as-202 launch vehicle operational flight trajectory
TRANSCRIPT
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NASA TECHNICAL MEMORANDUM
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N.U.4 TM X- 53470
June 3, 1966
A S - 2 0 2 L A U N C H V E H I C L E O P E R A T I O N A L F L I G H T T R A J E C T O R Y By Pamelia B. Pack
Ae ro -As t ro dynamic s Laboratory
NASA
George C. Marshall Space Flight Center, Hzlntsvzlle, Alabdma
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. TECHNICAL MEXO-UM X-53470
AS-202 LAUNCH VEHICLE OPERATIONAL FLIGHT TRAJECTORY . BY
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Pamelia B. Pack
George C. Marshall Space F l ight Center
Huntsville, Alabama
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This repor t presents the launch vehicle operational f l i g h t t ra jec- tory for the Saturn IB vehicle AS-202. Included is a discussion of the predicted t ra jec tory , the mission objectives, and cons t ra in ts , p lus the associated dispersion study, and tracking summary.
A successful f l i g h t w i l l place an Apollo spacecraf t i n to a lob-type t ra jec tory and w i l l a i d i n determining the performance of the launch vehicle and spacecraf t subsystems i n preparation of manned o r b i t a l m i s s ions.
NASA - GEORGE C. MARSHALL SPACE FLIGHT CENTER
NASA-GEORGE C. MARSHBIJ, SPACE FLIGHT CENTER
TECHNICAL MEMDRNBDuM X-53470
June 3, 1966
AS-202 LAUNCH VEHICLE OPERATIONAL FLIGHT TRkTEcTORY
BY
Pamelia B. Pack
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TRAJECTORY SECTION FLIGHT MECHANICS B R A "
FLIGHT TEST ANALYSIS DIVISION AERO-ASTRODYNAMICS LABORATORY
TABLE OF CONllENTS Page
SECTION I. INTRODUCTION
SECTION I1 . MISSION DESCRIPTION 2
A. Mission Objectives . . . . . . . . . . . . . . . . . . 2 B. Mission Constraints . . . . . . . . . . . . . . . . . . 2
SECTION 111. VEHICLE CHARACTERISTICS
A. Configuration . . . . . . . . . . . . . . . . . . . . . 3 Bo ~s.......................... 4 c. Propulsion . . . . . . . . . . . . . . . . . . . . . . . 4 D. Guidance 4 . . . . . . . . . . . . . . . . . . . . . . . .
SECTION IV. TRAJECTORY DESCRIPTION
. . . . . . . . . . . . . . . . . . . . A. Powered Portion 5, 6 B. Ballistic Portion . . . . . . . . . . . . . . . . . . . 7
SECTION V. DISPERSION ANALYSIS
A. Three-Sigma Dispersions . . . . . . . . . . . . . . . . 7 B. Other Dispersions. . . . . . . . . . . . . . . . . . . . 8
SECTION VI. TRACKING ANALYSIS 9
APPENDIX A Mass Characteristics . . . . . . . . . . . . . . . . . . 112-113
APPENDIX B IGM Constants and Pre-settings for AS-202 . . . . . . . 114-116
APPENDIX c Three-Sigma Tolerances . . . . . . . . . . . . . . . . . 117
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LIST OF ILLUSTRATIONS
Figure
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Launch Vehicle AS-202
Commanded Pi tch Att i tude Versus F l igh t T ime
Alti tude Versus Fl ight T ime
Fl ight P ro f i l e - Alti tude Versus Range
I n e r t i a l and Earth-Fixed Veloci t ies Versus Fl ight T i m e
I n e r t i a l and Earth-Fixed Path Angle Versus Fl ight T ime
Dynamic Pressure Versus F l igh t Time
AS-202 Ground Trace
Vehicle Fixed Acceleration Versus Fl ight T ime
S-IB Stage and Camera Capsule Impact Envelope
S-IVB Stage Impact Envelope
AS-202 C-Band Radar Elevation Histor ies
AS-202 Telemetry Elevation Histor ies
AS-202 GLOTRAC & ODOP Elevation His tor ies
AS-202 C-Band Radar S lan t Range Histor ies
AS-202 Telemetry S lan t Range His tor ies
AS-202 GLOTRAC & ODOP Slant Range His tor ies
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LIST OF TABLES
Table Title
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~ o m i ~ l Sequence of Events
Trajectory and Orbit Fact Sheet Significant Parameters at Selected Events
S-IB/S-IVB N d n a l Trajectory in Space-Fixed Parameters (Metric Units)
S-IB/S-IVB Naninal Trajectory in Space-Fixed Parameters (English Units)
S-IB/S-IVB Naninal Trajectory in Earth-Fixed Parameters (Metric Units)
S-IB/S-IVB Nominal Trajectory in Earth-Fixed Parameters (English Units)
S-IB/S-IVB Nominal Trajectory in Space-Fixed Plumbline Components (Metric Units)
S-IB/S- Iw N d n a l Trajectory in Space-Fixed Plumbline Components (English Units)
S-IB/S-IVB N a n i ~ l Trajectory in Earth-Fixed Plumbline Components (Metric Units)
S-IB! S-IVB N&M~ Trajectory in Earth-Fixed Plumbline Components (English Units)
S-IB/S-IVB Nominal Attitude History S-IB/S-IVB Aerodynamic Parameters (Metric Units) S-IB/S-IVB Aerodynamic Parameters (English Units)
S-IB Retro Space-Fixed Parameters (Metric Units)
S-IB Retro Space-Fixed Parameters (English Units)
S-IB Retro Earth-Fixed Parameters (Metric Units)
S-IB Retro Earth-Fixed Parameters (English Units)
S-IVB Coast-to-Impact Trajectory &n Space-Fixed Parameters (Metric Units)
S-IVB Coast-to-Impact Trajectory in Space-Fixed Parameters (English Units)
and Coast-to-Impact Trajectory in
and Coast-to-Impact Trajectory in
and Coast-to-Impact Trajectory in
and Coast-to-Impact Trajectory in
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12- 14 15
16- 19
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28-31
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36 - 39 40-43
44-47 48-5 1
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Tab le
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Title S-IVB Coast-to-Impact Trajectory in Earth-Fixed Parameters (Metric Units)
S-IVB Coast-to-Impact Trajectory in Earth-Fixed Parameters (English Units) Nominal Camera Trajectory
Dispersions Due to Three-Sigma S-IB Stage Propulsion Tolerances at Outboard Engine Cutoff
Dispersions Due to Three-Sigma S-IB Stage Propulsion Tolerances at Outboard Engine Cutoff
Dispersions Due to Three-Sigma S-IB Stage Non-Propulsion Tolerances at Outboard Engine Cutoff Dispersions Due to Three-Sigma S-IB Stage Non-Propulsion Tolerances at Outboard Engine Cutoff Three-Sigma Vehicle Performance Envelope at Outboard Engine Cutoff
Three-Sigma Vehicle Performance Envelope at Outboard Engine Cutoff
Dispersions Due to Three-Sigma S-IB Stage Propulsion Tolerances at Guidance Cutoff Signal
Dispersions Due to Three-Sigma S-IB Stage Propulsion Tolerances at Guidance Cutoff Signal
