apollo experience report guidance and control systems command and service module entry monitor...

8/8/2019 Apollo Experience Report Guidance and Control Systems Command and Service Module Entry Monitor Subsystem

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N A S A T E C H N I C A L N O T E 6

APOLLO EXPERIENCE REPORT -GUIDANCE A N D CONTROL SYSTEMS:C O M M A N D A N D SERVICE MODULEENTRY MONITOR SUBSYSTEM

Ben Reinu, Jr., and Herbert G. Patterson

Lyndon B. Johnson Space CenterHozcsto~z,Texas 77058

N A T I O N A L A E R O N A U T t C S A N D S PA C E 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 A N U A RY197


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1 Report No.

NASA TN D-7859

SERVICE MODULE ENTRY MONITOR SUBSYSTEM7 . A u t h o r b )

Ben Reina, Jr . ) and Herbert G. Pat terson

9. Performing Organization Name and Address

Lyndon B. Johnson Space Cente rHouston, Texas 77058

3. Recipient's Catalog No.I. Government Accession No.I

8. Performing Organization Report No.

JSC 5-40510. Work Uni t No,

9 1 4 -50 -1 7 -0 8 -72-11. Contract or Grant No.

13. Type of Report and Period Covered

4. Title and Subtitle I 5. Report Date

17. Key Words (Suggested by Author(s1)

Visual Displays

Loading Rate- G Force

-Trajectory Controlt

APOLLO EXPERIENCE REPO RTGUIDANCE AND CONTROL SYSTEMS: COMMAND AND

18. Distribution Starement

S TA R S u b j e c t Category: 12 .

I 6. Performing OrganLation Code

19. Security Classif. (of this report1

Unclassified Unclassified

20 . Securiry Classif. ( o f This pdge) 21 . NO. ot Pages 22 . Price

k2.756

2. Sponsoring Agency Name and Address

National Aeronautics and Space AdministrationWashington, D. C. 20546

15. Supplementary Notes

The JSC Dir ec tor waived the use of the International System of Units (SI) for this Technical Note,bec ause , in hi s judgment, the use of SI Units would impa ir the useful ness of the rep or t o r resul tin excessive cost.

The conceptual as pec ts of the command and ser vic e module entry monitor subsyste m, togetherwith an interpr etation of the displays and their as sociated relationship to entry trajectory control,are pres ente d in this document. The entry monitor subsyst em is described, and the problemsencountered during the developmental phase and the first five manned Apollo flights ar e dis-cussed in conjunction with the de sign improvem ents implemented.

16. Abstract


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APOLLO EXPERIENCE REPORT

EDITORIAL COMMITTEE

The mate ria l submitted fo r the Apollo Experience Repo rts( 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 Cen te r consisting of the following mem be rs :Scott H. Simpkinson (Chair man), Richar d R. Baldwin,Ja me s R. Bates , Wil l iam M . Bland, J r . , Aleck C. Bond,Robert P. Burt , Chr is C. Cri t zos , John M. Eggleston,E. M . Fields, Donald T. Gregory, Edward B. Hamblett, J r . ,Kenneth F . Hecht, David N . Holman (Edi tor /Secretary) ,and Carl R. Huss. The pr ime reviewer for th i s repor tw a s Chr i s C. Cri tzos .


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CONTENTS

Section

SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .SYSTEM DESCRIPTION AND DESIGN CONCEPT . . . . . . . . . . . . . . . .

Design Concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Software Concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

DEVELOPMENT AND QUALIFICATION . . . . . . . . . . . . . . . . . . . . .Prototype Hardware Development . . . . . . . . . . . . . . . . . . . . . . .Qualification Histor y . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

FLIGHT HISTORY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Apollo 7 Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Apollo 8 Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Apollo 9 Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Apollo 10 Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Apollo 11 Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

CONCLUDING REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

REFERENCE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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TABLE

Table Page

14ENTRY MONITOR SUBSYSTEM MISSION PROBLEMS . . . . . . . . .

FIGURES

Figure Page

2

3

1 Entry monitor subsystem block diagram . . . . . . . . . . . . . . . .2 Entry monitor subsystem . . . . . . . . . . . . . . . . . . . . . . . .

3 Entry monitor subsystem scr ol l pat tern s for lunar - re tu rne n t r i e s

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(a ) Lunar 3500 -nautical-mile r ange limit patte rn . . . . . . . . . . .(b) Lunar nonexi t range l imit pa t t er n. . . . . . . . . . . . . . . . . .

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HPOLLO EXPERIENCE REPORT

GUI DANCE AND CONTROL SYSTEMS:

COM MAN D AND SERVICE MODULE ENTRY MON ITOR S UBSYS TEM

B y B e n F!eina, J r. , an d He rbe r t G. Pa t te rso nLyndon B . J o h n s o n S p ac e C e n t e r

S U M M A RY

The design and development of the command and service module entry monitorsubsystem resulted in a conceptual component tha t appea red t o be capabl e of p er for mi ngfunctionally and of meeting miss ion requi reme nts . However, the conver sion of thecomponent from design to h ardware ihat would perform as designed wa s a f rus t ra t ingtask . Fr om the beginning of the entr y monitor subs ystem qualification test program,hardware problems appeared that continued throughout al l manned Apollo flights.Although some of t hese prob lems were due to poor workmanship, most wer e caused byvari ous component pa rt s not performing as designed. Even though modifications to thehardwar e wer e made throughout the Apollo Prog ram, component par t deficiencies wer enever eliminated completely.

INTRODUCTION

The e nt ry monitor subsyst em (EMS) was incor porated into the Apollo commandmodule (CM) to provide a backup mea ns of moni torin g guidance and navig ation (G&N)contro lled change in velocity (delta- V) and of providing thr us t term ina tio n to the st abi li-zation and co ntro l sy st em (SCS) controlled delta-V, as well as to provide a visual meanswhereby G&N entr y traj ect ori es could be monitored and compared with predicted t r a-jectory curves. This repo rt includes a dis cus sio n of EMS develo pment pro blem s, aswell as EMS problems encountered during the first five manned Apollo flights, and areview of the maj or modifications performed to resol ve those problems.

SYSTEM DESC RIPTIO N AND DE SIGN CONCEPT

D e s i g n C o n ce p t

The command and ser vi ce module (CSM) EMS design concept evolved fro m thesingle requirement for providing a redundant, manual capability to evaluate and contro lthe entry t rajectory (ref . 1) by mean s of visual entry-co rridor- verific ation displays.The evaluation as pe ct s wer e prim aril y concerned with monitoring the path of entr yaccording to a se t of prescri bed entry curves to determi ne whether the ensuing


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tra jec tor y would produce conditions jeopardizin g the safety of the crewme n. The con-tr ol as pects provided the knowledge required to assum e manual contr ol at any tim e andto complete a safe entry . The concept was fur the r expanded to achieve maximumtraje ctory design flexibility in the prim ary guidance, navigation, and contro l sys-tem (PGNCS) (normally autom atic) and to achieve mechanization and use simp licity inEMS design.

The design concept f ac to r s necessitated the definition of all system l imits thatcould affect crew safety in te r ms of entry performance constraints . The sys tem shaving cr i t ic al inter faces with the entr y t ra je ctory ar e pr imari ly the thermal protect ionsys tem (heat shield), the operating time-limited sys tem s, and the flight crew . Thelimi ts were defined by entry range (time) and acceleration cons train ts. Thes e con-str ain ts, the vehicle aerodynamic configuration, and the EMS resp onse -time c har act er -is tic s dictated the trajectory-s haping limit patterns on the flight-monitor scr oll . Tosha re s ens ors and electronics , the EMS display w a s mechanized to per form powered-flight functions in addition to the requir ed en tr y functions.

