatureand total pressure. These values indicated that the NPSP at first burnESC was 10.8 Name (22.8 psia). The NPSP then decreased during poweredflight to a minimum value of 8.3 N/omé (12,0 psia) at first burn ECO.At the minimum point, the NPSP was 3.9 N/omé (5.6 psi) above the required.Throughout the burn, the NPSP had satisfactory agreement with the predicted.The NPSP at the end of fuel lead prior to second burn was 8.0 N/eme (11.6psia) which was 3.7 H/em2 (5.2 psid) above the required. The NPSPincreased rapidly after ESC such that it was above the required leval

7-36

7 24a 8

#25 ws: | | g= 20 Ye PREPRESS INITIATED, -96 30a Wy [Ho TERMINATION OF PREPRESS, -75 =8 FIRST MOTION, 0.38 ySs i S-1WB ENG START SEQUENCE COMMAND, 577.28 4,3 I Ss20 Ver | Js

200-100 0 Too 200 goo. 400. 500 600RANCE TIME, SCCONDS

~ T—T TT T_T 405 S-I¥B ENG START SEQUENCE COMMAND, 577.28 a

S26 |] S-I¥B VELOCITY CUTOFF COMMAND, 747.04 a. | I | vo2 : ACTLAL DATA 36S

Be — { a 3

2a PREDICTION poe “

= i =320 4 em i 4 5

2 ry be718 ¥Y | 1 L | -

550 5/6 590 630 630 680 670 590 710 730RANGE TIME, SECORDS:

5 30 TT G= LHg TANK CONTINUOUS VENT VALWE OPEN ON, 806.25 40a

= 96 Talik BLOWDOWN TO CYS RANGE, 871.25 doy

B20 | wg

g's el ado =o]=E== 20 S

s10 i 23 750 1750 «2750 3750 4780 5750 6750 7750 8760 9750 10,750 a

RANGE TIME, SECONDS

0:30:00 00:00 1:30:00 2700700

RANGE TIME, HOURS :MENUTES:SECONDSFigure 7-28, S-IVB LHz Ullage Pressure - First Burn and Orbit

730700 3:00:00

7-37

Lig TANK REPRESS CONT2O1 YALE OPEN ~ OM. 11,487.70Lig Teak REPRESS TERMIWALION, 11,526.69S1vB ENG RESTART COMMAND, 11,614.69S.I¥B ECO COMBAKN, 11,830.33Lig TAUK VENT VALVE OPEN, 22,024,2<ddidg

30 a3 I 08Bs gg pa 32 a rb» €

2! 2 =So L a11,350 i140" 11,459.00 175550 11,650 17,850ANGE TIE, SeooNDS

sea Ye: __sTa70 bo10 RANGE TIME, HOURS:SNUTES:SECONDS£ wk3

gs THK VENT a8. - ne

2” 4 ie 3S10 811,609 21,680 11,700 71,750 11.800 11,860 17,909 11,880 12,000RANGE TIME, SECONDS

SieoieEBLE 70:00ANGE TINE, HOURS: INUTES:SELONDSEn 22 tUPCRESTART ATTEWPY ajuly - a3 Vows PA a 82 10 aaa 2

g \ 10g3 324 o =11,0007 78,000 18,000_ 23,000 27,0007 aT cooRANGE TIME, SECONDS

aEC a 6:50:08 T3050) BraOT00RANGE T:ME, HOJRS: MINUTES SECONDS

Figure 7-29. S-IVB Li2 Ullage Pressure - Restart Attempt

7-38

during the attempted restart. At second burn ECO the NPSP was 6.0 N/cm2(8.7 psia) which was 1.6 N/cm? (2.3 psi} above the required. The pumpinterface total pressure at the end of fuel lead was 22.5 N/emé (32.6psia). Figures 7-30 and 7-31 summarize the fuel pump inlet condittonsfor first and second burns, respectively.

7.10.2 S-1¥B LOX Pressurization System

The LOX pressurization system provides tank pressurization by threemethods. Prior to launch, GHe from a ground source is used. Afterengine start, for either first or second burn, GHe from eight cold Gilestorage spheres located in the LHp tank is warmed by a heat exchangerand used for LOX tank pressurization. During orbital coast (parkingorbit} ambient GHe, from two high-pressure LOX tank repressurization GHestorage spheres, is utilized to supply engine restart ullege pressurerequirements for second burn.

The oxidizer system performed adequately, supplying LOX to the enginepump inlet within the specified operating limits throughout first burnand responded normally during the attempted restart. The availableNPSP at the LOX pump inlet exceeded the engine manufacturer's minimumat all tines

LOX tank prepressurization was initiated at -166.5 seconds and increasedthe LOX tank ullage pressure from ambient to 28.3 N/cmé (41.1 psia’within 16.3 seconds as shown in Figure 7-32. Two makeup cycles wererequired to maintain the LOX tank ullage pressure before the ullagetemperature stabilized. The pressurization contro] pressure switchcontrolled the pressure between 26.3 and 27,7 N/cm® (38.2 and 40.2psia). At -98 seconds the LOX tank wllage pressure increased from 27.2lo 29.7 N/em@ {39.8 to 43.1 psia) due to fuel tank prepressurization, LOXtank vent purge and LOX pressure sense line purge. This caused the vent/relief va.ve to open, dropping the pressure down to 29.0 N/eme (42.psia). The pressure remained at this level until liftoff

During $-iC boost there was a relatively high rate of ullage pressuredecay necessitating tuo makeup cycles from the cold helium spheres ashown in Figure 7-32. This decay is partially due to utlage coolingand partially due to tank geometry response to the vehicle axiaacceleration. There 1s, correspondingly, a step rise in pressureproportional to acceleration changes at S-1C Inboard Engine Cutoff (TECOand Outboard Engines Cutoff (OECO}. There are also indications thatthe tank vent/relief and/or relief valve may have experienced intermittentleakage during the period from 105 seconds to S-IC cutoff. Althoughdata are not conclusive, it appears that approximately 0.18 kilograms{0.4 bm) of ullage gas was lost from the tank during this period. Ullagepressure at S-IC staging was 28.8 N/om2 (41.8 psia).

7-39

$-1¥B FIRST BURN STDY OPEN, 580. 82S-IVB ENG START SEQUENCE COMMAND, 577.28

S-IVB VELOCITY CUTOFF COMMAND, 747.04

20

¥ 25E15 + 7 |3 ACTUAL 22& 10 | otis i

hic &S$ bose 7 s}=Sse st St y

XL bs 3REQUIREDa 135 ro

3 2a#2 30/—4 - SzeBy weays ak3p 2 A poy2 PREDICTION Lo, =

5 la

. E -418Ge 23 - be

re L. 229 +8Ee 2 - ee55 esSE | 422 9SB a Ee 422 25

20 oa 406 BO too s1z080 608

TIME FROM FIRST ESC, SECONDS

yp600 62550 SCGS 700 725 750

RANGE TIME, SECONDSFigure 7-30, S-IVE Tuel Pump Inlet Conditions - First Burn

7-40

S-IVB ENG RESTART COMMAND, 71,614,69-IVB STCV OPEN, 17,622.92

S-1VB ECO COWMAND AND START OF Tz, 11,630.33

20

~ b 25515 i =2 pe oe& 10 - — 1 a

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¥ L -18te 23 of aa {| ZuEg | L -g2ogeSe 22 ‘ Ba

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TINE FROM SECOND ESC, SECONDS

v ¥v. 13:13:36 9:13:40 3313548 3:13:48) 3:13:52

RANGE TIME, HOURS:MINUTES :SECONDSFigure 7-31. S-IVB Fuel Pump Inlet Conditions - Restart Attempt

7-41

30

hyNeL ao

~ 2 Yg LOX PREPRESS INITIATED, -166.55 Lip TANK PREPRESS BAND, -96 to -75Bou AZ NARE-UP CYCLES 1 ANO 2°DURING S-IC BOOST, 80 AND 1252B oof LW S-1C/S-11 SEPARATION COMMAND TO FIRE L 30ee SEPARATION DEVICES & RETRO MOTOR, 149.088 ‘YS-11/S-IVB_ SEPARATION COMNRND TO” FIRE SEPARATION3 DEVICCS & RETRO MOTOR, 577.083 YVS-1¥G ENG START SEQUENCE COMMAND, 577.28x 15 ~ -3 20

19,200-100 T0200 3000 «00

RANGE TINE, SECONDSye

a b+ oe£5 a5 f-—+Bo PREDICTED3 WVS-[¥B ENG START SEQUENCE COMMAND, 577.283 on Y_W75-1uB Veloce CUTOFF conan, 747.04" GP 30S “550 570 590 610 630 50670 so0_710. 730750

RANGE TIME, SECONDSwpe5S . —ela Roope > fl rs 40ge t “st d£8 as 4

S PATTITUCE PITCH MANEUVER, 3207.303 WATTITUCE PITCH MANEUVER, 5427.30= oot tt F 30a 750° (1780 2750 375) 4750 5750 6750 7750 8750 9760 10750

RANGE TINE, SECONDS

0:30:00 Fahrer Tate ee aaah 700:00

Figure 7-32.

RANGE TIME, HOURS: MINUTES : SECONDSS-IVB LOX Tark LIlage Pressure - First Burn and Orbit

UCX

ULLAGE

PRESSURE,

LOX

ULLAGE

PRESSURE,

ILLAGE

PRESSURE,

alLOX

psia

psia

psia

During the S-IT boost there was a continuing pressure decay, but at amuch reduced rate, No maxeup cycles were required. Although ullagecooling continued during thts period, the major cause of the decayagain appears to be response to the vehicle acceleration; thereis a step rise in pressure at S-II engines No. 2 and 3 out, and atS-II cutoff.

The LOX tank sIlage pressure was 28.8 N/om? (41.8 psia) at ESC,satisfying the engine start requirements as shown in Figure 7-32During the start transient the ullage pressure decreased to a minimumof 24.2 N/on2 (35.1 psia) before the pressurant flowrate became Targeenough to increase the ullage pressure. During burn the ullagepressure cycled four times. The ullage pressure was sufficient to exceedthe minimum NPSP requirements during powered flight

The LOX tank pressurization flawrate variation was 0.16 to 0.18 kg/s(0.35 to 0.40 Ibm/s) during aver-control, and fram 0.11 to 0.14 kg/s(0.24 to 0.37 Tbm/s) during under-contral system operation. Thisvartation is normal because the bypass orifice inlet tenperature changesas it follows the cold heliun sphere temperature. Heat exchangerperformance during first burn was satisfactory

The LOX tank ambient repressurization system satisfactorily raised thetank ullage pressure Fron 25.7 to 27.7 W/om2 (37.3 to 40.2 psia) i19.7 seconds. Helium consumption during repressurization was 2.4kilograms {5.3 |bm) with the sphere pressure decaying from 2124 to 1327N/em2 (3080 to 1925 psia}. The tank ullage pressure was 28.7 N/eri2(41.6 psia) at second ESC, satisfying the engine start requirements

The LOX tank pressurization system operated noninally considering theboundary conditions during the attempted restart shown in Figure 7-33.The regulator discharge rose rapidly to 276 N/cm2 (400 psia} and renainedat that level until the system was turned off at second ESC +16 secondsThe helium flowrate was 0.16 kg/s (0.35 lbm/s) and the total mass ofhelium used fron the cold helium spheres was 1.3 kilograms (2.9 Ibm)

The LOX NPSP calculated at the interface was 17.2 N/em® (24.9 psi] atfirst burn ESC, The NPSP decreased after start and reached a minimumvalue of 14.5 N/en? (21.1 psi) at 26 seconds after ESC. This was 0.10N/om2 (0.15 psi) above the required NPSP at that time. The NPSP therincreased and followed the LOX tank ullage pressure for the durationof first burn.

The LOX pump static interface pressure during first burn followed thecyclic trends of the LOX tank ullage pressure. Values ranged from24.1 H/ome (34.9 psia) at 26 seconds after FSC to 28.8 N/em2 (41.7 psia)imnediately after first burn ESC, The NPSP calculated at the engine

7-43

LOK TANK REPRESS CONTROL VALVE OFEW, OW, 11,367.69LOK TANK REPRESS TERMINATION, 11 AC}. 39)SoIVENG RESTART COMMAND, 11,619.635-1¥8 EO) COMMANE AND STARZ OF Ty) 11,630.23Lox Tak YENT VALVE ae, 22,083¢30

,-g [_— Log

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P20 i

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_

m6 —aE.a) So2rane TIME, SECONIS

8:08:00 zo be6:10:90 6:18:00 6:18:00 6:27:00

RANGE TINE, FOURS: MZMUTCS:SCCONOS

aL s8HUe1 remwrvarion|2 OF OECAY AND THENCorUkeen boo

a0)

Lox

ULLAGE

PRESSURE,en?

LOX

ULLAGE

PRESSJRE,

asta

a+000 75,000 99,000" 23,000.27. 3,000RANGE TIME, SECONDS

8:00:00 6:05:00 7500700 B:m0:004:30:00 6:39:00 Brae 7:30:00RANGE TIME, HOURS: MINUTES :SECONDS

Figure 7-33. S-I¥B LOX Tank Ullage Pressure - Restart Attempt

7-44

interface wag 18.5 N/on2 (26.8 psia} at second burn ESC, At all timesduring second burn attempt, NPSP was above the requirec level. Figure7-34 and 7-35 summerize the LOX pump corditions for the first burn ancthe second burn restart attempt, respectively. The run requirementsfor first and second burn were satisfactorily met as previously preserted

The cold helium supply was adequate to meet all flight requirements. Atfirst burn ESC the cold heliun spheres contained 151 kilograms (332 bm)of helium at @ pressure 2027 N/oné (2940 psia}. At the end of the166.52-second engine burn, the helium mess had decreased to 124 kilograms(273 Thm) at a pressure of $51 N/om2 (1380 psia}. Following first burnECO the cold helium sphere pressure data indicated a decay of approximately2.0 Néeméfmin (2.9 psi/min} as shown in Figure 7-36, resulting in anapparent pressure at second burn ESC of 593 N/om? (860 psia). The causeof this decline is still under extensive investigation; areas of interestinclude the cold helium system coroseals ard joints, the LOX pressuriza-tion module, and the cold helium sphere pressure transducer. Figure 7-37shows these arcas of interest. The conoseals will te changed to 775aluminum coated with teflon throughout the cold helium system for AS-503and subsequent vehicles. Conoseal re-torquing rcquirenents are also beingconsidered.

During the restart attempt the LOX pressurization system was activatedfor approximately & seconds, dropping the cotc helium sphere pressureto 872 Nfom? (830 psia}; mass usage during this period was 1.1 kilegrams{2.4 Ibm}, Following the restart attempt the sphere pressure datacontinued to decrease, reaching 379 K/em2 (55C psia) by 22,023.30seconds. From this time until loss of stage power the cold helium sphereswere dumped overboard through the LOX tank vent valve. inaccuracy inthe sphere pressure data (negative pressure data during the dump as shownin Figure 7-36 invalidates mass catcutations during this period.

After the staye failed to restart a decision was made to vent the LOXtank and cold helium spheres in an attempt to gain additional datapertaining to stage safing. Tre LOX tank vent valve and the coldhelium shutoff valves were commanded open at 22,023.30 seconds as shownin Figure 7-33. Tre cold helium pressure began a stracth blondown tozero psia and the LOX tank ullage pressure began ta blowdown from its 23.8N/on2 (34.5 psia) level. However, at 22,047 seconds with the ullagepressure at 16,5 N/om2 (23.9 psia}, the pressure decay abruptly terminated,and pressure remained constant until 22,061 seconds at which time theblowdown abruptly resumed and continued as long as data was available.During the period when the ullage pressure remained constant, both thetank vent valve and the cold helium shutoff valves remained open. Thereason for this plateau is presently under investigation.

There is « noticeable change ir the rate of pressure decay occurring at22,085 seconds which is coincident with all of the LOX liquid temperature

7-45

WY S-1V8 ENG START SEQUENCE COMMAND, 577.28S-I¥B FIRST BURN STDY OPEN, 580.52S+1VB YELOCITY CUTOFF COMMAND, 747.04

LOX

PUMP

INLET

6

2435

223 20 so ACTUAL 30

a” 18g 25= 16x n“14

r 202 = REQUIRED10 L b15

mS65

= 40 - + 60Z2 4, PREDICTIONS eo~2 30 FLIGHT DATA | — 45,

28 eT 40xe 8 + b 357 20 30

94ge thasge 90 ~ 23

ne 8 -3055SE og i 9 2 408010020. —«t0 160 180

TIME FROM FIRST ESC, SECONDS

875 600 625 650 675700 725 750RANGE TIME, SECONDS

Figure 7-34, S-IVB LOX Pump Inlet Conditions - First Burn

7-46

LOX

NPSP,

psia

Eypsia

LOX

PUMP

TOTA

LINLET

PRES

SURE

TEMPERATURE,

°F

LOX

NPSP,

N/cmé

LOX

PUMP

TOTAL

INLE

TPR

ESSU

RE,

N/cm

2ey

eKLOX

PUMP.

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TTE

MPER

ATUR

:a

18

15

le

45

40

8

30

26

20

94

92

90

a8

66

YVS-1VE ENG RESTART COMMAND, 11,614.69WW S-1V8 STOV OPEN, 11,622.93‘YS-1V8 ECO Command and Start of T7, 11,630.33

Rendon,

REQUIRED

17

J

30

25

20

15

60

50

F 40

30

+294

298

306

46 8 10 12 14 16 18 20TIME FROM SECOND ESC, SECONDS

Figure 7-35. $-IVE LOX Pump Inlet Conditions - Restart Attempt

740 713d Tis4B a: 137623:13:42 3:73:46 3:13:60 4:13

RANGE TINE, SECOKDS

TAT

154

LOX

NPSP,

psia

LOX

PUMP

TOTA

:IN

LET

PRESSURE,

psia

-302

OXPUMP

INLET

TEMPERATURE,

°F

LUM

PRESSU

RE,Wen?

LAGE

peSSPE,

Wren?

ax

SoINWLS START SLQUFHS cxMhaND, 577.28cIUecoctTy GOFF Coben, iy 04S18 ENT AESTAET CSOMBND, 1192863So128 ECO CoewAND au START Or 72, 14,686.21STAHL DF-cOlis HELIUM achey 72402838

fee gfe g

aes

: oe

a afy one

Ry aA Boe> 12,110, 16,000 29,209 24,000‘inte, sFeem$

DOE COS Taa-oe TREO TOHOnL TNE. HOSP YUTLsSEOMN

2308

er

eee COLD We SOHEKE eerssuREsoo a

carve 3200of | ressurt |RI PRLSSLRE DaREDICTED soe

=O Teak uueace eessuae | paeortre az = i F, «vu™egDnzTE

SiR 7 Teaco TT,c00" 11,608 17,700 Mesbo0 2, da

akge TIME, seconcs

D3. samt ao000 nese aero So00:00 g:09-09ANSE TIME, nouRS #anuneS se6ON=Figure 7-36. S-IVB Cold Helium Supply Decay |

7-48

HELIOM

PRESSU

RE,pat

a

measurenents converging at 91.7°K (-294.5°F). This is the saturationlevel correlating with 12,2 N/om2 (17.7 psiz). The ullage pressure wasindicating 12.5 H/en? (18.2 psia) at this time, suggesting that LOXboiloff is responsible for the change in the rate of pressure decay,Although a definitive analysis has not yet been performed, the lengthof the pressure decay (approximately 11,600 seconds were required to dropthe ullage pressure to 0.7 Nem[1 psia]) suggests that the vent pathjs highly restricted from its nominal 103 r2 (16 ine). At O.? Nene(1 psi} the flowrate through an opening of this size would be sufficientto vent the tank completely in 500 to 1000 seconds.Investigation of both the ullage pressure plateav at 16.5 Nom? (23.9psia} and the vent line restriction is continuing.

7.1] S+IVB PNEUMATIC CONTROL SYSTEM

The pneumatic contret subsystem provides supply pressure for all stagepneumatically operated valves with the exception of J-2 engine valvesbut including the engine start tank vent valve. A pneumatic powercontrol module regulates filtered ambient heliun flowing from theambient helium sphere loaded to a pressure of 2136 68.9 N/on2 (3100 4100psia) at 294°K (70°F). The module regulates pressure down to 338 £17 i/cm2(490 425 psia) for operation of the following:

a. Hg directional control valve during ground procedures.

b. Propulsive vent shutoff valve during powered fligat.

c. LOX and LHy 111 and drain valves during ground procedures.

d. d-2 engine GH start subsystem vent-relief valves.

8. LOX and LHp turbopumps turbine purge module.

f. LOX chiTtdown pump purge module control.

g. LOX and Lilg prevalves.

h. LOX and LHy chilldown shutoff valves.

i. LOX tank vent-retief valve.

J. LH propulsive vant valve.

The pneumatic control and gurge system performed satisfactorily duringal] phases of the mission. System performance was nominal during boostand first burn operations. The AS-502 stage incorporated the redesigned

7-59

pneumatic actuation control modules, and experienced no discernibleleakage as opposed to earlier stages which had significant degrees ofJeakage. Pnelmatic control bottle temperature. pressure, and regulatoroutlet pressure are shown in Figures 7-38 through 7-40, Battle massesat various pertinent times are shown in Table 7-9.

7.12 S-1VB AUXILIARY PROPULSION SYSTEM

The APS controls the vehicle attitude during S-I¥B operation, and position:the propetlants in the stage during first burn cutoff, parking orbit coast,second burn restart, and cutoff transient. Nitrogen tetroxide (N204) andtionomethy) hydrazine (MMH) are the APS propellants. These propellants arehypergolic and require no ignition system. The APS system is composed oftwo modules Iocated 130 degrees apart on the aft skirt assembly. All require~ments are suoplied from within the moduies except the electrical power sig-nals wnich are required from the stage. Each module contains three ablativelycooled, 667 Newtons (150 Ibf) thrust, attitude control engines; and oneablatively cooled, 311 Newtons (70 Tbf) thrust, ullage positioning engine.The attitude control engines control S-1¥B rolt during engine burn andpitch, yaw, and rol] during orbital coast. The ullage positioning engineFires to assure the presence of liquid propellants at the J-2 enginepump inlets during engine chitldown and restart, and’ to settle thepropellants prior to propulsive venting to prevent the loss of liquidpropellants through the vent systems

The APS pressurization systems demonstrated nominal performance through-out the Flight and met control system demands as required until APSpropellant, depletion. The regulator outlet pressures were maintainedat 134 N/on® (195 sia}. The APS ullage pressures in the tanks wereapproximately 132 N/ané (192 psia).

The oxidizer and fuel supply systems performed as expected during theflight. The propellant temperatures measured in the propellant controlmodule were as expected. The maximum temperature recorded was 317°K(110°F), The bulk temperatures of the propellants in the bladder rangedfrom 306 to 311°K (90 to 100°F). The propellant supply pressures werenominal at approximately 131 N/cm2 (190 psia) during the mission.

The APS engine performance was as expected with the exceptions noted inthe following paragraphs:

The propellants were depleted in both modules as a result of disturbancesinduced during First burn and after attempted restart. Because of thefailure to restart, the LOX and LH2 residuals were much greater thanpredicted, As a result of this condition and spacecraft separationthe vehicle's moments of inertia were larger, and the Center of Gravity(CG) further aft than predicted. With the CG Tow, the APS pitch andyaw moment arms were shorter than expected, resulting in longer APSburn times.

PHEUMATIC

SPHERE

PRESSURE,

N/en?

aTI

CSP

rER:

URE,

RAT

ERULATOR

EPR

ESSU

RL,

SCHAR

PHEUMATIC

2:

Bae PUMP PURGE {NITIATEC, 746.9ENG PUMP PURGE CFF. 1345-85

zea0

i)2mm 4

“ + 3000

2900 > 2900

b 28001900

oop TOMLbaon

1890 7 L 600

1790v5 120

+ t00309-4} ~ 0

Soe pm tee fe ei 8iv ro2ts L 40400

b $50x a

00 f 40me Yo00 ecco! 3c00) 4000 5000 6090

0:00:00 a

Figure 7-38,

RANGE TIVE, SECONDS

S-IVB Pneumatic Control Performance - First Burn

7-52

PRESSURE,

psia

ERE

UBTIC

SPH

anel

RATURE,

°FPREUMATIC

SPNE2E

NPTE

'SSURE

,PHELMATIC

REGLLATOR

TSCHARGE

PR:psia

HELI

UYBOFTLE

PRESSURE,N/

m?om

LIUM

BOTT

LEATURE

Hel TENDER

HELI

UMRE

GULA

TOR

PRESSURE,

N/em

d

1900)

2700

a1800 -— Sa e600

25001700 <p t-

2400

1600300 a0

[real40

aS - Pet

625009 1,287.79‘Y START OF Tg, 11,287. 560

bat | Le350 a “

480

300 aan6000 7000-000 9000 10,000 17,000 «12,000

RANGE TIME, SECONDS

1:45:00 2:00:00 2 02:45:00 3:00:00 3:75:00

RANGE TIME, HOURS:MINUTES: SECONDS

002

HELIUM

BOTTLE

PRESSURE,

psia

HELIUM

BOTTLE

TENPERATURE,

°FGULATOR

PRESSURE,

osia

HELIUM

:

Figure 7-39. S-I¥B Pneumatic Control Performance - Coast Phase

7-53

S-1¥B ENG RESTART COMMAND, 11,674.69S-1VB ECO COWMAND ANC START OF T7, 11,630.33

18002600

&3 1790 ~ ; 4 C8ge [petOL 2400 Bs

24 1600RE

1500 + 2z00280

f 40

ax pet Ea4 270 -~- = — Be

5S ps ge=e 2602s

250 “19at4By n | seoB33 my Tt Besa= [seo ga25 aso - #5"

258 [BEESBE 00 ay 838= 711,806 77,900 12,309 12,709 13,190 13,500

RANGE TT¥E, SECONDS

a 20:00 3:25:00 30:00 3:35:00 40:00 3:45:00RANGE TINE, HOURS: MINUTES -SEcONDS

Figure 7-40. S-IVB Pneumatic Control Performance - Restart Attenpt

7-54

Table 7-9. S-IVB Pneumatic Helium Bottle Mass

BOTTLE MASSTIME

kg Tom

Liftoff 3.87 8.54First Bura ESC 3.85 8.48First Burn ECO 3.85 8.483500 sec 3,60 7.946500 sec 3.57 7.86

Second Burn ESC 11,614.69 sec, (03:13:34.69) 3.43 7.56

Second Burn ECO 11,630.33 sec, (03:13:50.33) 3.43 7.5617,500 sec, (04:51:40) 3.42 7.54

The propellants in Module No. 3 (at Position 1) were depleted first asshown in Figura 7-41. The fuel was depleted at 21,953 seconds, while theoxidizer was depleted at 22,053 seconds. The fuel was also depleted firstin Module No, 2 (at oosition III) (22,602 seconds) as shown in Figure 7-42.The oxidizer was depleted at 22,634 seconds. The reason the fuel was deple-ted first in both modules was that the propellants were loaded for a 1.65 EMRto 1.0 EMR while the attitude control engines normally operate at a 1.60 EMRduring minimum impulse ait pulsing. This EMR was dropped even furtherdue to the high injector temperatures causing oxidizer vaporization andreduced oxidizer flow, Tne fuel load for the flight was maximum. Table 7-10presents the APS oxidizer and fuel consumption at significant events duringthe flight. The APS helium bottle oressures and propellant quantitiesderived from the helium bottle pressures are presented in Figures 7-43 and7-44,

The engine chanber pressures were normal and ranged from 64 to 69 N/cm?(93 to 100 psia) during the initial portion af the flight. However, atapproximately 6000 seconds, the chamber pressure of the first pulse in aseries of pulses on engine No. IT; was 41 N/cmé (60 psia). The chamberpressure increased to @ nominal value as the series of pulses continued.Figure 7-45 shows an example of this phenomenon. This initial Towchamber pressure has been attributed to a high injector temoerature asa result of heat soakback of the injector resulting from heavy APS dutycycles following first J-2 burn.

Engine No. 11] and IIlyy had similar low chamber pressures following thejong steady state burns on these engines after attempted restart. The

7-55

0 10530a Caubiced

0 © soe Fu,Dace Pee sionLEE wititioos wv1s |yeg70

8 untEAE Comin, 91128> ZA

|

Gata antiNPe o | gs2a mime teMaea eats Crs |Sz fo Rit arena. ms 20 =z LEB SEGEEDHae zae a50 52 so §Bo é

o 00 B00 1200 7,07 0028000ANGE TIME, 52CO¥DS

0:03:00 Tobe r00:0 —a:ebe00 300-00 Se00;00 —GF0D:36_PeatsO0TUNBE “INE, HOUAS-MINUTES-$2¢00S

Figure 7-41, S-IV8 APS Propellant Predictions - Module No. 1

xe20see obecat

xe We wikieigon ansoxrorzee0 160i oe com sigact ammo, 17 fsfacta igmate raeITE awwr 3 |g

_« HLRLatT | aS 3¢ srapcho! 2 il aneron. 2.608 :- ADS MOGLEWO PILEBET 2,64E so

a5gBx

®ja

ww

9 aC 005.90075,IE Teme, Sec0K05

oxo) Wade © Er Toso Feo{SE TINE, WOURS MINUTES SECOFigure 7-42, S-IVB APS Propellant Predictions - Module No. 2

Taos

7-56

Table 7-10. $-1VB AS Pronellant Consumption

MODULE AT POSITION T MODULE AT POSITION [11Tame Peaton ‘OxoTzER FUEL, OXTOTZER FUEL

ko (ibm) ka (bm)} kg (te) ka (br)

Initial Load 05.6 (188.5)] 56.5 (124.4)] 86.0 (189.4)}66,6 (124.7)

First d-2 Burn 0.7 (1.6) 0.5 (1.0} 0.7 (1.8)] 0.5 (1.0)Roll Contra]

J-2 ECO to End of 12.8 (28.1) 9.8 (21.6) 6.6 (1dE)] 5.0 (TL)Fiest APS UNaging

Ist and 2n¢ Earth 20,5 (45.3) 12.8 (28.2)}13.0 (28.7) 8.1 (18.0)Revolutions

Restart Preparatiors |22.9 (St.5} 17.0 (37,4)/23,4 (51,5) 417.3 (38.1)

Attempted J-2 28,7 (63.1) 16.4 (36.2)]42.3 (93.1) 25.7 (56.5)Restart toPropel lant.Depletion

injector temperature of engine No. IIIT increased to a maximum temperature of 406°k (270°F) during this period. Although the injector tempera-tures of engine Na. Ij] could not be obtained at this time because ofJoss of the measurements, it is thought that its temperature also exceeded382°k (280°F), Both engines had pulses with chamber pressure as low as38 N/on2 (55 psia). These were the initial pulses in a series. As thecocler propellants upstream of the injector reached the engine thechamber pressures increased to nominal values.

7-57

Soar REI cay

ao3 feronn e02 iso :: nog5 wre.Saw ae éZim weSn vee §

we 3 6Tayo Pesca

30:99 $500 49 6.3800 6 O-00ofoo Habae Va0-90 fons Gad wo acuae. Fe

Tigure 7-43. S-IVB kelium Bottle Pressure - Module No. 1

6,SAR gues come, 67 eehe Besta aia, 1604 99

a) vm 2z brs. 3

RnPETAR

a10

00 nett. woast

HD aoa Geb STOTT TOT TT A aE Fay NWS TOT TINHANMGE “TRE, moves

Figure 7-44, S-IVB Heliur Bottle Pressure - Modute No. 2

secewws

The APS ullage engine operation was normal with exception of the slightlyextended clamber pressure decrease time of the Module 2 engine (followingthe second ullage burn cutoff, it required approximately 1 second longerthan the Module 1 engine }. This extended pressure decrease time did notoccur at first ullage burn cutoff.

-| fre MS

37.9 N/em2 (55 psia) 63.43 N/om? (92 psia)—

0 1 2 3 4TYPICAL HEAVY APS DUTY BURN TIME, SECONOS

4:39:13 4:39:14 4339:15 4:39:16 4:30:17RANGE TIME, HOJRS:MINUTES : SECONDS

TYPICAL EXAMPLE OF INITIALLYLOW CHAMBER PRESSURE{NO TEMPERATURE DAIA ON INJECTORWALL OR PROPELLANT MODULES)

Tag tien? (65 psia) 63 Nene (92 psia}—

0 1 2 3TYPICAL HEAWY APS DUTY BURN TIME, SECONDS

et6:16:10 @:to:11 6:15:12 6:15:13

RANGE TIME, HOURS :MINUTES :SECONOS

INJECTOR WALL TEMP = 357°K (182°F) 22,500 TO 22,550 SECOXIDIZER MODULE TEMP = 204°K: (87°F) 22,500 TO 22,550 STCFUEL NODULE TEMP = 304 °K (87°F) 22,500 TO 22,550 SEC

Figure 7-45, S-I¥B Chamber Pressure, APS Engine No. 2 - Modute No. 2

7-89/7-60

SECTION 8

HYDRAULIC SYSTEMS

8.1 SUMMARY

The primary purpose of the hycraulic systems on each of the three boosterstages is to provide vehicle thrust vector contral during powered flightby controlling the thrust vector angle and direction af each of the move~able engines on command from the guidance and control syster. In additionthe S-IC stage hydravlic system atsc operates the engine control valvesduring engire start and shutdown operations. Since the S-IVB stage hasonly one encine, the Auxiliary Propulsion System (APS) modules provideroll corrections during $-IVB stage flight

In generat, the hydraulic systems performed satisfactorily in that thevehicle remained stable during al] portions of quidance-controlled poweredflight. flo hydraulic system problems occurred during $-IC powered flightS-II hydraulic systems performed within predicted limits, and operatedsatisfactorily until 280 seconds. At this time, the S-If engine No. 2 yawactuator delta pressure transducer began to deviate significantly fromexpected values. The engine No. 2 yaw actuator showed a cryogenic effectbetween 280 soconds and 419 seconds, From 319 seconds until engine No. 2cutoff, both the pitch and yaw actuators showed apparent side loads fromthe engine. After engine No. 2 cutoff, the yaw actuator performance in-dicates that it locked up. The engine No. 3 hydraulic system performed nor-mally until engine shutdown when the system pump stopped operation and thepressures decayed. The engine No. 1 and engine No. 4 hydraulic systens per-formed normally throughout 5-11 powered flight. 5-IVB hydraulic system per-formed within predicted limits during liftoff, boost, and first burn.During engine restart preparation and restart attempt, the system failedto produce hydraulic pressure. System temperatures observed during S-IVBfirst burn indicated the existence of a cryogenic fuel Teak which Ted tothe freezing of the hydraulic fluid and system blockage. During therestart attempt, measurements indicated that both the nain and the auxil-jary hydraulic pumps cavitated during operation and virtually no systempressure was produced.

At the present time there are no planned modifications to any of the stagehydraulic systems. The hydraulic system problems that occurred duringthis flight were related to the engine failures.

8-1

8.2 S-IC HYDRAULIC SYSTEM

The hydraulic system supplies high-pressure fluid, RJ-1 (kerosene}, froma ground source to each of the five engines to control the engine startingsequence ana to the four outboard engines for ground checkout of theThrust Vector Control (TVC) system. During engine operation, high-pressurecontrol fluid (RP-1) is supplied from the No. 1 fuel discharge of theturbopump assembly through the filter manifold to the servo valve andactuators for TYC. The fluid returns througa the checkout valve to theNo. 2 fuel inlet of the turbopump assembly. Hydraulic power is also usedto close the engine control valves for engine shutdown.

The S-IC stage incorporated four gimbal actuators of the Moog model(60884500-1} "and four of tre Hydraulic Research model (50M04050-1). Analy-sis indicates that all actuators performed as commanded during the flight.The maximum actuator deflection was equivalent to 0.7 degree engine gimbatangle at the initiation of the vehicle roll program. The average hydraulicsupply pressure was 1296 N/cm? (1880 psia) and operated in a small bandwithin the operating limits as shown in Figure 8-1. The temperature asdepicted by the return actuator fluid was 205°K (90°F) and operated withina narrow band. The maximum hydraulic engine valve opening pressure was1392 Yen? (2020 psia) and the maximum supply pressure to the actuatorswas 1378 N/om@ (2000 psia).

8.3 S-IT HYDRAULIC SYSTEM

A complete, separate, and identical hydraulic system for each outboardengine provides power for gimbating. The major System components includean engine-driven main pump, an auxiliary electric motor-driven pump, twoelectrically controtled, hydraulica?ly powered servoactuators, and an ac-cumulator reservoir manifold assembly. During $-IC powered flight, $-I1hydraulic lockup yalves are closed, holding the engines in a nullposition. After S-IC/S-1I stage separation, @ signal unlocks theaccumulator Tockup valves! releasing high-pressure fluid to each of the twoservoactuators. This fluid provides gimbaling power prior to main hydrav-lic pump operation, The main hydraulic pump, driven directly from theaccessory drive pad of the engine LOX pump, provides actuator powerduring S-11 powered flight.

S-If hydraulic system performance was essentially normal throughout theFlight on engines No. 1 and 4, and on engine No. 3 until premature shut-down. System supply and return pressures and reservoir volumes were with-in predicted ranges. Reservoir fluid temperatures were close to the maxi-mum predicted. Launch pad redlines were met with ample margins at liftofffor at7 four systems.

Several anomalies were apparent on engine No. 2 system throughout theflight. Actuator and reservoir temperatures of engine No. 2 leveled off

8-2

ea 1 1 I ene

170 4 - a/ | | 28ta - b4 ree oF

2 seg t 2200E can | |! |go 2000E oxo C: L 1003 ze

* 00 i Uf f alajeves opeversora use {rea f=

I! aceon

te eso saves acttvarsStan Stistusrs20

ines weracn RTD- TI ‘27 ~ WHE rmg LG ; 8

= - Lt 20 Tio GREG 5A 2g T «2

- {| z] poeay{ [une oncearconae La}—4

2a) : idaee

ANGE TIME, $1¢0908Figure 8-1, S-IC Hydraulic System Performance

18-3

early during S-11 boost and then decayed while the otrer three systemsrose normally. The differential pressure measurement (indicative of force)on the yaw actuator of engine No. 2 exhibited unusual ramp excursionsBoth the pitch and yaw differential pressures ircreased suddenly at319 seconds. At engine No. 2 cutoff, the reservoir velume dropped tozero, and the yaw actuator held its position

Uydraulic system performance for engines Ko. 1, 3, and 4 is shown inFigure 8-2. Reservoir volumes were within the predicted range prior tocutoff of the respective engines. At engine No. 3 cutoff, the engineNo. 3 reservoir yolume increased in a normal manner as a result of accumu-lator depletion followirg spindown cf the main hydrauTic pump. Reservoirtemperature increase was close tc the predicted rate of increase. Tempera-ture data for all four reservoirs and actuators were Tost at 415 secondsas discussed in paragraphs 13.3 and 19.2.2. Accumvlator pressures (indica~tive of system supply pressure) were within the predicted range of 2280to 2620 N/ené (3300 to 2860 psia) prior to Engine Cutoff (ECO), The de-cay in engine ho. 3 accumulator pressure was normal and resulted from theaccumulator depletion. The minimum reservoir volume was 13.0 percent offull versus the redline of 3.0 percent and was within the noninal pre-dicted bands. The hydraulic fluid minimum pressure was 2410 N/cm?(3500 psia). The highest of the three systems fluid temperatures was ap-proximately 306°K (90°F) at 415 seconds and all three were showing anormal upward trend at that time. These were well within the predictedlimits as shown in Figure 8-2. Engines No. 1, 3, and 4 actuator force:were Well below the predicted maximum of 84,509 Newtons (19,000 Ibf}. Theraximun tensile force was 57,800 Newtons (13,900 IbF), which was exertedby the pitch actuator of engine No. 1. The maximum force in compressionwas 36,000 Newtons (8000 Ibf) which was exerted oy the yaw actuator ofengine No. 4,

Prior to engine No. 2 cutoff, hydraulic parameters from all four enginesfollowed predicted values except engine No. 2 reservoir and actuator fluidtemperatures. Figure 8-3 shows that these temperatures leveled off between270 and 290 seconds and then decayed. This fs unlike the normal, con-tinuous rise characteristic shown by the engine No. 3 reservoir tempera-ture and is attributed to an abnormally low temperature environment inthe vicinity of the engine No. 2 hydraulic system.

At 280 seconds, the engine No. 2 yaw actuator differential pressuremeasurement started a positive ramp increase as shown in Figure 8-4. Thiindication was apparently not a measure of increasing pressure, but theresult of cryogenic fluid coming in contact with the transducer. Thieffect has been reproduced in tests performed at MSFC and at the S-I1stage contractors’ test facility. The transducer utilizes two Bourdortubes to sense pressure difference. In the tests. LN2 was sprayed on theactuator and caused the instrument to show a pressure difference asshown in Figure 8-4 although no pressure was applied to the actuatorThe indication rose, peaked, and then decayed apparently beceuse one of

8-4

S-II AYDRAULIC ACCUMULATORS, UNLOCK, 151.38WY SHIT STAGE ENG ND. 3 OUT, 414,78

® T Tene}4 Fj ENG ND. 3B so - t oh. +

ce * /meorere MAXIMUM

a

5 é0 IW

~ 3000 75. |ZTaenueTe nat | s000 =

3 2500 - 350 32

= # 2000 K [ 3000 Bs52 PREDICTED NININUM 332 1 > 2800 33ES 1530 38£ | L woo 8

yo00 i ~

350 T=F | 160

22 x5 3Be rapDieey TEMP. RISE DIPS ar ple 2g

ve 300 z : eae so BS

Bars aes r® 35250 W YZ fo 7a

oo 150 200 250 300350 400 50, 500-550 tnRANGE TIME, SECONDS

Figure 8-2, $-I1 Hydraulic System Performance

as

310 T T T

yt NO. 3 RESERVOTR Lag

TBae.yo LENS_NO. 2 yaw ACTUATOR 4

24 3°Se Be2s gsoe 290 ' 28oh L-7 Zs so. '2 FITCH ASTUATOR ogEa bso aitwe 280 Z pot a=ok ENG NO. 2 RESERVOIR n=25 li czgee 270 80g #22° MEASUREMENTS ee

LOST ,415 2260 L i Ww ~10 200 230 260 230320350 380410 aag) 70500

RANGE TIKE, SECONDS

Figure 8-3. S-II Engines No, 2 and 3 Hydraulic System Temperature

the Bourdon tubes was cooled first and then tke cther. when both tubeshad been cooled to the same temperature, indications are that transducerwas again able to measure differential pressure.

The ramp increase continued until 319 seconds when the apparent force on theyaw actuator had reached 80,000 Newtons (18,000 Ibf). At 319 seconds, theengine No. 2 pitch and yaw actuators showed a step increase in pressure whichwas apparently caused by a side load an the engine. The apparent forceon the yaw actuator rose rapidly to 10,000 Newtons (22,700 Ibf) as aconstant force of approximately 21,0C0 Newtons (4700 Tbf) was added to thetemperature-induced indication. At the save tine, the force on thepitch ectuator increased to 32,000 Newtons (7000 Tbf), These constantforces remained until engine No. 2 cutoff. The total yaw actuator forceindication continued to increase fror 101,000 Newtons (22,700 Tbf) andreached a maximum cf 125,06C Newtons (28,000 Ibf} at 336 seconds, whenit startec a decay which reached 13,000 Newtons (3000 Ib) at 416 secondsIn engine static firing tests conducted at MSFC, these actuator hydraulicpressure changes and engine novements have been reproduced. In the tests,a LOX-rich Augmented Spark Igniter (ASI) mixture vatio was used. Thicaused the main injector to be eroded. This, in turn, made a hole in theaft portion of the thrust chamber. Expanding gases from this hole createda side thrust which caused effects very similar to those observed iflight as shown in Figure 8-4. The engine failure analysis is describedin detail in paragraph 6.3,

140 TSENG NO. 2 PERFORMANCE SIFT, 319 L- 30WS-II ENG NO. 2 OUT, 412.92

420 ~L- 25

EXPECTED CRYOGENIC EFFECT100 - + 4

\ ENG NO. 2 + 20YAW ACTUATOR]

20' \ his

= 50 { . 22 N I 2

‘ ‘ I Liog* Ti && £= . =B 20 ~ if J rs 6= ENG NO. 20 NY 2

PITGH ACTUa

-20 -

-A0

“60 % Yi 280 300° 320 349 360 380 400 225 440 460 480

RANGE TIME, SECONDS

Figure &-4. S-11 Engine No. 2 Actuator Forces

At 423 secends, the yaw actuator force increased to 98,000 Newtons(22,000 Ibf), it then started to decay and finally leveled off at-35,000 Newtons (-8C00 Ibf) (tensile force) at 478 seconds, where it re-mained for the duration of S-II flight. The pitch actuator force droppedto -22,000 Newtons {-5600 1bf) at 415 seconds and then decayed to-31,000 Newtons (-7000 Ibf) where it remained to the end of S-IT flightAfter engine No. 2 cutoff, both the pitch and yaw actuator pressuremeasurements indicated a force which tended to force the engine inboard.Tris incicated force is stil] under investigation

Tre pitch actuator continued to respond to quigance commanas os shown inFigure 8-5 until 480 seconds when the accumulator was empty of fluid.Shartly after engine No. 2 cutoff, the yaw actuator failed to respond toguidance commands. The performance of the yaw actuator incicates that itjocked up at this time. The probable causes for this were closure of the

8-7

. I. PNA

NeistoxPOSITION 1

2PI

TCH

ACTUATOR

POSITION,

deg

ENG

NO.

VSI ENG NO. 2 QUT, 412.92

STOROSITIONL-

2 COMMAND

POSITION,

deg

ae2

-4 Vva0 412 «4144641 420422 ak

RANGE TIME, SECONDSFigure 8-5. S-II Engine No. 2 Actuator Conmands and Positions

ENG

NO.

2YAW

ACTUATOR

hydrautic lock valve within the actuator or seizure of the piston. Thelock normally closes when the difference between the actuator supply andreturn pressures decays below 900 to 1200 N/cm2 (1300 to 1700 psid). Alow differential pressure could have been caused by highly viscous fluédwithin the actuator or the supply and/or return lines to the actuator,resulting from a low temperature condition. Piston seizure could haveresulted from cryogenic fluid spraying on the actuater cylinder. TheNo. 2 hydraulic reservoir level also appeared to have dropped sharply from9 percent to zero at ECO. Hydraulic system No. 3’shows the expectedreservoir level at this time. The zero volume indication was probablydue to a complete loss of reservoir fluid caused by a rupture somewherein the low pressure side of the system. This is based upon a corresponding

8-8

rapid decay in reservoir pressure to zero, accompanied by a slow decay inaccumulator pressure which is on the high pressure side of the systemNormality, the reservoir pressure decays to approximately 35 N/cmé (50 psia)after ECO. The cause of the rupture has not been defined, however, anabnormally ow temperature environment could have subjected a portion ofthe system to extreme strains and eventual rupture. The gradual decayin engine No, 2 reservoir volume before engine No. 2 cutoff (Figure 8-6)was probably the result of accumulator gas cooling which resulted in aincrease in accurulator oi] volume and produced a decrease in reservoirvolume

8.4 $-1VB HYDRAULIC SYSTEM (FIRST BURN)

Engine gimbating 1s accomlished by an independent, closed-loop, hydrauliccontrol system consisting af an enging-driven main pump, an auxitiaryelectric motor-driven aums, two electrically controlled, hydraulicallypowered servoactuators, and an accumulator reservoir. Buring S-IC andS-H1 powered flight and coast, the auxiliary pump is operating to positionthe J-2 engine in the null position and to pressurize the system. Themain hydraulic pump, driven directly from the accessory drfve pad of theengine LOX pump, pravides the orimary actuator power during S-I¥B poweredFlight.

The S-IVB hydraulic system performed witain the predicted limits afterLiftoff with no overboard venting of system fluid as a result of reservoirfluid expansion. Just prior to start of propellant loading, the accumula-tor was precharged to 1570 N/cm? (2200 psia) at 277° (40°F). Reservoir

100

we YS-11 ENG NO. 2 OUT, 412.92= ENG NO. 3 OUT, 414.18S Boa ENG NO. NY

BE 69]

Sy 40; /

— 26

6 ENG NO. 2—7 Re 200° 230 «260 299 «320380389410 440 a7 S00

RANGE TIME, SECONDS

Figure 8-6. S-IT Engines No. 2 and 3 Hydraulic Reservoir Levels

8-9

oi] level (auxiliary pump off} was 78.5 percent at 270°K (26°F). Therewere no thermal cycles prior to launch. The low Gaseous Nitrogen (GN2)temperature of 278°k (40°F) resulted in a low accumulater precharge pres-Sure of 1517 N/ené (2200 psia). This caused appreximately 66 percent ofthe reservoir volume to enter the accumulator after the auxiliary pumpwas turned on. This effect, coupled with the cil volume shrinkage in thereservoir, resulted in a comparatively low reservoir oil level of 17 percent. Table 8-1 shows minor pressure level variations and compares theliftoff, first turn, parking ortit, and engine restart attempt systempressures.

During S-IC/S-II boost a11 systew fluid temperatures rose steadily (Figure8-7) when the auxiliary pump was operating and convection cooling wasdecreasing. AccumuTator gas and actuator cylinder temperatures remainedlow since they are located on the extreme ends of the systen. Thesupply pressure during the first burn was nearly constant at 2482 N/cm2(3600 _psia) as compared to the allowable of 2344 to 2517 N/cm® (3400to 3650 psia}. The maximum actuator torque resulting from vehicleatet tude comand during first burn was in pitch at 12,639 Nem (111,853Ibf-in

System temperatures did not rise normally during the latter portion offirst burn as shown in Figure 8-7, The fallowing characteristics werenot experienced on previous flights:

At 688 seconds the yaw actuator suddenly started to lose temperatureat the rate of approximately 0.28 °K/s (0.5 °F/s).

b. At 707 seconds the pump inlet oi? temserature suddenly jumped 16.6°K(30°F) in 14 seconds and then decreased 8.3°X (15°F) at cutoff.

¢. The wain pump discharge line temperature began to rise normally atthe start of engine burn. At 670 seconds it suddenly started todrop and then laveled off at ECO.

Table 8-1. S-I¥B dydraulic System PressuresT wie6 ono|

Ly SLLRENBLE GURTNGeecssumes Wer’ Were fish wes Reee Te hit eg MeNeysten x1 zue 34g 2344 ve 2517

(3506; | f36c0) i 13500 te 26801wu tOF Sy zie

|

ze00 se 489 238e s6 2507aainy

|

Getuy 2109) (zis (28eu e 3650Reservoir C41 WW via 4% | os 116 e9 126amen Tames jen |e) ses ee 285)tux Purp Air Tank |283! 296 258 Poa

aes) dar | (430; tayfox So Motse ie

|

1g 34,0 1 20fey ioe ae

The values Faveeen corrected to che 293 °< (67 “Th

8-10

E15 ENG STARL SHOUFHCE cOMPERG, 577.28SON LOECTY cuTRF ree, 3E sot

weec

10

220 ~ 1 =: _ . TP dy+ we - ~ bes é -# io ure ter xPee popes ia “aneBOTS F maT eg2oe => be :5 ; 0?un 2 ivSone ee za koSonatense TSE, SCah

aaURCUELATOR By 1a:Ex on | pect pctuvun ot

oe xa

25iso

2anp 1

a + = - 4Se +<8 ¢

Bow +0 | |v= ec aan 0

RaiGE TAME, sconsFigure 8-7. S-IVB Hydraulic System Performance - First Burn

8-11

8.5 S-IVB HYDRAULIC SYSTEM (COAST PHASE)

During orbital coast there were no thermal cycles of the auxiliary fiydraulicpump.” After ECO, the pump inlet of] temperature increased from 349 to358°K (168 to 185°F) due to continued heat transfer from the LOX turbinedome to the pump manifold as shown in Figure 8-8. Ouring the remainder ofthe coast period this temperature decreased gradually alang withother system temseratures. System bloeddown required 57 seconds and ac-Cumulator srecharge pressure stabilized at 1062 N/cm (2150 psia).8.6 S-IV3 HYDRAULIC SYSTEM (SECOND BURN)

The auxiliary pump was activated to the flight mode at 10.822 seconds(793 seconds prior to second burn). The pump failed to produce any dis-cernible hydraulic pressure. There was current draw to the pump motorof approximately 12 amperes which is an indication that the pump wascavitating. Normal motor current is 45 amperes. A 6.9 N/cm2 (10 psi)reservoir pressure increase occurred approximately 250 seconds later ashown in Figure 8-9, but this was of short duration

A

aed Zp4¥ us a __|=

: “Yan ACTLATOR may

“ TCH ACTUATE % — +. 4 “: 7wl GA Yh :

We E10 START SEaUENCE cOWMRD, 377.78ME RESTANT Coma11pVes

20.860 Tha OT wane ONE “INE, SOR EES: SECM

Figure 8-8. S-1¥B Hydraulic System Performance - Coast Phase

8-12

vac

a

oo 8

40

RESSURE

pressu

re,wr

en?

HTC PUMP FLIGAT NUGE OK, “0,822.253 START GH, 11,614.49STURT GE T7 {£00}, 11,630.337A IMDRAULCe, UH FLEGAT MODE CFF, 11,634.18

70

219,200 17,000 1,209 11,800 11,800. 2 ara

ange TORE, SEDUIYS

gw 1 1 L iw 13:06:03 Caos 3306529 TID TE TasT6:40

FANE TIME, HOY

RITES S7CONESFigure 8-9, S-IYB Hydraulic System Pressure Ouring Attempted Restart

After engine restart conmand the main engine hydraulic pump failed toproduce any measurable hydraulic system pressure. At the start of LOX tur-bine spin (at approximately 11,622 seconds), after restart engine command,there was a small fluctuation in reservoir oi] pressure, the actuatormoved slightly and there was a 15°K (27°F) momentary drop in pump inletoil temperature as shown in Figure 8-10. This ind’cated that the mainpunp was turning and moving fluid but was unable to develop system pressure

It has been concluded that both the auxiliary and main hydraulic pumpfeilures in the period preceding and during Second burn attempt were dueto cavitation. S‘nce the reservoir of1 level and pressure were normal,it is believed that the condition was caused by localized freezing ofthe pump suction line hydraul‘c flud by an impingement from a leakingcryogenic line. “he pump suction Tine runs across the gimbal plane onposition IIT between the accumulator reservoir on the thrust structureand the main hydraulic pump on the LOX turbine dome as shown in Figure 3-11.If this Tine is subjected to cryogenic freezing, a blockage of of] wouldresult in the line {pour point of MIL-H-5606 of1 is 205°K [-90°F]’ whichwould starve the inlets of both pumps, lose inlet pressure head and pre-vent reservoir fluid from reaching the pumps.

Two attempts were made to start the auxiliary hydraulic pump by groundconmand with no success. The pump inlet and reservar of] temperaturescontinued to sink at approximately the same rates. The pump dischargeline temperature measurement which was located near tre system thermaswitch apparently varied with exposure to solar raciatior or earth'sshadow. This shows that abnormal system temperature deviations were nolonger present after the engine restart attempt.

YY Fux svtnaucse PIMP FLIGHT Moce O04, “0,872.2WD SLiewG FART Oh, 11,614.69Y stacr or ty (Fco},"11,630.33WY Auk VvbaRut <C EUPP FLIGET MODE OFF, 12,038.18aay

cr - - - x0

gp 0 ~ I to mm= sin Pep eater -5 wlaa = ~FSSa Toa im 53 “|pSRS = wo #Aap Lf a 2Q t 7 42 ; " ia= 200[--— Jos5" Trescmvote, z5 LOMITS 820 =

sesb Fume +"Lents Fraa | “2

2 Toomade on S& 260 oemArke as£ Vis aciuatan CHLIN gz ecrea stag? ATER ES #Eom t

gy? T T TSL . actuaroe910m 697 bs-c _

A iuataR PUSTOR p3y Yaw 1i f |

“610,300 ne 1170 Tao 1 6 7690vance “ie, scons

eeWZ 4 4 4 103:00:60 Daa Tea aN thes THT6 00

RANGE TENE, JouRs etre 5 sEcoans Figure 8-10. S-IVB hydraulic System Performance

Curing Attempted Restart

Figure 8-11, d-2 Engine Hydraulic Component Locations

8-15/8-16

SECTION 9

STRUCTURES,

9.1 SUMMARY

The AS-502 flight vehicle experienced considerably more structural activitythan the AS-501; however, this activity did not reach suffictent magnitudeto pose a threat to the launch vehicle structural integrity. Areas ofstructural activity included:

a. A-slightly more severe vehicle release transient than aAS-501 (paragraph 9.2.1).

b, Coupling of longitudinal structural dynamics with thrustoscillations (POO) (paragraph 9.2.3.1).

c. Limited amplitude S-IVB panel flutter (paragraph 9.3.4)

d. A shock response type transient which occurred at 133 seconds{Section aN

e. Inboard Engine Cutoff (IECO) longitudinal acceleration designvalue exceedance (paragraph 9.2.1)

f, Premature cutoff of two of the five S-IT stage engines(paragraph 9.2.1 and 9.2.2).

The transients, due to thrust buildup and vehicle release, resulted inmaximum Tongitudinal and lateral (pitch plane) dynamic load factors of40.4 and £0.08 g, respectively, at the command module. The maximumsteady-state bending moment condition, 9.89 x 106 H-m (7.29 x 106 Ibf-ft)was experienced at 66.6 seconds. The maximum longitudinal loads wereexperienced at 144.72 seconds (IECQ) at a rigid body acceleration of 4,8 g,Although the 4.8 g [ECO condition was greater than the 4.68 g designvalue, no mainline structural problems were encountered during thisphase of flight.