Dispersions Due to Three-Sigma S-IB Stage Non-Propulsion Tolerances at Guidance Cutoff S igna 1 Dispersions Due to Three-Sigma S-IB Stage Non-Propulsion Tolerances at Guidance Cutoff Signal
Dispersions Due to Three-Sigma S-IVB Stage Propulsion Tolerances at Guidance Cutoff Signal
Dispersions Due to Three-Sigma S-IVB Stage Propulsion Tolerances at Guidance Cutoff Signal
Dispersions Due to Three-Sigma S-IVB Stage Non-Propulsion Tolerances at Guidance Cutoff Signal Dispersions Due to Three-Sigma S-IVB Stage Non-Propulsion Tolerances at Guidance Cutoff Signal
Three-Sigma Vehicle Envelope at Guidance Cutoff Signal Three-Sigma Vehicle Envelope at Guidance Cutoff Signal Performance Partials Applicable at Outboard Engine Cutoff Performance Partials Applicable at Guidance Cutoff Signa 1
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TECHNICAL m U M X-53470
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AS-202 LAUNCH VEHICLE OPERATIONAL FLIGHT TRAJECTORY
Vehicle 88-202, having a suborbi ta l f l i g h t with planned recovery of its Comrrand Module (CM), is scheduled for launch i n the th i rd quarter of 1966,
The S-IB boost phase uses a pre-set t i m e tilt polynomial t o produce a t r a j ec to ry which closely simulates t ha t defined fo r the operational manned ea r th o r b i t a l Saturn IB. The booster propels the vehicle t o an a l t i t u d e of 56.5 kilometers and a range of 57 kilometers. After separa- t ion, a s ing le 5-2 engine provides th rus t t o reach a ve loc i ty of 6800 m/ see a t approximately 601.5 seconds a f t e r l i f t - o f f . Mode (IGM) is employed t o guide the S-IVB s tage i n the p i tch and yaw planes, while the Auxiliary Propulsion System (APS) provides r o l l control. S-IVB cutoff s igna l occurs a t an a l t i t u d e of 217.3 km and a range of 1595.4 Ian.
The I t e r a t i v e Guidance
Using the operational t ra jec tory as a base, the three-sigma vehicle dispers ion envelopes presented i n t h i s repor t were establ ished by using the root-sum-square (RSS) technique.
SECTION I, INTRODUCTION
The launch vehicle operational f l i g h t t ra jec tory for the Apollo/ Saturn 202 mission, the th i rd Saturn I B launch, is o f f i c i a l l y documented i n t h i s report . and the tracking data sumnary.
Also presented are the associated dispersion envelopes
AS-202, comprising an S-IB stage, S-IVB stage, instrument u n i t (IU), and payload, is scheduled f o r launch from complex 34 a t Cape Kennedy i n 1966, and w i l l maneuver from a 100-degree east-of-north launch azimuth t o a 105-degree f l i g h t azimuth. The Command Module (CM) is scheduled fo r a water landing within acceptable recovery area. planned a l t e r n a t e missions for AS-202.
There are no
Acknowledgements a r e made t o G. Wittenstein, J. Weiler, and B. Brothers for t h e i r assistance and cooperation; a l s o to Miss Jean Shelton fo r her contr ibut ion of the tracking portion.
SECTION 11. MISSION DESCRIPTION
A. MISSION OBJECTIVES
The mission of the Apollo AS-202 is t o demonstrate the com- p a t i b i l i t y between the Saturn I B launch vehic le and spacecraf t , and t o determine the performance of the launch vehicle and spacecraf t sub- systems i n preparation f o r manned o r b i t a l missions.
The planned mission p ro f i l e and support plans r e f l e c t the require- ments of the following primary objectives: (1) To demonstrate s t r u c t u r a l i n t e g r i t y of the launch vehicle and spacecraf t and confirm launch loads; (2) t o evaluate separat ion sequencing system; ( 3 ) t o ve r i fy operation of propulsion, guidance and control , and e l e c t r i c a l systems ; ( 4 ) t o f l i g h t test the Emergency Detection Sys tern (EDS) i n closed-loop configuration; ( 5 ) t o evaluate the thermal performance of the Coxmuand Module (CM) hea t sh ie ld ablator dur ing a high-heat load, long durat ion entry, and (6) t o demonstrate the mission support f a c i l i t i e s required f o r launch mission operations and CM recovery. For more de ta i led information on the AS-202 mission, see Reference 1.
B. MISSION CONSTRAINTS
Certain mission constraints a f f ec t ing the launch vehicle t r a j ec - tory shaping and establishment of the S-IVB terminal conditions have been considered: (1) The f i r s t s tage t r a j ec to ry w i l l c losely simulate the f i r s t s tage h i s to ry of the nominal Apollo-Saturn I B o r b i t a l mission. (2) The space vehicle w i l l have a launch azimuth of 100 degrees. (3) Guidance and s teer ing connnand r a t e s a re l imited t o one degree per second i n p i tch , yaw, and r o l l with a lapse of 10 seconds following l i f t - o f f before i n i t i a t i o n of maneuvers. ( 4 ) T i l t arrest w i l l be i n i t i a t e d before inboard engine cutoff and w i l l be maintained u n t i l guidance i n i t i a t i o n which occurs a f t e r Launch Escape System (LES) j e t t i son . ( 5 ) The S-IVB stage propellant loading w i l l be grea te r than 98,430 kgs (217,000 lbm). (6) S-IVB programed mixture r a t i o (PMR) s h i f t is t o provide a high s t ep (5.5) duration of 295 seconds. (7) A minimum of 10 seconds between S-IVB cutoff s igna l and S-IVB/CSM separat ion is required.
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SECTION 111. VEHICLE CHARACTERISTICS
A. CONFIGURATION
The launch vehicle , consis t ing of four major vehicle assemblies, the S-IB s tage , S-IVJ3 stage, I U , and payload, i s shown i n f igu re 1. A weight breakdown i s given i n appendix A.
The S-IB booster s tage of the mult is tage Saturn IB vehic le develops the th rus t necessary f o r the vehicle t o achieve the a l t i t u d e and ve loc i ty required by the upper s tages t o perform the mission and accomplish object ives . Eight H - 1 engines, uprated t o 890,000 Newtons (200,000 l b f ) , supplied by four each LOX and f u e l (RP-1) containers c lus te red around a center LOX container, are mounted i n a square pat tern. Four fixed inboard mounts cant t h e engines three degrees while t he four outboard engines a r e canted s i x degrees r a d i a l l y outward f r m the v e r t i c a l ax is of the vehic le and gimbal 58 degrees t o provide t h r u s t vector cont ro l f o r t r a j e c t o r y guidance- 7,100,000 Newtons (1,600,000 lbf ) .
