Apollo Experience Report Spacecraft Relative Motion and Recontact Analyses

8/8/2019 Apollo Experience Report Spacecraft Relative Motion and Recontact Analyses

1/22

N A S A T E C H N I C A L N O T E N A SA TN 0 - 7 9 2 0

0NOI

az+4m4z

9

CASE F I L ECOPY

APOLLO EXPERIENCE REPORT -SPACECRAFT RELATIVE MOTIONAND RECONTACT ANALYSESRobert E . McAdams, Chdrles J . Gott ,and A f m l a n d L. Wi'llidmsonLyndon B . Johnson Space CenterHouston, Texus 77058N A T I O N A L A E R O N A U T I C S A N D S PA CE A D M I N I S T R A T I O N W A S H I N G T O N , D. C. A P R I L 197


2/22

1. Report No.NASA TN D-7920

7. Authorb)Robert E. McAdams, Charl es J . Gott, andMarland L. Williamson9. Performing Organization Name and Address

Lyndon B. Johnson Space CenterHouston, Texas 77058

2. Government Accession No.

2. Sponsoring Agency Name and AddressNational Aeronau tics and Space AdministrationWashington, D. C. 20546

5. Supplementary Notes

7. Key Words (Suggested by Author(s))' Crew Safety Collision Avoidance* Vehicle Safety Accident Prevention' Flight Safety 'Collision Res earc hSpacecraft Separation' Separation P ara met ers

-. Recipient's Catalog No.

18. Distribution StatemectSTAR Subject Category:12 (Astronautics, General)

5. Report DateA p r i l 1975

9. Security Classif. (o f this report ) 20 . Security Classif. (of this page)Unclassified Unclassified

6. Performing Organization CodeJSC- 045218. Performing Organization Report No.JSC S-421

10. Work Unit No.942-22-20-00-72

21 . NO. of Pages 22 . Price'21 $3.25

11 . Contract or Grant No.

13. Type o f Report and Period CoveredTechnic al Note

14. Sponsoring Agency Code

8 . AbmactPotential collision o r accidental recontact problems between the spacecraft and other spacevehi cles (or components) existe d during most of the Apollo mission s. These prob lems wereidentified before each Apollo flight, and approp riate solutions for elimin ating o r minimizing thechance of collision were determined through relative motion analyses. This repo rt pre sents asum mar y of the identification, solution, and analysi s pro ces s for some of the more significantof these separation problems. In addition , the scope of responsi bili ty, the deve lopment of thestudy effort, and the comput er simulation development, which supported the anal ysis proces s,are described.


3/22

APOLLO EXPERIENCE REPORTEDITORIAL COMMITTEE

The mater ial submitted for the Apollo Experience Re por ts(a series of NASA Technical Notes) w a s reviewed and ap-proved by a NASA Editorial Review Board at the Lyndon B.Johnson Space Center consis ting of the following mem ber s:Scott H . Simpkinson (Ch air man ), Richard R. Baldwin,Ja me s R. Bates, William M. Bland, Jr . , Aleck C. Bond,Robert P. Burt, Chris C. Critzos, John M. Eggleston,E. M . Fields, Donald T. Gregory, Edward B. Hamblett, Jr. ,Kenneth F. Hecht, David N . Holman (Editorhecretary),and Car l R. Huss. The pri me reviewer for this reportwas Scott H. Simpkinson.


4/22

CONTENTS

Section Page

SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1INTRODVCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1DEVELOPMENT OF THE SEPARATION STUDY EFFORT . . . . . . . . . . . . 2COMPUTER SIMULATION DE VE LOPME NT . . . . . . . . . . . . . . . . . . . . 3SUMMARY O F SIGNIFICANT SEPARATION STUDIES FOR EACH

APOLLO MISSION. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5Apollol . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5Apollo2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6Apollo3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6Apollo4 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6Apollo5 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7Apollo6 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7Apollo7 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7Apollo8 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8Apollo9 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9Apollo10 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10A p o l l o l l . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 2Apollo12 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13Apollo13 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13Apollo14 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15Apollo15 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15Apollo16 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16Apollo17 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 7

CONCLUSIONS AND RECOMMENDATIONS . . . . . . . . . . . . . . . . . . . . 17iii


5/22

APOLLO EXPERIENCE REPORTS P A C EC R A F T R E L A T I V E M O T I O N AND RECONTACT ANALY SES

By R o b e r t E. M c A d a m s , C h a r l e s J . Gott, a n d M a r l a n d L. W i l l i a m s o nL y n d o n 6 . Johnson S p ac e C e n t e rS U M M A R Y

The use of a separa ble space vehicle during the Apollo P ro gra m req uired thatpre mis sio n planning include th e definition of separ ation pro ce du re s that would avoidaccidental recontact and ens ure maximum crew safety. Potential collision o r acciden-tal recontact pr obl ems between the spa cecra ft and oth er spa ce vehicles (or components)exist ed durin g mos t of the Apollo mission s. These pro ble ms we re identified beforeeach Apollo flight, and app rop ria te solut ions f o r eliminating or minimizing the chanceof collision we re deter mined through rela tiv e motion analy ses. More than 50 individ-ual separ ation pro ce du re s we re designed, analyzed, and documented fo r each of thefinal Apollo flights. The mo re significant of these se par ati on stu die s and the acciden-tal recontact pr oblem s associated with them are presented in this r epo rt; informationis given fo r each of t he Apollo m iss ion s.

After ini tia l planning fo r the Apollo Pr og ra m was underway, it became obviousthat a la rge, well -organized, and well-managed separa tio n study would be required tosupplement the tota l m issio n planning effort. The development of t hi s ef fo rt and thescope of the res pon sib ili tie s involved in meeting the requi rem ent are discussed in thisrep ort . In addition, the computer simulation development that made possible the solu-tions to the many separation pr oblem s encountered is summarized.

I N T R O D U C T I O NThe primary reason for a spacecraft-separation study w a s to identify potentialcollision or accidental recontact problems between the s pacec raft and oth er s pacevehicles (or components) and to dete rmine , by performing relative motion analys es,

appropriate solutions to eliminate or minimize the chance of coll ision. The ident ifica-tion of problem areas wa s accomplished by analyzing and verifying that the p rem iss ionnominal and contingency operati onal spacecraf t sepa ratio n seq uen ces and pro ced ure swere free of accid ental recon tact prob lem s or, i f no separa tion procedur e existed, bydefining proce dures that were free of acc identa l recontact pro ble ms. In addition, thesep ara tio n study eff ort included the verification that integrate d separ ation s yst emhardw are did not have inherent recontact problems, partic ularl y with respe ct to motionoccurring immediately after the separation event; for example, the mech anical ejection


6/22

of the lunar module ( L M ) fr om the spa cec raf t lunar module adapt er (SLA). The following objectives wer e included in the sepa rati on study eff or t: identification of safe spaccr af t operational lim its (e. g. , the maximum permi ssible spac ecraf t attitude ratesdur ing LM extr acti on fr om the SLA); conf irma tion of prev ious ly defined oper atio nalseparation sequenc es and pro ced ure s fo r subsequent miss ion s; validation of re comm endations for mission rules; investigation of th e se nsiti vitie s of separ ation dyna mics f o rvar iou s spacecraf t and spa ce vehicle configuratio ns; and g raph ic presentation andillustration of the rel ative motions of v ari ous s pace vehicle s (or components) in aseparation proc edur es handbook fo r each Apollo mission. The re su lt s of the sep arat iostudy efforts were requ ired and used in flight planning, hard ware design, cre w proce-du re s and checklists, sepai-atttioii t i~hf i ic juz~ ,ed-timc ~ ~ i s s i c aL?PPO,P~,a ~ destflighmission evaluations. A l l resul ts were officially documented and published as formalseparation procedures (internal notes) at the NASA Lyndon B. Johnson Space Cen te r(JSC) (for merl y the Manned S pace craf t Cent er (MSC)). A significant study effort wa snecessary to provide the required analy ses fo r the spa cecraft elemen ts and sys tem smodes of operation in the Apollo Pro gr am .