Theust oscillation-structural dynamic response coupling (POGO) wasevident during the 110 to 140 second region of S-IC range time. Thelongitudinal dynamics of the launch vehicle induced lateral acceleration:of 6.65 Gpeak at the Lunar Module Test Article (LTA). Oscillations inthe First longitudinal mode during the 110 to 140 second time period

9-1

exceeded that exverienced during AS-501 flight by approximately a factorof three. Maximum response occurred in the 5.2 to 5,8 hertz bandwidth,

Fin bending and torsional modes compared well with analytical predictions.Fin vibrations exceeded the range of the accelerometers but the modalfrequencias did not coalesce and flutter did not occur, $-IC, S-I¥B,and Instrument Unit (IU) vibrations were as expected. S-I1 stage vibra-tions were as expected, except that forward skirt vibrations exceededthe sine and random criteria at liftoff, No adverse effects were nated.S-IVG forward skirt experienced limited amplitude pane) flutter. Thestress amplitudes encountered, due to flutter, were about three times higherthan those of AS-204 but were still within a tolerable level.

A pronounced transient was evident at approximately 133 seconcs in manyMeasurements on the vehicle. The transient was also reported by MarnedSpacecraft Center (MSC) to be present on many of the spacecraft measure-ments. Details of this transient are discussed in Section 9A.

9,2 TOTAL VEHICLE STRUCTURES EVALUATION

9.2.1 Longitudinal Loads

The vehicle longitudinal dynamic response due to thrust buildup andrelease ts shown in Figure 9-1, The axial cynamic loads derived fromstrain gage data are shown at the S-IC intertenk and forward skirt. Twoupper stations, where astrongut cowfort is cf prime concern, are givenin terms of acceleration. The simulated response at the command moduleis based on measured forcing functions ard is presented in lieu of themeasured accelerations which were not available.The measured accelerations shown at the IU were filtered to eliminatelocalized high frequencies (higher than 6 hertz) so that overall vehicledynamics could be nore accurately established. A frequency analysis ofthe filtered data indicated a predominance of 3.8 and 4.4 hertz oscilla—tions. Oscillations observed ir the axial load plots are not as pro-founced as in the acceleration data because of the low frequencylimitation (2.4 hertz) of the telemetry system from which the straindata was obtained, The 3.8 and 4,4 hertz oscillations are preciselyas predicted for the first two longitudinal modes at the time of vehicleTaunch.

In general, the AS-502 vehicle longitudinal dynamic response amplitudesat launch were greater than those experienced on the AS-50). “aximumresponse at the command module was approximately 0.4 g (simulated) onAS~502 and 40.2 g (simulated) on AS-501.

This increased response was due primarily to changes in the controlledrelease device characteristics. The time required to clear the controlledrelease device was 1.17 seconds on AS-501 and 0.54 second on AS-502 witha steady-state acceleration of 1,24. The most significant change wasthe reduction of the number of release rods from 16 for the AS-S01 flight

9-2

VEHICLE STATION{INCHES} (METERS)

4208.8 107.8

98.836.0

91.3

84.882.831.378.875.2ng68.6rt

— 3519, 64.0I 2337 60.6

46.9RE

39.135.6

5.35.7

Figure 9-1, Longitudinal Structural Dynamic Response Due to ThrustBuildup and Release

to 12 (€ instrumented) for AS-502. Revised rod lubricating techniquesadopted on AS-502 gave rise to the lower release rod naximum force shownir Figure 11-3 (Section 11) which also served to increase the releasetransient, Still another minor change in the release parameters was theslight decrease in release rod preload shown in Figure 9-2.

In spite of the changes in the controlled release device which increasedthe dynamic transient due to launch, the response accelerations and cor-responding loads throughout the launch vehicle primary structure werewell within predicted values. This phase of flight posed no tareat tomaintine structural integrity,

The vehicle longitudinal acceleration measured during flight is shown inFigure 9-3. The rigid body component of this acceleration was essentiallynominal throughout S-IC burn. This acceleration information, along withpreflight predicted weight distribution and aerodynamic drag, was usedas a basis for longitudinal Toads computations. The maximum Yongi tudinaldynamic response subsequent to the release transient was experiencedbetween 110 to 140 seconds and is illustrated by snadina in Figure 9-3.This aspect of the flight will be discussed in naragraph 9.2.3.1 of thisreport. The longitudinal loads which cxisted at the time of maximumaerodynamic loading (maximum bending moments) and at maximum compression(TECOS are shin fn Figure 9-8. These Figures illustrate the exceltentagreement between measured loads and Toads computed from measured flightparareters. The rigid body longitudinal acceleration at time of maximumbending moments (66.6 seconds} was 1.93 g. The maximum longitudinalloads experienced during flight occurred immediately prior to TECO

5 T t

oa - 4

Lx10-6,

Ibe

SS

owroe

RELEASE

POD

FORCE

x10-6

A ‘

sues

ot 7 LI

OLI00HH ARM RETRECT:

“3.0 “20 shat ac 1.6

BANG: TIME, SECON

Figure 9-2. Slow Release Rod Loads During Release

9-4

Teco, 144.72 T 7 T T TT T

eco, 148.41 yi +FO -— recake so} STON 92.58 eres| Li.p g(3250 TH} Toner Tu TaRL

OSCILLATIONSfeck permen oon (BY ~< 4-1

a ne farenal COMET To ae a

| we ef1 ot ve pet1 rotpo ef fee af penob oeLfpoa

RANGE TIME, SECCHDSFigure 9-3, Longitudinal Acceleration Time History

(144,72 seconds) at a rigid body longitudinal acceleration of 4.8 g.Although the 4.8 g JECO condition was, as predicted, greater than the4,68 g design value, no mainline structural problems were expected orencountered during this phase of flight. The higher g level was theresult of a longer inboard engine burn to attain LOX Tevel cutoff

Figure 9-§ shows longitudinal acceleration time histories at the S-ICcenter engine gimbal block, S-IC intertank region, and the command moduleduring S-IC thrust cutoff (OECO}, Due to absence of accelerometer datafrom the command module, only the dynamic simulation results are presented.

At approximately 413 seconds range time, two of the five S-TT enginesprematurely cutoff. The longitudinal loads corresponding to this two-encines-out condition are shown in Figure 9-6 and are well below design.

9.2.2 Bending Moments

The lateral transient due to thrust buildup and release is typified tythe S-IG intertank and forward skirt load time histories shown inFigure 9-7. The 1 to 2 hertz oscillations apparent in this data aredue to the lateral dynamic response, or “twang”, at vehicle release.The low frequency load buildup and decay subsequent to 1.8 seconds is a

9-5

mertcLe San

1G WENT (Le6G.E1, LOAD FACTOR = 5,996

L-GiLcuL Teo Fane ve csaneParasJy resco stato bata

ANTAL

Lav

a10-4

,IBF

Figure 9-4, Longitudinal Loads at Maximum Bending Moment andInboard Engine Cutoff

result of the programmed yaw maneuver. The pitch plane dynamic simulationresponse at the comand module during launch is also presented in Figure9-7. The maximum response acceleration was found to be +.08 9. Neesureddata at the command module was not available due to questionable flightdata.

The conditions which existed during the high aerodynamic Yoading phase offlight were such as to cause near minimum lateral loads. This isTllustrated by a comparison of the maximum AS-502 flight bending momentswith design values shown in Figure 9-8. The lateral Toad factor is alsoshown, The 9.89 x 106 N-n (7.29 x 108 Ibf-ft)} maximum bercing mement inthe S-IC LOX tank at 66.6 seconds was only one third of the design value.The $-11 engines which prematurely cut off (Engines No. 2 and 2) were inthe upper outboard location and could be expected to preduce relativelylarge bending moments. A conservatively high calculaticn cf thesebending moments, shown in Figure 9-8, indicates values well within designenvelopes, A comparison with measured strain cage data at one station1s shown in the figure.

Ta VEHICLE STATION

(INCHES)4265.4

3890.33781.0

3594.6

—— 3340.33258.63222.63100.62962.9

e320T= 27020— 2648.4an 2579.0-—— 2387.0

Sea sah

Come +] | Eee i—si2.0ttt

418

Figure 9-5,

9-7

(METERS)107.8

46.943.6

39.135.6

29

Longitudinal Acceleration During S-I¢ Thrust Cutoff {OECO)

PAI F STATIK, i,

ea 3oo.6 3 ¥YEAUCLE STATION,

emt wt weg5

: 18aa | 4* ; i3 | aes. | 22? 2g | oea2 —- S

' cag

t a2||p

Ll i °

TTI | UE

Figure 9-6. Longitudinal Loads Subsequent to S-11 Engines Out

9.2.3 Vehicle Dynamic Characteristics

9.2.3.1 Longitudinal Dynanic Characteristics. Frequency versus rangeime for the first longitudinsl mode is compared with the analytica

prediction and data from the AS-807 flight in Figure 9-9, Modal ampli-tude versus range tine is also shown, Oscillation in the first longi-tudinal mode was observed on at] longitudinal low frequency instrumenta-tion in the maximum Gpeak regions shown in the lower part of Figure 9-9The frequency data agree well with both the analytical prediction andthe AS-501 data. The data shown in Figure 9-6 of NPR-SAT-FE-68-1 Saturn¥ Launch Vehicle Flight Evaluation Report AS-501 Apollo 4 Mission isincorrectly identified as being on the IU, when in reality it was obtainedfrom the command module: therefore, it cannot be compared to the AS-50data shown in Figure 9-9 of this report.

The AS-502 vehicle experienced closed loop structural/propulsion coupling{POGO; in the letter part of the S-IC flight. The POGO pheromenor is aclosed loop interaction of the three essential interactive vehicle systems:the vehicle structure, the vehicle suction propellant feed system, andthe engine system,

3-8

VEHICLE STATIONCINCHES} (METERS)

4285.4 107.8

3890.3 oR.a7al.a 96.€

3594.6 91.3

fe v4 oe —— 2340.3

:

39.135.6

23.119819.1155

yee

9.35.7

Figure 9-7. Lateral (Pitch) Structural Dynamic Response During

Thrust Bui tdup and Release

9-9

BENDIN

GMRT

410-

6,Ye

Tn10-

8,tw

DINE

a

PEHECLE Sealey

pmo ROS Se9 Bit Wes Ise Nel oma ama gTate wail @ope wom

* Tr OEE SECON25TINGwatesto Ras»®a |aim

ereto[REBs «}—-4+—--P-4_ |Z | Zp coe -4,sexswoios

|

oni vara. (Srna Bae

65 rayaaserEna

MET Ar IPE AT B

vewrae 1108, inDono MAD eG Me BA HES Ue TAAope7 imoe oe ee ee ee

a»

‘0wsrat tow

» 8

“‘0

8jueas

a .—[T= elas wp. et Senha ;

Figure 9-8. Maximum and S-II Engines Out Bending Moment

9-10

weet

PENDIN

GMOM

ENT3 1

6-6FADING

MOMENT

440-

6,tears

TaeASE ataLysis ———asesoz veasunco @: : ~ AoP

weQMEAY,

He

i

rn L

Beep —teasso=

as-f02——30

'}

28 \-i|

ge 7;* fINSTAyEnT weir A610

ais 7i

0 \ — _i \\. \yN.she

ii 2m omANGE Tne, skemuos

1

Figure 9-9. First Longitudinal Modal Frequencies and AccelerationsDuring S-IC Powered Flight

9-17

The buildup of longitudinal cyramics started about 110 seconds range time.The structural oscillations reached a maximum level of £0.18 g at theS-IC gimbal plane and 40.65 ¢ at the command iedule at 125 seconds, thendecayed to a negligible level ty 14G seconds. This buildup is a resultof the coalescence of the first longitudinal frequency of the vehiclewith the first LOX line frequency in the frequency range of 5.2 to 5.4hertz. The frequercy coalescence and the resulting structural resnonsebuildup is illustrated in Figure 9-10.

All five F-1 engines were essentially in phase on AS-502 and shoved achamber pressure increase from 110 to 125 seconds and then a decreaseback within noise level arplitudes at approximately 138 seconds. Figure9-11 shcws the charber pressures on the individual engines and also 4compesite chamber pressure for the five F-1 engines.

A major difference observed in the propulsion data of the two vehicles(AS-501 and AS-502) is that all engines exhibited oscillatory thrust atthe save frequency and essentially in phase on AS-502; whereas, theAS-501 engines did not show the same engine-to-engine thrust frequencyThe buildup in the structural acceleration response at the S-IC gimbanlane and the F-T engine chamber pressure is shown in Figure 3-12,Modal amplitudes during the 110 to 140 second region of S-IC rangetime exceeded those experjenced during AS-501 flight by apsroximately afactor of three as shown in Figure 9-9. Amplitudes at differant vehiclestations reached peak levels in this time period at various range timesas indicated in Table 9-1. Maximun response occurred in the 5.2 to 5.5hertz bandwidth, The longitudinal dynamics of the launch vehicle in-duced lateral accelerations of 0.65 Gpeak at tre TA. The maximum levelsmeasured in the spacecraft LTA during the 120 to 135 second region were:

a. Pitch 0.68 Gpeak

be Yaw 0.30 Gpeak

¢. Longitudinal 0:70 Gpeak

The combination of steady-state longitudinal acceleration of 3.7 g nlusthe dynamics at 125 seconds created a relatively critical compressionloading condition for the upver poction of the vehicle. This can bebest illustrated by a commarison of the combined Tvad (Nc) at 125 secondswith TECO loads.

. P APRNem ape * aie 7 Oe

FREQUENCY,

Hz

5.6 OBSERVED FREQUENCY OF 7CHAMBER PressuRE [J

[OSCILLATION ‘7sa va

5.2 i oa

5.0 PREDICTED FIRST MODE: LONGITUDINAL FREQUENCY

+ 0.4

48 |ENVELOPE OF DYNAMIC —J

[RESPONSE AT S-ICGIMBAL PLANE 0.3

/ 0.2

b0.1

096 100 10 720 130 740 750

RANGE TIME, SECONDS

Figure 9-10. Longitudinal Oscillation Trends, 110 to 140 Seconds

9-13

DOUBLE

AMPLITUDE

ACCE

LERA

TION

,g

so | 1

a : 5 Tt,eetBa) v.23) os

ef T wzaaa - se

eg T we

Epa o*

a 10Seoot f 5#4 i ta >ace a a we?= ava

c0ve,

t0003i SMe commas commen ouessine

1 | | ] i 16Lod tahal

a io 2

‘HAMHE

R BRES

SURE,

984

Chante

PRESSE, H

ien?

2 a atMOT VE Ve Teo TE Te TTTTET

nce Time, sEeanos

Figure 9-11. S-IC Naximum Individual Engine and CompositeChamber Pressure Oscillations

O14

ACCELERATION,

g

°

2

15 L 20

10b

% R10 &

d Y 2g° og

= “a, a Yer L-10~10f/-- Ys

Pr -2015

no 120 130 140

RANGE TIME, SECONDS

Figure 9-12, F-1 Engine Chamber Pressure and Structural Acceleration Responseat S-IC Gimbal Plane During Time of High Long{tudinal Oscillations

O18

Table 9-1, Maximum Modal Accelerations at 5.5 Hertz for 110to 140 Seconds Range Time

INSTRUMENT VEHICLE S-1C RANGE ACCELERATIONSTATIONS TIME {GPEAK}METERS (IN. ) (sec)

PITCH£61-118 16.76 (660, 130 9,03E59-118 21.67 (853 132 0.04A4-320 37.08 (1460 132 2.036-603 82.55 {gee 134 3.19GA7015 88.60 (3488, 134 2.65cAOOOTA 97.41 (3835 134 2.35LAGOT2A 107.21 (4221) 130 0.49

LONGITUDINAL

£90-115 2.92 (118) 120 0.40E83-115, 3.67 (121) 123 0.16E82-115 3.15 {iza} 124, 0.18£57-115, 5.31 {209} 120 0.12£92-117 5.72 (225) vig 0.43£93-119 19:6) (772) 120 0.18E58-115, 21.62 (851) 128 0.146R7011 88.60 (3488) 125 0.70

YAW

47-603 82.55 (3250) 133 0.04GA7013 88.60 (3488) 133 0.10LAOOTIA 107.21 (4221) 138 0.12

Where:‘ircumferential compression shell load (1bf/in.)xial Toad (1bf)

8M = Bending monent (in.-Ibf!ocal vehicle radius (in.)age ullage or venting pressure (Tbf/in.2

The combined conpression loads attained the largest relative maximum,compared to IECO loads, in the S-IVB forward skirt and IU. At the 1U,Spacecraft interface, the limit compression load at 125 seconds wa1103.30 fom (630 1bf/in.) compared with 1015.74 N/m (580 Tbf/in.) atIECO.” the mininum predicted ultimate capability at this time i1523.60 H/em (870 Ibf/in,) based on a maximum heating trajectory tempera-ture prediction. AS-502 flight was such that maximum heating was notexperienced, and the minimum capability was only degraded to 2084.01 N/cm(1190 Ibf/in.}, The minimum factor of safety maintained through thicondit‘on for AS-502 was well in excess of 1,5, The axial load diagram,steady state plus dynamics, for this condition is presented in Figure9-13, Lateral loading at this time was sTight.

Figure 9-14 shows a comparison of normalized flight data with analyticallypredicted first longitudinal mode shapes. Mode shape data from the LManalysis have been included on one of the shapes for ccmparison purpases.

9.2.3.2 Lateral Dynamic Characteristics, Oscillations in pitch and yawware detectable throughout S-IC pawered flight. The frequencies of theseoscillations agreed with the aralytical predictions as shown in Figure9-15. The first three vehicle pitch modes were detectable throughoutfirst stage boost. The fourth mode came in very strangly near the end offirst stage boost. Spectral analyses were performed ta determine modaFrequencies using five-secord time slices. Te obtain raximum accelerationlevels, magnetic tane data were filtered using digital filters set at themodal frequency range. The yaw direction acceleration levels were Tessin the 110 to 140 second region than the pitch or longitudinal levels:however, the first moce response in the yaw direction was generally higherthan in the pitch direction throughout first stage flight. Modal datafor the second yaw mode were generally hidden by a two cycle noise whichexisted in the yaw instrumerts throughout first stage flight. The noiseappeared in the S-IC stage and [U measurements. A comparison of normalizedflight data with analytically predicted pitch and yaw mode shapes ispresented in Figure 9-16.

9.2.4 S-IC Fin Dynamics

Vibration levels on the fins were high at liftoff and in the high dynamicPressure portion of AS-5C2 flight, but the modal frequencies did notcoalesce and there was no evidence of flutter. Figure 9-17 shows thatvibration levels were simflar to those of AS-501, Levels observedexceeded the range af the instrumentation; however, the range of the‘instrumentation was below design levels.

9-17

N ‘ayo Wixy

9-18

VEHICLE

STATION,

in.

3600

3000

2500

2600

1500

1000

500

3VE

HICL

ESTATION,

m30

8070,

6050

40.20

==

125

SECONDS

STEADY

STAT

ELO

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DYNAMIC

LOAD

AMIE

ican

FACTOR

=0.

86

@STATION

106,

68METERS

401

50

1c

Figure

9-13

.Longitudinal

Load

sat

lime

ofMaximum

Longitudinal

Oscillation

49L gO favOT Twiry

THEW IW ARYSIS“X TERAAsAVSIS — WEASURED “e Bi ¥51BRALYSTS FRERRENEY = EI {u2)MEASURED FREQUENCY

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. [ASSESS |¢ ° °3 3 1 ay a Ve{RMALL2ED HPLITUDE NORMALIZED asPLeTUcEa7 cH seeon0s ar 138 Sec0s03Figure 9-14. First Longitudinal Mode Shapes During S-1C Powered

Flight

Measurement range ‘increases were requested after AS-501 flight todetermine fin vibration levels. These are to be incorporated into AS-503instrumentation. Maximum bending moments on the two instrumented fins(Figure 9-18} were approximately §0 percent greater than those encounteredduring AS-501 flight but were far below design load. Since maximumvehicle angle-of-attack in the high dynamic pressure reqien occurred inthe pitch plane, it is deduced that all four fins experienced approximatelyequal angles-of-attack, zerodynamic Toading, and bending moment,

9.3 VIBRATION EVALUATION

9.3.1 S-IC Stage and Engine Evaluation

The S-IC structure, engine, and componert vibration measurements (locationsof measurements are shown in Figures 9-18 and 9-19) taken on the S-IC stageare summarized in Figures 9-20 through 9-22 and Table 9-2.

9.3.1.1 S-IC Stage Structure. Vibration levels in the thrust structureat liftoff were similar to static firing levels but were lower during theremainder of AS-502 flight. This trend was expected and is similar toAS-601 flight deta. The intertank structure and forward skirt structure

9-20

nan

Figure

9-15.

Late

ral

Moda

lFrequencies

and

Accelerations

During

5-IC

Powered

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Figure 9-17, S-IC Fin Vibration Response and Bending and TorsionalModal Frequencies

9-22

F10-192,E0118)

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Figure 9-1B.

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9-23

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LOX TAKK FORWARD SKIRT

Figure 9-19, S-IC Vibration and Strain Measurements at Intertank,LOX Tank, and Forward Skirt.

9-24

9-25

LEGEND

MMM

STATIC

FIRING

DATA

ENVELOPE

--O-

AS~501

FLIGHT

DATA

ENVELOPE

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AS-502

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9-20.

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S-IC

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Envetope

9-26

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9-22

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9-28

Table

9-2.

S-1¢

Stag

eVi

brat

ion

Summary

AREA

MONITORED

MAX

LEVEL

rms

RANGE

TIME

(SEC)

REMARKS St

ructure

Engine

Components

Forward

Skirt

E18-120,

£19-120,

£47-120

Intertank

Structure

E20-118,

E21-118

Thrust

Structure

£23-115,

E24-115,

£49-115,

E79-115,

£80-

115

Combustion

Chamber

£36-

101,

E26-

103,

E36-104,

536-105

Turbopump

E41-102,

E42-102,

£42-103,

£42-104

Propellant

Delivery

System

LOX

Feed

Line

and

Pressure

Yolune

Compensator

Support

Bracket

£25-118,

£26-118,

£27-118

Cold

Heli

umLi

neE50-116,

ESt-

116

Engi

neAc

tuat

ors

E30-101,

£30-102,

€31-101

E31-102,

£32-101,

£32-102,

|€33-101,

£33-102,

£34-701,

£34-102,

£35-102

6.9

8.7

13.4

12.4

26.7

10.1 6.7 14.60.

2

5.7

720.5

[136.0

j-0.2

9.0

17.0 ia

ximu

rGrms

leve

lof

13.4

1s1.27

Grms

higher

than

static

firing

levels

but

the

spectra

are

very

sin

show vibration levels similar to AS-801 flight data. The jevels for launchand throughout AS-502 flight were lower than static firing levels.

9.3.1.2 -1 Engines, Four of the five F-1 engine combustion chambervibration measurements yielded valid data. Overall rms Tevels were higherthan static firing data. Al) five vibration measurements on the com-bustion chamber of AS-501 were invalid. The engine turbopump vibrationJevels, although slightly higher during AS-502 flight, were generallycomparable to static firing and AS-501 flight levels

9.3.1.3 $cIC Stage Components, The responses of components on the S-IC,engine actiatore- cold heliumTine, and propel lent delivery system aresummarized in Figure 9-22 and Table 9-2. The engine actuator measurement:showed amplitudes similar to static firing date and somewhat higher thanAS-501 flight data. Generatly, cold helium line measurements shovedJevels lower than static firing levels and similar to AS-501 flight levels.A11 AS-502 flight cold helium data were invalid before 3 seconds. Measurc-nents taken on the propellant delivery system show data similar to thestatic firing and AS-S01 data. The constant level throughout flightindicates that the vibration was a result of flow dynamics and not af-fected by acoustics. On the vibration isolated panels, the only validdata obtained was from one measurement at launch. This measurement in-dicated very low level vibrations as expected and therefore is not shownin Figure 9-22.

9.3.2 S-I1 Stage and Engine Evaluation

S-TI structure, engine, and component vibration measurements evaluatedon the S-IF stage are summarized in Figures 9-23 through 9-26 and Table

9-3. AS-501 data shown for comparison have been updated since releaseof the AS-501 flight evaluation report. Spectral analyses indicated thatthe basic AS-502 spectral shapes were the same as for AS-501. The energyis concentrated in narrow bands with major peaks occurring near 100 hertz.

9.3.2.1 S-II Stage Structure, The trends were as expected on the aftskirt, thrust cone, and interstage; however, forward skirt in the liftoffenvironment on the flame bucket side away from the tower exceeded thesine and random criteria. These exceedances occurred at frequencies near100 hertz on both AS-501 and AS-502, The exceedances were more proncuncedon AS-502 but no adverse effects were noted on either flight.

9.3.2.2 S-1l Stage J-2 Engines, The 1 S-11 engine vibratien measure-ments (combustion domes, LOX pump, and LHg pumps) for AS-501 and AS-602were considered invalid because of amplifier saturation at frecuerciesabove 3000 hertz.

9.3.2.3 S-I] Stage Components, Results were within design levels exceptfor the forward Shift containers which, Vike the forward skirt structure,axceaded the design criteria for about 1 second during the liftoff period.No failures occurred in the affected equipment.

9-29

Ywoes Yanna i sersen.

Jarame Dug 1 ne 5 crsTeaco santa Taantban tal = aEa_i amr Sms, frtn ge

three jooLets t

27 rede veka, atin FoalsLt ett ‘

pres fas uae CeveLe sya a

Figure 9-23. S-IT Stage Structure Vibration Envelopes at ForwardSkirt Stringers and Interstage Frames

9-30

re sri, ocWp pene stan seo

June ganasGone Beem

Fredsete, ancva oe

" ff

ecm ame uate

EEE

7

ree ave ioe Lcaecs

Figure 9-24, S-II Stage Structure Vibration Envelopes at Aft SkirtStringers and Interstage Frames

$-31

9-32

Fi

gure

9-25

,S-II

Stag

eStructure

Vibration

Envelope

atThrust

Cone

Long

eron

s,En

gine

1Beam,

and

Engi

ne|Gimbal

Pad

Peeves penne on ne, .36prea ne Wriey came ges Aye ae mer,

= I —." —-4-t-

MetTEAH

Figure 9-26. S-I1 Stage Component Vibration Envelopes

9-33

9-4

Tabl

e9-3.

S-IE

Vibrations,

Vehicle

Structure

ZONE

NO.

OFMEASUREMENTS

S-E1

501

AND

502

STATIC

FIRINGS

VEHI

CLE

OVERALLGRMS

TRANSONIC

Saul

MATHSTAGE

Forw

ard

Skirt

Containers

80,7+2.1

AS-501

AS-502

Forward

Skirt

Stringers

101,

6-4.

8AS-501

AS~5

02

Aft

Skir

t10.1-31.7

AS-501

AS~5

02

Interstage

Inte

rsta

geNo

tInstat].

AS-501

AS-5

02

Thrust

Structure

Containers

2.2-15.8

AS-501

AS-5

02

Thrust

Structure

Long

,4.1-12.3

AS-501

AS-502

Engine

Beams

5.4-15.4

AS~501

AS~5

02 Gi

mbal

Pad

8.

8

AS-501

AS-502

9.3.3 S-IVB Stage and Engine Evaluation

Nine vibration measurements were made on the structure, twenty-two atcomponents and six on the engine. Measurement locations are shown inFigure 9-27. The maximum composite (50 to 3000 hertz} vibration levelson the structure, forward components, aft components, and engine aresummarized in Figure 9-28 and Table 9-4. For comparison purposes, thevibration levels are shown with measurements taken during AS-501 flight.

9.3.3.1 S-IV8 Stage Structure and Components, The maximum vibrationlevels measured on the S-[78Seructureware <Tightly. lover on AS-802_ thanon AS-501, Forward component maximum vibration levels were greater onAS-502 than measured at similar locations during the AS-S01 flight. Themaximum vibration levels measured at the aft components were 70 percentof those measured at similar locations during the AS-501 flight.

9.3,3,2 S-1VB Stage J-2 Engine, The maximum vibration levels measured‘on the engine were almost identical to those measured during the firstS-IVB burn of the AS-501 flight,

ACOUSTIC, EXTERVAL Feta spiceRD STERNAL 08.1

"APS HOOULE NO. 1 ACOUSTICS EXTERKALFAD & AFTLrg TURBOPL™®

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Figure 9-27, $-IVB Acoustics, Vibration and Dynamic Strain Measurements

9-35

ELERAT

ION,

Come

SCCELL

RATTON

,Com

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Figure 9-28. S-IVB Stage Vibration Envelopes

9-36

9-37

Table

9-4.

S-I¥B

Vibration

Summary

AREA

MONITORED

MAX

LEVE

L.(Grns)

RANG

ETINE

(SEC)

RENARKS

Stru

ctures

Component

(LHg

Tank)

Engine

Component

(Fad

Skirt) Se

para

tion

Plane,

Pos

II-

Thrust

Fietd

Splice

Pos

I(1f)

-Th

rust

Field

Spli

cePo

sI

-Th

rust

Field

Spli

cePo

sI

-Pitch

Field

Spli

cePos

I-

Yaw

Fiel

dSp

lice

Pos

IThrust

Fiel

dSp

lice

Pos

11-

PiFi

eld

Spli

cePos

11-

¥

LH2

PUPr

obe

Inpu

t-

Radial

Gimbal

Point

-Th

rust

Gimbal

Point

~Pi

tch

Gimbal

Point

-Ya

wCo

mbus

tion

Chamber,

Dome

-Thrust

LOX

Turb

opum

p-La

tera

l

PUEl

ectr

onic

Pane

lInput

-|

Thrust

PUEl

ectr

onic

Pane

lInput

-Radial

PUEl

ectr

onic

Pane

lRe~

|sp

onse

-Radial

EBN

Range

Safe

tyPanel

Input

-Th

rust

EBN

Range

Safe

tyPanel

Input

-Radial

EBM

Range

Safe

tyPa

nel

Response

-Radi

alBattery

No.

1Input

-Thrust

ae ame6 tt 4.

3

10.6 5.5

3.2

Inva

lid

2.2

2.1

Bg 89 89 80 86 83 745,

739

739

589

5R6 78 EA

The

maximum

vibration

due

either

tosound

ipre

ssur

eat

liftoff,

turb

ulen

ceat

maximum

ldynamic

pressure,

orto

J-2

engi

neop

erat

ion

Themaximum

vibration

occurred

during

J-2

engine

star

ttransient.

9-38

Tabl

e9-4.

S-IV

BVibration

Summ

ary

(Con

tinu

ed)

AREA

MONITORED

MAX

LEVEL

(Gri

ns)

RANGE

TIME

(SEC)

REMA

RKS Co

mpon

ent

(Fwd

Skir

t)(Cont)

Comp

onen

t(aft

Skirt)

Comp

onen

t(T

hrus

tStructure)

Battery

No.

1Input

-Ra

dial

Battery

No.

7Input

-

Tangential

Sequencer

Panel

Thrust

Sequencer

Panel

Radi

alSe

quen

cer

Panel

Radial

Swit

chSelector

Pane

lInput

-Th

rust

Swit

chSelector

Pane

lInput

~Ra

dial

Swit

chSelector

Pane

lResponse

~Radial

APS

Wod-

1Aft

Atta

chPo

int

Input

-Thrust

APS

Mod-

1Af

tAttach

Poin

tInput

-Radial

APS

Mod-

1Fw

dAt

tach

Point

Input

-Radial

Inpu

t=

Inpu

t-

Resp

anse

~

HeBo

ttle

Input

-Th

rust

HeBottle

Input

-Pitch

HeBo

ttle

Input

-Yaw

Li2

Feedline

Input

-Thrust

42

1.9

72 78 78 78 83 80 80 75 7 589

589

Due

toa

loose

conn

ecte

r.

9.3.4 S-IVB Stage Forward Skirt Dynarics

Sixteen dyramic strain gage weasurerents were placed on the AS-$02/S-IVEforward skirt for the purrase of investigating possible panel fluttercharacteristics during the supersonic flight regime, These measurementswere located at vehicle station 79.64 weters (3135.8 in.) and placed ap-proximately every 22.5 degrees around the circurference of the skirtsection as shown in Figure 9-27, The measurement nurbers were SQ086-426through S0101-426, Positioning of each strain gage was such that it wasmounted 10.16 certimeters (4 in.} forward of the panel trailing edge.This location was ckosen because data was obtained at the same point duringearlier wind tunnel tests. The envelopes cf the naximur and rinimur cor-posite (0 to 800 hertz) strain levels measured at these locations areshown in Figure 9-29. The envelopes from the AS-204 flight are alsc shown.

The time history of the composite dynamic strain levels from most of themeasurements followed the same trend as the acoustic levels measured onthe forward skirt shown in Figure 16-18 (Section 16).

Angle-of-attack and differential pressure across the paneis are important.

parameters to consider when assessing panel flutter severity. Figure 11-10

{Section 11) shows the angle-of-attack history for AS-502 and Figure 9-30

shows the differential pressure history. For angles-of-attack smaller

than two degrees, all the panels are assumed to be buckled due to axial

loads alone. At larger angles-of-attack, a load relief ts experienced on

the windward side of the vehicle and a higher axial load exists on the

Teewsrd side. The panets would be most susceptible to flutter near thecritical buckling load of the panel.

The differential pressure time histories across the panels at station:80.49 meters (3196 in.} and 78,99 meters (3110 in.} are shown in Figure9-30. The dynamic strain gage location is approximately midway betweenthese stations. These differential pressure loads were calculated byusing the internal compartment pressure measurement 00051-411 and externaaerodynamic data obtained from the AS-204 flight. The presence of apressure differential across a panel will tend to decrease the flutterpotential and/or suppress the resulting panel flutter stress amp? itudes.Angles-of-attack greater than about two degrees will decrease the pressuredifferential loading on panels on the windward side of the vehicle makingconditions more favorable for flutter to occur.

The flight data showed a random response during the liftoff, Nach 1 andWax ( regimes. This response is typical and results from engine acousticand inflight fluctuating pressures. In the regime after Max Q,77 seconds s t < 92 seconds, measurement numbers $-90, $-92, S93, S-94,5-97, 5-100, and S-101 show a complex periodic waveform and a correspondingincréese in'strain amplitude. These waveforms are characteristic ofbuckled panel flutter and indicative that panel flutter did occur duringflight. Typical examples of these strain time histories are shown fn

9-40

QQ

1s-n02

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+5208

260

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189

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TIME,

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00:11:00

00:72:00

03:10:00

93:18:00

RANGE

TIME,

NOURS:

NINUTESS

ECONDS

14,800"

11.700

Figure

9-29.

S-IVB

Forward

Skirt

Dynamic

Strain

as1

ey Bswitorg ure]

—

Ne

:STATION 7,994 PRTERS.

B sos | Siem

|

_t

Sol.z

1

é13 8

Es - e

© iN12

gaa §

10ts y 3

»“

ST TT Ee TT 88no maeE?

Figure 9-30. Pressure Differential Across $-IVB Forward Skirt Panels

Figure 9-31, $-100 and S-101 exhibited the highest strain awplitudesduring flight. These values were 1750 y in./in. and 2000 in./in. ofStrain and occurred at 88 seconds (2.2 5 Mach < 2,4) and 80 seconds(1.7 «Mach < 1,9), respectively, The flutter frequency was approximately150 hertz for 5.100 and 300 hertz for S-101. An mis strain history forthese respective flutter frequencies is shown in Figure 9-32, The differential pressure across the panels at this time was approximately0,83 N/om2 (1.2 psid).

It is concluded from these data that pane} flutter did occur on the above

S:1VB forward skirt panels. The pressure differentia) across the panels

Suppressed the stress amplitudes to a tolerable level for the AS-502

Fright, Although these “Tuster amplitudes tended to be sunpressed by the

very high aP, they were about three times higher than those measured on

28-204, Detailed analysis and evaluation of these data will provide

guidance for future action with respect to vent area criteria and/or

structural fixes.

9.3.5 Instrument Unit Evaluation

Eight measurements were used on the IU for monitoring structural vibration

at the upper and lower interface rings. Twenty measurements were used tomonitor 1U component vibration levels. For comparison purposes, the IUStructure and component measurements are shown with those taken duringthe AS-501 flight. Figure 9-33 shows the Grms time histories of thesemeasurements.

9-41

9-42

pana

SER

RANGE

TINE

=7:

DsRANGE

TIME

~82

SECONDS

RANGE

TIME

=92

SECONDS

MEASUREMENT

NO.

S-94

L

noemeae

4AA

AAA

MhwweTi>orcas

rn

aed

rieata

NEGLIGIBLE

100

panna]ANAMal

RANGE

TIME

=68

SECONDS

RANGE

TIME

=80

SECONDS

NEGLIGIBLE

MEASUREMENT

NO.

10T

Petcncccnmneneonmma|

—

ftwninrlniamirinbatyiie|

—fauna

oenmeannnnnl

RANGE

TIME

=69

SECONDS

RARGE

TIME

=81

SECONDS

RANGE

TIME

=93

SECONDS

MEASUREMENT

NO.

92Figure

9-31

.Time

Stices

ofDynamic

Strain

Output

Showing

Wave

Forms

STRAIN

AMPL

ITUD

E,micro

in.fin.

140

120

100

80

60

40

20

MEASUREMENT NO. 100

MEASUREMENT NO, 101

ro

1 20 30 a0 GCCRANGE TIME, SECONDS

0.1 OR 1.0 2.0MACH NUMBER

Figure 9-32, S-IVB Forward Skirt RMS Strain Amplitudes

9-43

Grms

10

1 T 4sTRUCTURE | £3 AS-501AS-502 P12

| a : F 10

i ~~be

t ee

ee o r4

2m, | L >

L 4 to. 060 Ed 700 124060RANGE TIME, SECONDS

10-— . 1 4COMPONENTS Ea As-s018 as-s02 Fle

r 10

6 +8

4 6

Fa2

r2

9 Lo

Figure 9-33,

60 80 100 120 140160RANGE TIME, SECONDS

Instrument Unit Vibration Envelopes

9-44

Gpeak

Gpeak

9.3.5.1 Instrument Unit Structure, The structural vibration levels atliftoff ware higher on the AS-5OT. At Mach 1/Max Q, the levels werehigher on AS-502. After S-IC powered flight, the levels became negligible.

9.3.5.2 Instrument Unit Components. The instrument unit componentvibration datsindicated«broader range of data than that of the structurevibration measurements. This is due to different response characteristicsof the various components. The AS-01 component vibration Tevels exceededthose of AS-502 from liftoff to approximately 56 seconds range tire. Fror56 seconds range time through maximun in-flight load to approximately120 seconds range time, AS-502 component level exceeded those of AS-501.The vibration levels during S-IT and S-IVB powered flights were negligible.

There were no vibration induced malfunctions of the ST-124¥-3 inertiaplatforr on AS-S02. Available data indicates that the ST-124V-3 compositevibration levels at liftoff on AS-502 were very near these of AS-5C1 orthe inertial gimbal. The composite levels of the inertial gimbavibrations are shown in Figure 9-34.

9-45

VIBRATION,

Grms

‘ INERTIAL DOWNRANGE AXIS

* oilsabioth id Tl srboie saiahada

scl

INERTIAL VERTICAL AXIS

4

Bs2 > 7«| =

0 bas“ |,

ta

cinlainhainbaad

4 INERTIAL CROSSRANGE AXIS

te j

ate

i,

a Te seete0 ites tlabasctevnt fe-10 0 20 40 60 80 100 120 140

RANGE TIME, SECONDS 9090501

#46502

Figure 9-34, AS-502 Versus AS-501 Inertial Gimbal Vibrations

9-46

SECTION 9A

133 SECOND TRANSIENT

9A.) SUMMARY

At approximately 133 seconds abrupt changes of strain, vibration, andacceleration measurements were indicated in the S~IVB, Instrument Unit{IU}, Spacecraft/Lunar Module Adapter (SLA), Lunar Module Test Article(LTA}, and Command and Service Module (CSM). Photographic coverage,Airborne Light Optical Tracking System (ALOTS), and ground camera filmshowed pieces separating from the area of the adapter. There were noknown structural failures noted on the launch vehicle

Al] data from both the launch vehicle and spacecraft relevant to this133 second anomaly have been investigated by a joint task group at theManned Spacecraft Center (MSC). The results of this investigation arebeing published in a separate anomaly report by MSC.

QA.2 INSTRUNENT UNIT

Attempts were made to establish a timeline correlation showing where thedisturbance started and its propagation through the vehicle. A timetineof events from each data signal correct to the nearest millisecond wouldhave been necessary in order for the timeline to be valid. Problersassociated with the accomplishment of this task are enumerated below:

a. The primary problem is the interpretation of the start of the event.These differences vary from 1 or 2 milliseconds to as much as 20milliseconds. Forty-inch-per-second oscilTograms were used to ald inthis determination.

b. Delays inherent in the transducer, telemetry link, and receivingstation were identified with a step function input at the 10 percentand 90 percent levels; however the input to the vartous transducerswas not necessarily a step function, but a slow changing DirectCurrent (DC) level followed later by a high frequency shack. ThisCould cause additional undetermined delays up to 20 millisecondsdepending upon the Inter Range Instrumentation Group (IRIG) channetand the type of input.

c. Signals that should have had close correlation did not always showcorrelation.

SA-1

d. Submu7tiplexed data had a time span over which the event could haveoccurred.

Despite problems with establishing a meaningful timeline, an IU timelinewas wade. Figure 9A-1 shows that the single sideband measurements (wit!the prefix E) responded earliest to the disturbance, This is expectad because the transient is readily discernible due to tie low level noise andthe BC signal level prior to the disturbance and the viga (3 kilohertzresponse of the vibration measurements. Note that the first measurementto indicate a change was on a Radio Frequency (RF) assembly rather thana hard mounted transducer. It would be expected that the hard mounted trans-ducer would note the shock first. This indication, coupled with relation-ships of times of vibration measurenents on the upper and lower mountingring, illustrates the difficulty in attempting to locate the source of thedisturbance by the use of a timeline. As shown in Figure 9A-2, measurement E14, which is on the upper ring, has a higher G level and precede:E16 which is on the Tower ring; however, £39, wnich is on the upper ring,has a higher G level but occurs after E41 which is on the Tower ring,Also, £43, on the ST-124M-3 bracket, preceded all the upver and Towerring measurements as shown in the timeline.

An attempt was made to see which axis was affected first but this resultedin the same type of inconsistencies as listed above. The investigationresulted in a list of all the IU related measurements that exhibited aresponse at the time of the disturbance. This list is given in TableQA-1. The indication column of Table 9A-] briefly describes the dataresponse to the disturbance for eaca signal. A blip is defined as a smallexcursion from the nominal that returns shortly to the nominal.

The tolerance column indicates the accuracy of the time but does not accountfor the problems previously discussed. Tne adjusted time colum takes intoaccount. transducer, telemetry, transmission, and receiving station delay:in order to show when the transducer actually detected the event, Each ofthe parameters will not be discussed separately but will be grouped formore meaningful presentation.

98.2.1 Mechanical Versus Electrical Disturbance

A discussion of the nechanical versus the electrical effects at the timeof the disturbance points toward a positive mechanical primary event withmost secondary electrical effects readily understood. Table 9A-2 show:the disturbance to be primarily a mechanical shock because the equipmentmentioned oscillated at their respective resonant frequency. Measurement:46 and A7 were limited by the telemetry to a 25 hertz response. Theequipment would not respond at their resonant frequencies if the disturbancehad been purely electrical.

The ST-1244-3 servo loops, in reaction to the shock, drew a higher thannormal current from the 56 volt power supply. The supply in turn placeda larger current demand than normal (greater than 9,5 amperes) on the

9A-2

Je13-607 YIGRAII04 RADIO FREQUEMY ASSEMBLY LOCATION 9

{43-603 VIBRATION ST-174 GRAGKET, LONGITUDINAL LPPER LEFT LEG LOCAT:ON 217-602 VIBRATION $¥-124 SUPPORT, LONGITUDINAL LOCATSON 21E14-603 VIBRATION UPPER MOUNTING RING, LONGITUDINAL LOCATION 18[37-603 VIBRATION ST-124 SUPPORT, LOWER RIGHT RASE, LONGITUDINAL LOCATION 21 16-503 VIBRATION LOWER MOUNTING RING, LONGITUDINAL LOCATION 18E28 803 {GRATION LVDC/LVOA TANGENT LOCATION 1941-03 VIBRATION LOWER MOUNTING RING, LONGETUDINAL LOCATEOR 21

[44-603 CUTPUT, Z GYRO SERVO

38-693 VIBRATION UPE2 MOUNTING PING, LONGITUDINAL LACATLON 7

|e2-e03 VIBRATION ST-124 INERTIAL GIMBAL X-AKIS

€1-603 VIBRATION St-124 INERTIAL GIMBAL Z-AXIS3-603 VIBRATION ST-124 IYERTIAL GIMBAL Y-AXISF9-603 VIBRATION ST-12d SUPPOST, PERPENOTCIL AR LOCATCON 24Ni4-603 1-124 X ACCELERATION SERVONiS-603 ST-124 ¥ ACCELERATION SERVOH95-603 OUTPLT, ¥ GYRO SERVO26-603 VIBRATION VOC/LYDA, LONGITUDINAL LOCATION 19H10+603 $T-12¢ 7 ACCELERATION PICKUP

H12-603 S1-124 ¥ ACCEL ERATION PICKUPH13-603 $T-124 2 ACCELERATION SERVO

MEASUREMENTS:

HY6-602 Z ACCELERATION, AlH46-603 OUTPUT, Y GYRD'SCRYO

[1-601 SUBLIMATOR Hg0 FLOWRATE20-602 x ACCELEKATION, a1133-602 COE MONITOR ANGULAR VELOCITY, ROLL GROUP NUWGER 11-300 EDS ELIA PRESSURE PITCH, O°6ALLD3-00 EOS ELTA PRESSURE YAM G-BALL[62-101 yaw acTuaror 2osrT4oN

| 7-603. YAM ACCELERATIONH27-603 Y ACCECERALLON, B1

6-603. PITCH ACCELERATION26-602 ¥ ACCELERATION, 42

[H11-603 ST-124 X ACCELERATION PICKUP

1Gt-101 PITCH ACTIATOR POSITION

[92-663 LONATTUDIWAL AcCELERATION1 ~ 133.27 133.28 135.29 733,30

RANGE TIME, SECONDS

Figure 9A-1. First Instrument Unit Measurements Affected by the133 Second Disturbance

A-3

spuoras¢¢|

42PaLON

B4ayIeUL

SqUAAQ4Sat[4e7

youOL3e201

-z-yEaunByy

ato“eee

BOTS49

1 woTdtsoa

DAA

Table 9A-1. List of IU Measurements With 133 Second Response

iw OATH ADOUSTEDMeRsUROHENT - rouerawe rine Time

NUMBER MEASUREMENT NAME, FISICA sec) (a3 sce) (093 seco)92-602 Lorgi tudinal Accelerometer “soo FP) +z, -8 oz 0.20926-603 Piten Accelerometer searing 15 0.295 0.28927-603 Yow Accelerometer searing 35 0.293 0.2881-601 Accustie 23° OFF Position f

to Position Rings $20 om 0.29e202 Flight Contre! Computer (FCC)

Hethanol Water (W/W) ExitTenperature +aup 2, 63} 0.207 0.705216-01 1 Anbient Temperature inflection #1000 D te 0.1) oteo770.602 FCC Temperature Blip so, -100 0.3 0.82-601 Thermal Conditioning Syscem

{TCS} 642 Tewperature = B10 va, 10 0.3 03plo-s03 Gos Bearing System (388) ON2

Regutator Pressure Inlet Steps down 450, -a9 0.3 0.3Die-601 S-IVE M/W Exit Pressure Rings. re ee ce)28-601 WIN Puro Inlet Pressare Steps UpsLevels bt +80, -100 bt oe25-60 Gig RequlatorInlet Pressure

(tts) = he +80, 100 0.3 0343-01 Sublinator Veter trlet Pressure + Sten 42, 23 0.207 0.2051-608 Inertial Z Axts GinbaT Vebratton 12 6 PP Rings 25 0.262 0.2802.603 Inertial X éxts GinbaT Webratton 12.6 PF Rines 25 oz 0.2793-603 Inertial ¥ Aas Ginbal Yebration 206 PP Rings 35 o.2ez 0.2807-608 Longitudinal $t=12@M Support,

Vibrator 23.6 PP Rings 45 0.2% 0,278es.609 Perpendicular ST-124K Support

vibration. 316 PP Rings 95, p.2az 0,200514-603 Longt tuding Upper Mounting

Ring Vibration 47.6 OF Rings 45 0.276 3.27416-602 Long {tudinal Lower Mounting

Ring Vibration 326 PF Rings 45, v.27 9.276rre-6oe Long’ tudinal Radte FrequencyTEED Aesenoiy Vibration

42 Panel) 96 PP Rings 25 om 0.27226-603 Longitudinal aigital Computer

Data Adapter Vibration 4.5.6 PP Rings!) 45 oa 2828-603 Tangent Digital Corputer Data

Aéapter Vibration 5.56 PP Rings 25 oe 0.275£37-603 Low Right Base ST-128 Support

Vibra tien 7.56 PP Rings| 45 oz over39-603 LLongt tudina] upper Mounting

Ring Vibratton 16 6 pP Rings 45 oa 0.277ee1-603 Langstudina Loner Move ing

ing Vibration 2G rr Rings 35 oz? 0.273.608 ‘Upper ST-124M Bracket vibration 136 PP Rings 25 piers 0.273H-60r eo Flowrate Rings 8 o.ae 0.2842-608 Bypass W/L + mip 2, 93 ous 0,3873-601 Cold Plate Inlet Coolant

Location § + aMp 42, 09} 0.356 0.358F4-603 Cold Plate Inlet Coolant

Location 20 + ap 42, -nx oes 0.087F5-603 ST-126M Shroud tnlet Coolant] + Blt 2,33 ows os

SA-5

Table 9A-1. List of IU Neasurements With 133 Second Response {Continued}

ian ayoPeasarers

|

easurvent ae rocntion

|

Tagpanct | TomefTrécete

|

vgn tartel Comber vane v8 aterego

|

told Plate Location 4 one we, es sesscaca

|

paaoterrcmprter NM ete

|

3p ao 6.298eoece

|

tuseit cootone, et «50, -100 asrro-ecr

|

settelet cootant = atin ss0, 100 onFiceos

|

st-t2eshoud cuulant ei seo. oiso) pieeh Acuntor Position uns a os a2GI-162, | piten Actuator festtion + Mis 8 0,384 038gsiviaa

|

itch etuotor Fsitin bein 8

|

ose

|

ossfick

|

piten Retintor Posten ote vs

|

os

|

oserete

|

aw Actator Position =muip tian |

vo,

-e

|

oer

|

anecece

|

mau dctator Pasieton sone 3} owe

|

easants

|

Yaw acturtar Psttion vanip tmny]

sot

oe

|

ase

|

oceanconan

|

Yow Aotutar Position toe toin}[

v0,

| ose

|

aleswneiot

|

peck tetntor ove treent

|

"atte | os

|

osssor pitch fetuntos delve urea oa

|

onvecaea—

|

Pitch feantr vove corsent om

|

oerie

|

pitch fetntor delve Geren ass

|

0.34wea] aw tctntor shee torront | Blip aa

|

ott

|

cansvests [fav retanter Yeve Overt, os

|

osevesiai Yow Potato YelCoron o.oo

|

0.208vane |Petar Yeon Crvent ola

|

0.308inossoa

|

striae aeeteronter rekup

|

eins aa

|

oe

|

orwnrsgca

|

steioe t peoteroreter tsar

|

nis 1a

|

oz

|

casosna

|

str1at deetermeter Pekan se cn

|

oes

|

o.zeevonens

|

st-124 2 decetermeter Serve ea oes

|

omerrecen| saree x pecelermter serve

|

bie be zm

|

ozoins S112 fentoroter serve

|

6ttp ‘se ae

|

c-200inessor | teeteronter, Sonal thnged| 38 zea

|

oanvea-soa

|

& seceereeter, a1 Samal thet

|

a ouzee

|

comvetcsoa | secteromtera get crvaed

|

38 acs

|

a.z07ves.eca

|

fecetermeter, £1 Signet angel] 38 com

|

ozsvancers

|

1 feceeronten, #2 risen

|

38 ozan

|

o.ze9verso

|

¥ decleroneter, 82 siwatrn [ae ere

|

0.088vas.cea |ePrattor “enn v2, | oe

|

omevae.sea

|

ante a2 mz, servo fm Sunpty

|

= mip ssoate

|

oa

|

osFsyro Phy S120" teccinstor {v2 -e

|

ota

|

oo1 5yre vekaps ST towursion [ve Jn

|

os faeY fjre Pehups S120 tht a

|

oan

|

atutnut 2 So Seve figs 7542

|

x0 eo

|

omims.cea

|

cutest 6yro Serve ange sot

|

0 oz

|

ozovassiea

|

cutnot Bye Serve rings ie

|

20 outes

|

oasessez

|

nr rmosue A teeter | nee [2 a

|

o's | oe 3A-6

Table 9A-1. List of IU Measurements With 133 Second Response (Continued)

cance utaY toausieneeasuRenexT sur Touswance Time TNEMER PFASURERENT NAME INDICATION sec) (933 see+) (093 sece}2b-ie Pr Retiactive Fewer, FT Televeter (Stayed Un) +2, a 0.388 0.357327 6b RE Power Guteut, F2 Telewter | asp 42, 08 nas 0.3m38-602 PF peflactive Power, F2 Teleneter |- alip 42,| cas ose29-602 RF Power Output, FI Teteneter [+ offset 42, 8a na 04330-602 RE watlacttve Power, PI “eleneter + offset 7, a oss 0.397a1 -602 RF Power Cutput, SI Teteneter [+ offset +2, 3 0.309 0.a022-402 BF Rotlective Power, SI Teleneter |- 8112 +50, -100 os o.3091175-13 State Phase Error + Bie a2, 6a o.zsr o.a56376-693 Comrand and Communication System

(CcS) Autoratic Gein Contrel (AGCines ey sein +2, 22 ost 0,587

el 603 Summation Gyro Currents + Blin +2, 29 oer o.se6stream)

62-603 Syrmation Accelerometer currents |+ B1ip +2, 2 oss 0.308tersran3-808 YeTt, $6 EDC Surply +51 +2, 03 oa 0.4806-603 Yerts 280 Yolt Arpere (A)Inverter, Phase AB - B50 v2, a3 oan 3.310nes VeTt, 260 A inverter, Phase cA [4 Blip wz. ga foes o.a6eviz-en 6C11 Sus voltage “Be 2, | ase one13-691 e021 Bus Yortage +p $60, 100 03 0314-60 BNE Bus VoTtage - Blip +0, ton 08 0.818-601 6010 Battery Current, + Blip +2, 83 0.30 0.34417-601 6020 Hattery Curcent + Ele +50, 100 a 0.318-60" 6000 Battery Current + mip +2, 2 o.ser 0.3503-60 6041 bus Voltage + Offset, 450, -109 td 0.320-60 6040 Gattery Curent: + Blip +2, 63 0.295 o.200ver-003 Bata Adapter WY Sunwly on 8, a] oa 0.376Re.602 angular Velocity Pitch Control + Elin 6 oss 0.33rse6ae Angular Yatectty Yaw Control ~ Blip 8 0.308 0.305Ron? Angutar Yetoc*ty Rall contrat - Tip 8 0.297 0.20887-602 Angular velee*ty Pitch Fmergency

Detectfen System (FOS) GroupNusber 1 spare + ulin sso <0 0.3 2.2Ra-698 Angular velocity Yaw EOS Graun

aber 1 + our 8 0.322 0,321a Angylar Veloc! ty Yaw EDS Group

anber 2. (Spare + Bip #100 08 a312-602 Angular Velocity Roll EDS Groupharber’ 2 (Reference) + Bin 8 0,309 0.208Ra3-02 Angular Velocity Piteh EDS Groun

pumer 3 (Reference up Fa 0.30 a.36Rya-602 Anqutar Velocity fav EUS Groun

Munber 3 {Comane) woicy $80 oie etsis-60 angular velocity Roll EDS Groun

umber’ 3 (Spare) + oltp v2, 1] ose 0.34033-602 EOS Monitor, Angular velocityGroup hunker’? = BNP +2, «8 0.206 0.285R34-802 EDS Monitor, Angular Velocity

R Group Nimber 2 = Bite ve, 80 o.san o.ae9 9A-7

Table 9-1. List of IU Measurements With 133 Second Response (Continued)

way TATA |AO9USHED

easuncnent ‘ Touepance TIME TIMEaunt MCRSUREMENT Wan, Inplcarron taSrey

|

(133 Shoes [aat teeny36-662 EDS Wontar, angular veracityWGroup tmber 3 + Bite oz, aa] o.s39

|

oe36-602 Angular Velocity, Piten (EDS

nut) + Bup ca s.s o.ara37-602 Anqutar Yetoc'ty, RoV ¢EDS

Input) + Eup 6 ons

|

asiaga-6n2 Anqutar VeToctty, Yaw (EDS.

Input) + Ble 8 baa 0.371-800 FOS «P Q-Ba1l Bees nec

‘Then Returns 8

|

oa

|

0.2863-800 EUS 4P Quba1 Goes Nec

Then Retuens a] ozee

|

a.ae5038-900 EDS q-baNl Summation Goes NEG

Then Returss

|

42, 93

|

0.31

|

0,280Ho-6n3 Gutdance Computer Operattars

Fine Sinbat Argiesx (R071) 7 Are #0 o.2e0e

|

0.280"in Geet Tate¥ (item) 40.8 Are $80 0.26

|

0.262Hin crrset2 (yaw) “5.5 Are +0 nase

|

o.asoewin cfFeetLadders

kon Pulses $80 6.265"

|

0.265Piten Pulses $50 c.sase

|

a.695Yam Pulses $50 c.zeor

|

0,280

‘Times not accurate due te iffiovlty In tiring LVN date

DAB

Table 94-2, 133 Second Transient Survey, IU Stage

MEASURENENT FREQUENCY

NUNBER NAME RESONANT

|

133 SECOND{HERTZ} (HERTZ)

46-603

|

Pitch Body Fixed Accelerometer 190 *

7-603

|

Yaw Body Fixed Accelerometer 100 .

