s- george c. - nasa

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NASA TECHNICAL MEMORANDUM

July 31, 1967

A P E R F O R M A N C E S T U D Y FOR THE A P P L I C A T I O N O F THE S A T U R N V T O H I G H E N E R G Y E A R T H E S C A P E M I S S I O N S By RonaldG. Toelle

Ae ro- A s trodynamic s Laboratory

GPO PRICE $

NASA CFSTI PRICE(S) $

Hard copy (HC) s- - George C. Mdrshall Microfiche (MF) / L d

Spme Flight CenteT, ff 653 July 65

Hmtsuille, Aldbdmd

- V z x * &?A 37 L

(NASA CR OR TMX OR AD NUMBER) (CAT~GORY)

TECHNICAL MEMORANDUM X-53639

A PERFORMANCE STUDY FOR THE APPLICATION OF THE

SATURN V TO HIGH ENERGY EARTH ESCAPE MISSIONS

Ronald G. Toe l l e

George C . Marshall Space F l i g h t Center

Hun t sv i l l e , Alabama

ABSTRACT

The r e s u l t s of a performance survey f o r t h e a p p l i c a t i o n of t h e product-improved Sa tu rn V launch veh ic l e s t o v a r i o u s escape ene rg ie s a r e p re sen ted . Two upper s t a g e (S-II/S-IVB) p ropu l s ion systems (5-2 and J-2s) were i n v e s t i g a t e d . Exchange r a t i o s of payload w i t h r e s p e c t t o v e h i c l e parameters ve r sus C, ( twice the energy per u n i t mass) a r e presented f o r a C3 range of 0 t o 125 km2/sec2. The e f f e c t of boos t e r v a r i a t i o n s and proposed v e h i c l e improvement f o r d i f f e r e n t miss ions can be mapped i n t o the payload through t h e j u d i c i o u s use of t he exchange r a t i o s . f o r va r ious e a r t h escape and i n t e r p l a n e t a r y miss ions . The r e s u l t s of this performance survey a r e presented g r a p h i c a l l y .

These da t a a r e p r i m a r i l y f o r use as a guide t o payload p lanning

NASA - GEORGE C. MARSHALL SPACE FLIGHT CENTER

NASA - GEORGE C. MARSHALL SPACE F L I G H T CENTER

Technical M e m o r a n d u m X - 5 3 6 3 9

J u l y 31, 1967

A PERFORMANCE STUDY FOR THE APPLICATION O F THE

SATURN V TO HIGH ENERGY EARTH ESCAPE M I S S I O N S

BY

R o n a l d G . Toe l le

F L I G H T MECHANICS AND PERFORMANCE ANALYSIS SECTION

AERO-ASTRODYNAMICS LABORATORY ADVANCED STUDIES O F F I C E

RESEARCH AND DEVELOPMENT OPERATIONS

LIST OF ILLUSTRATIONS

F igure

A- 1

A- 2

A- 3

A-4

A-5

A-6

A- 7

A- 8

B- 1

B- 2

B- 3

B-4

B-5

B- 6

B-7

T i t l e - S a t u r n V Launch Vehicle , Apollo Conf igu ra t ion ........ S a t u r n V Launch Vehicle , MSFC Nose Cone Conf igura t ion

Launch Vehicle Performance f o r SA-516 Apol lo Con- f i g u r a t i o n vs C 3 ..................................... Launch Vehicle Performance f o r SA-516 MSFC Nose Cone Conf igura t ion ve r sus C3..............................

Optimal Thrus t t o Weight a t L i f t o f f ve r sus C3 f o r Both Conf igura t ions Using Standard 5-2 Propuls ion i n S-IIIS-IVB Stages . ................................... Optimal Thrus t t o Weight a t L i f t o f f v e r s u s C3 f o r Both Conf igura t ions Using J -2s P ropu l s ion i n S-11/ S-IVB Stages ......................................... Percentage Ra t io of S-IVB F i r s t Burn P r o p e l l a n t s i n t o O r b i t t o To ta l S-IVB Capaci ty ve r sus C 3 . . . . . . . . . . . . . .

Optimum Parking O r b i t A l t i t u d e f o r S - I 1 I n s e r t i o n I n t o Parking O r b i t versus C3 .........................

E f f e c t of To ta l F i r s t Stage Thrus t on Payload ve r sus C 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

E f f e c t of F i r s t S tage S p e c i f i c Impulse on Pay ve r sus C3....................................

. . . . . . . oa d . . . . . . .

E f f e c t of F i r s t S t age Dead Weight on Payload vs C3.. .

E f f e c t of To ta l Second Stage Thrus t on Payload vs C3.

E f f e c t o f Second S tage S p e c i f i c Impulse on Payload ve r sus C 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

E f f e c t of Second Stage Dead Weight on Payload v e r s u s C 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

E f f e c t of T h i r d S tage Thrust on Payload vs C3 ........

Page

8

9

1 0

11

12

1 3

14

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19

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iii

LIST OF ILLUSTRATIONS (Continued)

Figure

B-8

B-9

B-10

B-11

B-12

c- 1

c- 2

Table

1

2

T i t l e

E f f e c t of Third S tage S p e c i f i c Impulse on Payload vs c3 ................................................ E f f e c t of Nose Cone Dropped a t E a r l y A l t i t u d e versus C3 f o r MSFC Nose Cone Conf igura t ion .................. E f f e c t of ShroudfNose Cone Dropped a t Various A l t i - tudes ve r sus C3 f o r MSFC Nose Cone Conf igura t ion ..... E f f e c t of Axia l Force C o e f f i c i e n t on Payload vs C3. . .