As an aid to the re ad er , wher e nec ess ary the orig inal units of mea sur e have beeconverted to the equivalent value in the Systgme Inte rnat iona l d'Unit6s (SI). The SIunits a re writ ten first, and the original units are written parenthetically there after .

DEVELOPMENT OF THE S EP AR AT ION ST UDY EFFORTDuring ea rl y phases of the manned spac ecraf t pro gra ms, only a smal l separatiostudy effort w a s established . Proj ect Mercury and the Gemini Pro gra m consisted of

flights that involved relatively simple se para tion proce dur es fo r a sm al l number ofmodules o r components. However, the Apollo Pr og ra m consis ted of flights thatinvolved sep arat ion pro ced ure s of gre at er complexity than those of earlier programs,bec aus e of a n inc rea se in the number of modules and components requ iring sepa ratio nThe possible abo rt or alter nate m iss ion s fu rt he r complicated the magnitude of theanalysis required. Therefore, the separation study effort was greatly expanded forthe Apollo miss ion s to en sur e that all accidental r econt act p robl ems would be identifieand resolved or reduced to an acce ptab le lev el of probability befor e eac h flight.

The study of potential reco nta ct pro ble ms began durin g 1964 as part of the Apolltraj ecto ry definition effo rt being per formed at MSC. Var iou s MSC orga niza tion s, witthe assistance of the prime and support cont racto rs for the Apollo Pro gram, wer eoriginal ly involved in the identification of th e potential col lis ion pro bl em s and thedefinition of cor rec tiv e action to avoid th es e prob lems . In addition, gene rali zed separation studies wer e performed by a support cont racto r. In November 1966, the scopeof thi s support was alter ed fr om stud ies of a gene ral nature to studies directly applicble to specific Apollo missi ons . During th is tim e, no sing le MSC organization hadovera ll responsibility fo r the analyses. Conseqilently, cre w procedures, separationtechniques, contr ol modes, and operational req uir eme nts we re obtained fr om a varietof sou rce s. Because of th i s fact, it was recognized that efficie ncy and m issio n operatio ns could be improved by a ssig ning the functional respons ibil ity of perf orming theApollo separati on and rec onta ct ana lysi s to a single organization . Because the problewa s basically operational in nature, the Flight Operations Dir ect ora te (FOD) at theMSC was select ed for th is ta sk in June 1967.

2


7/22

The scop e of the sepa rati on ana lys es was defined officially to include the followingobjectives.

1. Evaluate operationally the separ atio n mode s defined fo r each Apollo mis sio n.2 . Verify separ atio n techniques, identify accidental recontact problem areas,establish a sequence of events, and recommend pro ced ura l o r tech nica l changes when-

ever necessary .3 . Develop the neces sar y analyti cal proce dures and techniques required to pe r-f o r m a detailed analysis.Th e Apollo Joint Separation and Recontact Working Group, composed of r epre -

senta tives fro m the pr im e and support contract ors and of appr opria te MSC personnel,was established in May 1968. At these working group meetings, which werech ai red by MSC personne l, the knowledge and cap abi lit ies of the prime and supp ortcon tra cto rs we re used by seeking the ir technical exp ert ise to aid in decisionmaking andby using their capability to supply most of the input data required for separation anal-yses. In addition, the working group assi st ed each organization in recognizing andunderstanding the pr obl ems faced by the other org aniza tion s; achieved the goal ofcoordinating the activ ities of the different participating organization s in the separationstudy effo rt; and minimized duplications, er ro rs , and confusion.

The sepa ration and recontact effort for the Apollo Pr ogr am remained with theFOD through the fina l Apollo mission . A broad base of ex perie nce was developed byplacing the respon sibility fo r the se paration and recontact effort with one organization.Th is expe rie nce proved advantageous in solving the Apollo sepa rat ion problem s.

COMPUTER SIMULATION DEVELOPMENTThe use of separ able sp acec raft modules o r components for the Apollo Pr ogr amrequi red the development of p roced ures and ma neuvers to ensur e a safe and prop erseparation. The ana lysis included nonnominal si tuations to redu ce the possibility of

accidental recontact o r collision should a contingency occu r. The rec ont act could haveoccurred durin g the actual separation o r ejection process o r subsequently d urin g theresu lting relati ve motion of the separa ted components. The analy sis required the useof th re e- and six-degree-of -freedom co mpu ter pr og ra ms so that all Apollo separationsequencing could be fully and accura tely simulated. To bet ter understand th e complex-ity of the problem and the s yste m modeling required to perf orm the a nalys is, the fol-lowing se para tion events are given in the ord er of occ urr enc e during a typical lunarlanding mi ssion :

1. The launch escap e tower (L ET) fro m the command and ser vic e module (CSM)2. The CSM from the SLA pane ls and the Saturn IVB (S-IVB)3 . The SLA panels fr om the S-IVB

3


8/22

4 . The LM from the S-IVB5. The ser vic e module (SM) scien tifi c ins tru men t module (SIM) bay door fr om

the CSM6. The CSM from th e LM

7. The LM ascent st age fr om the LM descent st age8. The LM from the CSM9. The subsatel lite fr om the CSM

10. The experiment instrument booms from the CSM11. The docking ring and probe a dap ter fr om th e CSM12. The CM fr om the SMOne of the primar y probl ems that had to be solved was the inadequacy of th eavailable thre e- and six-degree -of -freedom computer p ro gr am s to provide the simula-

tion capability requi red to analyze the many Apollo separ atio n even ts. The ref ore , theevolution of the required simulation capabilities was important to the ov erall experi-ence gained fro m the Apollo P rog ram in the se parat ion study effort.

Originally, the compu ter simulation capability wa s a multivehicle, three-degree-of -freedom program that had been modified to perform simple separationstudies. To inc rea se the efficiency in running general parame tric studies, separationcomputer pro grams to be run on an analog compute r sys tem we re developed inmid-1966. Late in 1966, the analog capability evolved into a hybrid computer conceptin which the advantages of a digital and analog system were combined into one com-puter. In mid-1967, a detailed review of t he Apollo se par ati on and recont act simul ationcapability resulted in a decision to continue the support con tra cto r's hybrid prog ramin geffort then in pro gr es s and to tra nsf er thi s capability t o the MSC hybrid sys tem whenthe effort w a s completed. During the sa me period, MSC personn el had expanded andmodified the orig inal two-vehicle, three-degree-of -freedom digita l pro gra m to accom-modate simultaneously as many as eight vehic les; th is expansion, combined with theaforementioned planned tr an sf er of the support contr act or' s hybrid pro gram effort toMSC, was expected to provide all the simulat ion capability ne ces sar y to analyze theseparation sequen ces for futu re Apollo missions.

