apollo experience report abort planning

8/8/2019 Apollo Experience Report Abort Planning

1/37

APOLLO EXPERIENCE REPORT -A B O RT P L A N N I N G

by Charles T . Hyle? Charles E . Foggatt?and Bobbie D, Weber 4

Mdnned Spacecra, Center ,

Houston, Texas 77058c e

N AT I O N A L A E R O N AU T I CS A N D S PA CE A D M I N I S T R A T I O N WA S H I N G TO N ,D. C. J U N E 1972


2/37

TECH LIBRARYKAFB, NM

1. Report No. 2. Govern ment Accession No.

NASA "I D-6847

APOLLO EXPERIENCE REPORTABORT PLANNING

Charles T. Hyle, Charles E. Foggatt, andBobbie D. Weber. MSC

9. Perfo rming Organizatio n Name and Address

Manned Spacecraft CenterHouston, Texas 77058

2. Sponsoring Agency Name and Address

National Aeronautics and Space AdministrationWashington, D. C. 20546

5. Supplementary Notes

-3. Recipient's Catalog No .

5. R e m r r n a t PJune 1972

6. Performing Organization Code

. - .

8. Performing Organization Re port NO.

MSC S-307-

10. Work Unit No.

924-22-20-00-72. - - . . -. . . -

1. Contract or Grant No.

13. Type of Report and Period Covered

Technical Note~ _ _ - - _ _14. Sponsoring Agency Code

The MSC Dir ec tor waived the u se of t he International System of Units (SI) fo r thi s Apollo

Experie nce Report, becaus e, in his judgment, the use of SI units would impair the usefulne ssof the re port o r resul t in excessive cost.- ~

6. Abstract

Definition of a prac tica l return-to- earth abor t capability was requir ed fo r each phase of an Apollomission. A description of the bas ic development of the complex Apollo ab ort plan is presented inthis pap er, The pro ce ss by which the retu rn-t o-ea rth abor t plan was developed and the constraining factors that must be included in any abort procedure are al so discusse d. Special emphasis isgiven to the descript ion of crew warning and escape methods for each mission phase.

~ .. - _ _7. Key Words (Suggested by Author(s)) 18. Distribution Statement

* Contingency Mi ssio n PlannfngLunar Trajector ies

* Space Flight Safety

9. Security Classif. (of this report) 20. Security Classif. (of this page) 21. NO. of Pages 22. Price

None.None 35 ] $3000'

~_ _

*For sale by he National Technical Inf orm ati on Service, Springfield, Virginio 22151


3/37


4/37

CONTENTS

Section Page

SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ABORT PLANNING PROCESS . . . . . . . . . . . . . . . . . . . . . . . . . . .

. .LAUNCH PHASE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Constraints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Maneuver Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Return to E arth . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

TRANSLUNAR INJECTION . . . . . . . . . . . . . . . . . . . . . . . . . . . .Constraints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Maneuver Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Return to Earth . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

TRANSLUNAR COAST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Constraints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Return to Ear th . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Apollo 13 Tra nsl una r Coast Abort . . . . . . . . . . . . . . . . . . . . . . .

LUNAR-ORBIT INSERTION AND LUNAR OR BIT . . . . . . . . . . . . . . . . .Constraints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Maneuver Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Return to E ar th . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

LUNAR DESCEN T AND ASCENT . . . . . . . . . . . . . . . . . . . . . . . . .Descent-Orbit-Insertion-Maneuver and Powered-Descent-

Maneuver Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Lunar Module Abort to Rendezvous and CSM Res cue . . . . . . . . . . . . .

9

9

10

12

13

13

14

15

15

15

16

18

21

22

23

Powered Ascent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 4


5/37

S e c t i o n Page

TRANSEARTH INJECTION A N D TRANSEARTH COAST . . . . . . . . . . . . . 25

C o n s t r a i n t s . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25Maneuver Moni to r ing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25Return to Ear th . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 6

CONCLUDING REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

i v

........1.111111111111111.1111.11.11 . . . I I


6/37

TABLES

Table Page

I ABORT PLANNING CONSTRAINTS . . . . . . . . . . . . . . . . . . . 3I1 SPACECRAFT DISPLAYS USED FOR ABORTS . . . . . . . . . . . . . 6

I11 PRIMARY CONSTRAINTS DURING TRANSLUNAR INSERTION . . . . 10IV PRIMARY CONSTRAINTS DURING LUNAR-ORBIT INSERTION . . . . 16

FIGURES

Figure Page

1 Abort planning p ro ce ss . . . . . . . . . . . . . . . . . . . . . . . . . 22 Saturn-Apollo vehicle components . . . . . . . . . . . . . . . . . . . . 5

3 Typical tim e history of launch-vehicle dynamic var iab lesfollowing an actuator failure . . . . . . . . . . . . . . . . . . . . . 7

4 Ground-control t rajectory l imits . . . . . . . . . . . . . . . . . . . . 7

5 Apollo spacecraft launch abort modes . . . . . . . . . . . . . . . . . . 8

6 Ground-monitoring display of near -in ser tio n abo rt

decisions fo r Apollo missio ns . . . . . . . . . . . . . . . . . . . . 97 Apollo mission profile . . . . . . . . . . . . . . . . . . . . . . . . . . 9

8 Basic crew-maneuver-monitoring technique . . . . . . . . . . . . . . 1 1

9 Malfunctions duri ng TLI . . . . . . . . . . . . . . . . . . . . . . . . . 1 2

10 Aborts during TLI using an earth-f ixed thrust-vectorattitude alinement . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

11 Apollo 11 trade-off display . . . . . . . . . . . . . . . . . . . . . . . 1 4

1 2 Operat ional abo rt plan from launch to TLC . . . . . . . . . . . . . . . 15

13 Visual LO1 attit ude check . . . . . . . . . . . . . . . . . . . . . . . . 1714 Effects of at t i tude-referenc e fai lu res on pericynthion al t i tude . . . . . 1 7

V


7/37

Figure Page

15 P r e m a t u r e LO 1 shutdown tra ject orie s . . . . . . . . . . . . . . . . . . 18

16 P r e m a t u r e LO1 shutdown trajector y cl as se s . . . . . . . . . . . . . . 1817 Abort AV r equ i r emen t s fo r a stable el l ipse . . . . . . . . . . . . . . 19

18 Minimum AV requirements for LO 1 mode I a b o r t s . . . . . . . . . . 19

19 Sum mar y of minimum-fuel abo rt capability as a functionof LO1 burn t ime . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

20 Th e LO 1 mode I1 f i r s t -burn AV as a function ofLO1 burn t ime . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

21 Time between burns and the pericynthion altitude followinga nominal first burn for an LO 1 mode I1 abor t . . . . . . . . . . . . . 21

22 The AV for midcourse cor rec t ions as a function of TEIburn t ime at crew takeover for various pi tch dr if ts . . . . . . . . . . 25

23 Operational abort plan from lunar a r r iva l to TEC . . . . . . . . . . . . 2 6

vi


8/37

ACRONYMS

A GS

APS

CDH

CM

c SI

CSM

DO1

DPS

A H

Ah

A V

EDS

E M S

E P O

g

G&N

GTC

I MU

LES

LM

LO1

L PO

LR

MCC

abort guidance system

ascent p ropuls ion sys t em

constant differential height rendezvous maneuver

command module

coelliptic sequence initiation rendezvous maneuver

command and s erv ice module

descent o rb i t inser t ion

descent propulsion syste m

altitude difference between PGNS and LR al t i tude determinat ions

differential height

chara cter i s t ic veloci ty change

emergency detect ion system

entry monitoring system

ear th park ing orb i t

load factor (force divided by weight)

guidance and navigation

guidance-determined thrus t command

iner t ia l mea surement uni t

launch escape system

lunar module

lunar-orbi t insert ion

lunar park ing orb i t

landing radar

midcourse cor rec t ion

vii


9/37

MSFN

PDI

PGNCS

PGNS

P37

RCS

RR

RTCC

s-IC

S-II

S-IVB

SPS

B

TEC

TEI

TLC

TLI

Manned Space Flight Network

powered descent initiation

pr im ar y guidance, navigation, and control sy st em

primary guidance and navigation system

command module computer pr ogr am number 37, abort prog ram

react ion control system

rendezvous ra dar

real-time computer complex

Saturn IC, f i rs t s tage of the Saturn V launch vehicle

Saturn 11, second stage of the Saturn V launch vehicle

Saturn IVB, third s tage of the Saturn V launch vehicle

ser v ic e propuls ion sys te m

burn t ime

t r anse a r t h coa st

t ra nse ar th in jec tion

t rans lunar coas t

t ranslunar inject ion

viii

L


10/37

APOLLO EXPERIENCE REPORT

A B O RT P L A N N I N G

B y C h a r l e sT. Hyle , C ha r l e s E . Fogga tt ,*a n d B o b b ie D . We b e r

M a n n e d S p a ce c ra ft C e n t e r

S U M M A RY

To att ain the high confidence lev el requ ired of th e safe ty asp ec ts of th e ApolloP rog ram, a pract ical return-to-earth abort capabil i ty w a s provided fo r each of th evari ous mission phases. The development of su ch an abor t capability is especiallycomplex becaus e of th e myria d of potential mis sion contingencies that c an be identified.Tra ject orie s that sat isfy the requirements and constraints of a safe re turn to ear th andthat extend to all possible preabort condit ions a r e analyzed. The spacecraft systemcapabilities, ground-support - equipment capabilities, flight-crew perform ance, andother operational conside rations and limitations ar e then superim posed. The resultan tinteraction ha s led to the development of sev era l distinct abor t techniques and powered-flight monitoring procedures. These techniques and procedures ensu re that a saferetu rn-t o-e arth capability exis ts throughout the spe ct rum of an ticipate d off -nominalmission conditions.