E f f e c t of Launch Azimuth on Payload vs C 3 . . . . . . . . . . . .

Nominal Var i a t ion of Axia l Force C o e f f i c i e n t w i t h Mach Number f o r Apollo Sa tu rn V Conf igura t ion ........ Nominal V a r i a t i o n of Axia l Force C o e f f i c i e n t w i t h Mach Number f o r Sa tu rn V MSFC Nose Cone Conf igura t ion

LIST OF TABLES

Propuls ion C h a r a c t e r i s t i c s f o r Sa tu rn Launch Vehicle SA-516 ............................................... Weight Summary f o r Sa turn Launch Vehicle SA-516 Apollo Conf igura t ion wi th Standard 5-2 Engines i n S-II~S-IVB Stages . . . ................................. Weight Summary f o r Sa tu rn Launch Vehicle SA-516 Apollo Conf igura t ion w i t h J-2s Engines i n S-111s-IVB Stages ............................................... Weight Summary f o r Sa tu rn Launch Vehicle SA-516 MSFC Nose Cone Conf igura t ion w i t h Standard 5-2 Engines i n S-111s-IVB Stages ......................... Weight Summary f o r Sa tu rn Launch Vehicle SA-516 MSFC Nose Cone Conf igura t ion wi th J-2s Engines i n S-II~S-IVB Stages ....................................

i v

Page

25

26

27

28

29

32

33

34

35

36

37

38

A Z

c3

PLD

‘AT

DEFINITION OF SYMBOLS

D e f i n i t i o n

l i f t - o f f weight a t ground i g n i t i o n

f i r s t s t a g e t o t a l sea l e v e l t h r u s t

f i r s t s t a g e s e a l eve l s p e c i f i c impulse

f i r s t s t a g e dead weight

second s t a g e t o t a l vacuum t h r u s t

second s t a g e vacuum s p e c i f i c impulse

second s t a g e dead weight

t h i r d s t a g e vacuum t h r u s t

t h i r d s t a g e vacuum s p e c i f i c impulse

t h i r d s t a g e dead weight

ins t rument u n i t

t h rus t - to-we igh t r a t i o

J- 2 s i m p 1 i f i ed propuls ion sys tem

launch azimuth

twice the energy per u n i t mass o r hyperbol ic excess speed squared

n e t payload

a x i a l f o r c e c o e f f i c i e n t

V

TECHNICAL MEMORANDUM X-53639



SUMMARY

The r e s u l t s of a performance survey f o r a p p l i c a t i o n of a product- improved Sa tu rn V launch v e h i c l e t o va r ious escape e n e r g i e s a r e presented i n t h i s r e p o r t . Two upper s t a g e (S-II/S-IVB) propu l s ion systems ( 5 - 2 and J-2s) were i n v e s t i g a t e d . The nominal f l i g h t p r o f i l e was S-IC/S-II/S-IVB t o a one-hundred-nautical-mile parking o r b i t w i t h r e i g n i t i o n of t he S - I V B s t a g e t o i n j e c t the payload t o the des i red energy v a i u e . Because the f i r s t - b u r n p r o p e l l a n t of t h e S-IVB s t a g e opt imized t o z e r o f o r h igh energy v a l u e s , s p e c i a l f l i g h t p r o f i l e s were i n v e s t i g a t e d t o extend the payload c a p a b i l i t y t o h igher energy va lues . These a re S - I 1 i n j e c t i o n i n t o a h ighe r c i r c u l a r o r b i t t han nominal, and a s i n g l e burn S-IVB s t a g e o u t of o r b i t .

Exchange r a t i o s of payload wi th r e s p e c t t o v e h i c l e parameters ve r sus C3 ( twice t h e energy pe r u n i t mass) a r e presented f o r a C3 range of 0 t o 125 km2/sec2. improvements f o r d i f f e r e n t miss ions can be mapped i n t o t h e payload through j u d i c i o u s u s e of t h e exchange r a t i o s . These d a t a a r e p r i m a r i l y f o r use as a guide t o payload p lanning f o r var ious e a r t h escape and i n t e r p l a n e t a r y mis s ions . The r e s u l t s of t h i s performance su rvey a re presented g r a p h i c a l l y .

The e f f e c t of boos te r v a r i a t i o n s and proposed v e h i c l e

I. INTRODUCTION

During the p a s t decade, t h e m a j o r g o a l o f t h e n a t i o n a l space program has been t h e e x p l o r a t i o n o f n e a r - e a r t h space w i t h a prime goa l o f l and ing two men on t h e moon and r e t u r n i n g them t o e a r t h by 1970. A s t he Apollo program comes n e a r e r t h i s goa l , t h e i n t e r e s t o f NASA p lanne r s t o send l a r g e payloads beyond t h e earth/moon system has g r e a t l y inc reased . A q u e s t i o n o f t e n asked i s "What i s the payload c a p a b i l i t y of the Sa tu rn V launch v e h i c l e t o h igh energy missions?" Th i s i n v e s t i g a t i o n i n d i c a t e s t h a t t h e Sa tu rn V launch v e h i c l e , developed f o r t h e Apollo program, has t h e c a p a b i l i t y w i t h minimum modi f i ca t ion ( s losh b a f f l e s i n t h e S - I V B s t a g e ) o f p l a c i n g space probes t o the o u t e r reaches o f t h e s o l a r system.

