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........ ........ ......... ......... ......... ........ & Y + 4 T .......... .......... ........... ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... NATIONAL AERONAUTICS A N D SPACE ADMINISTRATION .. ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ...........

MSC INTERNAL NOTE NO. 68-FM-172

July 17, 1968

THE DETERMibJATION OF ATTITUDE DEVIATION LIMITS

TRANSLUNAR INJECTION MANEUVER FOR TERMINATING A NON-NOMINAL

By Charles T. Hyle and Alexander H. !Treadway,

Flight Analysis Branch

........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... ............ ........... MISSION PLANNING AND ANALYSIS DIVISION .... ..-. .~

MANNED SPACECRAFT CENTER H 0 l J STON ,TEXAS

i-

.-- . - n 0 9 (ACCE T E P L 35 (THRU)

(PAGES) (CODE) 5 P

ZI -- ' (NASA CR OR TMX OR A D NUMBER) (CATEGORY)

https://ntrs.nasa.gov/search.jsp?R=19700026429 2020-04-23T02:08:15+00:00Z

MSC INTERNAL NOTE NO. 68-FM-172

PROJECT APOLLO

T H E DETERMINATION OF ATTITUDE DEVIATION L I M I T S FOR TERMINATING A NON-NOMINAL

TRANSLUNAR INJECTION MANEUVER

By Charles T. Hyle and Alexander H . Treadway FI ight Analysis Branch

July 17, 1968

MISSION PLANNING AND ANALYSIS DIVISION

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNED SPACECRAFT CENTER

HOUSTON, TEXAS

PfiECEDING PAGE GLANK NOT FILVED .

CONTENTS

Section

SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . FACTORS AFFECTING T L I ATTITUDE LIMITS . . . . . . . . . . . .

Guidance and Control Malfunctions (S-IVB) . . . . . . . . . C r e w Safety . . . . . . . . . . . . . . . . . . . . . . . . Alternate Missions . . . . . . . . . . . . . . . . . . . . Abcrt Planning . . . . . . . . . . . . . . . . . . . . . . Communications and Lighting Constraints . . . . . . . . . . Platform Gimbal Lock (LV and SC) . . . . . . . . . . . . . S-IVBDisposal . . . . . . . . . . . . . . . . . . . . . . Predictable 4ttitude Variations . . . . . . . . . . . . . . General Information . . . . . . . . . . . . . . . . . . . . Summary Remarks . . . . . . . . . . . . . . . . . . . . . . Description of Figures . . . . . . . . . . . . . . . . . .

CONCLUSIONS . . . . . . . . . . . . . . . . . . . . . . . . . REPEBENCES . . . . . . . . . . . . . . . . . . . . . . . . . .

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50

iii

TABLES

Table Page

I LUNAR ALTERNATE MISSIONS AND AUDWABIE AV . . . . . . . 10 I1 ALLOWABLE ATTITUDE RATES FOR LUNAR ALTERNATE MISSIONS . . 11

I11 ALLOWABLE ATTITUDE DEVIATIONS FOR LUNAR ALTERNATE MISSIONS . . . . . . . . . . . . . . . . . . . . . . 1 2

I V APPROXIMATE PERICYNTHION ALTITUDE FOLLOWING AN ALTERNATE MISSION MIDCOURSE . . . . . . . . . . . . . . 13

iv

FIGURES

Figure

1

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4

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6

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8

9

10

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Typical a t t i t u d e excursions during t r ans luna r i n j ec t ion due t o slow deviat ion S-IVB malfunctions . .

Launch vehicle a t t i t u d e at TLI i gn i t i on . . . . . . . . Velocity change through t h e t rans lunar i n j ec t ion

b u r n . . . . . . . . . . . . . . . . . . . . . . . . . Apogee a l t i t u d e and perigee a l t i t u d e through the

t rans lunar i n j ec t ion burn . . . . . . . . . . . . . . I n e r t i a l ve loc i ty , i n e r t i a l f l ight-path angle , and

i n e r t i a l azimuth through the t rans lunar i n j ec t ion b u r n . . . . . . . . . . . . . . . . . . . . . . . . .

Lati tude, longi tude, and a l t i t u d e through t h e t rans lunar i n j ec t ion burn . . . . . . . . . . . . . . . . . . . .

Launch vehicle a t t i t u d e through a t y p i c a l t rans lunar i n j e c t i on maneuver . . . . . . . . . . . . . . . . . .

Maximum allowable S t t i t ude deviat ion t o avoid en t ry after TLI . . . . . . . . . . . . . . . . . . . . . .

Osculating apogee a l t i t u d e as a function of S-IVB platform d r i f t rate

( a ) Pit.ch d r i f t rate . . . . . . . . . . . . . . . . . . (b ) Yaw d r i f t r a t e . . . . . . . . . . . . . . . . . . . Allowable dr i f t r a t e s f o r avai lable a l t e r n a t e mission

AV . . . . . . . . . . . . . . . . . . . . . . . . . Entry midcourse time requirement . . . . . . . . . . . . Communications and l i gh t ing considerations . . . . . . . Alternate mission AVMcc as a function of a constant

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18

19

20

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22 23

24

25

26

27 d r i f t r a t e f o r various TLI burn times . . . . . . . .