I N T R O D U C T I O N

Although the maj or objecti ves of the Apollo Pr og ra m a r e to land men on the moon,explore the lunar s urfa ce, and ret urn the men safely to earth , the safety of the flightcre w has always been of paramount importance. The stringen t requ irem ent for crewsafety dictates the ne cessity of a s much o r m ore contingency planning for abort s i tuations as is provided for the nominal miss ion . The contingency planning to ensu re thatthe crew can always abort the mission and ret urn safely to ea rth is accomplished whenan adequate cr ew warning technique and a method of es ca pe have bee n defined. An ade

quate cr ew warning technique is difficult to achieve during powered flight but ha s beenaccomplished by having the crew o r ground-control person nel (o r both) monitor crit icalspacecra ft sys tems and para mete rs . The method of escape from a contingency situation is often dete rmin ed by the amount of pa ra me te r deviation fro m the nominal allowedby the monitoring techniques and paramet er-devia tion limits. A method of e sca pe isre fer red to a s an abo rt mode.

*Present ly a manage ment tr ain ee with the Wicks Corporati on, Saginaw, Michigan.


11/37

Definition of a pract ical monitoring technique and abort mode for the r ea lm ofpotential contingencies is the res ult of an ite rat ive an alysi s cy cle in which the effects ofvar ious cons t ra in t s a r e cons idered . Such constraints may be caused by either hardware or operat ional l imitations. The cha rac ter is t ics of e a r t h o r moon t ra jec tor i es (orof combinations of e art h and moon traj ect ori es) d et er min e to a great extent the set ofinitial conditions upon which the ha rdwar e and operational const raint s a re superimposed.

However, the t raje ctory chara cter i s t ics a r e als o funct ions of operat ional constraints(for example, a speci fic lunar-landing-si te req uirem ent), The achievement of a soundcontingency return-to- earth abort plan is therefore obtained by considering the interaction of many c ons tra int s and by satisfyi ng the obj ectiv es of an adequate cr ew warningand escape capability.

Because of th e dive rsity of the co nstra ints encountered during the various phase sof a lunar miss ion, independent abort plans must be pr ep ar ed fo r each of s eve ral di scr et e mission parts . Basical ly, these mission pa rt s include the powered-fl ight phases(launch, translunar injection (TLI), lunar-orbit insertion (LOI), lunar descent andascent , and tra ns ea rth injection (TEI)) and the coast phases (eart h parking orbi t (EPO),translunar coast (TLC) , lunar parking orbi t (LPO), and tr an se ar th coast (TEC)) .

Acknowledgment is made to W . Bolt for contributing the section on lunar descentand ascent and to R . Becker and J . Alexander for the section on lunar module abort torendezvous and command and ser vi ce module res cue .

A B O RTPLANNING PROCESS

For th i s paper, an abor t is defined as the recognition of an intole rable situ ationand the performanc e of the act ivi t ies nece ssa ry to terminate the mission and retu rnthe cre w to earth. An al ter nate mission is a continuation of the flight, usually with le s s

ambitious objectives than were originally planned.

The pr oc es s used to formulate asound abort plan is shown in figure 1. Inthe abort planning pr oc es s, the spacecraftand launch-vehicle hardware capabilitiesand the missio n opera tional cons train ts andobjec t ives a reused as a basis for s imulatedabort t raj ect ori es to invest igate displayadequacy, entry conditions, ti me requi rement s, landing points, and so forth.Throughout this pr oc es s, the f l ight crew,

flight - control personnel, safety personnel,and other responsible groups conduct reviews and discuss and modify the simulationresu l t s un t il al l groups a re satisfied. Thetotal abort plan fo r a part icular missioneventually con sis ts of sev er al detailed document s. Each document is concerned with aunique responsibility, but al l a r e based onthe sa me assumptions and a r e consistent

_ _G r t r a i n t rand 1 Trajecto?

considerations s imulat ionr Abort plan~ 1

~-- .-

Hardr+are Targeting Abort report

l a u n c h v e hi c le - Attitudes hl iss ion ru le s

- 1

Spacecraft l l a l f u n c t i o n r Crew proceduresOperational AV Recovery pian

Sequencing Real-lime program

Flight l imi t s

hlodificatton I

_ _Joint panel , l lode

meetings st y it cho ve r

\ l o r k i n q g ro up s 1 isplaks I i n i l i iI

Crew s imula t ions

Real-timeprogram

Fi gu re 1. - Abort planning pro ce ss ,

2


12/37

with the other documents, Among the documents that most closely re flect the totalabort philosophy a r e the Abort Report , the Mission Rules, the Crew Pro ce du re s Manual,the Data Pr io ri ty Contingency Techniques Document, and the Flight Limits Document.

Development of the abor t plan re qu ir es that detailed consid eration be given to twotypes of constr aints . Although the se const raint s a r e generally classifie d as ei ther hardware o r operational, each category is composed of a var iet y of conside ration s (table I) .

TABLE I. - ABORT PLANNING CONSTRAINTS

ISystem o r category I ConstraintsHardware

Launch vehicl e(S-IC, s-11,and S-IVB)

Spacecraft(command andser v ice module

(CSM) , andlunar module(LM))

Propulsion

Computer

Guidance and cont rol

Structur a1

Displays

Propulsion

Computer

Guidance and control

Optics

Structural

The r ma l

Aerodynamics

. .-

Emerg ency detection sy st em (EDS) andredundant capabilities

Redundancy, update, and tar get ing

EDS, gimbal lock, and hardo ver

EDS and shutdown sequence

Caution and warning, switc h configuration ,at t i tude, and entry -monitoring- sys tem(EMS) quantities

Typ e, pe r f o r manc e, backup c apab ilitie s ,and duty cycles

Targeting, display, sto rag e, and navigation

Performance, procedures, backup sy ste msautopilots, i nertia l platf orms, gimballock, and hard over

Navigation, s extan t, tele sco pes , and fieldof view

Couch su pp or ts , landing, and CSM/LMinterface

Prote ction limitations, py rotechnics, andheat shield

Stability and tr i m and lift/dragcharac te r i s t ics

3


13/37

TABLE I. - ABORT PLANNING CONSTRAINTS - Concluded

-_ ~

r r p i s t e m o r category I ConstraintsHardware - Concluded

Window/c r ewgeometry

Consumables

Sequencing

-.

.. __ ._ - - . .-

Tra jec tory

Lunar - landing s i te s

Al te rna te miss ons

Landing andrecovery

Communications andtracking

Environmentalsur roundings

Human fact ors

P ro c edu re s

Range safety.

Crew visibi l ity to horizon, manual takeover, reticles, and CSM and LM blockage

Elec tr ica l power, environmental syst em,and propulsion

Att i tude requirements , act ivat ion t imes,and procedu res

~ _ .

Fre e- r e tu rn miss ion , o rb i ta l a l ti tude ,launch windows, flight ti me s, and entr yco r r i do r

Launch windows, lun ar -orbit inclinations,earth/moon geom etry , and lighting

Object ives, lunar operat ions, andphotography

Geography, lighting, communications,logist ics , and medical support

Systems monitoring and ground targetingand command capability

Atmospheric properties, winds, lighting,weather conditions, radiation, andme teo r i t e s

Crew schedules, c rew accelerat ion and decelerat ion tolerances, crew and ground-cont ro l respon se t imes

Separation techniques, recontact avoidance,s impl e and rel iable for t rain ing proficiency, and mission-to-mis sion carr yove r

Land-impact avoidance and launch windows-~

4


14/37

Malfunctions in any of the hardware i t ems mus t be cons idered as additional cons t r a i n t s . Obviously, so me of thes e const rai nts a r e meaningful only duri ng one part icula r mission phase; however, other con strain ts always exist . Because of the la rgenumber of cons t ra in t s that mus t be considered, abort plans and techniques must bekept as s imple as possible .

The constr aint of t im e crit ica lity has meaning both in monitoring (crew warning)and in ab ort (escape) mode analysis . Some contingency situatio ns allow mo re ti me thanoth ers f or making the abo rt decision (monitoring) and fo r making the r etu rn to e arth(abor t method).

LAUNCHPHASE

Constraints

Becaus e of potential launch-vehicle breakup and subsequent explosion possibili ties,the launch phase is probably the most dan gerous pa rt of any manned spa ce flight. The

atmospheric environment is a major operat ional constraint that requires special escape-hardware design considerat ions, The launch escape syste m (LES) is the mechanismthat provides this atmospheric es cape capabil i ty. The LES is shown with other majorSaturn-Apollo vehicle components in figure 2 . The planned o r nominal launch trajector y al so deter min es the kind of environment fro m which the available hardware mustescape. After the space cra ft exits the significant pa rt of the atmosphe re, malfunctionsthat might occur a r e s ignificant ly less t ime c r i t i ca l ; that i s , s t ruc tu ra l b reakup wi th theassoc iated overp ressur e and f i r e hazard is less probable becaus e of the reduced aer odynamic loading. Every majo r sys tem o r category l is ted in table I influences the launchabort plan in so me way.

aunch escape system

Lunar

module

S a t u r n P.Apollo vehicles

Figure 2 . - Saturn-Apollo veh iclecomponents.