£7-603 |S7-124 Leg Vibration,Longitudinal 100 100

£9-603

|

ST-124 Leg Vibration Tangential

|

About(Stiffened in Radial Direction)

|

140 140

#10-603

|

Z Accelerometer Pickup 40 30-40

H11-603

|

X Accelerometer Pickup 40 30-40

H12-603

|

¥ Accelerometer Pickup 40 30-40

4-502

|

Pitch Rate Gyro 17-21 WV

R5-602

|

Yaw Rate Gyro 17-21 W

R6-602

|

Roll Rate Syro 17-21 W7 “Measurement Bandwidth Limited at 26 Hertz.

6D10 bus. A momentary short in the 6D91 bus may have a1so contributedto the problem (see paragraph 94.6). These could result in transientelectrical effects throughout the IU and may account for some measure-ment changes that vere smal1 and unrelated to the mechanical disturbance.

Analysis of the 56 volt power supply regulation and efficiency factorproves that a surge of greater than 10 amperes on the 6D10 battery caneasily happen. Worst case condition of ST-124N-3 servo loop drain on thesupply would be close to 9 amperes, Normal system operation showsapproximately a 1.4 ampere drain. With a 65 percent efficiency, the inputcurrent to the 56 yolt power supply from the 6D10 bus would be 4.3 amperes.

For normal operation there is an input current of 4.3% amperes at 28 ¥DCassuming 65 percent efficient output current.

1, = (22449) (0.65) = 1.4 amperes

9A-9

Assuming the input current did go to 14 amperes at 28 VDC, the outputcurrent was;

1, = 428204) (0,65) = 4.55 amperes at 56 vnc

Therefore, this is normal operation under vibration conditions.

When the current surge on the 6010 bus with its conseouent voltage dropoccurred, there were excursions of a lesser nature noted on the 6030 andS040 buses. These fluctuations were apparently caused by other IU ponersupplies reacting to the changing voltage on the 6D10 bus. As a resultof these fluctuations, transients were induced throughout the IU asevidenced by fluctuations in many measurements,To discount the possibility of a measuring system problew causing the‘indications observed, an analysis of the various measurerent paths wasperformed. It was proven that no commonality of measurement discrepanciesexists. Measurements are powered by various buses, and are not signal-conditioned in any one measuring rack. Multiple cables, multiplexers,and telemetry transmitting equipments are involved.

Figure 9A-3 is a composite of vartous electrical indications in responseto the disturbance. The electrical effects are first noted on the 6040bus at 133.211 seconds. Also, it can be noted that an electricaldisturbance occurred at 133.261 seconds at the 6D10 bus.

The 6030 and 6040 bus current increases are considered to be due to loadsharing in the Launch Vehicle Data Adapter (LVDA) and Flight ControlComputer (FCC). When the 6D11 bus voltage dropped, the 6030 and 6D40buses assumed the Toad momentarily but returned to’ normal when the 6011voltage came back up.

9A.2.2 Pressure, Flowrate, and Temperature Measurenents

AN] of the 11 flowrate measurements and 5 of the pressure measurementsindicated a change at the time of the disturbance. The threshold for theflow measurements is Tow enough that they would readily measure accelera-tion changes due to the hammer effect. The pressure measurements alsoare susceptible to G-level changes of the magnitude noted.

It is not known whether the temperature measurements that changed werereal temperature fluctuations or ter‘ tary electrical effects, The levelsinvolved were so smaTl they are negligible.The sublimator inlet water pressure transducer, 043-601, sensed a suddenchange in pressure beginning at 133.307 seconds. This change was either@ plus pressure applied to the positive pressure port connected to the

SA-10

T3504

Ta3.3

Ta2-2

TasT

RANGE TINE, SECONDS

GAT)

a0

Composite Electrical Effects“3.

128

Figure 9A.

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water system or a negative change in pressure applied to the minus pressureport which was open to the internal IU compartment. The pressure increased9.069 N/emé {0.1 psid} from 133.3 to 133.8 seconds as shoyn in Figure 9A-4,Since the IU compartment shares a common pressure environment with theS-IVB forward skirt and the SLA; and ALOTS film showed pieces separatingfrom the area of the SLA, the question of measurement validity was raised.The conclusion that the measurement was valid and did in fact detect aPressure change at about 133 seconds was based on the following facts.a. The 0,5 psid pressure transducer has been used previously on sevenSaturn IT misstons (SA-4 through SA-10) with seven transducers used pervehicle. Tt was also used on the uprated Saturn I first stage onAS-201 through AS-204 with at least four transducers used per vehicle.Failure occurred on only one transducer, It is suspected that thetransducer may have been electrically unconnected,b. The transducer has been tested by MSFC under vibration and shock levelswhich significantly exceeded any IU vibration and shack requirements.No DC shift occurred during any of these tests.

¢. The maximum friction between wiper and pot would be two resolutionsor 0.66 percent full scale or 0.03 psi3 however, friction betweenwiper and pot is never considered to be a contributing error whenthere is @ known vibration present on the transducer.d. The effect of temperature and acceleration would not have produced anymeasurable amount of error. The smaliest measurable error would be9.33 percent full scale or 0,015 psi representing one resolution(wire turn) of the transducer’s potentiometer.

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Figure 9-4, Sublimator Inlet Water Differential Pressure

9A-12

ie. Vibration is not considered a contributing error because vibration will

not produce a change in DC level.

9.2.3 Radio Frequency (RF) Measurements

Several telemetry system pover measurements exhibited leve! changes atapproximately 133 seconds range time. The DS1, DP1, and DF] power outputschanged by approximately 0.8, 0.15, and -0.15 watts, respectively. Thecorresponding reflected powers changed 0, 0.05, and'0.1 watts, respectively.The OF2 power measurement levels did not change even though they showeda slight transient.

Only DFI showed a characteristic which may be directly attributed to aslight impedance change causing an electrical mismatch in RF componentsbeyond the transmitter output; however, the same impedance change couldhave improved the Voltage Standing Wave Ratio CUSWAY on DST and OPIresulting in the observed changes. Acoustic noise, vibration, or antennastructural stress changes could have caused a slight impedance change

The causes of tae observed level changes and transients have not beenidentified.

9A.2.4 ST-124N-3 Stabilized Platform Subsystem and Control SubsystemAnalysis

9A.2.4.1 Platform. Of the data available, the accelerometer pickupmeasurements seemto provide very significant insight into the phenomenon.These measurements indicate that a disturbance began at approximately133.3 seconds and continued for a period of about 100 milliseconds. Thenature of the deflection indicates an impulse type disturbance; howeverbecause of the short duration of the input, the accelerometer responsewould be of a transient type regardless of the nature of the input.

Factors that indicate a physical disturbance in the platform area are:

a. The platform required additional current load with servo disturbancesseen. At least part of the additional current drawn from the 5010 buswas due to increased requirements of servo loops.

b, Platform accelerometers showed disturbances of approximately the sameratios as seen during vibration. This would be highly improbable inthe case of an electrical disturbance.

©. Phasing was not the same for the three accelerometer loops. For anelectrical disturbance, the phasing would very likely be the same.

d. There was no irregularity on the 56 volt power supply that would savecaused the servo disturbance sean.

9A-13

The platform gimbal angle analysis indicates the following changes :a. Pitch. An aporoximate 0.8 arc-minute shift at 133.280 seconds.b. Yaw. An approximate 4.5 arc-minute shift at 133.350 seconds.¢, Rall. An approximate -4.7 arc-minute oscillation at 133.262 seconds

returning to null.

NOTE: LVDA times on these parameters cannot be accepted as factual dueto the problems in establishing LYDA times.The yaw and pitch transients responding to vehicle movement corroboratethe angle-of-attack changes that occurred as measured by the Q-Bat]. Thedelta pressure changes indicate that the vehicle nose moved in both pitchand yaw attitudes and then returned to null. The ladder outputs and gimbalangles indicate vehicle movement in pitch, yaw, and roll. (The Q-Balldoes not measure roll.)

In both pitch and yaw the vehicle assumed a slicht offset; whereas in roll,there were oscillations after which the vehicle anc gimbal showed realign-ment. The offsets in pitch and yaw did not affect guidance and navigationas evidenced by nominal S-IC stage cutoff conditions.9A.2.4.2 Control Rate Gyros. All of the rate gyros that were not toonoisy to evaluate exhibited @ response at the time of the disturbance.The effects on the rate gyros at this time are summarized below:

a. Pitch. & slow buildup of 0.7 degree peak-to-peak signal level at §hertz beginning at approximately 120 seconds range time until 133.32seconds, The signal level reached 4.1 degrees peak-to-peak over 0.7seconds tapering off ta less than 0.5 degree by 136 seconds. A 17-21hertz gyro resonance noise was superimposed on the 5 hertz signalsof 0.5 degree peak-to-peak during this period,b. Yaw. Steady-state 0.26 degree peak-to-peak signal level at 5 hertz

before the disturbance becoming negligible afterward. The signal levelreached 1.4 degrees peak-to-peak over 0.4 seconds of the distursance.A 17-21 hertz noise was evidenced but hardly discernible.

¢. Roll. Slow buildup of 2.0 degrees peak-to-peak signal level at 5 hertzfrom 120 seconds until the observed disturbance, The signal levelpeaked at 8.2 degrees peak-to-peak over 0.4 second diminishing to 1.0degree peak-to-peak after 136 seconds, A 17-21 hertz noise wasevident throughout this period of 9.8 degree peak-to-peak.Rate gyro signal indications at 133.3 seconds tend to indicate thedisturbance te be a physical vibration and/or resonance close to the rategyro package as opposed to an electrical transient. The initial disturbance18 seen as a buildup and not an immediate peak as an electrical transientwoutd produce.

DATA

9A.2,5 Structures and Dynamics

Because the IU was not instrumented with strain gauges, a corventtonalstrain analysis cannot be performed; however by deductive reasoning, thestructural integrity of the IU during the 133 second transient can becetermined.

a. Any IU structural collapse would result in 2 catastrophic failure.This has been demonstrated by structural testing.

b. The IU is an entarglement of electrical cables and EnvironmentalControl Subsystem tees) Plumbing. Any structural failure would mostlikely break an electrical cornector or cause a leak in the ECSplumbing. Since no malfunction occurred in the IU during the expectedlifetime, no structural failure occurred,

c. The vehicle trajectory performance was nominal assuring that no abnormalstructural deformation occurred which would influerce the ST-124M-3stabilized platform subsystem,

d. The IU structure has been qualified for vehicle loads which exceededthose experienced on AS-502.

e, Vibrational responses experienced during the 133 second transient wereof shorter duration and no worse than those occurring during the max-‘imum inflight load condition of AS-5OT or AS-502.

In conclusion, no IU structural failure occurred on AS-502 since it wouldhave been detected in the electrical or ECS areas, and no known anomaliesexisted.

9A.3 S-IVB STAGE

The S-I¥B stage experienced an unusual load distribution at 133 secondsas indicated by strain gage measurements at forward skirt stringers.

Thirty two axial strain gages were installed on external hat stringers.The gages were located at vehicle station 3145 of the forward skirt andstation 2821 of the aft skirt as indicated in Figure 94-5. Eight measure~ment locations at each station were approximately equally spaced aroundthe circumference. The recorded data from the strain gages were determined to be valid by the S-1¥B Data Qualification Review Board. Thesedata were reviewed thoroughly and in detail and no electrical anomalieswere detected. In particular, the unusual strain changes which occurredat 133 seconds and which remained through ECO at about 145 seconds werefound to be valid by the Review Board,

The strain histories for the eight side-mounted gages on the aft skirtare presented in Figure 9A-6, At range time zero seconds, all measuredstrains were adjusted to the computed correct strain corresponding to the

9A-15

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STRENGER NO. 1 SECTION A-A

Figure 9A-5. S-IVB Strain Gage Locations

1g axial load condition preceding liftoff. The maximum and minimum de-sign strain envelopes shown in Figure 9A-6 were calculated from designconditions and include the effects of maximum expected aerodynamic gustsand wind shears. It 1s to be noted that the aft skirt strain traces in-cate no unusual strain changes at 133 seconds. It appears that thestructure distributed loads from the forward skirt uniformly to the aftskirt,

The strain histories for the eight side-mounted gages on the forward skirtare presented in Figure 9A-7. the histories for the eight top-mountedgages are shown in Figure 9A-8. A number of strain measurements expert-enced load shifts at 133 seconds, These strain shifts at 123 secondsare summarized in the polar bar chart of Figure 94-9. In the chart, negativevalues indicate increases in compression; positive values indicate de-creases in compression. The shift at stringer 81 was 0.00120 in./in.which corresponds ta an increase of the stringer axial compression loadof approximately 15,124 Newtons (3400 1bf). Stringer 95 experienced an

9A-16

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9A-17

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98-18

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9A-19

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increase of 3781 Newtons (850 Ibf) in compression, and stringer 27 anincrease of 2669 Newtons (600 lbf) in compression. Striagers 27, 54, al,95, and 108 experfenced relatively small changes in strain at the topmounted gages. These gages reflected the conbined effects of axial loads,local stringer bending, and location shifts of the apoTied axial stringerloads. Changes in the structura? character of forward skirt to instrumentunit interface could result from changes in stiffnesses or Toad paths ofstructure above the $-IVB stage forward skirt.

A review of the possible causes for the sudden strain changes (Figure 9A-9)at 133 seconds resulted in the following observations:

98-20

The strain changes were not thermally induced because:

(1) The changes were too rapid to be from thermal effects.

{2}. Ten temperature transducers in the forward skirt did nat revealany sudden or drastic changes in skirt temperatures

The strain changes were not inertially induced since vehicle acceler~ometer recordings were normal at 133 seconds,

The strain changes were not induced by sudden changes in grass massor engine thrust as indicated by the continuity and constant slopeof the plotted curve from accelerometer readings.

The strain changes were not induced by sudden changes in body bendingmoments because:

(1) Airloads were negligible at 133 seconds

(2) Engine gimbaling was negligible at 133 secends

(3) Gody bending would be transient from engine gimbaling; whereas,the strain disturbances were substantially steady state from133 seconds to IECO.

The strain changes were not induced by faulty strain gage systemelectronics because:

(1}, The data were carefully reviewed by electronic specialists withrespect to steady voltages, shorts, gage debonding, wiringidentifications, data transmission, data reductin, etc. Thedata were evaluated as being valid.

{2} The pattern of strain changes at 133 seconds was not the typeexpected of faulty electronics which would produce offscale orzero readings, or result in all similarly wired gages shiftingin the same direction by the same amount.

The S-IVB stage forward skirt did not fail and cause the strain changesobserved at 133 seconds because:

{1) The strain gages continued to respond in normal manner through-out the remainder of powered flight.

(2) The applied flight Toads and temperatures did not exceed designloads and temperatures.

(3) A detailed stress analysis using measured total strains in theforward skirt stringers indicated 2 minimum margin of safety of78 percent.

9A-27

g- The strain changes at 133 seconds in the S-IVB forward skirt stringersappear to be due to a change in load path through the skirt

h. The strain changes at 133 seconds involved eight strain gages; how-ever at 107 seconds, there was an indication of a similar sudden shiftof strain From the single measurement of side gage 66 at stringer 95.This strain shift is shown in Figure 9A-9, The cause af this localizedincident remains to be determined.Causes of these strain changes have been investigated by a Zoint taskgroup at the MSC where all relevant telemetry data from the AS-502 vehicleand spacecraft were reviewed and analyzed. The results of that investiga-tion will be published in a separate report by MSC.QA.4 $-I1 STAGE

The S-II stage vibration and acoustic measurements do not indicate ageneral response to shocks near 133 seconds as has been observed on theInstrument Unit, Only two measurements, E83-219 at vehicle station63.348 meters (2494 in.}, and E81-219 at vehicle station 62.433 meters(2458 in.), on the forward skirt, showed appreciable shocks occurring at133.8 and 134.1 seconds. The energy appeared to be concentrated atfrequencies of approximately 400 and 800 hertz.9A.5 RF SYSTEMS

A transient in RF signal level at. approximately 133.3 seconds was observedat Grand Bahama Islapds on IU VHF telemetry links DF1, DF2, DS},and DP] as shown in Figures 15-1 and 19-2, Section 19.5. No effects werenoted at this time on the Cape Tel 4 (TEL 4) recorded data. The mostTikely cause for these transients was that a piece of debris from the SLA,shown by the ALOTS film, fell past the IU at this time and momentarilyshielded the antennas from the ground receivers. To be effective, thispiece would have to:

a. Be fairly large; at least 0.186 to 0.279 m@ (2 to 3 ft2),

b. Fall fairly close to the antenna,¢. Be of a conductive material.

Figure 94-10 shows how passage of a piece of debris from the spacecraftcoincides with these transients. Times on this figure were derived fromthe ALOTS film.

9.6 EMERGENCY DETECTION SYSTEM (EDS)Although Taunch vehicle EDS indications were normal, there are reports ofanomalies indicated in the spacecraft. Most significant of these spacecraftanomalies was an indication that one of the three abort Tines noted automatic

9A-22

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abort in the EDS logic. Two of the three circuits must give this indicationbefore an automatic abort fs initiated. This single abort indicationaccurred at the time of the 133 second transient and remained throughoutthe duration of the flight. There were other apparently erratic indication:in the spacecraft which occurred subsequent to the transient primarily duringthe period from 133 seconds to S-II ignition. ATl functions in these cablesare not monitored,

Figure 9A-11 is a simplified diagram to illustrate the interrelationshipbetween the IU and spacecraft EDS power and auto abort interfaces.

Power to operate the IU relays is always derived from IU batteries, althoughthe signa? to actuate the relay may come from the spacecraft. The samephilosophy halds for spacecraft networks. As a result, the launch vehicleEDS buses 6091, 6092, and 6093 are fed to the spacecraft. Conversely,spacecraft ES buses No. 1, 2, and 3 are fed to the IU. Since the Taunchvehicle buses are powered by the three operational IU batteries, a shorton an EDS bus would be reflected in the measurements associated with the IUbus which furnishes power to that particular EOS bus, A current spikemeasured in the 6DI0 battery bus which occurred at 133.3 seconds (Figure13-19, Section 13), could be attributed to a momentary short in the 609

gA-23

bus; however, a more likely cause was the additional current drawn bythe platform servo system to stabilize the platform (see paragraph 9A.2.1).Since the 6091 bus power is transmitted to the spacecraft through thesame cable that contains the line which indicated an abort condition,it fs conceivable that the same disturbance which caused an abort indi-cation could have caused a momentary short on the 6091 bus and theresultant transient in the 6010 battery.

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DA-24

SECTION 10

GUIDANCE AND NAVIGATION

10.1 SUMMARY

10.1.1. Flight Program

The performance of the guidance and navigation system was as predictedfrom liftoff to 412.92 seconds. S-I] engine No. 2 cutoff at 412.92seconds was followed by cutoff of S-If engine No. 3 at 414.18 seconds.When engine No, 2 cut off a discrete signal was recognized by the IUindicating an engine failure, However, due to a ground rule thatonly single engine failures be considered, no program action wastaken for engine Nox 3 failure. The change of vehicle accelerationwas detected. With the discrete signal and loss in acceleration theprogram entered a guidance mode where guidance and navigation compu-tations were made based ona thrust change which was 50 percent ofthe total actual change. This mode (artificial tau) Tasted until theTl sensed an acceleration change due to S-II Programned Mixture Ratio{PMR} shift. Guidance computations responded to variations in altitudeand velocity caused by the decrease in thrust during the S-IT burnperiod, The control system responded well to the guidance commandsfor the renainder of the boost period. Due to the two-engine-outperturbation, flight path angle and velocity were not optimum atthe time guidance commanded 5-1¥8 Engine Cutoff (ECO). This resultedin an overspeed of 48.9 m/s (160 ft/s). A parking orbit which wasacceptable though off nominal was achieved.

AN1 orbital guidance maneuvers were satisfactorily performed. IUcommands were properly executed for S-IVB restart, but the engine didnot reignite. Since acceleration test conditions were not met, TimeBose 7 (17) was initiated and a cutoff command was issued to theS-IVB stage.

10.1.2 Instrument Unit Components

The Launch Vehicle Data Adapter (LDA) and the Launch Vehicle DigitalComputer (LYBC) performed as expected for the AS-502 flight. TheST-124N-2 inertial platform and associated equipment performed asdesigned, A transient occurred in dynamics at approximately 133seconds (see Section 9A). Outputs of the servo loops indicated the

10-1

disturbance at this time and that the platform responded properly toall transients. Qutputs of the accelerometer servo loops indicatednominal performance. The accelerometers correctly measured vehicleacceleration throughout the boost phase, Telemetry signals from thegyro servo loops indicated that the inertial reference was maintainedthroughout flight untfl the yaw gimbal reached its limit (460 degrees)at 2,112.4 seconds as a result of the loss of vehicle attitudecontrol at 22,023 seconds. Loss of attitude control was due to fueldepletion in the Auxilfary Propulston System (APS) module at position Iat 21,953 seconds ard in the module at position {I at 22,602 seconds,The module at position III wes not able to offset the LOX ventingforces after 22,023 seconcs.

10.2 GUIDANCE AND NAVIGATION SYSTEM DESCRIPTION

This subsection describes the function of the IU components and thebasic flight program.

16.2.1 Instrument Unit System Description

A block diagram of the Navigation, Guidance, and Control System isshown in Figure 10-1.

The LVOC is a high-reliability general-purpose random-access digitalcomputer which contains the logic circuits, memory, and timing systemrequired to perform arithmetical operations necessary for navigation,guidance, and vehicle flight sequencing. The LYDC is also used forprelaunch and orbital checkout,

The LVDA is. the input/output device for the LYUC. It is designed totransform signals to be compatible with the receiving equipment,perform minor computations, and provide temporary data storage. theLYDA contains the power supplies used by the LVDC,

The ST-124N-3 platform system is a three-gimbal configuration withgas bearing gyros and pendulous integrating gyro accelerometersmounted on the stable element to provide an inertial space-fixedcoordinate reference frame for attitude control and navigation measure~ments (see Figure 10-2}. Vehicle accelerations and rotations aresensed relative to the stable element, Gimbal angles are measuredby resolvers which have both fine and coarse outputs, Inertial velocityis obtained from measurements of the angular rotation of the acceler-ometer measuring head. The data are in the form of encoder outputswhich have redundant. channels.

10.2.2 Flight Program Description

The flight program controls the LYDC from Guidance Reference Release(GRR) or initiation of Tg until the end of tre mission. The program

10-2

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10-3

FLIGHTPATH

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Figure 10-2. Platform Gimbal Configuration

performs seven primary functions: navigation, guidance, event sequencing,attitude control, data management, ground command processing, andhardware evaluation. The program can be describec in twa parts, bocstroutines and orbital routines,

10.2.2.1 Boost Routines, In general, the boost routines perform!navigation and guidance, event sequencing, and attitude control, Boostnavigation encompasses the computations and logic necessary to determinepositon, velocity, and acceleration of the vehicle during poweredflight phases.

10-4

The boost guidance is divided into two distinct modes, pre-IterativeGuidance Mode (pre-IGM) and Iterative Guidance Node (16M). Pre-IGMis used from initiation of Ty to S-IC OECO plus 42 seconds. Programmedcommands include a yaw maneuver for tower clearance, voll to alignthe vehicle to the flight azimuth, and tilt comands, The yaw maneuveris initiated at IU umbilical disconnect sensed by LYDC (T]) plus1,0 second, Rott and tilt are initiated when the vertical componentof space-fixed position changes by 137 meters (449.5 ft) approximately11 seconds after T], A time backup of T] plus 13 seconds is providedin case of an accelerometer failure, After the tover is cleared andthe vehicle fin-I/fin II axis is aligned to the flight azimuth, theroll and yaw commands are set to zero for the remainder of pre-IGM.Tilt comands are computed from one of four third-degree time-tiitpolynomials, Nominaily, tilt arrest tfme is T) plus a preset timeand guidance commands are frozen until Tnttiation of 16M

The IGM guidance scheme is a modification of the multi-stage, three-dimensional form of IGM, IGN is a near optimal scheme based on a flatearth optimum steering function for planar motion of a point-massvehicle, The approximate thrust vector steering function is imple-mented in both the pitch and yaw planes. IGM is implemented in twoFlight modes, boost-to-parking-orbit IGM and out-of-orbit 16MHowever, only the boost-to-parking-orbit mode is applicable to AS-502,since the S-IVB stage engine did nat reignite, IGM is initiated atT3 (S-IC OECO} plus 42 seconds. Based on the state vector at ‘nitia-tion of ICM, guidance conmands are computed and implemented to steerthe vehicle to preset termnal conditions. Alternate logic andbackup procedures are provided for thrust level shifts and vehiclestaging. These procedures are discussed in more detail in paragraph10.4.2 along with applicable portions of the Flight Program Evaluation,

The Steering Misalignment Correction (SMC) compensates for the misalign-ment of the thrust vector and for the time lags in output of the steeringcommands, The SMC terms are only used during active IGM guidance.SMC is not used for any computation cycle in which the reasonablenesstest on both velocity words faits, turning rates are limited, or whenboth crossover detectors for the gimbal angle readings are determinedto be bad.

Event sequencing is accompl {shed by the switch selector routine on astored table basis. The routine determines if it is time to issue aswitch selector command, verifies that no switch selector stage washung, verifies that the correct address is sent to the stage switchselactor, and issues the read comands,

Attitude control is accomplished in the minor loop support section ofthe major loop on an interrupt basis. The minor loop support routineincludes calculations of such parameters as steering rates to be appliedin the minor loop. Limiting of the ladder outputs is accomplishedwhen necessary, and backup and failure paths are provided in caseginbal angle discrepancies occur,

10-5

10,2.2.2 Orbital Routines. The orbital program consists of two inter-ruptable monitor Foutines, The first is the Instrument Unit HardwareEvaluation Program (HEP), and the second is the Telenetry ExecutiveProgram (TEP). Navigation, guidance, event sequencing, attitudecontrol, and ground conmand processing are initiated on an interruptbasis from etther HEP or TEP, During orbital flight and when thevehicle is not aver a ground station, the HEP routine is exercised.That is, the computer will be engaged in addressing the ComputerInterface Unit (CIU), compressing CIU and LYDC data, and executingcomputer self-test. Once the vehicle acquires a ground station, TEPis entered as the program major loop. This routine orovides timesharing telemetry of compressed and real time data. In addition,command system data and various special data are telemetered on aninterrupt basts. Data from the LVDA are telemetered automatically.The orbital guidance routine controls the computation of the commandedvehicle attitudes, This routine is initiated at Ts (LVDC sensed firstS-IVB engine cutoff) plus 15 seconds for the parking orbit and reenteredat Ty for the waiting orbit. Orbital navigation encompasses thecomputations necessary to determine position, velocity, andacceleration in the space-fixed coordinate system during earthorbit. These computations are carried out in an indirect fashion,making use of mathematical models of the earth, its atmosphere, andthe vehicle. A routine for switching the C-band antenna as a functionof position is included. This routine is also entered upon exit fromthe minor Toop at B-second intervals,

Event sequencing in orbit is accomplished in exactly the same manneras in the boost phase but with the added capability to receive timebase updates and special outsut sequence commands from ground stations.

Attitude control for orbital operation is accomplished in the samemanner as in the boost phase with the exception of the rate of entryinto the minor loop. The orbital minor loop #s entered 10 times persecond, The First and fifth pass are the attitude update passes(cycled through twice per second), and the remaining eight passes arefor attitude hold (cycled through eight times per second) to minimizedrift problems.

Ground command processing is accomplished by the Command Rece‘verinterrupt with the Digital Command Subsystem {DCS} routine, The DCSroutine processes ail ground commands, provides data and mode veri fi-cation, and supplies the necessary information to the various affectedroutines, Flight program differences between the AS-501 and AS-502flights are described below:

a, One M/F smoothing filter was used for all stages on AS-502.

b. Anew curve fit for the orb{tal vent model was incorporated.

10-6

c. New orbital guidance maneuvers were planned for Tg.

d. Tp was initiated by the inboard engine cutoff interrupt instead ofa fixed time in Ty.

@. Five additional CCS commands were provided.

f. The Apallo Standaré Coarcinate System was implemented.

g. The second S-IVE cutoff parameters were representative of translunartrajectories.

h. ho artificial tau steering mode was used for the second S-1¥2 burnpertod.

The roll ladder limit was 7 degrees from Ty +0 to Ty +600 seconds,and 15.3 degrees fer all other flight times. The pitch and yawladcer limits were increased from 2,5 te 7 degrees from T7 +15 toT7 +600 seconds,

10.3 GUIDANCE INTELLIGERCE ERRCRS

The postflight guidance hardware error analysis is based on comparisonsof the ST-124N-3 platform measured velocities with the postflighttrajectory established by tracking. Figure 10-3 presents comparisonsof the platform measured velocities with correspondirg values fromtre final postflight trajectory (ONPT). Comparisons were made hysubtracting guidance values from trajectory values, The differencesshown for the pitch plane, (range and altitude) are well within theaccuracy of the data compared. The range velocity difference waswithin + 0.1 més (0.3 ft/s) for the SIC and S-IT flight pertod,At S-IVB velocity cutoff comand the difference was -0.3 m/s (-0.98 ft/s).The altitude velocity difference increased to a maximum of 0,45 m/s(1.49 ft/s} at about 280 seconds and changed slope. At S-IYB cutoffthe difference was -0.7 m/s (-2.3 ft/s).

The crossrange velocity difference was well within the accuracy ofthe data and 3 sigma hardware errors, At S-IVB velocity cutoff commandthe crossrange velocity difference was -1.85 m/s (-6.07 ft/s). Severalattempts have been made to establish errar terms that would producethe velocity differences shown, The curves have been simulated towithin 20.1 m/s {0.3 ft/s} for all three components. A more completeerror analysis w'll be performed and published as a classified memoran-dum for linited distribution, An estimate of acceleration bias asso-ciated with the guidance accelerometers was made using plots of thetelemetered velocity outputs during orbit. Although venting wasessentially continuous, the general slope of the oscillating curvegives a reasonable estimate of the acceleration bias, Bias errorsdetermined by this method and by calculations in the postflight OrbitalCorrection Program {OCP} indicate error magnitudes within 41,0 x 10-4 m/s2{3.3 x 10-4 Ft/s2) of the preflight measurements for the platform S/N 13,

10-7

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10.4 NAVIGATION AND GUIDANCE SCHEME EVALUATION16.4.7 Inertial Platform and Navigation Parameter Comparisons

ST-124N-3 platform measured velocities and LVDC velocities at severalflight event times are shown in Table 10-1, along with correspondingvalles computed from the postflight trajectory data, Trajectory datawere smoothed through the transient areas to be compatible with thevelocity differences shown in Figure 10-3. No discrepancy was notedbetween data telenatered from the accelerometer pickoffs and theaccumulated velocities from the LYDC,

Table 10-1. Inertial Platform Velocity Comparisons

TELEMETERED GUTDARCE POSTFLIGHTEVENTS VELOCITY*

}

ACCELEROHE: ER COMPUTER TRAJECTORYms (ft/s) ms {4/8} nfs (ft/s)

SIC x 2482.00 2482.00 2482.35BECO : (8143.04) (8343.04) (8144.19)Sensed by LVOC y 712.95 12.95 “13.12148.41 sec : (242.49) (42.49) (-43.04)

i 2252.75 2232.75 2292.68(7328.30) (7325.30) (7328.07)

Sell Cutoff j 3627.50 3627.50 3627.29Sensed by L¥OC, (41901 .25) (11901 .25) (11900.56)576.33 sec -10.80 1-19.80 12,25

(-35,43) (-35.43) (-40.19)i 6648.95 6648.95 | 6648.87

(27814.14) {21814.14) (21813.88)

5-1V8 Velocity k 2144.85, 3144.85 3144.18Cutoff Comand . (10317-75) {10317.75) (10315.45)747,04 sec ¥ 1.78 1.75

“ (74) (5.74)i 7659.05, 7659.05

{25128.12) (25178112) (25177. 70}

Parking Orbit 2145.75 3145.76 344.97Insertion, (10320. 70) : (10318.14)757.04 see 1.75 0.27

. (5.78) (-0.89)i 7661.45, : 7661.17

(25135.99) (25135.99} (25135.07} *These coordinates are as defined in the Apollo 13 coordinate system:

R= Altitude velocityCrossrange velocityDownrange velocity

10-9

Navigational positions and velocities determined from the preflighttrajectory, postflight trajectory, and telemetered LVDC data are shownin Table 10-2. At S-IC/S-I1 separation, the guidance data were in verygood agreement with the postflight trajectory values. The differencesBetween the actual and preflight data reflect nonstandard flight condi-tions and S-IC performance. An apparent yaw bias that built up duringS-IC burn contributed significantly to the 21.94 r/s difference incrossrange velocity at $-IC {LVDC sensed) 0ECO. The vehicle driftfn yaw resulted froma small thrust misavignment and slight off-nominalengine verformance, It should be noted that pre-IGM guidance doesnot provide path guidance and that the relatively large crossrangevelocity difference posed no problem for 16H guidance.

There was good agreement between quidance and postflight trajectory atS-I1/S-IVB separation, S-IVB cutoff, and parking orbit insertion.However, due to the perturbation caused by premature cutoff of S-IIengines No. 2 and 3, the agreement between preflight trajectory positionsand velocittes and either guidance or postflight trajectory values isnot as good. Measured velocity gain due to thrus+ decay after first cut~off of the S-IVB stage engine was 2.50 m/s (8.20 ft/s) compared toa predicted value of 2.26 m/s (7.41 ft/s) (see Section 7.4}. Thevelocity outputs of the guidance accelerometers from orbital insertionto 1g were curve-fitted with tine polynomials. The velocity polynomialswere differentiated and then evaluated to determine measured accelera~tions, A Root-Sum-Square (RSS) of the acceleration components is shownas a solid line in Figure 10-4 along with the predicted acceleration(dashed line), taken from the final operational trajectory (preflightpredicted), and the programmed vent model. The guidance accelerationcomponents were also adjusted for estimated bias, and the circledpoints represent adjusted accelerations. The thrust produced by theContinuous Vent System (CYS) based on venting parameters is shownin Figure 7-7 (Section 7.5).

Oscillations of the measured accelerations appeared to be out of phasewith the predicted values, This Was probably due to some combinationof inaccuracies in the predicted values, curve fits of measured data,off-nominal propellants onboard at insertion, and off-nominal orbit,Further investigations are being made to varify the oredicted accelera-tions, The measured accelerations, with minor adjustments for bias,were used in establishing the postflight orbital trajectory (seeFigure 4-6 and paragraph 4c3.2)) Figure 408 1 sintlar to Figure 10-4jn its comparisons of accelerations in parking orbit, except thatFigure 4-6 includes an adjustment for predicted drag,

10.4.2 Flight Program Evaluation

The flight program performance was as expected for the flight pertur-bations experienced, The navigation scheme functioned correctly in allphases of flight, control calculations were corract, and orbital operation

10-10

10-17

Tabl

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Guidance

Comparisons

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RANGE TIME, HOURS: MINUTES: SECONOSFigure 10-4. Predicted and Measured Accelerations in Parking Orbit

was as expected, Guidance schemes functioned as designed. Investtga-tions have demonstrated that an overspeed of 48.9 m/s (160 ft/s) wasa direct result of premature shutdown of the to S-II stage engines.

When the premature $-IT outboard engine shutdown was detected by theFlight program, bit 15 was set in mode code word 25 (MC25). This bitwas improperly identified as inboard engine out in the program docu-mentation. Since this bit is not used in any of the program logicflow, the improper description had no effect on the operation of theflight program; only an erroneous indication in real time telemetryresulted, This problem has been corrected for future Flight prograns.

Pre-IGM guidance was nominal. The yaw maneuver for tower clearance wasinitiated at 1.9 seconds and properly executed before pitch and rollcomands were initiated at 11.1 seconds. The yehicle was rolled from its90 degree launch azimuth to the 72 degree flight azimuth by 31.3 seconds.Upon completion of the yaw and roll maneuvers these guidance commandswere set to zero for the remainder of pre-IGM guidance made. No backupnodes were required for the S-IC stage, and tilt arrest occurred at 140.9seconds. From this time the pitch cenmand (Chi ¥) was frozen until IGMinitiation at 191.0 seconds.

10-12

A roll bias of -0.5 degree was observed from approximately 30 seconds toInboard Engine Cutoff {IECO}. A negative yew bias was noved at approximately 15 seconds and slowly reduced to zero by 70 seconds, then graduallyincreased from 88 seconds to IECO. The yaw and rol? biases were probablythe result of allowable thrust deviations during S-IC burn. The pre-IGMguidance does not provide path guidance corrections and the vehicle yawdrift was the major contributor to the 21.94 m/s (71.98 ft/s) off-nominalcrossrange velocity at S-IC (L¥DC sensed) OECO.

The performance of closed loop guidance (16M) was nominal until 412.92seconds, when the S-IT stage engine No. 2 shut down, followed by cutoftof engine No, 3 at 414,18 seconds. The flight program detected an S-Ioutboard engine failure based on both an external discrete and an ace?eration test. Although the engine faitures occurred in the sane computa:tion cycte that the PMR shift was expected, the program properly identifiedthe change in acceleration as an engine out. At this time two modifications were made to the nominal S-II guidance equations

a. An artificial tau made was entered in which preset (versus measured)acceleration values were used as guidance inputs. Tau is a calculatedIGM parameter representing time-to-go required to burn al) remainingvehicle mass at a constant mass flow rate. Tau is the product ofaverage exhaust velocity (preset value) and the reciprocal of totalacceleration,

b. All parameters based on time of S-I1 fuel depletion were adjusted toreflect the longer burn time expected on four engines. This modi fica-tion resulted in a change of the time-to-go to S-II stage cutofffron 101.0 seconds to 126.3 seconds.

The second engine failure that occurred was not acted upon because a basicground rule in designing flight program backup modes for hardware failureswas to protect only against single engine failures. No logic was pro-vided to check for multiple engine failures through acceleration changesThus the guidance parameter adjustments which revised time-to-go forS-II stage cutoff underestimated the time by approxinately 45 secondsVehicle guidance remained in the artificial tau mode until ProgrammedMixture Ratio (PMR) shift occurred at 490,8 seconds, At PUR shift, a secondartificial tau mode was entered to smooth the transition to the lowerthrust evel. At termination of the second artificial tau mode (510.2seconds), due to quidance sensed PMR shift, the measured accelerationswere lower than the artificial values used. Guidance commanded nose-upattitude in response to the lower acceleration. Seven and one-halfseconds later, a chi freeze was properly initiated in preparation forS-II/S-1VB staging.

10-13

Because of contro? considerations at S-1I/S-IV3 staging, the flight pro-gram normally freezes the guidance commands 5 seconds prior to S-II stagecutoff. For AS-502 the freeze period was entered as specified, but theactual remaining burn time exceeded the time-to-go, which resulted in a88 second period of constant attitude hold rather than the nominal 5seconds. This constant vehicle attitude led to a buildup in radial veloc-ity which, in turn, caused an altitude overshoot,

The perturbation duc to loss of two SII engines had two effects on thevehicle performance: the reduced acceleration resulted in lower thannominal velocities, and the lengthened attitude freeze led to higher thannominal altitude. Thus at 28,2 seconds after initiation of guidancein the S-I¥B stage, velocity was 172.1 m/s (564.6 ft/s) lower thannominal and altitude was 9,1 kilometers (29,850 ft) higher than nominal.

The IGM flight program normally steers the S-IVB stage to preset endconditions of radius and vector velocity optimally. Steering angles toachieve the vector velocity end conditions are added to steering anglesto satisfy the radius constraint in order to form the total commandedyehicle attitude. For AS-502, the highly perturbed state vector whichresulted from the two S-II stage engine failures led to inconsistentsteering angles to achieve velocity and radius end conditions. Thecomposite of the two pitch steering angles is shown in Figure 10-5. Thistotal commanded vehicle pitch attitude is not rate limited. The corra-sponding rate limited conmands are presented later, As can be seen fromFigure 10-5, the commanded attitude on AS-502 was a pitch down to reducethe radius. The radius was corrected, but a resultant negative radia?velocity was achieved. A nose-up pitch command was given to compensatefor the excess radial velocity. By 712.3 seconds, when terminal guidancewode was entered, pitch attitude had changed 46 degrees nose-up to 13degrees above the local horizontal. The sensitivity of IGM to changesin acceleration increases as the desired terminal parameters are approached.A terminal guidance schewre (chi bar steering) is required which uses onlythe velocity constraints. Curing chi bar steering, altitude constraintsare set to zero,

With IGN calculations now constrained only to component velocity in theterminal guidance mode, 3 continued nose-up pitch was commanded to achievethe desired radial velocity. Chi commands increased +96 degrees untilS-IVB stage cutoff. Vehicle pitch attitude was rate limited through the35-second duration of terminal guidance. S-I¥B stage cutoff occurredwith pitch attitude approximately 49 degrees above the local horizontal.S-IVB stage cutoff occurred 88.0] seconds later than the nominal predictedtime and with total velocity 48.9 m/s (160 ft/s) greater than desiredterminal velocity.

10-714

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TIME FROM GRR, SECONDSpo

250 300, 350 490, 450 500 560, 600 680 700 750,

AANGE TIME, SECONDSFigure 10-8, Composite Pitch Steering Angle (Not Rate Limited)

Referenced to Local Horizontal

Figure 10-6 is a plot of the radial velocity (aky) te-be-gained versusthe horizontal velocity (sZy) to-be-gained. The circle about theorigin represents the cutoff loop which is entered when total veloctty-to-be-gained falls below approximately 65 m/s (213 ft/s). Entry intothe cutoff loop {s made only from calculations of component velocity-to-be-gained. Once in the loop, cutoff 1s effected as a function ofthe magnitude of total velocity, The difference in the actual andnominat velocities-to-be-gained is a result of pitch down attitudecommands to correct altitude during the initial portion of S-IV8 guidance.The pitch down attitude tended to increase the radial velocity-to-be-gained rather than decrease it. This increase led to an inconsistencyin the components of velocity-to-be-gained, and, as a result, bothcould not be driven to zero simultaneously, The radical attitude changesnear S-IVB cutoff finally reduced the velocities-to-be-gained sufficientlyto cause entry into the cutoff loop.

10-15

INCREMENT OF HORIZONTAL590 1000 3500 2000 2800woo 7

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=100

200 ‘“a0 € 00 Due 300 a a) 70aINCREMENT OF HORIZONTAL VELOCITY, nis

Figure 10-6. Actual and Nominal Velocities-to-be-Gained

On entering the cutoff loop, the AS-502 total velocity was in excess ofthe desired value of 7790.9 m/s (25,560.7 ft/s), and cutoff signal wasimmediately given at 747.94 seconds with an overspeed of 43.9 m/s (160ft/s). This overspeed did not represent an error in the guidance programoperation but merely reflected the radial velocity-to-be-gained constraintwhich had to be satisfied before entry into the high speed cutoff loop,AN] commanded attitude outputs were rated limited to 1.0 deg/s. Thislimit serves to make the vehicle less responsive to program attitudecommands. Table 10-3 gives the times the commands were rate limited.Note that they were limited for about two-thirds of the S-IVB activeguidance.

Table 10-4 summarizes the start and stop tines of the modes used to com-pute guidance commands. Note that from the $-[I-engines-out until $-1VBCutoff, the guidance was either in an artificial tau mode, rate limited, orfrozen for all but 62 seconds.

Commanded attitude angles (rate limited) during the boost phase of flightare summarized in Figure 10-7, Both actual and commanded attitude (ratelimited) angles during S-IC burn, S-IT burn, and S-IVB first burr areshown on expanded scales in Section 11.

19-16

Table 10-3. Rate Limited Steering Conmand Times

START TIME STOP TIMESEC SEC415.4 419.1

510.2 514.0

582.9 644.1

699.0 747.0

Table 10-4. Start and Stop Guidance Commands

EVENT TeH PHASES ARTIFICIAL TAU

|

CHI BAR CHI FREEZE(SEC) (SEC) STEERING (SEC)

(SEC)START STOP [START STOP

|

START STOP

|

START STOP

First 191.0 418.4 1415.4 490.8

|

=== wee Poem o16H

second

|

490.8 517.7 [490.8 510.2

|

wr- vrs

|

517.7 582.9IoM

Third se4.@ 747.0 [564.8 594.3 1712.3 746.41 746.4 762.316H

Fourth 11628.4 11630.31GH

IU commands were properly executed for S-IVB restart but tne engine didnot reignite. The Flight program immediately initiated T7 through negativeresults from the acceleration test as programmed. The fifst event scheduledin Ty was a cutoff conmand to the S-IVB stages this was executed. A stmu-Jation using actual AS-502 flight data and a nominal S-IVB second burnthrust revealed that an acceptable waiting orbit could have been achievedfrom the perturbed parking orbit had the $-I¥B engine reignited.

10.4.3 Orbital Guidance

Al] orbital guidance and sequencing functions were performed correctly.The vehicle response to the guidance commands in Tz was sluggish because ofthe inertial characteristics of the vehicle resulting from no second burnof the S-IVB stage and early separation of the spacecraft on ground conmand.The predicted attitude timeline for the AS-502 flight is shown in Figure10-8, Comparison of parking orbit attitude commands shown in Section 11with the attitude timeline yields good agreement.

10-17

10-18

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dddddqdad

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PARKING ORBIT INSERTION, 669 (PREDICTED)INITIATE 180° ROLL TO PLACE POSITION IIT DOWN (7; + 90), 749 (PREDICTED)

INITIATE 20° PITCH DOWN MANEUVER (T, + 2460), 3119 (PREDICTED)INLTIATE 20° PITCH UP MANEUVER (T, + 4680), 5339 { PREDICTED)INITIATE 80° ROLL TO PLACE POSITION I DOWN (T, + 5040), 8699 (PREDICTED)INITIATE RESTART SEQUENCE AND START OF Tg, 11,076 (PREDICTED)INITIATE MANEUVER TO SEPARATION ATTITUDE (T, + 20), 11,478 {PREDICTED}LY-LTA/CSM PHYSICAL SEPARATION, 11,908 (PREDICTED)INITIATE MANEUYER TO POST SEPARATION INERTIAL ATTITUDE (1, + 600), 12,328{PREOTCTED)

240

“TOTTI

TAINahsal \-180

-240® 2000 4000 6000 sooo 1000s 1zpoo.(14p00

RANGE TIME, SECONDS VV

00 00 00RANGE TIME, HOURS :MINUTES : SECONDS

Figure 10-8, Attitude Timeline - Liftoff to SpacecraftSeparation

10-19

Vehicle control was exercised until 22,112.4 seconds, at which timeattitude reference was lost. Partial control was Tost at 21,953 secondsdue to loss of fuel tn the APS module at position I. The opening ofthe LOX vent valve at 22,023 seconds produced a thrust which overcamethe remaining thrust capability of the APS module at position III.Loss of fuel in the APS module at position III occurred at 22,602 seconds.

The minor Toop performed as expected during flight. The rate at whichthe yaw gimbal changed when the APS ran out of fuel was insufficient tocause the program to switch to backups or set the guidance fatiurediscrete.

10.4.4 Orbital Routines

The TEP routines performed as expected. This was proven by the realtime and compressed data telemetry which has been reduced. No problemswere found with the LYDC self-test or in CIU processing. All DCScomands received by the flight program were processed correctly.10.4.5 Event Sequencing

AVL program sequencing was proper. All switch selectors were issued intolerance. The Environmental Control System logic monitored the thermatswitches and opened and closed the control valve properly.

10.5 GUIDANCE SYSTEM COMPONENT EVALUATION

10.8.1 LVDC Performance

Data analysis indicated the LVOC perforned as predicted for the AS-502mission. No valid error monitor words and no self-test error data havebeen observed that indicate any deviation from correct operation.

10.5.2 LYDA Performance

The LVDA perfarned satisfactorily for the AS-502 mission. With theexception of the error monitor word related to the interrupt processorno deviations from correct operation have been found. Two error monitorwores were observed which indicated apparent disagreements in theTriple Modular Redundant (THR) Orbital Check Ready (OCR) latch. Onedisagreement occurred within 0 to 30 seconds prior to 8618.2 secondsand the other at approximately 9218 seconds. For each OCR latchdisagreement a TMR Interrupt Control (INTC) error monitor word indi-cation might also be expected. However, the flight program inhibitsthe OCR interrupt and processes it on a cyclic basis. Since the OCRinterrupt is inhibited, no disagreement between logic channels canbe sent to the INTC latch. The apparent dfsagreement in the OCR latchwas attributed to a difference between rise delay times for the THRinterrupt input logic channels, Two error monitor words were observed

10-20

in compressed data which indicated apparent disacreerents in the INTClatch. Thase disagreements occurred within 0 to 30 seconds prior to14,984.7 and 15,044.7 seconds.

The real time and interrupt countdown processers performed satisfantort ly.

10.5.3 Ladder Outputs

The ladder networks and converter amplifiers performed satisfactorily.No data have been observed that indicate an out-of-tolerance conditionbetween channel A and the reference channel converter-ampli fiers.

10.5.4 Telemetry Outputs

Analysis of the available analog telemetry buffer and flight controlcomputer attitude error plots indicated symmetry between the bufferoutputs and the ladder outputs. The analysis of the available L¥DCpower supply plots indicated satisfactory performance of the powersupply telemetry buffers.

30.8.5 Discrete Outputs

No valid discrete output register words (TAGS 043 and 052) were observedto indicate guidance or simultaneous memory failure.

10.5.6 Switch Selector Functions

Switch selector data indicate that the LVDA switch selector functionswere performed satisfactorily. No error monitor words were observedthat indicate disagreement in the TMR switch selector register positionsor in the switch selector feedback circuits. No wode code 24 words orswitch selector feedback words were observed that indicated a switchselector feedback was in error. In additton, no indications wereobserved to suggest that the B channel input gates to the switch selectorregister positions were selected.

10.5.7 ST-124M-3 Inertial Ptatform Performance

The inertial platform system performed as expected with the only davi~ation being in the gasbearing supply system. This deviation isdescribed in Section 18.5,

The XK, ¥, and Z accelerometer servo loops maintained the accelerometerfloat within the measuring head stops (#6 dagrees) throughout liftoffand MAX Q (see Figure 10-9), The accelerometer encoder outputs indicatedthat the accelerometers measured the vehicle acceleration properly.

The X, Y, and Z gyro servo loops for the stable element functioned asdesigned. At no time during liftoff and MAX Q were the operational

limits reached, There were several perturbations observed in the servo

10-21

10-22

“ANAT.

oo +

3LaWOYTT 99¥ 7 -

Ea

F

29 +

ANH TdYaLIWOVT] a0" x

=|

te

9dnydIdF

‘USL INOUTTAD

Figure

10-9.

24

56

7RANGE

TIME,

SECONDS

Accelerometer

Pickup

Signals

During

Liftoff

gyro Toops which cannot be fully explained at this times however, thegervo loops responded to these perturbations in the proper manner andmaintained the stable reference. Periodic ievel changes in the servosignals are attributed to servo loop deadband lim't cycling associatedwith vehicle attitude changes.

The fine backup ginbal angle resolver output indicated nominalperformance -

There were no vibration-induced malfunctions of the ST-124M-3 InertialPlatform on AS-BO2, Tre effects of vibration on the gyro and accelerometerservo output signals are illustrated by the graphs of maximum vol tagevariations shown in Figure 10-19, These voltages were sufficient tomaintain inertial reference and prevent a vibration-induced malfunctionof any accelerometer or gyro. This figure also ilTustrates the largevibration experienced at 133.3 seconds.

Avatlable data indicate that the ST-124N-3 composite vibration levelsat liftoff on AS-502 were very near those of AS-501 on the inertia

gimbal, However, the structural vibration was lower on AS-502. The

Yibrations at Mach 1 and MAX Q were higher on AS-S02. The compositelevels of the inertial gimbal vibrations are shown in Figure 9-34 inSection 9. The vibration experienced at 133.3 seconds fs not showndue to its short duration.

10-23

2 ALL {..2* aL¢ ol Me aa2 6 T

ee LIZ IN .E bh= 1Be Uh Is |ACU

aL NYle

2 Boos {oe |g o heepo}

e ies2 1.0

* Bos 7A t

20 40 60 80 100 120 140 160 180RANGE TIME, SECONDS

'V TRANSIENT AT 133.3 SECONDSFigure 10-10, Envelope of Naximun Deviations of the Gyro and

Accelerometer Servo Amplifier Outputs

10-24

SECTION 11

CONTROL SYSTEM

11,1 SUMMARY

The AS-502 Flight Control Computer (FCC), Thrust Vector Control (TYC}, andAuxiliary Propulston Systen (APS) satisfied all requirements for attitudecontrol and stability of bending and propellant slosh modes in both theboost and orbital coast modes of operation. During liftoff, all vehicleclearance requirements were met satisfactorily with less than 20 percent.of the clearance margins being required. The programmed 1.25 degrees yawmaneuver to provide adequate tower clearance and the 18 degrees rollmaneuver to the 72 degree flight azimuth were satisfactorily initiatedand executed.

The wind-biased pitch tilt program was satisfactorily initiated andexecuted. The control system was required to correct for a steady-stateroll attitude error of anproximately -0.5 degree through first stage boost.This roll torque was not observed on AS-501, as the attitude error wasessentially null after about 60 seconds.

S-IC/S-IL separation was satisfactorily accomplished, as was second planeseoaration. Control system performance was consistent with events whichoccurred during S-II boost. The performance shift of engine No. 2 at319 seconds was evidenced in the TYC as well as in the FCC parameters.However, this varformance shift caused no control problens and resultedonly in a new steady-state trim condition.

The Steering Misalignment Correction (SMC) was initiated at 212.0 seconds.The performance shift at 319 seconds had a negligible effect on SMC, andthe maximum SHC's in pitch and yaw prior to engines No. 2 and 3 shutdownwere -0.7 and 0.2 degree in pitch and yaw, respectively. The FCC and TYCresponded satisfactorily to the perturbations caused by the shutdown ofengines No. 2 and 3 at 412.92 and 414.18 seconds, respectively. Thisshutdown resulted primarily in a large pitch plane disturbance duringwhich the pitch rate built up to a maximum of 2.8 deg/s (nose-up) and thepitch attitude error reached a maximum of 13.4 degrees. A maximum enginedeflection of 5.95 degrees was required to stabilize the attitudeexcursions. The maximum SHC's required at this time to compensate forthe thrust vector misalignment were -13.2 degrees in pitch and -0.6 degreein yaw. The nose-up trim condition resulted in a 7.4 degree pitchattitude error at separation.

At S-IT/S-IVB separation, the guidance coxputer switched to the S-1VBcoast mode for 0.3 second. The 7.4 degrees pitch attitude error causeda@ full-on APS pitch-engine firing of 0.3 second duration to correct theattitude. At 0.3 second after separation, the guidance computer switchedto the S-IVB burn mode. The pitch attitude error was trinmed out by theTW after S-IVB stage J-2 engine ignition. Contral system performarcewas nominat for the reriainder of S-IVE first burn.

Orbital attitude control requirements required considerably more APSactivity than anticipated. The APS system was required to overcome a50 degrees nose-up fron loca: horizontal attitude and a 1 ceg/s rose-uangular rate to align the vehicle along the local horizontal. ire vehiclewas subsequently exercised through a sequence of four maneuvers asfollows: 180 degrees roll, 20 degrees pitch down, 20 degrees pitch up,and 180 degrees roll. The pitch and roll maneuvers were planned toproduce information on the S-IV6 restart bottle repressurization andpropellant stosh excitation vnile qualifying these maneuvers formanned flight. Each maneuver was executed as planned. LHsloshingwas not appreciable during any of the maneuvers. Significant LOXsloshing existed at the initiation of each pitch maneuver; however,the initial amplitude was not sustained due to high damping,

An auxiliary hydraulic pump failure prevented the $-IV stage J-2 enginefrom being centered at the time of S-IVB Engine Start Command (ESC). Theengine position at ESC was approximately 1.8 degrees in pitch and -2.3degrees in yaw. Appreciable attitude errors resulted from this engineposition during restart attempt; however, vehicle control was maintainedby the APS systen following the switch from thrust vector to coast modecontrol,

Subsequent to spacecraft separation (Section 12) the APS system waintainedcontrol until module © fuel depletion at approximately 271,553 secondsVehicle attitude rates began to build up significantly feltawing rodule 111fuel depletion (22,602 seconds) and continued to increase as indicated hyreduced radar data unt'? a tumble rate of 180 deg/s was recarded by theninth day following launch.

11,2 CONTROL SYSTEM DESCRIPTION

Figure 10-1 (Section 10) shows the interconnection ard signal flaw pathsof the control components as they relate to the guidance ccmpcnents.

Vehicle attitude correction is accomplished in accorcance with the require-ments of the guidance system through attitude error signals. These signalsare generated by the Launch Vehicle Digital Computer (LVOC) and LaunchVehicle Data Adapter (LVDA). During S-IC stage burn, attitude steeringcormands are the result of the preprogrammed yaw and roll maneuvers andthe time tilt pitch program. At the initiation of Iterative GuidanceMode (16M), attitude steering commands become the result of guidancesystem computations,

Angular vate ‘nputs are present when the control system has responded toattitude error commands or forces acting on the vehicle. Commanded vehicleattitude changes during powered flight are limited to rates of 1 deg/s orJess, depending upon requirements of the guidance system.

Control system outputs are valve currents (Iy) ta first, second, and thirdstage engine actuators and relay currents to the APS.

The vehicle engine, actuator, and nozzle arrangements and axts definitionsare shown in Figure 11-1.

The AS-502 Flight Control Computer, which is essentially identical to theAS-501 FCC, is an analog computer which generates the proner conmands forthe S-I¢, S-IT, and S~1VB stage engine actuators and S-IVB stage APS. Ingenerating the engine commands, the FCC processes and combines attitudeerror signals from the LYDA and angular velocity sfgnals from the Contro}-EDS Rate Gyros/Control Signal Processor {CSP}. Ta inaut channels acceptthe contro? signals. ach channel amplifies, scales, and filters it:respective signal according to a predetermined time-variable set of gaifactors.

The FCC also provides $-IVB stage attitude control commands to the APSThis control is provided to the roll axis during 5-1V3 stage burn and toall three control axes during coast

The Control-EDS Rate Gyros/CSP used on AS~502 were essentially identicalto those used on AS-501. The Control-E0S Rate Gyras/CSP combinationprovides angular velocity signals to the FCC for dynamic feedback. TheControl-EDS Rate Gyros contain nine rate gyros, three in each axis.