Because of rigid sch edu les and th e lar ge volume of productivity tha t wa srequi red, the or iginal decision to move all the hybrid capability to MSC was al te re d toretain a portion of the capability at the support contract or facility and to divide thedes ire d capabilities between the two hybrid sys tem s. The altered plan a lso providedf o r the retention of the ent ire set of hybrid capab il it ie s in a digital program that wasdeveloped fo r verification. Initially, the purpose of th is digital progr am w as to pro -vide a verification base f or the two hybrid p ro gr am s and to provide an ana lysis capa-bility fo r time- criti cal problems. However, as the Apollo Pr og ra m ente red theoperational phase, the hybrid cap abili ties could not provide the req uir ed output becau seof the inc rease in tim e-c rit ica l probl ems and the constra ints of the associate d

4


9/22

schedu les. Thus, the us e of these pr ogr am s diminished to such apoint that all maintenanceand updating effo rts we re termina ted; and all furthe r analy sis was performed by usingan all-digital simulation. Termination of t.he hybrid cqxibility a!!owed the c6ficeiiti-a-tion of all re so ur ce s on the all-digital program, named Dynamic Analyzer f or Separa-tion and Recontact, which proved advantageous for solving the sep arat ion ana lys isproblem.

SUMMARY O F S I GNIF ICANT SEPARATION STUD IESFOR EACH APOLLOMISSIONMany separat ion studie s were performed during the Apollo .Prog ram. The moresignificant of the se analyse s, the accidental recont act probl ems identified, and thesolutions to the se p roblems fo r each of the Apollo missi ons are described in this

repo rt to emphasize the importance of comprehensive detailed proj ec t planning. With-out th is detailed planning, some of the problem areas presented and discussed herewould have remained unidentified and unsolved; they could have become critical real-tim e proble ms, jeopardizing cre w safety and mission suc ces s. The following dis cus-sion of the more significant separation problem areas emph asi zes the impact of t heseproblem areas on cr ew safety and mission success . The discus sion demo nst rat es that,through detailed compre hensiv e planning, thes e problem situa tions can be identifiedand resolved before the mi ssion ra the r than during real time o r aft er the flight.

Apollo 1The first Apollo mission was an unmanned ballistic mission performed to assessthe maximum tota l heat rate on the command module (CM) at supercircular entry

velocities and to evaluate the Saturn IB (S-IB) launch vehicle.Poss ib le re contac t situ ations between the CSM and S-IVB, between the CM and

SM, and between the CM and S-IVB were investigated befor e the mission. Resu ltsindicated tha t the CSM sepa rat ion fr om the S-IVB could be perfo rmed with SM reactioncontrol system (RCS) eparation maneuvers as small as 11 secon ds in duration andstill remain free of recontact problems.

A pa ram etr ic CSM separation study indicated that separa tion distan ces variedonly slightly when the CM and the SM were separated at attitu des between 0" and l o o " ,plus-X axis above the local horiz ontal plane. The refo re, an attitu de of 60" above th elocal horizontal plane w as selected fo r the first Apollo mission p rima ril y becauseseparation at this attitude required very little CM orientation to achieve the entryattitude.

A determinat ion of when the S-IB should be shut down to allow the launch escapevehicle (LEV) to perform a safe abor t se paration of the CM fr om th e S-IVB w a s theobjective of another analysi s. The resu lt s of this analy sis indicated that, to ensur e noreco nta ct between the abor ting LEV and the S-IB, the time fo r bo ost er engine cutoffenable should be 40 secon ds aft er lift-off. Th is booster engine cutoff tim e wa s acceptedby the United States A i r For ce Ea ste rn Tes t Range safety personnel.

5


10/22

Apo l lo 2N o separation stud ies were performed fo r the Apollo 2 mission because it was atest of the launch vehicle only.

procedures were required.The spa cecr aft was not flown; therefore , no separation

Apol lo 3The evaluation of the S-IB launch vehic le and a test on the pe rfo rma nce of the

spacec raft heat shield during a high-heat-load, long-duration en tr y w e r e iiie pi-iiiieobject ives of the Apollo 3 mission.

Separation s tudi es of the CSM and S-IVB ind ica ted that a 0.91-m/sec ( 3 ft/sec)separa tio n maneuver would be sufficient to preclude recon tact.showed that, fo r abor t and nominal ent ries , SM separatio n fr om the CM at 60" abovethe positive local horizontal would yield maximum separa tion cle aran ces. In addition,the study showed that, if the CM serv ice propulsion sys tem (SPS) failed to ignite at thetime of the first burn, acc identa l recontac t could be avoided by extending the CSM RCSplus-X translation to 7 5 seconds, by initiating the S-IVB engine cutoff, and by ensuringthe S-IVB attitude con tro l during venting.

The CM/SM study

Apol lo 4The unmanned Apollo 4 missio n used t he S-IVB to inse rt the sp acec raft into acirc ular Earth parking orbit and to perform a simulated translunar injection maneuverto place the spacec raft on a highly elliptical Earth-intersecting traject ory.Only two separat ion s, CSM/S-IVB and CM/SM, we re as sociated with the Apollo 4

missio n; however, se ver al accidental recon tact problems wer e identified. Durin g thelaunch phase, i f an ab ort occu rred involving an SPS ignition failur e, recon tact betweenthe CSM and the S-IVB would be imminent. Thi s problem could be avoided by com-manding the S-IVB ullage maneuver off upon S-IVB abort shutdown.

Fo r the nominal CSM/S-IVB separa tio n sequence, an inoperat ive SPS could resultin recontact between t he spacecr aft and fragme nts of th e destru cted S-IVB aft er theplanned bulkhead-reversal test. The solution fo r avoiding thi s problem wa s to estab-li sh warning time s, fo r the cre w and ground control , that would provide sufficient ti meto enable a two-jet o r four-jet RCS separation to achieve a safe separation clearan ceand avoid recontact with the debris.

Another potential problem area conce rned the possib ility of the LEV recontactingthe Saturn IV boost er i f a launch phase abo rt occurr ed. If the booster were to cut offat approximately 42 seconds aft er lift-off, the 243.8 -meter (800 foot) con strain t onlateral separ ation dista nce between the LEV and the boo ster would be violated. There-for e, the recommendation that the booste r enable cutoff setting be 30 secon ds o r lesswa s implemented to ensur e adequate sep aratio n clea ran ces.

6


11/22

Apollo 5The Apollo 5 mission w a s flown pri mar ily to test the LM in a near-Earth-orbitalenvironment for verifiration nf LM sys tems . After orbit insertinn, the aerodynamic

shro ud wa s sep ara ted fr om the S-IVB and then the fou r SLA panel s we re opened. AfterS-IVB/LM separ ation , LM ab or t staging using the as ce nt propulsion sy ste m w a splanned.

The nOXin31 aerodyn amic shroud jettison proce dure w as foilnd to be free of recon-tact problems, and recommendations we re made for shroud jettison p roced ures to beused in the event of an ab or t missi on. The S-IVB passivation expe rim ent pro ced ure swe re analyzed, and the re su lt s were used to define the safe ran ge of operat ional atti-tu de s tha t would avoid potential coll ision s during the exper imen t. The LM abor t stagingsequence was free of rec ontac t probl ems as planned.