Maneuve r Monitor ng

Ground - cont r ol and crew - monit o r ingactiv ities mu st include distinction betweennominal and off -nominal flight performancein real t ime. The deg ree to which nominalflight conditions may be allowed to det eriorate is determined by the escape capabili tyavai lable at the t ime. There fore, the monitoring requir ements and abort methods areclosely related. The interact ion of t h e s etwo fac ets determi nes both the l im its fo rmaneuver monitoring and the adequacy ofa given escap e technique. Th re e types ofcontingencies are of concern, and all maybe considered t he res ul t of launch-vehicleproblems.

The f i rs t type of contingency concernsp r e ma t u re o r l a te (e i ther ac tua l o r apparent)thrust terminat ion of the Saturn I1 (S-11) o r

5


15/37


16/37

---

50

A n

The au tomat ic sys tem s - complemented by the c rew who ar e aided by onboard displays,window views, and physiolo gical cu es - com pri se an adequate sys tem fo r contendingwith rapi d launch-ve hicle deviations. The adequacy of th is sy st em is ensured by compute r s imulat ions of the most probable vehicle fai lu res . Thes e s imulat ions a r e made toes tab l i sh the requi red l imi t s for an au tomat ic o r manual abort . For the mos t par t ,rapid deviations, such as an engine hardover, would resu lt in an LES abort . Typicalre su lt s of th is type of contingency a r eshown in figur e 3 in t h e f o rm of rate andatt i tude excursions as a function of timefr om l if t-off.

The thi rd type of contingency cons i s t s of those malfunctions with effect s thata r e not immediately obvious to the c rew ;these mal funct ions may be re fe r red to asslow deviations. Slow deviations a r ecaused by at t i tude-reference problems o rguidance fai lur es and a r e not as t ime c r i ti ca l as the malfun ctio ns of th e f i rs t o rsecond contingency type s. The re su lt s ofsuch slow deviations usually appe ar on theground-monitoring displays as deviatedflight conditions, when compared to nominal tra jec tor y conditions. To provide thecre w with the required abort decision, theground controller must know the extent towhich the trajectory can be allowed todeviate before a subsequent abort pro cedu re would violate a crew o r spacecraf tconstraint . Such constraints as the1 00-second crew-procedure t ime al lotment o r the 16g human-endurance limitfor en t ry dece le ra t ion force s a r e used toestabl ish these s low-deviat ion t ra jecto rymonitoring limits (fig. 4).

P a r am e t e r s f o r d isp l ays such asthose shown in figure 4 we re developedaf te r many s tud ies . The par ame ter s selected we re general ly the mo st s ignif icantvari able s relat ing the equations of motion

Abort l imi t

Pi tch rate. P i t c h e r r o r

- 2 1 I 1 1 I I , . . I I -178 80 82 84 86 88 90 92 94 96

Flight time. se t

Figure 3 . - Typical tim e histor y oflaunch-vehicle dynamic variables following an actuator failure.

50 I

A n \s-ICiS-11 Maximum entry load fac tor

Time 0 1 free

300 000 I ta l t i tude

,ertion

staging

'10 1 - 10 4 1'2 ;b 20 24 28 x IO 3

Ine r t ia l ve locity. V i . Itlsec

Figure 4. - Ground-control t raj ecto ryl imi t s .

to known constraints. This development task has not been s im pl e because t he motiondescribed by orbi tal and f l ight mechanics re qu ir es mor e than the two dimensions avai lable for displays. In som e instances, two displays wer e required fo r completeness(for example, velocity compared with flight-path angle (or altitude rate) and altitudecompared with range) . The de si re to have a reasonably common terminology in thedecisionmaking displays fo r ea se of understanding was another fa ctor used to sele ct thepa ra me te rs and constra ints on the display shown in f igure 4 as well as on most oth erd isp lays . Constraints such as the LES per formance l imi t s we re s tud ied and d iscardedon the basis that the validi ty in thei r us e as cons t ra in t s w a s ambiguous.

7


17/37

Retu rn to Ear th

Fo r an Apollo spa cec raft launch, the abo rt mode s that have evolved fr om th e contingency planning pr oce ss can be brief ly de scri bed as follows. In a mode I abor t , theLES ower per fo rms the t ime-c ri t ical esc ape fr om an impending launch-vehicle explosion durin g atmosp heric flight. A mode 11 abort may occur af ter the vehicle leaves the

a tmosphere , where the re is little chance of aerody namic loads that can lead to launch-vehic le break up and explosion. A mode II abort c ons ist s simply of a sep ara tio n of thespacec raft and the launch vehicle, followed by sp ace craf t orientation to entry attitudeand a subsequent landing in the Atlantic Ocean.

As the f light pro gr es ses and the inert ial veloci ty V . increa ses, controll ing the1

ti me incre men t between launch-vehicle cut-off and a re t rograde maneuver was de t e rmined to be the most effective mean s available of controlling the spacec raft landingpoint. Thi s launch abo rt technique is r e f e r r e d t o as a mode I11 abort .

If a contingency should ar is e in the last 2 minute s of the launch phase , when t r aj ectory conditions are still suborbital , the space craft propulsion sys tem can providethe t ransi t ion to an orbi ta l t rajec tory. Such a procedure, called a mode IV abort , isnot strictly an abort in that the procedure does not produce an immediate return of thecr ew to earth . Becaus e initiation of e ith er of the other two abo rt modes that can occurin this f l ight regi me (mode I1 o r 111) could result in a wide range of undesirable landinglocations, a pr im ar y advantage of the mode IV p rocedu re is that landing-site selectionopportuni t ies a r e provided. That is , af te r an orb i ta l s ta te is attained, the spacec raftmay tra vel through par t of a revolution until a desired landing area is approached andthe cr ew may then perf orm the usual entry maneuver. If sufficient propellant remainsafte r the t ransi t ion maneuver and if the original contingency do es not req uir e flightterminat ion, som e mission object ives maysti ll be accomplished. The four launchabort mode s a r e shown in f igure 5.

The la st few seconds of a launch a r eparticularly critical because the ground-control personnel must advise the crew atlaunch-vehicle cut-off whether the tr aj ectory is sat isfactory (a st at us of go) o rwhether a mode I11 o r mo de IV abort procedure is required (a st at us of no-go).The import ant go/no-go decis ion is faci litate d by the u se of a ground-monitoringdisplay such as that shown in figure 6.

After an EPO is achieved, emergency entry req uir emen ts may be sent tothe spacecraft by the ground-control personnel. Also, in thi s situation, the cre wmay use previously pre par ed deorbi tda ta cards .

(24 000 f t l s e c ~

d' Switchover criteria

Fi gu re 5. - Apollo spacecraft launchabort modes.

8


18/37

TRANSLUNAR INJECTION

Cons t rai nt s

Tran slun ar inject ion is the powered-flight maneuver made by the S-IVB thatsend s the command and ser vic e moduleand lunar module into a t rans lunar t ra jectory. This event is noted as a part of themission pro file shown in figu re 7. Be

21.0 2i.8 2i.6 2i.4 2412 2510 28 2d.6 27 . ;~ lo 3 cause TL1 begins in an at anInertial velocity, V i. ftlsec of 100 nautical mil es, the previou sly dis -

Figure 6. - Ground-monitoring displayof near-inse rtion abort decision s forApollo missions.

Figure 7. - Apollo mission profile.

cuss ed launch-vehicle-type contingenciesthat would have required an abort duringlaunch may now, becaus e of m or e flexibleconstraints, be reduced to alternate-missionsi tuat ions. That is , off-nominal S-IVB performance is l ikely to require thrust termination but is unlikely to requ ire an immediatere tu rn to ear th . Therefore , the CSM has i t sentire propulsion capability intact and available for other uses .

Because the spacecraft is in a pass ivero le dur ing TLI, po tent ia l p roblems a r e le sslikely to occur in the spa cecraft than in theS-IVB. However, much pr em is si on effortis expended in making the warning andesc ape techniques completely inclusive,even for unlikely situations. Although allconstra ints in table I a r e generally applicable, certain constraining fact ors for possible contingencies apply only during theTLI mission phase. These cons t ra in ts a r eshown in table III.

9


19/37

- -

TABLE III. - PRIMARY CONSTRAINTS DURING TRANSLUNAR INSERTION

Vehicle Constraint

Hardware~ ~~

Launch vehi cle Guidance and control, ine rti al refe ren ces , and gimbal lock

Spacecraft

Computer

Communications

Guidance and control, ine rti al ref ere nce s, gimbal lock, manualsteer ing, and displays

Computer - disp lays and prog ram s

Optics - window visibility

Propel lants

Separation and sequencing

Operational

Al te rna te miss ion s

Minimum orb ital altitude

Entry corr idor

Recovery a reas. -

Maneuver Monitoring

Within the context describe d, the pri ma ry objective (aft er cre w safety) when amalfunction develops is to pe rf orm an al tern ate mission. Therefore, a l lowable deviated

fl ight condit ions must be determined in advance to ens ure that the desi red al ternate-missi on capability will exist. Consideration must al so be given to the provision ofreasonable ini t ia l condit ions f o r perfor mance of an abort maneuver. These requir eme nt s have been fulfilled by the development of a crew-monitoring procedure that includes appropriate S-IVB shutdown limits.

The c re w mu st be able to monit or and eval uate TLL without ground sup port , beca us e the S-IVB sec ond bur n ca n occ ur out of th e Manned Space Flight Network (MSFN)

10


20/37

t racking range. General ly, TLI occ ursnear Aus t ra l ia in the We s t Pacific Ocean.A schematic of the basic crew-maneuvermonitoring technique (fig. 8) shows that anabor t can be per formed f or a t t i tude-ra teproblems, fo r at t itude-deviat ion problems,and for spacecraft sys tem problems. Because S-IVB fa i lu res normal ly resu l t in anal ternate mission, only a cr i t i ca l spacecraf t sys te m problem is l ikely to requirean abor t . Al l fou r of th es e monitoringru les , as well as numer ica l l imi t s , wereestabl ished af te r s tudies showed that theal te rnate object ives and possibly c rewsafety could not be ensu red without them.