System description and components. - The EMS (fig. 1) provides information tothe flight crew f or monito ring the PGNCS-controlled entry performance and the PGNCS-controlled delta-V, fo r providing thrus t term ination signa ls through the SCS-controlleddelta-V, and fo r manually controlling entr y after a PGNCS fai lur e. Th e E M S a l sodisp lays to the flight crew v ery high frequency (vhf) ranginginfor mation between the CMand the lunar module. The sy ste m co ns is ts of both har dwa re and soft ware . The soft-war e a spec ts of the EMS refer to the fligh t-pa ttern limit- line generation and to develop-mental techniques and pr oced ures . The EMS hard ware in cludes the components andelectronics required to dr ive the major displays. The hardware consis ts of t h e s c r o l l(entry acceleration load fact or G and inerti al velocity V plot ter) , a roll atti tudeindicator (MI), nd a delta-V/range-to-go digital display counter.

Physically, the EMS (fig. 2) cons is ts of two basic asse mblies : the entry monitorcontrol assembly (EMCA) and the ent ry monitor s cr ol l ass emb ly (EMSA). The EMCAcontains the electronic components and is composed of inte grated ci rcu its ; circu it boa rds ;

down light

RJEC . eaction jet engine on-off control GD C= gyroscope display couplerAV velocity change S P S - service propulsion system

Figure 1. - Entry monitor subsystemblock dia gram .

a range in tegra to r ; an acce le romete r ; adelta-V/range -to-go counter logic; a pulsesc al er ; power suppl ies , re lays , and switches;an d all associated wir ing. The EMSA or G -Vplot ter assemblv consi s ts of a sc ro ll of M y-la r tape o r film imprinted with rays calledG-onset, G-offset, and range-pote ntial l ines.The s tepper motors , se rv omotors , geart ra ins , and electronic components nece ssaryto d r ive the sc ro l l are enclosed within theEMSA.

The EMS display is composed of fivefunctional components essential to entrytrajectory monitoring and flight control.The five components a r e th e MI, he entrythreshold indicator (0. 05G), the corr idorverification indic ator, the delta--V/range -to -$40 Counter, and the flight moiiitor O r Scrol l .

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F u n c t i o nselectorswitch

i d7

UPPela n n u

i n d i c a t o r / a nnu n c i a t o r n u m e r i c d i s p la y

assembly

- S c r o l lassembI(replaceable)

-F ive- po si ons ew switch

Figure 2. - Entry monitor subsystem.

Th e M I s an ins tru me nt used to display the angu lar position of the lift vecto rabout the re lativ e wind vector of the vehicle. The indicator disp lays continuous ro ta -t ion in e i the r direct ion.

The ent ry threshold indicator is a lamp that is illuminated when the vehicleencounte rs a threshold aerodynamic accelerat ion level , normally 0.05G. Thi s dis -play provides a visua l indication that the xtmosphere of the Ea rt h ha s been encounteredand indicat es the initiation of entr y. The threshold lam p is extinguished at any timeduring en try when the accelerat ion load factor falls below the p res cri bed level of 0.02G.

The co rr id or v erification indicator co nsi sts of two lamps, one of which is i l lumi-nated at a presc r ibed t ime (normal ly 1 0 seconds) after the entry threshold is r e a c h e d .The lam p i l luminated depends on the measured ac celerat ion level . The two lamps a r e

identified with the required lift orientat ion during the ini t ia l a tmospheric penetrat ionand are used t o ver i fy and, i f necessa ry , cor r ec t the o rien ta t ion .

Th e delta-V/range-to-go counter is a digi ta l display that in dicates the predictedvelocity of the sp ace cra ft. The counter is located d ire ctl y below the flight monito r.

The f l ight .monitor o r s c r o l l is the major component of the system. T h e s c r o l lp rov ides a rect i l inea r presentat ion of G as a function of V. The display is crea ted

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with a Mylar tape that ha s a monitoring pattern printed on the fro nt of 90-m illimet erfilm and a dye-encaps ulated e mulsi on bonded on the back. The tape i s contained inthe s cro l l assembly and is driven at a speed horizontally proportional to velocitychange. The ver tic al axis is driven to a position prop ortional to the acceleration loadfactor. The scr ib e removes the emulsion and cr ea te s the entry t ra jectory G-V tra ceas the tape tra nsl ate s horizontally and the scrib e moves verti cally. The tra jec tor ytrace is compared by the pilot to the permanently displayed monitoring pattern thatdefines the limiting G-onset and G-offset r a te s throughout entry.

Hardware requ irements . - In implementing the hardware requi rement s for EMSvelocity and range-to-go functions, which are the two most important calculationsperformed, a computing technique called the digital-operational method was used.Th is method comb ines some of the ch ar ac te ri sti cs of both digita l and analog methods.The overall computational concepts are those of analog oper ation al syst ems , in whichinformation flows continuously fro m one circu it to an other as received. Each circui tenti ty perfor ms a mathematical operation (such as multiplication, integration, andaddition) on the input data. The digital chara cte ris tics of the EMS res ult f rom the fac tthat the system variables ( i .e . , velocity and distance) are represented in discrete orpulse form. The dis cre te nature of these var iab les allows the use of a hardware mech-

anization with integrated circ uitr y and thus redu ces siz e, weight, and powerconsumption.

Sof twareConcept

Description and concept. - The EMS provides a rec tili nea r display of the totalaerod ynamic -acce leat io n load factor as a function of t he i ne rt ia l velocity magnitudeduring the entry flight phase. The display mechanization provides a trace of actualG and V and famili es of pr edete rmine d G -V pro file s that a r e used by the pilot toassess the chara cte r of the entry traj ect ory on the basis of the cr i te r i a used to developthe predetermined G-V profiles. The G-V profi les are not t r a jec to r ie s nor a r e the

profiles intended to repre sen t traje cto rie s, even though it may be possible in someinstances to maneuver the vehicle s o that the a ctual G-V conditions correspond to thoseof a given profile. The profiles actually repr es en t the limiting ra te a t which the valueof G may be inc rea sin g o r dec rea sin g at a par tic ula r velocity and the G conditionfor which the pilot. may maneuver the vehicle and avoid violating the criteria used todeve lop the slope of the profi le. A famil y of pro file s is used to present the limitingslo pes because the profil es can be conveniently and meaningfully inte rpr ete d by thepilot. There fore, the required information can be presented with two easily obtainedvariables , G and V, without res ort ing to the genera tion of accel eratio n ra te s.

Flight -pattern limit -line generation and developmental cr ite ri a. - The proceduresa r e outlined for the thre e types of flight l imits incor porat ed into the EMS softwa re.

flight l imi ts and the associated pr ocedu res f or the ir development may be generated toany criterion that can be defined as a function of velocity and accel era tion vari ab le s.The procedures presented herein per ta in only to cr i t er ia that have been specif ical lydefined fo r the Apollo lunar mis sio ns. The E M S sc r o l l p a t t e rns f o r Apollo lunar-retur n entr ie s a r e shown in f igures 3(a) and 3(b) .

Th e

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(a) Lunar 3500-nautical-mile rang e limit patte rn.

r R a n g eguidel ines G-onset l ines-

LNonexit Orb i t a lvelocity i n d i c a t o r 1 Velocity, Wsec x 13l i m i t l i n e s

(b) Lunar nonexit range limit pattern.

Figure 3. - Entry monitor subsystem scro l l pat terns for lunar-return entr i es .