The t o t a l s t age t h r u s t is nominally r a t ed a t
The S-IVB s tage is powered by a s i n g l e 5-2 engine i n s t a l l e d on of the s tage and having ra ted t h r u s t of 890,000 Newtons Pi tch and yaw control during powered f l i g h t is maintained
the cen te r l ine (200,000 lbf ) . by def lec t ion of t h i s engine. Roll cont ro l i s achieved by the Auxiliary Propulsion System (Aps) cons is t ing of two modules, each having three engines of 670 Newtons (150 l b f ) thrust .
The Instrument Unit (IU) houses the ST-124-M-I11 i n e r t i a l platform, the Launch Vehicle Digi ta l Computer (LVDC), and the e l e c t r i c a l equipment required f o r launch vehicle performance; and i s t h e second f l i g h t prototype of the Saturn IB/V Instrument Unit.
Apollo spacecraf t 011, as the AS-202 payload, cons is t s of the
Although Launch Escape System (LES), Conrnand Module (CM), Service Module (W), and a Spacecraft-Lunar Excursion Module (LEK) adapter (a). the mission is unmanned, the systems a r e designed t o be control led or monitored by astronauts . cabin containing spacecraf t controls and displays, observation windows, access hatches, and environmental controls .
The crew compartment is a three-man pressurized
Reference 2 contains more de t a i l ed information on the vehic le configuration.
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B . MASS
Shown i n Appendix A a r e the mass c h a r a c t e r i s t i c s used t o generate the AS-202 launch vehicle operat ional f l i g h t t r a j ec to ry . data were extracted from reference 3 .
These
C. PROPULSION
The propulsion cha rac t e r i s t i c s used t o generate the data presented t i n t h i s report were obtained from references 4 and 5 . Plots of t h rus t , flow r a t e s , and s p e c i f i c impulse f o r both f l i g h t s tages w i l l be published a t a l a t e r d a t e f o r f l i g h t evaluat ion purposes.
D. GUIDANCE
The i n e r t i a l Saturn I B guidance system performs navigation evaluations, issues d i sc re t e commands, i n i t i a t e s c e r t a i n guidance and cont ro l functions, and issues s teer ing commands t o guide the launch vehicle t o the pre-specified terminal conditions. These functions a r e accomplished by means of the AS-202 Fl ight Program which i s s tored i n the Launch Vehicle Dig i ta l Computer (LVDC) (see reference 6 ) . Inputs t o the LVDC f o r navigation and guidance evaluations a r e the pre-specified constants for the S-IB s tage and S-IVB guidance modes (see reference 7), and the i n e r t i a l acce le ra t ions , i . e . , ve loc i ty increments, sensed by the ST-124-M-I11 s t ab i l i zed platform accelerometers during powered f l i g h t .
The navigation subroutine which i s i n the LVDC Fl ight Program combines g rav i t a t iona l acce lera t ion evaluations with the i n e r t i a l p l a t - form data t o compute the pos i t ion and ve loc i ty vectors with respect t o an earth-centered, i n e r t i a l coordinate system. The s t a t e var iables as determined by the LVDC navigation subroutine a r e then used by the LVDC guidance subroutine t o compute a t t i t u d e commands.
The LVDC guidance system subroutine ca lcu la tes the vehic le a t t i t u d e commands i n the Euler Angle System ( X x , X , X z ) . The commanded a t t i t u d e angles a r e compared with the ST-134-M-I11 gimbal angles ( 8 x, 8 y, 8 z ) which represent the vehic le ' s a t t i t u d e . The a t t i t u d e e r ror angles a r e then u t i l i z e d by the cont ro l computer t o ca l cu la t e motor swivel commands.
S-IB s tage guidance cons is t s of a three-segment t i m e function polynomial i n the p i tch plane of the form:
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The above polynomial i s evaluated approximately every .9 seconds, which is the major loop computation cycle of the LVDC during S-IB operation u n t i l active guidance.
S-IVB s t age guidance is the I t e r a t i v e Guidance Mode (I@€) For more information on the IGM formulation fo r two f l i g h t stages.
f o r AS-202, see Reference 8 . f o r the AS-202 mission a r e found i n Appendix B.
The target ing constants and pre-set t ings
SECTION I V . TRAJECTORY DESCRIPTION
A. PCXJERED PORTION
Shown i n table 1 is the nominal sequence of events for AS-202, and i n tab le 2, a t r a j ec to ry and o rb i t f a c t sheet . Table 3 gives some s i g n i f i c a n t parameters f o r certain events. The above tab les serve as a sumnary of the following descr ipt ion.
Guidance Reference Release (GRR) s i g n a l occurs approximately five seconds before launch canmit, a t which time the LVDC begins its computations and the platform becomes i n e r t i a l . Vehicle first motion occurs approximately 0.2 seconds before Instrument Unit (IU) umbilical disconnect. ( T B ~ ) is i n i t i a t e d . The launch vehicle w i l l r ise v e r t i c a l l y fo r 10.2 seconds i n order t o clear t h e launch f a c i l i t i e s , and w i l l then simultaneously begin i t s p i t c h program and r o l l maneuver.
A t the moment of I U umbilical disconnect, t i m e base one
S-IB s tage shaping was ionstrained t o be c h a r a c t e r i s t i c of t he operat ional Saturn I B o r b i t a l mission. I n addi t ion t o t h i s , the p i t ch program w a s shaped t o minimize aerodynamic moments during the period of highest pressure. The tilt program was not biased f o r wind as the expected launch date occurs during a low-wind period. A tilt arrest of 60 degrees is prograxned a t 136 seconds a f t e r l i f t - o f f t o i k u r e ample damping t i m e f o r various sequences. along with the nominal S-IVB stage cammanded a t t i t u d e h is tory .
The f i r s t s tage tilt program is shown i n f igure 2,
AS-202 i s the second I B f l i g h t u t i l i z i n g the three-second t i m e d i f f e r e n t i a l between inboard engine cutoff (IECO) and outboard engine cu tof f (OECO) se lec ted a f t e r coordination with in t e re s t ed labora tor ies (reference 9) as the optimum solu t ion t o performance penal t ies imposed by the six-second carry-over from the Saturn I vehicles. IECO and OECO a r e predicted t o occur a t 140.87 and 143.87 seconds a f t e r f i r s t motion.
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Approximately 3 . 8 seconds are required f o r 5-2 engine th rus t buildup. This portion of the t r a j e c t o r y a l s o includes ul lage rocket burn f o r propellant sea t ing a f t e r separation.
S-IVB mainstage operation begins when the 5-2 t h r u s t l eve l reaches 90 percent of i t s nominal value. seconds a f t e r l i f t - o f f , i s continued u n t i l 172.75 seconds, allowing s u f f i c i e n t t i m e f o r the LES tower and ul lage rocket casings t o be je t t i soned . The Saturn guidance system, the I t e r a t i v e Guidance Mode (IGM), i s implemented a t t h i s t i m e t o guide the vehicle t o desired terminal conditions.