11.3. S-IC CONTROL SYSTEM EVALUATION

The AS-502 control system performed salisfactorily during S-IC poweredflight with mast parameters near oredicted. The 1.25 degree yaw biatower clearance maneuver was executed as planned and resulted in adequatetower clearance.

Vehicle liftoff acceleration was greater than that of AS-501, Simulationwith measured stow release forces and theust verified this result. Ac-caleration was greater than AS-501 because 12 lubricated slow release rodswere used instead of 16 non-lubricated rods. Less than 20 percent of theavailable clearances were used during liftoff.

The vehicle performed within flight dynamic constraints throughout flightIn the region of high dynamic pressure, pitch angle-of-attack peaked at3.1 degrees, and pitch engine deflection peaked at 0.47 degree. Absenceof any divergent bending or sloshing frequencies in vehicle motion indicatesthat the bending and slosh dynamics were adequately stabilized, Thecontrol system sufficiently rejected two prominent disturbances near theend of S-IC flight: the 5.3 hertz longitudinal oscil¥ation that coupledinto pitch (see Section 9.2.3.1), and the transient at approximately

1-3

”ROLL

ORX

AXIS

39cagB82

“PITCH

ORY

AXIS

S-IC

_STAGE

ur

Engines,

Actuators

and

Nozzle

Arrangement

Figure

11-1.

133 seconds (see Section 94). Vehicle dynamics pricr to S-IC/S-II firstplane separation were well within reouired levels.

11.3.1 Liftoff Clearances

Positive clearance existed between the vehicle and the mobile launcherstructure, as shows in Table 11-1. Vehicle clearances were at least80 percent of those available. Positive clearances resulted fromafavorable combination of vehicle-system misalignments. The ground winddirection was 132 degrees east cf north, and the magnitude was approxi -mately 53 percent of the design wind. The launch ground wind +s comparedwith the design grourd wind in Figure 11-2. The launch ground wind hada Steacy state magnitude of 7.5 m/s (24.6 ft/s) at the 18.3 meters (60 ft!Tevel, anc 12,5 m/s {41.0 ft/s) at the top of the launch tower. The com-bination of offset center of gravity, thrust imbalance, and thrust mis-alignment in yaw cancelled the yaw manent from inboard engine cant.Table 11-2 compares the vehicle misalignments measured during flight withpreflight measurements.

Liftoff vertical motion and soft release forces are shown in Figure 11-3Liftoff accelerations were greater than predicted. The slow releaseforces reduced from continuous analog recorder measurement provide anexcellent match between simulated vertical motion and the vertical motionobserved by the liftoff cameras. The AS-502 vertical rise was ahout 0.5secend faster than the AS-501 because the AS-502 had 12 lubricated slowrelease rods as corpared to 16 non-lubricated rods on the AS-501. Figure11-4 shows that the actua? (photographed) trajectory of the S-IC thrust

Table 11-1. Summary of Liftoff Clearances

POTENTIAL TATERFERENGE WacoTeyED MIN ACTWAL‘ueanance™ |cLeaRayces|

WEnTCLE. a0uN0 EQUIPMENT erin.) enfin.)Tiras Structure WotasounPowe a2 7 sa

an eanThnat Structure Holédoun Post Hood w.16 22.06

(ao) ©)Thrust, structure Lirtott switches variable * 12.7Insulation 6)Engtoe 6e11 Noldaoun Post yey 38.10 03-22(85'3) (isa) aso}service Module SM Stag Arn Versable 191.6 ed

(49.8)SL Stage S-1V8 Forward Swing varteble 101.6 “

ane (40.6)ST/Se1¥B latarstage SH1V8 ATL Setng Arm variable a] ”

SHI Sage S11 Foneard Swing Am Yarfable 116.86 ”46.3)

ST Stage S11 tntemedtate Sving vartable 116.06 “am (46:3)Fle Tp Swing Am 262.79 499.58 70),0¢

(39.68) a3) (276.0)WSeiten remains on stritar plate For Soalgna.condvtaos Conara data indicates clearance —no quantitative dats avatlable.

ns

tw Ren, Ais

ee Tgauay onsaeSieIne he te)

ned

de |

20 acTER

Moa Bob

{180 DHGy si 4

i Att yt fa

mr, i |

WHE S200, m7Figure 11-2. Surface Wind Speeds

Table 11-2. AS-502 Misalignment Summary

PRE-CAWNEH WERSURED LaurePARRIETER Perch ‘a ou parce

|

_¥aW ROLFthruse Misarfanment, ee waar

|

yoo at,a0e 0.9

|

e617

|

9.003Inboard Engine Cant, deg case

|

tie ~

|

oa

|

ooa6 -Servo Amp GFfset | oo ec - - : :enceng :Yericle Stack*ng ant PadMisaTiwrent, dag stood 0.00 6.18 -o.oo0 -t.nag 0.272

Peak Sore Release Force 348,000 (78,200) 326,000 £73,300)per Rod, W {1b#),ine 95 Percentile Envelove 7.5 wis (28.6 ft/s) atsm oo ies tenetThrust+to-de:ght Ratio 1249 1.289 + These vector measurement urcertainty.r+ A onsitive solarity was usec to determine minimum fin tip/umtilice? towerClearance. A negative polayity was aed tp deternime venicle/Gse clesrances.

7

uw *NOTLOW TOTLYaA

VERTICAL

MOTION,

in.

02

4

LWA.

AS-502

FLIGHTS

CAMERA

AND

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SIMULATED 4

+

|__|

[L-—ESTIMATED

AS-801

__|

SLOW

RELEASE

FORCES

Liz

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-502

3/-

——PR

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RELE

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4

k=

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1

‘SPE

CIFI

CATI

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SLOW

RELEASE

1Forces

0bane

0o

N\\

00

12

30

0.08

0.10

0.15

TIME

FROM

HOLDDOWN

ARM

RELEASE,

SECONDS

VERTICAL

MOTION,

m

SL EOL “O04 3S¥312a Y3d IONOF

N GOL ‘doy asvai3y Yad 39YOd

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0.36

1.36

2.36

3.36

RANGE

TIME

,SECONDS

‘VYHOLDDOWN

ARM

RELEASE,

0.36

Figura

11-3.

Liftoff

Vertical

Moti

onan

dSoft

Rele

ase

Forc

es

t100

structure (position 1) was almost vertical, with a maximum lateral driftof less than 20 centimeters {7.9 in.) after 180 centimeters (59 in.]vertical rise. The simulated data also shown compare favorably withthe camara data.

Figure 11-4 also presents the clearance between engine bell 4 and theholddown post at position I and shows that the simulated vertica’ motioof the bell is almost a duplicate of the actual motion. The motionpicture camera monitoring Ground Support Equipment (65E} operation con~firmed that the horizontal motion of engine bell 4 was less than 15centimeters (5.9 in.) after 600 centimeters (236 in.) vertical rise

Figure 11-5 shows that the combination of the yaw bias and wind blowingtoward the tower resulted in a clearance of 10.0 meters (32.8 ft} betweenS-IC Fin tip A and the top of the tower. Flight data were taken from atower camera located 426,7 meters (1400 ft) due east of the mobilelauncher.

Center engine translation and the exhaust plume angles “or each of thefive S-IC engines during the first 220 meters (722 ft) of verticaflight are shown in Figure 11-6, Center engine translation was amaximum of 7 meters (23 ft) south and 1 meter (3.3 f+} west. Maximumdeviations of the plume were 1.5 degrees north and 0.6 degree east forengines No. 3 through 4 and 1 degree north and 0.7 degree west forengine No. 5. The exhaust plume angle is the angle between the exhaustplume of an engine and the vertical, taken at the engine gimbal pointFor both translation and plume angle, positive motion is considered tooccur toward the north and the east.

11.3.2 S-IC Flight Dynamics

Table 11-3 lists maximum control parameters during S-IC burn, Dynamicsin the region between liftoff and 20 seconds occurred prinarily from theyaw bias maneuver, the start of the pitch tilt and roll maneuver, andthe end of the roil maneuver, As shown in Figure 11-7, during this timespan, maximum pitch attitude error of 1.07 degrees and engine positionof 0.52 degree occurred at 13.3 seconds and 12.9 seconds, respectively,The maximum yaw attitude error of -1.28 degrees and engine position of-0.49 degree occurred at 3.5 and 3.2 seconds, as shown in Figure 11-8These were the largest yaw dynamics encountered during S-IC flightFigure 11-9 shows toe maximum rot] dynamics during S-I¢ flight were -1.37degrees attitude error at 13.0 seconds and 1,44 deg/s attitude rate at13.4 seconds, Dynamics in this region occurred as predicted

In the region between 30 seconds and 140 seconds, maximum dynamics werecaused by the pitch tilt orogran, differences between the wind-bias windand actual wind magnitude, ahd wind shears. Figure 11-7 shows thatmaxinum pitch dynamics were 1.14 degrees attitude error at 70.5 seconds,-1.03 deg/s attitude rate at 60.2 seconds, and 0.47 degree engineposition at 70.9 seconds. The maximum pitch angle-o*-attack of 3.1 degreesoccurred at 51.7 seconds, the time of maximum difference between the wind-

11-8

11-9

An g §3

SONODAS *3NIL TONE

0.77

0.36

CAMERA

DATA

—----

SIMU

LATE

DWITH

Wo ANANAOVTESIO TYIELUIA

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DATA

HORIZONTAL

DISPLACEMENT,

in.

4a

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DISPLACEMENT,

in.

10-80

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11.83

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Figure

11-4.

:30

°0.3

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00

60HO

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DISPLACEMENT,

cm

Holddown

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AT132

DEG

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20624

0

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0.x °, cant26 win unos wsSi Sigettes® "NST, tut ORION AL O:SPLACEMEST. €

Figure 11-5. Liftoff Trajectories of Fin Tip A

Table 11-3, Maximum Control Parameters During $-IC Burn

aREPETER LUIS) FATEH vane) aa Tre FOUL CONSTRATT]uatuse Error ces 1 ae] as oar 153egular Rate dears saa foe

|

ose

|

ose

|

oat ewEngine cerlection deg ase ie -nee ao cre 5.16inversgehsagte-of-Attack ag ae] oe} asf] = :eReRegron)Harmal Accelerazion mis? 0.59 Te B 81.2 -

(sss?) 11.94) Jizat a7! tae: (aeg.cy = - 2cynam'c Pressure {U) Wee 4.76 (3.45) at FE? secones,UBFr- 0= Product Wedegy cm? 10,2 (4.8) at 68.5 sceores (ib Fae102)

1-10

W110

SGNOD3S '3WIL IONE

12.5

.

99 4.3

6.1

36.

ENGINES

7THR

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032.8

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0

120

{

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J

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re|SS

a70.1

TRAJECTORY

OF

CENTER

ENGINE,

mS-1C

PLUME

ANGLES,

deg

400

w *as1y TWOLLY3A

300

200

{——t—~t

100

Figure

11-6.

$-IC

Center

Engine

Trajectories

and

Plum

eAn

gles

aa *3STY TWOLLURA

PITCH

ATTITUDE

(POSITIVE

NOSE

UP},

PITCH

ATTITUDE

ERROR

(POSITIVE

AVERAGE

PITCH

FNGINE

POSITION

(POSITIVE

STEER

NOSE

DOWN),

deg

NOSE

UP},deg

PITCH

RATE

(POSITIVE

NOSE

UP),

deg/s

deg

|| M™Y PREM ye imevner, 1.9END YAW MANEUVER, 9.8Y sesiw PrTcH avo ROLL maueuveR, 11,1eof & 2O POLL maxewver, 31730- § MACHT ACHIEVED, 60,5

vyvv

MmOCCURRENCE OF Hix DYNAMIC PRESSURE. 75,2FLIGHT CONTROL COMPUTER SWITCH POLHG. 1, 108.66FLIGHT CONTROL COMPUTER SNITCH POINT N21 120.38END PLIGH MANEUVER, 140.8SIC Teta {SOLENOID activation stout), 166.72SIC O€¢o SENSED BY LvDC, 142,41

-sob

ap

Lei

[TPN Aang

1.5

1.0

0.5 i-

0 Wi ¥ mi

—|

vl wvi iv Re

|

Q 20 40 60 80 100 120 140 «160RANGE TIME, SECONDS

MEASURED oe = SIMULATED

Figure 11-7, Pitch Plane Dynamics During S-IC Burn

12

BEGIN Yaw MANEUVER, 1.9END YAW MANEUVER, 9.8BEGIN PITCH AKD ROLL MANEUVER, 11.1TEND ROLL MANEUVER, 31.30MACH 1 ACHIEVED, 60.5OCCURRENCE OF MAX DYKAMIC PRESSURE, 75.2FLIGHT CONTROL COMPUTER SNITCH POINT NO. 1, 105.66FLIGHT CONTROL COMPUTER SWITCH POINT 40. 2, 120.84END PITCH MANEUVER, 140.9S-IC TECO (SOLENOID ACTIVATION SIGNAL), 144.72S-1C OECO SENSED BY L¥DC, 148.41

ets 7angget! AEge oo Pr us223 MANYase cowArma POTN« 1ga3, Gase 'MEASURECtare’ - SIMULATED:

10eS 0.6

Ee lseB= “0.5 MEASURE

eS ao

Be. os TT28a ° 1 1B55 2 age2 os smuuareo

#2 ,..wivl vv wy 200 «620 640 «60 8HSCOOSC1z0 sad

RANGE TIME, SECONDSFigure 11-8. Yaw Plane Dynamics During S-IC Burn

11-13

BEGIN YAN MANE JVER, 1.9END YAN MANEUVER, 9.8BEGIY PITCH AND ROLL MANEUVER, 11.1END ROLL MANEUVER, 31.30MACH 1 ACHTEVED, 60.5OCCURRENCE OF MAX DYNAMIC PRESSURE. 75.2FLIGHT CONTROL COMPUTER SWITCH POINT NO. 1, 105,66FLIGHT CONTROL COMPUTER SWITCH POINT NO, 2, 120.84ENO PITCH MANEUVER, 140.9S-IC FECO (SOLENOID.ACTIVATION SIGNAL), 144.72

Y S-1C GECO SENSED BY LvDC, 148.47

ge ° Fecomeatogofee arriveS228 -10cee: L_/S avenueEgSE a } ~gEE® |233 ay

zag ¢B28 . SIMULATEDBag ost PM | tmonisnatasSEE -1.0 —ifes |

a hb

ae 2

efg 1

Se" 0

ee

Bee. pete]age" 2 staarea MEASURLD:

gbo8 04

zShE 0.6 we y YY YW y “200 0 2 49 60) 80 slog 120 14d(160

RANGE TIME, SECONDSFigure 11-9. Roll Plane Dynamics During S-IC Burn

Tel4

bias wind and actual wind. Yaw dynamics during this region were causedby the wind, The maximum yaw plane component of anglc-of-attack shownin Figure 11-10 wes 1,6 degrees at 84.5 seconds, the time of maximum yawwind, Figure 11-11 shows winds encountered during flight as given byJIMSPHIERC and Q-Ball reduced data. Q-Ball reduced winds were determinedusing Ball angles-of-atteck in pitch and yaw, Observed Mass PointTrajectory (OMPT) data, and vehicle characteristics data, Maximum pitcwinds in the high dynamic pressure region were 26.8 n/s (87.9 ft/s) at77.8 seconds from JIMSPIIERC data as compared to 20.2 m/s (66.3 ft/s) at77.8 seconds from Q-Ball reduced data, Yaw winds were 12,9 m/s (42,3 ft/s)at 84.3 seconds from JIMSPHERE data as compared to 13,5 m/s (44,3 ft/s}at 84.5 seconds from Q-Ball reduced data. It is necessary to introducea time varying roll moment into the simulation to match the flight dataThe rol] moment would have to reach a value of 228,000 N-m (168,200 Ibf-ft!at S-IC Outboard Engine Cutoff (QECO)

Significant longitudinal oscillations occurred during the latter part ofS-IC flight. These oscillations were coupled into the pitch, yaw, androll planes. The largest coupling occurred in the pitch plane, a:evidenced by the pitch accelerometer and pitch rate gyro outputs. Fre~quency spectrum analyses of pitch ratc, command to the actuator, andactuator position indicate that the 6.3 hertz oscillation was present inthe rate gyro signal but was absent from the conmand to the actuatorControl system attenuation of 32.4 decibels at this frequency cffectivelyreduced the signal sent to the actuator magnetic amplifiers. The smallresponse of the engine position at the 5,3 hertz frequency was duc todirect vibration of the actuator by the engines. Maximum vibration ofthe actuator was 0.0085 degree peak-to-peak.

A transient occurred in dynamics at approximately 133 seconds as dis-cussed in Section 9A. Oscillogram traces shown in Figures 11-12 and11-13 show the effect on the control system. The transient was observedin pitch and yaw IU accelerations and rates but not in the nitch and yawengine positions. The transient was not obvious in pitch accelerationand rate measured in the S-IC stage.

The transient due to Inboard Engine Cutoff (LECO) was used to determinethe inbeard engine cant of -0.36 degree in pitch and -0,26 degree in yawOther misalfgments were thrust vector angles required to match flightdata. The equivalent thrust vector misalignments were 0.0 degree inpitch, 0.17 degree in yaw, and 0.093 degree in roll, Table 11-2skmmarizes AS-802 misalignments.

Dynamics from 14C scconds to separation were caused by tilt arrest andTECO. Table 11-4 lists $-IC dynamic end conditions.

Observed ard predicted slosh frequencies are shown in Figure 11-14.Predicted slosh frecuencies are syster modes associated with slosh in theclosed Toop control system. Because the tanks are coupled through thevehicle and control system, it is difficult to associate them with a

1-15

PITCH

PLANE

ANGLE-OF-ATTACK,

{POSITIVE

NOSE

UP),

deg

YAW

PLANE

ANGLE-OF-ATTACK

(POSITIVE

NOSE

RIGHT

TOTAL

ANGLE-OF-ATTACK,

Y BEGIN YAW FANCUVCR, 1.9END YAW MANEUVER, 9.8)BEGIN PITCH AND ROLL MANEUVER, 11.1END ROLL MANEUVER, 31. 30MACH 1 ACHIEVED, 60.5OCCURRENCE OF MAK DYNAMIC PRESSURE, 75.2FLIGHT CONTROL COMPUTER SWITCH POINT NO, 1, 105.66FLIGHT CONTROL COMPUTER SWITCH POINT NO. 2, 120.84END PITCH MANEUVER, 140.9S-1C IECO (SOLENOID ACTIVATION SIGNAL), 144.72S-1C OECD SENSED BY LYDC, 148.47

-——— MEASUREO (JIMSPHERE) ————MEASURED (Q-BALL)

4

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RANGE TIME, SECONDS

Figure 11-10. Free Stream Angle-of-Attack During $-1C Burn

11-16

BEGIN YAW MANEUVER, 1.9END YAM MANEUVER, 9.8BEGIN PITCR AND ROLL MANEUVER,END ROLL MANEUVER, 31.30MACH 1 ACHIEVED, 60.5OCCURRENCE OF MAK DYNAMIC PRESSURE, 75.2FLIGHT CONTROL COMPUTER SWITCH POINT NO. 1, 105.66FLIGHT CONTROL COMPUTER SWITCH POINT NO, 2, 120.84END. PITCH MANEUVER, 140.9S-IC TECO (SOLENOID ACTIVATION SIGNAL}, 144.72S-IC OECO SENSED BY L¥DC, 148,47

——— NEASURED (JIMSPHERE}~-—-—— CALCULATED FROM ONBOARD DATA - Q-BALL

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RANGE TINE, SECONDS

Figure 11-11. Pitch and Yaw Plane Wind VelocityDuring S-IC Burn

V-17

11-18,

During S-IC TransientOscillograms of [U Control SensorsFigure 11-12.

RANGE TINE, SECONDS

LuPITCH

TuYA

WIU

PITCH

RATE

ACCE

LERA

TION

ACCELERATION

133

NOTE:

134

135

136

MAXIMUM SENSOR READING (+0.5g3 +10 deg/s)

FULL SCALE CORRESPONDS APPROXIMATELY TO

137

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NOTE: FULL SCALE CORRESPONDS APPROXIMATELY TO MAKINUM SENSORREADING (45 deg; #C.5 gi +10 deg/s)

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ACCELERATION

LA

i HUI | HLL133 $34 135, 136 137 138

RANGE TIME, SECONDS

2ATE S-

ICPITCH

Figure 11-13. Oscillograms of Engine Position and S-1C SensorsDuring S-IC Transient

1-19

*Table 11-4. $-1C Dynamic End Conditions

PARAMETER UNITS PITCH YAH ROLL

Attitude Error deg -0,42 0,84 0.75

Attitude Rate deg/s 0.06 0,005 -0.70

Average Actuator deg 0.08 -0.16 0.10Position * Conditions at separation command, 149.08 seconds.

particular tank. An attempt was made to do this, and se they are labeled“coupted". They are different from the individual tank frequencies whichshould not be present during flight. Presence of other modes was detectedin the tanks, as was expected, These were bending and control systemmodes. Predicted curves for these are also shawn where required.

Propellant slosh amplitudes in the $-I¢ tanks are shown in Figure 11-15,Measured peak-to-peak amplitudes were derived from opposing pairs ofliquid level probes in the pitch and yaw planes. Simulated data shownas a comparison were derived using first mode slosh models in the pitchand yaw planes. The figure shows the oscillating slash wave peak-to-peakmagnitudes in addition to low frequency effects due to non-zero lateralaccelerations during flight. Maximum peak-to-peak slosh amplitudes inthe fuel tank were -0.3 meter {-0.98 ft) in pitch at 65 seconds and 0.25meter (0,82 ft) in yaw at 2.5 seconds. $-IC LOX peak-te-peak amplitudesteached ~0.3 meter (0.98 ft) in pitch at 72 seconds and 0.15 meter{0.49 ft} in yaw at 19, 28, and 85 seconds. Simulated data for the S-ICtanks showed close agreement in phasing but measured sTosh amplitudeswere larger, which may indicate tess slosh damping than predicted,Figure 11-16 shows propellant slosh in the S-II and $-IVB tanks. Ampli-tudes shown for the S-11 slosh give only the oscillating slosh wavepeak-to-peak amplitudes. Static, or low frequency osci}lations, havebeen removed from the data. Only S-IVB LH2 stosh data are available.There was poor agreement between measured ard simulated slosh amplitudesin the S-IT LOX and S-IVB LHe tanks, Because the S-IT LOX probe waslocated very near the center of the tank, it was difficult to correlatemeasured and simulated data using only first mode simulation data,Maximum peak-to-peak amplitudes in the plane of the Propellant Utilization(PU) probes were 2.7 degrees at 90 seconds for the S-IT LOX tank and1.4 degrees at 95 seconds for the S-11 LH2 tank. Maximum $-I¥B LH2 sloshwas 0.084 meter (3.30 in,} peak-to-peak at 85 seconds. There was noevidence of unstable buildup.

Peak-to-peak engine response to propellant slosh is shown in Figure 11-17.The response was derived by passing measured and simulated engine deflec-tion time histories through bandpass filters, retrieving only slosh

11-20

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frequency components. Since frequencies of signiftcant slosh modes laywithin a relatively narrow band, the engine responses shown were due toall tanks collectively. Maximum pitch engine resnonse to slosh was=0.12 degree at 13 secorids. Maximum yaw engine response to slosh was0.1 degree at 11 seconds. The disagreement between measured andsimulated engine response to slosh shown on Figure 11-17 tends tosupport the indication of smaller slosh damping than predicted. Thesmall actuator activity at slosh frequencies other than at the time ofknown disturbances confirmed that slosh was adequately stabilized.

Peak-to-peak engine response to first bending mode was determined bypassing measured and simulated engine deflection time histories throughbandpass filters, retrieving only bending frequency components. Maximumresponse to first bending was 0.026 degree at 85 seconds in pitch and0.02 degree at 90 seconds in yaw. Results indicate that bending dynamicswere adequately stabilized throughout flight.

11.4 S-TT CONTROL SYSTEM EVALUATION

The S-II stage attitude control system performance was satisfactory.Analysis of the magnitude of modal components in engine deflectionsrevealed that vehicle structural bending and propellant sloshing hadnegligible effect on control system performance prior to cutoff ofengines No. 2 and 3, The maximum values of control parameters occurredjn response to shutdown of engines No. 2 and 3. The shutdown of engine:No. 2 and 2 caused attitude errors of 13.4 degrees in pitch, 2.7 degreesir yaw, and 4.2 degrees i roll. Attitude rates for pitch, yaw, androll were 2.8 deg/s, -C.8 deg/s, and -2.2 deg/s, respectively. Theresponse at other times (such as S-IC/S-II separation and initiation ofIGM guidance) were within expectations.

11.4.1 Attitude Contre] Dynamics and Stability

Attitude control commands were computed in the minor Toop section of theLYDC. For 42 seconds following S-iC Q&CO these cormands were heldconstant. Significant events occurring during thet interval were S-IC/S-1separation, S-I] stage J-2 engine start, second plane separation, andLaunch Escape Tower (LET) jettison. The attitude control dynamicthroughout this interval indicated stable operation (see Figures 11-18theough 11-20) Simulated data shown for comparison in these figure:generally leads actual data by about 2.6 seconds because of a 2-secondbias in telenetry data used as input commands to the simulation and a0.6-second bias in the actua? data. The maximum control excursionsoccurred in the roll axis following S-IC/S-I1 separation when 2,0 deq/:attitude rate and -2.3 degrees attitude error occurred, as shown inTable 11-5. Steady state attitudes were achieved within 10 seconds fromS-IC/S-I] Separation. The principal attitude error of approximately0.5 degree for the roll axis was maintained until cutoff of engines No.2'and 3, Absence of roll rate during this time indicates a combinationof center of gravity offsets and thrust or engine misalignments producinga constant roll torque, Similar roll offsets existed during AS-501 S-Istage flight.

11-26

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WY

s-IC/S-HI SEPARATION COMMAND, 149.08S-II SECOND PLANE SCPARATION COMMAND, 179.06INITIATE IGM PHASE 1, 190.95FLIGHT CONTROL COMPUTER SWITCH POINT NO. 3, 269.76FLIGHT CONTROL COMPUTER SWITCH POINT NO. 4, 339.76SIL ENG NO. 7 OUT, 412.92;S-IE ENG NO. 3 OUT, $14.38:FIRST ARTIFICIAL TAU INITIATE, 415.4‘V FIRST ARTIFICIAL TAU TERMINATE,SECOND ARTETICIAL TAU INITIATE, 490.8¥ SECOND ARTIFICIAL TAU TERMINATE, 510.2

v INITIATE CHI FREEZE, 517.7S-IT ECO SENSED BY LDC, 576.33

Figure 11-18. Pitch Plane Dynamics During S-II Burn

11-26

YW S-1C/5-11 SEPARATION COMMAND, 149,085-17, SECOND PLANE SEPARATION COMMAND, 179.06INITIATE IGM PHASE 1, 190.95FLIGHT COMTROL COMPUTER SW: TCH POIKT NO. 3, 209.76FLIGHT CONTROL COMPUTER SWITCH POINT HO. 4, 339-76S:IT ENG 80. 2 GUT, 412.925S-11 ENG 80. 3 OUT, 414.18;FIRST ARILFIGTAL TAU INITIATE, 415.4

W FIRST ARTIFICIAL TAU TERMIKATE,SECOND ARTIFICIAL TAINCTIATE, 490.8SECOND ARTIFICIAL TAU TERMINATE, 510.2INITIATE CHE FREEZE, 517.7

‘Y S-1F ECO SENSED BY LYOC, 576.33* NOF CORRECTED FOR A 0.6 SECOND TIME BIAS IN TELEMETRY DATA‘NOT CORRECTCD FOR A 2-0 SECOND TIME BIAS IN SIMULATION INPUT

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W-27

S-IC/Sc1L SEPARATION 2OMPAND, 149,08SeIT SECOWD PLANE SEPARATION ‘COMMAND, 179.05INITIATE 13M PHASE T, 190.95FLIGHT CONTROL COMPLT:R SWITCH POINT NO. 3, 209.76FLABHF CONTROL COMPUTER SWITCH POIN 40. 8. 339.78S-HT ENG NO. 2 OUT, 412,92;S-H1 ENG NO. 3 9UT, 414.185FIRST ARTIFICEAL TAU INITIATE, 415.4FERST ARTIFICIAL TAU TERMINATE,SECOND ARTIFICIAL TAU INITIATE, 490.8SECOND AMILFICIAL TAU TERMINATE, 510.2INITIATE CHI FREEZE, $17.7S-II ECO SENSED BY VDC, 576.33

dd

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MEASURCD DATA (NOT CORRECTED FOR A 0.6 SECONDTIME BIAS IN TELEMETRY DATA)| SIMULATED DATA [OT CORRECTED FOR A 2.01 SECONDTIME BIAS IN SIMULATION INPUT)

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Figure 11-20. Roll Plane Dynamics During $-11 Burn

11-28

Table 11-5. Maximum Control Parameters During S-II Burn

PARANETER Sale/s-lt | ge| sO. 2 83) TERMINATE 2no S11SEPARATION

|

aytStrrow

|

ENG FAIL

|

ARTIFICIAL TAU| CUTOFF

Pitch PlaneAttitude Error,deg 0.5 “19 13.6 32 nABody Rate, deg/s] 9.9 1.0 28 10 0.1Iwerage CnbalAngle, deg 0.24 0.9 5.95 4.37 3.5Slosh Conponentof yerage GinbalAngle, deg 0.01 0.06 0.18 9,02 :

fay PlaneAttitude Error,deg “0.8 “0.4 27 0.2 0.0Body Kate, deg/s 0.3 oO “0.8 “0.15 0.2Average GinbalAngle, deg 0,83 -0.23 0.62 0.19 anStosh Conponentaf Average Gimbalfogle, deg 0.03 4.05 0.8 0.01 -

Poll PlaneAttitude Frror,deg 2.5 08 4.2 “07 0.2Body Rate, deg/s 2.0 “04 222 oa 02Average GimbalAngie, deg -0.49 0.23 0.39 0.1 “ole

IGM was initiated at 190.95 seconds, and the flight contro] computerreceived thrust vector control commands to pitch the vehicte up.Following IGM initiation, a -1.9 degrees pitch attitude errar, 1.0 deg/spitch rate, and 0,9 degree pitch gimbal deflection occurred, Theseresponses were similar to those obtained during AS-501 flight. Theeffects of steering misalignment corrections (initiated at 212.0 seconds)and flight contro? gain switch point 3 had no noticeable effect uponattitude control performance. The thrust reduction on engine No. 2 at319 seconds {see Section 6.3) had a minor effect on attitude control,

11-29

Subsequent to a drop in thrust in engine No. 2 @ discrete signal wareceived by the IU from the S-I1 stage that indicated thrust not OK onat least one engine, During successive LVDC computation cycles, aceceleration decrease due to cutoff of engines No. 2 and 3 wes confirmedThis confirmation and the external discrete caused the guidance systemto enter an engine-fa¥'ure-mode artificial tau, The principal dynamicsfollowing cutoff of engines No, 2 and 3 occurred in the pitch axts andwere attributable to a pitch disturbance moment which acted on thevehicle as a result of the two engines out. The pitch rate built up toa maximum of 2.8 deg/s before the control engires could be repositionedto counteract this moment.

Pitch axis control was stabilized within 16 seconds from time of theengine cutoff. The pitch attitude error reached a maximum of 13.4 degreesand remained above 6 degrees until final engine cutoff prior to $-I1/S-IVBseparation, This attitude error was reouired by the contro] system inorder to keep the centrol engines positioned to counteract the effects ofengines No. 2 and 3 hein inoperative. With only two outboard engines(engines No. 1 and 4} operating, the yaw and rol? axes control systensinteracted, A domfnant mode with a period of 20 seconds and dampingfactor of 0.7 was apparent in their responses following the shutdown ofengines No, 2 and 3.

The effects of terminating the second artificial tau mode at 510.2 secondswere most apparent in the pitch axis when a 1 deg/s pitch-up responserate occurred. This was a consequence of the change in conmand angle atthts time, Seven seconds later 2 cai-freeze mode was entered. Afterthis time the vehicle rates were reduced to Tess than 0.2 deg/s for therenaining 40 seconds of the S-II stage flight, At S-II stage enginecutoff (prior to S-IVB separation), the vehicle attitude errors (excentpitch), and attitude rates were at or near null

Steering Misalignment Correction mode became operative at 212.0 seconds,as shown in Figure 11-21. Its principal effect upon the contro] systemwas to introduce a low frequency mode {approxinate"y 0.05 hertz) belowthe rigid body control mode {approximately 0.11 hertz). During steadystate attitude control operation the magnitude of SMC is a measure of thrustmisalignment. and center of gravity offsets. Following cutoff of engine:No. 2 and 3 the SMC angles increased to a maximum of -0.85 and -13 degree:for the yaw and pitch axes, respectively. The low frequency mode presentin the yaw and roll axes vehicle responses shown in Figures 11-19 an11-20 is also apparent in the SMC yaw command angTe corrections shown érFigure 11-27,

At 319 seconds a load increase of 32,000 Newtons (700C Thf) appeared onboth the pitch and yew actuatars of encine Ke. 2 (see Section 8.3). Thiswas precisely the time engine No. 2 chamber pressure dropped, signifyinga reduction Of thrust in the engine cf 33,806 Newtons (7600 Tbf). (SeeSection 6.3.) Figure 11-22 skows the load pressures, input commandcurrents, and the actuator positions for the pitch and yaw actuators of

11-30

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After cutoff of S-II eng‘ne No. 2, the yaw actuator no longer respondedto commands (reference Section 8, Figure 8-5),

11.4.2 Liquid Propet-ant Dynamics and Their Effects on Flight Contra

Estimates of liquid nvopellant dynamics were extracted from the fine nasprobe liquid sensor measurements. Assuming planar wave propel tant slosh,the LOX probe measures 98 percent of the pitch component and 20 percentof the yaw component. The Lig probe measures 16 percent of the pitchcomponent and 99 percent of the yaw component.

Estimated LOX and LHz slosh amplitudes during S-II boost are shewn iFigure 11-23. These data were extracted from the fine mass probe liquidsensor measurements by a data processing technique specially developedfor this purpose. The amplitude plots show periodic biasing (non-sinusoidal) which should be ignored because it is a consequence of theprocessing technique. The top two plots are for the Lt2 slosh amnlitudesat the probe and wall, respectively. The bottom plet is for the LOXsurface angle, The slost modes were excited after S-I1 stage J-2 enginestart, IGN initfation, cutoff of engines No. 2 and 3, and termination ofartificial tau cuidance mode, as indicated in the figure. The largestLig slosh amplitude was 15 centimeters (6 in.) at the probe, andoccurred after cutoff of engines No. 2 and 3. The LOX sTosh amplitudewas more sustained than Lig sleshing throughout S-II boost.

S-IVB stage slosh date are shown fn Figure 11-24. Lp sloshing wasdiscerrible at S-II engine ignition (149.8 seconds) and at engines No. 2and 3 cutoff (412.9 and 414.2 seconds). The LH2 slosh quickly damedcut (within 15 secends). This corresponds to an estimated damping factorof ©.26. The S-1VB LH2 slosh frequencies were near the natural uncoupledfrequency as shown in Figure 11-25. LOX sloshing in the S-IVB tanks wasnot discernible.

The slosh data were also analyzed for frequency content. Figure 11-25depicts slosh frequencies estimated from fine mass probe measurenents inthe S-II stage liquid oxygen and liquid hydrogen tanks. The Liz frequencyplots show that sloshing occurred near the natural (uncounted) LH2 slosh

11-33

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saNGE TUM, secoves19 Bruce" TRUST, 153.03 Y str eno aw, 2 aur, 42.92

je ; 2 SLE ENG 8025 DUTY a1.Figure 11-25, Predominant Slosh Frequencies During $-II ‘Burn

11-35

SLOSH

AMELITLDE

fp-o)

frequencies. The Lig sloshing was initially excited during S-IT stageJe2angine ignition {149.8 seconds) end initiation of 1am {191.0 secondsand then oscillated independently with slow decay due to damping. Theobserved LOX frequencies were in the range of 0.4 to 0.6 hertz. Thesefrequencies were also evident in the S-II pitch and yaw gyros. The LOsloshing was apparently coupled with the flight contra} system response.it is theorized that snal1 amplitude (less than 0.02 degree) attitudelimit cycling of the Flight control system was forcing the LOX sloshdynamics. A small amplitude attitude limit cycle was predicted ky theanalog simulation of the S-IT stage flight dynamics control systemAttitude control timtt cycling is possible because of the quantizationof flight control data by the LYDA and the non-linearity of the LVDA-LDC. The results of simulation of Lig amplitude at the wall are plottedin Figure 31-23 for comparison with flight cata. The cornarison indicate:good agreement during the initial excitation of the slosh modes; however,the decay characteristics of the simulation show mare attenuation. Thisindicates that less damping of the slosh modes occurred in flight thananticipated by the simulated flight dynamicsThe effect of liquid propellant stcsh non the control system wasestimated from the component of the slosh mades in the engine deflections,presented in Figure 11-26, The engine deflections were analyzed usingbendpass filtering.

The largest slosh component magnitudes of 0,18 and 0.15 degree occurredwith the yaw and pitch ginbaT angles following cutoff of engines No. 2and 3. The yaw magnitude was approximately 30 percent of the total yawgimbal angle at this tive (see Table 11-5).

S-IL stage J-2 engine No. 1 pitch and yaw actuators position data werebandpass filtered at five selected Flight times to detect the presenceof any structural bending mode frequency components. The results indicatenegligible engine deflection due to structural bending of the vehicle.This result was to be expected and indicates the flight control computerelectronic filters were effectively attenuating any bending mode effectsin the rate gyro or other control system inputs.11.5 S-IVB CONTROL SYSTEM EVALUATION

The S-IVB Thrust Vector Control System and APS provided satisfactorypitch, yaw, and roll contro! during S-IVB first burn and parking orbit.buring restart attempt the auxiliary and main hydraulic pump cavitated,precluding operation of the T¥C system. ‘With this exception, the vehicleattitudes correlated well with actual commanded attitudes, and demandson the control system were well within the capabilities of the system,

11.5.7 S-I¥B Control System Evaluation Sefore and During First Burn

From 118 to 133 seconds during S~IC burn, a 5.5 hertz oscillation atapproximately 276 N/cmé (400 psid) peak-to-peak was observed on the S-IVBPitch actuator differential pressure. Figure 11-27 presents a comparison

11-36

11-37

S-IC

/S-I

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COMMAND,

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NO.

2&8

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Figure

11-26.

$-IT

Engine

Deflection

Response

toPropellant

Slosh

3Bap (4-4)

NOILO31430 MYA

400 6002 a0Be 200 th i ta ; Re=e 200 58

Ee 9, 0 SeBO ono -200 82 +200

2

= 40 soo

a KALALKAd TA wa PA 200 52ES oo o ESR -200 200 &2 Us 120 122 la 126 128 130 132 134 136 138 2RANGE TIME, SECONDS

Figure 11-27. S-1VB Pitch Actuator Cxcitation During S-IC Burn

of AS~502 and AS-501 pitch actuator differential pressure during thistime interval. The §.5 hertz oscillation was not evident on the yawactuator differential pressure data, Section 9 contains a discussionof the POO phenomenon which caused this.

At 877.1 seconds (S-11/S-IVB seperation command), there was a 0.3 secondAPS firing by the pitch engine at position I (Ip}. The pulse wasterminated at approximately 577.4 seconds by the S-1VB Burn Mode ONswitch selector command. This command places the flight control computertn the S$-IVB burn mode, which enables J-2 engine gimbaling for nitch andyew control during S-1VB burn, and enakles the APS to control in roll.Because the flight control computer was in the S-IVB coast mode for this0.3 second interval and the pitch attitude error signal was of sufficientmagnitude, the Ip engine fired as conmanded. This is a normal occurrenceon Saturn V. AS-801 had a simflar unsynchronized yaw/roll pulse on theAPS ergine No. 4 at position 1 {Iry) during staging.

The conditions on Saturn ¥ are not the same as on Saturn IB, in whichthe flight control computer goes directly into S-IVB burn mode fromS-IB burn mode and cannot go to coast mode and trigger the APS unlesscommanded to coast mode.

The TVC responded satisfactorily to flight contro] computer commandsduring first burn. The maximum attitude errors and rates occurred atS-II/S-IVB separation. A summary of the maximum values of criticalflight control parameters is presented in Table 11-6.

11-38

Table 11-6. Maximum Control Parameters During S-IVB Burn

S-1T/S-1VESEPARATION, BEGIN S-1¥8

PARAPITER SUID. INITI- CHI FIRSTATION, AND BAR CUTOFFART. TAU

Pitch Attituce Errar, de 79 -1.0 -0.9

Yaw Attitude Errar, deg -2.3 -1.0 -0.8

Rol] Attitude Error, deg -0.9 0.5 0.4

Pitch Rate, deg/sec -3.5 Mw 1.0

Yaw Rate, dag/sec Ww -0.2 0.1

Roll Rate, deg/sec -0.6 0.0 0.0

Pitch Actuatar Pos., deg 6.7 0.2 0.2

Yan Actuator Pos., deg 1.3 -0.5 -0.6 The large attitude error at S-II/S-IVB separation resulted in a pitch

actuation position requirement of 6.7 degrees, but the separationtransient was within the capabilities of the T¥C system. The S-I¥B pitch,yaw, and roll dynamics, including command angles, attitude errors,angular rates, and actuator positions, are presented in Figures 11-28,11-29, and 11-30, respectively. The pitch and yaw effective thrust vectormisalignments were +0.25 degree in pitch and -0,4 degree in yaw. Thesteady-state roll torque was 54 N-m (40 Ibf-ft). The powered flight APS‘impulse requirements are included in Table 11-7.

LOX and LHz slosh parameters are presented in Figures 11-31 and 11-32,respectively, LOX sloshing exhtbited the coupling found in prevtousflight data, LH2 sloshing indicated a large slosh wave following S-IT/S-IVE separation due to the large control systen transient which occurredat that time. The slosh wave was highly damped by the deflector locatedin the area of quiescent surface level, Variations in LH2 sloshingparameters were also noted at approximately 645 seconds, due to the nose-up command which occurred at that time. The LOX and LH2 sloshing did notsignificantly affect the control system operation during S-IV8 burn

11-39

Sc1Ve ENGINE START SEQUENCE FITCH comaNDAose LP

comaND, 577.28 ATTITUDE, 648.02Y IWITIATE L6H, PHASE 111, $84.78 YY INITIATE cu FREEZE, 746.41] aMTirrcia. Tau Teemte, 504.3 SCONE VELOCITY cinoFe coreuNo,“$0 T T C7: i ATT

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RANGE TIME, SEconDSFigure 11-28. Pitch Plane Dynamics During S-IVB First Burn

11-40

ddq S-INB ENGINE START SEQUENCE COMMAND, 577.28 WY PITCH COMMAND NOSE UP ATTITUDE, 644.02

INITIATE 16M PEASE 111, $84.78 ‘V INITIATE CHI BAR STEERING, 712.3ARTIFICIAL TAD TERMINATE, 594.3 WW INITIATE CHI FREEZE» 748.43S-1vB vELOCITY CUTGPF cOWMAND, 747.04

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Figure 11-29, Yaw Plane Dynamics During S-IVB First Burn

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11-42

1-43

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11-7,

APS

Impu

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Requirements

MODULE

MODULE

APS

ENGINE

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Figure 11-31, S-I¥B LOX Slosh Frequency and Height at the Probe06 a

_

PREDICTED secon WOE __ |===os Pe ee oo° °04

° ==oof eee wep een; PREDICTED FURST HOME

0.2°

at | -° © 88-502 FLIGHT paTA16 6

@ HEIGHT WITHOUT ATTENUAT2 ON,Te (at Peobe) we© erowrtH arTewarron

(ar pease)2+ a

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° a @ ale |,0 eWay

Figure 11-32,RANGE THM, SECONDS.

S-IVB LHy Slosh Frequency and Height

yeaa

11.5.2 Control System Evaluatior During Parking Orbit

Evaluation of existing crbital data indicates that the APS attitudecortrel perfarmed satisfactorily during orbit.

Pitch, yaw, and rol] control attitude errors, angular rates, and APSpulses following first burn are shown in Figures 11-33, 11-34 and 11-35,respectively.

Orbital attitude control requirements were much larger than on previousflights, which can be attributed primarily to initial conditionsexisting at S-IVB first and second cutoff.

Initial conditions at S-IVB first cutoff (80 degrees nose-up fram thelocal horizontal and 1 deg/s nose-up angular rate} required considerablymore APS activity than anticipated. Following activation of the APS,the pitch engine fired F411 on for approximately 43 seconds orimarilyto correct for the rate condition. At 762 seconds (Ts +15 seconds) thepitch engine was conmanded to verform the relatively large maneuver ofaligning the vehicie with the local horizontal. At 837 seconds (Ts +90seconds) the S-1VB began a sertes of maneuvers planned to produce informa-tion on the S-I¥B restart bottle repressurization (Section 7.5 andFigure 10-8) and propellant slosh excitation while qualifyina theseManeuvers for manned flight (see Mission Plan}, This sequence ofmaneuvers consisted of a 180 degree roll, 20 degree pitch dovn, 20 degreepitch up, and a 180 degree roll, At 837 seconds (Ts +90 seconds) theS-IVB was commanded to roll countercTockwise 180 degrees to alignposition III toward earth,

The 20 degree pitch down maneuver was initiated at 3207 seconds (T5 +2460seconds) for the purpose of evaluating propellant sloshing during amaneuvering period. The commanded and actual vehicle attitudes, attitudeerrors, angular rates, and APS puises for pitch are shown in Figure 11-36Following initiation of the maneuver, the pitch attitude error increased.as expected, to the -2.5 degree attitude error limit, which establisheda maneuvering rate of 0.3 deg/s. APS engine Illy fired to establish therequired maneuvering rate, which was maintained until the vehicle attitudeapproached the commanded attitude, when engine Ip fired to reduce theangular rate to the desired orbital sitch rate,

High frequency (19.5 to 22 sertz) rate signals having a maximum amplitudeof 0.5 deg/s peak-to-peak were experienced, particularly on the roll rategyro, commencing at 331] seconds and terminating approximately 15 secondslater. This time correlates with the time period of high frequency{ansroximately 5 to 6 firings per second) APS engine firings required toterminate the pitch maneuver. The high frequency rate signals appearedto terminate as the APS engine Firing frequency reduced to a lower level(less than 4 ffrings per second). Correlation of the high frecuencyrates and APS firings indicates that the high frequency signals wererelated to the APS engine firings, The high frequency rate signals,

11-45

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Teo Ora: 25RANGE TIME, 4OURS :MINUTES: SECONDSLOSS CF SECM, THSERTIOR SHIP S-2VB YELOCITY cUTaFE ccMMAND, 747.08MERCOUISITION OF 'SIGHAL, CANARY I5LAnIOS

SY

HANEIWER 1D LECAL HORIZONTAL, 762.9mmm ToIcaTES THOIVEOLAL PULSES INITIATE 12) DES ROLL TCTD interes FULL ov PuLse PIACF POSITION TT] DOvt1, 37.3Figure 11-23. Pitch Control Dynamics Following S-IVB First Burn

11-46

. neers aTTITUE

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‘LOSS OF SIGNAL, IHSERTION SHIP‘+ ACQUISITION OF SIGNAL, CANARY ISLANDSMMM TWOLCATES NDE VIOURL Fuses COO TAOICATES FULL Om PUISe

Y 5-190 VELOCITY CUTOFF COMMAND, T47.u61G NRMEUVER TO LOCAL HORIZONTAL, 762-3Y CITENTE 180 GEG ROLL 70 PLACE POSI™LON (11 DOW, 837.3

Figure 11-34, Yaw Control Dynamics Following S-lv8 First Burn

W-47

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Figure 11-35. RoTl Control Dynamics Following S-IVB First Burn

11-48

PITCH

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9:51:40 9:56:00 0:56:06 0:88:00 1:00:60

RANGE TIME, HOURS: MINUTES: SECONDS(MMB INOLCATES INDLYIOUAL PULSES INITIATE 20 DEG PITCHCT INoIcaTEs FULL Of PULSE DOW MANEUVER, 3207.9Figure 11-36. Pitch Control Dynanics During 20’ Degree

Pitch Down Maneuver

11-49

which were not of sufficient magnitude to affect attitude contro] systemoperation, appeared to be similar to those experienced on previous flightsand during other intervals of the AS-502 missicn. The APS responseappeared normal during this maneuver.

LH sloshing was not appreciable during the 20 degree pitch down maneuver.Significant LOX slosh existed at the initiation of the maneuver; however,LOX sloshing was not sustained cue to the high LOX damping,

The 20 degree pitch up mareuver to the local horizontal was initiatedat 6427 seconds (T5 +4680 seconds). The cormanded and actual vehicleattitudes, attitude errors, angular rates, and APS pulses are presentedjn Figures 11-37, 11-38, and 21-39. APS pitch engine Ip fired to removethe existing orbital pitch rate and establish a pitch up rate. The pitchattitude error increased to the attitude error limit of 2.5 degrees,establishing @ maneuverine rate of -0.3 deg/s. As the actual vehicleattitude approached the commanded attitude, APS engine IIIp fired toremove the maneuver rate and establish an orbital pitch rate (nose down).High frequency oscillatiors were also observed during this maneuver,particularly on the pitch and roll rate gyros, Again, the high freouencyrate signal (0,2 deg/s peak-to-peak for approximately 25 seconds) existedpredominantly during irtervals of high frequency APS engine firings(approximately § firings per second) and damped out at approximately4 firings per second. The APS operation appeared normal during the 20degree pitch up maneuver.

LHg sloshing was negligible during the 20 degree pitch up maneuver, aswas true during the pitch down maneuver. Significant LOX slashing occurredat the maneuver inttiation but was not sustained due to large LOX damping.

A distinct difference in pitch attitude errors was noted between thepitch down and pitch up maneuvers (see Figures 11-36 and 11-37). Thepitch up maneuver exhibited considerably more overshoot than the pitchdown maneuver. This was attributed to the greater impulse required toterminate the pitch up maneuver and establish the orvital pitch ratethan for the pitch down maneuver. Only part of the maneuvering vateduring the pitch down maneuver must be removed due to the requirement.for the orbital pitch rate, whereas all of the maneuvering rate must beremoved and an orbital rate established in the opposite direction forthe pitch up maneuver. APS impulse requirements to initiate and terminateeach of the pitch maneuvers were as follows:

Pitch Down

Initiate TerminateModule at 4775 Nes Module at $026 NesPosition III (1060 lbf-s) Position 1 (1130 1bf-s)

Pitch UyInitiate Terminate

Module at 10,097 N-s Module at 9B75 NesPosstion I (2870 1bF-s) Position 111 (2220 1bf-s)

11-50

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Figure 11-37.

3 0 BBO 5700 FBI 5900-6000 16D Azad 6300 E400RANGE TIME, SECONDS

4200 1738200 1:42:00 7:46 40RANGE TIME, HOURS MINUTES: SECONDS

CIVIOUAL PULSES.ILL_ON PSLSECEG PITCH WP MANEUVER, $427.3DEG ROLL TO PLACE POSITION’ 1 DOWN, $787.3

Pitch Control Dynamics During Pitch and Roll Maneuvers

1-81

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734300 1:38:00 az 1ia6:40RANGE T-ME, HOURS: MENUTES SECORES.

MBINOICATES 1HpLvTOUAL PULSES!CAINDICATES FULL ON PULSE

TNITIATE 20 BEG PITCH UP MANEUVER, 9427.3INITIATE 180 DEG ROLL TO PLACE POSITEON I DOWN, 5787.3

Figure 11-38. Yaw Control Dynamics During Pitch and Roll Maneuvers

11-52

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130700 1234200 1:38:00 zRANGE TIME, HOURS :MINUTES :SECONDS

MEINDICATES INDIVIDUAL PULSESTD IMOIGATES FULL ON PULSE,

INITIATE 20 DES PITCH UP MAREUVER, $427.3INITIATE 180 DEG ROLL TO PLACE POSITION I DONN, 5787.3

Figure 11-39. Roll Control Oynamics During Pitch and Roll Maneuvers

76:40

11-53

Figures 11-37, 11-38, and 11-39 show pitch, yaw, and roll control dynamics,respectively, during the 180 degrees counterclockwise roTT maneuver torealign position I toward earth at 5787 seconds (T5 +5040 seconds). Nohigh frequency (19.5 to 22 hertz) oscillations were exerienced duringinitiation or recovery phases of this maneuver.11.5.3 Control System Evaluation During Restart AttemptPitch and yaw commanded and actual attitudes, attitude errors, angularrates, and actuator positions during S-IV3 restart are shown in Figures11-40 and 11-41. The roll attitude error, angular rate, and attitudecontrol system engine firings during this interval are shown in Figure11-42," An auxtliary hydraulic system failure (see Section 8.6) preventedthe J-2 engine from being centered at the time of S-IVB engine startcommand. The engine position at ESC was approximately 1.5 degreesin pitch and -2.3 degrees in yaw. Disturbances experienced during chill~down were relatively small but increased to a significant magnitudeduring a short time interval when the thrust built up to approximately44,482 Newtons (10,000 Ibf}. This can be seen on both pitch and yawrates at 11,626 seconds, The apparent discontinuities in yaw and pitchattitude errors at 11,630.33 seconds were due to the shift from Té te TyAt T7 initiation the cormanded attitude was set at the instantaneoucutoff attitude, thus nullify?ng any attitude errors which were presentat S-IVB main engine cutoff. The thrust vector control system did notprovide pitch and yaw control during the short time interval that thisystem was active due to the aforementioned lack of hydraulic systerpressure,

11.5.4 Control System Evaluation After Restart Attempt

With failure to achieve restart, the LVDC went into Ty and initiatedthe planned maneuver to the attitude desired for snacecraft separation.The spacecraft was actually separated by ground command to the spacecraftJust after initiation of the maneuver. This caused the S-IVB/1U centerof gravity to move further aft than normal due to the large LOX massstil] onboard. This condition reduced the available APS contro? moment,thus requiring longer APS firings and greater propellant consumptionthan normal to complete the separation mancuver and to control propellantsloshing.

The pitch and yaw channels of the APS were activated at approximately11,633.8 seconds (T7 +3.5 seconds) after which the pitch engine atposition III (Itt) and yaw engines [yr and [111] care full on. Theditch engine remained on for approximately 67 seconds. The extendedcontrol engine firings were attributed to initial rates at S-IVB cutoff(pitch -0.5 deg/s, yaw 0.45 deg/s), the spacecraft separation maneuver(11,650 seconds), spacecraft separation, propellant sloshing, and LOXventing, The attitudes, attitude errors, and body rates about the threeaxes are shown in Figures 11-40, 11-4, and 11-42,

Vi-88

PSTCH

ANGULAR

RATE

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29),

deg/'s

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tty11,500 T1600 11,700 17,860" T7900

RANGE THE, SEcoHOS. vy yyy wy 3

ane 3:18:00 Ts Fie:2aANGE TIME, HOJRS: MINUTES: SECONDS

SM rorceres tno]VIOWAL PULSES 5 Wotcares FULL on PULSE$190 ENGINE RESTART comma, 11,614.69 INITIATE 5/C SEPARATION MANEUVER, 11,650.33S:lyB CD TNTERGUPT,, START OF TIME @ASE 7, S-I¥B/CSH PHYSICAL SEPARATION 11.667.82¥i,850.a3; LOX VENT” OPEN, 11630.51; 4H, TERMINATE PITCH MEMEUVER, 17,50VENT OREN, 11,630.70. TERMINATE ROLL RAMEUVER, "11,374

1 Lox vewr ¢.0sé0, 11,800.28Figure 11-40. S-IVB Pitch Attitude Errors and Rates

‘During Attempted Second Burn

11-95

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RATE

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SY

S-1yo/cow PuvsieaL SEBRRRTOM, 11 86F 821 630.33: a8 YEN on, 11630. 915 Dp TERMINATE PEYCH weve, 11,59YEHT geen, 11,630.20 TERMINATE ROLL NWMCUVER, "11,274W Loe veNr cLoséa,

2600.28

Figure 11-41. $-IVB Yaw Attitude Errors and RatesDuring Attempted Second Burn

11-56

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ANGE TIME, HOURS: MIWUTES: SECONDSGH eocATTS NOT TOUAL PuLsES CS voreares rune on rust

5-108 ENGINE RESTART GOMUND, 17,6°4.68 IMITIATE 5/¢ SEPARATION HENEUVER, 11,630.33198 ECD INCERRUPT, START OF THE BASE 7, sonvbycow PHrSlcAL, SEPARATION, 11 667.317,689.53; Lox VENT ‘OPEM, 11630, 81: Uy ‘Tepowaty PITCH maveLver, 15,389MEAT BEN, 1° 630, TERMINATE ROLL MIMEUVER, 11,724Vila

ENT CLOSED, 11,640.28

Figure 11-42, S-IVB Roll Attitude Errors and RatesDuring Attempted Second Burn

11-57

APS engine chanber pressure was used to determine the times when theAPS modules ceased to function. At 21,953 seconds, the APS module at posi-tion I showed total fuel depletion, followed by oxidizer depletion at 22,053seconds. At 22,053 seconds, the yaw angular rate began to diverge due to aLOK vent initiated at 22,023 seconds. Fuel depletion occurred in the moduleat position IIT at 22,692 seconds; oxidizer depletion occurred at 22,634seconds. At this time the pitch and roll rates began to increase andreached anproximately -8.5 deg/s and 1.0 deg/s, respectively, at 22,800seconds. The vehicle continued to roll, resulting in osciliatina pitchand yaw angular rates. Deoletion of the APS propellants and subsequentJoss of attitude control is reasonable when considering the relativelylarge and unexpected demands on the control system, particularly followingS-IVB first cutoff and following the attempted restart and spacecratseparation. The large LOX mass remaining following spacecraft separationreduced the APS control moment arm which resulted in greater propellantconsumption for maneuvers and slosh induced disturbances. The attitudedynamics for this time period are shown in Figure 11-43, The offscalemeasurements in pitch, yaw, and roll shown in Figure 11-43 are due to LYDsoftware limitations,

Pitch and yaw attitude rates began to increase after depletion of pro-pellant in the APS module at position III as shown at the top of Figure11-44. Corresponding APS firing commands during the pronellant depletionin this module are shown in the center of Figure 11-44, $-LVB/IU tumblerazes obtained from reduction of radar tracking data for a period of ninedays following launch ate presented at the bottom of Figure 11-44, Therate buildup to 180 deg/s as indicated on the ninth day 7s attributed topropellant venting.