The following ite ms can be notedabout the TLI monitoring technique.

1. The T LI ignition w i l l be inhibitedif the launch-vehicle attitude before ianition is mor e than 10" fro m nominal, asdetermined by horizon reference.

2. The S-IVB engine wil l be cut offby the crew for pi tch o r ya w ra te s of10 deg/sec o r g rea t e r.

Burn attitudecheck (h0riZOllf

Ignition

I

Alternate missionor abort

A

i Display and I1 keyboardcheck IIL---T----J

r-lackup cut-offI I

Figure 8. - Basic crew-maneuvermonitoring technique.

3 . The S-IVB engine wi l l be cut off by the c re w with the ab or t handle fo r attit udedevi atio ns of 45" o r m or e f ro m the nomina l a t ti tude.

4. A backup to the S-IVB guidance cut-off sign al w i l l be perfor med by the crew ifthe S-IVB h a s not sh ut down at the end of the predi cted burn t im e plus a 2a dispersionof 6.0 seconds and i f the nominal inertial-velocity display by the spacecraft computerhas been achieved.

The crew is provided with preflight t abl es of at titud e and compute r-dis playpar ame ter va lues a t d i s c re te t imes dur ing the TLI burn . These tab les provide fo r bo thnominal monitoring and crew manual steering. If a launch-vehicle inertial-platformfa i l u r e o ccu r s be fo r e TLI o r i f a t t i tude-re fe rence s igna ls a r e los t dur ing TLI , the c rewmay ass um e manual control of the burn with the hand cont rolle r. Typical pro ble ms thatthe monitoring concept can prevent o r avert a r e shown in f igure 9.

11


21/37

v

1. Nominal2. Loss of iner t ia l a t t i tude"3 . Lossof at t i tude-er ror s ignal '

4. Loss of at t i tude-ra te s ignal a5. Loss of attitude-command signal '6. Pla t form dr i f t0.13 deglsec

* 10 se c after TL I igni t ion

h

0 40 80 120 160 200 240 280Ground-elapsed time from ignition, sec

Figure 9. - Malfunctions during TLI.

Return to Earth

After the investigation of possiblespacecraf t sy s tem fa i lu res tha t might occurdurin g TLI, no known single-point fai lur eso r constraint f act ors were found that mightreq ui re the c rew to abor t the miss ion dur ingTLI. However, t raje ctory analyses havebeen ma de that indicate t he feasibility ofperforming a simple abo rt maneuver byusing the horizon of the e ar th as a re fe rence .The decision was reached to develop thismethod of aborting duri ng TLI to protectthe c rew in the event future system-fai lureanalyses dictated a need for a TLI abortcapability.

A constant attitude of 5 " w a s selectedas optimum by trading off re tu rn tim es fo rab or ts at e i ther the s t ar t or the end of theTLI maneuver, that is , providing sufficient

320 t ime f r om ab o r t t o en t ry f o r ab o r t s a t t hebeginning of TLI and dec re as in g the re tu rnt ime as much as poss ib le for abor t s at theend of TL I. A del ay t ime of 10 minutes w a sselec ted fo r the following two re aso ns.

1. The 10-minute delay provided the cr ew with sufficient time to orie nt the sp acecraft fo r the maneuver.

ana lyses re qui res the c rew to c a r r y a chart 2 206 6,4s im i l a r t o f i gu re 10 on board the spacecraft gfo r u se in conjunction with onboard disp lays $ 204- 5. 6

ance. Although thi s abo rt proce dure was 2 198 2 3.2 ,' /./

/'

/ -10- L

eliminated as unnecessary on the late r *P 196 '2.4 / / / /0+' l 2Apollo mi ssion s, it was d iscussed her e to - 0indicate the deg re e to which abor t plans E! 194- 1.6 ~ 4 / : - ~ - Earth-fixed attitude aline ment

wer e developed. ? 192 .gc

5 190- 0 1 . 1 1 1

12


22/37

This technique is part ial ly faci l i ta ted by the fact that t racking ships a r e placed suchthat the acquisition of s ign al a l w a y s occ urs within 10 to 20 minutes after TLI.

Traje ctory analy ses indicated that , for any TL I posi t ion (geographic) , communications could be established soon er by completing the TLI maneuver than by shutting downthe S-IVB and coasting in an EPO. Thus, a syste matic check of the spacecraft sy st em scan be accomplished ea rl ie r by continuing the TLI maneuver. This procedure was

adopted at t h e r i s k of increas ing the re turn- to-ear th t ime in or de r to maximize miss ionsuc ces s probabi li t i es if the fai lur e indication should be negative.

After the decision was made concerning the bes t possi ble co ur se of action following fai lur e indications during TLI, the abor t planning then consisted of determining,through traje ctory analysis and considerat ion of the operat ional constraints , the ear l ie stprac t ical t im e that the mission could be aborted should the fai lure indicat ion be confirme d. Among the fac tor s considered wer e the t ime required to perfo rm the malfunction check, S-IVB sepa ratio n, orbit determination, abort-maneuve r computations, andinert ial-platform al inement. The resul tant abort t im e is 90 minutes following nomina lTLI cut-off. Implementation of this abort technique re qu ir es that ground contr oll erscompute, duri ng the EPO, a retu rn-to -eart h solution timed fo r 90 minutes followingTLI. This solut ion res ul t s in a return-to-earth t ime of le s s than 18 hou rs and a r e t u r nto one of five contingency landing sit es that a r e geographic l ines located nea r convenient recovery-stag ing a r e a s throughout the world. The 18-hour ret urn w a s consistentwith the ava ilab le t ime l imi t es tab li shed f or an assumed pres sur e-s u i t compr essormalfunction. The prob able re cov ery a r e a would be the Atlantic Ocean beca use of theoperat ional constraint that T LI should occur ov er the Pacif ic Ocean.

T R A N S L U N A R C O A S T

Cons traint s

The abort planning for TLC ( if a nominal TLI and a nominal transposition andCSM docking with the LM a r e assumed ) co nsi sts p ri ma ri ly of det ermi ning which of th eavailable propulsion s ys te ms should be used fo r the abort. With the LM and CSMdocked, the total vehicle is capable of firi ng two independent main propulsion sy st em s,with each vehicle having a pr im ar y guidance, navigation, and control syst em (PGNCS).In addit ion, the se rv ic e module react ion control s yst em (RCS) is avai lable for maneuvering, and two independent communications sy st em s a r e available. Also, t he commandmodule (CM) computer contains an abo rt pro gr am (P37) capable of computing targetingp a r a m e t e r s f o r t h e CM PGNCS. With tra nsf orm ati ons , the CM abor t pr og ra m can beused in the LM. The fol lowing constraints a r e considered in planning abort t r aje cto rie s .

1. The maximum total f l ight t ime considered is within the spacecraft s yst emlifetime.

2. The maxim um entry velocity considered is limited by the spacecraft entryheating constrain ts .

3. The abort maneuver magnitude is within the s er vi ce propulsion sys tem (SPS)capability.

13


23/37

4. The abor t maneuver is targeted to conditions within which the crew and spacecraf t are expected to ope rat e at an optimum (that is , a t des ign limits).

Th re e methods of effecting a r e t u r n t o e a r t h f r o m TLC involve maneuvers that a r eperformed before the spac ecraf t rea che s pericynthion and that provide a d i r ec t r e t u rnto ear th (d i rec t - re turn abor t s ) ; maneuvers tha t a r e per formed before the spacecraf trea che s pericynthion and that provide a r e t u r n t o e a r t h a f t e r t he spacec r a f t p a s se sbehind the moon (circumlunar aborts) ; and maneuvers that are per formed after thespac ecraf t rea che s pericynthion and that provide a re turn to ea r th a t tha t t ime (pos tpericynthion ab orts) .

Because the flexibility afforded by multiple systems and maneuver choices doesnot allow for a sim ple contingency plan, othe r guidelines wer e developed. The following guidelines wer e among thos e developed.

1 . Circumlunar abor t s a re not to be perfor med outside the lunar sphe re ofinfluence.

2. The LM descent propulsion sys tem (DPS) will not be used instead of the SPSfo r a di rec t - re turn abor t . (Also, disp osin g of the LM befo re the bur n would be a real-t ime dec is ion .)

Return to Earth

With continuous ground trackin g available d uring TL C, the contingency warningdecision can be made by the ground-control personnel. With key fac to rs and potentialproblems a l ready identi f ied , abor t p lans fo r TLC ar e reduced essen t ia lly to rea l - t imedec is ions. These dec is ions a r e facil itat ed by u se of th e type of d ata displa y shown inf igure 11. Invest igat ions helped reduce candidate displays to the s imple nece ssi t ie sshown in figure 11. The TLC var iab les for SPS abo rts to the pr im ar y contingency landing si te located in the middle of the Pacif ic Ocean a r e also shown in f igure 11. Fromtrade-offs , poss ib le abor t so lu t ions a r e se lec ted for a rb i t ra ry t ime s (usua lly co incidingwith crew awake times ) throughout TLC. These solut ions a r e also analyzed before themis sio n to provide the cre w with navigation sighting schedule s that would as si st them

(in the event an abort is necessary) in performing the required navigat ion maneuvers

11 lo3Ground-elapsed time at landing. hr to allow a safe re tur n within the entry c or

l o t ;5 /49

/73 r idor a t the des i red landing s i te .