T i l t arrest ,which began 136
This i s the second f l i g h t u t i l i z i n g a Programed Mixture Ratio (PMR) s h i f t t o increase performance capab i l i t y by s h i f t i n g power levels. The Propellant U t i l i za t ion (PU) system provides i n - f l i g h t propellant mass h i s t o r y and s igna l s f o r t he f u e l tank pressure s t e p and deplet ion log ic ,
A t PU ac t iva t ion (6.8 seconds a f t e r 5-2 i gn i t i on ) , the system senses the LOX overload and commands the engine t o burn a t the high engine mixture r a t i o of 5 . 5 : l . Fl igh t of the S-IVB s t age w i l l continue a t the high mixture r a t i o u n t i l the measured LOX e r r o r i s removed. me mixture r a t i o s h i f t f o r t h i s t r a j e c t o r y i s expected t o occur a t 4 5 0 seconds of f l i g h t t i m e . The PU system w i l l then command the engine t o burn a t the reference mixture r a t i o (RMR) of 4 . 7 , s t r i v i n g f o r simultaneous deplet ion of LOX and LH2 fo r maximum s t age performance.
The I G M w i l l s teer the vehicle i n i t s normal mode u n t i l approximately 46 seconds before expected S-IVB cutof f . A t t h i s t i m e , t he 'iT mode i s entered,and IGM the rea f t e r enforces only the ve loc i ty vector end conditions. This serves t o desensi t ize the guidance system t o anv la rge osc i l l a t ions i n t h r u s t and flow r a t e s . i n t e rac t ion , a d i g i t a l f i l t e r f o r F/m has been implemented f o r AS-202, replacing the-digital f i l t e r fo r the AS-201 (see reference 10). Until the t i m e of X mode, navigation and guidance calculat ions were being made on the average of every 1 .7 seconds. Due t o reduction i n the number of computations performed by the LVDC, a t t h i s t i m e , the computation cycle i s reduced to every 1.4 seconds u n t i l cu tof f .
To avoid any s losh guidance
Guidance Cutoff Signal (GCS) is sen t by the LVDC when the i n e r t i a l veloci ty reaches 6800 m/sec. a ve loc i ty and t i m e function polynomial f o r the desired cutoff ve loc i ty .
Actual cutoff t i m e i s solved by
The conditions a t Guidance Cutoff Signal are t i m e - 601.55 sec; space-fixed ve loc i ty - 6799.63 m/sec; space-fixed path angle - 86.004 deg.
Shown i n f igures 3 through 9 is a graphical display of the AS-202 t r a j ec to ry .
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B. BAIJ;ISTIC PORTION
Retro-rockets are igni ted a t separa t ion of the S-IB/S-IVB t o dece lera te t h e S-IB stage. subsequent b a l l i s t i c f l i g h t t o impact, is presented i n t ab le s 15 through 18. S-IVB ul lage burn, and 5-2 engine i g n i t i o n s t age a f t e r separation. presented in f igure 10. i n t a b l e 23.
This port ion of the t r a j ec to ry , with
-0 motion p ic ture cameras, recording S-IB/S-IVB separat ion, are e jec ted from the S-IB
The S-IB and camera capsule impact foo tp r in t is The nominal camera capsule t r a j ec to ry is. presented
Following S-IVB Guidance Cutoff Signal , there is a 10-second coas t , a f t e r which the CSM is separated from t h e S-IVB stage. coas t t r a j e c t o r y i s presented i n tab les 19 through 22 with impact foot- p r i n t i n f igu re 11.
The S-IVB
SECTION V. DISPERSION ANALYSIS
A. THREE-SIGMA DISPERSION ANALYSIS
Since a deviat ion from the mean i s s t a t i s t i c a l l y probable, i t is necessary t o determine,within reasonable limits, a vehic le envelope which w i l l descr ibe the f l i g h t of AS-202. developed by introducing perturbations of a 3 0 magnitude i n t o the nominal p ro f i l e . Propulsion and nonpropulsion per turbat ions w e r e considered f o r each s tage of t he launch vehicle. a r e considered t o be independent and a r e assumed t o exhib i t a normal (Gaussian) d i s t r i b u t i o n and have equal p robab i l i t i e s of occurrence.
Off-nominal t r a j e c t o r i e s w e r e
Individual per turbat ions
In determining the dispersed f l i g h t cor r idor , t he e f f e c t s of each of the off-nominal parameters on s i g n i f i c a n t t r a j e c t o r y var iab les w e r e determined. The root-sum-square (RSS) technique w a s then used t o determine probable plus and minus var ia t ions as a function of t i m e f o r each of the var iab les considered. per turbat ion e f f e c t s by squaring each r e su l t i ng parameteric dispers ion and ex t rac t ing the square root . Thus,
This technique accumulates t h e various
A parameter = (Perturbed Parameter1 -1 Nominal Parameter I
-Rss = t w . The e r r o r sources considered a re presented inAppendix C. The .
7
deta i led r e su l t s of t h i s analysis a r e presented i n tab les 24 through 3 9 . S-IB dispersions , both propulsive and nonpropulsive, a r e shown i n tab les 24 through 27 a t outboard engine cu tof f with the t o t a l envelope presented i n tab les 28 and 29 . Tables 30 through 33 give the nominal parameters with dispers ion de l tas a t Guidance Cutoff Signal (GCS) due t o S-IB stage e r ro r s . S-IVB per turbat ions a r e given i n tab les 34 through 3 7 . A t o t a l vehicle envelope a t GCS is shown i n tab les 38 and 3 9 . Dispersions due to platform tolerances a r e not presented here, bu t w i l l be presented a t a l a t e r date . d r i f t s were used i n generating the nominal t r a j ec to ry .
Actual measured platform alignments and gyro
The predicted three-sigma envelope shown i n tab les 38 and 39 i s applicable a t S-IVB/CSM separat ion except f o r ve loc i ty due to t h r u s t decay uncer ta in t ies . i s k.65 m/sec.
The revised ve loc i ty envelope a t S-IVB/CSM separa t ion
Signif icant r e s u l t s a t Guidance Cutoff Signal a r e as follows:
Time = +19.8 sec , -18 .5 sec
Radius Vector Magnitude = +263 meters, -103 meters
Space-Fixed Path Angle = - . 0 0 2 deg.
F l igh t Performance Reserve (FPR) = 1650 lbm.
An FPR of 1650 lbm i s necessary t o achieve the mission under three-sigma considerat ions, and is within the performance capab i l i t y of the AS-202 launch vehicle.
A spec ia l note needs t o be made on the r e s u l t s of the S O X and +Fuel (LH2) cases for the S-IVB s tage . These cases cause s h i f t s i n the t i m e that the PU value w i l l come off the s top. As w a s mentioned e a r l i e r , the primary reason f o r a Planned Mixture Ratio s t e p i s f o r performance purposes; however, f o r AS-202, the nominal s t e p time i s non-optimum; hence, a gain i n res idua l appears i n the r e s u l t s of the - fue l and +LOX cases. The associated loss i n nominal useable res idua l i s approximately 6 8 kg ( 1 5 0 lbm).