11.6 INSTRUMENT UNIT CONTROL CONPONENTS EVALUATION

11.6.1 Control-EDS Rate Gyros/Control Signal Processor Analyst

The analysis of the Control-EDS Rate Gyros/CSP indicated that the performance of this combination was normal. The highest detected ratesoccurred between §9 and 61 seconds and were excursions of apprex imately1B hertz which reached peak amplitudes of -3.5 and -5.1 deg/s in thepitch ond roll axes, respectively. Analysis incicates that the rateswitch filters in the CSP performed nominally.

11.6.2 Flight Control Computer Analysis

The FCC performed normally throughout boost and coast phases of flight.Analysis of the argular velocity and attitude error signals indicatedthat these signals, as telemetered from the FCC, were similar to thesame signals telemetered from the originating component.

The maximum FCC output current was 46.5 milliamps and occurred at 419seconds, as evidenced on the four pitch 50-milliamps servo amplifiertelemetry signals. This resulted from the premature cutoff of twoS-IT_ stage engines. This value represents 93 percent of the currentavailable from the FCC servo amplifiers

11-58

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DécREASING CHAMBERPRESSURE

<TTERO AFTER 22,655 SEC£2,000 22,100 27,200 27,300 22,470 22,500 22,600 22,700 22,200 22,900 23,000

RANSE TIME, SECGNDS

SES

FIRIRGS

6206743 6217-00 @15100 319200, Biz3: 20RANGE THME, HOURS:HINUTES: SEcotIDS

MMM Wro1CATES INOTVEDUAL PULSES? Loss OF APS FULL 1 MODULE 1 AT POSITION T, 21,953.TNDICAIES FULL ON PULSE [LOSS OF APS FUEL IN MODULE 2 AT PosiTION TiT, £2,602

* LOSS OF ST6AAL, HAMAIT+ ACQUISITION OF SIGNAL, CALTFORNIA

Figure 11-43. Pitch, Yaw, and Roll Dynamics at Loss of Control

11-59

+200we \wef 1

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20 4 —4 :L i

RANGE TIME, -ORY.

oo:o0s00 #8000 Seroa00 Tap DOS W9eRANGE TIME, OURS MTKUTES SECONDS

Figure 11-44, S-IVB/IU Tumble Rate History

11-60

SECTION 12

SEPARATION

12.1 SUMMARY

S-IC retro motor performance was satisfactory with negligible thrust im-balance in the pitch and yaw planes. The S-IC retro motor data indicatethat some parameters were either above normal or possibly above the maxi-mum limits but caused no problem, S-1C/S-II separation and associatedsequencing occurred as planned with adequate clearance between stages.

Available data indicated that the S-II ullage motors operated as expected.The loss of S-II main engines No. 2 and 3, with a consequent longer stageburn time, caused loss of ullage motor data on the post S-11 cutoff on-board tape recorder playback. Photographic coverage and simulation analy-sis showed that no problems were encountered during second plane separation.

The S-II retro motors performed as desired; minor deviations from nominalperformance in no way impaired their function. The S-I¥B ullage motorsperformed satisfactorily, with performance values very close to thosepredicted and within design specifications.

The S-I¥B separated from the S-II faster than predicted, primartly becauseof the 40 percent loss of S-II tailoff thrust. The loss of tailoff thrustresulted in a moment on the separating S-II stage. The pitch attitudeerror caused the S-IVB engine to be gimbaled 6.5 degrees during separa-tion; however, this resulted in no clearance or control problems.

Spacecraft separation was initiated by ground command to the spacecraftduring the maneuver to separation attitude. There were no Service ModulePropulsion System (SPS) engine bell clearance problems during spacecraftseparation; however, there may have been a momentary interference betweenthe Command and Service Modules (CSM) and the Spacecraft Lunar ModuleAdapter (SLA) panel at the separation plane. Any momentary interferencewas not detrimental to the separation. Therefore, initiation of space-craft separation during the maneuver to separation attitude did not re-sult in any significant problems.

A sunmary of separation events and times of occurrence is given inTable 12-

12-1

Table 12-1. Separation Event Times

ACTUAL TIME {SEC} PREDICTED TIME (SEC)

EVENT RANGE TIME RANGE TIMETIME BASE + TIME BASE +

LVOC Interrupt (S-IC OECO Sensed),Start of Time Base 3 {T3) 148,41 - 147.56 -

S-II Ullage Notor Fire Signal 148.87 0.45 148.06] 9.50

S-IC/S-1I Separation Command 149.08 0.67 148.26] 0.70

S-IC Retro Motor EBW Fire Signal 149.10 0.59 148.43 0.87

S-IS/S-II Physical Separation 149.14 0.73

S-I Engine Start Command 149.76 1.39 148.96 1.40

S-II Second Plane Separation Conmand} 179.06 30.68 178.26 30.70

J.VOC Interrupt (S-I] ECO Sensed),Start of Time Base 4 (Tq) 576.33 = 817.69] --

S-IVB Ullage Motor Burn TimeInitiation (Avg. of 2) 876.98 0.65 $18.39) 0.70

75 Percent Ullage Tarust 577.07 0.74

S-II/S-I¥B Separation Command 577.08 9.75 518.49 0.80

S-II Retro Mator Fire Command 577.08 9.75 518.49 0.80

10 Percent Retro Thrust 577.11 0.78

S-1I/S-IVB Physical Separation 877.13 0.80

90 Percent Retro Thrust 877.13 0.80

S-I¥B Engine Start SequenceCommand 577.28 0.95 518.69] 1.00

S-II/S-I¥B Separation Complete 578.07 1.74

S-I¥B £C0 Interrupt, Start ofTime Base 7 (T7) 11,620.33 = V,728.09| --

Spacecraft Separation Command 11,666.02 35.69

S-T¥B-IU/CSM Physical SeparationComplete 11,667.82 37.49 11,908.09 180,00

12-2

12,2 S-IC/S-I1 SEPARATION EVALUATION

12.2.1 S-IC Retro Motor Performance

Ignition signal to the retro motors occurred at 149.10 seconds. The per-formance of all S-IC retro motors was satisfactory, although a review ofthe data, shown in Table 12-2, indicates that some parameters were eitherabove normal or possibly above the maximum limit, Computed effectiveimpulse values based on combustion pressure data were apparently above themaximum limit shown in the model specification but caused no problemsThe average effective pressure and average effective thrust were generallyhigher than normat, although all were within the upper limit allowableaccording to the model specification. These higher than normal data arebeing given further study but are bevieved to be caused, at least partially,by differences between static and flight measurement conditions whichcould cause higher than "real" pressure records, If, on the other hand,the pressure transducers were reading actual chamber pressure, the retrorotors and the stage attachnent hardware were structurally adequate towithstand the higher loads. A test program has been initiated to comparethe chamber pressure transducer used in flight and <ts environment withthe transducer used and environment encountered during the qualificationtesting. The qualification testing established nominal performance and3 sigma limits.

Thrust imbalance in the pitch and yaw plane was negligible, Figure 12-1shows thrust versus time for the retro motor with the highest maximumthrust (fin D, position Lv) and for the retro motor with Towest maximumthrust (fin B, position IIL}.

12:2,2 SIT UVage Motor Performance

The S-IC onboard camera system which viewed the S-IC/S-1] separation pro-vided evidence that all four S-II ullage motors fired.

This evidence agrees with the thrust buildup observed from $-I1 telemetrydata. No additional information about performance can be determined fromtelemetry data, since the quality of S-II ullage motor chamber pressuredata was inadequate for detailed analysis because of telenetry attenuation.This attenuation was due to S-IC retro motor firing. The ullage motor data25 normally obtained fron onboard tape recorder playback after S-I1 cutoffDue to the longer than planned burn time, this taped information was notplayed back.

12.2.3 S-IC/S-II Separation Dynamics

S-IC/S-II separation and associated sequencing was accomplished as planned.Subsequent S-IC and $-I1 dynamics provided adequate positive clearancebetween the stages. The predicted and measured dynamic pressures at sepa-vation were 0.1123 and 0.1189 N/em2, (23.46 and 24.84 Ibf/#t2), respectively.Dynamic conditions at separation fell within estimated end conditions andwell within staging limits.

12-3

Table 12-2. $-IC Retro Motor Performance

PARAMETERvero woron [PRETag eppecnnee] yom] errecruve [ro cere

Peessune Mbutse uwpluse TaRUSTae Ment tes ae NGE, tester caves?! ces! dee?

Fink Post 3.09} 1161 209.935 266,02 398,959(1528) ius,ieo}

|

(58,586)

|

(69,692)ros 1 [o.see tige zaa,ao¢ 262,204 495,008

a7) tea,sa6) (58,466) (91.053)Fin #- Pos 11 |0.686 1196 2e2.373 262008 405,057

(738) 65,420) (58,366) (91,239)pos tit 6.668] 1163 z7a,ase 255,991 395,199

1887) 462,354) (57,349) (@e,é24)fin c- Pos mt 9.560 [rat ze.ai9 265,750 492,0581 a7en, ie7.tz7) (59,743) (90,888)

posty [user] 149 253,369 258.811 398,023(asery 3.04 Gaia 7.230}

Fin = Pos tv [0.a56| 1207 zes.te2 266,007 406,994078} wer83a) (60,512) (3t-496)

tos. 0.605 1173 zars21 57.005 398,765cet} (63.in)_ (57,377) (as,648)

Average oss | uve 200,090 201,905 ao0,2nano} 64,850) 58.882) 3,57)

Nomina? 298.1"GIF} Motor 0,633] 1120 ‘io Seve 297242 351,055

(1836) (35.605) (87,913)~3u Linge 277.9%Gor Hotor o.ss1e] 1009 No spec 242,172 352,426

oath (54,443) (79,209)7p Linty 321, 9|Gao Motor” “To.ssueed 1208 No spec |252,507 |440,096

(9362) {36,766) (98,938)

¥ at 221.9% (120'F)¥ At 271.9°K (30°F) Effective Burning Tine = The effective burnieg tine fs the Interval fron at~taiment of the initial 75 percent of maxinum pressure on the ascending por=Uon of the pressure trace to the sane level on the decay portion of the pressure trace.

Average Effective Pressure - The average effective pressure 4s the pressuretine integral betwaen the linits of effective burning tine divided by the ef=fective burning tine.* Total Impulse - Total impulse is the area under the thrust-tine trace fromzero tine untit the thrust returns to zero.Effective tnoulse - The effective smgulse is the area under tae thrust-tinecurve, between the limits of affective burning time.Average Effective Thrust - The average effective thrust 1s the effective impulse divided by the effective burnirg tine.

12-4

“0 : r 130

ate VSN aoe cf 15-590 5-501 4 [Rees fast ex famicnaae

; 1 1 avensse a5-sov 5f i 7 RETRC MOTCRS THRLS™it

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old oo oF oF nn On 1.0 12 ve

TOME FROM TOADTION, SECONDSTaa

RANGE TIME, SECONDS.

Figure 12-1, 5-IC Retro Motor Thrust

The first plane separation was monitored by accelerometers and rate gyroson each of the two stages, Separation rate transducers (extensometers )provided relative separation rate and distance data. In addition, motionpicture film provided a visual indication of the clearance between thetwo stages as they separated. For evaluation purposes, first plane sepa-ration dynamics were caiculated using a computer program which took intoaccount F-1 thrust decay, S-IC retro motor thrust, $-I1 ullege motorthrust, initial trajectory conditions, engine gimbal angles, and mass properties. The simulated first plane separation dynamics and separationdistances agreed very well with the actual data.

Figure 12-2 shows separation distances and relative velocities of the twostages and their respective contributions to the total. These velocitiesare changes in velocity magnitudes from time of physical separation, Theplot for separation distance also shows the point where the S-IC stageclears the J-2 engines, which extend beyond the separation plane by0.41 meter (16 in.). Very close agreement between the AS-501 and AS-502flights is seen. The separation time for the AS-802 flight was sTightlyJonger than for AS-501 because only four ullage motors were used on AS-502instead of the eight motors used on AS-501.

12-5

WY S-IC/S-11 SEPARATION COMMAND, 149.08 SECQV S-IC/S-11 PHYSICAL SEPARATION, 149.14 SES

I = 40

w RELATIVE

. al Ss 20 25 4 CONTRIBUTIONLig8 eer =4-- SBo 0 &

Eo S-1c po 85z Pd CONTRIBUTION --20.2 , Sq -30 =

“12 -40

57 A5-501 i re

48-502© CALCULATED .. hia7 4b - z8 / hiea as son 17 gz 3 \. RIO =

5 7 re os- / 5gs 2 ‘ 2s 4 A Fe 8a LZ 5 502 2= S-1C STAGE CLEARS: 55 J-2 ENGINE 5 4 3

LZ

2 eT 0149.0 149.2 149,149,618. 150.0PANGE TIME, SECONDS

Figure 12-2. S-IC/S-II Relative Velocity and Separation DistanceDuring First Plane Separation

12-6

Figure 12-3 shows lateral clearance and longitudinal accelarations forthe separation, The minimum clearance was calculated to be 1.33 meters(52.4 in.) between engine No. 1 and the S-IC stage. The longi tudinaacceleration indicated that physical separation occurred approximately0.1 second earlier than AS-501, This was due to the retro motor thrustrising to full thrust 0,1 second earlier. Good agreement between calcu-lated results and flight data existed for this flight.

S-IC angular dispersions during $-10/S-II separation are shown in Fig-ure 12-4, S-IC attitude deviations after separation were derived byintegrating the measured angular rate. Figure 12-5 presents the angulardispersions of the S-II stage during separation. No significant differenceexisted between the AS-501 and AS-502 flights.

12,3. S-[1 SECOND PLANE SEPARATION EVALUATION

Photographic coverage provided the only means of adequately monitoringsecond plane separation. However, the dynamics of both the second stageand the separating interstage were calculated using a computer. Thesecalculations utilized appropriate initial trajectory data, postflightmass characteristics, and J-2 engine plume characteristics obtained fromflight data, The only flight data from film analysis available were rela-tive velocity and relative displacement. All other data are catculatedresults,

The relative separation velocities, the relative velocity contributionof each body to the total, and relative separation distance between theta bodies are shown in Figure 12-6. Very good agreement is seen betweenAS-502 and AS-501 flight data. The velocities are the changes in veloc’~ties from time of physical separation and are calculated results. As wa:the case for the AS-501 Flight, better agreement in relative velocity databetween calculated and flight data was obtained by using an electricaldisconnect force of zera pound, The relative separation data also indi-cate very good agreement between AS-501 and AS-502 flight data. The sepa~ration was complete when the interstage passed the bottom of the J-2 en-gines and was calculated to have occurred at approximately 180.12 seconds.

Figure 12-7 presents the angular dispersions af the S-II stage duringseparation. Attitude errors remained near zero for both flights duringsecond plane separation.

The lateral clearance between the interstage and the engines was computedand is shown for each engine in Figure 12-8, The figure shows the lateralclearance, i.e., the clearance projected in the Y-2 plane, versus thebody station on the interstage at which the least distance occurred. Anarrow indicates direction of increasing time. There was a minimum clear-ance of 1.07 meters (42 in,) between engine No. 1 and the interstagerings at vehicle station 43,79 meters (1724 in.}. The separation planeis located at vehicle station 44.70 meters (1760 in.}. Figure 12-8 also pre-sents the calculated body rates of the separating interstage, which arecompared with those calculated for AS-501,

12-7

AXIAL SEPARATION DISTANCE, ft

¥ 500-100 058 9 2.S oiapd 1 i 1 i i gea TIME. AY POINT WHERE SEPARATION Suag ss WAS COMPLETE (0.41 m) pisog

25 LageBS 1s pase5 108 -0.4 0.30.2 -0) 0 O01 02 0.3 04 05 9

AXIAL SEPARATION OISTANCE, m

a° = ‘AS 502

—-— AS 501" 1 120© AS 502 CALCULATED

‘S-1C OECO SENSED BY L¥OC, 148.41

Ei S-I1 ULLAGE TRIGGER, 148.87 100

‘S-IC/S-I] SEPARATION COMMAND, 149.08

% ‘S-1C/S-11 PHYSICAL SEPARATION, 149.132. go SS$ 2> 20 >g bogE 5d Fag3 82 g2 f- 20 Z5 . Sell 3

50 of& / z

f-20

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Se1C-20 Z jo148.0 148.8 149.2 «1496 150.0

RANGE TIME, SECONDS

* THE DISTANCE BETWEEN SEPARATING S-IC STAGE AND THE J-2 ENGINE NOZZLE** AS-S01 IS COMPARED 10 AS-502 FROM S-IC OECO SENSED BY L¥OC

Figure 12-3, S-IC/S-I] Clearance Distance and Longitudinal AcceTerationDuring First Plane Separation

12-8

12-9

suid weoNyHOLS 315

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12-4,

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158188

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149.75

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OCK,

152-38

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90PE

RCEN

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RUST

.153.08

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han

dYaw

Angular

Dynamics

Following

S-IC/S-II

Separation

12-10

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502

-—~

4S501

wOudT 300.

}i

‘(dl 3S0N JALLISOd

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145

150

185

160

40145

150

155

160

RANGE

TIME,

SECONDS

RANGE

TIME,

SECONDS.

3A1L1S0d}

WVS-IC/S-I1

PHYSICAL

SEPARATION,

149.14

AYS-U1

HYDRAULIC

ACCUMULATORS,

UNLOCK,

157.38

WYS-II

ENGINE

START

COMMAND,

149.76

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90PERCENT

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153.08

Figure

12-5.

S-II

Angular

Dispersions

During

$-IC/S-IE

First

Plane

Separation

RELATIVE

VELOCITY,

m/s

RELA

TIVE

LONGITUDINAL

TRANSLATION,

m8

10

*

T T—— AS 502 (FIL DATA) 60-—+— AS 501 (FILM DATA).

© CALCULATED | , ow

| Ttfondjon

S-11 CONTRIBUTION

bo | (CALCULATED)t— + + t +t Q

INTERSTAGE CONTRIBUTION |{CALCULATED} |

¥ t LisAS 502 (FILM DATA)

>> AS 501 (FILM DATA) typ

© CALCULATED fiSV SECOND PLANE SEPARATION

COMMAND, 179.06 12

po sLeSEPARATION COMPLETE ?AT APPHOX 180.12 i1 ol Lio

| 4t Z rs

1 ot #4: L¢4

: LF L 4

i wei Here |. rz

Lat a179.2 179.4 179.6 179.8 180.0

RANGE TIME, SECOKDSFigure 12-6, Interstage/S-II Relative Velocity and Separation

Distance During Second Plane Separation

32-11

RELATIVE

VELOCITY,

ft/s

ITUDINAL

TRAN

SLAT

ION,

ft

RELATIVE

LO!

12-12

AS-502

so

AS-5OI

VSECOHD

PLANE

SEPARATION

COMMAND,

179.05

2

ii

ae

‘a

et

bap

Sep“(df JSON 3ATLISOd)OWNS JOLTLIV HOUTA

“(LHOTY 3S0N JALLISOd)YOULa FONTHWA

q

I

s/Gap

Lua AOE MYASS]

AS~501

&AS-502

30eoapt

“(dM 3SON JATLISOd)LV 408 HOLTa

-2. 17

0175

780

185

190

RANGE

TIME

,SE

COND

S

“(AHOIY ISON 3AILISOd)

2¥,

170

178

180

185

730

RANGE

TIME

,SECONDS

Figure

12-7.

S-II

Angular

Dispersions

During

Second

Plane

Separation

NEAREST POINT BODY STATION, ft

we a 5 lo 15 wwgua + 4 1 r 3Ze TIME ENGINE 4 gvSa 12 ENGINE 3 r* Syee paper] Bs£2 1.0 ENGINE. ? ‘7 z%& 0.8 ENGINE 2 5

0 7 2 3 4 5NEAREST POINT BODY STATION, m

= TtBS pps 502 (CALCULATED) —,Zu AS 501 (CALCULATED) —-—|

eeBE o 4Ee,e 4

BS> 3

=> Q -

ze |

BEEs 35 V9.2 Wed W796 1998 180.0

RANGE TIME, SECONDS

VY LOCATION OF INTERSTAGE RING, 4.05 m (13,3 ft)

‘W/POINT WHERE SEPARATION WAS COMPLETE, 4.98 m (16.3 ft)YSECONO PLANE SEPARATION COMMAND, 179.06

‘{YSEPARATION CONPLETE, 180.12

Figure 12-8. Lateral Clearance Distance and interstage Body Rate:During Second Plane Separation

12-13

12.4 S-11/S-IVB SEPARATION EVALUATION

12.4.1 S-IT Retro Motor Performance

The four retro motors mounted on the S-II stage performed satisfactorilyand separated the $-I] stage from the $-IVB stage. The pressure buildupfor all four retro motors began within 0.02 second of each other at577,08 seconds. The thrust and chamber pressure arofiles for the fourwotors were very sinilar, and the maximum difference in burn times was0.04 second. Prior to the flight a correlation between thrust and chamberpressure had been obtained from previous ground tests of similar retromotors. This correlation was used in a computer program to determine theretro motor torust from tae chamber pressure data during the AS-502 flight,

Table 12-3 presents the performance parameters for the individual motors.All parameters were within the nominal performance limits except for theburn time total impulse for motor A, which was slightly greater than thenominal maximum value but which had no detrimental effect on motor per-formance.

Table 12-3. $-II Retro Motor Performance

j SPECLEZCAT CON <1MITS,waToR aevavanecree . . : AT 288.9 °K (60 °F)pos tv-1 Joos ‘-1:a] eos cert pos tz reef nner

urn Time! see 86

|

ase

|

otsn 1.54 1.93 67 1.38Average Warn “ime Chanber

|

119g nso ive

|

206

|

ng) 1065.Pressure? W/cm? (psia) (asa;

|

ciate

|

cams)

|

cash

|

car) (154s)Maximum Thrust 34 (TSE) te1.487

|

178,370 7 172,815) 1 177.229) 152,129£40,800;

|

(39,236) | (39,300)| (43,700)

|

(40,000); ¢43,420;

|

432,209)average burn Time Tnruses

|

161,786 4285

|

,187,271] 152,520] 159,868] 138.292

|

175,416NODE (36,971)

|

(25,494)

|

(25,356)

|

(35,536)

|

(25,9393| (30.490)

f

(396835)BurnTime Total imputse?

|

750,768

|

239,986

|

237,477] 260,279/ 244,630| 250,435

|

232,897aes (Ibfes) (26.875)

|

(53,951) | (53,387)

|

(56,265)

|

(62.995)] (56.300)

|

(52,330) Bura Tine - Defined in Section €.2.1.30 af Thiokol Model Soect*leation TEMS-T1Burs Tine Average Charber Pressure - The average chanber pressure during burn tine is the areeunder the pressure tine curve over the buen time, divided 3} the burn tine.Naxinun Thrust - The Aighest thrust developed by the racket motor under any Hormel operatingcondition excluding iget tion.Burn Tine Average Thrust = The everege thrust during burn time 16 the burr time total inpuisedivioed by the Bure tine.Gurn Time Tetal Impulse - The ares under the thrust-tine curve over thé burs tine.

12-14

A tarust orofile for the retro motors is shown in Figure 12-9. The smooth-ness of te thrust traces for motor B and motor D is due to the fact thatthe chamber pressure data for these two motors were not good on the tapeused for the computer program and therefore had to be entared in tabularform from other telemetered data.

32.4.2 S-IVB Ullage Motors

‘JVage motor performance was satisfactory. The ullage motor ignitioncommand was given at 576.98 seconds, with the jettison command at589.08 seconds. These times, relative to engine start command, were veryclose to predicted. Table 12-4 presants the individual motor performanceparameters. A comparison of tnese data with nominal performance Timitindicates that voth motors performed within design specifications.Figure 12-10 presents the thrust profiles during burn.

12.4.3 S-I1/S-1¥B Separation dynamics

Separation of the $-IVB stage from tae S-TI stage was accomplished0.07 second faster than predicted time for a nominal separation (0.99 sec-ond). This was caused by lower than nominal $-II tailoff thrust levedue ta two engines out. Physical separation occurred 0.05 second afterthe separation command (established from extensometer data). Engines No, 2and 3 out resulted in a moment on the separation S-II stage which causeda positive S-II pitch rate and caused the critical separation point tobe on the position T side of the S-I¥B thrust structure. This moment re-duced the clearance distance to approximately 15.2 centimeters (6 in.)in the direction of positian I.

Table 12-1 contains significant times and events for the S-II/S-IVBseparation. Figure 12-11 presents the axtal separation history and rela-tive velocities between the two stages during S-II/S-IV8 separation. Alsoshown are the predicted and actual AS-S01 separatfon histories.

Figure 12-12 shows the longitudinal acceleration for the S-II and S-IVBstages. The reconstructed acceleration histories were obtained fromS-II and $-I¥B accelerometer data.

The angutar rates for both the S-IT and the S-IVB stages are presented‘in Figures 12-13 and 12-14. The S-II rates were all approximately zeroat physical separation. The S-I1 pitch rate increased to almost 2 deg/sby separation complete. ‘The yaw rate increased to 0.95 deg/s by separa-tion complete and the roll rate remained approximately zero throughoutseparation, The S-IYB rates were all small with pitch and yaw rates Tessthan £0.2 deg/s.

The path of the interstage lip during separation is shown in Figure 12-15,The closest approach point was a point on the S-IVB engine bell at posi-tion 1, The 7.6 degree pitch attitude error existing at S-I[/S-IV8 sepa-ration caused the $-I¥B engine to be gimbaled 6.5 degrees in the direction

12-18

Table 12-4. S-1¥B Ullage Motor Performance

SPECIFICATION LIMITSMOTOR A

|

MOTOR B

|

“aT Z94,1°K (70°F)

PARAMETER uNIT

|

(Pos 11

|

(Pos tv- II) =) MAXIMUM

|

MINIMUM—— —__—Burn Time! sec 3.81 3.80 4.0 3.64Average Burn Time

|

N/cm2 672 690 74 627Chamber Pressure?

|

{psia) (975) (1001) (1075) {910)

Maximum Thrust? 15,684 18,916

|

17,344 14,185N(bf) (3626) (3578) (3899) (3189)

Average Burn N 15,088

|

15.449

|

16,847

|

13,745Tine Thrust* (bt)

|

(3383)

|

(3473)

|

(3786)

|

(3090)Burn Time Nes 57,333

|

58,703

|

60,45)

|

55,603Total Impuise® {ibf-s) (12,889)

|

(13.197)

|

(13,590) (12,500) } Burn Time - Time beginning when the pressure has risen to 10 percent ofthe maximum chamber pressure and ending when the pressure has drappedto 78 percent of the maximum chamber pressure.2 Average Burn Time Chamber Pressure - The area under the pressure-timecurve during burn time divided by burn time.3 Maximum Thrust - The highest thrust developed by the rocket motor underany normal operating condi tfon excluding the first 0.20 seconds ofoperation.

* Average Burn Time Thrust - The burn time total impulse divided by theburn time,5 Burn, Time Total Impulse - The integral of thrust with respect to timebetween the burn time limits.

of position I during separation. The gimbaled engine and S-11 momentresulted in a total clearance of approximately 1.60 meters (63 in.) outof @ total available clearance of 2.06 meters (81 in.), presenting noclearance or control problems.

12-16

5g

MOTO

RTHRUST,

103

#Last

250—

%0200 L-—T voragrm woroR ow

pre SO ETE to150 |i

woroR ¢ oh100 Ty 1s-1t

\ 20

“ [= ‘° r=

|

SIC F6E871.0_ S778 78D 3B S80 S72-E EOD.0ANGE TIME, SECONFigure 12-9. $-II Retro Notor Thrust

2»

8 40|aun yer

6as

m7 Ftv Itt20

le28

10

208

1s‘

10’

2 08

oO [meet aV0 0 40 2.0 3040 5.06.0 7.0 Bo 83 10.TEMFROM ULLAGE MOTOR IGNITION, SECONDS

Bree san 3 aARIGE TIME, SECONDSFigure 12-10. $-1VB Ullage Motor Thrust

W2-17

E780

MOTOR

TIWRL

UAGEMOTOR

THRU

ST,

1031b

“ _|5Eo ,wgee SEant

AS-15 EoD senseoeF eaeLec, $76.33 8g 05-11/5-14SEPARATION 232

Y stadt. 87708 oLtS-1¥8 ENGINE eTAET 0Seovence conmano, ¥ Tan28 zgw

EE9 30

gee?

ESS 10Bae

4 wy g 570 380 590600

RANGE TINE, SECONDSFigure 12-13. S-IL Angular Dispersions Ouring S-II/S-I¥B Separation

eeas

38 settge, >

seuyste mystea EB S15Saath88 é,‘S-1¥B ENGINE START ~SEMENCOMBO 577.28 wy

WY 5-15/5-1¥8 SEPARATIONfelEae seo — on|

S-1v8

Ley vya7 Ore 378.0 Tae 379.0

RANGE TIME, SECOHIS

YAWANGULAR

RATE

(POSITIVE

NOSE

RIGHT)

dea/s

Figure 12-14. $-II and S-IVB Angular Dispersions During

S-IL/S-IVB Separation

12-20

2,5 S-IVB/IU/SPACECRAFT SEPARATION EVALUATION

Tz was established at 11,630.33 seconds. At Ty +20 secands the prepre-grammed maneuvers to separation attitude were initiated. However, at71,866.02 seconds spacecraft separation was initiated ty ground comandto the spacecraft.

The $-1VB attitude errors and rates during S-IVB/It/Spacecraft separationare shown in Figures 11-40, 11-41, and 11-42 for pitch, yan, and roll,respectively. The S-IVB pitch, yaw, and roll attitude errors at space-craft separation were -7.0, +6.5, and -7.0 degrees, respectively. TheS-IVB pitch, yaw, and roll angular rates during spacecraft seperationwere #0.4, -0.8, and +1.3 deq/s.

Telemetry from the spacecraft and the S-1VB/IU indicated that unexpecteddisturbances were applied to toth vehicles imediately following physicalseparation. The spacecraft pitch rate increased from 0,3 deg/s ta 1.83 deg/:in the nose-up direction over a period of 0.1 second following physicaseparation, The 1.53 deg/s rate change may be attributed to a momentaryinterference between the Spacecraft Lunar Vodule Adapter (SLA) panel locatedat position I ard the spacecraft at the separation plane. The S~[V8/IUpitch rate decreased fram 0.38 deg/s to 0.1 deg/s nose-up pitch rate in thethe same time interval. This rate change may be attributed to the abovemomentary interference between the vehicles coupled with asynmetrical forcesand moments resulting from: delayed deployment of the SLA panel locatedat position I. The 5-IVB/IU pitch rate increased from 0.1 deg/s to 0.4 deg/snose-ve 1,0 second after physical separation. This may result from thedeceleration of the SLA panel at position III on hitting the deploynentposition stop prior to the SLA panel at position I. The SLA panelsreach an angular rate of approximately 40 deg/s within 0.09 secondof physical separation and are fully deployed by 1.3 seconds, Longi-tudinal acceleration impulses in the aft direction were detected by theIU accelerometer inmediately following physical separation and 1.0 secondlater.

Spacecraft attitude rates and linear acceleration data provided by MSCwere utilized to reconstruct the relative motion of the service moduleengine bell with respect to the $-I¥B/IU during spacecraft separation.The resultant relative motion as a function of time from physical sepa-ration ¥s shown in Figure 12-16.

Even though there were no SPS engine bell clearance problems there mayhave been a rorentary interference between the CSM and a SLA panel at theSeparation plane as discussed above. However, the momentary interferencewas not detrimental to the separation. Therefore, initiation of spacecraftseparation during the maneuver to separation attitude did not result inany significant problems.

12-21

12-22

ATHOF

|etteStaceue

b-15.2 em th tn.)Figure

12-15,

S-11/S-I¥B

Sepa

rati

onClearance

bes “yossvers3s

‘ates Ye

Figure

12-16.

Relative

Notion

During

Spacecraft

Separation

SECTION 13

ELECTRICAL NETWORKS

13,1 SUMMARY

Each stage has its own electrical system which includes the batterypower supply, power distribution systems, switch selector, etc. Thestage electrical system supplies and distributes power to all electrical/electronic equipment on the stage. Details of stage electrical systemsare presented in the individual stage descriptions

In general, launch vehicle electrical systems performed satisfactorilyduring powered flight. Stage deviations are detailed in the individuastage descriptions.

S-IC stage deviations included:

a. Battery No. 2 voltage drop at about 158 seconds which continuedfor 11 seconds,

b. Power supply ‘bus 1086 voltage drop From a nominal 5 vde to zero vdelasting from 160 seconds until loss of telemetry signal

S-II stage deviations included:

a. A 34-ampere current spike at 412.7 seconds on the main battery.

b. Instrumentation battery positive currént spikes of 47 and37 amperes with corresponding voltage drops of 1.2 and 0.5volts beginning at 414.2 seconds.

The only S-IVB stage electrical deviation was that the PU static invertersconverter 5 vdc exceeded the upper limit of 5.1 vde by fram 30 to 90millivolts; however, no adverse effects resulted from this deviation

Instrument Init deviations included:

a. Acurrent surge of 9.5 amperes, lasting 400 milliseconds, at133.3 seconds on the 6D17 bus, with a corresponding voltagedrop of 1.8 volts in the 6010 battery,

b. A current step of 1 ampere at 22,112.4 seconds on the 6011 bus.

13-1

13,2 S-IC STAGE ELECTRICAL SYSTEM

The S-IC stage electrical system is composed of five 28 vde batterieswhich supply the various systems and components through the powerdistribution system as shown in Figure 13-1. The system includesseven 5 vde instrumentation measuring power supplies, distributors,and a switch selector.

The electrical system performed satisfactorily during $-IC poweredflight. Battery voltages and currents are shown in Figures 13-2 and13-3, Both battery No. 1 and battery No. 2 stayed within design limitsof 26.5 to 32 vdc (as read on buses ID10 and 1D20, respectively}, andbelow 64 amperes for battery No. 1 and 125 amperes for battery No, 2(as read at each battery), until approximately 168 seconds. At thistime, approximately 19 seconds after S-IC/S-I1 stage separation,battery No. 2 voltage dropped to 25.8 vdc and the current rose above175 amperes, the upper limit of range on the ammeter. After 11 seconds

000 oe

[eS Be |ane Unc PANG SOUND GATTERY eaTTEREary a | mT |

T_Tt J r |a] |aane = |

prey wanepeon men, |Lomos ons CONTRA | ‘SEPRAATIOO LIGHTSaman=, asa| |ae ae =a Bes necweg, ate 82esage Se tagiome Fes iar srnvrie rerShow Eu, on L Jee neeaie, ®STEN aaa

Breve Figure 13-1, SIC Power Generation and Distribution Systems

Block Diagram

13-2

* PPER LIMIT BATTEPY VOLTEEL 9 a ~T

ir eee Ioarbore ==LLLLEEEEELy ibe Leg

2 ASB) FLIGNT DATA D

; f ae HULApf: 246# Lum sarreR vouzanz i i

“ T T Tsscsor iar oara |!a0 RS i

5 +2 as-508a FLIGHT ATA — 4

woh neste =

oh erencesate te eaa

aioe TIME, sconsFigure 13-2. S-IC Stage Voltage (on Bus 1010) and Current

(at the Battery)WePER LEM; wae vaTAGE

TESBC) LILA DE

vouts,

de

Peter OAT

LOWER LIM!T BP1VERF VOLTAGE

PREDICTED, Son FIG ORTH

CURRENT,

ams

$-b01 LIGHT Oar

a o Es ny ten ToANGE TINE, SECONDS

Figure 13-3. S-IC Stage Voltage (on Bus 1020) and Current(at the Battery)

1343

the battery voltage and amperage returned to normal. A somewhat similarsituation occurred on AS-501 S-IC battery No. 1 immediately after -S-IC/S-II separation. Tape recorder performance was not affected sincethe voltage remained above the minimum required by the tape recorder,Investigation is underway to determine probable cause of this deviation,and corrective action to be taken,S-IC stage batteries No. 1 and 2 outputs in ampere-minutes and as apercent of rated capacity were well within desicn limits, as shown inTable 13-1, Batteries No. 3, 4, and 5 were not instrumented to givethis information.

The seven 5 vde measuring power supply voltages for instrumentationvaried from 4.96 to 5.04 véc curing powered Flight, This was withindesign parameters of 5 30,05 vde. At approximately 160 seconds,11 seconds efter S-IC/S-II separation, measuring voltage power supplybus 1D86 dropped from a nominal 5 vdc to almost zero vdc and remainedat this level until loss of telemetry signal. This deviation is beinginvestigated,

All S-IC switch selecter channels functioned as commanded by theInstrument Unit.

Separation and retro motor Exploding Bridgewire (£BH) firing units werearmed and triggered. Charging time and voltage characteristics of theEBM firing units were within design specifications. Separation andretro motor ignition charging time and voltage characteristics werewithin required parameters.

Table 13-1, S-IC Stage Etectrical System Battery PerformanceDuring Flight

cousin7104 |CAPACITY* AMP-MINS:BATTERY 4(AMP-MINS) yay Tom ACT Ja

expectcn |"+t J8~

Battery No. 1 (Qperational] 640 aa 32.00% 5.0Battery No. 2 1250 384.2 a04.gree

|

31.6(instrumentation)

Batteries No. 3, 4, 5 1250 Not instrumented to give(Optical Instrumentation) 1250 sonsumptton data

640

* Aupere-minutes ratings are approximate values.** 0 to S-IC/S-11 separation,*** 0 to 210 seconds.

13-4

13.3 S-EI STAGE ELECTRICAL SYSTEM

The S-I1 stage electrical system contains four 28 vde batteries, twoof which are connected in series to furnish 56 vdc to the recircula-tion inverters, The batteries furnish power to the various stagesystems and comoonents through the power distribution system, ashown in Figure 13-4, Five 5 vdc power supplies furnish measuringvoltage to the telemetry and other instrumentation. The five Lzrecirculation pump inverters convert 56 yde to 42 vac, 3-phase, 400 hertzpower for the ac induction motors on the LH recirculation pumps. Thesystem includes various controllers and a stage selector switch

The S-II electrical power system performed satisfactorily throughout allphases of the AS-502 flight. Operation of the batteries, power transferswitches, and LH2 recirculation inverters was normal.

All bus voltages stayed within specified limits during the prelaunch andflight periads. Voltage and current profiles for the main bus are shownin Figure 13-5. Main bus current stayed within specified limits except

zee pee

ast NcTaTI On aegtacu.arion PEcIRCaLeTTo4barter anrcesr SETTER EC. | BATTERY 0.2

nsreunenr2 stom rat tones RECIRCULATIONws (eee bus (2311) us (2061) U5 (208!)

BoC 2avoe evn 500atta 3-2 ENE IeWITION 5 Leg RECIRCULATIONusr aUMCHTT Fa PRESS RAZATION PoMeTWERTERS

TELEMETRY PROPELLANTFIVE syGc ewe MANAGEMENTSaPOLiFS SEPARATIONCAMERA SYSTEM 5-2 ENGINE CONTROLRATE. GYRO HEATERS CowhIND CESTRLCTcommun BesT=ucT wesel ELECTALCAL,EMERGENCY SEQUENCE COVTROCTIS¥STEM eseRGeNe

[ONE OF THO} BERET EG susreM(One oF 140]

Figure 13-4. 5-11 Power Generation and Distribution SystemsBlock Diagram

13-5

for ar instantaneous current spike of 34 amperes at 412.7 seconds, ap-Proximately the time given for J-2 engine No. 2 cutoff. Since therewas no significant change in bus voltage at this time, the soike wasprobably caused by one of the following:a. Awild data spike, indicating false data.

b. Instantaneous short in the stage wiring or bus loads. The possibilityof wire damage in this area is indicated by voltage spikes on othermeasurements and loss of measurements routed by vehicle station 42.57meters (1676 in.) and stringer 108.

Instrumentation bus voltage and current are shown in Figure 13-6,Instrumentation bus current stayed within specified limits except fora 47-anpere positive spike at 414,2 seconds which lasted for 1.2 secondsand a subsequent 37-anpere spike of shorter duration, with correspondingvoltaga decreases of 1.2 and 0.5 volts. The time of this deviation corresponded approximately with cutoff of J-2 engine No. 3, The cause wasprobably shorting of distribution wires in the region of vehicle station42,57 meters (1676 in.) and stringer 108, where loss of measurements indicatesdamage to wire cables.

Voltages and currents profiles for recirculation and ignition busses arepresented in Figures 13-7 and 13-8. Recirculation and ignition buscurrent stayed within specified limits.Battery temperatures (see Table 13-2) remained well within the predictedrange. S-II stage batteries consumption in ampere-hours and as a percentof rated capacity are given in Table 13-2. All four battery ampere-houroutputs were well within design Timits. The Five 5 vde power suppliesprovided proper measuring voltage to the telemetry and other instrumentation.

Table 13-2. S-I] Stage Battery Consumption:

BATTERY oesranarion

|

creacrty™| consumPrion

|

rencent [__TENPERATURE(REFERENCE)

|

(RNP-HR)

|

{AMP~HR)

|

CONSMER Fay AIN

Main an 35 8.83 244 310°k 304°K38°F)

|Gaer)

Instrumentation 2021 3B 10.60 30.3 311K 4303.5°KGover)|

Grrr)Rectrewiation fo. 1

|

2951 % 8.17 14.8

|

302%

|

300°k(5°F)|

Goer)Recirculation No, 2 2051 35 6.21 14.9 301°K |298,5°%

and @3eF)

|

Garr)6

* Anpere-hour ratings are approximate values.

13-6

ae HE

Lowes LIMops ¥OLzaGe

Ey

‘cvual

‘une

aT,ane

s

af zl"1060 -a "TAO 145 150155 Yao NO ——aea TaN 42 x9 SIO "S730abe THPE, SECOSns

Figure 13-5, S-11 Stage Main DC Bus Yoltage and Current

UPPER Lindy aud woaFace

2

LONER LIMIT BUS. WTS]

sarorcten

no«a

acrUALolsh asa 15 YAS 1e0 155 Teo Yeon mg ——-ez0 "49950580 60‘soe Time, stconas

Figure 13-6. 5-II Stage Instrumentation Bus Voltage and Current

13-7

“ TTT] %“ +

8 LM ws Yua of

DLLELT.Yi Y

w0LI8,

ac

Fy

”

Fy 7 +Lose LI 8:5 «

2a PET

semuaoo

URRERT

,aps

on

bea) <0 DGD 1M

BANE TIME, sEvaNESFigure 13-7. S-II Stage Recirculation DC Bus Voltage and Current

2 rT TtUPPER LIMIT BUS YOLTAGEa J

0 Os

a

ASTUAL

8

VOLTS,

ve

a 26 &

‘\ ose LITT @us WoL TAGE2a it80 4b 20 D140 145 sD 18S SO 16S 70178180RANGE TIME, SEvOWDS

Figure 13-8. S-IT Stage Ignition DC Voltage

13-8

Performance of the switch selector was satisfactory during the flight.The switch selector tclenetry output measurement was not available duringportions of the flight due to a malfunction in the measurement; however,events indicated that the switch selector correctly sequenced the S-Istage as commanded by the IU computer. The LH recirculation inverter:operated satisfactorly within acceptable limits during the J-2 enginechiTldown period.

Performance of the eiéctrical portion of the separation system was

satisfactory during the flight. EBW firing units charge and discharge

responses were within the predicted tine and voltage vimits.

13.4 S-IVB STAGE ELECTRICAL SYSTEM

The $-1VB stage electrical system is composed of three 28-vde and one 56-vde batteries which supply the various stage systems and components througthe power distribution system, as shown in Figure 13-9, Two S-vdc excitation modules furnish measuring voltage to instrumentation measurement trans~ducers and signal conditioners, Seventeen 20-yolt excitation modules supplysignal conditioning power for event measurements, excitation power for temperature and voltage measurements, and amplifier power to various measurement:that use a de amplifier. The static inverter/converter converts 28 vde from

zo zoe zene sencerons Toque a asate. onEo. 2 airtel wa. arro,

"om BU 9 ronan aie wa. E a wi wt arrBuswe

Taito Be Tete weesree rwirerarcmearee |a-2 nce suacugar soonaso means tts Noe teeeecore CCM sreren

|faamutan meuston|

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‘uo eee Ear le sare tg cr4co0mse exer fine gacraran are itheEA ita pace seensot eet108 essaste hicas sreeesa errson

Figure 13-9. S-IVB Power Generation and Distribution SystemsBlock Diagram

13-9

forward battery No. 2 to 116 and 2 vac, 400 hertz, single phase, andto 5, 22, and 117 vde power for the propellant utilization system. TheLOX and Lz chilldown inverters convert 56 vde from aft battery Yo. 2to &6 vac 400 hertz for the chilldown pump motors. Power is routedthrough various distributors and a switch selector.

The $-1¥B stage electrical system performed satisfactort ly during Firstburn and through restart attempt. Al] systems responded normally to theInstrument Unit commands.

Battery voltages, currents, and temperatures are presented in Figures13-10 through 13-13, Battery temperatures stayed below the 347°K (165°F)limits for the powered portion of the flight (the 347°K limit does notapply after insertion into orbit). The highest temperature reached was341° (154°F) on forward battery 1, unit 2, during the sixth revolution,as observed on PCM data from Hawaii (see Figure 13-10). All battery ten-peratures were within their specified limits through the sixth revolutionThe output voltage of forward battery No. 1 was 18 vdc and decreasingaccording to sixth revolution Hawaii data. The other three batteryvoltages were nominal through the sixth revolution, $-IVB stage batteryconsumption in ampere-hours and as a percent of rated capacity are givenjn Table 13-3. Forward battery No. 1 furnished 6 percent more anpere-hours than the battery was rated for. The other three battery ampere-hour outputs were well within design limits.The two 5-vde excitation moduies provided proper excitation voltage at3,40,03 vde to the transducers through fifth revolution Hawadi data,(27,680 seconds). The sixth revolution Hawaii data (33,380 seconds }showed" both voltages offscale high at 5.50 vde. However, this isattributed to a signal conditioning drift rather than an actual cut-of-range condition. The 17 20-vde excitation modules performed satisfactorily.The switch selector decoded the Instrument Unit signals property andactivated the desired relays, valves, etc. at the proper times, throughthe sequencer. The LOX and Lp chilidown inverters performed satisfactorilyand met their toad requirements, The 5 véc PU static inverter/converterexceeded the upper limit of §.1 vde and ranged from 5.13 to 5.19 vde.However, no degradation of mass calculation occurred since ratios ofvoltage levels are utilized in the calculations.

All EBW firing units responded as expected to their respective commands.The ullage motor ignition £BW firing units were charged at 482.88 secondsand fired at 576.98 seconds. The ullage motor jettison EBM firing unitswere charged at 586.09 seconds and fired at 589.08 seconds, to jettisonboth ullage motors. The Secure Range Safety EBN firing units were notcharged or fired, since they were not required for AS-502.

13-10

132-11

UPPER

LIMI

TBATTERY

VOLTAGE

we

WY,UE

16

22 30

iff]

ty

2

P SLIUA

.LO

WER

L:HIT

BATT

ERY

VOLT

AGE

ren

me

crue|

tt

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oS

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ETI

ME,

102

SECO

NDS

50

0

sdue + 113Ean9

0 z

WwL1

1L

OD:16:40°00:69:00

0:28:20

a2%21:40

wa:20200

05:0

0:00

,06:28:29

7:46

:40

RANG

ETIME, M

OLRS: M

T UTES

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11,8

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TIME,

103

SECONDS

ca

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04-2

680

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—0ar

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NGE

TIME

HOURS

:MINUT

ESSECONDS.

Figu

re13

-10.

S-1V

BStage

Forw

ard

Batt

ery

No.

1Voltage,

Current,

and

Temp

erat

ure

31

9

Ao “RUNYERMAL

MIdhIA

TAL

5 <== precicreD2 . r1 ieee (init eae Wot| PT3 hy Ye

AG Li> 1 LOWE LIMIT BATTERY ¥OLTAGE : 1ali ee Bare Lr 7

' : 4t iapo ' i i ul7 ; 7 T 17

a [i 1 4 i 'B |a ' ' i '> fr l tel5 Fa4ep toea it8 1 ' ifts 1 \ : ft8 ! 1 ' ‘ft

ay Yas) 660 708 Feh VT S80 TseMTSRANGE TIME, SECONDS

00:00:40 00:11:00 90:13:00 03:12:30 03:14:10 09:10:0000:09:10 00:12:06RANGE TIME .NGURS :MINUTES:SECONDS

220 1 |Fas — 150

#0 pf ee Ee3

aos

BZ me& 305 t =z \ r zgF309 oe295 10° + % We ap

RANGE TIME,107 scconos.

bas92:13:20 04:26:40 06:40:08 08:53:20

RANGE T(E HOURS: MERUTES

:

SECONDSY TRANSFER TO INTERNAL POWER ‘VPs VALVE HARODVER POSITION OFF 11,627.59Y pu activate on 595.30 PU ACTIVATE ON 11627.50

Pu AcTIvaT 7. ‘© pu ACTIVATE OFF 11,631.70vy © OFF 48.15 PLEINVERTER AND OC POWER OFF 11,641.80‘VW FU VALVE HARNOVER POSITION ON 11,574.69

Figure 13-11, S-IVB Stage Forward Battery No, 2 Voltage, Currert,and Temperature

13-12

13-13

Siar * LRM Yo IRNL

Tere

Linas

BATTER

YWoU

TaBE

UU

Yes

LOWER,

Li<T

BATTERY

YOUTAGE7

i

oOo1e320

Dari0s00

LL

1J

05:00:00

06740509

06:20:30

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0:c0

PRNG

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¥INJTES:

SECONDS.

40:0

0

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S-1¥0

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Solve

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ULLAGE

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SECCNDS

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TIME;H

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Figu

re13-12,

$-IVB

Stag

eAf

tBattery

No.

1Yoltage,

Current

and

Temperature

——~ ACTUAL, WITTPREDICTED nr 2

7066

¥ UPPER LIMIT BATTERY, VOL TAGE:"62

a 5a

* 8450

i LOWER LIMIT BATTERY VoL TAGE 4106 .

1i& T* 6c

ge 40

r

SS PES SET 12RANGE TIME. 107 SECONDS

00:16:40 00:50:00 0:28:20 02:35:00 03:20:0¢RANGE TIME HOURS :MIHUTES: SECONDS

= 385 140u

gf 120gz ----- 100 ¢& oe §= wt] foot

5 10 1s 26 5 30 35RANGE TIME, 103 SECONDS

02:46:40 05:35:20 08:20:00RANGE TIME, HOURS

:

MINUTES :SECONDSVY TRANSFER TO INTERNAL POWER Y seve esc, 11,614.68Y 5-1ve ESC 577,28 FUEL CO PUMP OFF, 11,613.90

LOX Co PUMP OFF, 576,88 LOX CO PUMP OFF," 17,614.09FUEL CD PUMP OFF, 578.48 ‘Y Aux HYD PUMP OFF, 11,634.18

‘VY AUX HYD PUMP OFF, 751.37

Figure 33-13. $-IVE Stage Aft Battery No, 2 Voltage, Currentand Temperature

13-14

Table 13-3. $-IV3 Stage Battery Consumption

wap CONSUMP 108;ACITY CAMP-HRS} PERCENTBATTERY 1

| AMP-HRS)"Taxtwun [ACTUAL HRg] USAGEEXPECTED? 6.53 9.34]

=Fud No. 1 350 279 268 370) 106

Fwd No. 2 25 13 13.4] 43.1] 52.4

Aft Ho. 1 300 55 63| 137] 45.7

Aft No. 2 78 a4 20,6 20.6] 26.4

NOTES: 1. Ampere-hour ratings are approximate values2. Predicted usage tased cn maximum expected values for 6.5

hour flight.3, Actual usage for 6.6 hours based on available flight data

Total usage through sixth revolution (9.3 hours)

13.5 INSTRUMENT UNIT ELECTRICAL SYSTEM

The Instrument Unit electrical system includes four 28-vde batterteswhich supply the various IU systems and components through the powerdistribution system, as shown in Figure 13-14. A S-vde power supplyconverts 28 vde from an auxiliary power distributor to closely requlated5 vde for use as a signal conditioning reference voltage and also tosupply various IU transducers. The 56-volt power supply converts 28vde from battery 6D10 to 56 vde for the ST-124-M gyro and accelerometerservoloaps and for the accelerometer signal condittoner. The platformac power supply converts 28 vde from battery 6D10 to 26 vac, 400 hertz,3 phase and 20 vac, 4.8 kilohertz for the ST-124-M gyro, and to 20 vacat both 1.92 kilohertz and 1.6 kilohertz for electrical support equip-ment. Power is routed through six distributors. The switch selectoractivates the circuits necessary to execute commands received.

IU battery voltages, currents, and temperatures are presented’ inFigures 13-15 through 13-18. In general the IU electrical systemperformed satisfactorily during the flight. The only major differencefrom AS-501 data was the 133 second anomaly mentioned below.

The 6D10 battery performed as predicted except for a current spike at133.3 seconds and a current step at 22,112.4 seconds. The 133.3 second

anomaly involved 2 400 millisecond current increase of 9.5 amperes with

13-15

gent emer ee

a a we qoofan] | [ae | Lad)

ate |

nessun:easuens oes nLsuraR

cae Figure 13-14. Instrument Unit Power Generation and Distribution

Systems Block Diagram

@ corresponding 1.5-vde voltage drop, as shown in Figure 13-19. SeeSection 9A, °133 Second Transient," for a discussion of this anomaly andits implications. The 22,112.4 second current step of 1 ampere resultedfrom the gimbal torquer motor trying to drive the yaw gimbal off the 60degree mechanical stops, where it had hit due to loss of effectivenessof one auxiliary propulsion system medule in combination with the openingof the LOX vents

The 6D20, 6030 and 6040 battery performance was normal, although 6030current was somehwat higher than predicted,

All battery temperatures were within normal limits. The maximum observed‘temperature was 321°K (118.4°F) on the 6030 battery at 9.4 hours,Instrument Unit battery consumption in ampere-hours and as a percent ofrated capacity is given in Table 13-4. At the time data were lost, battery6030 had furnished 1 ampere-hour more than the battery was rated for, andthe other three battery ampere-hour outputs were well within rated limits.The 5-vde power supply operation was satisfactory. The 56 volt powersupply operation was nominal except for the effect of the 6010 battery

13-16

ER LIME SETTER WLTARE

=INCT gaT-ER¥ yo_Tece

enn "||mnt

ero\| Ae tee

nm TT St >itcnet— . : : we

- i | : i —ifr08Vie sas er SS NRTWES aeaewee

BANE TAME, SECONDS 4 123

ate wa rary m8 RANGE TCMESHEURS MIM ES :SLCONUS

Figure 13-15, IU Battery 6D10 Voltage, Current, and Temperature

2 Tinearibae waagai2a2 afge Lose LIMIT EATERY vtrage

fe potgy hi Zag Prevrcten| + eeos Lue(ae|zo CSBas mE |

20 pmfai : a4 :gas fo i | : pet2eaoaf j ee ify | 90S5 :Ep te | I | 10g

fal | L ak sofotaeieneneeaefuer THRE, SecOmDs = 108

cooo: 1640 aan Gera30 OFsaa 26:49ANGE TIME-HOURS MINUTES: SECONDS

Figure 13-16, {U-Battery 6D20 Voltage, Current, and Temperature

13-17

J Pit uateELsal2 2 LUA“al TOWER LIMIT ATT ERY VOLTAGE:

a ACTUAL:

PREDECTEDZ

iJ —12 3 2S ESS eT ase as ieeas 20 1 DNETT ARENCE T:ME, SECONDS 4 103

mo:t6e10 D505 eo Wrserd Ga: Ze ecROSE 12H HOURS MI WITES SECONDS

Figure 13-17. IU Battery 6030 Voltage, Current, and TemperatureRx 1] WeheR LIT BATTERY VOLTAGE

WLS,

de

LEER LUMIeALTERY VOUT

erence |i2

gm pegSan oeBe 70Basa 5reed 1123 FS 67 8 9 wo 1s Ws 1s i617 18 18 20 Bae 2°27 28°93 rRaRGE TAME, SECONDS x 199

00: 1640 0520 ceoeaoYores17:50:00ANGE TIME=HOURS MINUTES: SECONDSFigure 13-18. 10 Battery 6D40 Voltage, Current, and Temperature

13-18

spike already discussed (see Figure 13-19). The effect of the 6010battery spike was also felt on the platfarm ac power supply (seeSection 9A).

There is no indication of any discrepancy in distributor performance.

The switch selector functioned in a satisfactory manner throughout theflight. Al] commands to the switch selector were received properly andno complement commands were necessary.

Table 13-4, Instrument Unit Battery Consumption

TARE 10K

CAPACITY * ANP-HRS}) PERCENTBATTERY(AMP-HRS) Thaxtwss Tartuxrs] USAGE

expected [6.5 |9.4

epia 350 NA wa

|

295] 8a

6020 asc NA na

|

310

|

88.5

6030 360 NA wa} 361

|

100

6040 350 NA na

|

205] 87.2 *— Based on 350 ampere-hours at a 35 ampere discharge rate.** Total usage through 9.4 hours.

wrt Based on the decrease in terminal voltage.