G-onset l imi ts: The cri ter ion for the G-onset flight l imi ts is to not exceed aspecified maximum G-level. Specifically, the dG/dV li mi ts mus t be det erm ine d a s afun cti on of G and V at the most a dve rse vehicle lift atti tude (full negative lift , wherelift @ = 180") so that, when this condition is reached, the pilot can delay for 2 secondsand then initiate a maneuver to the best vehicle lift attitude ( f u l l positive lift, @ = 0").The ensuing t ra jectory then w i l l result in a peak G of 10. It is not necess ary to com-plete the maneuver to the best lift attitude when the pullout a t 10G is achieved.10G does not r epr ese nt a rigid limit ( i . e . , devi ation s of 10 to 15 percent above 10Gare toler able) , the analy sis model conside rs only off -nominal (not acceptable) devia-t ions in the most sensi t ive para meter s . The par ame ter s are the vehicle lift-to-dragratio , initial atti tude, and maneuver response. A l l other pa ramet e rs a r e des igna tedas acceptable and are a zero-inclination conic (latitude = 0", azimuth = 90" ) and the1962 U.S. Standard Atmosphere . In general, the procedure used to ar ri ve at thelimiting flight conditions is to work backward from the known ter min al c ri te ri a (in thisca se , 10G) to determine the acceptable ini tia l condit ions that a r e the f l ight l imits .

Because

Nonexit G-offset l imits : The crite rion fo r the nonexit G-offset l i mit s is to main-ta in t r a jecto ry condit ions that do not result in exiting fr om the atmosphere, as definedby a selected minimum G-level (G = 0.20), o r do not exceed a specified maximumrange .G-onset l im its in that the limiting dG/dV rat io as a funct ion of G and V must be

The requ i rements fo r meet ing th is c r i t e r io n are sim ilar to those needed fo r the

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determined.($ = O"), and the maneuver is ma de t o the best lift attitude (+ = 180"; full negative l i f t )incorporat ing the sa me pi lot-response and vehicle - respons e requ ireme nts .f o r the nonexit G-offset l imit, the cr i te r ia are mor e cr i t ic al , and the analysis modelmust be constructed s o that a wors t -case mode l res u l t s .

In this case , the most adver se vehicle lift attitude is ful l posi t ive lift

However,

The procedure f or obtaining the f l ight l i mit s fo r the G-offset l imi ts is to workbackward from the defined terminal conditions to determine acceptable initial conditions.The specified minimum G-level is used to define a termi nal condi t ion in the sa me man-n e r as the maximum G-limit wa s used to define an initial condition. However, thecr i t e r i a inc lude a range l imit , and this l imit does not provide a defined te rmin al con-dition.limit criteria and with other flight safety cr i te r ia is generated.

To include the ran ge cri teria, a limiting G-V profile that i s compatible with the

Maximum range G-offset l imits : The cr i te r io n fo r the maximum range G-offsetl i m i t s is to maintain traje ctor y conditions that do not exceed a specified maximumentr y range. Unlike the nonexi t l imits , thi s cr i te r ion wil l al low an atmospheric exitfollowing initial entr y. The flight l im its must be deter mined as a func tio n of G andV such that ran ges up to the specified maximum can be obtained consistently and suchthat no rang es in ex ce ss of thi s value can be achieved. Thi s se t of flight l im its incor-por ate s the limiting G-V profile as defined by st ep s 1 t o 4 of th e nonexit pat ter n and,in fact, may be the sa me profile as long as the range l imit , the analysis model , andthe flight mode are the same as thos e of the nonexi t pat ter n.l imi ts generated fr om the l imit ing prof ile are determined in exact ly the sa me manne ras those generated from the nonexi t l imits . Th is generat ion includes f l ight l imit s thatoccur at G-levels less than those of th e limitin g profile.

In addition, all flight

The only difference between the methods used t o obtain the Go ff se t l i mits andthose used to obtain the nonexit l im its oc cu rs in defining the acc eptable ter mi na l con-di t ions at G = 0 .20 fo r veloci ties less than the velocity a t which the limiting profilebeco mes tangent to G = 0 . 2 0 . Fo r the nonex it l imi t s , G = 0.20 wa s the lowest allow-able load; consequently, the ter min al conditions could be obtained direct ly. F or themaximum range G-offset l imits , the range mu st be constrained to a maximum magni-tude; consequently, the allowable conditions at G = 0.20 mus t be compatib le withth i s range.

To dete rmin e the allowable conditions, the ma nner in which the tra jec tor y is t obe shaped throughout the entry flight must be spe cified. The exact t ra jectory shaping,o r f l ight mode selected, is par t of the design model. The flight mode sele cted is basedon a constant G-level during the supe rcirc ular f l ight region to a velocity a t which fullpositive lift (G = 0 . 2 0 ) is implemented and maintained to exit. Following exit, a z e r o -lift atti tude ($ = 9 0" ) is maintained to an altitude of 25 000 f e e t .

The t ra jectory data to def ine acceptable termi nal condi t ions are separated intofour g roups as follows.

1. The l imiting t ra jecto ry prof ile

2 . The maximum trajec tory

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3. F u l l positive lift, constant-G trajectories

4. Limiting profile, constant-G trajectories

The maximum trajector y def ines (1) he minimum exit velocity from which the rangelimit can be achieved and (2) the corresponding maximum G-level that can be encoun-tered and the range limit that can sti l l be achieved. The maximum trajec tor y se ts thelower velocity limit f or which te rm in al conditions must be defined. The lower velocityand the velocity at which the limiting trajec tory be comes tangent to G = 0.20 set thetwo extr eme s in exit conditions.

Verification and simulation of the flight limit lin es and range guidelines. - Proce-d u r a l and simulat ion exami natio ns designed to verif y and evaluate the adequacy of theEMS flight limit lin es and the range guidelines have been perfo rme d. The purpose ofthe procedural examination was to verify the proce dure s and dat a that r elated to thederivation of the E M S flight limit lin es and range guidelines fo r the two types of lunarsc ro ll patterns (3500-nautical-mile range limit pattern and nonexit range limit pattern).All derivation procedures conformed to al l other exis t ing procedures . Data werechecked to en su re consi stenc y throughout the development of ea ch se t of lines . Onee r r o r wa s found in the development of the 75-nautical-mile range guideline. The er r o rcaused this guideline to rep resen t a range that was incorr ect by 2 to 4 naut ical milesfo r G-levels between 2 . 5 and 3 . 0 and between 7.0 and 8.0. However, in an actua lent ry, the e r r o r developed at the 75-nautical-mile range guideline would be eliminatedas the G-V trac e approached the 50-nautical-mile ran ge guideline; thus, the er r o ractually would not r epr ese nt a se ri ou s compr omise with the EMS backup rangingcapability. N o other er r o r s or inconsis tencies were found. The f l ight l imit l ines andthe r ange guidelines, with the one exception noted, wer e co rre ctly developed and werevalid for all applicable flight conditions for Apollo entries.

Two independent manned sim ulati ons we re per form ed to veri fy and evaluate the

flight l imit l ines and the range guidelines on the two lunar sc ro ll pat ter ns. The purposeof the f i r s t s imulation w a s to evaluate the G-level and the range attained on the sc ro llpat ter ns of the design models afte r a violation of the s cr ol l pattern had occur red.pr ima ry intent was to evaluate the human aspects, as opposed to exact nu meric alre su lt s that could be handled digitally.

Th e

The results of the first simulation indicated that the G-onset and G-offset lineson both sc r o l l pa t t e rns a re valid for all cas es in which the pilot delay time in recog -nizing a violation is less than the d esi gn value of 2 seconds. Simulation results a l soindic ated that the EMS monitorin g ta sk of judging the sl op es of two li ne s (one of whichis dynam ic) within 2 seco nds of the a ctu al tangency is a feasible pilot task but req uir esclose attention and a cr iti ca l pilot atti tude.