V

Figure

1 4

15

16

17

18

19

20

21

22

23

24

25

26

Deviation from nominal l i n e of apsides as a function of constant S-IVB platform &if% r s t e s through TLI . . . . . . . . . . . . . . . . . . . . . . . .

as a funct ion of S-IVB platform p i t c h d r i f t r a t e "MCC fo r two midcourse correctiorr times . . . . . . . . .

as a function of TLI burr! time f o r various p i t ch "MCC drif ' t r a t e s . . . . . . . . . . . . . . . . . . . .

as a function of t h e ro t a t ion of the l i n e of "MCC apsides f rom a nominal TLI cutoff or ien ta t ion . . .

Osculating apogee a l t i t u d e s as a f-mction of the ro t a t ion of the l i n e of apsides f'rorn a nominal TLI cut off o r i en t a t ion . . . . . . . . . . . . . . .

Alternate m i s s i m AV as a function of a constant MCC d r i f t r a t e f o r various TLI burn times . . . . . . .

Deviation from nominal i nc l ina t ion as a function of constant S-IVB platform d r i f t r a t e s through TLI . .

AVMcc as a function of S-IVB p l a t f o m yaw drif 't r a t e s

f o r two midcourse times . . . . . . . . . . . . . . AVMCC as a function of TLI burn time for various yaw

d r i f t r a t e s . . . . . . . . . . . . . . . . . . . . L'VMcc as a function of delta inc l ina t ion from a

n m i n a l TLI cutoff or ien ta t ion . . . . . . . . . . . Osculating apogee a l t i t u d e as a function of d e l t a

i nc l ina t ion from a nominal TLI cutoff or ien ta t ion . Geocentric i nc l ina t ion time h i s t o r i e s fx various

yaw d r i f t rates . . . . . . . . . . . . . . . . . . I n e r t i a l ve loc i ty time h i s t o r i e s f o r various p i t ch

dr i f t rates . . . . . . . . . . . . . . . . . . . .

28

29

30

31

32

33

34

35

36

3;

38

39

40

vi

Figure

27

28

29

30

31

32

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35

I n e r t i a l f l ight-path angle time h i s t o r i e s f o r various p i t c h d r i f t r a t e s . . . . . . . . . . . . . . . . . .

Alti tude time h i s t o r i e s f o r various p i tch d r i f t rates . . . . . . . . . . . . . . . . . . . . . . . .

I n e r t i a l ve loc i ty time h i s t o r i e s f o r yaw d r i f t rates of 0 and .1 degree pe r second . . . . . . . . . . . .

I n e r t i a l f l ight-path angle time h i s t o r i e s fo r yaw d r i f t rates of 0 and .1 degree per second . . . . . .

Alti tude t i m e h i s t o r i e s f o r yaw drif’t r a t e s of 0 and .1 degree per second . . . . . . . . . . . . . .

Pitch a t t i t u d e t i m e h i s t o r i e s i n i n e r t i a l system f o r various p i t ch d r i f t r a t e s . . . . . . . . . . . . .

Pitch a t t i t u d e t i m e h i s t o r i e s i n l o c a l v e r t i c a l system f o r various p i t ch d r i f t r a t e s . . . . . . . . . . . .

Yaw a t t i t u d e time h i s t o r i e s i n i n e r t i a l system for various yaw d r i f t r a t e s . . . . . . . . . . . . . . .

Yaw a t t i t u d e time h i s t o r i e s i n l o c a l v e r t i c a l system f o r various d r i f t r a t e s . . . . . . . . . . . . . . .

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v i i

THE DETERMINATION OF ATTITUDE DEVIATION LIMITS FOR TERMINATING

A NON-NOMINAL TRANSLUNAH INJECTION MANEUVER

By Charles T . Hyle and Alexander H. Treadway

SUMMARY

This report contains an ana lys i s of t he t rans lunar i n j ec t ion (TLI) maneuver and various f ac to r s a f f ec t ing ea r ly terminat ion requirements. A co r re l a t ion between vehic le a t t i t u d e and maneuver terminat ion require- ments i s drawn.

measurement un i t ( I M U ) platforms and t ranspos i t ion and docking (T&D) communications and l igh t ing cons t r a in t s , t o t a l a t t i t u d e l i m i t s which insure safe operations and maximum opportunity for a circumlunw a l t e rna te mission are b6' and -43.4' i n p i t ch and 52' and -66' i n yaw. Because of t h e p o s s i b i l i t y of inconsis tent and unfamiliar IMU alignments a t T L I i g n i t j on md because s impl ic i ty has always served bes t , a t o t a l a t t i t u d e deviat ion of 245' i n either p i t ch or yaw i s recommended as the TLI s hut-down c r i t e r i a.

It was found t h a t , disregarding a gimbal locked ST-124 o r i n e r t i a l

TNTRODUCTI ON Because the T L I maneuver w i l l occur without bene f i t of ground

t racking , a real-time evaluation of t h i s l a rge t h r u s t maneuver m u s t be made so le ly by the crew using onboard displays. I n reference 1, it was determined that. the two indeFendeQt spacecraf t a t t i t u d e reference frames and the associated information displayed on t h e two f l i g h t d i r ec to r a t t i t u d e ind ica tors ( F D A I ' s ) provided t h e most e f f ec t ive maneuver monitoring displays. inary TLI shutdown l i m i t s of 15' a t t i t u d e deviat ion from the expectrcr or 1 0 deg/sec i n r a t e the f a c t t h a t launch vehicle malfunctions character ized by r a t e buildup do not recover and eventually lead t o a spacecraf t E t ruc tura l break at about 70 deg/sec.