13-19

57.0

56 VOLT POWERSUPPLY

VOLTS,

de

ooh V\ VAAL AN

55.5

VOLTS,

dc

_RK

io

CURRENT,

amps

+6010Nt| 30

132 133 134 135RANGE TIME, SECONDS

Figure 13-19. Currents and Voltages Associated with 133.3Second Transient

13-20

SECTION 14

RANGE SAFETY AND COMMAND SYSTEMS

14,1 SUMMARY

Data indicated that the redundant Secure Range Safety Cormand Systems(SRSCS) on the S-IC, S-[I, and S-IVB stages were ready to perform: theirfunctions properly an command if flight conditions during the launchphase had required vehicle destruct. The system properly safed theS-I¥B SRSCS an comand from Kennedy Space Center (KSC). The performanceof the command and communications system in the Instrument Unit (IU) wasvery good.

14.2 RANGE SAFETY COMMAND SYSTEMS

The SRSCS provides a means to terminate the flight of the vehicle byradio command from the ground in case of emergency situations inaccordance with range safety requirements. After successful insertioninto earth orbit, the system is deactivated (safed) by ground command.Each powered stage of the vehicle was equipped with two commandreceivers/decoders and necessary antennas, The SRSCS in each stage wascompletely independent of those in other stages.

‘Three types of SRSCS commands were required for this unmanned flight asfollows:

a. Arm/fuel cutoff - Charging of the Exploding Bridge Wire (EBH)firing unit and thrust termination.

b. Destruct - Propellant dispersion by firing of the EBW.

c. Safe - Command system switched off.

During flight, telemetry indicated that the conmand antennas, receivers/decoders, and destruct controllers functioned properly and were in therequired state of readiness if needed. Since no arm/cutoff or destructcommands were required, all data except receiver signal strengthremained unchanged during the flight. At 5889 seconds the safingcommand was initiated, deactivating the system. Both S-IVB stagesystems, the only systems in operation at this time, responded properlyto the safing conmand.

14-1

The Launch Vehicle-Lunar Module Test Article/Command Service Module (L¥-Lim/CSM) was inserted in earth orbit with the range safety system armed, whichis normal. The safing command which normally follows shortly afterinsertion was not sent, however, since the S-IE and S-IVB extended burnPeriods resulted in insertion occurring further downrange, leavingInsufficient time before Loss of Signal (40S) to safe the vehicle atthe normal post insertion time. The system was safed during the firstorbital pass over KSC,

RF performance of the system is discussed in paragraph 19.5.3.2.14.3 COMMAND AND COMMUNICATIONS SYSTEM

The IU Command and Communication Systen (CCS) is a phase-coherent receiver-transmitter system capable of establishing a communication link betweenthe Unified S-Band (USB) ground stations and the IU! of the Ssturn ¥launch vehicle, The operational requirements of the CCS include commandup-data and downlink telemetry. Turnaround ranging, although notmandatory, can be performed. “Specifically, the CCS receives anddenodulates conmand up-data for the guidance computers in the IU,transmits Pulse Code Modulated (PCM) missian contro] measurementsoriginating in the S-IVB and the IU to the USE ground stations forprocessing, and coherently retransmits the pseudo raridom noise rangeGode that is received from the USB ground stations. The CCS physicallyconsists of a transpendor, power amplifier, and antenna system.The performance of the CCS command functions was, in general, verygood, despite the vehicle anomalies.The command portion of the CCS transponder performed flawlessly, asindicated in Table 14-1. Fifty-two flight commands and 600 test wordswere transmitted by the ground station, and all were received by thevehicle.

The Mission Control Center-Houston (HCC-H) Command History shows thatattempts were made to send two additional flight commands from Carnarvonduring revolution 3. The first occurred at 16:12:16 Greenwich MeanTime tant) and the second at 16:12:30 GMT,

Records indicate that Carnarvon was unable to transmit either command.The 70 kHy subcarrier was off when the first attempt was made(Carnarvon was sweeping in order to acquire the downlink), and thetransmitter was off during the second attempt.

14-2

Table 14-1. CCS Command History, AS-502

FLIGHT COMMANDS TEST WORDS

RECEIVED AND RECEIVED ANDISTATION PASS SENT VERIFIED SENT VERIFIED

Re rmuda 2 340 340

Carnarvon, 3 2 23

Canberra 1 133 133

Hawai i 3 7 7a 15 15

IGuaymas 1 127 1273 9 9

frotal 52 52 600 600

Adequate signal strength for good telemetry was achieved throughout mostof the mission, Exceptions occurred at:

a, Bermuda during passes 2, 3, and 4, where it appears that the CCSreceiver remained connected to the acquisition antenna, causing lowsignal strength.

b. Carnarvon during pass 3, as previously described.

c. Handover problems at several stations caused some data loss. Themost notable example occurred during Guaymas first pass, when handovervas attempted before Texas was ready.

RF performance of the system is discussed in paragraph 19.5.3.1.

14-3/14-4

SECTION 15

EMERGENCY DETECTION SYSTEM

15.1 SUMMARY

The space vehicle Emergency Detection System (EDS) was flight tested inthe automatic abort closed-loop configuration on AS-502. Launch vehislemeasurements indicated that no EDS limits were exceeded and the systemfunctioned properly. There were some anomalies indicated in the space-craft. {See paragraph 15.4.)

15.2 SYSTEM DESCRIPTION

The AS-502 EDS configuration was the same as AS-501, with the exceptionthat the avtomatic abort mode was active. There are two parameterswhich are monitored for automatic abort, These are: angular overrateand two or more S-IC engines out. These are deactivated by the InstrumentUnit (IU) switch selector prior to S-IC inboard cutoff. After automaticabort deactivation, overrate and S-IC thrust are manual abort parameters.The remaining manual abort parameters are:

Angle-of-attack (aP).

Launch vehicle attitude reference failure.SI thrust.S-1VB thrust,S-I¥B propellant tank pressures (orbital phase only).Vehicle attitude, attitude rates, and attitude error (spacecraftsensed).

>eaoge

Figure 15-1 is a functional diagram of the AS-502 EDS.

35.3 SYSTEM EVALUATION

15.3.1 General Performance

The excursion of the parameters monitored by the EDS sensors remainedwithin EDS limits for proper time periods throughout flight, with theexception of premature cutoff of $-II engines No, 2 and 3.

15-1

weiBetg(BUOLZoUNy

S03*L-g1

aunBi4

TT

II-$

‘a5ONE,

ns

0193735HOLIPS:

440409

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——

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28.

laceyTyna

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WEASASLIGAYS

NUDA

Nunays

NOTIVEREaSLivisaovds,

15-2

15.3.2 Propulsion System Sensors

Three thrust OK sensors are used in each of the F-1 engines and two areused on each of the J-2 engines. The F-1 thrust OK switches are votedtwo out of three to give indication of engine out, and the J-2 thrustOK switches are vated one out of two in the logic circuitry. Al] thrustmeasurements from the launch vehicle indicated proper operation of thethrust OK logic. Table 15-1 shows the thrust switch operation times.

On marned Saturn ¥ vehicles the S-I¥B propellant tank pressures will bemonitored by the flight crew during the orbital phase of flight.Although no provision was nade for pressure display in the Block I space-craft, the sensors functioned properly. and the tank pressures remainedwithin acceptable limits.

18.3.3 Flight Dynamics and Control Sensors

The angle-of-attack dynamic pressure product is sensed by a redundantQ-Ball mounted atap the Launch Escape Tower (LET). One output is displayedin the conmand modute and telemetered from the spacecraft; the other out-put is routed to the IU from which it is telemetered. The maximum aPrecorded during the AS-S02 flight was 0.60 Nene. 86 psi) at 66.5 seconds.The preliminary Saturn V abort limit is 2.21 N/eme (3.2'psi), Figure 15-2.

A failure of the launch vehicle inertial reference is indicated whenthe platform gimbal angles are displaced excessively for a given incrementof tie, The limits for AS-502 were such that an angular displacement inexcess of 0.4 degree must occur in at Teast three minor computation cyclesof a major computation cycle. Reasonableness test failures must thenoccur 15 times during the next second before an inertial referencefailure is considered to exist. The maximum gimbal displacenent during asingle minor computation cycle for AS-502 powered flight was 0.144 degree.This represents 36 percent of the rate required for a failure indicationas stated above.

Angular rates are sensed by three rate gyros in each axis, The outputsof the gyros are fed through filters to rate switches. The rate switchsettings for AS-502 were +4 10.49 deg/s in the pitch and yaw axes and +2041,5 deg/s in the roll axis, When two of three rate switches in any oneaxis indicate an overrate, an overrate indication is given to the space-craft; and, prior to overrate auto abort disable, an automatic abort isinitiated,” The maximum angular rates (unfiltered) measured on AS-502during the period in which the overrate automatic abort was active were asfollows: 42,5 deg/s in the pitch axis, £0.5 deg/s in the yaw axis, and$5.0 deg/s in the roll axis, There was no indication of any rate switchClosures on flight records,

15-3

Table 18-1. Performance Summary of Thrust OK Pressure Switches

ME CLOSEDSTAGE encrne

|

swrtck teeTe, (RANGE.TIMEbec)

S-IC 7 7 ~1.64 148.601 2 -1.64 148.601 3 “1.73 148,602 1 “1.31 148.602 2 “1.31 148,60z 3 “1.23 148.603 1 -1.64 148.693 2 -1.68 148.693 3 “1.56 143.604 1 a 148,604 2 -1.33 148.574 3 -1.33 148,575 1 -2.00 144,905 2 -1.92 144.906 a -2.00 144,90

S-II ? 1 183.70 576.491 2 183,79 576.43z 1 183.77 412,882 2 153.74 412,92

3 1 183.77 aaie3 2 183.74 at4 1 153.70 576.434 2 183.74 576.455 1 182.70 576.495 2 153,71 576.51

S-1¥8 1 1 582.03 747.27Lip 1 2 582.03 747.27

15.3.4 EDS Sequential Events

The sequential events pertinent to the operation of the EDS were normal.Table 15-2 lists the discrete event times for AS-502 and Table 15-3 liststhe switch selector event times for AS-502.

15.4 INTERFACE CONSIDERATIONS

Although launch vehicle EDS indications were normal, there were reportsof anomalies in the spacecraft. These anomalies apparently are relatedto the transient events which occurred at approximately 133 seconds and arediscussed in Section 9A.

2.5

TIT TT ETLITTTiT iy L355PRELIMINARY SATURN Y ABORT LIMIT :

2.21 N/omé (3.2 psid)Lf. Phot pq ayyjp

2.0

b 2.5

1s

y } 2.0¥ 2

Be Pie

b1.0

0.5 WW

bas

LALA

20 40; 60 80 100 129

RANGE TIME, SECONDSFigure 15-2. Q Ball AP Versus Flight Time

15-5

Table 15-2. Discrete EDS Events

DISCRETE MEASUREMENT DISCRETE EVENT ON/OFF

|

RANGE TIME (SECK73-602 ECS or WAN. Cutoff of

|

on 20.7LV Engines Armed

K74-602 Ets or MAN. Cutoff of

|

on 40.95LY Engines Armed

kei -602 EDS S-IC One Engine Gut) On 144.89Re2-602 EDS $-IC One Engine Out| on 14.89K79-602 EDS S-1C Two Engines on 148.64

loutk80-602 EDS S-IC Two Engines On 148.63

loutKa8-602 S-IC Stage Separation

|

orf 149.99k57-603 Q-8a71 on Indication ort 150,71

(+5021)K58-603 Q-Ball on Indication ofr 150.76

(+6041)k87-602 LET Jettison "A" On 184.78

LET Jettison "A" orf 184.81«87-602 LET Jettison "B" on 184.98

LET dettison "BY orf 188..3¢«75-602 EDS or NAN. Cutoff of

|

On 11,666.IL¥ Engines from S/¢

K76-602 EOS or MAN. Cutoff of

|

on 11,666.

L¥ Engines from S/¢

15-6

Table 15-3. Switch Selector EDS Events

FINCTION

Start of Time Base |Auto-Abort Enable Relays ResetMultiple Engine Cutoff EnableLaunch Yehicle Engines EDSCutoff Enable

S-IC Two Cngines Qut Auto-AbortInhibit Enable

S-1C Two Engines Gut Auto-AbortInnibit

Excess Rate (P, ¥, R) Auto-Abort Inhibit Enable

Excess Rate (P, ¥, R} Auko~Abort Inhibit

Start of Time Base 2Start of Time Base 3Q-Ball Power OFFLET Jettison "A" OnLET Jeteison "8" On

RANGE TIME FRON BASE (SEC)stage |TIME

(sc)

|

nomtma.

|

actual

|

sewraTioN

- 69

|

1) 40.0 a --1M set

|

ty 45.0

|

4.95

|

-0.08sere

|

14.65

|

1) 414.0

|

13.96

|

0.08io) ac.66

|

Ty +90.0

|

29.97

|

-0.03

tw {138.98

|

1) 4139.3 1135.26

|

-0.08

rw

|

136.97

|

Ty 4135.5

|

135.48

|

-0,02

w 136.34

|

ty #135.7

|

135.65 0.08

rT 136.64], -138.2

|

135.85] -0.05

~ yaa.gs |Tz 40.9- waa |13 40.9w 180.76 |Fy 42.4ww |te4.77 |Ty 436.4wf 184.98 T3 +36.6

1-7/15-8

SECTION 16

VEHICLE PRESSURE AND ACOUSTIC ENVIRONMENT

16.1 SUMMARY

The vehicle internal, external and base region pressure environment wasmonitored by a series of differential and absolute pressure gages. Thesemeasurements were used in confirming the vehicle design external,internal, and base region pressure environments. The flight data weregenerally in good agreement with the predictions and compared well withthe AS-501 data. The pressure environment was wetl below the designlevel.

The vehicle internal and external acoustic environment was monitored bya series of microphones positioned to measure both the racket engine andaerodynamically induced fluctuating pressure Tevels. The measuredacoustic levels were generally in good agreement with the liftoff andinflight predictions and with AS-501 data. The S-IC stage internalacoustic levels at liftoff and during flight were somewhat lower thanstatic firing levels. No detrimantal effects due to the acoustic Tevelshave been determined at this time

16.2 SURFACE PRESSURE AND COMPARTMENT VENTING

16.2.1 S-IC Stage

External and interna? pressure environments on the S-IC stage wererecorded by 43 measurements which were located on and inside the enginefairings, aft skirt, intertank, and forward skirt, Representative datafrom a portion of these instruments are compared with the AS-501 flightdata and predictions in Figures 16-1 through 16-4. The ambfent pressurehistory of the AS-S02 flight is approximately 0.10 N/cn2 (0.15 psi}greater than the history based on AS-501. The predictions are based onaut able wind tunnel data and the 48-hour Observed Mass Point TrajectoryNPT) .

The AS-502 S-IC engine fairing compartment pressure differentials areshown in Figure 15-1. The AS~502 pressure data were generally Jess thanthe data for AS-501. This was expected as a result of removing the baseflow deflectors on AS-S02. However, the agreement {s good and the trend:are the same.

16-1

16-2

Figure 16- 1, S-IC Engine Fairing Compartment Pressure Differential

RANGE TINE, SECONDS20 Too 120

FAIRING

PRESSURE

DIFFERENTIAL

(Piat-Pamb),

N/em2

YER STA3.83 m

150,79 inj

SEROUD 0,

aN

2

FAIRING

PRESSURE

DIFFERENTIAL

(Pint-Pamo)

,N/em2

YEN STA

PREDICTED

HRD OT—

T40

FAIRING

PRESSURE

DIFFERENTIAL

{Pint-Panb),

psid

FAIRING

PRESSURE

DIFFERENTIAL

{Pint-Pamb),

psid

SHROUDa)

g 37F ‘eoso-0e | 4 gea2 oakSe ecfeticeFga8 28aEm 85gee Laps ates 0

25 288a v=PREDICTED =

“4

“4 i

‘ ‘SHROUD af)

wen eel I,0050-108 |*ys

g 2 lobar pooss.t0e| 25. suRSt 25EY +4 - toe

Bo perpen 9 Be

ze ' eeest Aust i age

z ' at“ i “— T

4 SHROUD &

4 veut Ze L@ 32 6 poosb-is

[

“yeZz 2 (268.50 in.) 'Do0s8-108 | 2

Ee a | Be

2. || 3"ae =| 08sgs tl #3

Zo CRUSH. - ge

= 4d DICTED

a a a CDRANGE TIME, SECONDS

Figure 16-2. S-IC Engine Fairing Pressure Loading, Sheet 1 of 2

16-3

SHROUD Oy+ YH sta ‘he2 } Boo7ra-riof

*

ge2 2beuast (145.67tn.) 0082-710 5ge bz2. a 2.BE 0 - aeet t Nas. s0z ee

= i . PREDICTED ae

“ ! vt i

4TROD

he Gom7-110f *g 2 | —finss.67 tn. pooss-t10] =By bos : 2222 AS-502 ! gSSs + oneBe

|

HV" + stee

|

cus assay SOL 725Be | PREDICTED zz . ae

“4

4° - 4g 10su 2 Wa] 2a3 ues 282

25 o hf ets 08geo ri Begé rN tee= crus 4 ' ge

o1cr 48

20 a e810) ab Ta

RANGE TIME, SEcaMDs

Figure 16-2. S-I¢ Engine Fairing Pressure Loading, Sheet 2 of 2

16-4

~ The $-IC engine fairing pressure loading is shown in Figure 16-2. TheAS-502 and AS-501 data agree very well in magnitude and trend. Theshrouds experienced a crush loading over almost the entire flight becausetheir geometry is intended to deflect the air stream away from the S-ICengines.

The S-IC engine, intertank, and forward skirt compartment pressure dif.Ferentials are shown in Figure 16-3 as a function of range time. Due tothe flow deflector removal the AS-502 engine compartment pressure differential was less than that experienced on AS-501 throughout flighthig resulted from the compartment being vented inte the shroud baseregion, which also experienced slightly lower pressures on AS~502 ascompared to AS-501, as will be shown in paragraph 16.3. The intertankand forward skirt pressure differentials show good agreement as a func-tion of range time. The predicted bands ware derived analytically usingmaximum and minimum leakage areas. The intertank and forward skirtpressure differentials show a drop between 60 and 70 seconds on botflights. This wes associated with the vehicle passing through Machand wes probably the result of a normal shock moving rapidly down the sideof the S-II and S-IC stages. On each Flight Mach 1 occurred between 60and 61 seconds.

The S-IC engine, intertank, and forward skirt compartment pressureloadings are shown in Figure 16-4. The engine and intertank pressureloadings agree well with the AS-501 data. The forward skirt loading wagreater on AS-502 throughout. flight but presented no problem since themaximum value of approximately 0.25 Nfem2 (0.36 psi) was well below theaerodynamic design value of 1.38 N/om2 (2.0 psi). The predictions werebased on wind tunnel data and predicted internal pressures

16.2.2 S-II Stage

Surface pressure and compartment venting analyses were conducted usingthe AS-502 final OMPT and angle-of-attack data obtained from the S-ICFlight Control Conditioned Data Tape (Q-ball}. Atmospheric data wereobtained from the final Meteorological Data Tape (Met Tape).

The external flow field at a discrete point on the S-II stage wasanalyzed by means of the semi-empirical digital computer flow fieldprogram. Since the flow field program assumes that the velricle has aClean configuration the flow disturbances created by the existence ofmultiple protuberances cannot be predicted. The internal pressure wasanalyzed by means of a multiple venting digital computer program.

Comparison plots of the pressure loading, acting across the forward skirtwall, are presented in Figure 16-5. The design, AS-502 flight, and post-flight prediction data are presented in the form of maximum-minimum databands. AS-501 flight data are also shown for comparison. Both flightand predicted pressure loadings were obtained by taking the differencebetween the respective external pressure values and an internal pressurewhich wes measured at vehicle station 74.53 meters (2934.25 in.) and

16-8

16-6

Figure 16-3, S-1C Compartment Pressure Differential

RANGE TINE, SECONDS

FORWARD

SKIRT

COMPARTHENT

PRESSURE

OLFFERENTIAL

{Pint-Pamb),

N/em2

INTERTANK

COMPARTMENT

PRESSURE

IFFERENTIAL

(Pint-Pamb),

N/cm2

ENGINE

COMPARTHENT

PRESSURE

DIFFERENTIAL

(Pint-Panb),

N/em2

60

PREDICTEDtof

30

709

5(1813.78 in.ab

004

(hh |oa} Ope |0092-113

YEH Sra18.55 n

730.31 in.)

Tao

FORWARD

“SKIRT

CONPARTMENT

PRESSURE

OIFFERENTIAL

(Pint-Pamb),

psid

fi.0

INTERTANK

COMPARTMENT

PRESSURE

DIFFERENTIAL

(Pint-Pamb},

psid

ENGINE

COMPARTMENT

PRESSURE

DIFFERENTIAL

(Pint-Pamb],

psid

“p-9LaanB

LdBuypeoy

aunssaigquaugseduo)

31-sAGINEcaNPARTHERTPRESSURE Touatas{rint-Poxt)Wet

9

EXGINEcOnPARTAGATPRESSURE aaztne(PinePent)Rene

i

iaott.

1 @soN0n3s

"LL20H

a

on

ICaPARINENTPRESSURE TOHBISGrineevespate

FoqwngoskzaeencssuRe UORLIME(PtatePent|sNon?

aataecempantoenPRESSURE Tnabine(PtatoPexe)psi

treRtanecowARTIENTwRESSURE NADINE(PinesPant)Wend

rene

*

8

soosTL

7604

oa

wa

abangarresin LUADINE(Praeovest]get

nietacOnPARTEERTPRESSURE Gabthiattertysete

PRES

SURE

LOADING

(Pext-Pint),

N/en

®Nyon?

&

FORWARU SKIRT

CRUSH oa

BURST

Sree. ES:a: me

ESIGNerr

PZMEASUREVENT VEHIOLE'STA m (in), L-Di

T_As-501 DATA

‘DIOS-219 3.88 (2514.96)———-—D109-219 63.75 (2509.84)

(CEyAs-502) 6130-219 63.51 (2500. 39, “10EGCOIPOSTFLIGHT PREVICTION OIT1-213 62.99 [2479.92

Sc1C/S-H_INTERSTAGE

Des-120 YEHr 99-120 STA 37.64n

(1981.891n,)DIOI-120 YEH STA

Z 36. 85m: {1480.785n.) Lic

CRUSH AS-502 AS-501

0

BURST| L-1.0

|_ POSTFLIGHT ESIGNPREDICTION Ro LITF-2.0

o 20 40 60

Figure 16-5.

80° 106 120,140" —“T60 «180200RANGE TIME, SECONDS

$-II Compartment. Pressure Loading

16-8

SSURE

LOADING

(Pext-Pint),

psid

Pre!

PRESSURE

LOADING

(Pext-Pint),

psig

assumed to be uniform within the interstage. The flight and predictedvalues were in fairly good agreement and well within design limits,

Comparison plots of the pressure toading, acting across the interstagewall, are also presented in Figure 16-5. The design, AS~502 flight, andpredicted data as well as the AS-501 flight, data are shown. The pre-dicted loadings were approximately 1.0 N/cm? (1.5 psi) higher than thecorresponding flight values, however, both fell within the design limitsThis apparently large discrepancy was primarily due to the method andvalue of the discharge coe“ficient used te predict the internal pressure,The method was identical to that used ‘or the forward skirt analysiswhere fairly good agreement was obtained. However, since the boundarylayer is considerably thicker on the aft skirt region an increase in theventing efficiency, with a corresponding lower internal pressure, mayresult. Note also that a uniform internal pressure was assumed to existthroughout the aft skirt compartment for purposes of inflight ventinganalysis.

Comparison of the AS-502 flight and postflight predicted pressure loadingacting across the LHz sidewall insulation at vehicle stations 54.2 and59.8 meters {2733.86 and 2354.33 in.) is presented in Figure 16-6. TheFlight data for vehicle station 48-0 meters (1889.76 in.) is to be conesidered unreliable pending further investigation and is shown in thifigure for reference only. AS-501 flight measurements for vehicle station59.8 meters (2354.3 in,) are also shown and compare well with the cor-responding AS-502 measurement,

The predicted internal pressure histories were computed by means of themultiple venting digital computer program using 2 math model to simulatethe Liz sidewall insulation venting, The math model was developedempirically by matching S-11-1, S-II-3 and S-II-4 ambient blow down testdata.

16.2.3 S-1¥B Stage

Pressures on the S-IYB stage were measured by one internal transducerin the forward compartment and 21 external and 3 internal measurementsfor the aft compartment.

Figure 16-7 shows the predicted internal minus ambient pressure differ-entials for the forward compartment together with flight data for bothAS-501 and AS-502. The vent area for AS-502 was 0.097 m2 (150 in.2) ascompared to 0.129 m2 (200 in.2) for AS-501. With the trajectories flown,the smaller vent area on AS-502 should have resulted in higher internapressures than AS-501 and correspondingly higher pressure differentials.The lower internal pressures for the first 60 seconds of the AS-502flight, 0.034 N/en2 (0,05 psid), are attributed to instrumentationaccuracy (40.52 N/cm2, £0.75 psi).

16-9

AS S02 DATA BAND FORYEH STA 59.8 m AND 54.2 01AS 502 DATA YEH STA 48.0 m

Q2ZZZZz4 FS 502 POSTFLIGHT PREDIGTED© AS S01 DATA AT VEH STA 59.8 n

18I 2.0

10 —~iN bio &

o, 0 eR 35 E0 o ¢z aay z

2-05 gé f-1.0 820 #5 DESIGN 2= AP= -3.5 Nom? (-5 psid) b-2.0 #2 ¢ : =2.57] +2 YEHSTAz . 2.8m -Bor) (2854.33 in. b-3.0

. 54.2 mast (2]391a6 tn.)

$8.0= (1883.76 in.) f -4.0

-3.0 1. 1o 2 30 75 yo 12 750

RANGE TIME, SECONDS

Figure 16-6, S-II Lig Sidewall Insulation Differential Pressure

Figure 16-7 also shows the absolute pressure history in the forward skirtfor a time interval centered on 133 seconds. The abrupt pressure dropat this time is thought to be associated with the 133 second anomaly asdiscussed in paragraph 9A.2,1,

Figures 16-8 and 16-9 show predicted and measured pressure differentialsand pressure loadings for the aft compartment as internal -minus-ambientand internat-minus-external respectively. The flight data fell withinthe predicted band during the critical flight. period and the maxtmumbursting and crushing pressures, 2.0 and 0.55 N/en@, (2,9 and 0.8 psidrespectively) are well below the design values, 3.23 and 1.54 N/cm(4.69 and 2.24 psid).

16-10

spew

Figure 16-7. S-IVB Forward Compartment Differential Pressure

we we Tb Ee

parnera, @ = 0

7

fonse “IE, S00

Figure 16-8. $-IVB Aft Skirt and Interstage Differential Pressure

6-11

oreNE A

} | i

i

ene TM, ems

Figure 16-9, S-1VB Aft Skirt and Interstage Pressure Loading

16.3 BASE PRESSURES

16.3.1 S-1C Base Pressures

Static pressures on the S-1C base heat shield were recorded by 11 measure-ments, 2 of which were heat shield differential pressures. Representativedata from a portion of these instruments are compared with AS-501 flightdata and predictions based on the 48-hour OMPT. The predictions includethe effects of the base flow deflectors since wind tunnel data on a con—figuration without flow deflectors were not available.

The S-IC base pressure differentials are shown in Figure 16-10, Ingeneral, the agreement is good. The AS-502 base pressures were slightlyless than AS-501 up to approximately 20 kilometers (10,8 n mi) altitude.Beyond this altitude only small differences existed. The lower pressurebelow 20 kilometers (10.8 n mi) was a result of removing the base flowdeflectors on AS-502.

The S-IC base heat shield pressure loadings are shown in Figure 16-11,Again the AS-502 data were less than AS-501 Selow 20 kilometers(10.8 n mi}. The flow deflector removal on AS-502 lowered the basepressure ta which the engine comoartment vents. These heat shield dif-ferentials were well within the 1.38 N/cm2 (2.0 psid} design differential.16.3.2 S-IT Base Pressures

The postflight oredictions of the S-Ii base heat shield aft face staticpressures are evaluated from a semi-empirical correlation of base

16-12

16-13

Figure 18-12, S-IC 3ase Pressure Differentials

ALTITUDE, kn

or

10 20 a0 0 50

BASE

PRES

SURE

DIFFERENTIAL

{Poage-Panb)

,H/

en?

BASE

PRESSURE

O1FFERENTIAL,

fPbage-Pamb},

N/ome

ASE

PRESSURE

DIFFERENTIAL

ain

Wa

oR

As.

i

501

PREDICTED |

PREDICTED |

ne

Vx

ao

Nees

AS-502

IL |1

2.81 m

VEH STA

{110-63 in.)

Vex STALJ

2810(110.63 in.)

|

IEe

veH STA2.81 0

(110.63 in.)

a

OX0038-106

{Pbase-Pamb),

psid

JASEPRESSURE

DIFF

EREN

TIAI

®Oo F-1

0036-106}.

J

2

1

0

“BAS

EPRESSURE

DIFF

EREN

TIAL

(Poase-Pam),

paid

T L

? T

——

BASE

PRES

SURE

OLFF

EREN

TIAL

{Pbase-Panb),

psid

ALTUTUDE, ni

HEAT

SHIELD

PRESSURE

LOADING

HEAT

SHIELD

PRESSURE!LOADING

Nyeme,

(Pint~Phase),

nt-Phase},

N/cm2

ALTITUDE, 9 mi15 20 25t L n

YEH STA2.81 m

(110.63 in.)

o

T =| YEH STA Bis

- | 2.81 m Napie»(110.63 in.) SS)

0047-106

Figure 16-11.

ALTITUDE, ke

S-IC Base Heat Shield Pressure Loading

16-14

psid

HEAT

SHTELO

PRESSURE

LOADING

(Pint-Pbase),

osid

HEAT

SHIELD

P2ESSURE

LOADING

(Pint-Phase)

,

pressures and heating rates derived fron the hot flow scale model testresults and AS-501 flight data. The effects of the interstage on thebase pressures are accounted far in the analysis. The postflight pre-dictions of the thrust cone region pressures were based on the AS-50flight data.

Figure 16-12 presents the predicted and measured static pressures on theaft face of the base heat shield. It is noted that the pressure distri-bution on the base heat shield was more uniform prior to second planeseparation, The pressure reached a peak value during interstage separa-tion because the J-2 exhaust plumes were confined by the interstageresulting in high impingement pressures and increased reverse mass flowrates.

After engines No. 2 and 3 cutoff, the pressure distribution on the baseheat shield aft face became highly unsymmetrical. This was becausepositions IT, II1 and IV on the heat shield no longer exeertenced directreverse flow impingement, The pressures at these positions drozved tothe ambient level while the reverse flow impingement pressure wa:maintained in position 1 quadrant.

Figure 16-13 presents the predicted and measured static pressure on theheat shield forward face and on the thrust cone surface. The sressuredistribution was more uniform in this region as compared to that on theheat shield aft face. The pressure rise resulting from S-IC/S-IT inter~stage separation was also apparent in the thrust cone region; however,it was not as pronounced as that on the base heat shield aft face. Notethat the pressure in the thrust cone region dropped by a factor of 20after interstage separation.

Figure 16-13 also shows that a considerable pressure rise existed in thethrust cone and heat shield forward face region immediately prior to andduring J-2 engine No. 2 cutoff at 412.92 seconds. Note that the maximumindicated pressures were above the pressure transducer range of 0.0689 N/em?(0.1 psi} and that they were also above the pressures recorded prior to andduring interstage separation. Therefore, it is concluded that the measuredpressure rises could not result from the J-2 engine exhaust reverse flow.Since the base heat shield aft face/thrust cone region is open to theatmosphere, and because the measured pressure rise appears to be uniformthrough the base heat shield forward face region, it appears that a verysudden and substantial mass injection into this region occurred causingthe measured pressure rise. For further details see paragraph 6.3

Figures 16-14 and 16-15 present an overall view of base region pressureinstrumentation together with the flight data recorded prior to andduring the engines No. 2 and 3 cutoff time interval. It is seen thatthe recorded pressure rise on the thrust cone surface was considerablysmaller than the pressure rise on the base heat shield forward surfaceand that a relatively small pressure rise was also-recorded on the heatshield aft surface at this time, This was probably due to enginegimbaling associated with the cutoff of engines No. 2 and 3

16-15,

16-16

©(ZA

os-son

nrovcte

—Tases

ohv0

nara

Figure

16-13.

S-II

Base

Heat

Shield

Forward

Figure

16-12,

S-II

Heat

Shield

Aft

Face

Pressures

Face

and

Thrust

Cone

Pressures

ok Hass

16-17

80207yee

028

org]

“aS?can

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relPSUR

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008

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414

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Figure

16-1

4.S-IL

Base

Heat

Shield

and

Thru

stCone

Pres

sure

s

PRESSURE,

N/em?

ay cae

02 z

TigrebeetsAmisite4 0 1

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Figure 16-15.

zPRESSURE,

/en'

Ah

PRESSURE,

Nyon?

PRESSURE,Nyon

oy 0

08.06| -

06%> 0151-206 -peace |messut |oy Theo

222

eyMot °TeaPAGE TINE, SEUONDS

PRESSURE,

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PRESSURE

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ANSE TIME, SECM

6TTpigaez08Pa LWA p08PRESSURE fw] Nya

Wy b.04

ol f-.03408 tae ata

ANGE TIME, Secowas

S-I] Base Heat Shield Pressures

16-18

PRESSURE,

ps

~ 16.4 ACOUSTIC ENVIRONMENT

16.4.1 External Acoustics

The external fluctuating pressure environments for the AS-S02 vehiclewere recorded by nine instruments located on the instrument unit,S-I¥B forward and aft skirts, S-11 forward and aft skirts, and S-ICintertank, aft skirt and fin D. Representative data for theseinstruments along with AS-501 data and applicable prediction curvesare shown in Figures 16-16 through 16-19.

Both the Digital Spectral Analys*s program (Ravan) and acoustic analyzerdata have been used in determining the plots shown in Figures 16-16through 16-19. Due to incomplete data, overall pressure levels wereobtained by integrating the 1/3 octave band fluctuating pressure leveltime histories for each instrument. Ravan data have primarily been usedto determine time slices where maximum aerodynamic fluctuating pressureoccurs. Acoustic analyzer overall levels were computed at these timeslices to verify the Ravan results. These data remain relatively quick-Took and may be revised after further analysis.

The AS-502 externa? acoustic environment at liftoff is shown in Figure16-16. Both Ravan and acoustic analyzer data are plotted for comparisonwith predicted and AS-501 curves. Agreement is generally good, Thescatter of Ravan, acoustic analyzer, and AS-50] data points is below6 decibels except at vehicle stations 2.06 and 63.98 meters (87.10 and2518.90 in.) where an approximate difference of 10 decibels existedThese differences are still under investigation.

VEMICLE STATION, 1

007 raya Tala tsrd.8 1968.5 238.2 56.9 3149.8 2849-90“ TT 2 5-301. Lon . : _, 8 832502 av

Ze 1 aS-s02 ACSTIE ANALYZERcn i ee PREDICTED.eel +

ae

gx

a w 20 30 40 30 ey 0 w 0

VEMICLE STATION,

Figure 16-16. Vehicle External Overal} Sound Pressure Level at Liftoff

16-19

Liftoff sound pressure spectral densities are compared with AS-5G1 datain Figure 16-17. Frequency characteristics appear similar fcr both AS-501and AS-502 with the exception of an apparent level shift in three of theS-IC stage instruments. Instruments 8005-200 and E0CC4-200 are locatedat the same vehicle station but on opposite sides cf the S-Il stage.Data from both measurements were similar. The maximum levels indicatedat launch were 148 decibels and 183 cecibets for 60005-200 and B0004-200respectively. Therefore, data are shown for B0004-200 only. In thisreport B0013-426 on the AS-501 is compared with B0025-426 on the AS-502flight, since they are at the same location, and appear to indicatenormal ‘operation,

vevall fluctuating pressure levels are shown in Figure 16-18. 4S-502acoustic analyzer and Ravan data are both compared with predicted andAS-501 data. Generally higher levels are indicated on the AS-507.Notably excepted is the IU instrument which shous pointwise drans of uoto 10 decibels. Further analysis appears necessary. The AS-501 overalllevel for instrument B0002-115 terminated at 90 seconds due to instrumentfailure,

A 5 hertz component shawed up in the data for two measurements on theS-II aft skirt and S-IC intertank section, shown in Figure 16-19, Theappearance of this § hertz component requires investigation because thelowar frequency limit, where the telemetry system began to attenuate thedata signal, was 50 hertz.

The Occurrence of the 133-second transient in the acoustic data channe’sis similar to other data records which are discussed in Section 9A, Thesignal Increase at this time was significant and requires further analysis.Pressure Spectral Densities (PSD) for AS-802 at maximum aerodynamicfluctuating pressures are shown in Figure 16-19. Acoustic analyzer datawas used in the PSD calculations while Ravan data fixed the time at whichmaximums occurred. Similar distributions are evident between AS-~801 arcAS-50? data with sone level shifts noted.

16.4.2 Internal Acoustics

16.4.2.1 ScIC Stage. The S-IC stage intertank internal acoustic datafrom a single measurement is shown in Figure 16-20. Liftoff levels weresomewhat Tower than those for static firing and were similar to those forAS-501. Flight levels were much lower than static firing levels andgenerally the same as flight levels on AS-5C1.

16.4.2,2 S-I1 Stage. Two internal microphones were located on the S-ITstage as foTTaws:Measurement Saturn ¥ Station Azimuth RadiusAveaNumber m Cin.) (deg) om Cin.)2001-206 Thrust Cone 42.6 (1677.2) 133.50 (138)B002-219 Forwarc Skirt 63.3 (2892.1) 1744152. (178)

16-20

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16-1

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16-17.

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16-18.

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16-18.

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16-19.

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28 WT ENMICATEAi

= We Te Th 0pease TIME, stccs

Figure 16-20. §-IC Internal Acoustic EnvironmentFigure 16-21 presents the measured overall decibal levels versus rangetime. Also indicated in the figure are the raximum expected levels forliftofF and inflight acoustics, which are values obtained from Saturn ¥Vehicle Acoustic Environment, R-PAVE-SVE-64-191, August 10, 1963.

The measured acoustic levels were well under the naxinum expected values,particularly in the thrust cone area. The thrust cone internal acousticlevel was also well under the external acoustics measured on the inter-stage, as indicated in Table 16-1. The differential at liftoff was 13to 19 decibels between the aft external and internal measurements. “hisdifferential was somewhat larger than expected and is under investigation,

The forward internal measurement 1s considered valid. The forwardexternal measurement is considered questionable, and is under investigation.

16.4.2.3 S-IVB Stage. The $-1VB acoustic environment was measured atfour positions, internal and external on the forward skirt and internaland external on the aft skirt.

Envelopes of the composite, 50 to 3000 hertz levels time histories arepresented in Figures 16-22 and 16-23. The shading between the externaland internal responses indicates the structural transmittability for

16-27

71 WOE: AScSa: Oma ASI nr 52316 cars

ATcoapanEN

Osoa Figure 16-21, $-II Internal Acoustics

sound pressures at liftoff and for boundary layer pressure fluctuationsin the transonic portion of flight, Also presented are data “ror theS-IVB acoustic measurements of the AS-501 flight.

The AS-502 forward skirt levels were in yeneral equal to or less thanthe AS-501 levels. Also the transmission loss for the AS~502 forwardstructure was greater than for the AS-501,

Table 16-1. S-II Acoustic Noise |evels Comparisoncf AS-501 and AS-502 Gata

. SWERALL ave roaeaRD Si2PT ART LVTERSTAGE/AFT SetaTVERT o - —1. - _——

INTER, ExTERySL ETERNAL(3002-225) (goad § Bovs-20n) 3021-206)

AS-902

|

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7

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125

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193.5-147,0) 146.6

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136.0 10 125 150 153 ass 155“Questionable data, ander invest*catian

Ve-26

AS~502EZ AS-501

170AMPLITUDE NEGLIGICLE, AFTER

2 760}. SIC SEPARATION 4

2 sof [exes = INTERNAL

al

3 130

$120 ~ ~

B10 SS NOISE FLOOR 200 20° 40 GO BD WO 120 740

RANGE TIME, SECONDS

F gure 16-22. $-IVB Forward Skirt Acoustic Levels

GS3 As-5a2ZA AS-501

AMPLITUDE NEGLIGIBLE AFTERS-I¢, SEPARATION

EXTERNALINTERNAL

169

150

140

130

120S

SOUND

PRESSURE

LEVELS

,db

L NOISE FLOOR,-20 0 20 49 60 BO 19D 120 140

RANGE TIME, SECONDS

Figure 16-23. 5-IVB Aft Skirt Acoustic Levels

16-29/16-30

SECTION 17

VEHICLE THERMAL Ef IRONMENT

17,1 SUMMARY

The AS-802 $-IC base region radiation and total heating rates were moresevere than those measurec on AS-5C1 flight; also the base heat shieldand engine gas temperatures were greater and increased more rapidly.Loss of M-31 to the level of the crushed core on the base heat shieldwas visually observed on this flight via the television cameras whichviewed the heat shield. The S-IC forward skirt thermal environment afterS-IC/S-II separation was higher than design.

Tre eerodynamic heating ervironrent on the S-IC forward skirt, inter-tank, engine fairings and firs was slightly higher than AS-501 data butwithin the predicted bands. The AS~502 trajectory was a higher heatingtrajectcry than that flown by AS-501.

The effect of the prevature shutdown of engines No. 2 and 3 on the S-IIheat shield and base region envfrorment was minor. With the exceptionof abrupt spikes due to the engine anomalies, the base reqion thermaldata compared favorably with AS-501 data.

Protuberance induced heating effects on the S-1I stage were generallybelow the design and postflight predictions. The temperatura datafrom AS-502 was in general very similar but slightly higher than thatfor AS-S01.

The AS-802 $~I¥B aerodynamic heating environnent was slightly moresevere than the AS-501 but temperatures were well below the design values.

17.2 S-IC BASE HEATING AND SEPARATION ENVIRONMENT

Thermal envtronments in the base region of the S-I¢ stage were recordedby 39 measurements which were located on the heat shield, F-1 engines,and base of fin D, This instrumentation included 6 radiation calorimeters,19 total calorineters, and 14 gas temperature probes. Representativedata from a portion of these instruments are compared with AS-501 Flightdata in Figures 17-1 through 17-3.

The calorimeter radiation and total heating rates measured in the baseregion were more severe than those measured during the AS-501 flight.This increase ts attributed to tre removal of the flow deflectors on the

W7-1

AS-602 vehicle. The removal of these flow deflectors allowed theexhaust gases to recirculate into the base at a lower altitude. An‘increase in heating had been anticipated from model test results.

The radiation hunp which was observed at adout 30 kilometers (16 n mi)altitude in the AS-501 base flight data was also present in the AS-502flight data. The AS-501 flight evaluation indicated thet the radiationhump may have been caused by a combination of increasing view factors,diminishing afterburning with increase in altitude, plume interactionregions and hot gas recirculation. AS-502 video taves from S-IC baseTV cameras indicate that the hot gas recirculation was the main causeof the radiation hump. The TY data shows that hot recirculated gasreached the base heat shield at an altitude of 12.2 kilometers (6.59 mi)which correlated with the rise in the radiation heating rates as shownin Figure 17-1. This recirculated exhaust flow is fully established inthe base region at 18.5 kilometers (9.99 n mi} which correlates wit thepeak in the radiation heating rate data. The TV data also indicatedthat after 36.6 kilometers (19.8 nmi), the base region was clear,in¢icating a Significant reduction in the burning of the fuel richexhaust gases in the area. This fact correlates with the dropoff inmeasured heating rates.

Results from the total and radiation calorimeters indicate that acorvective cacling rate was measured by the base heat shield instrumentsto an altitude of 19.8 kiloreters (10.7 n mi) and then changed to asmz11 corvective heating rate at the higher altitudes (Fiqure 17-1}.However, the average calorimeter effective wall temperature is 350°K(170°F) which, wren compared with the lower gas temperature curve inFigure 17-1, irdicates that the calorimeters should experience convec-tive heating at altitudes above 10 kilometers (5.4 nmi), A differenttrend is nated in Figure 17-2 for the f-1 engine nozzle extensioncalorimeters near the nozzle Tip, where convective heating was presentfrom liftoff to 2 maximur value of 11.36 watt/em@ (10.0 8tu/ft2-s} atan altitude of 18.3 kilometers (9.88 1 mi]. Comparison of the averagecalorimeter wall temperature with the gas temperature curve in Figure 17-2also indicates that the engine nozzte extension calorimeters receivedconvective heating throughcut the flight but at altitudes above24.4 kilometers (13.7 9 mi) was negligible. No appreciable differencewas noted in the canvective heatirg measured by the calorimeters onthe heat shield or engines when comparing the two Saturn ¥ flights. Thebase heat shield and engine gas temperatures were greater and rose moretapidly than those measured an the AS-5C1. This, however, is notreflected in the total and radiation calorimeter data.

The total heating rates measured by the celorimeter on the base ofFin D are compared with the AS-501 flight data in Figure 17-3 and showthe tvo to be approximately the same, The initial rise in heatingoccurred at 12.2 kilometers (6.59 n mi] which correlated with therecirculated exhaust gases reaching the heat shield as observed on theTV cameras. The second rise and peak in the curves occurred at aboutthe time that Flow sepavation occurred and was probably caused by thehot gas from the plime moving up the side of the vehicle.

17-2

“LedoanBed

JUDURIOALAU

(euvsylPLaLus

2BaHaseg

3T-s“Brdk

aanBugbua

L-4queauddLAUZ

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op

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IASTERPERATURE,°KTOTALAGATIAGBATE,RAGIATEONHFATINGRATEsationnatok

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09

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waSnLLWoe

TOTALIEATINGBATE,RAORSTECHHEATINGRare, CASTENPERATLAL,°FBusteesBears

weauri

“agus

—-—AS-501FLIGHT DATA

—— AS-502FLIGHT DATA

ALTITUDE, mi

a 5 0 18 20T 12

1e | |eosena |a 1 10| := 1 ai tm Eg2% i 8&es 4 7 ; mi eeag /(° eea ~ + 235 [ S 285 az2 att x 4+

aheffSaeee _ ESP.

1% 10 20 30 40 50 66"

ALTITUDE, kr

Figure 17-3. S-IC Base of Fin 0 Total teating Rate

The heat shietd tenoerature history is {llustrated in Figure 17-4.Average cold side and bondline (honey comb/ti-31 interface} temperatureswere 345 and 458°K (161 and 364°F), respectively, versus 333 and 380°K(139 and 224°F) for the AS-501 flight. This increase may be attributedto the deletion of the base flow deflectors on the AS-502 vehicleresulting in higher total heating, However, as indicated by the dashedHines showing predicted maximums, the data is well within design limitsexcept for the measuramants CO32-106 and C033-106. These bwo thermo-couples were located 4.22 meters (166 in.) radially fron the heat shieldcenter on Position Line 11. At 96 seconds, the sudden increase intemperature could be caused by local loss of M-31 to the level of thecrushed core. A similar failure occurred on the AS-501 vehicle. Notethe response of €033-106 in Figure 17-4 indicating that i# suddenlybecame exposed to the base region gas but did not fail. Loss of M-31is substantiated by the correlation of the output from C032-106 conparedin Figure 17-4 to a reconstructed curve. This curve was computed usingthe AS-502 radiation and gas temperature data and the desiqn heat trans-fer coefficient assuming 0.407 centimeters (0.16 in.) of M-31 was Tost

17-4

Figure 17-4. $-IC Heat Shield Thermal Environment

17-5

fron the aft surface (to the level of the crushed core} at 95 secondsThe reconstructed temerature of the forward surface using the sameassumptions is also shown in Figure 17-4. M-31 loss to approximatelythe level of the crushed core was visually observed on AS-502 va thetelevision cameras wich viewed the heat shield. These cameras alsoshowed tearing of the inboard flame curtain fiberglass cloth protectivecovering. This was expected and does not indicate flare curtaindegradation.

‘4-31 insulation loss occured at 110 seconds on AS-507 and 95 seconds‘on &S-502 which appears to be a random process and may occur at almostany time during flight. The effects of this loss at various times wascomouted and the results presented in Figure 17-5. ‘ote that the lossof H-31 at 20 seconds would result in outboard engine cutoff temperaturesof B63°K (1094°F) at the aft honeycomb face sheet and 535°K (50°F) atthe forward honeycomb face sheet. The maximum allowable temperaturefor the forward face sheet is 423°K (302°F], However, this is inter-preted as a design goat and testing showed the heat shield capable ofwithstanding temperatures above this level. Actual limits for ef therFace sheet ave not defined at this time but are being examined todetermine whether a potential problem exists,

Correlation of computed temperatures with heat shield data, as shownin Figure 17-4, for the first 70 or 80 seconds of flight was not good.Examination of the data indicaced that the actua’ M-31 conductivitytust differ from the published data. Laboratory testing has shownthat the M-31 insulation is hygroscopic and the wet conductivity wouldtherefore differ fron the published dry value. The heat shield tem-perature at the M-31-honeycomb interface for both AS-501 and AS-502increased rapidly after engine ignition to approximately the boilingpoint of water. Thereafter, the temperature followed the saturationcurve for water as ambient pressure decreased with altitude until about70 seconds, and then the temperature increased as would be exnectedwhen the water was completely boiled avay. The fast rise tine of theinterface tenverature early in flight is only partially accounted forby the increased conductivity of the wet N-31, If it is assumed thatthe transmissivity of the #-31 is greater than zero, this difficultycan be resolved. Hence, there are four factors that acting togethercould account for the observed temperature history,

a. Increased conductivity early in flight due to the presence ofwater in the N-31

b. Transpiration cooling due to water boiloff in the M-31.

c. N31 transmissivity greater than zero.

4. Decrease in conductivity later in flight due to outgassing asaltitude is gained.

Analyt'cal verification of these factors is now under way.

1-6

It is apparent that the heat shield is capable of withstanding the thermalenvironment with the flow deflectors removed. The bandline and cold sidetemperatures were below levels to which the panels have been subjectedin corb‘ned therma?-acoustic qualification testing.

T oF #-31 1058 AT” VARTOUS RANGET SHICEO TENPFPATURES AT OECD

eiTIMES aN.

1209

1109oucy coma

0 APT FACE T9009COMPUTED

00

79

ATECD.

ATURE

ATQE

CO,

60WONyom

-—FORWARO FACE. nCOMPUTED, 509

TEMP

ERALW

Tene

ace —LIMCItnG FneMARD FACE 200TEMPERATURE PER CELSPECIFICETION

30“20 u 2 40 60 #0 yoo zoADSC«ORANGE TIME, SECONDS

Figure 17-5. S-IC Thermal Environment Effect of M-31 Losson Heat Shield Temperatures at OECO

Temperatures measured on the engines were very close to the AS-501 flightdata. Temperatures under the insulation on the qimbal actuator and onan outrigger remained below 313°K (704°F). Temperature on a fuel dis~charge line reached 315°K (107°F) and the ambient ai temperature underthe engine cocoons ranged from 293 to 328°k (68 to 131°F) at the endof flight. The time-temperature data from che thermocouples located‘on the heat exchanger bellows, turbine exhaust manifold, and nozzleare plotted in Figure 17-6 for both flights.

VW-7

zone1300

a5-302

Feige? ona8.802 Lae" bara cizeeat 1600

noe 2-101 BEAT EXORANGE ~ forefnedEnvauST aN FOLD BELLOW

# ant—Jas-s02 re00*: sli ota zgg c:37-10? 2g “aa 45-801 2= no . FllGHT 9328 Eg cize-191 é= _. ary bzaue

as-501 FLtowt oatal530 c132-107* HEAT EgonauGER 400

DiS LLake 36-1

T hOBZLE300 EXTERN.

109:=20 é 2 Bd Bi et we 1SRANGE TIVE, SECONIS

Figure 17-6. Thermal Environment, Temperature Under InsulationOn Inboard Side of Engine No.1

17.2.1 S-IC/SIE SEPARATION

As shown in Figure 17-7, gas tenperacures during separation were similarto those measured on AS-501. Two spikes in gas temperature were pre-Sent, corresponding to separation and J-2 engine ignition. The factthat the first spike peaked at a lower temperature on AS-502 than onAS-501 may be due to the elimination of four of the eight S-II ullacemotors on AS-802. The second spike is related to J-2 engine ianition.Data from the separation extensometers indécate that the separationrate for AS-502 was slightly slower than for AS-501.

The forward skirt skin temperatures measured during separation areshown in Figure 17-7 and are similar to those experienced durinc theAS-501 flight. The LOX tank dome temperatures during separation areshown in Figure 17-7. These temperatures were considerably lower thanthose measured on AS-501 which were as high as 473°K (392°F). This isarimarily due to the fact that the LOX dome thermocouples were coveredwith an insulating material on AS-502 thereby isolating then fron thethermal environment causing them to read near the actual dome temper-ature. Three of the four measurements show a temperature drop atSeparation which is attributed to residual LOX impinging on the

17-8

dome. Most of the pressure data in the forward skirt area duringseparation was lost due to the Radio Frequency (RF) blackout, however,pressures of 0.857 N/cm@ (1,243 psia) were measured as compared to a5.18 N/em(7.52 psia} waxtmun on AS-501,

Several ‘vinor failures occurred after separation an both AS-501 andAS~502 which are attributed to the separation enviroment. Theseinclude a shorted casle on AS-501, loss of two LOX tank ullage pressuremeasurements on both AS-501 and AS-502, failure of a 0X vent andrelief control pressure line an AS-502, and failure to eject three ofthe four S-IC cameras on AS-502 as discussed in oaragraph 5.9. Suf-ficient instrumentation was not available on AS-502 to determine theseverity of this environment.

Assuving that these failures were due solely to overheating, extensivestudies were made that indicate a separation environment of about twicethe design values would be required to cause failure of the aluminumtubing associated with the LOX tank ullage pressure measurements and 1480. ExTRAPOATED

S20 THERMAL ENVIROWALNS UPPER 2000COMPARTMENT aMGIENT ATR TEMPERATURE ‘epuge LMT

TSOHOURING S-NC/S-2T STAGE SEPARATION ‘QF TRANGOUCER1280°K

ACE 120 (asa) 18001080 0.03652120

U S-1/S-11 SEPARATION, 148.06}

. 1200= eo. a

i 52 #S-501 FLIGHT OATA oe &© 450. #

5-802 FLCGHT OATA|460. 800

20

Me Fags STERANGE TIME, SECONDS

Figure 17-7. S-IC/S-IL Separation Thermal Environment, Sheet 1 of 2

17-9

ToT Tod ~st suerstonss verse JE weSPN Topcbe PF LsSiS SAFi ay i

JP S-1C/S-11 SEPARATION, 149.08tea: a8 Ba ne

5 anc BT

ay ~ -| 4

we a . Ji i a

ra t |a oa 1s is fi ‘i Ts an

wert St, stats

28 pra salesaa ronareh She me! T Vs36/51 sefearION, 19.08elas saicrSocl ~SEPARATI® rh] ‘

20 iess Pt i

P=SoH af.

Ez 1 |

ae -i 4

199 —-a. }| tg

i cansie L ae1 = : te ve is

Figure 17-7.

TEER

S-IC/S-II Separation Thermal Environment, Sheet 2 of 2

17-10

the contra’ pressure line, This closely correlates with the environmentrequired to match LOX dome thermocouple data from AS-501. This severeenvironment may exist only in isolated locations in the $-IC forwardskirt. “his condition cou’d exp’ain the fact that on"y a few fai ‘uresoccurred. Efforts to more accuratey define the separation environmentare continuing.

17.3. 5-31 BASE HEAT SHIELD AND SEPARATION ENV: RONMENT

‘The post flight predicted total heating rate values are based on hotFlow scale model test data. These data include the effects of thecold turbopump exhaust gas injection in the J-2 nozzle on base heatshield total heatina rates. The postflight predicted radiation heatingrates were analytically obtained by means of a digital computer programwhich uses the method of totel hemispherical emissivity derivatives, tocorpute the incident radiation, Engine performance, i.e., EngineMixture Ratio (E42), chamber pressure, temperature, and gimbalingeffects on convective heating rates are taken into account.

The AS-502 flight base region total heating rates nave been normalizedto the 295°K (71°F) cold wall condition.

Figure 17-8 shows the moximun-minimum band of total heating rates tothe aft face of the base neat shield taroughout S-II boost. The post-Flight prediction and AS-501 flight data are also shown for comparison.It is seen that, initially, the postflight prediction heating ratesare higher than’ the flight data, otherwise, good agreement was obtainedInitially, the AS-501 actual heating rates were higher than the corres-ponding AS-802 values. The sharp increase in heating rates duringinterstage separation was due to the interstage-cxhaust plume impinae-ment whith resulted in a higher reverse mass flow rate, increased 92temperatures, pressures, and hence heating rates

A small abrupt increase of base heat shield total heating rates wasabserved at approximately 319 seconds in the vicinity of the No. 2engine, as shown in Figure 17-8, Sheet 2. This figure shows that atthe same time the heat shield temperature qradient also increased ir

this region. It should be noted that at 319 seconds a slight dropin No. 2 engire chamber pressure occurred. Since no changes in engineginbaling were observec during this tine period, it appears that theonly possible cause for this increased heating could be an abnormalchange of the engine performance, for example, an increase in the EMRUnder norma] operations this would result in a corresponding increasein chamber pressure. The opposite trend was observed in this casefor further details see paragraph 6.3.

Figure 17-8 sheet 2, shous the AS-502 flight radiation heating rateto the base heat shield in the vicinity of No. Z engine. It is seenthat a slight abrupt increase in heat rate occurred at approximately319 seconds. This increase is insignificant with respect to the baseheat shield total heating rete. However, it is noteworthy that thi

17-11

increase accurred at the time when:

a. No. 2 engine charber pressure registered a slight decrease.b. One surface temperature transducer and two total heating ratetransducers in the same region showed an increase.

Analysis of engine performance and gimbaling data for this tine periodindicates that an EMR increase is the most probable cause for theincreased radiation heating rate.