The second simulation was performed to evaluate the range guidelines and vari -ous backup ranging techniques that could be used in the e vent of a PGNCS malfunction.Re su lt s of the study indicated that use of the E M S range guidelines provides the bestmethod fo r backup ranging.

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The r es ul ts of th e EMS flight limit line and r ange guideline veri ficat ion andsimulation s tudies are summarized as follows.

1. The EMS G-onset l imit lines , the EMS 3500-nautical-mile rang e limi t lines,and the EMS skip limit l ines were co rre ctl y developed. There fore, the scr ol l patter nsinvestigated can be used to monitor PGNCS performance during all Apollo entries.

2 . The E M S rang e guidelines, with one exception, w er e cor re ctl y developed.The one exception can be accounted fo r durin g manual ranging so that, in the event ofa PGNCS malfunction durin g entry , the r ange guidelines can be used in backup ranging.

3 . The EMS cri ter ion fo r manual takeover is based on the pilot's evaluation ofa tangency condition between the G-V tr ac e and the flight limit lines .

DEVELOPMENT AND QUAL1F ICAT1O i l

Pro to type Hardware Deve lopmen t

Complete system development and compatibility testing w a s accomplished on aprototype EMS breadboard assembled on tw o chassis with an independently mountedscro l l a ssembly.closed-loop sy st em perfo rmanc e in t e r m s of magnitude of unwanted signals and inte r m s of EMS susceptibility to externally generated signa ls and to internally ra diatedsignals i n the immediate environment fo r a wide rang e of operat ing conditions. Theprinc ipal benefits of the test ing wer e confirmation of filter design optimization anddemonstration of acceptable e lectr ical noise and t ransient suppression.

The test ing enabled evalua tion of comp onent compat ibilit y and

Q u a l i f i c a t i o nHistory

Qualification test prog ram. - Much difficulty occurred with the EMS during thefi rs t attempt at qualification testing. The original qualification te st was abortedbecause of repeated fa ilur es encountered durin g the vibration and tempe rat ure t es ts .The test unit was modified and the tes ts re peate d. The modifications wer e directedat what were considered the pr ima ry cau ses of fai lures: intermit tences in the el ec-t ronic components package and scr ol l assembly electroluminescent short c i r cui tsto ground.

After the minor modification period, the qualification te s ts were attempte d;however, EMS fa ilu re s continued. The fai lur es recor ded du ring the performa nce of thetes t fe ll mainly into four cat egori es.

1. Problems with the s cro l l assembly

2 . Susceptibility to humidity

3 . Intermittent connections on the ter min al boa rd s beca use of vibration

4. Repeated prob lems with the elec trolum inesc ent lighting panels

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The most severe problems were associated with the scroll assembly and the RAI e lec-trolu minesc ent lighting. Much redesi gn w a s required for solution of these pro blems.Fo r example, elimination of the electroluminescent lights on the sc rol l assembly andon the RAI requir ed the addition of incandescent lamps .

Delta-qualification design and proof test program. - One production EMS (qualifi-cation tes t unit 2) wa s subjected to the full ser ie s of qualification te st s to cert ify theflightworthiness of EMS design impro vement changes that had been incorporat ed a fte rthe original EMS qualification tests. Only two of t he cha nge s we re signi fican t:(1) he use of incandescent lighting for the scr oll and fo r the R A I and (2) the improved(mo re flexible) initialization of the range-to-go int egr ato r. Numero us fai lu re s we reencountered during the delta-qualification design and proof t es ts ; however, the onlyenvironment that adversely affected the EMS w a s the salt-fog and combined oxygen-humidity porti ons of t e s t s with cor ros ive contaminants, oxygen, and humidity (CCOH).

Problems of minor significance, such a s e lec t r ica l overs t ress ing of some EMScomponents during the insulation resist ance tes ts per formed aft er the environmenttes ts , wer e resolved by modifying te st equipment and proced ures.

fai lur e to sc r ibe and a V- a x i s failure during therma l vacuum testing were cor rec tedby verifying the co rr ec t stylus spring for ce and by checking the film spool interferenc ebetween the r ol l of fi lm and the housing.

Such conditions as

Severa l problem areas wer e encountered that wer e the di re ct res ul t of CCOH.The qualification te st s revealed moist ure pro blems around the hermetically sealeddelta-V/range-to-go indic ator (causing dimming and flick ering of the elect rolumi nescen tsegment s) and around the potting mater ial enclosing the wire he ade r to the hermetica llysealed function switch. These situations were corr ect ed by a design change in whichthe delta-V/range-to-go display assemb ly w as epoxy-bonded to the front panel and byimprov ements in manufacturing proce ss c ontrol s and in inspection to ensure compati-bility with humidity. No other significant problems wer e discovered in tes ts on the

re ma in de r of the EMS compo nents .

Flammability tests. - Flammability tests were performed on a mockup unit con-taining E M S components made of mat eria l generically si mil ar to that used in the actualspac ecra ft configuration. The purpose of the te st s was to determi ne the str uc tu ralflammability cha ra cte ri sti cs of the EMS in 6 . 2 - and 16.5-psia oxygen environments.The tests were t o determ ine whether, in the event of an internal fir e, the EMS com-ponent uni ts would rupt ure and thus pe rmi t the s cat ter ing of flami ng contents and wouldemit f la me s through vents and thereby propagate the flam es to nearby nonmetallicm a t e r i a l s .

The flammability tes ts demonstrated that the EMS would not propagate flam es a sa result of an internal f i r e . However, minor material changes were required to meetnew Apollo flight mat eri al selection crite ria. A pott ing compound on one of t he EMScables was covered with an aluminum shield. A silicone heat-shrink sle eve wasremoved from a cable. Polyamidespot-t ies wer e replac ed with Beta cloth spot-ties. The function sel ecto r switch knobmade of tenite was re placed with a knob made of al umi num .

The nylon sleeving on a dust cover was removed.

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Centrifuge tes ts. - Throughout the original qualification test program and earlyproduction phases, one design deficiency wa s found that requi red majo r re design:The s cr ol l assembl y and the RAI electr olumine scent lights did not provide sufficientillumination to observe the sc roll . In addition, the electr olumine scent t abs aroundthe s cro ll and the RAI were susceptible to moistu re, and this susceptibility causednu me rou s e le c r olumine scent failure s .

To implement a new design for the scroll and RAI lighting, a centrifuge test was

The EMS usedperformed at the NASA Lyndon B. Johnson Space Cen ter (JSC) (f orm er ly the MannedSpacecraft Center (MSC)) to es tablish the exact lighting req uirement s.was a modified engineering prototype with incandescent lights for sc ro ll and RAI illumi-nation. The sc ro ll pat ter ns used with the EMS we re approved for Apollo miss ionorbital and lunar profiles. The subjects included thre e engineers from MSC, twoastrona uts, and thre e test pilots employed by the contra ctor.observed the EMS sc ro ll and functions throughout the G-profile.

Each test subject

The test r esu lt s were determined from the c omments of the subjects during andafter the test r uns and from the subjects ' abilit ies to detect traject ory violations duringthe runs . The maximum G-level reached in these tes ts w a s 9.5G.

In all cas es , the subjects found that immediately after shifting thei r ey es fromthe entr y tr ac e to another part of the instr umen t panel, it w a s difficult to visually r e -acquire the t rac e. Each subject recommended use of a colored t rac e. The corr idorverification indicator w a s sufficiently bright that no proble m exi ste d under any of th eacc eler ati on conditions. The 0.05G light wa s found to be too bright when viewed with-out the filt er. A l l subjects recommended that a fil ter be used. The subject s reco m-mended that the number s on the range-potential lin es be placed in a vert ical columnins tea d of in a staggered arrange ment on the scro ll pattern. The G-scale on the leftside of the display was al so found to be too wide; it wa s obs tru cti ng too much of t heu s e f u l t r ace .