For a quick-look ana lys i s , a group of se lec ted p a r t i a l s are given i n t ab le s 40 and 4 1 , which a r e appl icable a t outboard engine cutoff and GCS, respectively.
B . OTHER D I S PERS IONS
Control engine-out capab i l i t y is almost nonexistent fo r a minus three-sigma performing vehicle; however, analysis reveals capab i l i t y
8
i n t h i s case from approximately 130 seconds a f t e r first motion. is l e s s than a 50 percent probabi l i ty of a successful mission i f a cont ro l engine f a i l s earlier than 90 seconds of f l i g h t t i m e . r e s u l t s are based s o l e l y on performance capabi l i ty .
There
The above
SECTION VI. TRACKING ANATYSIS
This sec t ion presents the expected coverage of the AS-202 mission from primary t racking and telemetry sites. provided by C-band radars , ODOP uprange, GLOTRAC, and close- in o p t i c a l systems .
Tracking coverage will be
A m p defining the vehicle ground t r a c e and the coverage redundancy above an e leva t ion of f i v e degrees i s shown i n f igu re 8. be ava i lab le from a t least three s t a t i o n s through S-IVB cutof f .
Coverage will
A de ta i led assessment of the v i s i b i l i t y of the vehic le t o each ground s t a t i o n scheduled t o receive telemetry and t rack t h e onboard e l ec t ron ic systems is given i n f igures 12 through 17. range h i s t o r i e s are shown from l i f t - o f f through S-IVB impact. Antigua, the last ground si te t o receive data from the vehicle before impact, will t rack the S-IVB t o a f l i g h t t i m e of approximately 910 seconds.
Elevation angles and s l a n t
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11
TABLE 2
AS-202 TRAJECTORY AND ORBIT FACT SHEET
A. LAUNCH
Launch Complex and Pad
L a t i t u d e
Longitude (west)
Launch Azimuth
Fl ight Azimuth
B . TRAJECTORY HISTORY*
F i r s t Stage - S-IB Stage Pi tch T i l t I n i t i a t i o n
S-IB Stage Roll T i l t Termination
S-IB Stage Roll Angle
S-IB Stage Mach One
S-IB Stage Maximum Dynamic Pressure
S-IB Stage Pi tch T i l t Arrest
S-IB Low Level Sensed
S-IB Stage P i tch Angle a t T i l t Arrest
S-IB Stage Inboard Engine Cutoff (IECO)
S-IB Stage Outboard Engine Cutoff (OECO)
S-IB Stage Velocity a t OECO (Space Fixed) (Earth Fixed)
S-IB Stage Path Angle a t OECO (Space Fixed) (Earth Fixed)
34
28.521958 (deg)
80.561142 (deg)
100 (deg)
105 (deg)
10.2 (sec)
15.2 (sec)
5.0 (deg)
65 (sec)
79 (sec)
136 (sec)
137.87 (sec)
60.00 (deg)
140.87 (sec)
143.87 (sec)
2209.53 (m/sec) 1855.41 (m/sec)
64.46 (deg) 59.11 (deg)
.
1 2
TABLE 2 (CONT)
I '
I 1
s-IB Stage Al t i tude a t OECO
s-IB Stage R a n g e a t OECO
S - ~ / S - ~ W Separation
Ullage I g n i t i o n (S-IVB Stage)
s-IWS-IVB Separation s i g n a l
Retro-Rocket I g n i t i o n (S-IB Stage)
Second Stage
S-IVB Stage (Main) Ign i t ion Cumnand
S-IVB Stage Ullage Rocket Cutoff
S-IVB Stage 90% Thrust Attained
S-IVB Stage Launch Escape System Je t t i soned
S-IVB Stage Guidance I n i t i a t i o n
S-IVB Stage Guidance Cutoff Signal (GCS)
S-IVB Stage Velocity a t GCS
S-IVB Stage Path Angle a t GCS
S-IVB Stage Al t i tude a t GCS
S-IVB Stage Range A t G€!S
S-IVB Stage Lat i tude a t GCS (Geodetic)
S-IVB Stage Longitude a t GCS
C. INSERTION CONDITIONS (S-IVB/CSM SEPARATION)
Time
Space-Fixed Velocity
Path Angle of Velocity Vector (Against Local Vertical)
55.63 (lan)
55-68 (lan)
144.55 (sec)
144.75 (sec)
144.75 (sec)
146.15 (sec)
140.42 (sec)
149.95 (sec)
170.00 (set)
172.75 (sec)
601.55 (sec)
6799.63 ( m / s e c )
86,0036 (deg)
217.31 (km)
1595.40 (km)
23.6407 (deg)
65.5384
611.55 (sec)
6798.73 ( m / s e c )
86.1678 (deg)
13
.
TABLE 2 (CONT)
Altitude (Oblate Earth)
Ground Range
Geodetic Lat i tude (north)
Longitude (west)
Azimuth of Velocity Vector (Posi t ive East from North)
D. IMPACT DATA
S-IB Stage
Time
Lat i tude (Geodetic)
Longitude
S-IVB Stage
Time
Lat i tude (Geodetic)
Longitude
221.91 (km)
1657.06 (km)
23.4258 (deg)
64.985 0 (deg )
112.2241 (deg)
507.78 (sec)
27.4103 (deg)
76.3436 (deg)
1521.07 (sec)
4.7609 (deg)
31.4236 (deg)
* A l l values a r e space fixed where appl icable unless otherwise noted. T i m e s are referenced t o F i r s t Motion.
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1 I I I I I I
I I I I I
I I I I I I I I I I I I I
91
w l 3
I I
I I I I I I I
I I
I I I I I I I I I I
I I I I I I I
I I I I I I I I I I I I I I I I I I I I I I I I I
I I I I L I I I I I
I I
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I I I I I I I I I I I I I I I I I I I I I I I I
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I
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I I I I I I I I I I I 1 I I I I
I I
I I I I
I I I I I I I
\o
? I
m Q: -2
0
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0:
Q: m m N rl I
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m rl N I
P-
0 N N N P- I
rn 2 W
I I I I I I I
I
I I
I I I
I I
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I I I I I I I I I
.
92
8
.
L
B f l
rn
& 8 &
1
p r l m
6 % m m l - a 0 4 o m rl
= 6 6
ti f: 0 0 rl
k aJ PI
6 f: 0 0 d
k Q) PI
E z 3 1 PI
“1 N
m b
I
Q‘. N rl
I
b
m d
/3 m m 00
cr) o\ 0 rl
I
N r l b r l N \ D hl
0
r(
I
-
0
“1 m m
I -
0
m h
0
Y “ 1 N I
0 0
0 0
m b m m O N
I
5 H
3 I cn
W
!% 5
E E - 1 - 1 1
I O c n & -- PI
2 6 H Z
0
-
;2 N
0 0
\D m rl
I
m d N
n
H I cn
9 W
2 0 1-11
2 -
t h
!3
2
8
H I cn W
0 1-11
1-11 94
LAUNCH ESCAPE MOTOR
LAUNCH ESCAPE TOWER
COMMAND MODULE
- INSTRUMEWT UNIT- - #- GUIDAWCE SYSTEMS
TELEMETRV EQUIPMENT POWER SUPPLY BlCnERlES
AND INVERTERS
S-IPB SECOND STAGE 11-2 EWGIWE 200,OM LBS.TOTAL
THRUST LOX/LH2
S-IB FIRST STAGE &/H-1 EWGIWES 1,600,000 LBS.TOTM
THRUST . U X h P - 1
LAUNCH WEI6UT 1,312,000 LBS.