The abo rt maneuver information provided to the c rew in re al t ime for guidance-computer target ing co nsist s of a bort t ime,A V , longitude of the earth landing site, andl5 ntry t ime. This information is used onlyin the even t of a total lo ss of gro und-to-air

2 L u n a r s p h e re communicat ions.o f i n f l u e n c eI

0 1 ... j . , . I 1 > ' I I 14 8 12 16 20 24 28 32 36 40 44 48 52 56 60 64 68 72 Th e outbound porti on of the abort plan

Ground-elapsed l im e of abort. hr (from launch to lun ar ar r ival ) can be con-

Figure 11. - Apollo 11 trade-off display. c i se ly represen ted as shown in figure 12 .

14


24/37

Launch 1 EP O

Monitor ing groundl imi ts trajectory

Contingency-Mode Irocedure

option

'Early S - 1 p B staging

Of part icular interest a r e the relat ionshipsof th e various ab ort mode s to each other

~ and to nominal miss ion even ts. As an ex4ttitudek v i a t i o n = ample, maneuver monitoring occu rs nom15 deg inally during TLI, and an alternat e mission\ b u r n t i me =

is the pref er re d procedure r a the r than anabort at TLI plus 90 minutes , depending onthe se ver ity of the contingency.

T L I + 90 min Direct=el A p o l l o13 T r a n s l u n a r10-min S P S C o a st A b o r tIonboard)

It is approp riate to mention in thissection the Apollo 13 contingency that incapacitated the electrical power and maneu-

Figure 12 . - Operat ional abort plan fr om veri ng capability of the CSM during thelaunch to TLC. tra ns lun ar phase of that mis sion . As a

result, the TLC postpericynthion abort previously d escribe d w a s required to effect the

safe ret ur n of the flight crew to earth. The abort maneuver w a s initiated 2 hours af terpericynthion passage and w a s performed with the LM DPS as previously established bythe ab ort plan.

In addition to providing a previously prepared and reh ears ed return-to-earthtechnique, othe r asp ec ts of t he preflight ab ort planning wer e used during this emergencysituation. Fo r example, backup (to the CSM) life-sup port pro ce du re s and limitatio nsusin g the LM had been identified. Also, the u s e of window views of the cel est ial s p he reto obtain prescr ibed spacecraft at t i tudes for performing abort a s well a s midcoursecorrect ion (MCC) maneuvers ha d been developed, and th is concept w a s used because theguidance and navigation (G&N) sys t em w a s unavailable becaus e of e lec tri cal powerlimitations.

L U N A R - O R B I T I N S E RT I O N A N D L UN A R O R B I T

C o n s t r a i n t s

The LO1 bu rn , which is performe d by the SPS, t r an sf er s the spacecraft from TLCto an LPO. Pr em at ur e termination of the LO1 burn places the vehicle on an off-nominaltrajector y fr om which ei ther an abort or an alte rna te mission may resu lt. In the eventan SPS fai lure occ urs, the LM is required to re tu rn the CSM to ear th.

The development of feasible abort pro cedures fo r the LO1 mission phase m usttake into account many h ardw are and operational con straints. The majo r constraintsthat w er e included in the definition of an operatio nal LO1 abo rt philosophy a r e sum marized in table IV.

15


25/37

TABLE Tv. - PRIMARY CONSTRAINTS DURING LUNAR-ORBIT INSERTION

.

I Constraint category ConstraintI

Hardw are Engine duty cycle s

Optics - window visibility

CSM and LM autopilots

Iner t ia l re fe rence f rames

Propel lants avai lable

Propuls ion pressur iza t ion sys tem

Computer displays

No onboard abort proc es so r within lunar s ph er e of influence

Real- t ime computer complex (RTCC) pro ces sor (s ingleimpulse only)

Minimum LM activation times

MSFN tracking and RTCC solution requirements

Maximum total miss ion t im e - consumables.. - ~ . - . - . - - . .- . .. -~

Maneuver Monitoring

Because LO1 always occu rs behind the moon, the c rew mus t be able to evaluatethe pro gr es s of th e maneuver without ground supp ort. The recomme nded LO1 crew-monitoring technique is shown in figu re 8.

The preigni t ion space craft a t t i tude check, which is i l lus t rated in f igure 13 , ismade m or e d i ff icu lt by the presen ce of the LM. However, the horizon and sev era ls t a r s should be vis ible fro m the CM rendezvous window, and these ref ere nce s may be

used as a backup to the optics fo r the orientation check befor e ignition.craf t a t t i tude is not within * 5 O of nominal, the LO1 should be no-go, bec aus e lar ge rIf

the space

a tt i tude- ref e rence e r r o r s could resu l t in more se r iou s problems dur ing th i s c r i t i ca lmaneuver.

Although mainten ance of cr ew safet y is always the pr im ar y objective of th e monitoring proce dures , another important objective is the assurance that adequate abortcapability is provided and that the capability is compatible with poss ible r es ul ts of the

16


26/37

----

monitoring procedu res. This objective isaccom plished f o r LO1 by definition of soundproc edu res for the four types of p robl emspossible during LOI. The four problemtypes a r e guidance and control, non-SPSsy st em s (other than guidance and control),the SPS, and inadvertent SPS shutdowns.

A solution to the guidance and controlproblem is for the c rew to assu me manualcon tro l of th e LO1 man euv er, which is normally controlled by the PGNCS, and to complete the LO1 at the original ignition attitude.One of the mo st dangerous possibi l i t iesassociated with guidance and control probl e m s is spacecraft inert ial measurement

-50 -40 -30 -20 -10 0 10 20 30 40 50 unit (IMU) dri ft dur ing LOI. The cr ew can-X deq not detect a sm al l drift until an attitude

deviation builds up and appears on theFigure 1 3 . - Visual LO1 attitude check. secondary inert ial-at t i tude referenc e sys

t em. Because the drift could occur in eitherthe secondary reference s yste m o r the I M U ,

the crew is unable to distinguish th e erro neo us sy st em without using the backup attitudee r ro r need le s (a third inert ial referenc e system). Detection of the e r r o r make s possibl e a manual tak eov er and compl etion of th e bur n s o that the spacecraft can enter anLPO. Because uncorrected IMU dri fts in pitch can produce impact tra jec to rie s, attitude li mit s for which a takeover should be initiated were developed (fig. 14). A s is thecase for TLI , the ra te l imi t f o r LO1 is 10 deg/sec because la rg er r at es a r e not withinnormal system operat ion. If this l imit is exceeded, a crew takeover is initiated, andmanual completi on of LO1 at ignition attitude is performed. Non-SPS problem s req uir e

0 - 0 -Crew takeover-0 2 - 2 - Pitch dr i f t

Pi tch misal inement.04 - 4 -

Positive pitch drift

Negative misalinemen t

.14 - 14

.16 - 16 1 I I 1 I I I I I-30 -20 -10 0 10 20 30 40 50 60

Alt i tudeof per i lune, h pc, n. m i .

Figure 14. - Effects of attitude -referenc efai lur es on pericynthion altitude.

completion of LOI, because it is advantageous to be in the planned lunar orbit r at he rthan in a possibly undesirable lunar orbitin the event an abort is required.

Manual SPS shutdown occurs only ifc r i t ica l SPS subsys tem problems a r i se tha twould severe ly re st r i ct the future perf ormance of the engine o r jeopardize the safetyof the crew. If an inadv ertent SPS shutdownocc urs and SPS l imits a r e not exceeded, therecommended procedure is to initiate animmediate SPS restart . If t he r e s t a r t is unsuccessful and an abort situation exists, theLM DPS is used for the abort maneuver.

In sum ma riz ing LO1 monitoring, animportant objective is completion of the LO1bur n. Only when bur n completion is not possibl e should the SPS burn be terminat ed. Fo rthi s situation, the LM DPS is t he p r ima rysou rce fo r the return-to-earth maneuver.

17


27/37

Return to Earth

Lunar orb i t s tha t res u l t f ro m prema tur e LO1 thrus t t e rmina tion range f ro m h ighl jenerget ic escape hyperbolas (which resul t f rom early SPS shutdown) to stable lunare l l ipses (which resu l t f r om la te SPS shutdown). Th es e luna r orb its fall in to th r eegenera l c lasses .

1. E scap e t r a j e c to r i e s ( c l a s s I)

2. Unstable o r impacting el l ipses ( cla ss 11)

3 . Stable or nonimpacting el l ipses (cl ass 111)

The th re e cla ss es of lunar orbi ts a r e shown in f ig ure 15. Esc ape t r a j e c to r i e s( c l a s s I) a r e orbi ts with e nergi es suff icient ly high to cause escape f rom the lunarsp he re of influence. Esc ape is som eti mes caused by ea rt h ( third body) perturbat ionsacting on a spacecraft in a highly elliptical orbit at o r ne ar apocynthion. Unstablee l l ipses (c lass 11) a r e orb it s with apocynthion altitudes high enough for the s pacec raftto undergo lar ge pertu rbati ons by the ea rth, but not high enough for the spacecraft toescape the lunar s ph er e of inf luence. Impacting el l ipses (c las s II) a r e t r a j e c to r i e s t ha t

~ Approach hyperbola

YLOI urn7,No ni mpact e l ip rew

verburn

Impact ellipse

Figure 15 . - Pr em at ur e LO1 shutdowntra jec tor ies .