The principal benefits of the centrifuge t est ing were the confirmation of a day-glow orange background to improv e identi ficati on of th e tr ac e, a narrower G-scale ,an optical filter design for the 0.05G and service propulsion system (SPS) thrust lights,and the vert ica l placement of the range-p otenti al guideline num ber s. In gen eral , thechanges made as a res ult of the centrifuge tes t prog ram provided an improved flight-worthy scroll assembly.

Delta-vibration qualification tests. - The EMS wa s subjected to a to ta l of 10 min -2

Utes of r andom v ibr ati on at 0 .0 6 7 g /Hz with no anomalies occurr ing during or af terthe delta-vibration tes t. System perform ance dur ing prefunctional and postfunctionaltesting w a s normal except f or an identical failure occu rrin g in both functional te sts .The s tyl us tip penetrated the Mylar film to the extent that the G- and V- a x i s d r i v e rbecame inoperative. Analysis indicated that the sty lus penetration resulted fr omexcessiv e retrac ing of the scr ibe line, with other fa cto rs contributing to the penetr a-tion.anomaly was not considered to be environmentally induced.delta-vibration t es t provided additional confidence that the ac ceptan ce vibration tes t

level of 0.04 g /Hz does not degrade the flight hardwa re.

Because the re were no indications of pene tratio n during the vibrati on cycle, theTh e com ple tio n of the

2

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Design and proof tests for vhf ranging. - Design and proof te st s wer e also pe r-formed to cer tify t he fligh tworth iness of the EMS unit modified to inco rpo rat e vhfrangi ng within the EMCA, and of th e improved he ad er seal within the EMSA. Th emodified e lem ent s successfully completed all t e s t s .in a r e a s not associated with the modifications. However, no problem wa s attributedto the environments imposed.

Some fai lures were encountered

A hardwar e failure wa s detected before the random vibration; the G-axis ser vo-motor (styl us drive ) hesitated approximakely 3 0 to 60 seconds before responding anddriving from the 9G level to 0.22G. Subsequent ana lys is of the fai lur e establish edthat it wa s caused by a frac tur ed solder joint that had apparently been damaged duringthe modification of th e EMS qualification te st unit.

During the combined high- temper ature and vacuum environment, the EMSAsc ro ll dr iv e failed. Disas sembl y of the EMSA reveal ed that the inter nal fro nt platehad not been relief-milled (as specified) to provide space for the orange tape thatwa s added to produce a "colored" G-V tra ce. The cleara nce between the internal frontplate and the enclo sur e was insufficient, and the buildup of scr ibe -co at residue caused

fri cti ona l fo rc es beyond the capabili ties of the V- a x i s stepper motor.

Severa l fail ure s caused by ar. r r o r in the qualification test procedure and bynoise pickup of a wire within test-equipment cabling wer e encountered. Test pro-cedure changes wer e implemented to prevent the problem from recu rri ng. After thenoise-sensitive wire w a s rerouted within the test-equipment cable, pro per operationof t he EMS wa s verif ied . N o fai lur es were encountered eit her in the new header sealo r in any of the ci rcu it s that we re added or modified by incorporation of the v h f -ranging mode.

Microencapsulated-dye-coated scro ll engineering development tes ts . - Engineer-ing development te st s were performed to evaluate a microencapsulated-dye-coated

EMS sc ro ll . One of the most significant fai lur es encountered with the previous sc ro llass embl y occurr ed during the Apollo 9 and Apollo 10 flights.pen etr ate the emu lsion backing of the sc ro ll patter n and thus failed to produce a visi-ble t race.

The G-stylus failed to

Severa l fai lur es we re encountered with the originally used scri be coat through-out the qualification test phases.coat formula, to the scri be-coa t application technique, to defective stylus springs ,and to impro per s tylus t ips . For each failure, some form of cor rect ive action wasimplemented such as using sharp er t ips , increasing the spring force, improving thefor mula with the elimination of ur ea fro m the s cr ib e coat, and using different purg-ing proced ures. Although these minor correctio ns seemed to indicate a tentativesolution to the scribecoat problems, the contra ctor w a s directed t o investigate thescr ib e coat thoroughly. Ar ea s investigated included, but were not nec essa ril y limitedto, the following.

Some of th e failures were attributed to the scribe-

1. Basic pro ces s changes of the scrib e coat

2 . Length of bake and nu mb er of bake cyc les

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3 . Exposure time to air durin g handling and installation

4 . Effect and control of ambient humidity in the handling areas

5 . Measure o r indication of scri be-coa t har dnes s

6. Tim e limi ts on expos ure to vacuum, heat, and humidity

The scribe-coat investigation was terminated p rematu rely by the contra ctor becausehe thought that no greater improvement could be achieved with the existing scribecoat.

Concurrently with the scr ibe-co at investigation, a search for an a l t e rnatescrib ing concept w a s begun. A survey of the industry was performed , and a state-ment of work w a s awarded to develop a mor e functional and rel iab le type of r eco rdi ngassembly. The study w a s designed to integrate and evaluate available technology thatw a s related to the scroll recorder tape requirements and that could be introduced intothe existing flight scro ll assembl y with minimum hardware modifications. Thestudi es were performed sequentially : dye sel ection and evaluation, encapsulation ofselec ted dyes, formul ation of emulsion coatings, development of coati ng techniques,and envir onmen tal tes tin g of coated Myla r ta pe s.

In the beginn ing of th e EMS scro l l program, contractor conferences providedinformation and dye materials that enabled proceeding to the encapsulation studyimmediately. Encapsulation studi es were perfor med on each of six dyes. Becausesom e of the dy es were used in contr actor -furni shed paper and had been extensivelyinvestigated, encapsulation presen ted no probl ems.

The capsule coatings are self-contained capsule sy st em s; that i s , the encap-sulated dye is fi rs t applied to the Mylar s ubs trat e and then overcoated with a phenolicres in dispersion. On pres sure act ivat ion of the dry coating, the encapsulated dye isreleas ed and immediately re ac ts with the acidic phenolic r es in to develop a coloredprint.

A production EMS scroll assembly was powered by bench test equipment thatpermitted evaluation under laboratory conditions, lighted and unlighted, of variousscro l l t r aces using a coated scro ll. The existing scrib ing method w a s compared tothe contra ctor method. The con tra cto r method did provide a usable t rac e under theconditions demonstrated; however, the trac e w a s pale red and did not contrast w e l l .A da rk e r color of dye was used to improve the contrast. The new tr ace w a s a l soslightly narrower: 0.020 to 0 . 0 2 5 inch compared to 0.030 inch. A slightly widerscr ibe was developed as a result of this test.

On June 5, 1969, two contractor pilots and two astronauts participated in acentrifuge evaluation of a scr ol l incorporat ing the previously discussed changes. Eachpilot made two runs in the centrifuge, the fir st a t 5 . 8 G and the second at 9.4G; thelat ter w a s a maximum G-profile.the t raceb ackgrou nd contrast , and a reshaped stylus tip that wa s used produced a

The dark-red dye used in the new scro l l improved

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slightly wider trace. Based on the successful tes ts , it wa s concluded that thecontractor-developed method provided a visible and readable trace under normal-Gand high-G e ntr y conditions.