A product-improved S a t u r n V launch vehicle , des igna ted as SA-516, has been def ined f o r two c o n f i g u r a t i o n s . One c o n f i g u r a t i o n , t h e Apol lo geometry conf igu ra t ion , is a man-rated v e h i c l e ; i . e . , t h e Launch Escape System i s a v a i l a b l e f o r b o o s t e r a b o r t s . The second c o n f i g u r a t i o n , the MSFC Nose Cone c o n f i g u r a t i o n , is a n unmanned f l i g h t v e r s i o n . The nose cone is j e t t i s o n e d i n o r b i t f o r the performance d a t a p re sen ted , and no c y l i n d r i c a l payload shroud s e c t i o n i s d e f i n e d . When a mis s ion is d e f i n e d , the e f f e c t of t h e payload shroud upon t h e i n j e c t i o n payload can be ca l cu la t ed from the exchange r a t i o s conta ined i n Appendix B, as can o t h e r v e h i c l e p e r t u r b a t i o n s .

Two upper s t a g e (S-111s-IVB) propu l s ion sys tems were i n v e s t i g a t e d f o r each conf igu ra t ion , t h e f i r s t be ing the s t a n d a r d 5 - 2 engine propul - s i o n as p r e s e n t l y def ined f o r the Apollo program. The second system inves t iga t ed i s the 5-2 s i m p l i f i e d engine (J-2s) which d i s p l a y s a g a i n i n s p e c i f i c impulse while reducing the respect ive S - I 1 and S-IVB s t a g e dead weights . The performance c h a r a c t e r i s t i c s of the r e s p e c t i v e propul- s i o n systems are g iven i n Appendix C .

These performance d a t a a r e a p p l i c a b l e t o s i n g l e launch S a t u r n V ' s .

11. ASSUMPTIONS USED FOR PERFORMANCE CALCULATIONS

The fol lowing i t e m s l i s t the assumptions used f o r t he nominal p e r - formance c a l c u l a t i o n s :

(1) Conf igura t ion aerodynamic d a t a a re obta ined from r e f e r - ences 1 and 2 and a r e conta ined i n Appendix C .

( 2 ) Vehic le weight da ta are taken from r e f e r e n c e 3 and a r e conta ined i n t a b l e s 2 through 5 of Appendix C .

( 3 ) Launch from Kennedy Space Center (KSC), Pad 3 4 , g e o d e t i c l a t i t u d e = 2 8 ' 3 1 ' 1 7 . 5 0 6 4 " , and g e o d e t i c l ong i tude = - 8 0 ' 3 3 ' 4 0 . 8 8 6 9 " . F i r i n g azimuth = 7 0 " measured from n o r t h t o s o u t h over eas t .

( 4 ) A l l s t a g e s a r e f i l l e d t o p r o p e l l a n t c a p a c i t y and T/WO1 and t r a j e c t o r y shaping opt imized t o y i e l d maximum payload as a f u n c t i o n of miss ion energy.

(5) The v e h i c l e l i f t e d o f f w i t h a v e r t i c a l r i s e of twelve seconds . A cons tan t p i t c h r a t e i s i n i t i a t e d and executed u n t i l t h i r t y - f i v e seconds of f l i g h t when t o t a l ang le of a t t a c k i s s e t t o ze ro f o r t he remainder of t h e f i r s t s t a g e f l i g h t .

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(6) The f i r s t s t a g e exerc ised an engine shutdown sequence of one - fou r w i t h a four-second i n t e r v a l .

(7) Three and e igh t - t en ths seconds c o a s t is allowed between the f i r s t s t a g e f i n a l cu t -of f and second s t a g e i g n i t i o n . The atmosphere i s dropped from the c a l c u l a t i o n s a t second s t a g e i g n i t i o n .

(8) A programmed mixture r a t i o is used dur ing the second s t a g e burn t o inc rease performance and use more p r o p e l l a n t tank volume.

(9) The l a r g e S-IC/S-I1 i n t e r s t a g e is dropped t h i r t y seconds a f t e r S - IC f i n a l cu t -o f f on both conf igu ra t ions .

(10) The Launch Escape System is j e t t i s o n e d t h i r t y - f i v e seconds a f t e r S- IC f i n a l cu t -o f f f o r the Apollo c o n f i g u r a t i o n only.

(11) The nose cone f o r the MSFC Nose Cone c o n f i g u r a t i o n w a s j e t t i s o n e d i n parking o r b i t .

(12) Parking o r b i t a l t i t u d e equals one hundred n a u t i c a l mi l e s excep t f o r the s p e c i a l t r a j e c t o r y p r o f i l e s where the S - I 1 s t a g e p l aces a f u l l y loaded S-IVB/IU/Payload i n t o an optimum a l t i t u d e parking o r b i t as a f u n c t i o n of C3. The optimum a l t i t u d e is de f ined as t h a t which w i l l y i e l d t h e maximum payload whi le dep le t ing the S-IVB s t a g e p r o p e l l a n t s minus r e s e r v e s t o reach the s p e c i f i e d energy l e v e l .

(13) Boi l -of f and a t t i t u d e c o n t r o l l o s s e s , c a l c u l a t e d f o r fou r and one-half hour parking o r b i t coas t , remained cons t an t f o r a l l energy l e v e l s as l i s t e d i n t a b l e s 2 through 5 , Appendix C .

( 1 4 ) Upper s t a g e t h r u s t angies were opt imized v i a the s t e e p e s t a s c e n t technique over a r o t a t i n g 1960 F i sche r E l l i p s o i d e a r t h model w i t h the fou r th -o rde r g r a v i t y func t ion .