Apollo 6The Apollo 6 mission was sim ila r to the Apollo 4 missi on in that a translunarinjection bur n wa s simulated. Following tra nsl una r injection, it wa s planned that the

S-IVB at titud e propulsion s yst em would per for m a reori entat ion maneuver to the CSMSPS ret rog rad e burn attitude; then the CSM would sepa rat e and ret ur n to Earth . Thesep ara tio n ana lys is of t hi s proce dur e indicated that the S-IVB would probably reco ntactthe CSM if the S-IVB attitude propulsion sys tem failed to proper ly o rient the CSM to theco rr ec t burn attitude. A solution to the prob lem specified that the CSM, and not theS-IVB, should orie nt to the burn attitu de and that the CSM should perform a 10-second,four-jet RCS translation to achieve adequate clearance from the S-IVB. Th is solutionwas a lso recommended to preclude a possible recontact problem for an earl y tran s-lunar injection burn termination .

Another CSM/S-IVB reco ntact problem w as identified fo r an alt ern ate Ea rt hparking o rbi t whe re the S-IVB was t o be positioned above and in fron t of the CSM at thetime of the SPS burn ignition. A pro ced ure w a s designed to eliminate rec ontac t forth is case by prop er orb ital positioning of the CSM/S-IVB sepa ra ti on maneuver at onerevolution be fore the first SPS burn.

Apollo 7Apollo 7, the first manned Apollo mission, was flown to evaluate the crew/space craft operational compatibility, to evaluate a trans posit ion and simulated dockingexercise, and to evaluate a CSM-active rendezvous procedure.A l l CSM/S-IVB sep ara tio ns investigated wer e free of recon tact prnbl ems exceptf o r launch phase abor ts after the LET jettison in which the SPS thrust-vector-controlrate damping wa s used. Thi s recon tact possibility, which wa s identified to exist at anytime during the entire launch phase, could be proce dural ly avoided by req uir ing afaster cr ew r eac tio n in identifying and initiating an abort.

7


12/22

A relative motion an alysi s of the tran sposi tion and docking simulation exe rcis e,and of the CSM-active rendezvous, indicated that all sequences were free of accidentalrecontact problems.

Beginning with the Apollo 7 mission, the sepa ration attitude fo r the CM and SMwa s changed f rom an in-plane attit ude used on prev iou s Apollo mis si on s to an out-of-plane attitude. The new sepa ratio n attit ude wa s to be attained by iner tia lly holding thede orbi t burn attitude and yawing the CSM 45" out of th e or bi t plane. This procedurewould preclude a possible recontact between the CM and SM fo r any entry, reg ar dle ssof the aerodynamic lift profile of ei th er vehicle. The recommen dation was made andaccepted to perform tiis out-of-piane 5ivI jeiiisuri as s w u r i as possible &ei* deoi-bit burcutoff to incr eas e the separation dis tance s during entry .

Procedures were designed fo r a contingency in which nominal CSM se par ati onfrom the S-IVB could not be accomplished. If th is failure occ urr ed, only the CM woulbe sep ara ted fr om the SM/S-IVB configuration, and the miss ion would be terminated.

Postflight anal ysi s of the SM rel ati ve motion fro m the CM wa s made fo r thi smis sio n because of a deviation between the act ual SM tra jec tor y, which was dete rmineby rad ar tracking, and the predicted traject ory. The reconstructed trajec tory indi-cated a total relat ive sepa ratio n del ta velocity (AV) of a pprox imate ly 9.14 m /se c(30 ft/sec), which wa s much sma ll er than the predi cted 88.39 m/sec (290 ft/sec). Threduction in relative separ ation velocity occ urre d because the SM failed to rem ainspin stabilized throughout the complete sepa rat ion mane uver. However, the an al ys isindicated that, although the se par ation velocity had been reduced significan tly, thereduction did not create a recon tact probl em between the CM and SM dur ing entr y.

Apo l lo 8The Apollo 8 mis sion was the first manned flight in which the three-stageSaturn V rocke t boost er was used and the first missi on in which the cr ew orbited the

Moon. Therefore, new separation procedures and analyses were required for almos teve ry phase of the mission. Abort and alte rnate mi ssion plans wer e new and we re thooughly analyzed for identification of po ssib le recon tac t probl ems. The s am e kind ofCSM evasi ve maneu ver used duri ng the Apollo 7 mission to evade the S-IVB was notapplicable fo r the Apollo 8 mission because the Earth-orbital effects on relative motiocould not be used advantageously. The evasi ve maneuv er fo r the Apollo 8 missionwould be required after the trans lunar injection burn; theref ore, the separ ation tra-jec tory would not be pertu rbed significantly by orbita l effec ts. The CSM eva siv emaneuver was redesigned to c onsis t of a 0.46-m/sec ( 1 . 5 ft/se c) RCS trans latio n alonthe positive r adi us vecto r of the Eart h (away from the Ear th). The maneuver wa sdesigned to be initiated with the spa cec raf t located in a stationkeeping position15.2 meters (50 feet) ahead of and 12.2 meters (40 feet) above the S-IVB, with theCSM apex pointed toward the Earth.

The CSM minus-X RCS tr an slat ion should have produced an adequate, safe dis-placement fr om the S-IVB and avoided any re conta ct pr obl ems du ring the launchvehicle lunar targeting maneuv ers. However, the actua l stationkeeping maneuv erswe re not executed c orre ctl y; at the t ime fo r the nominal evasive maneuver, the sp acecr af t was not located in the co rr ec t stationkeeping position. Th is fac t wa s unknown to

8


13/22

the cr ew o r ground control at the time and, thus, the evasi ve maneuver wa s executedfr om the in co rr ec t rel ati ve position. The objective of satisfa ctor ily evading the S-IVBwa s not achieved, and th is fai lu re was soon confirmed through visual tracking r eport edcrew in real time, and executed satis facto rily to alleviate the unfavorable relat iveposition between the two vehicles. Postflight analysis of th is problem r evea led thatthe space craft did not or ient the full 180" in the pitch plane af ter initially separ atingfr om the S-IVB, which wa s nec es sa ry to null the sep ara tio n velocity completely andtherefore esta3lish a stationkeeping position. In addition, the spac ecr aft plus-X axiscould not be alined visually along the negative ra diu s ve ctor as originally planned fo rthe first evasive maneuver. As a resul t, the spacecraft attitudes were approximately15" and 45" in e r r o r fo r the first and second maneuvers, respe ctive ly.

hy the p r p w . -4 secnfic! grGflnd-cnmputec! evasi...e rnLqeu\'cr -3;;sdefined, relayed to the

It wa s evident that for fut ure mi ssi ons each stationkeeping and e vasive maneuve rshould be defined to include th e following: (1) specific spac ecraf t inert ial measure mentunit gimba l attitudes that are computed and simulated before the mission and updatedin real time and (2) isu al, out-the-window monitoring at ti tude s of the S-IVB o r anyot he r spac e vehicle in the vicinity of the spac ecraf t when mane uver s are planned.

The Apollo 8 mission was al so the first mission during which the SL A panels werejettisoned fro m the S-IVB. The prim ar y objective of jettisoning the fou r panel s wa s toprev ent the SM RCS thr ust plume s fr om reflecting off th e deployed p anels and onto theLM. At space craft separation, ih e SLA was pyrotechnically severed from the SM andinto fou r pa nels tha t we re hinged to the S-IVB. As he CSM trans lated forward, thefou r pane ls deployed o r rotated outward fr om the S-IVB and LM; and, a ft er openingthrough an angle of approximately go", they were spring 6jected from the S-IVB witha velocity of 2.44 m/sec (8 ft/sec) o r g r ea t e r.