Based on information ava i lab le a t t ha t time , prel in-

were suggested. The rate l i m i t i s based primarily on

The a t t i t u d e deviat ion l i m i t w a s b w e d on a conservative estimate of allowable ST-124 platform d r i f t from which a hybrid lunar a l t e rna te . mission could be flown. The conservatism was in+,roduced pr imari ly because of the des i r e t o keep the f l i g h t t r a j e c t o r y near t o nominal and abort and a l t e rna te mission plans m e general ly defined r e l a t i v e t o the nominal. Additional information per t inent t o t h e f i n a l s e l ec t ion of a t t i t u d e deviation l i m i t s has become ava i lab le . The remainder of t h i s

2

paper i s an examination of these f ac to r s and an e f f o r t t o def ine t h e most meaningful S-IVB T L I shutdown c r i t e r i a .

FACTORS AFFECTING TLI ATTITUDE LIMITS

The following fac tors d i r e c t l y o r i n d i r e c t l y a f f e c t t he choice of an a t t i t u d e deviation l i m i t f o r shut t ing down the TLI maneuver.

1. Guidance and cont ro l malf’unctions producing slow t r a j ec to ry deviat ions.

2. Crew sa fe ty .

3. Alternate missions.

4. Abort planning.

5 . Communication and l i gh t ing requirements during t ranspos i t ion and docking.

6. Platform G i m b a l Lock (launch vehicle and spacec ra f t ) .

7. S-IVB Disposal.

8. Predictable var ia t ions i n t h e nominal burn a t t i t u d e h is tory .

Guidance and Control Malflmctions ( S - I n )

Among the known S-IVB malf’unctions, those which can r e s u l t i n slow t r a j e c t o r y o r a t t i t u d e deviat ions, and the associateG c r i t i c a l i t i e s , i. e . p robabi l i ty of mission loss per mi l l ion f l i g h t s , a r e

1. Loss of i n e r t i a l a t t i t u d e . . . . . . . . . . . . . 2300

2. Loss of a t t i t u d e e r r o r commands . . . . . . . . . . 50

3. Loss of a t t i t u d e command s igna l . . . . . . . . . . 4

4. One actuator inoperat ive ( n u l l ) . . . . . . . . . . 1700

5. ST-124 platform d r i f t . . . . . . . . . . . . . . . 36 8

The f i rs t four of these malf’unctions are character ized by t h e low r a t e , l a rge a t t i t u d e excursion shown i n f igure 1. i s l e s s than 1 deg/sec and, therefore , probably not not iceable t o the

It can be seen t h a t the r a t e

3

crew, which i s , of course, a major problem f o r the slow deviat ion malfunctions. t e s t ) i n t h e Saturn computer, t he guidance f a i l u r e l i g h t i s lit fo r a tumbling platform. Therefore, the lo s s of i n e r t i a l a t t i t u d e reference i s the only f a i l u r e with a spacecraf t cue o ther than a t t i t u d e deviation. F r m reference 2 , these f a i l u r e s can occur i n e i t h e r a p i t ch o r yaw channel, and t y p i c a l deviations can reach 90' within 100 seconds of burn t i m e . Additional s tud ie s involving these malf'unctions a re being made by Boeing, but are not expected t o influence the S-IVB shutdown l i m i t s ince a t t i t u d e excursion by i t s e l f i s not dangerous. Analysis of t he f i f t h malfunction, a simulated launch w h i c l e platform d r i f t , Frovides a w a y t o examine some of t he actual cons t ra in ts . The next four sec t ions d r a w l a rge ly from T L I t r a j e c t o r y simulations of a constant ST-124 platform d r i f t .

Since the re i s a platform gimbal check (reasonableness

Crew Safety

A t y p i c a l TLI maneuver and the ign i t i on geometry a re shown i n f igures 2 through 7. Because the i n i t i a l o r b i t i s 100-n. m i . a l t i t u d e c i r c u l a r , a s ign i f i can t e r r o r i n t h r u s t d i r ec t ion would be rzquired t o cause an immediate vehicle en t ry . The primary crew safe ty consideration i s , therefore , t o always provide a safe perigee. References 1 and 3 both show t h a t platform d r i f t i n t he negative p i t ch (down) d i r ec t ion r e s u l t s i n immediate entry by the end of t he 328-second burn. per igee, which i s the ea r th o r b i t i n se r t ion GO - NO-GO c r i t e r i a , i s obtained by a negative p i tch r a t e of -0.175 deg/sec through the burn. pos i t ive p i t ch r a t e s produce lower perigees tnan t h e negative equivalents , pos i t ive r a t e s r e s u l t i n t he vehicle ending t h e burn postpericynthion and therefore a re not time c r i t i c a l . m a x i m u m allowable a t t i t u d e deviat ion t h a t would provide a 75-n. m i . perigee a l t i t u d e a t the end of t he burn would be (-0.175) (328) = -57'. Examining d r i f t r a t e s up t o 1. degjsec, however, shows t h a t l e s s than a 75-n. m i . perigee r e s u l t s i f d r i f t r a t e s between -0.2 and -0.4 deg/sec a re allowed t o continue t o -57'. From f igure & t h e m a x i m u m deviation insuring at l e a s t a 75-n. m i . perigee T a t he remote p o s s i b i l i t y t h a t a constant dr i f ' t of -0.175 aeg/sec occurred, terminating the burn a t a 52' a t t i t u d e woul3. mean a cutoff 42 seconds ear ly . m u s t be protected against a l l d r i f t r a t e s .