I C' I 12' SQACECRAFT
2:
SPACE VEHICLE
FIG. 1. APOLLO / SATURN I B SPACE VEHICLE
'95
Pitch Attitude Command Angle (deg
80
70
60
50
4 0
30
20
10
0 0 100 200 300 400 500 600
Fl ight Time ( s e d
FIG. 2. AS-202
NOMINAL PITCH ATTITUDE COMMAND
VERSUS FLIGHT TIME
Alt i tude ( k m )
300
200
100
0 0
' S-IB Ballistic Flight
\ \ \ \
500 1,000 1,500 Fl ight Time (sec)
FIG. 3. AS-202 LAUNCH VEHICLE ALTITUDE VERSUS FLIGHT TIME
97
n
€ * Y
a#
z - 0 0
c c .- - a M
t + c c CI, .-
0 0 cu
0 0 v
0
0 - o E o = G -
a# CI, c 0 a
0 0 0 * L
0 0 0 cu c
W (3 z a a n
a
z 3 0
(3
v) 3 v)
W > W a 3 I-
a
- 5 a W 4 u I W > I 0 z 3
4
cu 0 cu I
v)
-
a
a
4 . (3
LL -
98
~~~
Velocity (kmlsec) S - I Y B Cutoff
7
6
5
4
3
2
1
0
Space Fixed
0
0 0 S - I B Cutof f 0
1 4 b & 0 / /*E;rth 0 Fixe! I
1
m 0 loo 200 300 400 500 600
Flight Time ( s e d
FIG. 5. AS-202 SPACE FIXED AND EARTH FIXED VELOCITY VERSUS FLIGHT TIME
99
S - E B Cutoff
\ \ \ 1, I \ c
I /
Path Angle (deg from local ver t ica l )
90
75
6 0
45
30
15
0 0 100 200 300 400 500 600
Flight Time ( s e d
FIG. 6. AS-202 SPACE FIXED AND EARTH FIXED PATH ANGLE VERSUS FLIGHT TIME
II)
100
t
Dynamic Pressure (io3 N/m2)
FIG. 7. AS-202 DYNAMIC PRESSURE VERSUS FLIGHT TIME
101
102
0 P b I L)
0 E
I v)
0 3
A
!
!!!
k
b
b e
B P
ip >
w 0 4
t
0 z
a
Vehicle Fixed Acceleration (m/sec2
0 200 400 600 Flight Time ( s e d
FIG. 9. AS-202 VEHICLE FIXED ACCELERATION VERSUS FLIGHT TIME
103
n
z m Q) U Y
Q) U c 3-
W e 0 4 W > z W
I- o a e =E -
o
e a - a
W a I- v)
a
104
f
t
0 v)
105
c 0 0 a E
m
c - w I
0 V
m
kl I
v)
Y
n
4) 0 Y
c H -) I 0 .-
v), o 4 c
w I
0 w v)
4) -
- E
m L 4)
0-
333 0
0 0 cu
0 pc)
0 cu 0 r
0
wu, r-
aa
0 0 0 r
a
106
c
c clr a3 U Y
c 0 .- c
P= Q) - W
0 W
0 v)
0 M
0 cu 0
n CT o) U Y
c 0 .- c
E- o)
W -
c 0 0 I
I I I I I I
0 0 3 Y 0 0 0 0 CD v) * M
0’s Oo) Wu, Y-
o)
0 .E o+ d= Y,
0 *= O b cu
c
Y
0 0 0 Y
0 0 QD
0 0 W
0 0 d=
0 0 cu
0
108
4
n
Y E Y
0
c 0 a +
0 0 c 0 a
v) eJ
c
- 0 0 rc) r
0 0 0 r
0 0 v)
0
109
0 0
I I
m
v) v) I I
I .
c t 0
v) -
0
aa
cu
0 0 0 T
0 0 e
'0
110
4
* I
3 I I I I W
c 0 0 c 0 0
v) cu - 0 0 0 T
0 0 v)
111
APPENDIX A
I t e m S-IB Stage, Dry S-IB/S-IVB In ters tage ( Inc l . re t ro-rocket prop.) S-IVB Stage, Dry (at ground ign i t ion ) Instrument Unit a t G. I.
AS-201 MASS DISBURSEMENT
Item
S-IB Stage a t Ground Ign i t ion (G. I.) S-IB/S-IVB In ters tage a t G. I. S-IVB Stage a t G. I. Vehicle Instrument Unit (V. I. U. ) a t G. I. Spacecraft a t G. I.
F i r s t F l igh t Stage a t G. I. S-IB Thrust Buildup (10887 LOX, 3237 RP-1)
F i r s t F l igh t Stage a t Lif t -off S-IB Mainstage (611014 LOX, 270500 RP-1) S-IB Frost S-IB Gear Box Consumption (RP-1) S-IB Seal Purge S-IB Fuel Additive (oroni te )
S-IVB Frost S-IB I. E. T. D. (761 LOX, 1400 RP-I)
F i r s t F l igh t Stage a t 0. E. C. 0. Signal S-IB OETD To Sep. (579 LOX, 1061 RP-1)
F i r s t F l igh t Stage a t Separation S-IB Stage a t Separation S-IB/S-IVB In t e r s t age S-IVB Aft Frame S-IVB Ullage Rocket Propellant S-IVB Detonation Package
Second Fl ight Stage a t Ign i t ion S-IVB Thrust Buildup Prop (318 LOX, 111 LH2) S-IVB LH2 S t a r t Tank S-IVJ3 Ullage Rocket Propellant
Mass (lbm) 91550 6408 23264 45 02
i
Mass (lbm)
1003833 6408 254725 45 02 56906 -------------
1312249 - 88 15 14 - 1000 -722 -7 - 24 -2161 - 100 4267 22
-------------
- 1640 ------------- 425082 - 102641 -6408 - 25 -77 -4
315927 -429 -4 - 105
-------------
112
. APPENDIX A (CON")
8
I t e m
Second F l igh t Stage a t L i f t - o f f (90% Thrust) S-IVB Mainstage (189943 LOX, 35462 LH2) S-IVB Ullage Rocket C a s e s S-IVB Aux. Prop. - Power Roll Control Launch Escape System
* Second F l igh t Stage a t Guidance Cutoff S igna l S-IVB Thrust Decay (117 LOX, 27 LH2)
Second F l igh t Stage a t E. T. D.