A f t e r the centr i fuge tes ts , engineering development te s t s were performed b y thecontractor during June and J u l y of 1969 . Both the EMSA and individual segments of

the microencapsulated-dye-coated sc ro ll s were subjected to environments that mightadver sely affect them, and the EMSA and s cro ll seg ments performed successfully inal l the environments to which they were exposed. A few anomal ies occurred duringthe testing; however, the anomalies eit he r could not be attributed to the environmentsimposed o r wer e encountered only in off-limit conditions. The succe ssful completionof the design verification tes t program demonstrated that the coated sc ro ll s werecompatible with the selected Apollo environments. N o fa i lu res o r anomal ies wereencountered that would prevent the coated s cro ll s from producing an acceptable ent ryt r a c e .

FLIGHTH ISTORY

In eac h of th e first five manned Apollo flig hts, se ve ra l malfunctions of t he EMSocc urr ed. Many of the anomalies recorded wer e hard ware fai lur es, w her eas oth erswere c lass i f ied as operat ional pro cedures or equipment-operation pecul iarit ies. Thesignificant fact about these failures is that none of the m oc cur red du ring the qualifi-cation tests. A combination of mar gin al design conditions and poor workmans hipaccounted fo r the majority of the fail ure s.

A critical design review of the EMS w a s conducted after the malfunctions hadbeen encountered. Sever al major design improvements wer e implemented to improveEMS operation and to in crea se EMS reliabili ty. In addition to making hardwareimprovements , ex tr a car e during the modif icat ion cycle w a s requested of qualitycontr ol and manufacturing personnel. Despite the extr a precaution taken to ens uretrou ble-fr ee ha rdwar e, additional minor prob lems we re encountered with the redeli-vered units. Table I is a summa ry of the overa ll probl ems encountered during theApollo 7 to 11 missions .