A 75-n. m i .

Although

Using the -0.175 deg/sec rate, the

i s -52'.

Even so , t he t r a j e c t o r y

Alternate Missions

Among the many possible a l t e rna te missions from a non-nomina; T L I , those se lec ted by the Apollo Alternate Mission Vwking group t o be supported by p re f l igh t planning are described i n reference 4. a l t e rna te s involviyg the mocn require a burn duration of approximately 290 seconds, depending on the time of t he midcourse maneuver, as shown i n reference 5. This burn duration normally produces an apogee a l t i t u d e of about 30 000 n. m i . be made from a TLI which produces an apogee of about 30 000 n. m i . From

Those

A CSM flyby of t he mom ( a l t e r n a t e 4 ) can therefore

4

f igure 9 t h e m a x i m u m allowable platform d r i f t r a t e s which can r e s u l t i n a 30 000-n. m i . apogee a l t i t u d e are -0.18 deg/sec and +0.175 deg/sec i n p i tch and +C.125 deg/sec i n yaw. Respective deviations by t h e end of T L I are -59' : t57' and 41'. apogee a l t i t u d e , a c loser examination i s necessar j t o insure t h a t t h e d i rec t ion of t he deviated TLI i s within SPS correct ion CapabiliLy. the previous sec t ion t h e m a x i m u m allowable negative p i t ch l i m i t was -5.2'. Although t h i s l i m i t i s too r e s t r i c t i v e f o r t he -0.175 deg/sec case, t he flyby could s t i l l be accomplished by making the h y k i d midcourse correct ion at about 3 hows pas t TLI L s t e a d of 4 hours. Alternate lunar missions being planned include :

Because these l i m i t s provide only a 30 000-n. m i .

I n

DesignaticA

4

5A

Description

CSM circumlunar

Hybrid circumlunar with t h e LM (2000 f p s AV is saved f o r abor t )

5B CSM h m a r o r b i t

6 L u n a r o r b i t with LM; n? reiidezvous

7 Return t o nominal

For each of these a l t e rna te s , a contingency AV has been subtracted from the ava i lab le , and the r e s u l t , as shown i n t a b l e I , is an approximate AV fo r use as a "midcourse correction" from a deviating TLI . Using the midcourse AV (AVMcc) for each a l t e r n a t e with f igure 10, t h e maximum

allawable d r i f t r a t e i n both pos i t i ve and negative p i tch and y~.w can be establ ished as i n t a b l e 11. For constant drifts through T L I , t h e r e su l t i ng deviations a re shown i n t a b l e 111. Figure 10 is only approximate f o r a l t e rna te s 5B, 6 , and 7 but does serve t o demonstrate t he wide choice of acceptable deviations accordirA, t o object ives . It w a s assumed the m i d - course would be made at T L I plus 3 hours and the t o t a l f l i g h t time would not exceed 160 hours. a AVMCC of 7000 f p s wazi used f o r a l t e r n a t e 4, and 1000 f p w a s saved f o r

a possible delay i n making the midcourse. se lec ted f o r only t h e widLst acceptable s i t u a t i o n , t he a l t e r n a t e 4 deviations give -54' and +6i0 i n p i t c t and -66' and 52' i n yaw. the -52' l i m i t f o r negative p i t ch takes precedence. a t t i t u d e deviat ion presented i n t a b l e I11 could vary s l i g h t l y dcpending on the month of launch. from a 72' launch azimuth i s the nominal reference f o r t h i s r epor t . )

Table I V shows the va r i a t ion i n pericynthion. Also

Assuming t h a t l i m i t s should be

Again It i s noted the

(The February 2 , 1968 first oppoytunity in jec t ion

Abort Planning

Through several months of in-house MPAD and TRW discussions, prelim- inary abort procedures were defined and el-eptually agreed t o i n Apollo Abort Working Group meetings. One of these planned abort techniques cons is t s of shut t ing down t h e TLI i f t h e crew should need t o en te r immediately. Although such a t ime-cr i t ica l f a i l u r e has not been i d e n t i f i e d , t h e technique, re fer red t o as the f ixed a t t i t u d e abor t , i s provided f o r safet.y. If such abort were required following a slow drift of the S-IVB, normal dispersions i n Derfoming t h e a b o r t maneuver result i n an entry which i s no longer i n t h e V-versus-y entry in t e r f ace coryidor. entry midcourse maneuver; however, a minimum S-IVB burn i s required i n order t c have time t o perform such a midcourse. discussed s o far a r e k0.2 deg/sec. These r a t e s and t h e min-hum S-IVB burn t i m e provides the max imum deviat ion which insures t h e f ixed a t t i t u d e abort maneuver does not r e s u l t i n an en t ry out of t h e corr idor without allowing enough t i m e t o correct with the midcourse. study of t he S-IVB burn t i m e required t o generate t h e boundary shown i n f igure ll. This p i tch a t t i t u d e deviat ion i s (0.2 deg/sec) (230 sec ) = 46" and (-0.2'/sec) (217 sec ) = -43.4'. For t h e lunar mission, t h e f ixed a t t i t u d e abort may produce a land landing; however, f o r t he E mission the delay t i m e t o SPS i gn i t i on is es tab l i shed t o insure water landing. Allowing s ign i f i can t a t t i t u d e deviations m a y a l s o result i n land landicgs f o r the E mission.