An analy sis was performed to deter mine i f the panel jettiso ns would cr ea te anypotential recont act probl ems with the sp acecr aft fo r the nominal mis sion, fo r launch-phase o r Earth-orbital aborts, o r fo r alternate missions. The res ult s indicatedadequate sep aration clear ances for al l ph as es of flight, with the s ingle exception ofretrog rade mode I11 SPS abort s. For this particular abort, the essenti al retro grademode I11 burn would c ause the space craft to fly in and through the area in which thefo ur pa ne ls had been jettisoned . However, th e low probability of a mode I11 retrogradeab or t combined with the low probability of recontact was co nsider ed acc ept abl e.

Apollo 9The Apollo 9 mission was a 10-day, Earth -orbi tal mis sion that wa s flown to

dem ons tra te the combined oper atio nal capability of the CSM and LM to p erf orm selec tedfunctions of th e lunar landing mission.

The following nominal separations were analyzed fo r the immediate, close-in,and eventual recontact regions: CSM separat ion fr om the S-IVB, jettison of the fourSLA panels, LM ejection f ro m the S-IVB, LM undocking fr om the CM, LM staging,LM asc ent s tag e jett ison fr om the CSM, and CM/SM separ atio n.

9


14/22

Following transpo siti on and docking, the CSM/LM configuration was sp rin gejected from the S-IVB. Th is was the first Apollo miss ion in which the docked con-figuration used four compr esse d spr ing s to expel itself fr om the S-IVB. Originally,had been planned to use the CSM RCS th ru s te rs to withdraw the LM; however, pr im arbecause of jet plume impingement onto the LM, the withdrawal technique was rep laceby the spring-ejection method. Th is wa s the s ame kind of proble m tha t result ed in thdecisio n to jettison the f our SLA pane ls on the Apollo 8 mission. The LM ejectio nprocedure was evaluated and proved to be a success ful separati on technique fo rApollo 9 and subsequent missions.

r n r1ne per r~or r~ i i ~ ~ i~ ef a piaiiiied posiejeciioii CSTVT/LTVTai-iii-oi-biidi maneuver ioevade the S-IVB wa s based on prem issi on sep arat ion and reco ntac t anal ysis . Aminus-X RCS tran slat ion burn wa s to be performed in a pitched-down attitude, takingadvantage of the continuous propulsive venting of liquid hydrogen fr om the S-IVB andadvantage of the orb ital motion effe cts to produce the desi red sepa rati on cle ar anc es.

For LM jettiso n and subsequent asc ent propulsion sy ste m testing , the sp acec rawa s to be placed in a safe rel ativ e position with res pec t to the LM by ori enting theCSM/LM configuration to the asc ent propulsion sy st em burn attitude and then holdingthe configuration inerti ally stable. Before executing the burn, the CSM wa s to jettisothe LM, maneuver to a stationkeeping position, execute a pitchdown and yaw out-of-plane orientation maneuver, and then perform a four-jet minus-X RCS evasive maneuve r to obtain the desi red disp lacement between the CSM and the LM asc ent stag e. Thprocedu re would pe rmit as much as a 0.61-m/sec (2 t/sec) e r r o r in stationkeepingrela tive velocit ies and would still ensure a safe separation distance at the t ime ofas cen t propulsion syst em ignition. It was designed to avoid the repeti tion of p robl emexperienced duri ng the Apollo 8 mission because of stationkeeping maneuver er ro rs .

A new SLA panel separ ati on and reconta ct evaluation wa s based on new att itu deand resultan t velo citi es that encompas sed the highe r panel-deployment rates of 60 to74 deg/sec observe d in the Apollo 7 postflight anal ysis . The higher opening rate of74 deg/sec could re sul t in panel jettisons at a maximum attitude of 130" and at amaximum velocity of 4.27 m/sec (14 t/sec). Minimum valu es previou sly analyzedwere 90" and 2.44 m/sec (8 ft/sec). The new separa tion atti tude s and jettison veloc-ities were fre e of recontact proble ms for all mission phases except fo r the ret rogradmode III ab or t region, which was previously identified.

Apollo 10The Apollo 1 0 mission w as flown to eva luate the LM operati onally in lunar orbiwhile separat ed from the CSM. N o landing was a ttempted , but the LM did maneuver

into a low perigee orbi t s im il ar to the one fr om which a landing would be attempte d othe next Apollo mission.The Apollo 10 miss ion w as the first mission fo r which a separation procedur eshandbook was published in addition to the sep arat ion analy sis s umm ary document. T

procedur es handbook was originated prima rily to furnish the flight cont roll ers a con-venient separation reference during premission simulations and during the act ualmission. It proved to be a valuable aid and was published f o r each of th e rema inin gApollo missions.

10


15/22

Beginning with the Apollo 10 mis sion, a more conser vativ e transposition anddocking proce dur e wa s defined, with only min or changes fr om the one used on theApollo 7 and 9 missi ons. The initial separ ation velocity and the first nulling maneuvertance f ro m the S-IVB. The new trans posit ion and docking pro ced ure was found safe,and was used without change fo r the re maining Apollo m iss ion s..&?,by reduced in magn;,u& te sax,..re-R.CS p .Q p e ~ ~ ~ f i t2d te decrease the EeparatiQEdis-

Pr im ar il y to c onserv e RCS fuel and to eliminate the effects of e rr or s for smal lRCS maneuvers, a decision was made to use the CSM SPS fo r the post-transposition-docking-and-extraction evasive maneuver fr om the S-IVB. Th is w a s a new separationproced ure f or the Apollo 10 mission; and subsequent ana lysis revealed that, if theS-IVB liquid hydrogen propulsive vent did not close before LM ejection, a potentialrec ont act problem would ex is t between the S-IVB and LM af te r ejection. The ref ore , adecision was made to orien t the CSM t o the evasive maneuver attitude immediate lyafter ejection and to p erform a 5-second plus-X RCS transl ati on, if necessary.

During the nominal lun ar rendezvous, the LM des cen t st age was planned to bejettisoned in a posigrade direction 10 minutes before the ascent-stage insertionmaneuver, which was perfor med in a ret rog rad e direction 27" above the local horizon-tal. If LM staging were performed as planned at 10 minut es before the insertionmaneuver, then the LM descent stage would be ahead and above the ascent stage at theti me of the ins ert ion maneuver , and ther e would be no possibility of rec ontac t. How-ever , if staging we re executed ear ly, then the tr aj ec to ri es of the two vehicl es wouldint ers ect . The descent stage would have required 65 minutes to rea ch this intersectionpoint, and the LM as cent st age would have requi red 2 minutes to rea ch the sam e point.Consequently, if staging had been perfor med at 63 minutes before the in serti on maneu-ve r, the asc ent sta ge would have recontacted the des cen t stage 2 minut es after theinsert ion maneuver. For this reason, early staging w a s to have been avoided. Toensure that the descent stage was in a safe relative position fo r the ascent stage inser -tion burn, and thus avoid any accid ental recon tact proble ms, it wa s recommended thatstaging not be performed between 53 and 73 minutes before the insert ion maneuver.