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TABLE I. - ENTRY MONITOR SUBSYSTEM MISSION PROBLEMS

I

Mission

Apollo 7 (CSM 101)

~~~

Prob lem

An imprope r scr ibe driv e unit (EMCA) was replaced.

The replacement unit had ground and flight problems.

1. Range -to -go function

2. Delta -V/range -to -go co un ter j ump s of the"90 000" digit

The delta-V/range-to-go coun ter jumped 100.4 f t /secat Saturn IVB separation.

The midcourse correction (MCC) 5 velocity w a s notmeasured properly by the EMS.

The delta-V/range -to -go cou nter jumped du ringpost-MCC -5 t roublesho oting.

Scrol l scr ibe per turbat ions occu rred dur ing entry.

No G-V trace scr ibing w a s obtained during entry.

Intermit tent scr ibing was obtained during preentrytes ts .

Electroluminescent segmen t fa i lure oc curre d in thedelta -V/range -to -go numeric display counter.

Apollo 9 (CSM 104)

Apollo 10 (CSM 106)

Apollo 11 (CSM 107)

Apol lo7 Mission

Performance. - Both the delta-V and the range-to-go counter cir cu it s in theEM S malfunctioned bef ore lift-off of the Apollo 7 CSM. The fi rs t failu re involvedthe range counter during a self - tes t . In th i s t e s t , the coun ter is preset , then countsdown in response to a known s tim ul us fo r a specified time, and finally rea che s avalue near 0 miles. The syste m repeatedly failed the self-t est , both before flight

and in flight.

A delta-V/range-to-go coun ter malfunction, total ly independent of the rang ecounter failure, was noted during the prelaunch setup of t he delta-V/range-to -gocounter j u s t before lift-off. A "9" appeared i n the most significant digit of the co unterwhen a crewman switched the function se lec tor to the "delta-V set " position; how-ever, the operation w a s normal dur ing a r e p e a t of the procedure. The malfunction

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occur red se ver al t i me s during flight and, in all but one instance, coincided with aswitching operation.

Postflight testing and an alysi s indicated that the m ost probable cause of t he"90 000" digit-jump anomaly was a high-resistance cr im p joint on wire 18 of thedelta-V/range-to-go display. The high resist ance caused a voltage dr op of ap prox i-

mately 0.7 volt dir ect cu rre nt (dc) in the 4-volt dc input l ine. The voltage-drop levelis near the operatio nal threshold of the circuit .

During postflight testing, the range integrator malfunction w a s verified andrep eate d. However, du ring tem per atu re testing in the investigati on of the "90 000"digit-jump anomaly, t he range integrat or began functioning prop erly, and the malfunc-tion could not be induced again.

The most probable cause of the range integrator anomaly was the shor t c i rc ui tk gof welded lea ds in the range integ rator submodule. The sho rt circuiting caused thevelocity reg is te r to be latched at 37 000 ft/sec and, t hereb y, induced a 2.2-nautical-mile er r o r in the entry moni tor se l f - tes t and a gr os s e r r o r in r ang ing dur ing en t ry.

Postm issio n modification. - The postmission corrective action implemented r e -quired sold erin g of all wir e crim p joints in the delta-V/range-to-go counter and pe r-for min g of low-voltage m arg in tests of the counter-int egrated cir cui t module. Athermal cycl ing screening tes t w a s als o instituted as pa rt of acceptance testin g.

Entry postflight results. - Postfl ight assessm ent disclosed that the an omaliesencountered p revented accomplishment of some of the detailed flight te st obj ectives.The ranging accu racy could not be asse ssed because the range-to-go display did notopera te cor rec t ly as had been anticipated. The G-V was displayed by the ent ry moni-to r properly. The EMS G-V flight scr ol l tr ace correl ated well with the PGNCS tr ac eflown duri ng the entry trajecto ry. The velocity e r r o r of 286 ft/sec at 11 500 ft/se caft er 0.05G was within both the hardware and the total delta-V e r r o r budget of300 ft/sec . The total velocity e r r o r of 52 ft/sec at 4000 ft/sec w a s again less than the97-ft/sec budgeted delta-V er r o r. Despite the range integ rator anomalies, the EMSdemonstrated prop er sc ro ll operation and delta-V control during powered flight andentry.

Apollo 8 Mission

Performance. - The following performance anomalies in CSM 103 wer e observedby the Apollo 8 l i ight cr ew during the mission.

1. A -100.4-ft/ sec jump occ urr ed in the delta-V/range-to-go counter display atthe ti me of CSM sep ara tio n fr om the Saturn N B . (Postflight crew debriefing confirmedthat the jump actually was -100.4 ft/sec. )

2. A -6 .2-f t /sec e r r o r occurred in the del ta-V measured dur ing midcourseco rr ect io n (MCC) 5. The actual MCC-5 delta-V w a s -5.0 ft/sec, and the EMCA e r r o rshould have been 0.7 ft /sec maximum, as stated in the specification requi rement s.

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3 . A higher than normal EMCA acceleration drift rate occurred immediatelyfollo wing comp letio n of MCC-5. A counter display change fro m -6 .2 t o -6.9 f t / sec wasobse rved by the flight crew imme diate ly following MCC-5, jus t before the EMS wa sturned off.

4 . Erroneous fast counts as high a s -20 ft/sec occurred when the EMS modeswitch was cycled fr o m "standby" to "auto. I' The counts occ urre d during post-MCC-5

troubleshootin g performed by the flight cre w. The magnitude of the count inc rea sedwhen att emp ts wer e made to go through the switc h-clo sure points slowly durin gactuation.

5. Two negative acce lerat ion pulses occ urr ed on the EMSA scroll pattern dur-ing entry.

Postflight testing. - After flight, the EMSA passed a visual inspection and a com-plete functional te st , and no anomalie s we re noted. During functional testi ng of theEMS, including the EMCA, er ra ti c acce ler ome ter operatio n was noted.ometer null output w a s unstable for a period after the EMCA accel eratio n input axiswas placed on the horizontal.

The acceler-

Transient s ignals on the accel eromet er analog dc output were a lso observed.The transient s occur red only during the t i me following a repositioning of the ac ce le r-ome ter input axis between zero -g and approxim ately one-g conditions. Occur rence ofthe tr ans ien ts depended on allowing sufficient stabilization time before repositioning theaccelerometer input axis.

The linear acceleration ineasurement unit (LAMU) was physically removed f r omthe EMCA, and the sa me tra nsi ent behavior wa s produced by physically moving onlythe LAMU. During these t es ts , the LAMU output signals were disconnected from theEMCA electronic components and monitored to en sur e that the anomalies we re theres ul t of a n LAMU malfunction ra th er than a problem within the EMCA electronic com-ponents. Thi s EMCA disconnectio n isolated the probl em to the LAMU, and the LAMUwas subsequently returned to the vendor for fur the r analys is and te st . During vendortest ing, the LAMU continued to exhibit ext ran eou s outputs, which we re thought to becaused by an air bubble in the accele rome ter sen sor . Before the LAMU wa s opened,the validity of the herme tic se al wa s verified by leak te st . When the unit was opened,oil w a s found around the sensor /elec tronic unit dust cover. The leak path was trace dto the O-ring se al s under the two oil fill/seal sc rews on the sensor cover. The sensorw a s drained of oil , refil led, res eal ed with new O-ring s and se al s, and ret est ed. Thema s s of oil removed indicated that approximatel y 1 cubic cen tim ete r of oil had leakedfr om the unit. Retest verified prop er acc ele rom ete r operation and confirmed thehypothesis that the air bubble caused the extraneous acceleration outputs.

Postflight ground testing correlated the spurious accelerometer output to the

delta-V/range-to-go counter in the delta-V mode and to the G-V t ra ce in the ent rymode. The spikes on the G-V tr ac e wer e not of the sa me magnitude as the spikesencountered during the mission ; however, ground testi ng indicated that the air bubblehad varying magnitudes of effect, primarily a s a function of the history of rotationalmotion and acceleration inputs.

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Extensive diagnostic testing was performed in an effort t o duplicate and isolatethe spuriou s -100.4-ft/sec velocity that appeared in the delta-V/range-to-go displayduring Saturn IV B separation, as well as the spurious counting that occurred duringslow cycling of th e mode swi tch. With a positive acceleration, the delta-V/range -to-go display counts down if an accelerat ion r ever sal (negat ive accelerat ion) occ urs asthe display reache s 0 . 0 ft/sec and if the display jumps to -99.2 ft/sec. Thi s phenom-

enon was repeated by using both the sine torque and the ti l t tab le fo r accelerationr e v e r s a l s . T h i s t e s t w a s a l s o run on prototype unit 1, and display jumps to -9.2 ft/secas w e l l a s t o -99. 2 ft/ sec we re encountered. The display jump was caused by a logicrace that is inhe rent to the design of the delta-V/rang e-to-go disp lay.

The postflight testin g and an alysi s result ed in the identifi cation of two specificproblems within the Apollo 8 EMS. The first, air bubbles in the accele rometer , isconsidered an isolated case ; the second, the logic race, is inherent to the EMS designbut may o r may not exist in all units, depending on component ch ar act eri st ics ofindividual units. The manner in which these probl ems wer e reflected in the EMS pe r-formance dur ing the miss ion is specified as follows fo r each flight anomaly.

1. A -100.4-ft/s ec jump in the delta-V/range-to-go counter display occur redat the ti me of CSM separ ati on fro m the Saturn IVB.duplicated; however, the logic rac e, which w a s at least the majo r contr ibutor to thee r r o r , w a s repeated many t ime s dur ing tes t ing. During these tes ts , the displayconsistently jumped to -99.2 ft /sec; however, the te st s wer e performed with theApollo 7 (CSM 101) acce ler ome ter ra th er than with the Apollo 8 (CSM 103) acc el er -ome ter. In addition, circu it analy sis indicated that the logic ra ce would jus t asl ikely cause a display jump to -100.2 ft/sec ( the differenc e of 0.2 ft/ sec being likelyto have mission segments). Th is jump, coupled with the nor mal ac celerat ions duringSaturn IVB separation, could cr eat e the -100.4-ft/sec display. An alternative andequally probable cause of t hi s anomaly would be the comb ination of th e two sys te mproblems. Where as the logic race caused the display to jump fro m 0 . 0 to -99.2 ft/sec,the diff erenc e of 1 . 2 ft/sec was the resu lt of both natural accele ratio ns associatedwith the separ ation and s pur iou s counts caused by the air bubbles.

This anomaly w a s never exact ly

2. A -6 .2-f t /sec e r r o r occurred in the del ta-V measured dur ing MCC-5.Thi s anomaly can be direct ly a t t r ibuted to the accelero meter air bubble.

3. A higher than normal EMS acceleration drift rate occurred immediatelyfollowing MCC-5.