Trajectory classesI Escape

II Unstable

I l lStable

I I l l .~ II

I I I I1 1 I 1 I

0 1 2 3 4 5 6LO 1burn time, mi n

I I ~

I

Figure 16. - Pr em at ur e LO1 shutdowntra jec tory c lasses .

18

are suff icient ly perturbed for the spacecraftto impact the moon during i ts f i r s t approachto pericynthio n. Stable nonimpacting ellips e s ( c l as s 111) resu l t if SPS shutdown occ ur sdur ing the final port ion of the LO1 burn.Stable el l ipses a r e defined as orbi t s withpericynthion al t i tudes gr ea te r than 40 nautical mi le s. The regio ns of the typicalluna r-miss ion LO1 burn that will produceeach t ra jec tory c l as s a r e il lus t ra ted inf igure 16.

Because the c l as s 111 t r a j e c to r i e s a r estable lunar orbi t s with a pericynthion altitude in ex ce ss of 40 nautical miles , e i theran al ternate mission or an abort may resul t .In the event an abor t situation exist s, thereturn-to-earth maneuver would be s imi la rto the no rma l TEI burn and would occ ur onthe far si de of the moon. In th is region,the abort maneuver (an LO1 mode I11 abort)con sis ts of a single burn that is general lywithin the AV cap abi lit y of th e LM DP Sengine.

The abort AV requi rements fo r atypical s ta ble lunar e l l ipse a r e shown infigure 17 . The abort AV is a function ofthe de lay t ime and des i re d t ran sea r th f ligh tt ime. The AV that is required rapidlyi n c r e a s e s as the ti me of t he abo rt ignition


28/37

310 1 10

2 -~ I l l l l i I l I I

0 .4 .8 1.2 1.6 2.0 2.4 2.8 3.2 3. 6D e l a ytime, h r

Figure 17. - Abort AV r eq u i r em en t sfo r a stab le el l ipse.

2800

-400D P S AV

c dpd bl IIt y2000

Delay l ime , hr

~1600,A--

is delayed past the LO1 burn shutdown. A sapocynthion is pas sed in the ell ipse, however, the mo re op timum c i rcumlunar abor tsolutions beco me available, and the AV inc r e a s e s t o a minimum value just beforepericynthion is reached.

During the LO1 burn , if an SPS shutdown occurs before a c l a s s I11 t ra jec tory isreached, an abort is nece ssary because astable lunar orbit has not been achieved.The single- impulse abort requirements forshutdowns e ar ly in the LO1 burn a r e shownin f igure 18 f o r a typical mission. Thistype of ab or t man euv er (an LO1 mode Iabort) is a function of de lay ti me and des i red t ran sear th f l igh t t ime ( o r des i redlanding- site longitude). The AV that isrequired in cr ea se s proport ional ly with theti me delay before execution of t he abo rtmaneuver . By superimpos ing on figu re 18the LM DPS AV that is avai lable, i t isshown that the LO1 mode I abort capabilityd e c r e a s e s as the abort maneuver is delayed.

The development of oper atio nal abor ttechniques fo r the ea rly Apollo lunar m iss ions w a s greatly influenced by the cons t ra i n t s summar iz ed in tab le IV. The AVcapability of the LM w a s based on the useof th e LM DPS engine only, bec aus e the us e

of the LM asc ent propulsio n s ys te m (APS)engine res ul t s in control problems. Inaddition, t he t i m e con str ain ts of LM acti va

. ~L ion for a primary guidance and navigationO L20 40 60 80 100 120 140 sy st em (PGNS) DPS burn , combined with

LO1 burn time. se c the t ime cons t ra in t s fo r g round t ra jec tory -solution preparat ion , resu l t in a minimum

Figure 1 8 . - Minimum AV requi rements delay time of 2 hours fo r an abor t us ing thef o r LO1 mode I abor t s . LM PGNS DPS. The del ay ti me h a s no

effe ct on LO1 mode 111 aborts because mult iple abort opportuni t ies exist . (That is ,

the re a r e mul t ip le revolu t ions in lunar orb i t . ) However, a s can be seen f rom f igure 1 8 ,

the LO1 mode I abort capability is direct ly affected.

As indicated in the previous discussion, the l imitations caused by the real- t imecomputer complex (RTCC) single-impulse abort solution capability (table IV) bound aregion of th e LO1 burn f o r which an abo rt cap ability doe s not exist. In the abse nce ofsingle-impu lse abor t solutions, an operationa l multi-impulse abor t technique that usedthe existing ground computation pr og ra ms had to be developed. The technique considered w a s a two-impulse procedure that would use the existing real-time single-impulse

19

3


29/37

abort p ro ce ss or and thus avoid the development of a spec ia l ized f i r s t - impuls e pr oces so r. Moreover, the first impulse could be specified in a convenient manner by usingp re mi s s io n d a ta .

Use of the DPS fo r mult i - impulse operat ion necess i tate s th e considerat ion of addit iona l hardware res t r ic t ions . For example, multib urn operation of the DPS engine in

troduces constrain ts on the engine duty cycles , specif ical ly with reg ar d to burn durat ionand coast periods between the burns. A maximum coas t- t ime l imit between DPS burnsexists because of the pr es su re bui ldup in the LM supercri t ical- hel ium stor age tanks.In addition, the duration of th e f i r s t burn mu st not exceed engine rest ric tio ns and thusinhibit the DPS re st a r t capabili ty. Crew act ivi t ies and re st cycl es also wer e considered.

The f i rs t maneuver of the two-impulse procedur e must accomplish s ev er al goals .The maneuver must be small enough to allow a r e s t a r t of th e DPS and must leave enoughpropel lant to complete the return-to -earth maneuve r. The ini t ia l maneuver must alsor e s u l t i n a safe intermediate lunar orbi t that al lows sat isfactory delay t ime to thesecond impulse. Following an extensive anal ysis , th e first maneuver w a s planned as avar iab le AV burn (depending on the LO1 burn t ime) d irec ted down the rad ius vect or.

Selection of the AV value to be used f o r the cor rec t ive maneuver is based on atrade-off between total fuel expenditure and inte rmed iate delay time . Fo r LO1 burncut-offs e ar ly in the c la ss I I t r a j ec to ry reg ion , a la rg e value of A V is neces sa r y t oredu ce the apocynthion altitude (and thereby ea rt h pert urba tion s) to the extent that anadequate pericynthion for a s tab le e l lipse (c las s 111) resu l t s . Converse ly, fo r l a te cu toffs in the cl ass I1 t raje ctor y region, where the preabort e l l ipse approaches a c l a s s I11stable el l ipse, a correspondingly s ma ll corr ect iv e maneuver is requi red .

The total A V of the two maneuv ers in cr ea se s almost l inearly with the magnitudeof the fi r s t burn AV1, and a trade-off mus t be made between propella nt co sts and ti me

between DPS burns. Incr ease s in delay t im e before ini tia t ion of the f i rs t maneuver r educe the total AV req uir eme nts at the cost of incre ased delay t ime before the secondmaneuver and general ly lower pericynthion al t i tudes caused by e ar th perturbat ions fo rthe 1GiA:(2X'period orbits.

In view of t hes e consid erations, the final technique f o r determina tion of the c orrect ive maneuver is to define the minimum allowable AV

1nece ssar y to ob ta in a t r a

jec tory jus t ins ide the c l ass III t raje ctor y region. This value of A V1

will provide a

pericyn thion of app roxi mate ly 60 nautical mi le s and a t im e between b urns that is withinthe hel ium pres suriz at io n l imits . The value of AV, will thus decr eas e l inear ly to

I

0 ft/ se c at the st a r t of th e LO1 mode III region.

The f inal two-impulse abort proced ure (a mode I1 abort) f or shutdowns in thec l a s s I1 t ra jec tory region is as follows. The cor rec t ive maneuver is dire cted down theradi us vector. (The RTCC target ing is determined f rom pr emi s s ion da t a . ) T he co rrec t ive maneuver is pe r f o rmed as soon as poss ib le a f te r LO1 SP S shutdown (nominallya t 2 hours for LM activation and ground-based tracking). The magnitude of the c or re ct ive maneuver d ec re as es l inea rly with LO1 burn t i me for simplifica tion of rea l-t imerequi rements .

20


30/37

Because the ab ort capability duringLO1 is a function of earth-moon geometry,LO1 geo metr y, and so forth, the abort requi rements must be de termined fo r eachpart icula r mission. The total abort A Vfo r minimum-fuel re tu rn s that would have

been re quired following an SPS fa i lureduring the Apollo 1 1 LO1 burn is summar ized in f igure 19. The value of A V l is

shown in figure 20. The pericynthion altitude and the time between bu rns fo r theLO1 mode I1 intermediate el l ipse a reshown in figur e 21.

During the lunar-orbit phase, areturn-to-earth maneuver (a mode I11 abort)sim ila r to the nominal TEI burn can beinitiated on each revolution.

loo0rRO Oh

101 To t l e r q o n0 11 4 0 I 5 0 2 00 2 10 2 70 ? 30 ? 40 2 50 3 00

101 b u r n l i m e , Ig. m in sc c

- - Mode I S P S AvOnly )103 - - -- Mode JJ lCSM

- Modem

0:OO 0 : U O 1:20 2 : 0 0 2:40 3:20 4:OO 4:UO 5:20 6:OO 6:40101 burn time, tg, min:sec

Figure 19. - Su mm ar y of mini mum -fuel abo rt capability as a functionof LO1 bu rn t im e.

I 4 0 1 50 7 00 7 102

202

30 7 402

503

00101 l i l j r n l t i i r I l, rm n wc

Figure 20. - The LO1 mode I1 f i rs t-burn Figure 2 1 . - Time between bur ns and theA V as a functio n of LO1 bu rn ti me . pericynthion altitude following a nomi

nal fi rs t burn fo r an LO1 mode I1 abor t .