(15) F l i g h t performance r e se rves were c a l c u l a t e d equal t o .75 pe rcen t of the t o t a l v e h i c l e c h a r a c t e r i s t i c v e l o c i t y f o r C 3 = 0 ( l o c a l escape) t o 1 pe rcen t of t h e t o t a l v e h i c l e c h a r a c t e r i s t i c ve loc- i t y a t C 3 = 125 km2/sec2. This v a r i a t i o n w a s c a l c u l a t e d by us ing a 30 launch v e h i c l e e r r o r a n a l y s i s a t va r ious va lues of C,.

(16) F l i g h t geometry r e se rves were c a l c u l a t e d equal t o s i x t y meters per second.

(17) Net payload is def ined as the weight forward of the ins t rument u n i t (IU) a t f i n a l i n j e c t i o n .

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111. DESCRIPTION - EXPLANATION OF RESULTS

The r e s u l t s of t h i s s tudy a r e presented i n g r a p h i c a l form. The f i g u r e s a r e s e l f - e x p l a n a t o r y b u t c a r e is t o be used i n e x t r a c t i n g d a t a , as explained i n Sec t ion I V .

Figures A - 1 and A - 2 a r e drawings of t he Apollo c o n f i g u r a t i o n and t h e MSFC Nose Cone c o n f i g u r a t i o n , r e s p e c t i v e l y . It w i l l be no t i ced t h a t t o t a l he ight of the MSFC Nose Cone c o n f i g u r a t i o n has n o t been s p e c i f i e d because t h i s w i l l depend on the payload packaging procedure used. For the purpose of t h i s s tudy , no c y l i n d r i c a l payload s e c t i o n w a s assumed; i . e . , the nose cone i s a t t ached d i r e c t l y t o t h e ins t rument u n i t and i s j e t t i s o n e d i n parking o r b i t . The e f f e c t on payload of t he requi rement of a c y l i n d r i c a l payload f a i r i n g w i l l be d i scussed i n Sec t ion IV. Fig- ures A-3 and A-4 d i s p l a y the n e t payload a t i n j e c t i o n f o r t he Apollo and MSFC Nose Cone c o n f i g u r a t i o n s , r e s p e c t i v e l y , as a f u n c t i o n of C3 f o r t he two types of upper s t a g e propuls ion systems i n v e s t i g a t e d . The s o l i d payload curves of the graphs denote miss ion p r o f i l e s where the s-IVB s t a g e is s u b o r b i t a l l y burned f o r i n j e c t i o n i n t o o r b i t and r e i g n i t e d t o r e a c h the des i r ed C3 v a l u e . The dashed p o r t i o n of each curve denotes p r o f i l e s where the S-I1 s t a g e i n j e c t s i n t o an optimum a l t i t u d e c i r c u l a r park ing o r b i t and the S-IVB performs a s i n g l e burn t o o b t a i n the f i n a l C3 va lue . F igures A-5 and A - 6 a r e p l o t s of optimum T/WO1 a t l i f t o f f ve r sus C3 f o r bo th conf igu ra t ions w i t h 5 -2 and 5-2s p ropu l s ion , r e s p e c t i v e l y . F igu re A - 7 d i s p l a y s the r a t i o of S-IVB f i r s t burn i n t o park ing o r b i t p r o p e l l a n t s t o the t o t a l s t a g e p r o p e l l a n t capac i ty .

Figure A - 8 is a p l o t of optimum o r b i t a l t i t u d e ve r sus C3 f o r v a l u e s of C3 g r e a t e r than 105 km2/sec2 f o r J-2s propu l s ion and 125 km2/sec2 f o r s t anda rd 5 - 2 propuls ion systems. The weight sav ings r e s u l t i n g from removing the requirement f o r r e s t a r t c a p a b i l i t y from the S-IVB s t a g e have n o t been added t o the payloads shown, b u t any S-IVB s t a g e weight r e d u c t i o n is d i r e c t l y a d d i t i v e t o payload.

IV. EXCHANGE RATIOS

Exchange r a t i o s of payload w i t h r e s p e c t t o v a r i o u s v e h i c l e parameters are presented i n f i g u r e s B-1 t o B-12. These exchange r a t i o s a r e a p p l i c a b l e only t o f l i g h t p r o f i l e s where t h e S-IVB s t a g e i s s u b o r b i t a l l y burned i n t o park ing o r b i t and r e i g n i t e d t o r e a c h t h e d e s i r e d C3 v a l u e .

Figure A-7 shows t h a t the upper c3 l i m i t of t h e exchange r a t i o s f o r conf igu ra t ions w i t h J-2s propuls ion is c3 = 105 km2/sec2 and C 3 = 125 km2/sec2 f o r the conf igu ra t ions us ing the s t anda rd 5-2 propu l s ion sys tem. The payload e f f e c t s of t he v a r i o u s exchange r a t i o s a r e a d d i t i v e w i t h i n

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t h e s p e c i f i e d l i m i t s and a r e a p p l i c a b l e t o both conf igu ra t ions w i t h e i t h e r 5 -2 o r J-2s propuls ion systems. The excep t ion , n a t u r a l l y , is t h e exchange r a t i o s and e f f e c t s of t he nose cone and payload shrocd . F igu res B - 9 and B-10 a r e no t a p p l i c a b l e t o the Apollo conf igu ra t ion .