4. Erroneous fast counts as high as -20 ft/sec we re observed when the modeswitch was cycled from "standby" to "auto.troubleshooting performed by the flight crew.

The counts occu rre d durin g post-MCC-5

5.during entry.

Two negative acceleration (+G) pulses on the EMS G-V tr ace occur red

Ite ms 3, 4, and 5 can be direc tly attributed to the accel ero mete r air bubble.The sp ur io us counting wa s caused by the res idu al effect of MCC-5 on the accelerome-t e r. T h e air bubbles wer e dis tor ted as a function of ac celera tion and accel ero met er

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int ern al geometry; following accel eratio n, re tur n of the bubble to it s minimized pre s-s ur e shape damped fluid motion and resulted in spurious accele rome ter outputs. Theacce le romete r air bubble is considered to be the total cause; slow cycling of the modeswitc h simply inhibited th e dis pla y in the standby mode and allowed the disp lay ofspurious counts when in "auto. "

he foregoing conclusions were based on extensive testing and analytical effort

after the Apollo 8 mission. The proble ms that we re found resulted in positive co rr ec -tive action to preclude a recurrence on subsequent missions.

Postmission modification. - After fl ight, the following corre ctiv e action wasimplemented f or the Apollo 9 (CSM 106) missi on and subsequent mis sio ns.

1. An acce leromet er stabilization tes t was added to the EMCA acceptance te stprogram to verify the absence of air bubbles in the ac cel erom ete r damping fluid.

2. A verification te st wa s added to prove the integr ity of the plug-type O-ringseals.

3 . The EMCA acceptance vibration tes ts were revise d to include thre e-ax ist e s t s .

4. For the Apollo 9 (CSM 104) missi on and s ubsequent mi ssi ons , an ApolloOperat ions Handbook (AOH) change, dated Feb rua ry 4, 1969, was implemented toco rr ec t the erro neo us display of 100 ft/s ec. The AOH change moves the sett ing of thedisplay f rom 0 .0 to -100.0 f t/s ec, moves the mode switch to "standby, 'I and specifieswhen it is necessary to bias the counter.

---- ntry postfl ight res ults . - Even though the EMS encountered intermittent tran-sie nts contributing approximately 100 ft/sec to the velocity e r r o r during the Apollo 8flight, the E M S operation at the end of the en try tr a c e w a s within the er r o r budgetspecified in the CSM Operation al Data Book ( Pa rt I, Constra ints and Perform ance) .Postflight ass ess ment indicated that the gen eral operation of the EMS during entr y w a ssatisfactory, except for the two intermittent trans ien ts caused by the air bubbles. Thetra ns ie nt s affected both G and V channels. In addition, an ent ry pattern misaline-ment amounted to a constant 0 .44G e rr o r throughout the entry t ra ce , which, in it se l f ,did not degrade the usefulness of th e EMS. This type of discre pancy in the init ialsett ing of the sc rol l should have been c orre cte d before the ent ry phase.

By comparison to the G-time and V-time derived f ro m onboard pulse-integratingpendulous accelerometer data, overall scroll operation (G-V trace) was within the0.12G limit, except at the two points at which transient positive-G pulses occurred.The EMS ranging perform ance was sa tisfac tory. The EMS ranging accurac y fo r al una r- re tu rn en t ry is k23.04 nautica l mile s. Th e EMS range-to-go a t the end of the

en t ry t r ace (V = 4000 f t / sec) was 17.4 naut ical mi les . The range-to-go counter countsapproximately 25 nautical mile s fro m the end of the ent ry t ra ce to the drogue deploy-ment altitude. At drogue deployment, th e ran ge w a s -7.6 nautical miles, which i smuch more accurate than the *23.04-nautical-mile luna r-re turn ranging accuracyspecified. A furt her substantiation of the accu racy of the E M S was reported when thecrew verified that the range-to-go display read 50 nautical mi les jus t as the

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50-nautical-mile potential-range guideline was reached on the scrol l . Comparison ofthe delta-V at 11 500 ft/sec af ter 0.05G (V = 24 700 ft/sec) to the hardware del ta-Vbudget of 100 ft/s ec indicated that th e Apollo 8 EM S delta-V of 130 ft/s ec with the twot rans ien t s w a s not within the ha rdwar e e r r o r budget but was within the total delta-Vbudget of 300 ft/sec.

Apollo 9 Mission

Performance. - On the Apollo 9 (CSM104) mission, the entry monitor failed toscr ib e dur ing entry. Postfl ight tes t ing of the scr ol l assembly revealed a leak aroundthe base of one of the four scribe-glass-adjustment-screw cups. The neon leak rate

w a s approxi mately 1. 0 cm / sec . The leaking scr ew cup showed evidence of phy sicaldama ge, an indication of pos sibl e unit mishandling af te r the la st leak tes t.

3

Analysis of the scro l l af ter i t was removed from the s cro l l assembly disclosedthat the s cro l l scr ibed proper ly fo r the f i rs t f l ight test pat tern. The scro l l fa i led toscr ibe a t the s ta rt of the pr eentr y flight test pattern and began scr ibin g again during

the l a s t half of the pattern. Scribing of the film w a s prop er down to the initial setposition of the first entry pattern, but scribing failed fro m that point to ent ry untilj u s t before drogue deployment.

Photog raphs of the sc ro ll made through the use of s pe ci al lighting and photo-graphic techniques revealed that the acceleration and velocity drive ass embly thatholds the styl us fo r scrib ing the film functioned proper ly and would have indicated thepro per ent ry pattern had scrib ing occurred. Hardening of the scrib e coat duringflight prevented s cribi ng by the stylus.

T h e s c r o l l is susceptible to moisture, and subsequent slow drying of the mo istu recauses a hardening of t he film coat.

ambient air leaked into the unit before lift-off. During flight, reduction of cabin p re s-su re to 5 psia provided a slow vacuum dry as the moist air was expelled fr om theunit. After being soaked fo r 10 days at 5 psia, the s cri be coat hardened and preventedscrib ing. Postflight analy sis also revealed contamination to the styl us holder andbushing; the contamination caused a 2- to 3-second lag in the s t ylus response.

The sc ro l l w a s probably moisturi zed when

Postmission modification. - After fl ight, the following corre ctive action wasimplemented for the Apollo 10 (CSM 106) mission and subsequent missions .

1. The s tylus spr ing load was increased to 11 cunces.

2. The dimensions of the styl us holder and bushing were verified.

3 . Acceptance te st s wer e performed to verify repeated scri bing.

Entry postfl ight resu l ts . - Postflight as ses sm en t disclosed that, because of thescrib ing anomaly encountered during entry, the s cr ol l reproduction wa s not sufficientfo r an accur ate reading of the G-V t rac e. However, the scr i be did t ra ck the loadfac tor and velocity of the tra jec tor y accurately during entry. The G -V impression

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The ele ctr ical c i rc ui t analy sis was performed basically to determine whetherthe problem was gener ic and applicable to all the numeric d ispla ys. Each digit in thenumer ic display contains seven individual electroluminescent seg ments ; each segmentis controlled by se par ate but identic al logic. Any of the individual compo nents suchas re sis to rs , sil icon-controlled recti fie rs, and integrated cir cu it s could have causedthe fa i lure .

Each electroluminescent segment making up the digits is independently switchedthrough a logic network that activ ates a silicon-controlled rectifier to bypass the lightwhen it is not illuminated . The power sour ce is 115 volts at 800 her tz .

A reliabili ty revi ew revealed four ca ses of malfunctions sim il ar to that report ed.One malfunction involved a constantly illuminated segme nt. In each case, the cau sewa s identified as misro utin g of logic w ir es in the c i r cui t ry control l ing the s i l icon-control led rect i f iers . The misrout ing bent the wi re s ac ro ss termin al s t r ip s contain-ing sh ar p wire ends that punctured the insulat ion and caused ei ther shor t c i rcui ts t oground o r a &volt activation o r deactivation, respect ively , of the segment.

The affected c i rcui t ry wa s reworked in the proc es s of soldering crim p joints thatwe re involved i n the Apollo 7 anomaly. An inspection to detect misrouting was pe r-forme d; however, because of res tri cte d acce ss, the inspection wa s limited.

Although several other failure mechanisms existed in circuit elements and leads,the re wa s no associated fa i lu re his tory. A generic o r design problem was consideredunlikely because of the number of satis facto ry activ ation s demon strat ed by unit s in thefield.

During qualification testi ng, se ve ra l display se gme nts failed because of CCOHpenetration of the moisture seal. However, i n no case w a s a single segme nt blankedwhile the remaining segments wer e in full bril l iance. An improved moisture seal w a s

insta l led to resolve the mois ture problems.

Postmission modification. - No major postmission design change w a s required,because th e past histo ry of the nu meric display problems encountered during qualifi-cation testin g revealed that the most probable cause fo r the electroluminescent segmentfa i lu re was e i the r a faulty component controlling that p arti cul ar segment o r a workman-ship problem associated with wiring o r solder ing. The recommendat ion fo r th i s par -t i cu la r failure was to re ly on exis t ing acceptance screen tes ts and an addit ionalspa cecr aft te st f or v erifying il lumination and uniform intensity of all segments.

Entry postflight results. - Postflight assessment indicated satisfactory overallop er at io n of the EMS. All the test objectives were accomplished. Because the on-board d at a storage equipment was not turned on during entry, the accuracy of the EMSload facto r, velocity, and range da ta by comparison with the integrated pulse-integrat ing pendulous accelero meter d ata w a s not verified.

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CONCLUDINGREMARKS

The operational cha ract eris tics of the entr y monitor subsystem were great lyimproved with two desig n modifications: the changing of the crimp- join t ter mi nalconnections of the delta-velocity/range-to-go counter to solder connections and thereplac ing of the origi nal sc ri be coat with a new microencapsulated-dye pro cess . Thelesson learned from experience with the entry monitor subsy stem w a s to be wary of a

design that is relatively straightforward in itself but that is very difficult toimplement.

Lyndon B. Johnson Space Cente rNational Aeronauti cs and Space Administration

Houston, Te xas , September 4, 1974914-50-17-08-72

REFERENCE

1. Frank, A. J . ; Knotts, E. F. ; nd Johnson, B. C. : An Entry Monitor Systemfo r Maneuverable Vehicles. Pap er 65-495, Am. Inst. Aeron. Astronaut. ,2nd Annual Meeting, July 1965. (Also available in J . Spacecraft Rockets,vol. 3, no. 8, Aug. 1966, pp. 1229-12 34.)

22 NASA- L; ing lcy, 1975 S-405


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