Yosta 414m ssc, 148.7ABST . v ©cor-ace Y7 SEOMNO PLANE SEOARATISN, 179.06WY ENG WO.2 ANG HO. 9 our,212.92, 14.18

cb87-206 WY pee sale, b99.76~ci20-208 WH cco, 676.38

EZZ4 85-502 Flour pata

7 as-s0r preotertsT ts-s01 rurait ont

2B

nrrT) 350 400 aD SOD 550 aneRANE: TIME, SECOnES

Figure 17-8. S-II heat Shield Aft Face Heating Rates andSystem Temperature, Skeet 1 of 2

17-12

PEAT

ING

RATE,

Bea/Ft2es

Figure 17-9 presents the AS-502 flight and postflicht total heatirgrates to the thrust cone surface. AS-501 flight data is indicated forcomparison, It is seen that the postflight predicted total heatirgrates are in very good agreement with flicht data. Prior te secondplane separation, the AS-501 heating rates were slightly Tower than. thecorresponding AS-502 values, however, the Lasic heating rate trend isrepeated very closely during both flights.

Figure 17-9, sheet 2, has a comparisen plet for three indivicual thrustcone total heating rate reasuremerts for the AS-501 and AS-502 flights.It is seen that at approximately 225 seconds the AS-502 flight totalheating rate transducers indicate a gradual heating rate decrease, ascompared to the AS-501 data. Sirce no change of the J-2 enginedeflectians or performarce was ratec during thts time in flight whichcauld cause this reduction tr heating rate, it is believed that thiswas caused ty @ cryogenic leak in the base reaion.

2686-208,WV sc, 14.76 638-208, 701-206Vv SEDONG PLANE SEPARATION, 173.06

WW 3-2 ING No.2 xO 0. 3 our, 412.92, - co21-208a1a18

WY re SHIFT, 291.76 3-0°POS T

3 T T T TSESIGN VALUE 3.45 watt/en? I as-501 Fe:GHT DATA

BZA as-602 FLiGHT DATA. 2,g 2 - ESS3 PCSTFLIGHT PRECICTION s

Hor ce Fi &

g &Eo fomel =

¥ y --1 150 «200~=80~C«DSC«C«iSCSCSSSCSC«RANGE TINE, SECONDS

Figure 17-9. $-II Thrust Cone Total Heating Rates, Sheet 1 of 2

17-14

YEH STA47.4 m (1866.14 in.)

44.1 om | (666-208

821-208

f \e-0°POS 2

ui 0.4 0.5Sos (821-208 as-5a1 F 0.4ce I pis-502 F 0.3

Ee oer fr 9.23? a anee E o1= 0 o

oak 0.5Eu ce66-268 L oa

os Fr o.3ag 0? L 0.25 * Po i - 0.12 9 Soe °

joa 0.5ey C688-208 F 0.4oe 1 r 0.328 0? 0.252 —~ ~+towt_ Lo.= a

1s 200 225, 250 275 300RANGE TINE, SECONDS

Figure 17-9. S-Ii Tarust Cone Total Heating Rates, Sheet 2 of 2

V7-15

HEATING

RATE,

HEATING

RATE,

Btu/tt2-s

Btu/ft2-s

AEATING

RATE,

Btu/Ft2-s

Figure 17-10 presents an overall view of the thrust cone reating ratereasurement locations and their performance during the time period oriorto and after engine No. 2 cutoff. All the instruments indicate that thethrust cone experienced a sudden increase in heating rates and tem-peratures during a very short interval at the time of engine No. 2 cutoff.The increase appears to be spread over a very wide region and isunrelated to the location of the instrument with resnect to engine No. 2.

A carresponding plot of engine curtain gas temperatures and base heatshield forward face surface tenperatures is snawn in Figure 17-11. Itis seen that a1] but one of the gas temperature transducers exceededtheir maximum range of 422°K (300°F). Several discrepancies are observedjn Figure 17-11. These are:

a. The indicated heat shield surface temperature in the vicinityof engine No. 3 exceeded the recarded gas temperature value inthat region.

b. The surface temperatures appear to "veak out" at approximately370°k (206°F) while the transducer is set for a maximum yalueof 590°K (482°F).

c. The "peaks" have a different duration, and several of theindicated values finally droo below the value recorded prior toengine No, 2 failure.

4 similar trend was exhibited by the base region pressure instrumentation,as shown in Figures 16-13 and 16-14. Therefore, it appears that the baseheat snield forward face temperature transducers indicate the correcttrend, if not tre correct level.

In general, base region temperatures at lifto*f were colder for AS-502than for AS-501. This was due to controlling the oneration af theengine compartment conditioning system to a Tower temperature duringprelaunch operation. AS-502 ambient, structural, and componenttemaerature characteristics were similar to those for AS-S01 for theFirst 225 seconds of flight. After 225 seconds, data from AS-502ambient, strictural and component temperature transducers lacated in theregion forward of the heat shield and outboard of engine No. 7, as notedin Figure 17-12, show an abnormal cooling zrend ‘ndicating a cryogenicteas near No. 2 engine.

The normal S-11 boost temperature history trend far ambient temperaturesin the region forward of the heat shield is shown by the AS-501temperature data in Figure 17-12, The amb‘ent temperatures increasewith the interstage on and then decrease after interstage separationalong with a decrease in base heating. Ambient temperatures forwardof the heat shield leveled out and began increasing 250 seconds afterAS-501 liftoff. On AS-502, the tenperatures followed the same trendsas the AS~501 temperature data until 225 seconds after liftoff when,

17-16

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VW7-19

instead cf leveling out and increasing, they began decreasing at anincreasing vate as shown in Figure 17-12. AS-502 hydraulic systemtemperatures for engine No. 2, container 206A31 equipment mounttemperature and center engine beam temperature also showed coolingtrends between 250 and 290 seconds when compared with reszective AS-501data. Since AS-502 hydraulic system temperatures for all enginesexcept engine No. 2 appeared normal and the only container tenperaturewhich appeared abnormal was a container Tocated on the thrust coneover engine No. 2, it is believed that a cryogenic leak occusred nearthis engine during the flight.

In the region forward of the heat shield and below the thrust cone andLOX tank, flight data show soikes occurring in the heat shield Sorvardface temperature. The ambient temperature in this region, shown inFigure 17-12, exhibits a spike at 412.6 seconds, just prior to thetime of engines No. 2 and 3 cutoff. These tenperature sp‘kes correspondto the time of pressure increases discussed earlier in paragraph 16.3.2.Smaller temperature spikes were noted on the thrust cone, Thesetemperature spikes are believed to have been caused by release of hotgases from engine No. 2 just prior to its shutdown. Thrust conetemperatures also spiked at second plane separation due to the increasein base heating exptained earlier in this section.The temperature spikes indicated by structural temperature measurementswere probably more of an indication of transducer temperature rather thanactual structural temperatures.

Center engine beam temperature, and LOX tark external skin temperatureappeared to erratically fluctuzte between 412 and 455 seconds. Sincethe heat capacities of these structures ave large, it 1s doubéful thatthe recorded temperatures were actual structural temperatures. It ispossible that the transducers separated from the structure to whichthey were originally bonded, in which case, a very small change inheating and cooling would cause the fluctuations and these transducerswould then irdicate arbient temperature after debonding.Temperatures recorded during the AS-502 flight on the aft face of thebase heat shietc are compared in Figure 17-13 with design, AS-501 actualand AS-502 postflight predicted temperatures, The AS-502 data were wellbelow the design and compared favorably with the AS-501 data and AS-502postflight predictions. The lower temperatures in the 4S-501 band weredue to the added heat capacitance of a special steel transducer mountused on AS-501. The actual AS-502 tenperatures deviated from the actualAS-501 data at the time of engines No. 2 and 3 shutdown (approximately413 seconds) due to a decrease in base heat rates presented eartier inthis section. The naximum AS-502 actual temperature af 736°K (865°F}Occurred at 413 seconds, and is in agreement with the maximum of 741°K(876°F) recorded on AS-801. The design temperatures were calculatedusing the maximum design environment, and the postflight predictionswere based on postflight oredicted heating rates.

17-20

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S-II heat Shield Aft Face and Thrust ConeSurface Temperatures

17-21

Thrust cone temperatures were Slightly lower on AS-502 than on AS-501as shown in Figure 17-13 which presents a typical thrust cone surfacetemperature measurement. This was due to Tower base heat rates onAS-502. Design and AS-502 postflight predicted temperatures are alsoshown in this figure, and were based on the maximum design environventand postflight predicted heat rates

V7.4 S-I1/S-1VB SEPARATION ENVIRONMENT

During separation, the retro motor heating rates of the S-IVB structurefor AS-501 and AS-502, listed in Table 17-1 were similar, indicatingonly small differences in the plune impingement environment. Dat:from the calorimeters indicate that the heat flux to the J-2 enginewas somewhat higher on the AS-502 than on AS-501 {but three to six time:lower than that experienced on uprated Saturn 1 flights). Predictedheat fluxes are also shown in Table 17-1 and indicate that they werewithin the maximun envelope expected

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17-22

17.5 VEHICLE AEROHEATING THERMAL ENVIRONMENT

17.8.1 S-LC Stage Aeroheating Envi ronment,

The aerodynamic heating environments were measured using thennocoupleattached to the backside of the structural skin an the S-IC forwardskirt, intertank, engine fairings and fins. Generally, the aerodynamicheating environments and, consequently, che skin temperatures werewithin prediction bands, below design limits and slightly higher thanAS-501 flight data

Figures 17-14 through 17-16 show comparisons of AS-S0} and AS-502 skinterperatures and the heating rates derived from these temperatures .Post*light sinulated skin teuperatures are ‘ncluded for comparisonThe sicwlatign o* the “in skin cemperatures includes 0.284 watt/cm?(0.26 Bru/fte-s) for plune radiation, but did not consider effects offlow separation. Skin tenperatures on the forward skirt remainedat a nearly constanz level throughout powered flight as seen in Figure17-17. The slight downward trend of skin temperature until about80 seconds follows the trend of the forward skirt compartment gastenperacure. At 80 seconds, the free stream recovery temperaturebegins io rise rapidly causing the slight upward trend in skin tem-peratures from this po‘nt on.

Intertank temperatures closely followed the trend of AS-501 flight dataAs shown in Figure 17-17, intertank temperatures were within predictedvalues throughou: flight, and initial temperatures were slightly lesthan ambient due to the cooling effect of the gas in the intertankcompartmenz. The cool‘ng during the first 70 seconds of fight followedthe trend of compartment and free stream ambient gas temperatures. At70 seconds, the ‘ntertank area started to respond ta aerodynanic heatingand temperatures continuously increased until separation, reaching amaximum 343°K (158°F), The results of an integration of the calculateheat‘ng raves on the intertank and fin indicate that the AS-502 vehiclereceived a slightly higher tota’ heat input than AS-501

The aerodynaric heating to the body, fins, and engine fairings wainterrupted at approximately 110 seconds by flow separation. The flowseparation results from expansion of the F-1 engine plumes and, con:sequent'y, hot gases are recirculated up the vehicle side. The aastemperature and heat-transfer coefficient in the separated reafon areexpected to be less than those which would have been experienced ifseparation had not occurred, The temperature increases during separa:tion were most likely caused by the radiation from the hot recirculatedgas. The hot gas radiates energy because of the high emissivity of thecarbon partic’es present in the recirculated flow. The resulting changein the heating rates can be noted in Figure 17-14 through 17-16 between110 and 135 seconds. The increased heating due to flow separation didnot constitute a detrimental heating environment. Flow separation andsubsequent hot gas recirculation up the side of the vehicle has been

37-23

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17-27

Noted in AS-501 and AS-502 Flight optical data. Flow separation wasgoserved an the Satura I and I8 flights and was anticipated on theSaturn ¥.

flow Separation results from the expans‘on of the F-1 engine plumes atthe higher altitude. The plumes create what could be considered 2 salidwall to the oncoming Free stream. At the lower altitudes, the freestream flow can be deflected around the exhaust plume ty the extecralplume shocks but as the plumes increase in size, the free stream Flowcan no longer turn near the plume Surface. Consequently, the flow onthe side of the vehicle separates or begins to turn before it reachesthe plumes. Figure 17-18 iTlustrates the flow field which is obtainedafter separation occurs. Hot gas is fed into the separated reaionfron, the base region and plure interface. The base region hot gasresults from the ergine exhaust flows impinging upon one anothes andforcing some of the exhaust gases toward the base heat shield.Separation, once induced, will continue until outboard engine shutdown.As the plume diareter expands with altitude, the point of flowseparation moves forward along the vehicle.Flow separation on the AS-501 and AS-502 flights was first observedbetueen 105 and 110 seconds. Measurements have been made of the pointof flow separation for various flignt times and are shown in Figure17-18. Tt is noted in this Figure that the separation region extendedbeyond tne too of the S-I¢ forward skirt. just prior to stage separation,The observed blackness on the stage may be a carbon deposit rather thenpaint being burned.

Temperatures on the aluvinun portion of the fairings (forward of theheat shield) fell within a narrow band reaching a maximum of §73°K(872°F) at the end of powered flight as seen in Figure 17-19.Temperatures at the end of flight were about 40°K (72°F) hianer thanthose attained on AS-501. This was due to the higher heating trajectoryflown by AS-502 as seen in Figure 17-20 which compares aerodynamicheating indicator for AS-501 ond AS-5CZ. The initial rise in temperatureat zero seconds is attributed to iritia? incident plume radiation.Predicted maximum temperatures were not exceaded on the forward fairingand no severe temperature gradients were observed.Temperatures cn the titanium portion of the fairing are shown in thelower graph of Figure 17-19. Tenperatures followed the expected trendand are well below the predicted maximum. The slightly higher tenperaturesexperienced on AS-502 compared to AS-50? were to be expected due to thehigher aerodynamic heating trajectory flown by AS-502 and the slightlymore severe base environment. The narrow flight data band indicatesthe lack of any severe temperature gradients on the aft fairing.Temperatures on the thrust structure followed the expected trend andaTthough the prediction was slightly exceeded at 125 seconds, thispresented no problem as the temperatures were within design capability.Since aerodynamic heating effects in this area are thought to be less

17-28

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17-29

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17-30

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severe than the design levels it is concluded that radiation fromtheplume and burning gases in the separated flow region accounted for thefact that temperatures execcded the predictions in this area.

Temperatures on the wedge section of the fins were about 60°K (108°F)higher on AS-502 than on AS-SOT as shown in Figure 17-2) due to thehigher heating trajectory. However, the predicted maximum temperaturewas nat exceeded and the relatively narrow flight data band indicatesan absence of any severe temperature gradients on the fin wedge section.The initial wise in ternerature at liftoff was due to burning exhaustgases ard initial plume radiation.

Temperatures on the flat section of the fins were of the same order ofmagnitude ag those on the wedge section reachirg 622°k (680°F) at the eneof flight as chown in Figure 17-21. Predicted maximums were exceeded crlynear liftoff where a sharp rise in temperature was observed dle to turningexhaust gases and plume radiation. Convective cooling occurred untilabout 70 seconds when recovery temperature began to rise and cerodynamicheating effects caused a skin temperature rise. for the most part, highertemperatures were obtained on the fin flat section on AS-S02 than on AS-501again due to the higher heating trajectory flown by AS-502. No harmfultemperature gradients were observed an the flat section of the fins.

W7-31

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17-32

7.5.2 S-1t Stage Aeroheating Environment

Aeroheating rates on S-IT stage and its protuberances were analyzedusing the AS-502 Final Observed Mass Point Trajectory (OMPT) and angle-of-attack data obtained from the S-IC Flight Control Conditioned DataTape (Q-ball). Atmospheric data were obtained from the final meteoro-Jogical data tape.

The aeroheating rates on the S-EI stage cylindrical surface and it:protuberances were calculated by means of an aeroheating digital computerprogram. The program includes turbulent flow flat vlate heating ratetheory and real gas therrodynamic and transport properties for air

The heating rates to the protuberances were obtained by increasing thebasic flat plate neat fiux by the aporopriate experimentally determinedprotuberance factors. The aredicted aeroheating rates were corrected tothe calorimeter conditions for aurooses of direct comparison. Thesetransient calorimeter heating rates were determined by first computingtha sensor temperature corresnanding to the nominal aeroheating ratesand then applying the temoerature mismatch correction to obtain theFinal sensor tenperature. The finat indicated calorimeter heating rate:were obtained from the calorimeter sensor-body temperature differentialsand the calorimeter calibration curves.

In general, the postflignt prediction transient heating rates were higherthan the corresponding fligat values. The validity of protuberancefactors used, which may have contributed to these discrepancies, areunder investigation in arder to improve the prediction techniques.

The heating rate design values are omitted from these presentationsbecause they are not corrected ta the calorimeter conditions. Since thestructural surface temperature response plays a key role in transientheat flux predictions, the comparison of design heating rate values forthe structure with the calorimeter indicated heat flux is meaningless.However, it has been established that the AS-502 flight aeroheatingrates were considerably lower than the design valves

The corparison of AS-502 flight and postflight prediction of the aero-heating rates experienced by the calorimeters mounted on the LHz feed-line fairings are presented in Figure 17-22, AS-501 flight data is alsoindicated and it is seen that it is enveloped by the AS-~502 flightdata. Reasonably good agreement of flight and postflight prediction wasobtaired for the reasurements installed on the fairing boattail sectionHowever, the predicted heating rates to the fairing nose cone sectionare considerably higher than the corresponding measured values. There-fore, the experimentally obtained wind tunnel mode? test protuberancefactors will reauire further odification in order to obtain improvedcorrelation.

17-33

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The comparisons of AS-&02 flight and postflight predicted aercheatingrates experienced by the calorimeters mounted on the fairing conicalnose sections are presented in Figure 17-23. AS-50) flight data erealsa skown for comparisen. Again, the predicted data are considerablyhigher than the corresponding flight values as was shown to be the casefor the LFp feedtine fairings.

The comparisers of AS-502 flight and postflight predicted values of theaerodyramic heating rates sensed by the calorineters installed on tieinterstage structure and first plane separation fairing are presentedin Figure 17-23, AS-S01 flight data, which envelops the 45-592 Flightdata, are also included. The predicted values are higher than theCorresponding flight values during the maximum heating portion of theflight. A qualitative disagreament exists between fligat and predicteddata beyond 125 seconds. The flight data at this time shows decreasingheating rates which approach a constant level, It is believed thatthis phenemenon was caused by the scorching of the surrounding corkinsulation. The scorching of cork insulation, due to the aercheating,was also observed in the photographic coverage of insulation testsperformed on the X-15 flights. This phenomena was not accounted forin the analytical postflight prediction, heace the resulting discrepancy,Tt shoukd be noted that tre same trend was obsevved in the AS-501 Flightdata.

17-34

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Figure 17-23. $-II Ullage Motor Fairingand Aft Skirt Total Heating Rete:

Representative structural, fairing, and surface tenperature measurement:influenced by AS-502 aerodynanic heating are show in Figures 17-24 and37-25. Each plot gives the actual flight data alang with desiqn andpostflight predictions. AS-507 flight data is included, where appropriate,for comparison. Design predictions are based on S-Ii stage contractordesign heating trajectory with a high anale-of-attack. The postflightpredictions were based on the heating rates discussed earlier in thissection.

The temperature data fron AS-502 are in general very similar to thatfrom AS-501. A slightly higher temperature environment and slightlymore severe trajectory resulted in hiaher temperature and a largertemperature rise for AS-502.

Figure 17-24 presents typical interstage structural temperatures.Postflight predicted and flight values are in good anreement.

Feedline fairing temperature and ullage motor fairing temperature dataare presented in Figure 17-24. The postflight predictions of temper-ature for both these fairings are higher than flight data. Revisionsin the protuberance factors used to calculate the aerodynamic heatingare expected to result in predictions that will ‘tore closely match theflight data.

Figure 17-25 presents the data together with design and postflightpredictions for CO27-219 forward skirt skin temperature. The post-Flight prediction is only slightly lower than the flight data.

17-35

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isos 1M, SELON

Figure 17-24, 5-11 Structural Temperatures

LHy tank insulation surface temperature measurements for five differentloCatians are presented in Figure 17-25. The abrupt shift in themeasured values after the data dropout at approximately 150 seconds iunexplained so far, Up to this shift the postflight prediction andmeasured values were in good agreement. The significantly Tower valuesobtafned from measurement (893-218 were similar to data recorded orAS-601 and may be attributed to cooling due to venting of cold heliumgas within the ‘nsulation. Investigation into this and other anomaliesis continuing.

17.5.3 S-IVB Stage Seroneating Environment

The AS-502 aerodynamic heating environment was slightly more severethan AS-501. AT} temseratures were well below the design values forthe maximum heating trajectory. Figure 37-26 shows the data andpostflight simulation for four skin sensors and one stringer sensoron the forward skirt and data correlations for the LH tank measurements. All simulations use the design method of analysis exceptthat boundary layer transition is determined as the time at whichTwall/Trecovery = 0.5. The maximum recorded temperature was 384°K(232°F). The Liz tank measurenents were noted to have frost or iceto some degree at liftoff. Unlike AS-501, the frost appears to haveoersisted for this flight with the exception of one measurement whichrecorded the maximum temperature of 301°K (82°F)

17-36

450

STA63.6 m

z Gegesti) e_ 300a" we INTERWAL wr iy “ DESIGN . 3as TRANSDUCE! g i P20 SSZe 3c 2828 AS-501 ae

g = 300f 100 2 a

Ss” XposTeLIGHT é- PREDICTION 0

ee 250

HELIUM PURGEDTRANSDUCER” HOHEYCOMB_INSULATION

500os F400

aT oR

8 asop 2 8= nN x. &= DESIGN aBu PREDICTION 300 ae2 400 x 7 2".au LHe SuEe TANKWALL p20 ESSE 350 INTERNAL 3BE VA BS2g seal ae10 Le=i 300 4 ge2 POSTFLIGHT =

DROPOUT ——-{j] PREDICTION s= 260 po os \ ky.

200 35 BOE TOT208RANGE TIME, SECONDS

VurtorrYQ S-IG/S-T1 FIRST PLANE SEPARATION, 149.08

AY S-IC/S-11 SECOND PLANE SEPARATION, 179.06

891-218 AS 502 FLIGHT DATA— (892-218 AS 502 FLIGHT DATA

(893-218 AS 502 FLIGHT DATA~ C894-218 AS 502 FLIGHT DATA

——— (115-218 AS 502 FLIGHT DATA

Figure 17-25. S-II Forward Skirt Skin and Insulation Temperatures

17-37

OANA SeIRT Scum

were: SeaBOLe Dewey FLIT eTLies cewre Swe! simu.Laihit

son

Terme

aarueL

,

aiweenevane,©

Lrg Tan Say Tore: Ste Te Fattvit? SATOH S198

Figure 17-26. S~I¥B Aeroheating Environment,

17-38

Figure 17-27 shows data and correlatians for selected skin and stringerteriperatures on the aft skirt. All measurements were covered withKorotnerm and the analysis assumed 9 srotuberance heating factor of1.5. The maximum temperature recorded was 353°K (176°F} on the skinThe date and correlation for two adjacent measurements othe afinterstage, one on a stringer and one on the skin are also presentedThese measurements were also covered with Koratherm and the maximumrecorded temperature was 360°< (187 The data and correlation fortie measurement on the feadline fairing forebody is also shown. Themaxinum recorded temserature was 411°< (280°F)

Tae net heat transferred to the LH2 during boost was analyticallydetermined by twa mathods, A comparison between the results obtainedfron these two methods and the maximum and minimum design heatingvalues is shown in Figure 17-28. Curve 4 is a simulation using therecovery tenperature aad tan< wall heat transfer coefficient historiebased on tre flight trajectory. Initial LHz tank skin temperature:from fight data and maximum values of insulation thermal conductivity(k} as a function of teraerature, as determined from $-I¥B loading andacceptance firing test data, were used. The maximum and minimum designvalues of LHz veating are also based on the maximum and mivimum k curvesThe neat transferred tiraugh heat shorts (i.e., heating paths other thanthe cylindrical tank) was taken fram a recent $-I¥3 2rozel lant heatinganalysis. Curve B was calculated by integrating LHz bulk temeraturecnange resulting from aerodynamic eating during ascent, The Lllzheating values fall within tie design range.

17.5.4 Instrument Unit Aeroreating Environment

Tne Instrument Unit (IU) aeroheating environment was monitored by eightthernocauples mounted on the _inner surface of the ‘oneycomb structureon the low density, 49.7 kg/m (3.1 Thm/fts), core, Seven of the eightmeasurenents indicated temperature rises due probably to internaradiation or local convective heating during the first 30 seconds offlight, This is shown in Figure 17-29, The two sensors located nearposition Iv at station 82.47 meters (3247 in.) and station 82.14 meters(3234 in.) indicated increases of 5°K (9°F) and 8°K (14°F), resaectively

After 30 seconds, these measurements indicated a cooling trend. The IJcompartnent ambient gas temperature dropped to 283°K (50°F) at 80 seconds.After that tine the sensor output was somewhat meaningless since the com-partnent pressure was approaching 2.1 ¥/em2 (3.0 osia). The inner skitenperature indicated a maximum 369°K (189°F) at approximately 190 secondat the sensor Tocated near position I and a minimum of 339°k (151°F) nearposition If. The simulation in Figure 17-29 indicated a maximun externatemperature approxinately 24°X (76°F) higher than the inner sensortenperature for the no solar heating case. From 160 to 550 seconds, theeffects of solar radiation could be noted in the measured data. Thesimulation of the data was for no solar heating. The sensors located atpositions I and IV would have experienced the greatest solar heat flux;this was indicated in the measured data. These trends due to solarheating were noted in the AS-501 data

17-39

SeT s61Rt SeaWes SaaS DERTE AT son

ey

vWseeR

eTuRe

a,

sear seas

soe TRS CT Ten aT 7INES UEREIL Seg Suit asning

=

1 eno {VE Tete GS1 1caocusi 2 ee

B20, ~ ES -

3.

i | 7|

oy BET ar To weve SarRAAGE TINE, SECS

Figure 17-27. S-IVB Protuberance Aeroheating Environment, Sheet 1 of 2

17-40

APT INTERSTASE SHIN

500 FRTES SRL SOOTES UT ATE TET 7 TIne Genores Seve03 SaPCLE 10W cave L aocNm elo es 03So d3 wie 0s1 1

L300200 anf — + 7

2 awe

é : BE i roesob

p28ae

: a

208 roe

Leg FRRELINE FAIRING FoaEsnonORT SL ETRETEEUne Devores sensor somarscs | fae

hime tn €-28 GFBe 03 Twier a3

aaaota er ! °: aootEe ! Z5 o | Z

E te &3ce — Baa

6

2 -oe oe a5 ise 70 ranSe “IME, SCN

Figure 17-27, $-1¥8 Protuberarce Aeroheating Environment, Sheet 2 of 2

17-41

189 OTE GET STORTS THEUUED Th

2) bh! AO WATT-SLC (547estan "uaLues AS of isBU) Bo S20aoOsT PERN SEC. PE PRESS,

ptr

se

109 7 a yoo ODDSRAMEE TIME, SECEIDS

Ie

Figure 17-28. S-IVB Lig Heating During Boast

17.6 VEHICLE ORBITAL HEATING ENVIRONMENT

17.6.1 $-1¥B Orbital Keating

The orbital temperatures for the Aux? liary Propulsion System (APS)were determined by ten sensors mounted internally to the APS fairingon various components and propellant transfer lines and four sensorsmounted on the fairing.

Table 17-2 lists the maximun and minimum temperatures during the flightfor the APS measurements. The values for the AS-501 flight are includedfor comparison. It can be seen that all components, with the exceptionof the propellant transfer lines, remained within their allowable limitduring the mission. Thermal radiation from the APS engines caused fourline measurements to exceed the maximum alloweble temperature of 324°K(125°F}. Greater than expected APS engine operation was required dueto anomalous flight conditions which accurred about the time of theattempted restart of the S-IVB J-2 engines.Figure 17-30 shows a comparison of the fairing flight cata with resnectto the predicted desicn temperature envelore. The fairing data falwithin the analytically predicted maximum and ririrur values. Acorrelation of @ measurement located an the APS fairing is aiso shownFairly good correlation was achieved far the first two revolutions .After this time, it is believec that the Flight simulation parametersused to establish the vehicle orientation did not match the actualflight parameters.

17-42

17-43

Mo “FUNLva ada

SIMULATION

Or

EXTERNAL

T=

INTERNAL

ANNOY5-502

DATA

BAND

:

€41-602,

ca2-602

44-603

39-601

Iv+

375

Os,

043-603

40-601

c3e-601

|

c37-601

I

350

326

240

be200

b160

Lizo 80

300

278

Ay0

4080

120

160

200

300

400

500

RANG

ETIMC,

SECONDS

Figure

17-29.

IWInner

Skin

Temperatures

for

Ascent

60049

de ‘3UNHdW:aL

Tabl

e17

-2,

APS

Orbital

Temperatures

MEAS

UREM

ENT

nO.

[LAUNCHTEMPERATURE

°K(°F)

[MAXIMIM

TEMPERATURE

INFLIGHT

"X(°F)

|MINIMIM

TEMP

ERAT

URE

INFLIGHT

°K(°F)

AND.

DESCRIPTION

ORLO

CATION

MEASURED

wxHEASURED

—evcecy

win

MEASURED

excen

EXPECTED

as-s

ou[as-so2

|

cimtt

{asso

[as-soz

|rimes

|

urmt

t[a

s-so

r

|

as-so2

|

TIMEY

17-44

FAIRING

TOPE

Fore

body

O28?

Right

Side

0288

Too

C028

9Left

Side

INTERNAL,

“CO2SE

Fuel

Line

£029

5Fuel

Inlet

(Right

Side

)C0

296

Fuel

Inte

t(L

eft

Side

}0297

OxLine

296(73)

2932.a

286

(35'

sost

ea)

303¢

a6)

(326

¢125

)30

3tas

)

|

303(

86)

|324¢125)

303(

86)

|

303(

@6)

|a24(125)

30286:

303(

86'

303(

86:

303 (

86:

$35(143)

/1€,800

266(20)

303(86)

305(90)

326(127}

|16,850

|

266(20)

}303{86)

|

305(90)

beocai6)

266(20}

|300(8

0)308(88)

334(14

1)106

,800

|

266¢20)

|

303(86)

|

305/90)

28(130)

|16.850

|

266(20)

|302(84)

|

304(88)

16(110)

266{20}

|302(84)

}305(90)

11¢006)

266(20}

|305(9

0)|30

5(90)

0895)

|

=266(20)

|

300(20)

303(86)

|324(1

25)

|

316¢170}

pr3¢10

3)2s6(20)

|305(9

0)|30

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|

208/90)

fost9s

)266(20)

|293{6

8)|30

4¢¢0)

303(a6)

|324(125)

6029

8Ox

Line

(Left

Side

)60

299

OxIn

let

{Right

Side

)€0

300

Fue?

Tank

Supp

ort

C0301

OxTank

Support,

0302

Fuet

cont

.Mo

d.€0303'Dx

Cont.

Hod.

303¢ae)

|

320125)

303(86)

}324(125)

30a(s8)

|320(125)

303(

96)

+/32a(125)

303(

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3030

‘Time

atwhich

temperature

Vinit

tsexceeded

(available

date)

tHData

**Apparent

failure

during

boost

ot

a

al

“PERATURE,

*<

20a

esiAMLSTEIAL <#RE.ATIO

| peo sare | sonuca |L

rATURE

,

ayo coos co2ag

DESTGH HaKiMUM 200

too

tea

ne

‘. a0

"onoaniseo{1 Lsioo | [omerT———~_ fe

Fs 73,030 1,00 we.e00Ge TEM, SCONES

1 1 1.TOE Tana Saba rae Tae{GCE TENE, OURS: IATHDTESSeCcACS

Figure 17-30. S-IVB APS fairing Temnerature

17-45

Tigure 17-31 shows a correlation of a measurenent located on treoxidizer control nodule filter bedy housing. The measured and pre-dicted values compare fairly well and were within the ooerationallimits for the component. Thernal radtation from tha 89S engines(brought about by the heavy duty cycle imposed by the anomalous flightconditions mentioned previously) caused the flight data and analyticalvalues to diverge during the portion o* the missian after 13,000seconds -

17.6.2 IU Orbital Heating

The IU inner skin temperatures are shown in Figure 17-32 for the first36,009 seconds. The time szent in the earth's shadow is shown by thevertical shaded bars. The highest skin temperature experienced duringascent was 15°K (27°F) higher than that of AS-501. This was prinarilycaused by a 25 percent highes atmospheric density at the higher alt‘ -tudes during AS-502 ascent. Also contributing +0 th’s heating effectwere the higher initial temperatures and a 4 percent higher velocity forAS-502 from 60 to 150 seconds.

Based on the time of year far launch, AS-502 was predicted to be in asomewhat colder orbit than AS-501. The maximum inner-skin temperaturemeasured in orbit for AS-502 was 325°K (125.6°F) compared with 363°K(194°F} for AS-501, Figure 16-29 shows the inner-skin temperatures forAS-502. The large temperature excursions prior to 22,112 seconds werenormal, After 22,712 seconds, APS control was Test and LOX ventirgcaused high pitch and yaw rates, This motion of the vehicle tended tedamp out these temperature excursions.

WOE: VEMVURAILRE Doers LATTER Sages 9a FIGHT ni Luesce™ be MEMALOAS LEGITca03 24a: 104s SOT REESEAMET UCSL CORMEL ET UN CORELAT =O,

Temmenaune.=

“TeqPe

RATURE

[svoj | seca [sae |° sou 17,300 TS 72,009

ase TIME, seems

L L 1Tn F000 TaN Tae Tt a

MNGE TIME, -wURs PISUTES: SECONDS

Figure 17-31. S-IVB APS Pronellant Contral Module Temperature

17-46

it Inst shan 93.2fe mre Sem WI

HT & 1i tsar setateCSelSE

ane TaMr, Sec0MDE 103

poLaas PRT TM tatbda oss a0 ae eNaba:8boy aA 20(GE TIME, HUES HIAUTES:stctNaS

Figure 17-32. IU Inner Skin Temperature in Earth Orbit

7-47/17-48

SECTION 18

ENVIRONMENTAL CONTROL SYSTEM

18.1 SUMMARY

The S-1C canister conditioning system and the aft environmental condi tion-‘ing system performed satisfactorily during the AS-S02 countdown withonly one canister and one ambient temperature measurement dropping belowthe minimum requirement.

The §-I1 thermal control and compartment conditioning system maintainedtemperatures within the design limits throughout the prelaunch operations.

Temperatures monitored on the S-1VB aft skirt components were slightlycooter than on the AS-EC1 flight but within design limits. Temperaturesof all components mounted on the forward skirt cold plates were withindesign limits at liftoff.

The IU Envirormental Control System (ECS) performed well throughout theflight. Coolant temperatures, pressures, and flowrates remained withinthe predicted ranges ard design limits for the duration of the flightdata. One specificaticn deviation was observed which was expected. At11,670 seconds, the platform gas bearing pressure differential was 0.069Néem2 (0.1 psid) above the 10.7 N/cm2 (15,5 psid) maximum allowable andvemained there throughout the remainder of the flight period for whichdata is available (33,780 seconds).

18.2 SIC ENVIRONMENTAL CONTROL

The 5-1C stage ECS is usec tc centrol terperature in the instrumentattoncanisters of the forward skirt compartment, and thrust structure compart-ment during preflight operations. The conditioning and purge agent (airuntil -8 hours and 30 minutes, GN2 thereafter} 1s provided to the stagefrom a central ground supply.

The S-IC canister temperatures remained within the required Timits duringthe countdown except for canister No. 6, The temperature in this canisterdropped slightly below 288.7°K (60°F) during the last minute of thecountdown. It reached a minimum of 287.9°K (58.5°F) at -16.7 seconds(forward unbilical disconnect).

18-1

The other canister temperatures varied fron a maximum of 300.7°K (81.5°F)to a minimum of 289,2°K (60.8°F) during prezaunch activities as shown inTable 18-1.

Table 18-1. S-IC Environmental Control System Canister

Temperatures

MAXIMUM MINTHUNTEMPERATURE °K (°F} TEMPERATURE °K (°F}

Canister No. 1 C212-120[ 296.2 (73.4) 289.2 (60.8)Canister Ho. 2 €213-120 295.2 (1.6) 291.2 (64.4)Canister No. 3214-120] 300.7 (81,5) 298.9 (78.4)Canister No. 4 €215-120] 299.2 (78.8) 295.4 (72.0)Canister No. 6 217-120] 294.7 (70.7) 287.9 (58.5)

Extremes from -20 min to -16.7 sec The most difficult control condition occurs during J-2 engine chill¢ownwhich starts at -8 minutes and continues until umbilical disconnect.Within this t'me period the anbient temperature in the interstage(Forward compartment) area dropped as shown in Figure 18-1, All thetemperatures were above the 206.4°K (-90°F) predictcd minimum exceptfor (207-120 which reached a minimum of 168.2°K (-187°F). This tempera-ture probe is located under a J-2 engine and receives the maxinun effectof the cold helium.

Temperature plots during flight are not presented because the compart-ment and canisters are only conditioned until -16.7 seconds. A bandof canister temperatures versus time is shown ir Figure 18-2.

A characteristic aft compartment temperature is shown in Figure 18-3.Temperature extremes for 11 measurements are given in Table 18-2,All temperatures were within the required limits of 299.8 £11,2°K{80 420°F} except that C-202-115 was 3.5°K (6.5°F) below the minimunrequirenent prior to liftoff. This instrument is located at Position IIIfacing Position IY. This quadrant has two LOX suction ducts passingthrough it and C-202 is near the inboard suction duct. The maximumtemperature recorded wes 301°K (82.4°F) at instrument location €204-115.The flight batteries ore located 26.5 degrees from Position J towardPosition II, There is no telemetered instrumentation in this quadrant.

18-2

Figure 18-1. S-IC Environmental Control Systems Forward CompartmentCanister Conditioning System

18-3

a6

“p90

faS=H0e FLOAT DATA Banone

¥ e215- 120é [ e217. 320, Z .3 i gom y Ae := Wy é

ei 7

236,

501 aa 2 Ta “280 “a 2

TNE TIAL, SECIS

EE IS RTRANGE TIME, HOUS: MINUTES: SECONDS

Figure 18-2. S-IC Forward Skirt Canister Temperatures

Table 18-2, S-IC Environmental Contral systemAft Compartment Temperatures

[TIME AFT COMPARTMENT TEMPERATURES °K (°F)JSECONDS C107 toa C202 €203 C204 C205-240 298(77.0))293(68.0)1285.5(54.5)|290.5(63.5)| 301 (82.4) |298.5(77.9)

0 298(77.0)/293(68.0) |285.0(53.6)] 290.5(63.5) 301(82.4)|298,5(77.9)63 268 .0(23.0) |6 275.5(36.5):70 288(59.0)75 283.0(50.0)80 283(50.9))278(41.0)

18-4

905 1 U6)

me 203-78Ney Stai 36.9Hea tn

cz02-119ven stasn 7sz in)Ss

FEAT SHI

pus

a. <delye71s Ne y/ ue nyt‘ cathjen y),P adie Ins

Whe——.

—eesence onRectiaitPas ¢

POS CUE aMBLL Ste

Pos mos 11

75

Figure 18-3. $-IC Environmental Contra] System and CompartmentTemperatures, Sheet 1 of 2

18-5

11fore nto

40m Co a

|

10 Sia Ta a ETSpafae Gia eh *F yoo. +

z s

8

0mmBOpeeTe, sts

ool eee TT

»

* oe Spent ts at5 5E os a3° ae

zac}

es

285,“SET

- soz

Figure 18-3.

a)

TET

20

SUAENT

TES “TERE” FORE SaANSE TIME, SECIRDS 0)

too

at TE

aeRANGE TAME, OAS: AIMUTES SECONDS

S-IC Environmental Control System and CompartmentTemperatures, Sheet 2 af 2

18-6

Located on the battery container are four thermistors which control theanbient tenperature, near the batteries, within the limits of 299.8 +2.8°K(80 45°F). Figure 18-3 presents the temperature recorded at instrumentlocation 12K10, which is a hardwire measurement. The compartment tempera-ture was maintained within the required limits until LOX loading. AfterLOX loading until umb‘lical disconnect, the temperature decreased to291.9°K (66.0°F}. This is 2.2°K (4.0°F} below the compartment requirementsof 299.8 -5,6°K (80 10°F). All interface requirements of flowrate, pres-sure, and temperature were met by the ground support equipment.

18.3 S-I1 ENVIRONMENTAL CONTROL

The S-II stage Environmental Control System consists of two parts.

a. The engine compartment conditioning system provides a means ofpurging the engine and aft interstage area of explosive mixturesduring propellant loading operations, and maintaining proper tempera-ture control for stage components, The compartment purge is effectedby means of warm Gly and is operational only during the prelaunchperiod. The compartment vents have been designed to meet theseobjectives and to reTieve internal pressure during S-IC boost.

b, The thermal control system is designed to pravide both temperaturecontrol and an inert atmosphere for the electran{c equipment.

containers in the aft compartment. Graund equipment providesconditioned air for cooling during ground checkout, and GN2 forpurging and heating during and after propellant loading. Theconditioned gas is directed to the equipment containers throughducting and exhausts to the interstage area. The flow is fixedby orifices and is continuous until terminated 2t umbilicaldisconnect. During flight, thermal inertia and cortainer insulationwill preclude out-of-tolerance ecuipment temperatures,

To verify that the thermal control system provided adequate temperaturecontrol in the electrical equipment contairers, measurements of theequipment mounting surface temperature were made in all the containers,As shown in Tables 18-3 and 18-4, all the equipment mount temperatureswere well within tre required temperature limits during both theprelaunch and launch operations.

Equipment mount temperatures in the forward and aft systems were within7k (13°F) of their corresponding liftoff temperatures on the AS-501Flight and showed the same type of temperature profiles throughout theAs-£02 flight with the exception of container 206431, The equipmentmount temperature of this container was very similar to the AS-501 flightup to approximately 325 seconds. From 325 to 425 seconds, the containerequipment mount temperature dropped approximately 7°K (13°F) and continued

18-7

Table 18-3. S-1] Forward Thermal Control System AS-501 and AS-502Prelaunch and Flight Data

S211 fuNGH 5-501 eo TPMERT MOUS TEMP seston Tepe. eeet"s [temp ac staer] rene. av

|

rene, ar en

MCASIREMENT ant MUNICH vexinn [tata[OF BIB. ott.

|

Latsen

|

OF sic seostConeainar 220, C356. 200 Sa (140%

|

256 (24 wee ehGontainer 221, c3aG-zel

cx

{*F)] 34 (eo)

|

ee (at 282 (68)

|

ass festro] a (aay

|

es (0) ase (55)

|

2s 5)Sp 334 (140)

|

a5 :-20) 23 18} [20 Her)Container 225. €359-225 ¢k {FH 334 (140)

|

256 (ay Sut fee)

|

301 ae)ankerar 2a 8-227 oh UE oe (MD)

|

se Eh mae f52)

|

ani (60sGartsiner 224, Eoi0-z28 k (ert 13e (128!

|

350 (6) 230(62)

|

ma {8]BxTeno. at tscomece °K (FA) 320 (113)

|

S14 (03; Bia)

fF sys. Ghy Flow tars (lems) 302 {.e66)] 287 $ seh =SoC? Wane 5-502

Genteier gate eat-zey oy ene (ia) Tae (oy Paee ent Be By Te Bo sez)Eontasnar 225, C356-222 ck {oF soa frauy | ate {o) fet (sat

|

gas (sal

|

a: ier (30}Container 222, 361-222 °k (°F) x04 (140)

|

2p (0)

|

263 (50)

|

Zee (eo) [228 22 (48)Contsiner 225, Col l-z23 <x (>) sae (14D)

]

Zefoe0)

|

bar tae)

|

dap (aa)

|

2a 20 {4}Container 228) C358.228 age (nao)

|

256 15)" are fay.

|

28 an)

|

297 gr)Contatner 225, C359-225 gee ful Yaar cy ane cea) |ie)

|

3h Sos)Sontstver 227, Cal9-227 ass (v40)

|

zee (ay

|

2r9 (4a)

|

2a Fea)

|

ra zo fantfontatner 228: 510298 zs (125)

|

256 (0

|

297 (57)

J

2a '29) | 2a 32 Tem. at Biscomect *< tri] 320 (mas)

|

374 (tom: F316 (308),

|

319 ftogy| TT |Sy Oe Flow cars (Tours) 482 (666) [267 (-567;). 206 (025; 27s Leoey 2

to drop another 4°k (7°F) during the remainder of the flight. This compared to a drop of only 1°k (2°F) for the same measurement over the saretime span on the AS~501 flight indicates the presence of cryagenics inthe container area. Considering the proximity of container 206A31 toengine No. 2, together with the fact that none of the ather containerswere affected, would indicate a cryogenic line failure in the engineNo. 2 area. Plots of container 206A31 equiprent mount temperaturesduring both AS-501 and AS~502 flights are shown in Figure 18-4,

The engine compartment conditioning system maintained the compartmenttemperature Within the design limit throughout the prelaunch operations.The thrust structure wes below 256° (OF) at. launch which met the designrequirements. A comparison of AS-801 and AS-602 engine compartmenttemperatures at three time frames during both prelaunch and flight periodsis presented in Table 18-5. The measurement locations employed throughoutthe engine compartment are given in Figure 18-5. The interstage tempera-ture control thermistors were set to control to 275°K (35°F) in comparisonto the 278°K (40°F) value employed for the AS-501 launch. This resultedin slightly lower temperatures throughout the interstage at the startof engine chill for AS-502 than during the AS-597 launch; however, thetemperature history characteristics through engine chill and 5-IC boostwere quite similar to the AS-501 flight data, There were no indicationsof abnomatities through S-IC boost.

From the initiation of tanking tarough launch, there were no indicationsof hydrogen or oxygen in the engine compartment.

18-8

18-9

Tabl

e18-4.

S-II

Aft

Ther

mal

Control

System

AS-501

and

AS-5

O2Prelaunch

and

Flight

Data

CONTAINER.

AND

MEAS

UREM

ENT

NUMB

ER.

Aulung. 2g+ du,

notsAdan 249

Dest

anTemera

ture

Limes

faxt

rum

CF)

Ninigun

oF)

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ass a - pe1

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RANGE TIME, SECONDS:

Figure 18-4. Container Equipment Mount: Temperature

Table 18-5. S-II Engine Compartmert Temperature Data Comparisonof AS-B01 anc AS-502 Flights

[~ 55-204 ~ *i ease Te : fae ti

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18-10

Set ta Mase

179-26cane

eana.2e9

cane 29ene 120 SIC Lox Tae

Figure 18-5. S-II Engine Compartment Conditioning systemTransducer Locations

VSPA LOOKING -oRWARD

18.4 S-1VB ENVIRONMENTAL CONTROL

The forward skirt area of the S-I¥B is conditioned by the IU ECS. Theaft skirt and interstage ECS provide the following:

a. Thermal conditioning of the atmosphere, during ground operations ,around electrical equipment in the aft skirt.

b. Thermal conditioning of the Auxiliary Propulsion System (APS),hydraulic accumulator reservoir, and ambient helium bottle.

¢. Purging of the aft skirt, aft interstage and thrust structure,and the S-II stage forward skirt of oxygen and combustiblegases.

Tenperature-controlled air or GN2 is supplied at the rate of 3500 SCFMto accomplish the thermal conditioning. The Gkpurge is initiated atLOX loading and is continued until umbilical disconnect.

38.4.1 Ascent Powered Flight Phase

The temperatures monitored at launch for the aft skirt components, mountedon fiberglass panels, ranged from $.6 to 11.1°K (10 to 20°F) cooler thanthose recorded an AS-501, The forward skirt components, mounted on coldplates, ranged up to 9°k (16°F) wanrer than those on A5-501. AT terpera-tures at liftoff were well within the components’ upper and lower tempera-ture limits.

18-1

18.4.2 Parking Orbit Phase

The data monitored throughout the flight for the bridge modules, prozellantutilization, and static inverter assemblies appear to be valid. Thetemperature extremes recorded during the first 23,800 seconds of flightave presented in Table 18-6.

During orbit, al components remained within their temperature Timitsduring the first 24,000 seconds of flight, The cnilldown invertersexceeded their lover temperature limit at this time. This is not. con-sidered to be a problem because the chilldown inverters’ nomial operationperiod extends only over the first 16,200 seconds of a nominal Saturn VFlight. The propellant utilization assembly exceeded its low temperaturelimit at 33,400 seconds. This is attributed te the cold plates beinginoperative at this time and is not considered to be a problem,

Table 18-6. Forward and Aft Skirt Component Temperature*

=k (T HT SEP ok Co 7 er) ET seesonenenr 0 L bat ten ok (°F)

|

wan FLINT time CF) [Mu FLIGHT TENE FDAND DESCRIPTION

|

ex>ecTeD |__WEASURED LIMIT MEASURED Litt MEASUREDOR LOCATION j FESOH AS-0E FSBO

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ASST FESO

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AS-50E7

Lox crsttdoon

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275 — 300} 294

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559

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et“During First 35,800 Seconds of Flight.

18-12

Ground hold and orbital data, through the first 16,000 seconds of flight,ave presented in Figure 18-6 for the chilldown inverters. The result:of the postfligat analysis, simulating the environment experienced bythe LOX inverter, are also presented. As indicated in this figure thereis satisfactory correlation between the flight data and the analyticaresults.

LO aN LH=

AL © Lag INGERTER

VAL LOX INVERTER THTERAL

x0 eo

; mm pAR «; good 22 *e 8 Aa ad 2 ag . gBa 4 i gz POSTFLIGeT aa @ 40i maarhy 2 5

27 —: Pout ‘

} ies 2G

269 4a a a ToT TTTave TIME, SECOOS

5 Tab Tae TN Tao Fo

RANSE TIME, HOURS :MINCTES: SECONDSFigure 18-6. LOX and LH, Childown Inverter Temperatures

18.5 IU CNVIRONMENTAL CONTROL

Prior to launch, a purge duct, located circumferentially above the IU/spacecraft interface, uniformly distributes tenperature-condi tioned airfrom a ground supply prior to fueling, and tenperature-conditioned GN2subsequent to fueling. Four sensors, located in the IU, monitor thecompartment terperature and feed a signal to the ground equipment tocontrol the inlet gas temperature which maintains the compartment tempera-ture within the specified Timits, 290.2 to 295.8°k (63 to 73°F).

The Thermal Corditioning Subsystem (TCS) provides temperature-controlledheat sinks to which the electronics reject waste heat. A coolant pumpcirculates the 60 percent methanol-40 percent water by weight coolantfluid Methanol/water (M/W) to the IU and S-IVB TCS. Coolant distributionis controlled by fixed orifices within the IU to 16 coldplates and 4internally cooled components: the ST-124M-3 Inertial Platform Assembly,Flight Control Computer (FCC), Launch Vehicle Data Adapter (LVDA), andLaunch Vehicle Digital Computer (LVEC), The IU and S-IVB stage returnFlows converge before entering a heat exchanger assembly, which is com-posed of a preflight heat exchanger and a sublimator as shown in Figure18-7,

18-13

Pressure is supplied to the ICS through regulators and filters to a M/tlaccumulator from a 2068 N/em? (3000 psi), 2700 cm3 (165 in.3) sphere. TheM/W accumulator prevides for fluid expansion and contraction, compensatesfor fluid losses due to leakage, and iraintains coolant pump inlet pressureAn orifice regulatcr further reduces the GNz pressure to an acceptablerange. This pressure expels waler from the water accimulator to thesublimator. The GNz continually vents tu the Vehicle compartment throug}the orifice regulator.

Quring prelaunch operatian, M/M from a Ground Support Cooling Unit (GSCU)circulates te and from the preflight heat exchanger through the umbilical.A Nedulating Flow Control Valve (MFC¥) controls the onboard fluid tempera-ture by varying the arount of coolant flowing through or around the heatexchanger. “Continuous valve operation maintains a stable M/N temperatureof 288°+ 0.86°K (59 41°F).

In flight, heat expulsion is achieved in the sublirator. Water, which issupplied under pressure from the water accumulator, freezes upon exposureto the space environment and then sublimes, removing heat from the M/hcoolant.

At T) +5 secends, a switch selector command enables the sensor biasThis command places RI, shown in Figure 18-8, in paraltel with the elec-tranic contreller asserbly input, causing the MFCY to be driven to zerobypass, and all coolant is forced to flow through the sublimator. AtTy 475 seconds, a switch selector command disables the electronic con-troller assembly. The M/W temperature is measured by two thermal switche:whose outputs are sensed by the LYDA/LVDC. The LYDC is programmed tosense the condition of these discretes once every 300 seconds, startingat 7] +180 seconds. This allows sufficient time to elapse after littottfor the ambient pressure to drop low enough for the water sublimator tofunction, If the temperature af the coolant is above the upper temperature setting, 2€8.3°K (59.55°F), a switch selector command is issued toopen the water solenoid valve to allow sublimator cooling. If the coolanttemperature falls below the lower limit 288.1°K (59.16°F), a switchselecter comand is issued which closes the water solenoid valve, haltingsublinater operation.

AS-502 was the first vehicle to utilize sublimator water-feed regulationfor tengerature control fror liftoff. The system was verified on AS-208as an after-rission experiment. Previous vehicles utilized 4/l nodula-tion, as described for ground operation. ‘The change was made to eliminatethe undersirable minium sublirator cooling rate. AS-502 also had seven140-watt heaters, flown previously on AS-501, as a temporary fix toipreve the ECS heat balance.

18.5.1 Thenral Conditioning System

higher prelaunch purge-gas temperatures were provided on AS-502 toelininate the low compartment ambient temperature experienced on AS-B01

18-16

Figure

18-8.

Schematic

ofThermal

Ccntral

Usin

gWater

Coolant

Valve

following S-IVB LHz loading. This higher purge-gas temperature, coupledwith higher skin temperatures, resulted in a conpartment temperature atTaunch of 295°K (71.6°F) for AS-502, as compared with 287.4°K (58°F) forAS-50]. The average skin temperature at launch on AS-502 was 5.12°K(11°F) above the AS-501 average.

The available data shows proper functioning of the TCS, as shown in Figure18-9. The present indications are that the M/W control temperatureremained within the 280.2 to 293.0°K (45 to 68°F) control band.

Figure 18-9. Temperature Control Parameters

Figure 18-10 shows a sublimator water inlet temperature rise at 111 seconds.This was 71 seconds prior to sublimator startup and was related to apressure differential (D43-601) increase at 116.6 seconds. This perform-ance deviation has heen attributed to water droplets trapped in the tubing‘on the sublimator side of the water solenoid valve. By 111 seconds, theIU compartment pressure and sublimator cavity pressure, as indicated bythe zero differential reading of 043-601, had decreased to 0.276 K/cné(0.4 psia). This was approaching the vapor pressure of water at tnetemperature within the cavity. The warmest water droplets in thesublimator and Tine cavity would be those adjacent to the water valvewhich dissipates heat during the time it is closed. As these dropletschanged phase, the warm vapor was sensed by the temperature measurementand finally by the pressure differential measurement as the gaseous

18-17

molecules filled the sublimator cavity. Water, at the pressure and tem-perature indicated, will experience a volume increase of approximately50,000 times as it changes into the gaseous state. The pressuredifferential increase shown by 043-601 in Figure 18-10 resulted from aconstant vapor pressure in the sublimator and tubing while the decreasingIU ambient pressure was approaching zero.

jo

es

russes,

toe?

Figure 18-10. Sublimator Water Inlet Per*ornance

A step increase in the differential pressure measurement is noticeableat 133 seconds, which consequently recresents 2 corresponding decreasein the IU internal ambient pressure. If no change had occurred in theIU ambient pressure, a temperature increase of 4.4 to 5°K (9 to 10°F)would have had to occur at this time if the vapor pressure within thewater Tine had increased by the magnitude shown in Figure 18-10. Thetemperature and pressure decrease after 149 seconds indicated that allthe water within the tube had vaporized. This pressure decay ischaracteristic of the sublimator normal crying cycle. At 18 seconds,when the water solenoid valve opened, the pressure differential responseled the tenperature resporse as the vepar bled out of the water line,

Figure {8-11 shows the sublimator start-up characteristics during ascent.Not shown is the position of the MFCV, which began driving to the fullsublimator flow at 1] +5 seconds as progranmed. Full heat sink flow wasachieved at Ty +16 seconds.

: 182.29 seconds (first interrogation of the thermal switches), thecoolant temperature was 290.2°K (63°F), which was 1.94°K (3.5°F) above

18-18

WA CONTROL TEN? <15-e0 | T .Fae ] 66%2 a s r a2 fe - in de 8209.7 =2 =z saynaon sow on ofana t Ss

2 SUOLIMERATGR WERT REDE"TOW ~ I |Sap eh - —+

go4 - - 4

_ | - 4

i \ S08 Hg THLET PRESSURE 043-607

PRESSURE,

PRES

SURE

.psta

wo[weteR Fubw sonLet l 12

kate

i L o25a we 500 500 78 Boa

ANGE TIME, SECHDS s

oa | | |

FLowaTE,

Yos/tr

Figure 18-11, I AS-502 Sublimator Startup Characteristics

the control point, and the water valve opened. This was verified by theyalve position, increasing water inlet pressure, and water flowratemeasurements. ‘The sublimator inlet temperature reached a calculated293°K (68°F) during ascent, and a 10 kw (34,140 Btu/hr) heat rejectionrate was required to bring the coolant temperature back te the controlTevel. Subsequent sublimator cycling can be identified from the curvesin Figure 18-9. Seven water coolant switch selector operations inSection 2, Table 2-3, verify proper operation of the trermal cortralswitching system.

None of the water system deviations {i.e., erratic low-level waterflowrates and pressures) experienced on AS-204 were present in theAS-502 data. Fluid pressures and flowrates were within the desiredranges, considering the 5 percent possible measurement error, as shownin Table 18-7. Punp outlet pressure, D17-601, was erratic. However,this was a transducer problem and the weasurement was waived prior toJaunch.