Second F l igh t Stage a t Separation S-IVB Stage a t Separation v. I. u. Adapter (less ring)
Spacecraft
2. SPACECRAFT MASS BREAKD(IWN
Item 1
Command Module Service Module Less Propellant Service Module Propellant Adapter Ring Tota l a
Mass (lbm)
315389 -225405 -220 -4 -8521 --------------
81087 -28169 -45 02 -3713
113
APPENDIX B
AS-202 OPERATIONAL TRAJECTORY
I G M PRESETTINGS
t INITIAL
VALUE
28.88
UNITS
s ec
DESCRIPTION SYMBOL
*,--
T I i
'3 i
T3
Vex 1
Vex3
VT
'VT
'VT
'VT
XVT
+VT
%GT
'VGT
'VGT
'VT
..
..
E
Trl
114
T i m e from t i m e base 3 (T3) t o i n i t i a t e I G M guidance
299.25
129.9
322.733
Time-to-go for f i r s t IGM s tage sec
sec
sec
Time-to-go f o r second IGM s t a g e
Nominal value of m/m a t i n i t i a t i o n of second I G M s tage
4165.45 m/sec 5-2 exhaust ve loc i ty (go.1 ) for f i r s t I G M s tage SP
m/sec 4209.49 5-2 exhaust ve loc i ty (go.Isp) for second I G M s tage
6800 mlsec Terminal ve loc i ty f o r IGM equations. 5-2 cutoff veloci ty mangitude L
0.
6592538.9
0.
6783.4753
473.85 03
0.
0.
-9.17 139
0.
46.
Desired cutoff posi t ion vector
components i n IGM coordinate C
system m
m/sec
m/sec
m/sec
2 m l s
s ec
Desired cutoff ve loc i ty vector
components i n I G M coordinate
sys t e m
Terminal g rav i t a t ion vector
components i n I G M coordinate
system
V a l u e of T3i t o i n i t i a t e I G M A 7 guidance
6667 Mission dependent constant mu l t ip l i e r f o r terminal range angle equation
.
LVDC SYMBOL
Tr2
Tr3
TSMC
T2
TNOM
AP1
AP2
AP3
AP4
AP5
AP6
AP7
AP8
AP9
*---
FLAG
ACC
A,
A 1
INITIAL VALUE
0.
0.
19 0
APPENDIX B (CONT)
0.
R-ASTR
.409699
- .91221574
- .0028925398
.9 1 2 1 15 05
.40969873
-.013894577
.013859920
.0030542649
.99989925
236.1
2
. 0001
299.25
512.90
-1.28
sec
s ec
N/ D
m / s 2
DESCRIPTION
Mission dependent constants
for N and N equations 3 4
T i m e t o i n i t i a t e t h rus t misalignment correct ion
Mult ipl ier of %and B i n x x I
and X equations Z
Transformation matrix from
navigation coordinate system t o
the (X,, Y4, Z ) coordinate system 4
Time from time base 3 (T3) t o sample F/M for IGM s taging
Number of successive F/M decreases required for IGM staging c r i t e r i a
Tolerance fo r F/M comparison
Constants fo r updating second
s tage time-to-go for perturbed
EMR s h i f t t i m e
115
LVDC SYMBUL - PCO
7 10
*CO
INITIAL VALUE
30
558.64
30
133922
238.71
195.449
* LVDC symbol not defined
116
APPENDIX B (CONT)
UNITS
see
-
s ec 1 kg/sec
kg/sec
DESCRIPTION
Back-up t i m e for IGM staging
Constants for a r t i f i c i a l
T3 mode
a
Mass a t a c t i v e guidance i n i t i a t i o n
Average flow r a t e of f i r s t IGM s tage
Average flow r a t e of second IGM s tage
APPENDIX C
THREE-SIGMA LAUNCH VEHICLE PERTURBATIaS
I. S-IB STAGE PERFtXWANCE
A. Propulsion Group
Propellent Loading Mass Thrust and Flowrate I s p (flowrate) Mixture Ratio Fuel Bias
High Surface Wind Low Surface Wind High Ambient Surface Temperature Low Ambient Surface Temperature
B. Non-Propulsive Group
Non-Propellent Mass Thrust Misalignment (pi tch) Thrust Misalignment (yaw) Axial Drag Coefficient Headwind Tailwind Left Crosswind Right Crosswind
11. S-IVB STAGE PERFORMANCE
A. Propulsion Group
Prope 1 len t Loading Mass Thrust and Flowrate
Mixture Ratio (+ Fuel) Mixture Ratio (- Fuel) Mixture Ratio (+LOX) Mixture Ratio (-LOX)
ISP
B. Non-Propulsion Group
Non-Propellent Mass Latera l cg o f f se t s ( i n plane) La tera l cg o f f se t s (normal) Thrust Misalignment (p i tch) Thrus t Misalignment (yaw)
2. 3% 21.5% If: .9 sec +2000 lbm f u e l -1000 lbm fue l (Ref. (Ref . (Ref. (Ref.
c
f 304 lbm * .62 deg f .62 deg f 10% (Ref. (Ref. (Ref. (Ref.
* 1% * 3% * 3.12 sec (Ref . (Ref . (Ref. (Ref.
189 lbm, -122 lbm 2 .05 meters f .OS m e t e r s k .62 deg * .62 deg
117
. .
REFERENCES
1.
2.
3.
4 .
5.
6.
7.
8.
9.
10.
118
MSFC Flight Mission Directive, Apollo-Saturn 202 Mission, June 22, 1965 (U).
MSFC SA-202 Saturn Vehicle Data Book, May 15, 1965 ( C ) .
Memorandum, R-P&VE-VAW-66-49, Saturn IB AS-202 Final Predicted Mass Characteristics, June 6, 1966 (U).
Memorandum, R-P&VE-PPE-66-M-96, AS-202 S-IB & S-IVB Final Flight Propulsion Prediction, May 18, 1966 (U).
Memorandum, R-P&VE-PPE-66-M-102, Douglas Aircraft Final S-IVB 202 Propulsion Flight Prediction, June 2, 1966 (U).
MSFC-111-4-423-2, Launch Vehicle Digital Computer Equation Deftning Document (U) . Memorandum, R-AERO-FM-20-66 , S-IB Steering Commands , S-IVB IGM Pre-Settings and Guidance Terminal Conditions for AS-202 Vehicle, May 12, 1966 (U).
NASA, TMX-53266, Launch Vehicle Guidance Equations for the Saturn IB, SA-202, May 24, 1965 (U).
Memorandum, R-P&VE-PPF-65-M-142 (628), S-IB Stage Sequence Change, November 17, 1965 (U).
Memorandum, R-ASTR-F-66-1, AS-202 m/F Filter, April 1, 1966 (U).
.I
8
. .
APPROVAL TMX- 53470
AS-202 LAUNCH VEHICLE OPERATIONAL FLIGHT TRAJECTORY
Pamelia B. Pack
The information i n t h i s report has been reviewed for s ecu r i ty c l a s s i f i ca t ion . Defense or Atomic Energy Commission programs has been made by the MSFC Security Class i f ica t ion Officer. determined t o be unclassif ied .