Reference 6 contains most of t he associated d e t a i l s .

Normally t h i s would be handled by an

The m a x i m u m a l l a r a b l e r a t e s

Reference 7 r e su l t ed i n a

Communications and Lighting Constraints

The impact of deviating t r a j e c t o r i e s on communication and l i gh t ing c o n s t r a h t s can be examined w i t h a t y p i c a l a t t i t u d e p r o f i l e providtng optimum S-IVB high gain t racking and sunl ight angles as described i n reference 8. At-TLI-plus 15-minutes, t h e S-IVB woult p i t ch t o a p re f l igh t loaded a t t i t u d e , which it would hold throughout T&D-TLI-plus-2-hours. Since t h i s a t t i t u d e is optimized t o give favorable t racking f o r two hours, a small change ir i n i t i a l a t t i t u d e , as caused by platform d r i f t , could eas i ly cause loss of required telemetry by t h e ground. Even though the crew could hold proper a t t i t u d e from the spacecraf t , t h i s capabj l i ty is l o s t when they separate f o r T&D. Although it i s planned t o do T&D f o r a l l the a l t e rna te missions (reference 9 ) , t h i s object ive could be dropped fo r a l t e rna te 4.

Alternate 5A, however requires T&D and, as seen from t a b i e 111, could m e a n a -33' or a 413 e r r o r i n S-IVB ccxnmunication a t t i t u d e . This i s the i n i t i a l q r ro r and does not include addi t iona l e r r o r through T&D. Possible l i gh t ing e f f ec t s which depeni". OIJ the intended lunar landing s ight and a communications a t t i t u d e %--e C pic ted i n f igure 12.

6

Because the optimum communicatim a t t i t u d e and l i g h t i n g conditions vary, max imum a t t i t u d e deviat ion l i m i t i s not e a s i l y dztermined; however, it most l i k e l y would have t o be f a i r l y s m a l l - probably less than 30'. 3nly p a r t i a l t racking would be ava i lab le then. remain on t h e S-I'JB &id cont ro l t h e a t t i t u d e u n t i l t h e ground es tab l i shed tinat it w a s sa fe f o r T&D. The crew Could then execute a manual T&D as soon as possible . AVMcc required.

Possibly t h e crew could

Delay pas t TLI-plus-3-hours would increase i n t h e

Platform Gimbal Lock (LV and SC)

Launch vehicle.- The gimbal lock l i m i t on t h e ST-124 platform i n t h e Saturn i s 60° i n t he yaw plane. P r io r t o reaching t h i s l i m i t onboard log ic limits the maximum a t t i t u d e e r r o r command t o 45'. t h e malfunctions prodiicing the slow deviations a re a t t i t u d e e r r o r s igna l problems of some type , yawing beyond these l i m i t s s t i l l i s not l i k e l y t o change response d ra s t i ca l ly . the yaw plane a l t e rna te mission limits o r u n t i l a possible spacecraft gimbal lock limit.

Since most cf

Therefore t h e burn could be con t inued to

Spacecraft.- The th ree gimbal plztforms ( I M U ) i n t h e spacecraf t present ly cease t o provide va l id a t t i t u d e information when yawed pas t 85O. Many measures have been taken t o avoid such an occurrence. developments during meetings of t h e Apollo Data P r i o r i t y Midcourse Techniques Group have eliminated t h e requirement f o r t h e e n t i r e primary guidance and navigation cont ro l subsystem (PCNCS) f o r some lunar missions, which of course negates the requirement f o r t h e IMU. It i s f e l t t h a t t he s t a b i l i z a t i o n and cont ro l subsystem (SCS) provides an adequate backup system. With t h i s i n mind the re does not appear t o be much j u s t i f i c a t i o n f o r TLI shutdown t o pre-Jent I N gimbal lock.

However, recent

Since most of t h e malfunctions previously described can produce t h i s problem and i f f o r some reason gimbal lock m u s t be avoided, time ( a t t i t u d e deviat ion) would have t o be allowed f o r the spacecraf t t o separate and damp ra t e s .

If 3 seconds are provided from shutdown t o SPS ign i t ion , Only 55'

For a

If we had committed t o TLI with no PGNCS, t h e

yaw a t t i t u d e excursion from lift-off o r i en ta t ion could be allowed, considering a 10 deg/sec r a t e l i m i t , t o avoid IMU gimbal lock. t y p i c a l TLI plane change, t h i s means a yaw a t t i t u d e deviat ion of 42' ( i . e . , 55' - 13'). SCS e lec t ronic gimbal lock at 75' reduces the a t t i t u d e deviat ion l i m i t t o 32'.

In summary t.0 avoid IMU gimbal lock, an a t t i t u d e deviat ion l i m i t of about 42' i s required, however, s ince the crew has three hours t o midccurse time they could probably rea l ign . i s not a mandatory requirement f o r TLI shutdown.

Therefore, t h e gimbal lock avoidance

7

S-IVB Disposal

It i s noted t h a t even though pos i t i ve p i t c h rates during T L I r e s u l t i n t he spacecraft being post perigee a f t e r TLI, a low perigee such as allowed f o r a l t e r n a t e 4 may a l s o provide S-IVB disposal problems. Reducing t h e allowable pos i t i ve p i t ch l i m i t t o l e s s than 60° improves chances of a gradually decaying o r b i t f o r t h e S-IVB and may the re fo re prevent what would probak.Ly be a land impact.

