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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
15
18
19
20
21
22
23
24
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
A PERFORMANCE STUDY FOR THE APPLICATION OF THE
SATURN V TO HIGH ENERGY EARTH ESCAPE MISSIONS
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 pro- gram 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 .
2
(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 .
3
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 pay- load 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 param- eters 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
4
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 pay- load 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 pay- load 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
1 24
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
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 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
H2 i n S t a r t Tank
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
A PERFORMANCE STUDY FOR THE APPLICATION OF THE
SATURN V TO HIGH ENERGY EARTH ESCAPE MISSIONS
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