L U N A R D ES C EN T A N D A S C E N T

If an abort decision should be made aft er CSM/LM separati on for the lun ar-landing phase, rendez vous of the two vehicle s mus t be effected before the r eturn-toeart h maneuver. If descent procedure s have als o been ini t iated, the descent maneuversa r e monitored to maintain rendezvous capability.

21


31/37

Descent-0rbit-Inse rtion-Maneuver andPowered-Descent-Maneuver Monitoring

The desc ent-o rbit- inser tion (DOI) mane uver is th e first of th e two des cen t maneuve rs and occurs on the fa r s i d e of the moon. The DOI, a shor t re t ro grade maneuver ofapproximately 75 ft/sec, is performed with the descent engine and efficiently reduces

the orbi t a l t i tude fro m approximately 60 nautical mil es to 50 000 feet for the powereddesce nt initiation (PDI) .

The PDI maneuver is ini tia ted at the peri lune (50 000-foot altitude) of the DO1orbit, which is targ eted approximately 260 nautical mil es up range of the landing site .The powered descent re qui res a continuous thrusting of the de scen t engine fo r 12 minutes 36 seconds. During thi s maneuver, the thrus t direct ion and magnitude a r e modulated as neces sary t o br ing the LM to a hovering condition over the desired landingsi t e . The pi tch att i tude profi le is designed to allow the c rew visua l ass es sm en t of thelunar s urfac e during the terminal ph ases of the maneuver (from high gate, which is atapproximately 7000 feet al t i tude).

The DO1 man euv er is moni tored exc lusively by the c rew men becaus e of the position of the man euv er. An ove rbu rn of 12 ft /s ec (or 3 seconds) will c ause the LM to beon an impacting traje ctory before PDI. The maneuver is monitored fo r this impact ingcondition by compar ing the PGNS per for man ce with that of th e abo rt guidance sy st em(AGS) during the burn and by ra nge /ra te track ing with the rende zvous r ad ar (RR) immediately af t er the burn. If the maneuver is unsat isfactory, an immediate rendezvouswith the CSM is per for med using the AGS.

The powered descent is a complex maneuver that is demanding on both cr ew andsys tem s per formance . Therefore , as much monitoring a s possible is performed on theground to red uce c rew ac tiviti es and to use complex computing techniques not possibleon board the space craft . Obviously, t ime-c ri t ica l fai lu res and near- surfa ce operat ions

must be monitored on board by the c rew for immediate act ion.

The ground monitoring to detect G&N syste m pro blems includes direc t comparison of t eleme tered dat a fr om th e two guidance sys te ms on boar d the LM and data derived fro m the MSFN. The pri mar y guidance-monitoring sou rce is a comparison ofthe velocity component s, the AGS-minus-PGNS velo citie s, and the MSFN-minus-PGNSvelocities. In this manne r, an erroneo us system can be isolated by reference to thet h r ee s ou r ce s . L imi t s a r e estab lishe d on the MSFN-minus-PGNS velocity compa risonsuch that a degrading guidance sy ste m can be detected e ar ly enough that a maneuvercan be completed on the PGNS into a sa fe orbit (height of pe rig ee L 30 000 feet) withoutimpacting the sur face.

The per for man ce of th e total G&N sy st em is evaluated by monitoring the commanded thrus t magnitude (guidance-determined thr ust command (GTC)) . Nominally,the GTC decr eas es (approximately parabolical ly) fro m a n ini t ia l value near 160 percentto the throttleable level of 57 percent approximately 2 minutes before high gate. If theDPS engine prod uces off-nominal low thr ust , the GTC drops t o 5 7 percent la ter to guideto the desir ed posi tion and velocity. If the thrus t becomes excessively low, the tar get swill not be satisfied and the guidance solution fo r the GTC can dive rge. This dive rgenc ecan resul t in an unsafe t rajector y, one fro m which an abort cannot be sat isfactori ly

22


32/37

performe d because of excessive altitude ra te s. Hence, the GTC is monitored fordiverge nce, and an abor t on the PGNS is performed at the t im e of the detec tion ofdivergence.

The landing ra da r (LR)/PGNS interface is another area of pr im e conce rn duringthe descent . Without LR altitude updating, sy st em s and navigational e r r o r s a r e such

that the descen t cannot be safe ly completed. In fact , it is unsafe t o t r y to achieve highgate (the point where the crew can visually assess the approach) without altitu de updating. Thus, a mission ru le for real- t im e operat ion was establ ished that ca l ls foraborting the descent a t a PGNS-estimated altitude of 10 000 feet if altitude updating ha snot been established. In addition to the concern for the tim e the initial altitude updatingoccu r s , t he r e is al so co nce rn f o r the amount of altitude updating (that i s , the differencebetween PGNS and LR altitude determinations A H ) . If the LM is actually higher thanthe PGNS estima te, the LR will dete rmin e the discrepa ncy and update the PGNS. Theguidance then t r i es t o s teer down rapidly to achieve the ta rget s. A s a result of therapid changes, al t itude ra te s may incre ase to an unsafe level fo r abort ing the descent;that is , should an abor t be requ ired , the altitude ra te s could not be nulled by the ascen tengine in tim e to prevent s urf ace collision. The initial AH is monitored fo r accepta

bility before incorpo ration into the PGNS navigation. If the AH is unacceptable, it willnot be incorporated and an abort is required.

The trajectory is monitored fo r flight safety at all t imes. The prim e cri ter ionfo r flight safety is the ability to abor t the descent a t any time until the final decision tocommit to touchdown. Thus, flight dynamics li mits a r e placed on altitude and altitudera te to en sure that the vehicle maintains the capability to abort on the A P S until the lastpossible moment. The altitude and altitude ra te s a re monitored by both the crewmenand the ground; however, bec ause of communications dela ys, the ground only advi ses,based on projected trends, and the crewmen a r e responsible or protect ing againstflight-safety violations.

Lunar Module Abortt o Rendezvous and CSM Rescue

Fr om the beginning of the development of the proc ed ure s fo r the LM abo rt torend ezv ous and the CSM re sc ue of th e LM, the pr im ary emphas is w a s placed on thepreviously discussed powered-descent maneuver and immediate-postlanding phase,which a r e the most probable phases f or an LM abor t . A considerable amount of planning was als o done fo r fai lure s associated with DOI, fo r ca se s of no P D I , and fo rcorrect-p hasing LM asce nts before the nominal lift-off time .

The origi nal (beginning in 1964) LM-abort and CSM-rescue plans were extremelycomplex beca use of the li mited onboard capa bilities. For example, for an abort at any

time during the powered-descent maneuver,the LM was targeted for a constant inser

tion orbit. Therefore, several abort regions existed, and the rendezvous techniquesvaried for each. For abor ts ea rly in the burn, the fina l appro ach of t he LM to the CSMwas from above; for a l a t e r region, one and one-half revolut ions we re req uire d betweenthe coelliptic sequen ce initiation (CSI) and constant different ial height (CDH) rend ezvous maneuvers instead of the norm al one-half revolution; and, fo r late aborts , the LMapproac hed fr om below the CSM. Because of thi s complexity in the LM abor t plan, theresc ue plan was a lso complex. Th e CSM did not have th e CSI/CDH logic on board, andthe CM pilot had to depend eit her on the ground o r on the "m irro r-im age " technique

23


33/37

I I 1111II I

(that is , the method whereby the CSM app lies the LM-computed mane uver in the oppos i te di rec t ion). The pr i mar y re scue technique fo r bad-phas ing s i tua t ions was the six-impu lse technique, in which the CSM tr an sf er re d to a 20-nautical-mile cir cul ar orbi twith the f i rs t two maneu vers and then adjusted the phasing, becam e coel l ipt ic , and executed the termi nal phas e ( theoret ical ly with two impulses) with the las t four man euvers .

In early 1968, analyses w ere begun fo r the incorporation of sev era l powereddescent -abor t inser t ion orb i t s ( to vary as a function of abor t tim e regions). By lat e1968, t his work evolved into the variabl e-tar getin g concept, whereby the co rr ec t insert io n orbi t for an LM approach from a coe lli pti c dif fer ent ial height Ah of 15 nau tic almiles below the CSM could be targeted for all abor t t im es dur ing the first 10 minut es ofpowered descent . For an abort af ter 10 minutes, a constant 30-nautical-mile apoluneorbi t was targeted; however, an in-orbi t phasing maneuver (derived fro m onboard programs in conjunction with onboard charts) permitted the standard LM approach frombelow, although one additional revolution was req uir ed. Fo r the Apollo 12 missi on, asecond variable- target ing region ( through a two-revolution rendezvous) replaced thispost- 10-minute phasing region. The variable-targeting concept was originally thoughtto be unfeasible beca use of the softw are req uir eme nts involved; however, aft er a detai led analysis of the prec ise requi remen ts , the technique was deemed feasible andimplementation began in early 1969.

The variabl e targe ting led to much simplification and standardization of the ab ortand res cue plan. The sam e basic technique was now applicable for almost all cases inwhich the LM perfo rmed the rendezvous maneu vers. The re sc ue techniques, the ref ore ,were s tandard ized ; for example , fo r a CSM-active terminal phase, the CSM wouldalway s appr oach the LM fr om above. By th is tim e, the CSI/CDH lo gic had been placedon boar d the CSM, and an independent onboard ren dezv ous solutio n f o r the coellipticseque nce could be determin ed in the CSM. Thi s technique was a great improvement inth e CSM sup por t of an y ren dez vou s seque nce usi ng CSI/CDH logi c. Emphasis wasplaced on spacec raft independence because of t he uncertainty in the lunar potential andbecause nearly all re sc ue plans involved only one extern ally computed (ground) maneuve r. When cor rec t phasing existed ini t ia l ly, no external maneuver was required. Theaddition of very-high-fre quency rang ing capability to the CSM ensu red fur th er independence and confidence.