Predictable Aktitude Variat ions

Three th ings should be noted under t h i s top ic . The f i r s t i t e m concerns t h e propel lant u t i l i z a t i o n system which can, as observed on t h e AS-501 mission, change t h e expected a t t i t u d e h i s to ry of t h e launch vehicle considerably acd s t i l l be on a good i ra jec tory . The amount of excursion depends pr imari ly on t h e amount of l i q u i d hydrogen ava i l ab le . It i s f e l t t h a t t h i s i t e m should not inf luence l i m i t s e t t i n g because t h e a t t i t u d e excursion i s predictahle p r e f l i g h t , as i s t h e t ime t h e propulsion u n i t (PU) system is ac t ive . case and s ince t h e remaining S-IVB second burns, such as TLI w i l l occur primarily near t h e horizontal , a t t i t u d e excursions a r e expected t o be s ign i f i can t ly l e s s . Marshall- Space F l ight Center doc-ments such as reference 2.

cur ren t ly programed t o go t o t h e expected ign i t i on a t t i t u d e . moves t o a prestored gimbal pos i t ion regard less of t h e parking o r b i t dispers ions. i n a 100-n. m i . c i r cu la r o r b i t . After i g n i t i o n and ac t ive guidance b,?gins, compensation would be made f o r t h e a t t i t u d e . was a 100-by 1000-n. m i . a l t i t u d e one, t h e maximum deviat ion would only be about 6' , and, of course, t h e crew would expect t h i s .

Also, s ince t h e AS-501 mission was a spec ia l a t t i t u d e

These predicted p r o f i l e s are usual ly ava i lab le i n

The second item i n t h i s category concerns t h e way t h e S-IVB i s It always

This could mean it is aligned i n i t i a l l y as though it were

Even i f t h e parking o rb i t

The t h i r d item concerns other parking o r b i t dispers ions and t h e e f f ec t s on burn a t t i t u d e . Reference 1 0 contains a t t i t u d e h i s t o r i e s through t h e E and G mission second S-IVB burns. Dispersions include s t a t e vec tor , i gn i t i on time and thrust e r ro r s . general ly less than 5 O and are therefore excluded as shutdown l i m i t s .

The r e su l t i ng a t t i t u d e deviat ions a r e

General Information

Several changes from t h e monitoring sec t ion of reference 1 a r e described here for completeness. from t h e maneuver monitor program 47.

t h e components of ve loc i ty accumulated as measyed i n spacecraft cont ro l axes - i n e r t i a l ve loc i ty ( V I ) , a l t i t u d e r a t e (h) and a l t i t u d e ( h ) are

The DSKY parameters now ava i lab le a r e I n addi t ion t o L V X , AV and AVz -

Y ,

a

avai lable . a l t i t d e (ha) , perigee a l t i t u d e ( h ) , and time t o perigee ( t ) .

Also rout ine 30 w i l l enable t h e crew t o monitor apogee

P P

The de le t ion of program 1 5 eliminated a possible means of TLI cont ro l by t h e CMC and a l so removed a possible cue f o r cutoff as w e l l as a t t i t u d e e r ro r displays. chances of providing a meaningful cutoff backup a r e small s ince only a few seconds remain u n t i l f u e l deplet ion.

t .though t h e AV meter can provide a backup cue f o r cu to f f ,

Another s ign i f i can t chaiige requi res t ha t t h e spacecraft i n e r t i a l reference systems w i l l be aligned t o t h e launch pad alignment, as i s t h e S-IVB, f o r RTCC and ground cont ro l convenience. This means t h a t t h e i n i t i a l p i t ch a t t i t u d e of t h e FDAI d isp lays can be anything between 0" and 360' depending on i n j e c t ion opportunity.

Summary Remarks

An important ground r u l e being used by t h e Midcourse Techniques Group t o e s t ab l i sh a NO-GO f o r T L I e r ro r s which are l a rge r than c e r t a i n spec i f i ca t ion values. of l a rge TLI a t t i t u d e deviat ion l i m i t s i s somewhat contradictory t o t h i s ground rule.

re l ies on detecti i lg Saturn o r PGNS Acceptance

Perhaps t h e most s ign i f i can t i t e m not ye t discussed i s s impl ic i ty . Without regard t o s impl ic i ty , t h e widest acceptable l i m i t s f o r shut t ing down TLI a r e + 46" and -43.4' i n p i t c h and + 52' and -66" i n yaw. Recall ing t h a t a t t i t u d e deviat ion i s t h e d i f fe rence i n spacecraft FDAI a t t i t u d e and t h e nominal a t t i t u d e at a given burn time, plus t h e f a c t t h a t deviat ion can occur i n a combination p i t ch and yaw d i r ec t ion , it i s f e l t t h a t one l i m i t should be used. It i s the re fo re suggested t h a t f o r s impl ic i ty t h e translunar in j ec t ion maneuvers be terminated on a t o t a l a t t i t u d e deviat ion l i m i t of 45" i n e i t h e r p i t ch o r yaw. The primary constraint not met by t h e above l i m i t is the communications and l i gh t ing requirements. I f an a l t e r n a t i v e such as previously described i s not acceptable, smaller l i m i t s must be imposed and even then t h e communication and l i gh t ing requirements may not be achieved. t h e las t few seconds of T L I , as i n t h e launch phase, ac t ion by t h e crew on an a t t i t u d e deviat ion l i m i t may be dropped because slow deviat ions cannot s ign i f i can t ly jeopardize t h e t r a j ec to ry .