After staging, the descent-stage motion relative to the CSM was r etr ogr ade ; and,becaus e of the longer orbit al period, the descent stage would approach the CSM froma posigrade direct ion approximately 15 revolutions later. Real-time monitoring of thisrelati ve motion problem was performed t o determine if a CSM out-of-plane eva sivemaneuver was required. The crewmen saw the descent stage approaching, but it wasdeterm ined through real-ti me computations that the desc ent sta ge would pas s to therear of t he sp ace cra ft with adequate separation clearance. No CSM evasive maneuv erswere required.

Originally, the proc edure fo r LM ascent-stage jettison was planned t o be thesame as that used du ring the Apollo 9 mission, in which the ascen t propulsion sys temwa s allowed to burn to fuel depletion. At the crew's preference, the procedure waschanged to the extent that, aft er asc ent-s tage jettison, the CSM would mane uver fro ma position below the LM to one above it and then perform a radially upward evasiveman euver of 0.61 m/sec (2 ft/sec). Thi s would cause the s pacecr aft to tra nsl ate aboveand behind the LM and place the crew in a favorable relati ve position fo r observing theascent propulsion burn. Upon asc ent -st age jettison, however, a separ ation velocity,which was l ar ge r than predicted, was imparted to the CM and LM fr om the se veranc eof t he docking tunnel. Thi s caused the spacecraft to tra nsl ate rapidly below and ahead

11


16/22

of the LM, which resulted in sunlight interference in the CM windows when visualacquisition of the LM was attempted. Ther efo re, t he crew-p refe rred relati ve positionwas not attained because th e neces sary CSM maneu vers to achieve that position wer enot attempted. The CSM dr if ted below and ahead of th e LM instead of above and behindHowever, no recontact problems were detected. A separati on velocity that was higherthan expected was probably p roduced by higher-than-expected ga s pre ss ur e i n thedocking tunnel,

Apollo 11The Apollo 11 mission was flown to attempt the first manned lunar landing, toconduct exploration and scientific experi ments, and to r etr iev e lunar soi l and rock

samples.Fo r the Apollo 11 missio n, the LM asc en t stag e did not contain enough prope llantto achieve sol ar or bit with a fue l depletion burn as had been done on the Apollo 10mission. Instead, the separation procedure for the ascent stage was altered to leaveit in lunar orbit and place the CSM in a safe position t o continue the nominal ti me line.The jettison sequence fo r the LM wa s planned and executed with a local spacecraftpitch attitude of 4 5 " and a four-jet minus-X RCS transl ation fo r a A V of 0.30 m/sec

(1 ft/sec ), to avoid any possible recon tact.Possible ascent-stage recontact with the descent- stage plume deflec tor s durin g

lift-off from the lunar surface was investigated fo r thi s first lunar landing mission .Lunar slopes as steep as 4 5 " , which was the s tati c stabil ity limit of the LM, we re con-sidered . Resu lts of the ana lys is indicated that even f o r a worst-case ( 4 5 " ) slope, aminimum cleara nce of over 18 centimeters (7 inche s) wa s maintained between theasc en t stage and the plume deflec tor s of the desce nt stage.Immediate recontact problems associated with the LM RCS staging under nonnominal and altern ate miss ion conditions wer e al so investigated. Nominal LM staging

under abort-guidance-section contr ol could best be performed with narrow dead-bandl imits for ra te s less than 6 deg/sec; and, for rates of 6 deg/sec o r greater, stagingcould best be accomplished with a l-s econ d plus-X RCS di re ct ullage mane uver withoutLM digital autopilot or abort-guidance-section contr ol. No contr ol mode was availablethat would completely eliminate interst age recon tact for an inadvertent staging. Recontact would most likely occ ur between the ascent- stage fuel tank brace and the jet plumedeflector suppo rts approximately 10 seconds after inadvertent staging when narrowdead band was used and the ascent propulsion syst em igniiiion si gna l delay was 0 sec-onds. For nominal, powered-descent, abort -stag ing procedure s, a recontact situationexisted at the environmental control syst em lin es when the descent propulsion sys temtail-off thr ust was nominal or gr eat er than nominal. To avoid this problem, adecision was made to use plus-X RCS translation at LM staging.

Fo r CM/SM separat ion, a minus-X RCS separ ation burn t o fuel depletion wa splanned to incre ase the SM ent ry velocity and dec re as e the flightpath angle so the SMwould gra ze the Ear th atmosphere and "skip out" into an orb it with an apogee gr ea te rthan 926 000 kilometers ( 5 0 0 000 nautical miles). However, based on crew sightingsof th e SM and based on the a tmo spheric breakup, it wa s concluded that the SM sep ar a-tion sequence was not occur ring as predicted. Although th ere was no apparent

1 2


17/22

recontact problem, the fact that the Apollo 11 crew saw the SM when they should nothave and the fact that r ad ar tracking of the Apollo 10 SM al so indicated th is anomalyprompted a detailed study of the SM stability and traje cto ry deviation after separation.The purpose of the study wa s to determine if an unstable SM coi-ilc! r e s ~ l tn the devi&,edtra jec tor y and, if so , to dete rmi ne the ca use of the SM instability following separ ation .If possible, a new proc edur e that would ens ure a stable SM sepa ratio n burn would bedefined and verified. Par tic ula r empha sis wa s given to the effect of propellant slosh,which was shown to be capable of producing the erratic SM behavior that wa s actuallyobserved . Two dynamic m o d e l s of propellant slos h we re developed and wer e used toanalyze the SM jettison procedure. The re su lts led to the use of a new jettison proce -du re that would e nsu re stable SM at ti hd es during the separation burn. The newproc edur e involved reducing the RCS ro ll thr ust duration from 5.5 to 2.0 seconds andsetting the minus-X RCS burn to a dura tion of 25 seconds instead of a burn to propellantdepletion. Thes e changes we re used fo r the Apollo 1 3 missi on and subsequentmis sio ns. Sufficient tim e w a s not available to make the nece ssa ry SM hardwarechanges for the Apollo 12 mission; theref ore, SM separation was perf ormed using theburn-to- fuel-depletion sequence of the Apollo 11 mission.

Apo l lo 12The Apollo 1 2 mission was flown to continue lunar surface exploration and scien-tific ex periments and to re triev e lunar soil and rock sa mples from a different surfacearea than the Apollo 11 mission.For the Apollo 12 miss ion, the attitude propulsion sys tem burn, which was usedon the Apollo 11 missi on to tar get the S-IVB f or a preselecte d luna r impact point, wa sals o used as an evasive maneuver. Thi s action deleted the require ment fo r a space-craft SPS evas ive burn following transpos ition , docking, and extraction. The newevasive procedure was analyzed fo r the Apollo 1 2 mission and verified as being free

of recontact problems. After CSM/LM e jecti on fr om the S-IVB, the sp ace cra ft wa soriented to a viewing attitu de that would enable the c re w to visually monitor t he S-IVBevasive maneuver sequence. The spa cec raf t vieiving attit ude al so doubled as a backupCSM evasive maneuver attitude.