The exchange r a t i o s f o r j e t t i s o n i n g the nose cone and shroud weights (where d e s i r e d ) a r e s p e c i a l cases where much c a r e i s t o be used. I f t he nose cone is j e t t i s o n e d a t an a l t i t u d e be fo re o r b i t i n s e r t i o n , t he payload g a i n can be c a l c u l a t e d by m u l t i p l y i n g the percentage of payload g a i n f o r t he g iven C3 va lue from f i g u r e B - 9 and t h e nominal payload from f i g u r e A-4. I f an a d d i t i o n a l weight g r e a t e r than the nominal 2700- pound nose cone is t o be j e t t i s o n e d , t he e f f e c t can be c a l c u l a t e d by m u l t i p l y i n g the d i f f e r e n c e i n t h i s weight and the 2700-pound nose cone weight and the r e s p e c t i v e exchange r a t i o from f i g u r e B-10 and s u b t r a c t - ing t h i s from the p rev ious ly ca l cu la t ed payload. For a n increased weight j e t t i s o n e d i n parking o r b i t , the payload l o s s due t o shroud weight g r e a t e r than 2700 pounds i s then t o be s u b t r a c t e d from the nominal payload from f i g u r e A - 4 . These types of c a l c u l a t i o n s , when a p p l i c a b l e , m u s t be performed be fo re apply ing the o the r exchange r a t i o s . When e x e r c i s i n g a n increment i n launch azimuth ( f i g u r e B-12), a l l e f f e c t s of v e h i c l e p e r t u r b a t i o n s (propuls ion , weights , e t c . ) must be exe rc i sed b e f o r e c a l c u l a t i n g the payload increment due t o a launch azimuth v a r i a t i o n .

V. CONCLUSIONS

The r e s u l t s of t h i s s tudy show t h e Sa tu rn V th ree - s t age launch v e h i c l e capable of handl ing s i z e a b l e payloads f o r h igh energy escape mis s ions . The def ined v e h i c l e d i s p l a y s a c a p a b i l i t y of i n j e c t i n g approx- ima te ly 25,000 pounds i n t o a J u p i t e r probe t r a n s f e r miss ion (C, = 80 km'/ sec') us ing t h e s tandard 5-2 propuis i on sys tems . system f o r t h i s energy r e s u l t s i n a payload g a i n of approximately 3,400 pounds.

The J- 25 p r o p u i s i w i i

A l l miss ion p r o f i l e s d i sp layed w i l l probably r e q u i r e a r e l o c a t i o n of t he s l o s h b a f f l e s i n the S-IVB s t a g e because of t he two-burn propel - l a n t s p l i t v a r i a t i o n as d i sp layed i n f i g u r e A-7. Mission ene rg ie s which r e q u i r e a f u l l y loaded S-IVB s t a g e burned from park ing o r b i t a r e f e a s i b l e by i n j e c t i n g the f u l l y loaded S-IVB/IU/Payload i n t o an optimum a l t i t u d e o r b i t w i t h t h e S-I1 s t a g e . This type of p r o f i l e w i l l r e q u i r e on ly a s i n g l e burn S-IVB, and the weight saved by removing the r e s t a r t from the S-IVB can be converted i n t o payload.

5

The exchange r a t i o s d i s p l a y a need f o r g r e a t e r t h r u s t l e v e l s and h ighe r s p e c i f i c impulse va lues f o r a l l e n e r g i e s . candida te f o r a n increased t h r u s t i s t h e S - I 1 s t a g e , wh i l e t he b e s t candida te f o r a s p e c i f i c impulse i n c r e a s e is t h e S-IC s t a g e .

The most worthy

The J-2s p ropu l s ion system (now be ing i n v e s t i g a t e d by NASA p lanne r s ) shows a good p o s s i b i l i t y of i n c r e a s i n g t h e S a t u r n V payload c a p a b i l i t y . m e n app l i ed t o the S-I1 and S-IVB s t a g e s , payload g a i n s range from 6 percen t a t C3 = 0 ( l o c a l escape) t o 17 p e r c e n t a t C3 = 100 km2/sec2.

6

APPENDIX A

Performance Results

7

43

L e g e n d : A S e p a r a t i o n

4241 F i e l d Spl ice

154"

8'

3890 3841 3750 3595 3593

3340 3259 3223 3101

363'

260" D i o -

2387 396" Dia -

(Gimbal Sta)

FIG. A - I . SATURN Y LAUNCH VEHICLE, APOLLO CONFIGURATION

W

Legend: A Separation

Field Splice

I 1 I I I I I '

260" Dia 2032

2747 * 16 0 44" 2791

-L 2646 (Gimbal S t a 1 + 2519 -2307

- 396" Dia

1048 --.b 1 7 6 0 - 1 6 6 4 (Gimbal S t o 1 -& 4564 -k+ 1541

1 4 0 1

-

-

91 2 -

-365

b- 400 ( G i m b a l Sta 1 q?:;: F I G . A-2. S A T U R N 3L L A U N C H V E H J C L E

M S F C NOSE C O N E C O N F I G U R A T I O N

9

Net .Payload ( I O 3 Ib) T/WO4 = Variable A, = 700

V I I I

0 sb 160 - 0

C, (km2/sec2)

F IG . A-3. L A U N C H V E H I C L E P E R F O R M A N C E FOR S A - 5 1 6 APOLLO C O N F I G U R A T I O N V S C 3

10

ayload (IO3 Ib) T/WOf = V a r i a b l e A, = 70'

Net

11 0

100

90

80

70

60

50

40

3c

2c

10

0

FOR S A - 5 1 6

2700 Ib Nose Cone Jettisoned in Orbit

FIG. A-4. LAUNCH VEHICLE PERFORMANCE MSFC NUSE CONE C UN F I G U KATIUN

11

0 0 w II

-2

\

1 2

0 0 (c

II

N 4

! c i

13

I- m n \ z

n 0 0 0 0 z v) cr rc) (u

7 0

14

c 0

0 0 7

15

APPENDIX B

Exchange Ratios

1 7

I

t- v) 3

I a

w c3

I- v)

a

rc,

18

I -

0 0 0 0 0 0 W u- cu

0 a- - 0 - Q )

v) \ a E -L Y

13 V

0 0 r

0 v)

w cn J => e =E

V LL V w

- -

e t-

LL w

0

w LL

19

I

N 0

\ I I I I

I LL

20

N E

0

z 0

0 W v)

LL 0

I- o W LL L L W

21

I I

f

22

\ L

23

0 0) v)

1 I

2.5

2 .o

---m 0

/ w

8 0

0

FIG. B-9, EFFECT OF NOSE CONE DROPPED AT EARLY ALTITUDE

VERSUS C, FOR MSFC NOSE CONE CONFIGURATION

.5

0.