13-19

Table 18-7. IU TCS Flowrates and Coolant Pressure Data

, NEASURENENT FLIGHT VALUERrowmaTes UWEER ET Sica SPECIFICATIONmys (Goad mis (opal mys (gon)Coldplate § Fa-tor [0.000097 (0.87), b.00cos16 (0.60)

|

o,con02eT (0.685)Coldplate 20 4-603 Jo.oo0go3 (0.48) | 0.900032 (0.81)

|

o.c000261 (0.445)ST-120H-3 Shroud | F5-603 Jo.oonne3t —(*0.207))0.co0n139 (0.22)

|

o.cooovaa (0.222)Flight control

|

6-602 fa.noon309 (0.49)

|

o.0000378 (0.52)

|

o.coooes1 (0.445)Computercoléplate 4 Fr-eor fo.ogo (0.475) |o.c000a06 (6.50)

|

o.coooz81 (0.845)Lynas. Fe-s03 |o.oooosss (v0.77) [0.000873 (0.75)

|

o.cosa4sa (0.734)w re-soe —|o.onases(+9.4) Jo.ou0619 (9.8) {0.cov602 (9.55)so¥8 Fio-sot fo.onneys (7.85) 0.00052 (2.284) Fo, goa49 +6002

(rT 8.29)Srztaunea Fii-603 fo.coo0t2s —(+a.z00}/0.0000132 (0.21)

|

0.900018 (9.222)rowsTotal systen | serie fo.aoios §—(*17.25yf0.a0114 (18.05)

|

9.00109 + conesTo(17.3 +083)

Pressures New ~—~“Tpsta)_| Weg pata eaFuwp Inlet pea-eo [21.2 O62) [1a (16.5) 20.38

416.5 10.5)Pump QutTee or7-6or

|

28.2 (ay

|

aor (44.5)

|

17.2 to 28CErritic} (26 ra 34) Above Inlet

Seve Exit ois-sor [tas (26.53

|

20.7 (301

|

9.3 to 30.7(3.8 00 15.5)Below Pomp Outtet AeSpecified Flonrate

Watthin § percent possiBve measurenent error The TCS GNz sphere pressure decay is seen to be nominal and the excessiveGN2 consumption rate of AS-204 was not observed, as shown in Figure 18-12.

Figure 18-13 shows selected component temperatures for AS-502. Therewere greater fluctuations in component temperatures for AS-502 than AS-S01during the orbital phase of the flight, This was caused by the water-feed control method of temperature contra] used on AS-502 which alloned5°k (9°F) variations in the M/W temperature. The continuous contro)method employed on AS~&01 allowed variations of Tess than 1°K (1.8°F)in the M/W temperature. The variations in the component temperatureson AS-502 had no detrimental effect on the operation of any of thecomponents.

18-20

reco te |

fot

espeuse |

1

cae 120

fae:

renecR

aTURE

a0 TE,SECONDS 102TaeUa Baa awe TaD AS orga Oe

RAGE “OE, mRssMNUTESSENOS

Figure 18-12. TCS GHz Supply Pressure and Temperature

18.5.2 Gas Bearing Supply System

The Gas Bearing System (GBS) supplies regulated Gaseous Nitrogen (GN2)to the ST-12éM-3 Inertial Platform Assembly for lubrication of the gasbearings during preflight and flight operation as shown in Figure 18-7.During preflight operations, GN2 at 2068 N/cm@ (4000 psig) flows froma ground source through the GNo fill coupling on the umbilical andthrough the open GN2 solenoid valve to the 0.057 ms (2,0 ft8) sphere.Should the GNp storage sphere pressure drop below 637.3 N/am? (925 psig)during prelaunch subsystem operation, a signal from the low-pressureswitch would be initiated to shut off electrical power to the platform.The low-pressure switch is inactive during flight.

During system operation, GNo flows from the storage sphere at an initialpressure of 2068 N/em2 (3000 psig). After the GN2 is filtered, the gas

18-21

*

ee

ee7uRe

,

cloha. Cex Tene C920) |STADNIE 60 17H C55-6026 ASSN SDAP E73-6C3 |

|

Ge Earn site

resoen

eTute,

| t f foilOOo ee 6s We we ee ee Oe ee

awe Me, SECONDS » 10%

onsrSad OT 06a wer He basa eeaeaT OevouRs snnuTes-SE20n0s

BEAT OaeTaH VEOTees: 40

Figure 18-13. Component Temperatures

18-22

bearing pressure reculator drops the pressure to the acceptable value forges bearing operaticn, The GN2 flaws through the heat exchanger and asecond filter to the plotform gas bearing inlet. The heat exchangerthermally conditions the Giz. A line frer. the platform to the dome ofthe pressure regulator cuppiies the reference pressure required to controlthe pressure differential across the gas bearings.

The GBS pressure differential drifted above the 10.35 40.345 N/en®(15 40.8 psid) specification as expected. Figure 18-14 shows the differ-ential pressure to be €.069 N/om2 (0,1 psi) above the maximum specificationvalue after 11,670 seconds, The gas tearing pressure regulator issensitive to a chenging reference pressure, recuiring a series ofcalibration curves for each different reference pressure. The calibrationcurves generated during component acceptance testing (to modify the flightdeta) yield the following result:

The 10.35 N/cm2 (15.0 psid) at launch became 10.24 N/cm? (14,85 psid).

b. The 10.76 N/om? (15.6 psid} at 33,780 seconds became 1¢.85 N/a?(15.725 psid).

Moers 2

eo ¢

s ae SPECIETE °ori-602 se han 180

pL

aed i - 2s$ syecineo ne LeBU} Re anthe betcion :. _| #q cee a we

STS Seen1 a

1% 18 20 2 » ¥ «savce TIE, SECORGE «1038

weCrRANGE TINE, ORS MEMES: CONDE

Figure 18-14. GBS Pressure Regulation Performance

18-23

The fact that the 5 percent possible measurement error can account fora 0.69 N/omé (1.0 psid) correction further obscures the actual flightperformance data. The following facts should be noted:

a. The pressure differential was predicted to drift to 11.04 N/em2(16.0 psid).

b. The desired 10.35 N/on2 (15.0 psid) regulator set point is consistentlyin error because of the transducer sensitivity to the referencepressure.

¢, No conclusion about regulator performance can be reached until theflight data can be accurately assessed and transducer problems resolved.

The GN2 delivered to the platform was well within the 274.7 to 310.8°K{35 to 109°F) temperature requirement. The GBS GNp sphere pressure decayas shown in Figure 18-15 is seen to be near nominal. Calculations indicatea 0.388 SCFM usage rate (within the 0.3 to 0.5 SCFM allowable range)

venta useacesBhs

wu iss) TP

of}.orePaesstae eck

16ance TNE, sees « 102

St6)__ GSD” optosO) OLBGare GATE Gratiad SenANSE T29E, HOURS MERE: S049Figure 18-15. GBS Gllz Supply Pressure and Temperature

18-24

SECTION 19

DATA SYSTEMS

19,1 SUMMARY

The data system consists of the measurement systems, telemetry systems, taperecorders, RF systems and optical instrumentation. ‘There were 2753 telem-etered measurements active at the start of the AS-502 automatic countdownsequence. Of the 2758 measurements, 58 failed in flight, resulting in anoverall system reliability of 97.9 percent.

The Airborne Telemetry System operated satisfactorily, including preflightcalibrations and inflight calibration.

Tape recorder performance was good; however, due to the extended burn timeof the S-II and S-1¥B stages, the S-IC/S-I! separation data playback wasnot recovered from the S-II, S-IVB,and IU recorders. This was becauseinsufficient playback time was programmed to cover the anomalous situationcaused by the S-II two-engines-out condition.

Performance of the RF systems was good. Approximately 2 seconds of reatime data on all S-IC stage telemetry links were Tost due to a data drop-out at 146.0 seconds. This condition was also noted on AS-501 and appear:to be related to S-IC inboard engine cutoff (ECO)

Grourd camera coverage was good as evidenced by 84 percent system efficiency.Hovlever. only two of the six onboard film cameras were recovered. Threeof the cameras on the S-IC stage failed to eject and one of the S-II cameraswas not recovered due to a weak recovery beacon signal -

19,2 VEHICLE MEASUREMENTS EVALUATION

The measurement system is composed of the transducers, signal conditionersand power supplies necessary to transform the physical quantities beingmeasured into electrical signals suitable for telemetering, and of themeasuring distributors necessary to route these signals to the proner telem-etry channels.

The AS-5G2 measurement systems operated satisfactorfly. Lost measurementsdid not adversely affect vehicle postflight evaluation, except in isolatedcases such as J-2 engine vibrations, since sufficient data were acquiredto complete the evaluations.

19-1

There were 2788 measurements scheduled far AS-502 Twenty-six measurementswere watved/scrubbed prior to start of the automatic countdown sequence and4 measurements were not considered in devernining measurement reliability.Of the remaining 2758 measurements which were used for reliability determin:nations, 880 were on the S-[C stage, 951 on the S-IT stage, 597 on the S-IVB

Fifty-eight measurements failed in flight, resultingand 336 on the IU.in an overall measuring system reliability of 97.9 percent In addition tothe failed measurements there were 54 partially successful measurements,57 measurements with improper range and 2 questionable measurements. Four-teen of the waived/scrubbed measurements provided good data during flight.A summary of vehicle measurements is presented in Table 19-1.

19.2.1 S-IC Stage Measurement Analysis

There were 888 flight measurements scheduled for the S-IC stage. OF these8 measurements were waived/scrubbed prior to the automatic countdown se-quence, 15 failed in flight, 21 were partia:ly successful, and 19 hed improper range. Five of the waived/scrubbed measurements provided useful

Based upon 880 measurements active at the start afautomatic countdown and 15 failures during flight. the reliability wasdata during flight.

98.3 percent.

Neasurenents waived/scrubbed prior to launch, measurement failures (includ-ing partial failures), and measurements of improper range are summarizedin Tables 29-2, 9-3, and 19-4, respectively

Table 19-1. Yehicle Measurements SummaryESTE)

Suc sol STAGE INSTRUMENT

|

TOTALTASE Sings Domase ar) UNE VEHICLE

No. Scheduied 888] 958 604m) 60MM) 338, 275Gee*No. Waived/Scrubbed 8 7 9 9 2 26No. Failures 15 26 10 i c 58No. PartialSuccesses a 33" 0 a 0 54No. ImproperRange 19 24 8 8 57No. Questionable a a 2 1 2Measurement,Reliability 98.32 97.5% 92.3% 96.9% 1008 97.98

* Includes measurement faitures which were a consequence of enginemalfunctions.

** Fourteen of these measurements provided useful data during flight.

*** Four of these measurements are not wholly on the stage and are notused in calculating measurement reliability.

19-2

Table 19-2. Measurements Waived/Scrubbed Prior to Launch

meme seus.tovenr TLC comme eee| senses1

eat-na ascetersnvan, Lani zasinel estaiation not per aa-ver tus-1-t2, vais date atEas Separationconterss “emoersuoe, teriam Intet Manitate Trsdiese tatted at aasvar LIk-1-21, fstata,Tetaoneasters oration, Senta, Pitch Twanelecas failed, dni [28-1018 tyatta deta

coanr's ft congeot wit einen tutput 3 Failed 49 tnon ever valia daceyee fender aywore? tae stosn, pelea evel Fine, | seasuronont waged at saiver Uetz, valte atosha one niteaia teesRecrane fast sles uetta Laver cuore ensure: gagged at saver 81-325 ymdetash Sed nil'sate fives fg

woonerso rim cavers Tiner Ooerasion me code act var} aa ver LIR-(-26, portty waltViator Srecifrestie thes uetey aesuz2-12i Soros Licht aperation tmitcalinn sanow angle strate Mo'ver WIAGE-23, tata ant( Atpnc oe oporaeeLt. i - a

FHLwaecie 8 9 Side Lonitusianl Stain Spon strata slowene Defective “netalTatiorsorezia sT8 We cup Lengitactoat Seeain doen strain etamnt oetective sgtaltatiorsozane 578 96 otae Lonptasinal Stan deen stesta slomnt nofectse onetelatiorSa00-855 SAD Sede Lenatioat steain opm ctrate atamere vefeetive enstaltiorS028 614 S74 108 Top Lonltadteal Sean pen strata clement nefection intellattoeG7 ra tans cnsuzatioe surface Yenp oige aut at ester ocerst wate curtag rTigntcatecaveeims aneute cove torre Surface Tero Notas out oF secitic Peentver uate curteg fight[eter

” s-ive SAGE _ceai-aee tampassnzee tank Heettion @ rtsce'e Fon Produced dota suringtightcoss-roe toad “ant uae fis 20 tarcont nftex'e toa roduc dite from 69BHD sor=cds inphecxsoctar veapetagtag ti rasp Surface Renard at 119.2% to response to tomeratare chasesPREcassesor Tempt-g rrevate Bypass Line ensired ot 207.7% My repoine Ly Limparatare charges(Sascancion Teapeteriam aeoreas Sphere te. 4 Gos Dolannd tnyperattio ata nevoriftes to oe geetstecxnenns teas oritteer Tok siteeate You Shorted gel daring FREcare-sor Prossure-tmine Reytatur utter forTnate Tow vata reverifies to be gocease-401 Proscarsalen inte. Eventi date Mae 245 Ment 150 pags TustusSnr pelortteTeese TESt8sSeo tog second ore atleurtsear-s26 strc netyramic, rmwerd Skirt ronel 17| no sespoase oy eatt- tnubte to vecity RAC'S calibration

19-3

Table 19-2, Measurements Waived/Scrubbed Prior to Launch (Continued)

asa vensuaton rime erat cr esau sonsm

fona-cus

“T

Ftou Rate Colentate talet coolant

|

Fiavaeter vaitoated

|

trometer seared te wpncvie(loz 2) 1.33 62 Sah ery flew Erte to

|

promely at fiat? goa prentees1 tbe CP gon at ead date tor “essieder11 gt0017-851 seassure Conlang Yoattold In at Measurewiet siarat

|

taysccomet cecvides ayets* aeeaCoea ae BoE AT Fieercettret aoe ae L i

Tha crystal accelerometers and emitter followers used on AS-501 for 17engine area Vibration measurements were replaced with crystal acceler-ometers with integral emitter followers for this Flight. This change didnot completely eliminate hign amplitude, low frequency noise problemsnoted on AS-591. However, a significant improvement was noted in theamount of useful data obtained.

19.2.2 S-II Stage Measurement Analysis

There wore 958 flight measurements scheduled for the $-iI stage. Of these,7 measurements were waived/scrubbed prior to the automatic countdown se-quence, 24 failed in flight, 33 were partially successful, and 24 hadimproper range, Two of the Waived/scrubbed measurements provided usefuldata during flight, Based upon 951 measurements act‘ve at the start ofautomatic countdown and 24 failures during flight, the reliability was97,5 percent,

Measurements waived/scrubbed prior to launch, measurement failures (in-cluding partial failures), and measurements of improper range are sum-marized in Tables 19-2, 19-3,and 19-4, respectively.

Eleven temperature measurements which failed during flight indiceted apossible bonding problem. The bonding technique for these transducers isbeing investigated.

Erratic behavior of two temperature measurements on fuel turbine inletswas suspected to have been caused by melting of solder in the transcucerconnector, resulting in an open circuit to the dc amplifier. ECP J2-422effective on engines installed on A5~504 end subs, has been released, re-quiring that these wires be brazed to the connector.

Interaction between measurements on a common power supply which had beenAoted on AS~501 was not present on AS-502 even though events similar tothose which caused the interaction on AS-501 were experienced. This andstatic firing data from S-11-4 at Mississippi Test Facility (MTF) demonstratedthe effectiveness of the design change to eliminate interactions.

19-4

Several vibration measurements experienced short transtents. These tran-sients were sufficient to invalidate the data on two measurements. Exton-sive tests are under way with individual components and a complete stagesystem cenfiguratian to determine the cause, Transients similar to thosenoted on the flight data have been reproduced in a number of differentways; however, no conclusions as to the specific cause have been formulated.

It was also determined that al] engine vibration measurements saturatedtheir respective measurement channels. As a resuit, an engineering changeis being processed te change the range on all of these measurements.

The camera light battery monitor measurement decreased to 88 amperesat intersLage separation instead of the expected zero amperes. Theprobable cause was a cable short to the structure at interstage separation.Since cameras will nct be flown on future missions and this measurementwill not be used, no changes are conterplated,

The high percentage ef strain measurement failures experienced, with fivescrubbed measurements and one inflight failure, indicated a possible in-stallation protlem. tnvestigaticn indicated thet the present installationwas highly susceptible to moisture absorption, which could create the fail-ure mode experienced by these measurements. An engineering change ispresently being processed to moisture-prcof these instal laticns.

A number of measurements malfunctiored as a result of the failure modeof engine No. 2. Tests are being conducted in an attempt tc determirethe particular failure mode of the transducers and signal conditicniequipment. Thus far, it kes teer determined that similar failure indice-tions could be obtained with a diverse number of unexpected envircnmental

or failure condittons.

Fifteen temperature measurements fatled as a result of the engine No. 2failure mode at approximately 412.9 seconds. Of these, 12 hydraulic tem-perature measurements were believed lost due to at least ane wire from theconmon 428 vdc regulated power supply being shorted to its shield or tothe structure. Some of the wires from the power supply to the transducerswere located in the area of vehicle station 42.57 meters {1676 in.),stringer 108, where damage to wire bundles was considered tc have occurred.The other failures were considered to Fave been caused ty transducer opencircuits.

In addition to the 15 temperature measurement failures discussed above, fourPressure measurements did not exhibit normal shutdown characteristics atengine No. 2 shutdown. These measurements provided gcod data until engineNo, 2 shutdown; however, after engine cutoff, they did not indicate expectedvalues, Data indicated possible open transducers or oper circuits.An apparent failure occurred on the measurement monitoring switch selectortelemetry output. Investigation of the problem indicated that the wiringbetween the switch selector and the telemetry package, whick runs by ve~hicle station 42.57 meters (1676 in.), stringer 1U8, probably caused an

19-5

open circuit and was associated with the engine No. 2 failure. Switchselector operation was verified using another measurement.

AIT measurement failures asspciated with the engine No. 2 failure mode aregrouped separately in Table 19-3.

19.2.3 S-1VB Stage Measurement Analysis

There were 604 flight measurements scaeduled for the S-IVB stage. Of these,four measurements were not wholly on the S-IVB stage and nine measurement:were waived/scrubbed prior to the automatic countdown sequence.

During Phase I (liftoff to parking orsit insertion) there were ten neasure~ment failures in Flight. During Phase II (liftoff to 5-1YB/Spacecraftseparation) there were nine additional failures. One measurement whi clfailed in Phase I1 is considered questionable in Phase I. Eight measure-ments had improper range and two ware questionable. Five of the waived/scrubbed measurenents provided useful data during flight. Based upo591 measurements active at the start of automatic countdown and wholly onthe S-IVB stage and ten failures during Phase 1, the measurement reliability was 98.3 percent. Based upon 591 measurements active at she startof automatic sequence and 19 failures during Phase II (including the 10 PhaseI failures), the measurenent reliability for this period was 96.8 percent.

Measurements waived/scrubbed prior to launch, measurement Failures, measure-iments of improper range, and questionasle measurements are surmarized inTables 19-2, 19-3, 19-4, and 19-5, respectively.

Failure of measurements C10-493, 5162-403, and £209-401 is considered tobe associated with the engine failure to restart anomaly, and these measure-ments are grouped separately in Table 19-3

19.2.4 Instrument Unit Measurement Analysis

There were 338 flight measurements scheduled for the Instrument Unit. OFthese, two measurements were waived/scrusbed prior to the automatic count-down sequence. There were no failures in flight, either partial or totalSix measurements had improper range. Both of the waived/scrubbed measure-ments provided useful data during flight.

Based upon 336 measurements active at the start of launch and no failure:during flight, the reliability vas 100 nercent, Measurements waived/scrubbedprior to launch and measurements with imorover range are summarized inTables 19-2 and 19-4, respectively

19.3. AIRBORNE TELEMETRY SYSTEMS

The telemetry systems of each stage of tre vehicle operate independentlyof the other stages. These systems modulate the signals from the measure-ment system onto RF carriers for transmission to ground stations. The

19-6

Table 19-3. Measurement Malfunctions During Flight

Tet #0: Fase: Bod

fot crfeseorat engeo

19-7

viewer furore cic ARE eet aeLAREE__Pora

Tene atture cavie earn. |

ay es aL tit axe}: seceding eraneaver

Table 19-3. Measurement Malfunctions During Fligst (Continued)

veesueren exserno“ut em satu TRUSS sat ation eonsi veh [eet

408 cssartawnten o serts Tears Toone corercomes |rapvratie see ener iret He Lseate tome eter att extesraeena reiectis [attenhig se scores | sears foessoingpike uaent Uta [an “en sires Youre facie tance satma enserature-Fod Short BOrvezale, et Littate jane rcaley bign betwean

[fs iecnaet Pandinacrn: fmmepatirestentnUp fevtsateg ah Jens “ane eeanstng

Faeaery jee atae Tarte Pam, fon rene sexe new

|

soe[heat “ 1Te mTme, | Ta oR, ET- 1 ea _

ome 225 suet, faa,

|

rent, an Covet erase jaunt foneceonsatt —ramecrnemer wer bat onreeate, ne ret seers

|

2 vocefisieg case mse,

or24-apr Brass -UNz Sune Entersrone Coast pose yas ‘naestiqacion cintteving

sonny Jon aptentzr arse is lotta, ng era} tact [srt ttereaa

ccw.sarr [rene tage sees aiene

|

sree stent 9 sec-se

|

cry sesrds fcectatet ath engngcro |oempine sain tne Fase

|

agen cian 698 acoso secon acetates ontenemole |rsncemustion omer foe asgat tonutttteer| 494 scetn

|

ot snucsfecntice aert sotbts sedea ergioe

SRRTh MERARENENT FOTLIRES, SoePAE ‘

wor Taseatic. sen, rin

|

rateratacnt vit ze veces | ertnts one sable ont Testis 01d reraraninbyHage can

13-8

Table 19-3. Measurement Malfunctions Curing Flight (Continued)

LPR ei

rote) ero ettent tire th erne antty eet 8 scot, Feparent tansicersae Shae, raul Outer [aes nora ater 28 26 yecorte [29 seconde fonorent Eanedeceerene ere. stir, tnaenat, Hn fate nerate attar 62 ot socom |saconds pnneace transduce10.73 Tene, LOB Tak stage roti fom 1540 8 78 sounds ps scons

femcte emp, xox tone seta [ate ercnte after 8 26 atone 24 sacartsime Tap, Ue fae ste ia lege ates snotty eco 23 aconformate, raetcerhome Freee de cae set dats wesy fron 9 to 46 ta 88 ABE seen role anc

. * aSropt le at 42 seconds jtbone Fa tures

cane Hiravor, tet tap ston anpsttvze, tow | eo ea 38 secantASCs Magenshe elesos2-01 arianfan) Map ah epan, Yow ( seands Sone osabie dat

vets Sreat oc roes. ttcn fayeted 7 ttnes aner a |e sezenee lhactrant peenrearea § feed cit nati.eats ew ce ares. Seer 1, yeted @ mes atten em ea sacnts 9 secant faegnars sresureoza-r1é ——_‘seratn, totereank skirt, 2atu wend ot as emeetea| 12 seconds lise seconds [enperature comrnsation

wee i a -foo SORTA, PEAIREART FRILIAES, SH) SERGE

com.aea 14 Ful “wring tatet ane] Loos of 240 cacao second [reardater fare,Coors Hey Fuel “wotee caret Tem Lone oF dite 1710 sects 10 soronde eaneducer fate

cans FS Ler tomate Gitte Tate ene ot set secinse hi seconfraniucertt areceee2i9 i ressreaeuttorta. ton of 2 secanis fo secon rramduear tearcnp-sie reat Sr1e1d AF Suet mp “Lane of te 225 secon fies secnes firmtancer santoore heat shell AT Sort Tere wo eezends fat arinds [Trancdvcer rattepan tage we ate suet fe vacate on cece[Teasducer teller

[iowL

19-9

Table 19-3. Measurement Mal functions During Flight (Continued)

rat sustanee rime ove ar aun usally

“ ate Felure re saealy stand a08 peers |e cocoa

oe) 2 Tovervese Dut Plaid] 24%

soser

sunny sve | It arene

|

a8 arsende);cece agereere tat Flake 28

se

carsnty start

|

alt events

|

are sesame

19-10

banietoy

Lanta

mete

209 16

Lamia

Betustur,tr yitud'ral

vibration,inlet Flange,LongitudinalYibrotion, Foot PuTatet Fane,Feanitadiaa

Pump

Hibration, Fal PureFanaes sory: talinalibrabiva, Fuel eureFlaca, canqitatinalFibratinn, Ful sumeFlange, Vangituding!ibratinn, Laternal,Veliur Linewivration, Bending,PitehFibeation, Cesteuctsyster “ouetangPanel, LonzitudinalYinration, TelowsteyEeuinren: Paunt ingPanett Langitusiea

lard tom

lerescate, minh | secondlata Tow

lnrescate, nigh 2c +0 €9lent Tow seconds

lartscate, nigh }4 secondlara Towlyrrscaru, igh [o teraeu 60lard In secondsfirrseale, righ |e secondlant Towjprrsca'e, nich fo to 4 secondsaed 13potfscale, sigh [Us tenet

Tete level low | second

ets tevet lem |S secands

No valid data thequghout Flight:

ali data exceptfer tive noted.

a valid data: sheaunhout Flight:

alts deta exceat for tine rated

fio vatid sate errourneus Aint.

Halid Jets except for tive rotee

Walid feta exceat at 1ittofr

1a valid cata thenuthou T1Ght.

Wo valid vata after 5 seconds.

Table 19-4, Measurements ith Improper RangeFSSuREYENT time SATIRE HF Me |vormeet j Mesomrmst ere |e eC cen] TE we 79nks

[FREE Theeiitse, skh, aE VBTIA inte after 9 secondsFlush einted faveurTeta

coer-tog Hens best anietd— fotraecatey eign ictsd data except tor tine sates.Panacroy callcout-t1¢ tees, Suet Tank Stin forrecate, righ 109 s2conds |yci-d dota te 102 secondscrsteton [Tenp. seat shield foerseates elgh 50 to 9 HeTi data evtent for bine notee

Radiation Cal secsnds I

vowrt Jeress pire, noe foerscatey wigh 0.7 ue 2 fetid deta watee for time notecCortro? dalyc sptenls |

sateen, fiw, feefscara, mish Letsoet ane fetid cata oxsent for tine noteri ae hey lin tow iat secondseration, tin dy loetcate, migh |Littott, tox [esti data excent for tine notee,feaitieg tage “an Ton fod tescans

comi2 Jaibration, tin 2, jertscate, mint [Listorty tia featia uote exsent tar tire notesUsndtea Se ne tow wy and Ttseconds

rany.122 |sivratson, sin, fortseare, niger [Lisverts tox fratiu dato excugt tor tome notedTrailing aye ara ine iy ana ianecon

Srieration, tow tecate, sion fac seconds rater data except tor tle aoted,

19-11

Table 19-4. Measurements With Inproper Range (Continued)

PERSUREFENT ry TURE oF 4 orestamper

|

MEASUREMENT SuTLE

|

5pMMTURE OE[cag aresSon) STAGE

cc7e-219

|

FUE SeT Internal surf aFéscate, In Sarge 227.58 to 277 59°" (5 toTerp I Jato"F). Rane ¢rance $1123935-217

|

Lex Tan itage Texp ‘OFfeeate, Tom Bacge °99.82 to a2Z.uark {209t0 J00°F}. Deteted sree $1 3

ce37-217|

Lox tark Ul lage Tame OFescale, tow and subscesels [Lon Terk Ulage Tame aFeseale, Tow 1ceas-217 [Lox Tank uitage Temp OFfecate, Towcean-717 {Lox Tane ultage Temp [orfses1e, To447-217 Lex Tame Y1°age Town forfecate, ToeCaae-217

|

LOX Tame Ul'age “exp forfecate, Toe icaea.217 |ucx tanc Ul'age Tews forfscate, ve |casa-217 |0¥ Tank Ul:age Tew [orfscate, taecrm-29¢ [teat shiers as ortscate, tom | ange 1065.71 te ziee.ez"e {1600Recovery Tens eo J800F) Rana change 5- 21-3721-208 [Heat Shiewl fae 06 fSeate, Tow ange 1008.71 to 2199 H2"e (500Recovery Tea, 20 3500°F) Range change S-11-3ceap-21¢

|

tig tank Insutatiue [saturstes 13 <0 213 }ange 19,26 ta 30.93%(-425 toSubs Temp leetanen 130/200

|

seconds 190°) Rd aeLion oiennedsecondsceez-2oa_]e3 mat band 7 Sart |erfscate, tow ance 277.69 ta SHB.71°K (452 tnTena 00°F} No’ action atannedcotzor Jer Heat Exenanger [saturates at Farge Cte €R9.c8 Yen? [0 to{inter dress jprgtne start ea6 asia} buleted fram Seite?UCT8-202

|

2 Meat fxchanger

—

[saturates at fore subs| inlet Press lensing startdoré-208 [£3 ment Exchanyer

—

featurates at }Inet Press Jenene. startpove-rra [sa eat Cectamer

—

|saturates a }Inlet Fress longing start} :bote2c8 [5 Heat Cecmanger —|sasurates atInlet Press, longine stare ipia0-201 Jet PU value Intet —eatarates at Range ¢ to £89.48 weal ioPress jengire start ca6 esia} Fange cracae S-t1-4land $15.ovsn-roz [en eu vatve trlat [saturates at ;Press. engine startmreg.2¢3 £3 PU vaive Intet [saturates atPress eoine start |breo-2ct [ea PU valve Inet

—

|ssturates atPress lengine startbiac-208 ES ru verve tmlet

—

jsatyrotes atPress, angie start

19-12

Table 19-4, Measurements With Improper Range (Continued)

Exsuntnes? TMEASUREMEN : T_RATURE oF 1"Sunt MEAUNTHEN, FTL gesSeme creer THE REMARKS

S10 STAGE

Ecal-a0 ¥it-Thrust Strurt. at Lov TevelFe Dottle-thrust

eaez-a0 vit-thruse Steuce at Lov Yevelhie Dattte-Piten

2043-495 |vib Torus+ Struct at Low levelFe Bast Te-Yau

1057-493 |vib-Lry Trans vine jLow TevelTeS-Tutue

gvot-aga yin-are skirt Low tevet Pangt-‘heust,cice-es¢ yro-are setre Sw, Sel flow Tevel

Parel-ThrustE1e7-cah yib-Aee skiet Sw, Sel flow Tew

Panel -Rad.

£119-427 719-495 Hod, 1 sctach flaw tevelPt-Rad.

1 STAGE

cor7-6a3 |yiu ¢ Axts Inertial [150 percent [Litter'3iraay ST- 12aH

Eo14-533 Wh Joper tp Ring {150 percent —{R& secondsLana’

told |viv upper mtg Ping [150 percent 8 secandsPerp

026-613 vit cia, compstata flaw date eetAdapt Long

e027 603 ¥it Og. Ceewpytara how data calAdapt Per.

eozH-6U3 {Vit Dg. CompyDaza flaw data TevaI aeapt Tang.

Table 19-5. uesticnable Measurements

FEASUREENT FESSASUREKENT wexsunenent TITLE gufttiMlen REMARKS

SIE STAGE, PHASE T

209-401 | vinsComousttor paca prased and Fat ted during Phase 12Chanter Cone-Long. Jelipped

F211-201 | ¥ib-t92 Turbo Purp- fata biased ard Remained cunctianable during Phase 1}Lateral elippedL i * cons ideved vavid rensurerents by stage contrastor

19-13

different modulation techniques employed provide a means for efficienttransmission of a large quantity and variety of measured data requiringdifferent bandwidths and accuracy, There were 23 telemetry links used totransmit flight data on the AS-502 launch venicle: six on the $-IC stage,six on the S-IT stage, five on the $-1VB stage, and six on the InstrumentUnit. Performance of the telemetry systems was generally satisfactory

There were approximately 4.4 seconds of real time data lost on all S-1Ctelemetry links due to RF blackouts. Critical data were recovered, bow-ever, by airborne tape recorder playback. The $-[1 stage, $-IVh stage,and IU lost approximately 1 second of data on all telemocry links due toRF blackout during $-I¢/S-II staging. None o” these data were recoveredSince tape recorder dlayback did not reach this tine period due to a pro~longed recording >erfod brought about by S-II extended burn time and in-sufficient recorder playback time.

A summary of the telemetry system performance is presented in Table 19-619.3.1 S-IC Stage Telemetry System

There were six telemetry links used to transmit flight data on the S-ICstage: three Pulse Anplitude Modulation/Frequency Noduletion/FrecuencyModulation (PAM/FM/FM) links: one Pulse Code Nodutation/Frequency Modul a-tion (PCM/FM) link; and two Single Sideband/Frequency Modulation (55/FM)links, Transmission of data from all six links was generally satisfac-tory during flight with the exception of three significant cata dropcutperiods when data were lost from all links. These data dropouts cccurredat approximately 146.0 seconds (2 seconds), 149.2 secends (1.2 seconds),and 152.3 seconds (1,2 seconds), and corresponded with the RF blackoutperiods discussed in detaii in paragraph 19.5.1, Data from the AF] and AF2Tinks were recovered from the airborne tape recorder playback.

The interna: calibrator within the 270-chenne] multiplexer assemhly forPAM/FH/FH Tink AF2 initiated a sixth calibration step, approximately26 percent of full scale, after completion of each pregranmed five stepcalibration sequence. This resulted ir 83.3 millisecands of interrupteddata on each PAM channe] after each inflight calibration. No further hin-drance to data transmission was exhibited. This condition was noted prior toflight and a waiver granted.

During ground checkout, the frequency of the upper and lower band edges ofthe PCM/FH Tink RF assembly were found to be above their specified limitsby 2 kilohertz and § kilohertz, respectively, This out-of-tolerance con-dition was waived and produced no adverse effect on the PCM data transmitted.The signal strength from PAM/FN/FM link AFT became erratic and exhibiteda marked reduction in level at approximately 350 seconds. The incident andreflected power measurenents for this Tink Indicated an erratic mismatchin impedance between the Voltage Standing Wave Ratio (VSWR) monitor andthe RF multicoupler at this time. It is suspected that the problen isassociated with the cable interccnnecting these two components.

19-14

Table 19-6. Launch Vehicle Telemetry Links

Tine [PEI [oocaccton forae

|

EVMPt. vearomwance seas

av eat. [emvewsen s-16 [0-205 satsetactoryare ase.a HeRH/PH 9-098 natge Tie PTtanson (a2e)ser zane eewewer 0-395 fenae eiwr aa fooue ous veg Fawi

|

aas.0 |ssyre {a.205 123 3se 25.2 |sovea 0-385

or eas fomvenen |s-at [0-796 Satisfactorywe nee foamven [cee Joan vee te B Eon toewz zea. feawrenyat [sett |o-r80 a awr zane foo— [seit fon798 we Boarax 1-0ost gare fs fsst1 toeswee as. [soe b-795

cer) e5us eavensee |serva frurr duration satistacterycr euea fowrera [scr[rate auration Jorene same $383,tion spces

|

259.8 |rmyerzra [s-tve frum uration

|#™S

TEES Sareesco 232.9 [raven s-tv8 ]rutt duration Mae fovea 0cs aee.z [ssn setue [ruts avreeton

ary 2so.2 dewen tu frurt duration satistactoryare 2s. |rowseaven au [runt durasion pange Tine (oa)etbation (see)wr as7 [ssi fay fru aracion JTSser ass. [ocnyen fru rut doracton a2 ieprex 1.0vera err. |ocnyee Ley fran doractonvere zzez.s |ces tw runt duration

19-15

Programmed inflight calibrations indicated that all telemeter channelswere within the accuracies specified.

19.3.2 S-II Stage Tetemetry System

There were six telemetry links used to transmit flight data on the S-IIstage: three PAM/FH/FM links, two SS/FN links, and one PCM/FM Tink.Transmission of data from all six links was generally satisfactory withthe exception of expected dropouts during 5-IC/S-II staging when a1] datawere lost. These data were not recovered by airborne tape recorder play-back for the reasons discussed in paragraph 19.4.2,

Selected measurements and osciltograph recordings were evaluated to deter-mine the proper functional operatior of the telemetry equipment to theblack-box level. This selective review verified the proper operation ofthe telemetry equipment.

Four programmed irflight calibrations indicated that all telemeter channelswere within the accuracies specified.

Transients were noted on several channels of both SS/FM telemetry links.Comparfson of the data processed from all sites verified that the tran-sients were the result of a stage problem and not a data processing pro-blem. Studies are underway to determine the source of these transtents,

19.3.3 S-IVB Telemetry System

There were five telemetry inks used to transmit Flight data on the S-IVBStage: three PAM/FM/FM links. one PCM/FM Tink, and one SS/FM ink. Trans-mission of data from all five links was generally satisfactory with theexception of expected dropouts during S-iC/S-II staging when all datawere lost. These data were not recoverad ay airborne tape recorder play-back for the reasons discussed in paragrapa 19.4.

The performance of the PCM system was excellent. All multiplexers wereproperly synchronized and tneir outputs properly interlaced as confirmedby the reduced data. PCM was utilized as the prime data acquisitionsystem.

The performance of the PAM system was excellent. There were no systen orcomponent malfuncttons and synchronization was good.

Performance of the FM system was excellent. ‘The Voltage Controlled Os-cillators (VCO) performed well. VCO center and band edge frequencies werewell within their specified limits.

The $S/FM systems performed well during the mission, The SS/Rt transla~tor calibrated as specified, and all calibration signals were clean andeasily distinguishable for evaluation.

19-16

19.3.4 Instruvent Unit Telemetry Systems

There were six telemetry Tinks used to transmit flight data in the IU:one FX/FH/FM Tink, one PAM/FM link, one SS/FM link, one PCH/FM Tink (VHF),one PCM/FM Tink (UHF), and one POM/FM link which used the Cormand and Com-munications System (CCS) transponder. The three PCM links al] transmittedthe save data. Examination of available data indicated satisfactory per-formance of atl Tinks with the exception of expected drop out duringS-IC/S-11 staging when all data were lost. These data were not recoveredby airborne tape recorder playback for the reasons discussed in paragraph19.4. The reflected power in the FM/FM/FM link measured by the VSWR wasabove the specified 9 percent of the forward power; however, this measure-ment was above the specification prior to launch. Transmitted data wereof good quality and indicated low noise Tevels.

Performance of the PAM/FM and PCM/FM (VHF} Tinks was nominal. Al] datatransmitted were of good quality. Because of the abnormal flight profile,no satisfactory data were available to check out performance of theredundant UHF and CCS PCM systems,

The performance of the $S/FM link was satisfactory. The reflected powerfar this link was above the specified limits prior to launch but droppedto within specified limits after approximately 200 seconds of flight.

19.4 AIRBORNE TAPE RECORDERS

The airborne tape recorders recorded and stored for subsequent transmis-sion portions of the data that would otherwise have been lost due to flareeffects or visibility constraints at receiving stations. Not all of therecorded data were recovered due to factors discussed below. A summaryof vehicle tape recorders is presented in Table 19-7.

19.4.1 S-IC Stage Recorder

One two-channel magnetic tape recorder recorded data from the S~IC stageAF] and AF2 PAM/FM/FM telemetry links during S-IC/S-TI staging, The re-cord and playback commands were initiated on schedule as shown in Table19-7. Data were recorded for approximately 123.7 seconds. The durationof the airborne timer which initiated playback was 24.7 seconds, fallingwithin the design limits of 24 41.5 seconds. Airborne recorder playbackamplifier gain was within specified limits of 43 decibels of thecorresponding real time data,

The 3 sigma noise errors for recorded links AFI_and AF2Z were 2.73 percentand 2.08 percent of full scale, respectively. This compared with respec-tive 3.50 percent and 3.85 percent of futl scale noise values for theAS-501 flight.

19.4.2 S-II Stage RecordersTwo two-channel magnetic tape recorders recorded data from the BFl, BF2,and BF3 PAM/FM/FM telenetry links and selected discrete data pertinent to

19-17

Table 19-7. Tape Recorder Sumary

RECORN TIME PLAYBACK TIME

RECORDER peconpen (RANGE TIME, Sec) (RANGE TIME, SEC)START stop start _5T0P

LAUNCH PHASES10 Recorder AFI 50.15] 173.62] 173.82] 298.66

aF2S-I1_ Recorder BF 74.74] 169.78] 599,38] 699.60Nee BFZ 453.67| $87.59S-IT Recorder BFS 74.74] 159.78} 599.35] 699.60Ne, 2 BT 483.67] 587.695-1V6 Recorder cr 135.16] 193.78] 835.75] 908,75

tre 483145] 588.25crs

TU Recorder DF 135.95] 162.17] 836,45] 920.45pre 463.46] 595.75

ORBITAL PHASESuTVE Recorder wTPlayback at:

Tananarive 909.45} 2392.75 2397,95 2560.75

Guaymas 2560.95) 5at7.95 5418.35 5774.75

Tenanarive 8775.18] 7933.75 7939.97 8210.55Hawaii 8219.75] 10,442.75 to,a4z.95 flo,707.75Guaymas 10,707.35Hawati 16,929..72416 990.6271 967.42"

* Ground command.

19-18

the separation sequence. The discrete data were time division multiplexedby the BT1 multiplexer, and the playback wes transmitted on the BS1 5/FNtelemetry Tink.

The record and playback commands were initiated 2s shown in Table 19-7.The recorders and associated hardware performed 2s required, with the ex-ception of the record/playback capability due to the extended S-11 stageburn time.

Due to the extended S-II burn time the only receiver station to receiveplayback data was Bermuda: hence, this evaluation was limited to the flightdata received from that station.

Several problem areas were encountered because of te prolonged S-II powerecflight time. The normal record periods arc 85.04 secords for S~IC/S-IIseparation and 58 seconds for S-11/S-1VB separation folloved by a 100-sec~ond playback interval as determined hy S-11 stage onboard timer 3. Thistime sequencing was to ensure that the recorded calibrations occurringapproximately 18 seconds prior to S-IC engine cutoff and approximately18 seconds after S-II engine cutoff would be returned in addition to theseparation data. However, since the S-II stage burned longer than nominal,the record intervals for recorders 1 and 2 during S-I1/S-1¥B separationwere 103.93 and 99.48 seconds, respectively, rather than the nomina58 seconds. Tt should also be noted that recorders 1 and 2 stopped re-cording 17.2 seconds and 6.82 seconds after S-IT engine cutoff, respec.tively, instead of the nominal 23 seconds. The recorders were stopped atthese times because tape on the supply reel was exhausted. Additionally,Since the playback incerval was limited to approximate-y 100 seconds andthe recorders play back backwards neither the S-IC/S-17 separation datanor the calibration 18 seconds prior to $-IC engine cutoff were retrievedduring playback. No catibration data, therefore, were recrieved on taperecorder playback for comparison with the corresponding real time dat:

To ensure that these problens encountered on AS-502 would not recur on sub-sequent flights, Master Change Record (NCR) 5669 has been initiated todisable t'mer 3 by tieing back the output from the timer, thereby e]1imi-nating the 100-second limited playback interval. Although this would notrectify the failure to record the calibration cycle 18 seconds after S-Iengne cutoff, which is a function of the amount of tape in she supplyreel, it would ensure resrieval of S-IC/S-II separation data and the calfbration cycle 18 seconds prior to S-IC engine cutof?. Presently the taperecorders are equipped with sufficient tape to cover a burn time approxi-nately 40 seconds longer than nominal

The analysis of the tape recorder system was accomplished by evaluatingoscillograph recordings of continuous IRIG channels of BFl, BF2, and BF3plus certain PAM channels of Multiplexers Al, AZ, and A3. The presentrequirement is that the tape recorder playback shall be within 3 percentof che real time flight data.

19-19

The data on the continuous IRIG channels of BF1, BFZ, and BF3 varied from —the real time data only to the extent that additional noise was present onthe signal. Tre amount of noise present during playback above that pre.sent during real time data was 2 percent for ‘3F] and BF2, and 1 percentfor BF3. The PAM data atso displayed a nominal data level difference ofthe order of 2 percant for multiplexers Al, AZ, and A3. The comparisonof the calibration levels for the IRIG channels plus the PAM channels wanot performed since no calibration data were transmitted during paybackdue to the extended burn time,

The discrete measurements transmitted via the B71 multiplexer were analyzedutilizing a 3 percent trend tab and oscillograph recordings. These werthen compared with the PCM discrete tab for correlation between the twodata links. Analysis revealed that the 120 sample per second discretemeasurements correlated with the PCH data and the noise content did notexceed the 3 percent figure. The 12 sample per second discrete measure-ments transmitted via channels 24 through 27 of BT1 multiplexer, however,contained excessive noise not evident on the PCM data. Further analysisis required to resolve the source of the excessive naise. The cavibrationwas not transmitted during playback.

19.4.3 S-I¥B Stage Recorder

Three of the fourteen available tracks in the S-1VB tape, recorder wereutilized to record PAM/FM/FM Tinks CF], CF2, and CF3 during S-IC/S-II andS-I1/S-1YB separation. The recorder was commanded to play back the re.corded data after S-IYB first cutoff. Five of the 24 available tracks of theS-IVB tape recorder were utilized to record the PCN/FM telemetry link inorbit.

Tape recorder performance was generally good throughout the mission.S-IC/S-IT separation data were not recovered due to the extended S-11 burtime for reasons similar to those previously discussed on the S-II stagerecorders. The tape recorder recorded a1] analog data on fast and slowrecord commands and played it back for the t'mes specified with the ex-ception noted above. Although the orbit was perturbed as a result of theS-IYB overspeed at first burn cusofr, the ground track timeline over re-ceiver sites recording S-IVB tape recorder playback data apparently re-mained close to nominal since nearly 90 percent of the airborne recordeddata was received on the ground

19.4.4 Instrument Unit Recorder

One two-channel magnetic tape recorder recorded data from the IU DF] andDF2 telemetry links during S-1C/S-1] and $-11/S-IVB8 staging. The recorderresponded properly to record and play back commands. The quality of theplayback data received by Apollo Range Instrumented Aircraft (A/RIA) waspoor, probably because of poor signal reception. S-IC/S-II staging datawere not recovered due to the S-IT extended burn time for reasons similarto those previously discussed on the S-IT stage recorders.

19-20

19.5 RF SYSTEMS EVALUATION

The launch vehicle RF systems consist of telemetry, tracking (includingC-Band, ODOP, Azusa/Slotrac and CCS), conmand, and television systems trans-mission and reception. Not all of the data required to perform a totel RFanalysis were available for tais evaluation. 3ased on available data theoverall performance of launch venice RF systems was good. Measured flightdata with fev exceptions agreed favorably with expected trends. Telemetrypropagation was good, a5 was tracking performance. Due to the extendedS-I1 burn period tape recorder playback was not accomplished, and data lostduring S-IC/S-II staging were not recovered, as discussed in paragraph 19.4.Preliminary data indicated that the Command and Communications System per-formed well, Insufficient data were received to evaluate the video system.19.5.1 Telemetry Systems RF Propagation Evaluation

The telemetry transmission system provided the means for transmitting modu-lated measurement data from the vehicle to the ground receiving stations.The RF carriers were chosen to provide an appropriate ba‘ance of data trans-mission capability to handle the quantities and varieties of data originat-ing on the Saturn ¥ vehicle. The performance of the telemetry systens onAS~502 was excellent,

Gross main engine effects causing attenuation and signal strength flunctua-tions were observed between 120 and 145 seconds at Cape Telemetry 4 (TEL 4)and Central Instrumentation Facility (CIF) as predicted. Grand BahamaIsland {G81} also experienced these effects due to separation flow up theside of the vehicle. The average attenuation at TFL 4 was about 20 to 25decibels, which was less than the signal strength fluctuation experiencedby Saturn I and 1B vehicles and was comparable to AS-507

A transient in signal level occurred at 133.3 seconds but did not resultin any data loss. This transient was observed on the Instrument Unit OFI,OF2, DS1, And DP1 YHF telemetry Tinks at GBI and is stown in Figures 19-1and 18-2. No effects were observed at this time on the TEL 4 recordeddata. This transient is discussed in detail in Section 9A.

A drop of approximately 8 decibels was observed on IU Tirks DF, OF2, 0S1,and OPT at 141.2 seconds. This drop in signal level was indicative of apparent antenna ionization problems. Antenna recovery on these Tinks wasobserved at 165.5 seconds. A drop in sigral level fer 2 seconcs similarto that experienced on AS-501 occurred at 146.0 secancs and resulted iloss of signal from the S-IC VHF telemetry at TEL 4, CIF, and GBI. VHFtelemetry effects in the other stages were progressively less severe. Taperecorder playback data on the S-IC telemetry links at this time generallyindicated an increase in reflected cower and a decrease in forward powerThe time of this anomaly feltewed within 1.3 seconds of S-IC inboardengine cutoff and appeared to be directly related to this event.Staging effects at 149.1 seconds were as. expected, resulting in VHF tele-metry dota loss to al] supporting Taunch sites for approximately 1 second.

13-27

era

I

SGKUDAS"BAIL

JOR

S-II stage ignition effects on the VHF telemetry systems were observed atapproximately 162 seconds. 5-IC telemetry links experienced greater tan60 decibels attenuation at all sites and approximately 1 second data loss.S-I1 telemetry experienced attenuation up to 15 decibels, and the InstrumentUnit and the S-IVB stage had up to 5 decibels attenuation at TEL 4 and CIFNo effects were observed at GBI for the S-II, S-I¥B, and IJ telemetry Tinks.

Three drops in RF signal level occurred at 165.7, 168.1, and 169.5 secondson S-IC links AF1, AF2, and AF3. Duration of these drops were 2, 0.4, and1.6 seconds, respectively. Since these three links are transmitted throughthe same antenna system, it appears that this antenna system may have suf-fered an RF breakdown, This effect did not occur for tne other S-IC WHEtelemetry antenna system. The drops in signal level did not cause any lossin significant data since recovery to normal level occurred prior to taperecorder playback.

S-II second plane separation at 179 seconds resulted in approximately 25decibels signal degradation of the S-II, S-IVB, and IU telemetry signaltransmission to the Cape sites. Transmission to GRI was not affected,

Launch escape tower jettison did not result in any adverse effects to theRF telemetry transmission.

lonospheric effects were as observed on orevious Flights, posing no threatto reception of telemetry data. This phenomenon resulted in signal fluctu-ation to those ground stations loo<ing through the S-II exhaust plure andwas presumed to have been caused by interaction of the plume and ionosphericlayers.

Detailed analysis of S-IV¥3 and IU telemetry data during the orbital flightphase was at accomalished because of insufficient orbital flight data atthe time of tais report.

A summary of telemetry coverage from launch to approximately 28,800 seconds45 snown in Figure 19-3.

19.5.2 Tracking Systems RF Propagation Evaluation

The purzose of radio tracking was the determination of the vehicle's tra-jectory. The AS-502 vehicle carrfed severa? tracking transponders, as shownin Table 19-8,

Tracking performance throughout the flight was satisfactory, eccording todata received to date. No major anomaly occurred, although minor effectswere observed which are being evaluated to determine the potential impacton systems performance and possible impravenent for subsequent flights.

19.5.2,1 Offset Frequency Doppler (Q0UP;. The ODOP tracking system is aphase-coheren=, muTtistation doppler tracking system which measures posi-tion of the vehicle equipped with the ODOP transponder. The OCOF trans-ponder was carried in the S-IU stage of the vehicle, therefore CCOP tracking

19-24

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19-25

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19-26

Table 19-8, AS-502 Onboard Tracking Systems

vente ONBOARD ONBOARDTRANSMITTER RECEIVER

LOCATION SYSTEM FREQJENCY ERLQUE HY(MEGAHERTZ) (MESAUT RZ)

s-I6 pop 960 290Te Azusa/@lotrac 5000 5069. 199wu C-Band Radar 5765 Seae1 ccs 2282.5 2101.8

was limited to the flight of the firsl stage only. ODOP evaluation wasbased an limited data received from tie MARGO station and telecon with KSC.

The ONOP tracking system obtained doppler data for trajectory analysis fromvehicle liftoff untii S-IC/S-I] separation at 149 seconds. All groundstations maintained continuous track of the vehicle during this period.

The erratic data and large variation between predicted and actual ODOPtransponder received signal strength experienced on AS-501 from 35 to 45 sec-ands did nat recur on Lhis flight.

S-IC main engine flame attenuation on the ODGP transponder up-link signalstrength occurred from 80 seconds to S-IC/S-II separation. The flame cf-fects were sufficiently severe during the period from 110 to 145 secondste cause excessive phase modulation of the interrogator RF signal, thuscausing roisy data at the various ODOP receiving stations. This alsooccurrec curing the AS-501 flight and resulted in poor OLOP tracking analy-sis after 110 secorcs.

Phase Tock between the grounc interrogation station and the ODOP transponderwas lost at S-IC/S-II separatior due to staging flow field. Recovery wasmade shortly efter staging, at approximately 154 seconds, compared to192 seconds on AS-507.

19.5.2.2 fzusa/Glotrac. The Azusa/Gletrac is an interferometer trackingsystem usingdoppter and FM radar techniques to determine postion, range,range rate, and range sum data with a high cegree of accuracy. Becauseof the restricted number of ground stations, usage of the system is limitedto coverage in the Cape Kennedy area.

The performance af the Azusa/Glotrac system appeared to be satisfactoryand in acenrdance with nominal expectations. The Azusa Mark II statiortracked the vehicle from liftoff to 200 seconds. The signal was margiralfrom 200 to 237 seconds, at which time the onboard transponder broke phaselack. Bermuda accepted active interrogation of the transponder at 253 sec-onds and maintained track until 724 seconds. About 36 seconds of data were

19-27

lost during handover to Bermuda. All data from the Grand Turk Glotracstation were lost because of a local oscillator malfunction at the groundStation. Azusa/Glotrac tracking coverage fror launch to 11,720 secondsis shown in Figure 19-4.

19.5.2.3 C-Band Radar, C-Gand is 3 pulse radar system which used mono-pulse groundstationequipment and was used for precise tracking duringTaunch and orbit phases. Twa C-Band radar transponders were carried inthe IU to provide radar tracking capability Independent of the vehicleattitude. The transpender received coded or single-pulse interrogationfrom ground stations and transmitted a single-pulse reply in the samefrequency band, Each transponder radiated signals over a single transmit/receive antenna,

Insufficient data were received to compile a comprehensive analysis ofthe C-Band radar system throughout the mission in time for this report.However, available data indicated that performance of the C-Band radar waswell within the requirement of the mission with the following oxceptionsta, Data dropout occurred during second revolution at Bermuda and Redstonebecause of excessive azimuth rates and at Carnarvon because of site

transmitter nalfunction,

b. Undesirable transponder response occurred during third revolutionat Grand Turk Island and Redstone, with conplete data Toss atBermuda.

The C-Band radar tracking coverage from launch to 17,100 seconds ig shownin Figure 19-5,

19.5.3 Command Systems RF Evatuation

The AS~302 Command systems consisted of the Secure Range Safety CommandSystem (SRSCS) and the Command and Communications System (CCS).

19.5,3.1 Secure Range Safety Command Systen. The SRSCS provided a meansto terminate the flight of the-venicTe 5 radio command from the groundin case of emergency situations in accordance with range safety require-ments. Each powered stage of the vehicle was equipped with two comandreceivers/decoders and the necessary antennas to provide omnidirectionalreceiving characteristics.

Available data indicated that the SRSCS functioned properly during flight,and sufficient signal was received at all onboard receivers ta haveperformed the assigned function had destruct been essential.

19-28

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Figure 19-4. Azusa/Glotrac Coverage Sunmary

19-29

The S-IC camera impacted at approximately 7:10 a.m. EST and was re-covered at 7:25 a.m. by helicopter from the recovery ship. The capsulewas tracked by the SARAH beacon from impact time and recovery was made46.3 to 55.6 kilometers (25 to 30 n mi} from the ship. Inspectionshowed that two of the ground planes for the radio beacon antenna wereSheared off; three of the shroud lines in the drag skirt were broken;and the connector leaked, allowing water to enter the canera and wetthe film. However, the film was still usable.

Ouring recovery of the S-II cameras the beacon signal from one SARAHbeacon was not sufficient for radar lock on, resulting in loss of onecamera capsule. Drag flap and paraballoon, flashing light beacon, dyemarker, and shark repellant operations were all satisfactory on the re-covered capsule. The capsule landed within 13.45 kilometers (7.25 9 mi)Of the predicted impact point. The quartz window of the recovered cap-sule was shattered due to the impact attitude of the capsule, but therewas no apparent damage to film as a result of the broken window. The re-designed antenna deployment performed satisfactorily on this camera.Films from the TY cameras provided very good coverage.19.6.2 Ground Engineering cameras

In general, ground camera coverage was good. Seventy-five films fromthe Launch Comp?ex 39A camera system were received and evaluated. Sixitems had unusable timing, one was an unusable track, and one was a badexposure. In addition, five of the programmed cameras malfunctioned anddid not ecquire film data. As a result of the above failures, systemefficiency was 84 percent.

The Eastern Test Range tracking cameras provided good coverage with taeexception of an erratic track of the IGOR camera at Patrick Air ForceBase. In addition, these data were provided on 35 am film rather thanon the 70 mr film requested. The tumble rate of the S-IC after separationwould be difficutt to determine from 35 mm fitm.

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RANGE TINE, SECONDS.

‘ 4 ‘ 4. toh pole . 1 L 1 .1:13:20 316180 3220-00 9:23720 4:26:80 NWE:00;00"4:09:00 410604 4:10700 e130 A164 4720:00ANGE TIME, HOURS: MEWUTES:SECONDS

7 T = r .16,000 16,fo0 16,40 16,600 te,boo 17.500 17,000 17.400 17 800 7 So 18.0001 ve torravice me, Seconos. . 1 4 1. 44:26:40 4530500 543:20 736M) 4:40:00 Wr83120 41a5:40 4:50:00 S320 WEB TAD BODO 503.20

RANGE TIME, HOURS :MINUTES ;SECORUS

Figure 19-5. C-Band Radar Coverage Summary

19-30