Review of any information concerning Department of
This repor t , i n i t s en t i r e ty , has been
This document has a l so been reviewed and approved for technical accuracy.
APPROVAL :
G. Wittenstein Chief, Trajectory Section
Chief, Flight Mechanics Branch
CONCURRENCE :
Chief, Fl ight T e s t Analysis Div
&&- L E. D. Geissler Director , Aero-Astrodynamics Laboratory
Chief, Tracking & Orbital Analysis B r
I; . ; c' ',-A c 1;. -?.A
L e w i s L. McNair Chief, Projects Office
119
. MS-IP MS-IPL (8) MS-T, Mr. Roy Bland ( 6 ) MS-H
cc-P
DAC, M r . P. Dixson
MSC - FM, M r . Incer to FM-13, Mr. Bryant FM, M r . Ricks F'L, M r . C. F i l l ey EG2, M r . D. Cheatham EG22, Mr. J. Funk PM5, M r . R. McKann PD4, M r . A. Cohen FM2, M r . R. Ward PM2, M r . C. Perrine FA, M r . C. Kraft FC, M r . J. Hodge FC, M r . G. Paules I-MO-F, M r . Casey (3) FLY Mr. Thompson FM, Mr. J. Mayer FM3, M r . Huss FM3, Mr. Skopinski FM7, Mr. R. Nelson FM7, Mr. J. McAnulty FM3, M r . R. McAdams
KSC - K-DIR, Dr. Debus K-TEC, D r . Knothe (3) INS-1, D r . Bruns INS- 13, Mr . Je len TPR-72, M r . Bishop
HQTRS
Bellcorn, 1021, Mr. Sperry Bellcorn, 1124, Mr. Wagner OMSF-MAO, M r . Susson OMSF-MAR, M r . Lemke OMSF-MAT, M r . Savage OMSF-MB, M r . Winn OMSF-MAS, Unassigned OMSF-MO-1, Gen. Bolender
NASA Resident Apollo Spacecraft Project . f f i c e r MIT Instrumentation Laboratory Cambridge, Massachusetts 02142 Attn: M r . Edward Copps & M r . John Dahlen
NASA Langley Research Center Langley S ta t ion Mail Stop 1 1 7 Hampton, Virginia 23365 Attn: M r . A. Matson
Douglas Ai rc ra f t Company Attn: Mr. Webb Miller Huntington Beach, Cal i forn ia 92646
I B M Confidential Documentation Center Building 1 Sparkman Drive Huntsvi l le , Alabama 35807 Attn: Mr. J i m Chr i s to f f e l
S c i e n t i f i c and Technical Inf . F a c i l i t y (25) Attn: NASA Rep. S-AK/RKT P.O. Box 33 College Park, Maryland
b
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A d d r e s s e e s : DEP-T, e. R e e s
I-I/D-MGR, C o l o n e l James I - I / I B - E , Mr. V r e u b I - I / I B - E , Mr. C a l l a w a y I-I/IB-SIVB, Mr. G o d f r e y I-MO-MGR, Dr. Speer (2) I -MO-0, Mr. K u r t z I-MO-0, Mr. Naumcheff I-MO-R, k s . M c N a i r I-TO-DIR, Mr. Richard
I-E-MGR, Mr. B e l e w R-AERO-DIR, Dr. Geissler R-AERO-DIR, Mr. Jean
R-AERO-AT, Mr. Wilson R-AERO-ATA, Mr. E l k i n
DEP-A, k. Gorman
I-V-IGR, Dr. R u d o l p h
R-AERO-A, Mr. D a b
R-AERO-AD, Mr. Emley R-AERO-AD, Mr. Dunn R-AERO-ADD, Mr. Nunley R-AERO-D, Mr. H o r n
R-AERO-A, Mr. Holderer R-AERO-G, Mr. B a k e r R-AERO-D, Mrs. C h a n d l e r
R-AERO-DAG, Mr. T e l f e r R-AERO-DAP, Mr. Cr&n R-AERO-F, Mr. Lindberg (2) R-AERO-FF, MI-. Hagood (4) R-AERO-FF, Mr. Sheats R-AERO-F, Mr. Stone R-AERO-FKT, Mr. H a r d a g e (30) R-AERO-FMT, Mr. Wittenstein R-AERO-FMT, Mr. Weiler R-AERO-FKI', Mrs. Pack (5) R-AERO-FMT, Mr. B r o t h e r s
R-AERO-FMR, Mr. Leonard (2) R-AERO-E'MR, Mr. Crafts R-AERO-FO, Mr. H i l l (5) R-AERO-P, Mr. T e a g u e (10) R-AERO-P, Mr. M c N a i r R-AERO-P, Mr. Ledford (15) R-AERO-T, Mr. R e e d
R-AERO-Y, Mr. Vaughan R-AERO-P, Mr. Tinius
R-AERO-AU, Mr. Yo-
R-AERO-DD, Mr. R y a n
R-AERO-FF, Mr. M c N i e l
R-AERO-T, W. Crmmings
R-TEST-DIR, Mr. H e i m b u r g R-ME-DIR, Mr. K u e r s R-SAI, E a r l B u t l e r R-P&VE-DIR, Mr. Lucas R-P&VE-DIR, Mr - H e l l e b r a n d R-P&VE-XJ, Mr. G r i n e r R-P&VE-P, MI-. Paul R-P&VE-VOI, Mr. B y e r s R-P&VE-AVA, Mr. Denton R-P&VE-VS , Mr . Schulze
R-P&VE-SLY Mr. S h o w e r s R-P&VE-S, Mr. Kroll
R-P&VF.-S, MI-. H u n t R-P&VE-PTD, Mr. Hastings
R-P&VE-VAW, Mr. Marmann R-P&VE-VAN, Mr. Scott R-P&VE-PPE, Mr. McKay R-P&VE-PPE, Mr. Igou
R-P&VE-DIR, Mr. Palaoro (2)
R-QUAL-DIR, Mr. G r a u R-COMP-DIR, D r . H o e l z e r
PA-Mr. Jones I-MO-0, Mr. Ladner R-ASTR-DIR, Dr . H a e u s s e r m a n n R-ASTR-I, Mr. H o b e r g R-ASTR-F, Mr. B l a c k s t o n e R-ASTR-F, Mr . H o s e n t h i e n R-ASTR-F, Mr. Scofield
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R-ASTR-G, Mr. Mandel R-ASTR-G, Mr. Thamason R-ASTR-N, Mr. Moore R-ASTR-R, Mr. T a y l o r
R-COW-RRT, Mr. Crafts
R-ASTR-FO, Mr. Mink
R-ASTR-NG, Mr. Chubb
R-ASTR-S, W. Mack R-ASTR-M, Mr. B W h I-I/IB-TF, Maj. R m i n e k (2) R-ASTR-E , M r . Fichtner
R-ASTR-EA, Mr. Greer R-ASTR-IR, E-. B a r r R-ASTR-IR, Mr. E d e n NAA, Mr. T o o k e r (2)
R-ASTR-IRD, Mr. E l y
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