26

Note : Nominal Shroud Weight = 2700 Ib Dropped in Orbit

I I I

II II II

I I I

0 0 v

0

W

W

27

II I 1

7

v) >

z 0

t- z W

-I Q.

x Q.

-

Lr, 0

t- u W Lr, L w

28

APPENDIX C

Vehicle Data

31

t I

a

0" v z 0

w v

"0

z o O i 2 % L

32

0

Q, rc) I

to to I

0 U

0

W U

I

I

a

a

a

.. .c a3

33

TABLE 1

Propuls ion C h a r a c t e r i s t i c s f o r S a t u r n Launch Vehic le SA-516

Thrus t /Eng ine (1b)

S p e c i f i c Impulse

(set>

Mixture R a t i o

Wf /wo

Number of Engines

Engine Des igna t ion S tage

s-IC

s-I1

s-IVB

s-I1

s-IVB

-

1 , 5 2 2 , OOO+

204,080

228,915

189,113

~

2.27: 1 5 264.5;: F- 1

5 J- 2 425

422.1

426.47

5.0: 1

5 . 5 : l

4 . 7 : l

1 J- 2 205,000 427 5 . 0 : l

5 J- 2s 430.5

427.6

431.97

5 . 0 : 1

5.5: 1

4.7: 1

204,080

228,915

189,113

205 , 000 432.5 5.0: 1 1 J- 2s

$:Denotes Sea Level Values.

34

TABLE 2

Weight Surmnary f o r Sa turn Launch Vehicle SA-516 Apol lo Conf igura t ion wi th Standard 5-2 Engines i n S-II/S-IVB Stages

S t a g e I t e m Weight

s-IC Mainstage P r o p e l l a n t Capac i ty N2 Purge ( L i f t o f f t o Cu to f f ) S - I 1 I n s u l a t i o n Purge Gas F r o s t (Tota l ) Inboard Engine Thrus t Decay P r o p e l l a n t Outboard Engines Thrust Decay P r o p e l l a n t S tage a t Separa t ion S-IC/S-11 I n t e r s tage (small)

4,598,260 32

120 1 , 400 1 , 770 6,760

325,013 1 , 400

s-I1 Ul lage Rocket P r o p e l l a n t Thrus t Buildup P rope l l an t S-IC/S-11 I n t e r s t a g e (Large) Launch Escape Sys tern Mainstage P rope l l an t Capac i ty W/PMR Thrus t Decay P rope l l an t S tage a t Separa t ion S-II/S-IVB I n t e r s t a g e S-IVB A f t Frame (Separated w i t h I n t e r s t a g e )

2,720 1 ,836

8,300 970,000

3 60 93 , 031

7 , 682 48

9,220

s-IVB Ul l age P r o p e l l a n t

First Thrus t Buildup P r o p e l l a n t Eurn Ullage Rocket Gases

H2 i n S t a r t Tank

APS P r o p e l l a n t - Power Ro l l ( F i r s t Burn) Thrus t Decay P rope l l an t

Orbi ta l P r o p e l l a n t Below Engine Valve Coast H 2 + He Vented i n O r b i t

APS P r o p e l l a n t Used i n O r b i t LOX/Hydrogen Burner P r o p e l l a n t Oxid izer Vented in O r b i t

Second H2 i n S t a r t Tank Burn Thrus t Buildup P rope l l an t

Thrus t Decay P r o p e l l a n t To ta l Mainstage Capaci ty ( I n c l . Reserves+;) S tage a t Separa t ion Instrument Un i t

122 4

360 127 18 94

39 3,016

438 16

130

6 360

94 230,000

26,108 *, U J U I nrn

.L

Reserves c a l c u l a t e d as f u n c t i o n of mis s ion p r o f i l e .

35

TABLE 3

Weight Summary f o r Sa tu rn Launch Vehicle SA-516 Apollo Conf igura t ion w i t h J-2s Engines i n S-II/S-IVB Stages

S tage Item Weight

s-IC Mainstage P r o p e l l a n t Capaci ty 4,598,260 N2 Purge ( L i f t o f f t o Cu to f f ) 32 S - I1 I n s u l a t i o n Purge Gas 120 F r o s t (To ta l ) 1 ,400 Inboard Engine T h r u s t Decay P r o p e l l a n t 1 , 770 Outboard Engines Thrus t Decay P r o p e l l a n t 6,760 Stage a t Sepa ra t ion 325,013 S-IC/S-11 I n t e r s t a g e (Small) 1 , 400

s-I1 Ul lage Rocket P r o p e l l a n t Thrus t B u i l d u p P r o p e l l a n t S-IC/S-11 I n t e r s tage (Large) Launch Escape Sys tem Mainstage P r o p e l l a n t Capaci ty W/PMR Thrus t Decay P r o p e l l a n t Stage a t Sepa ra t ion S-II/S-IVB I n t e r s t a g e S-IVB A f t Frame (Separated w i t h I n t e r s t a g e )