It i s noted t h a t during

It i s emphasized t h a t t he probabi l i ty of a slow deviat ion malfunction i s extremely unl ikely; however, s ince they can occur as w e l l as cause ser ious problems , l i m i t s a r e required.

9

Description of Figures

Typical h i s t o r i e s of previously discussed parameters a r e included i n t h e figures. conditions, DSKY quan t i t i e s a t t i t u d e h i s t o r i e s and geometry va r i a t ions a re provided f o r both p i t ch and yaw platform drifts. t h e f igures involving a l t e r n a t e mission midcourse AV are conic so lu t ions and as a result are approximate.

I n addi t ion t o t h e figures used t o der ive l i m i t i n g

It i s noted t h a t

CONCLUSIONS

Disregarding a gimbal locked ST-124 o r IMU platform and T&D communications and l i g h t i n g cons t r a in t s , t o t a l a t t i t u d e l i m i t s which insure safe operations and maximum opportunity f o r a circumlunar a l t e r n a t e mission are + 4 6 O and -43.4' i n p i t ch and +52O and -66' i n yaw. Because of possible inconsis tent and unfamiliar IMU alignments at TLI i gn i t i on and because s impl ic i ty has always served best e i t h e r p i t ch or yaw i s recommended as TLI shutdown c r i t e r i a .

a -2otal a t t i t u d e deviat ion of +- 4 5 O i n

10

0 0 0 0 0 0 0 0 4 0 0 0 0 ,

0 0 0 0 0 0 0 0 0 0 0 0 0 0 0

I Y .- e 0

- m E E 0 E

.-

s E a PI e Y

11

v) CL

Y- . 0 0 I > a

l n r n o o o l n O N 0 4 s C Q o m 0 d o d o 0

O b 0 0

l n o l n l n o b O 0 l b l n Q O d . r b 0 r l d d 0 0

1 1 1 1 1

0 0 0 0 0 o o o o l n o o o o r l

12

W > a m m aJ z

v) Q. YI

5

0 0 a >

4

0 0 0 0 0 0 0 0 0 ~ O O O O d

m m 4 O G O d . . . . I I

14

0 OI

0 03

0 h

0 9

0 VI

0 d-

0 m

0 N

0 d

N . 0

rn W -0

0 L

5

2 a

W -0 a Y

-4.701

PiTCH WITH RESPECT TO LOCAL HORIZONTAL

20

10

GEODETIC LATITUDE, DEC

10

23 1 10 120 130 140 150 160 170 180 170 160

LONGITUDE, DEG

YAW, LOCAL HORIZONTAL SYSTEM DATA REFERENCE: MSC INTERNAL NOTE NO. 66-FM-70

Figure 2. -Launch vehicle attitude at T L I ignition.

16

17

B 0

18

3 N

0

N

1 I I I I i ro rc u3 L n w M N 4 3 c N m

0 N m

K

w 0 R

u 01 C ' I , 9 b c

c 2

R c

0 w

0 'F

0 N

0

U E m

20

0 co 'V

a

0 a I

F W

0

N 0 03 9 U N C I

U I+ I+ I I

I I I I+

I

sas/6ap ‘8 ‘ale4 W!d

22

0 .@ .08 .12 .16 .20 .24 - L 24 -.20 -. 16 -. 12 -.08 -.a IU pitch drift rate, deglsec

(a) Pitch drift rate.

Figure 9. - Osculating apogee altitude as a function of S- IVB platform drift rate.

-. 12 -. 10 -. 08 -.06 -.a -. 02 0 .02 .04 .06 .08 .10

Yaw dr i f t rate, deglsec

(b) Yaw dr i f t rate.

Figure 9. - Concluded.

24

0 cv

03 0

d 0

0

d 0 I

co 0 I

cv rl

I

0

E 0 .- cn u) .- E

L-

rc O

- a I

0 4

LL

9 rl

-5 D - I

7-

N CO U 0 i-l

26

z 0 0

27

0 l-4

a3 0

9 0

e 0

. aJ Y

0 2

I

U 0

I

9 0

I

a3 0 I

0 l-4

I

N l-4

I

m 0 -

m v) m

S 0 v) v)

.-

.- E

I

m l-4

0 0 0 0 0 0 0 0 -

3

0 N

a rl . N f-l . co 0

8

d 0 .

0

* 0 I

a3 0 . I

(v rl

I e

9 I 4

I . 0 N

e I

0 aJ s al

0

aJ . Y

m, Y

L 'c U r V

P Y .- € 0

m Q

cc Y

- a > v)

- I

I I 2800 -

I I t VI a. 1

- -

Figure 15.- AVMcc as a function of S-IVd platform pitch drift rate for two midcourse correction times.

-c 0

P v) 3 0

U .-

.- t > 0 L

y.

m v) m

I

9 rl

0 0 m

31

0 d

9

L .t-

Y 0 S 0

?

?

0 rl I

N rl I

m VI m 0 u I > U

I . b r(

32

"

-24 -20 -16 -12 -8 -4 0 4 8 12 16 20 24

re19 dq Relative rotation of l ine of apsides, Aw

Figure 18. - Osculating apogee altitude as a function of the rotatlon of the l ine of apsides from a nominal TLI cutoff orientation.

33

0 d

03 0 .