Apollo 13The Apollo 13 miss ion was flown to continue lunar su rfa ce exploration and ex per i-men ts and to collec t additional soil and rock sampl es. Analysis of the Apollo 13 launch

vehicle ope rational tra jec tor y indicated that the procedure fo r mode I1 abo rts of orie n-ting the CM to the entry attitude pr io r to SM jettison wa s not the best procedure toproduce maximum separation dista nces duri ng entry. The analy sis indicated thatorientation to the CM entry attitude befo re the SM separat ion was sat isfactory for anabort from a nominal booster trajectory. However, certa in other abort s (e. g., anabort resul t ing from a time-of-free-fall limit-line violation) would result in SM jettisonne ar the time-of-entry interface that, if performed with the CM oriented t o the en tryattitude, would reduce significantly the separation clear ance s. Ther efor e, a decisionwas made to jettison the SM immediate ly following the RCS abo rt ma neuv er before theori en tat ion of the CM to the en try attitude.

13


18/22

The lunar descent phase abort proc edures we re entirely different from thosepreviously planned for the Apollo 11 and 1 3 mis sion s because the CSM was to pe rfor mthe descent orbit insertion maneuver fo r the Apollo 1 3 missi on. The new procedureswe re evaluated and potential recont act pro ble ms we re identified. The most probableopportunity for a recon tact p roblem to develop would be f or a powered descent abortthat w a s initiated at approximately 10 minutes and 10 seconds after the fi rs t powereddescent initiation maneuver. If an abort occurred at this t ime, a recontact problembetween the CSM and the LM asc ent s tage would exist. Th is problem wa s resolv ed byspecifying tha t the LM execute an out-of-plane velocity component in its insertionmaneuver, o r that the CSM perform an out-of-plane maneuver at the tim e of LM ins etion.plane avoidance maneuvers, but only after the dan gers of an accidental collision hadpassed.

Subseqiieiit iliziieuvers during X X E ~ ~ Z V O EWOU!~! null the effects of the out-of-

When the Apollo 13 SM oxygen tanks ruptur ed durin g tra ns lun ar co ast , no em e rgency proce dure had been defined for th e exact situation in which the SM wa s ino pera-tive and unusable fo r re tu rn to Ear th. However, a premission separation procedurehad been defined fo r return ing to Ea rt h with the LM, jettisoning it 1 hour before entryand then jettisoning the SM before entr y interface. Thi s proced ure specified that theSM wa s to produce the s epara tion velocity requi red f or LM jettison by using its RCS.Because t his syste m was now inoperative, a technique w a s developed fo r using the LMRCS to produce the SM jettison velocity and using the LM docking tunnel pressure toproduce the LM separ ation velocity. The sep ara tio n attitu des f or the SM and LM werthe same as those defined in the premis sion a nalys is for LM jettison 1 hour beforeentry interface.

The SM jettison procedure wa s defined by the following: aline the LM plus-Xaxis along the positive rad ius vecto r, away fro m Eart h; yaw out-of-plane to the southof the CM groundtrack; perform an LM RCS four-j et plus-X trans latio n burn fo r avelocity of 0.15 m/s ec (0.5 ft/ sec ); j ettison the S M ; nd perform an LM minus-Xtranslation burn fo r a velocity of 0.15 m/sec (0.5 ft/se c) t o null the origina l maneuveTh is separat ion maneuver, perfor med with the push-pull technique, wa s to occur2 hours before entry interface and leave the CM docked with the LM.

The LM/CM sepa ration proc edu re that wa s planned to occur 1 hour before entrinterface was analyzed f or separation velocity values of 0.30, 0.61, and 0.9 1 m/sec(1.0, 2.0, and 3.0 ft/sec ). The re su lt s provided a parametric evaluation of the relative separation dis tan ces fo r vari ous LM separation velocity values. The evaluationwa s nec ess ary beca use of the uncertainty as socia ted with the predic tion of sep ara tio nvelocity result ing fro m se ver ing the p res su riz ed docking tunnel between the LM andCM. The LM jettison attitude wa s to be attained by alining the CM plus-X axis withthe positive r ad iu s and yawing 45" out-of-plane to the south, which wa s the sa me asthat used f or th e SM.

As the Apollo 1 3 mission progressed, other separation tim es and procedureswere evaluated. For SM jettison, sep ara tion t im e s of 3.5, 4.5, 5.5, 6.5, and7.5 hours before entry interface were assessed. It was discovered during these simlat ions tha t, because of the earlier jettison times then under consideration, the SMout-of -plane jettison would not re su lt in gr ea te r separ ati on cl ea ra nc es . Consequentlthe recommendation was made that SM jettison be perform ed in-plane to achieve themaximum separation displacement fro m the CM. The latest separation procedures,14


19/22

in which SM jettison was set at 4.5 hou rs before entry in terfa ce and LM jettison at1.0 hour be for e entry inter face, wer e furnished to the flight con tro lle rs and verifiedas being free of any recontact problems.The act ual SM and LM sep ara tio ns we re not performe d exactly as had been

planned. The SM jettison occurred 4 hours 40 minutes before entry, ra th er than theexpected 4.5 ho ur s bef ore ent ry inter face. The separ ation velocity of 0.58 m/sec(1.9 ft/sec ) achieved f ro m the push-pull technique wa s la rg er than the predicted0.30 m/sec (1t/sec). As a result, the separation dista nce s between the SM and theCM and LM we re gr ea te r than predi ction s made befor e separa tion. The LM jetti sonoccur red approximately 11 minutes earlier than planned and in an unplanned, butaccep table , attitude. The LM wa s jettisoned 64" out-of-plane to the north rat he r than45" out-of-plane to the south, and the pitch attitude of 64" above the local hor izo nta lshould have been 90". The actua l separ ation velocity was 0.76 m/s ec (2 .5 ft/se c),0.15 m/sec (0.5 ft/se c) higher than the highest value considered before jettison. Thi scombination of differences resulted in a decrease in total separation range from thatpredicted. At ent ry interface, the LM/CM clear ance was 1372 me te rs (4500 feet),approximately 1067 me te rs (3500 feet) cl os er than previously esti mated; however, norecontact proble ms were evident and none were rep orted.

Apollo 14The Apollo 14 miss ion was flown to continue lunar sur fac e exploration, to conduct

scientific experi ments , and to collect lunar surface sample s. During th is mission, theCSM encountered difficulty in performing a sat isfactory docking with the LM/S-IVBconfiguration after tr an slun ar injection. Because of thi s problem, plans for using thebackup evasive maneuver separ ation technique were considered. After severalattempts, a suc ces sfu l docking was finally accomplished, and the nominal post tr an s-position, docking, and extractio n evasi ve maneuver sequence wa s used. Using thesame procedures that had been developed for previous missions, all other separationphases of th e mi ssion were nominal and without incident.

Sev era l new contingency proce dur es were developed f o r the Apollo 1 4 mis sio nbut we re not requi red. The se pro ced ure s included a method fo r safely ret urning thecrew if the SLA ailed t o sep ara te from the CSM after the tran slu nar injection maneu-ve r. Th is method would involve flying the complete launch vehic le/sp acecra ft config-uration through a tr an slun ar coast, around the Moon, and to a tran sea rth coast; andthen executing a C M separation before Ear th entry. Specific separation pro ced ure swe re al so developed and documented to co ver an SM contingency of the type tha toccur red on the Apollo 1 3 mission after the trans lunar injection.

Apol lo 15The Apollo 1 5 mis sio n wa s flown to continue lunar su rf ac e exploration, to conductscientific experiments, and to collect lunar surface sampl es. This was the first

Apollo missi on to use the lun ar roving vehicle (LRV) to aid i n s urface exploration andsample collection and also the first mission to place a subsatell i te in lunar orbit .