2,720 1 ,836 9 220 8 , 300

970,000 3 60

89,931 7,682

48

s-IVB Ul lage P r o p e l l a n t 1 2 2 H2 i n s t a r t Tank 4

APS Propellant-Power Rol l ( F i r s t Burn) 18 Thrus t Decay P r o p e l l a n t 94

F i r s t Thrus t B u i l d u p P r o p e l l a n t 360 Burn Ullage Rocket Cases 1 2 7

O r b i t a l P r o p e l l a n t Below Engine Valve Coast H2 + He Vented i n O r b i t

APS P r o p e l l a n t Used i n O r b i t LOX/Hydrogen Burner P r o p e l l a n t Oxidizer Vented i n O r b i t

39 3,016

438 16

130

Second H2 i n S t a r t Tank 6 Burn Thrus t B u i l d u p P r o p e l l a n t 3 60

Tota l Mains tage Capaci ty ( I n c l . Reservesf:) 230,000 Stage a t Sepa ra t ion 25 208 Ins t rument Uni t 4,050

Thrus t Decay P r o p e l l a n t 94

-1.

"Reserves c a l c u l a t e d as f u n c t i o n of mis s ion p r o f i l e .

36

TABLE 4

S tage

s - I C

Weight Summary f o r Sa turn Launch Vehic le SA-516

i n S-II/S-IVB Stages MSFC Nose Cone Conf igura t ion w i t h Standard 5-2 Engines

Item Weight

Mains t age P r o p e l l a n t Capaci ty 4,598,260 N2 Purge ( L i f t o f f t o Cu to f f ) 32

F r o s t (Tota l ) 1 ,400 Inboard Engine T h r u s t Decay P r o p e l l a n t 1 ,770 Outboard Engines Thrust Decay P r o p e l l a n t 6,760 Stage a t Sepa ra t ion 325,013 S-IC/S-11 I n t e r s t a g e (Small) 1 ,400

S - I 1 I n s u l a t i o n Purge G a s 120

s-I1 Ul lage Rocket P rope l l an t Thrus t Buildup P r o p e l l a n t S-IC/S-11 I n t e r s tage (Large) Mainstage P r o p e l l a n t Capaci ty W/PMR Thrus t Decay P r o p e l l a n t S t age a t Sepa ra t ion S-II/S-IVB I n t e r s tage S-IVB A f t Frame (Separated w i t h I n t e r s t age )

2,720 1,836

970,000 360

93,031 7,682

48

9,220

s-IVB Ullage P r o p e l l a n t

F i r s t Thrus t Buildup P r o p e l l a n t Burn Ul lage Rocket Cases


APS P r o p e l l a n t - Power Ro l l ( F i r s t Burn) Thrus t Decay P r o p e l l a n t

Orb i t a 1 P r o p e l l a n t Below Engine Valve Coas t H2 + He Vented i n Orb i t

APS P r o p e l l a n t U s e d i n O r b i t LOX/Hydrogen Burner P r o p e l l a n t Oxid izer Vented i n O r b i t Payload F a i r i n g

Second Burn

H2 i n S t a r t Tank Thrus t Buildup P r o p e l l a n t Thrus t Decay P rope l l an t To ta l Mains t age Capaci ty ( I n c l . Reserves;';) S t age a t Separa t ion i n s i r u n i e r i i V u i i -

122 4

360 1 2 7 i8 94

39 3,016

438 16

130 2,700

6 360

94 230,000

26,lO.a 4 , n5n

-1,

Reserves c a l c u l a t e d as func t ion of mis s ion p r o f i l e .

37

TABLE 5

Weight Summary f o r Sa tu rn Launch Vehic le SA-516 MSFC Nose Cone Conf igura t ion w i t h J-2s Engines i n S-II/S-IVB Stages

S tage Item

s-IC Mainstage P r o p e l l a n t Capac i ty N2 Purge ( L i f t o f f t o C u t o f f ) S-I1 I n s u l a t i o n Purge Gas F r o s t (To ta l ) Inboard Engine Thrus t Decay P r o p e l l a n t Outboard Engines Thrus t Decay P r o p e l l a n t S tage a t Sepa ra t ion S-IC/S-11 I n t e r s tage ( sma l l )

Weight

4,598 , 260 3 2

120 1 , 400 1 , 770 6,760

325 , 013 1,400

s-I1 Ul lage Rocket P r o p e l l a n t 2,720 1 , 836 Thrust B u i l d u p P r o p e l l a n t

Mains tage P r o p e l l a n t Capac i ty W/PMR 970,000 Thrus t Decay P r o p e l l a n t 360 Stage a t Sepa ra t ion 89 , 931 S-II/S-IVB I n t e r s tage 7 , 682 S-IVB A f t Frame (Separated w i t h I n t e r s t a g e ) 48

S-IC/S-II I n t e r s t a g e (Large) 9,220

s-IVB Ullage P r o p e l l a n t

F i r s t Thrus t Buildup P r o p e l l a n t Burn Ullage Rocket Cases


APS P r o p e l l a n t - Power Rol l ( F i r s t Burn) Thrus t Decay P r o p e l l a n t

Orb i t a 1 P r o p e l l a n t Below Engine Valve Coast H2 + He Vented i n O r b i t

APS P r o p e l l a n t Used i n O r b i t LOX/Hydrogen Burner P r o p e l l a n t Oxidizer Vented i n O r b i t Payload F a i r i n g

Second Burn

H2 i n S t a r t Tank T h r u s t Buildup P r o p e l l a n t Thrust Decay P r o p e l l a n t To ta l Mainstage Capac i ty ( I n c l . Reserves;';) Stage a t Sepa ra t ion Instrument Uni t

1 2 2 4

360 1 2 7 18 94

39 3,016

428 16

130 2,700

6 360

94 230 , 000

25 , 208 4,050

-1.