The cu rre nt ab ort and res cu e plan has been changed somewhat because of thechange to the nominal plan (landing one revolution lat er relat ive to the ma in CSM/LMsepa ratio n). However, the or de r of the occ urr enc e of the regio ns (one-revolution o rtwo-revolution rendezvous) is the only significant change; the bas ic techniques a r e thes a m e . The current plan is by no means a sim ple one; but, compared to the plan in useapproximately 1 year before the Apollo 11 f l ight , the curr ent plan is considerablys imple r and mor e s tandard .

PoweredAscentDuring ascent from the lunar su rface , the LM ascent s tag e is steered by the

P G N S to effect the planned rendezvo us with the orbiting CSM. As describ ed in thedesc ent monitoring section, the guidance sys tem is evaluated to determine proper operation. Detect ion of e r r o r s o r malfunctions could be cause for ascent-maneuver completio n on th e AGS. A nonnominal orb it aft er cut-off caus ed by an ea rl y engine shutdown

24


34/37

- - -

could re qui re the CSM to perform the rendezvous to res cue the LM if the LM maneuvering capability were lost. The limi ting altitud e on CSM re sc ue of t he LM is 30 000 feet .

TRANSEARTH INJECTION AN D TRANSEARTH COAST

Constraints

The TEI burn is intended to t rans fer the spacecraf t f rom i t s nomina l LPO to anearth-return moon-centered hyperbola. At this t ime, the space craft cons ists solely ofthe CSM combination. Ther efore , the sol e propulsion so ur ce is the SPS because theser v ic e module RCS is inca pable of per for min g a burn as l a rge as tha t requi red for theTEI maneuver. The constrain ts (both hardware and operat ional) that were consideredfo r TE I a bo r ts a re si mi la r to those fo r LO1 abo rts , with the exception of LM-relatedlimitations.

Maneuver Monitoring

Both TEI and LO1 occ ur behind the moon, and the monitoring proc edu res andtechniques for both maneuvers a r e basical ly the same. Preigni tion at t i tude checks fromthe CM windows a r e per formed the sam e as for LOI. The major difference is that theguidance and control and syste m problems during TEI req uir e a continuation of themaneuver; that i s , guidance and control probl ems res ult in crew takeover and burn co mpletion at the ignition attitude, and SPS o r spacecraf t sys tem problems a r e ignored unt i lthe important TE I maneuver h as been completed. A backup to the CM PGNCS TEI cutoff is performed by the crew at 2 seco nds pas t the nominal cut-off ti me, followingconfirmation tha t the de si re d cut-off veloc-ity has been achieved (as shown by theentry-monitoring-system AV counter). Ifinadvertent ter minat ion of the SPS oc cu rsduring TEI, the engine is r e s t a r t e d , i fpossible, within approximately 30 seconds,o r a ground solution wi l l be required fora l a te r abor t a t tempt . Manual tak eov er ofthe TEI maneuver occ urs i f , as explainedpreviously fo r LOI, the cr ew confirms byus e of two independent refe ren ce sy st em sa 10" deviat ion from the f ixed inert i alburn attitude o r i f the TE I rat e l imit of10 deg/sec is exceeded. The attitude-deviation limit was selected with the aid

of the data presented in fi gu re 2 2 . Themidcourse correct ion required fol lowing a TEI maneuver that ha s been madere la t ive to a dr i ft ing ine r t ia l - re fe rence

- M C C a i T E Ic u t - o r rpius15 hrlP371 based o n f a s treturn 163 hr l

M C Cai T I 1cut-ot1 plus70 h r P371 bared on5106 return 187 hr l

-0 20 4b 60 80 lb 1% 140 16

TE I bcrn time at crew takeover, sec

1

180

platform is als o shown in f igure 2 2 . Figure 2 2 . - The AV f o r m idcou r secor rec t ions as a function of TE I bur n

In su mm ar y, the philosophy of T EI t ime a t c rew takeover f or var iousburn monitoring is that completion of the pi tch drif ts .

25


35/37

TEI burn is mandatory; that is , a manual shutdown is not t o be ini t ia ted fo r any CSMsys te m problem. If an e ar l y au tomat ic SPS shutdown occu rs , an immediate res ta r t isto be attempted. Only if immediate reigni t ion is not possible will an RTCC ab ortsolution be required.

Retu r n to Earth

The orb i t s tha t result f rom prema ture TEI th rus t t e rmina t ions a re s imi l a r t o t h eorb i t s tha t resu l t f r om LO1 underburns; however, the or b i t s tha t resu l t f rom pre matureTEI th rus t t e rmina t ions occur in rev ers eo r d e r and as a function of TE I burn t ime.Ther efore , the TEI phas e of the missionha s ab o r t cha r ac t e r i s t i c s s i m i l a r t o those 101-1 TEI TECof t he LO1 ph as e, with the additi on of t he

R a t e = 10 deg Rate i 10 degfollowing two fac ts pe culia r to the TEI sec sec

Attitude Attitudephase. Monitor ing deviation = deviationl imi ts

10" 10"1. The increased abor t AV capa- A b u r n A b u r ntime time

bility e xi st s bec aus e of th e us e of the SPS t1 0 se c + 2 secand the lack of the LM weight. Alternatf SP S

IRTCCI_ _ ~ ~

2. T he re is no backup propulsion M a r u a l Manualsys tem f o r a TEI abor t . h t i n g e n c y

completion completion

i rocedure

During TEC, abort maneuvers are ip t ion

initiated only if a much-faster-thannominal e ar th re turn is required o r i f achange in landing position is necessary. F; SPSIP37

A concise fo rm of the abo rt plan fro mlunar a r r iv a l to TEC is presented in Figure 2 3 . - Operat ional abort plan fr omfigure 2 3 . lunar a r r iva l to TEC.

C O N C L U D I N GR E M A R K S

Some of the planning perfo rmed to e nsu re the safe re tu rn t o ea rth of an Apollocre w in the event of a contingency situ ation duri ng a lunar- landing mission has beendesc ribed . In par tic ula r, the development of cre w warning and escap e methods f o reach missi on phase ha s been emphasized. The development was accomplished pr i

marily by providing powered-flight monitoring procedures and abort modes that a recompatible with har dwa re and operational cons train ts. Becau se of the interaction ofthese constraints and because so many sy st em s and components can malfunction, muchpremission effort is requi red to ens ure that abort techniques a r e avai lable during all

26


36/37

mission phases . Although probabilities a r e low that an abo rt plan will be put into effect,the high confidence level in the Apollo Pr og ra m is , in par t, bec aus e of the fac t that asafe, s imple, w ell-rehea rsed ab ort plan exists .

Manned Spacecraft Center

National Aeronautics and Space AdministrationHouston, Texas, March 8, 1972924-22-20-00-72

NASA-Langley, 1972-

31 S-307 27


37/37

II I1 I11111111111 111 1111m I11111111

N A T I O N A L A E R O N A U T I C S A N D S P A C E A D M I S T R A T I O N

W A S H I N G T O N , D . C . 2 0 5 4 6P O S TA G E A N D F E E S PA I D__

N AT I O N A L A E R O N A U T I C S A N DO F F 1 C I A L B U S I N E S S FIRST CLASS MAIL S PA C E A D M I N I S T R AT I O N

P E N A LT Y F O R P R I VAT E U S E 13 00 U S M A I L..

NASA451

If Undel iverable (Sect ion 15 8POSTMASTER: Posral Manua l ) Do Not Return

"T he aeronai~ticaland space activities of th e Uni ted SFntes shall b econducted so as t o contribute . . . t o the expansion of h u m a n knowl-edge of pheiiotirena in th e atniosphere and space. Th e Administrcrt ionshnll provide for the widest prncticable and appropriate disseminationof infor mati on concerning i ts activit ies and th e W S 7 4 1 t S thereof."

-NATIONAL AERONAUTICS A N D SPACE ACT O F 1958

NASA SCIENTIFIC AND TECHNICAL PUBLICATIONS

TECHNICAL REPORTS: .Scientific andtechnical information considered important,complete, and a lasting contribution to existingknowledge.

TECHNICAL NOTES: 1nformation.less broadin scope but nevertheless of importance as acontribution to existing knowledge.

TECHNICAL MEMORANDUMS:Information receiving limited distributionbecause of preliminary data, security classification, or other reasons.

CONTRACTOR REPORTS: Scientific andtechnical information generated under a NASAcontract or grant and considered an importantcontribution to existing knowledge.

TECHNICAL TRANSLATIONS: Informationpublished in a foreign language consideredto merit NASA distribution in English.

SPECIAL PUBLICATIONS: Informationderived from or of value to NASA activities.Publications include conference proceedings,monographs, data compilations, handbooks,sourcebooks, and special bibliographies.

TECHNOLOGY UTILIZATIONPUBLICATIONS: Information on technologyused by NASA that may be of particularinterest in commercial and other non-aerospaceapplications. Publications include Tech Briefs,Technology Utilization Reports andTechnology Surveys.

Details on the availability of these publications may be obtained from:

S C I E N T I F I C A N D T E C H N I C A L I N F O R M AT I O N O F F I C E

apollo experience report abort planning

Documents