9 0 .

U 0

0

(v 0

0

cv 0

b I

U 0 I 0

9 0

I

W 0 I

0 rl I 0

V aJ v)

aJ -0

. 3

. W 4 2 Y

L 'c -0

E, 0

m Q .

cc U - m > - I v)

E 3

51

aJ U

2

34

6

5

4

3

2 rn W D

- a, .-& 1

Q c 0

m .- Y

.E 0 - 0 S

m a, U

5

m W

.- Y - w -1 .- .-I - oc

-2

-3

-4

-5

-6 -.lo -.08 -.06 -.04 -.02 C .02 .04 .06 .08 .10

S-IVB platform yaw dr i f t rate, deg/sec

Figure 20.- Deviation from nominal inc! ination as a function of constant S-IVB platform dr i f t rates through TLI.

4800

4400

4000

3600

3200

1200

800

400

0

-400

-800 '. 1 -. 08 -. 06 -. 02 -. 02 0 $02 .&I .06 .08 .1

Plat form yaw drift rate, deglsec

F igure 21.- AVMcc as a f unc t i on of S - I V B plat form yaw drift rates for two midcourse times.

36

0 (v m

0 9) ul

0' +

f n 4 I-

3 - + 0

m m

37

4000

3600

3200

2800

m P cc

2400 0 H a k2 2000

> a,

L1 0 u E E

.2 1600 8

v) v,

E

'a 0

4.1

G' .= 1200 c - Q

800

400

0

-40 0 -8 -6 -4 -2 0 2 4 6 8 10

Relative delta inclination, Aire,, deg

Figure 23.- AVMcc as a function of delta inclination from a nominal TLI cutoff orientation.

38

600

560

520

480

440

400 .- E

360 10

l=

200

160

120

80

40

0 -6 -5 -4 -3 -2 -1 0 1 2 3 4 5 6

Relative delta inclination. Aire,, d q

Figure 24. - Osculating apogee altitude as a function of delta inclination trom a nominal TLI cutoff orientation.

U cu

0 9 m

0 co N

0 U cu

8 ul . n Y

0 cn

0

P m 7 - s

r

0 40 80 120 160 2co 240 280 320 360 400 TLI burn time, tb, sec

Figure 26. - Inert ia l velocity t ime histories for various pitch drift rates,

0 9 m

0 <u m

0 (u

o m 0 . - P

f

" i 0 3 Q n

I-

0 03

0

0 co c\1

0

(v

0 a,

0 0

0

L aJ P aJ m aJ V

e

.I

0 S m 0

0 L

44

45

0 9 m

0 (v

m

0 aJ (v

0 v (v

0 0 (v

0 9 4

0 (v rl

0 co

0 T

0 0 0 0 0 0 0 0 0 b 9 rn u - m (v rl 0 r( .-I 4 4 r( 4 r( d

' ! W ' U 'y 'apn,:llv

46

47

0 0 U

0 9 m

ln W

0 m, vl L

42

L

(v

U z U * 0. ln 3 0

m >

.- 0 a3 cu L

.-

I

m m C co ?!

3 m .- IA

0 U

0

48

49

0 0 P rr\

0 rl

0

N

0 Tr N

0 0 N

0 9 rl

0 m rl

0 co

0 Q

C?

0 d I

50

REFERENCES

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

Hyle, Charles T. : Preliminary Display L i m i t s and Crew Monitoring Considerations f o r T L I , LOI , and TEI. September 27, 1967.

MSC I N 67-FM-138,

Boeing Company: Saturn V F l igh t Mechanics and Dynamics Section: Saturn V Launch Vehicle Ehergency Detection System Analysis, SA-503. Boeing Doc No. D5-15555-3B, Apri l 19, 1968.

Treadway, Alexander: The Effec ts of I u Platform Pi tch D r i f t Rates on T L I f o r t h e Lunar Landing Mission. MSC I N 68-rn-113, May 9, 1568.

Huss, C a r l R . : Third and Fourth Lunar Landing Mission Alternate Missions Meetings. MPAD Memo 68-FM-H-32, March 18, 1968.

Duncan, Rocky D . ; Jenness, Martin D ; and Lostak, Dwyl R . : Preliminary Abort and Alternate Mission St,udies f o r 504A. Voluxe I.

Spacecraft

MSC I N 66-FM-4, January 6, 1966.

Contingency Operations Sect ions, TRW: Fixed At t i tude Abort Procedures f o r Aborts Occurring During Translunar In j ec t ion . February 1 5 , 1968.

MSC I N 68-FM-48,

Weber, Ec'Jbie D . : Midcourse and Entry- Following Fixed Att i tude Aborts from TLI . MPAD Memo 68-FM36-84, February 20, 1968.

Lu.iar Mission Analysis and Lunar Landing Branch: P re f l igh t Data f o r a Lunar Landing Mission - Volume I1 - Mission P r o f i l e . February 16 , 1968.

MSC I N 68-FM-41,

Huss, C a r l R . : Minutes of t h e F i f t h Lunar Mission Alternate Missions Meeting. MTAD Memo 68-1"~-~-34, ilarch 18, 1968.

Wiley, Robert F. : Preliminary Analysis of Monitoring an IGM-steered, Hy-persurfa:e-Targeted TLI Maneuver Using CMC Lambert Steer ing and t h e Saturn-Preferred IMU Alignment. MSC I N 68-FM-54, February 27, 1968.