15


20/22

The addition of a noth er int ers tag e oxygen quick-disconnect line nece ssi tat edstag e-se parat ion s tud ies of the Apollo 15 LM f o r the nomina l lift-off sequence, fo r theab ort s from main powered descent, fo r an RCS staging, and fo r an inadvertent stagingTh is high- press ure line would exert app rox ima tel y 4448 newtons (1000 pounds) of additiona l force between the st ag es when it was severed at sepa ration , substantially modi-fying the LM staging dyna mics . The Apollo 15 ana lys is indicated tha t the RCS andinadvertent staging dynamic s we re improved by the addition of t hi s int ers tag e oxygenline, and the se para tion sequence fo r the nomina l lunar lift-off staging precluded arecontact with o r without the additional oxygen line. However, the Apollo 15 ana lys isdid indicate the possibility that one of the plus-X RCS th ru st er sk ir ts could be in theward during staging. Th is situation would exist only for aborts initiated after descentengine shutdown durin g the main powered de sc en t phase, and the re con tac t problemcould only occ ur f o r negative pitch rates combined with equal positive roll rates on thLM during staging, which was not expected to oc cur .

- r C L4Lll of, aid recontact, one of the f o u r thrust plume deflectem as they cdlzipsec! eut-

New nominal sep aration proce dures we re requir ed fo r the Apollo 15 miss ionbec aus e of the addition of the SIM to the mis si on configuration. The SIM, which wasinstalled in one of the SM str uct ura l bays, contained scientific exper iment instr umentfo r use in lunar orbit, The struc tur al met al surface of the SM, covering the SIM fo rprotection, w a s designed to be se ver ed pyrotechnically and jettisoned bef ore ins tr u-ment use. A jettison procedure was defined and used successfully 4.5 hou rs beforelunar orbit insertion. One of the instr ument s contained in the scientific instr ument bwas a subsatellite that was designed to be jettisoned and left in lunar orbit. On theApollo 15 mission, the subsatelli te was jettisoned at the orb ita l ascending node, nor mto the ecliptic plane, toward the north, on revolution 73. This jettison attitude and th1.22- mIsec (4 ft/sec) jettison velocity ens ured adequate sep arati on cl ea ra nc es betwethe subsatel lite and the CSM for the Apollo 15 missi on.

New contingency jettison procedu res f or th e two ex perim ent instr ument boomswe re developed f or use in the event that one of the booms, which wer e deployed to alength of 6.10 m ete rs (20 feet), failed to retract satisfactorily into the SIM bay. Thebooms, which must be retr ac te d before all major SPS maneuvers, wer e designed to bjettisoned fr om the SM in the event of a n unsa tisfa ctory ret rac tio n.

Apollo 16The Apollo 1 6 mission wa s flown to continue lunar orb it and surfa ce e xperi menand fo r luna r sur fac e exploration and sam ple collections. The LRV and sub sat ell ite

we re als o used on Apollo 16.The Apollo 1 5 nominal s epar ation pr oc ed ure s we re used on the Apollo 16 mis sioexcept for the final jettison of the LM asc ent stage , which occu rre d in lunar orbit af t

ren dez vou s and docking on revolution 53. The CSM 0.61-m/sec (2 ft/ se c) ev asi vemaneuver fro m the ascent stage was executed at 5 minutes after the LM jettison andwa s made in a posigrade i nste ad of a retrograde direct ion as wa s planned fo r theprevious missio ns. Th is change ensure d that the CSM would be pointed away fr om thLM ascent stage fo r the evasive maneuver, even i f it wa s delayed as much as45 minutes.

16


21/22

No new contingency pro cedur es wer e developed fo r the Apollo 16 mission. Thesa me contingency or alter nate mission proc edur es developed fo r previous Apollomi ss io ns would have been used, if required.

Apollo 17The final Apollo miss ion was flown to continue lunar surfa ce s ampl e collect ions,to continue exploration , and to conduct or bita l and sur fac e scientif ic exper imen ts.The final lunar landing mission of th e Apollo P ro gr am used s epara tion proce dur esidentical to the previous mission with the addition of jettison pr oce dur es fo r the two

high-frequency antenna booms of the lun ar sounder expe rime nt. The se two ante nnaswe re nominally re tra cte d into th e scientific inst rume nt bay of the SM, but would haveto be jettisoned if a retraction failure occurred. No failure occurred, however, butjettison p roc edu res wer e defined and verified before the Apollo 17 mission. N o otheraccidental recontact problem areas we re identified o r encountered d uring this miss ion.

C O N C L U S I O N S A N D R E C O M M E N D A T I O N SF o r eac h of the Apollo missi ons , all separation procedu res fo r the nominal,

alte rnate , and abort mis sion tim e lin es we re documented and published in one docu-ment. The separat ion proce dures handbook became a valuable aid in quickly deter-mining the c or re ct separation sequencing to be used in the event a contingencydeveloped duri ng the mission. The handbook concisely describ ed each separ atio nsequence, pictorially depicted the vehicles being sepa rated in the ir c orr ect separationattitude, and graphically presented the resulting relati ve motions for each vehicle.For so me sep ara tio ns, the spac ecr aft out-the-window viewing attitud es we re providedso the c rew could visu ally tra ck what they were separating from.

The Apollo sp acec raft sep aration and recontact study effo rt prog resse d fr omsmall param etric separation analyses for the first two mis sions to ma jor evaluationsof all separation sequences fo r the remaining missions. The re su lt s and recommenda-tio ns of th es e an alys es we re obtained fro m many organiza tion s, both Government andcont ract or. A clos e working relationshi p was maintained among the personnel per -form ing the s epar atio n studies, the flight cre w support personnel, the flight con-tro lle rs, and the support contractors.

U se of t he spa cecr aft prime c ontra ctors later in the program helped make thetotal effort a su cc es s because of th ei r knowledge of the st ru ct ur es and su bs ys te ms ofthe s pac ecr aft . The establi shme nt of the Apollo Join t Separation and Recontact WorkingGroup, an informa l organization with represe ntati ves from the pri me and su pportcont racto rs, aided significantly in obtaining full cooperation fr om the v ariou s organiza-tio ns and individuals involved in the sep ara tio n activities. Fr om the experien ce gainedto date, it is recommended that, for future programs, the spacecraft prime contra c-t o r s be included in the major separation and recontact analyses, starting at the begin-ning of the program.

17


22/22

F o r large complicated sy st em s like those involved in the Apollo Progra m, theneed for t h i s type of an aly sis and re su lt s should be recognized e arl y in the program,and the overall responsibility should be assigned t o a single organization. Prepa rationshould be made to handle all interfaces h all operational modes, including failu resituations.

Based on the Apollo exper ienc e, it wa s very difficult and inefficient to developth e large, complicated computer tools at widely dispe rsed geographical locations, ondifferent computer hardware, and with sev era l groups involved.i-eijiiiring large, cijrnpkx c o i ~ p u t e r r ~ g r ~ ~ i n gd zn~!ysis ffcxts, c ~ ~ ~ ~ ! i d a t ismandatory fo r gaining efficiency in identifying and solving rec onta ct probl ems.

Fo r future act ivi ties

Lyndon B. Johnson Space CenterNational Aeronautics and Space AdministrationHouston, Texa s, October 31, 1974924-22-20-00-72

18 NASA-Lang l e y , 1 9 7 5 S-421

Documents

Apollo Experience Report Spacecraft Relative Motion and Recontact Analyses