Reserves c a l c u l a t e d as f u n c t i o n of miss ion p r o f i l e .

38

.

REFERENCES

1. " S t a t i c Aerodynamic C h a r a c t e r i s t i c s of t h e Apol lo-Sa turn V Vehicle ," NASA TM X-53517, Marsha l l Space F l i g h t Cen te r , September 16 , 1966.

2. " S t a t i c Aerodynamic C h a r a c t e r i s t i c s of t h e S a t u r n V Voyager (45' Shroud Length) ,I1 MSFC, R-AERO-AD-66-59, December 20, 1966.

3. " J e t t i s o n Weight Summaries f o r S a t u r n SA-516 Vehic le f o r Various Conf igu ra t ions , " MSFC, R-P&VE-AV-66-193, November 10, 1966.

4. B a t t l e , C . Tucker and Robert G. G o t t l i e b , "Opt imiza t ion of Mul t i - s t a g e Three-Dimensional Boost T r a j e c t o r i e s , " Raytheon Reports BR-2758-1 and BR-2758-2, May 1, 1964.

5. G o t t l i e b , Robert G. , Malcolm F r a z i e r and Richard Moroney, "Updating and Checkout of t h e BOP-01 Computer Program," Raytheon Report FR-66- 87-2, March 11, 1966.

39

APPROVAL NASA TM X-53639



Ronald G. Toe l l e

The information i n t h i s r e p o r t has been reviewed f o r s e c u r i t y c l a s s i f %cat ion . Review of any informat ion concerning Department of Defense or Atomic Energy Commission programs has been made by the MSFC Secur i ty C l a s s i f i c a t i o n O f f i c e r . This r e p o r t , i n i t s e n t i r e t y , has been determined t o be unc lass i f i ed .

This document has a l s o been reviewed and approved f o r t e c h n i c a l accuracy .

n. h&,% W. D . Goldsby, 4r.u Chief , Fl ight-Mechanics and Performance Analysis Sec t ion

H. F. Thomae Ch ie f , Advanced S tud ie s Off i c e

E . D. Geiss le r D i r e c t o r , Aero-Astrodynamics Laboratory

40

.

D I S T R I B U T I O N

D I R - D r . von Braun D r . R e e s

R - D I R M r . Weidner

R - A S - M r . Williams M r . Becker M r . C a r t e r M r . H u b e r M r . Spea r s (2) M r . Page M r . D a v i e s (2 ) M r . P e r r y M r . Madewell M r . D a n n e n b e r g M r . von Tiesenhaussen M r . N e i g h b o r s M r . Harris

R- P&VE D r . Lucas M r . M a r m a n n M r . M c K a y M r . Black Xr. Paul M r . G o e r n e r M r . O r i l l i o n M r . A l l e n M r . B a r r e t t M r . T h o m s o n M r . She l ton M r . N i x o n

R - A S T R D r . Haeus ermann M r . D i g e s u M r . D a n i e l

R- R P - D r . D r . M r .

IO - D r . D r .

S t u h l inger Shel ton Wood

R u d o l p h F. S p e e r

R-AERO D r . G e i s s l e r M r . Jean M r . H o r n M r . Thomae M r . S c o t t M r . E . T. D e a t o n ( 2 ) M r . C r e m i n M r . von P u t t k a m e r M r . G o l d s b y (15) M r . Toe l le (5) M r . Wood M r . Y o u n g M r . E l l i s o n M r . Jump M r . Summeral M r . Wheeler M r . Leonard X r . Smart M r . Perkinson M r . Fulks M r . M c N a i r M r . Lindberg M r . Baker M r . C a u s e y M r . W. V a u g h a n D r . H. K r a u s e

MS- I P M S - I L (8) cc-P I-RM-M MS-H M S - T (6) D E P - T

41

EXTERNAL D I S TRIBU TION

Technical & S c i e n t i f i c Informat ion F a c i l i t y (25) Box 33 Col lege Park, Md. At tn : NASA R e p . (S-AK/RKT)

MS C Houston, Texas 77058 At tn : M r . W. Stoney

M r . P. Thomas L ib ra ry (5)

A p p l i e d Analysis Inc . Holiday Of f i ce Center H u n t s v i l l e , A l a . 35805 At tn : D r . Nes l ine

M r . G o t t l i e b M r . H a l l

D r . Hu Nor th rop Space Labora to r i e s Technology Drive H u n t s v i l l e , A l a . 35805

M r . W. B. Tucker Lockheed Missiles and Space Co. Technology Drive H u n t s v i l l e , Alabama

M r . J . Winch M a i l S top AC-50 Boeing Co. P. 0. Box 1680 H u n t s v i l l e , A l a . 35807

M r . John E . Canady NASA - Langley Research Center Langley S t a t i o n , V i r g i n i a 23365 Mail Stop 215

NASA Headquarters At tn : D r . Seamans, AD

D r . Muel le r A . D. Schnyer, MTU(OMSF) D r . Mac Adams, R(0ART) Col. Burke, MTV (5) L ib ra ry (5)

.

42

EXTERNAL DISTRIBUTION (Continued)

NASA - Ames Research Center Mountain View, C a l i f . 94035 Attn : L ib ra ry ( 5 )

NASA - Langley Research Center Hampton, V i rg in i a 23365 At tn : L ib ra ry ( 5 )

.

43

s- george c. - nasa

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