CoA Note No.88 TECHNISCHE HOGESC
VLIEGTUIGBOUWKUND5 Kanaalstraat 10 - DELFT
2 0 nov.!S58
THE COLLEGE OF AERONAUTICS
CRANFIELD
ATMOSPHERE BREATHING ENGINES IN
ASTRONAUTICS
by
TECHNISCHE HOGt:>CHOOl
VLIEGTUIGBOUWKUNDE
K.n..l.tr.at10 - DELFT ^amm^_8Q^
OctöbsT, 1938.
T H E C O L L E G E O F A E R O W A U T I _C_ S C R A N F I E L D
Atmosphere Breathing Enp;ines i n Astronautics P a r t I . P l i g h t in t h e E a r t h ' s atmosphere
b y
-S.W. Greenvrood, B.Sc. (Eng.), M.Eng., A.M.I.Mech.E., A.P.R.Ae.S.
P a r t I I , P l i g h t i n t h e atmospheres of other p l a n e t s b y
-Dennis S. Carton, A.F.R.Ae.S,
The contents of t h i s note f a l l into tv/o s e c t i o n s . P a r t I considers the p o s s i b i l i t i e s arxL problems involved i n using ramjets as a poT/er source for one of t h e stages of a s a t e l l i t e launching vehicle or s i m i l a r p r o j e c t . I n comparing such a system vrlth rocket pcv/ered v e h i c l e s , consideration i s
given to both performance and mass of the various systems. Various t r a j e c t o r i e s are considered. This work includes a reassessment of p r o j e c t s that have been suggested elsevdiere.
The second p a r t examines the p o s s i b i l i t y of using forms of ramjet i n the atmosphere of other p l a n e t s . Because there i s insxjfficient knoY/ledge of t h e s e atmospheres, a study has been c a r r i e d out t o determine t h e approximate performance of a chemical rajnjet in atmospheres of Methane, Ammonia, Hydrogen ajnd Carbon Dioxide a t Mach 3. The vrork in P a r t I I i s o r i g i n a l , t h e r e being no previously reported papers on t h e subject known.
These s t u d i e s , which are n e c e s s a r i l y based on several simplifying assumptions, indicate t h a t a p p l i c a t i o n s f o r these engines may be expected to a r i s e i n a s t r o n a u t i c s , and t h a t t h i s i s a f r u i t f u l f i e l d for further s t u d i e s .
Peper delivered t o t h e Midlands Branch of the B r i t i s h I n t e r p l a n e t a r y Society on Saturday, 9ch November, 1957<
COMEIMTS Page Sxamnaiy PART I 1.1. Introduction 1.2. Engine Comparisons
1.3. Influence of flight programme 1.2f, Satellite vehicles
1,5. Conclusion
PART II
2 . 1 , Introduction
2 . 2 , Atmospheres of the Solar System 2 . 3 , Choice of Engine for Study
2,1^ Performance i n other Atmospheres
2 , 5 , Conclusions Referenc es Appendix I Appendix I I 1 i 1 2
3
4
5
5
5
6
7 89
10 14« i «
EAKr_I 1.1. I n t r o d u c t i o n
The rocicet engine has long been the favourite means of propulsion among those eartying out paper studies of possible space p r o j e c t s , and indeed i t i s one of the fev7 systems capable of operating i n the n e a r -vacuum of space. As i s v/ell knovm, i t s mor>t severe l i m i t a t i o n i s i t s high s p e c i f i c prcpel]ant consxjmption, and i t i s of i n t e r e s t t o explore the p o s s i b i l i t y of ^osing atmosphere-breathing engines, which are more e f f i c i e n t f ion t h i s point of view, during periods of f l i g h t i n p l a n e t a r y • atmospheres. I n t h i s p a r t of the paper we are concerned with f l i g h t i n the E a r t h ' s atmosphere, and t h e r e f o r e , p r i m a r i l y , with entry into E a r t h s a t e l l i t e oi'bits.
Several w r i t e r s have looked into t h i s problem, and t h e i r views w i l l be examined. That the scheme i s worth our consideration i s indicated by t h e p r a c t i c e adopted as e a r l y as 1947 i n the United States of launching experimental rocket-engine powered re-search a i r c r a f t from conventional a i r c r a f t powered by a i r - b r e a t h i n g p i s t o n engines. I f the p i s t o n engine 1-jas been found useful, s u r e l y t h i s -vvill prove t o be t h e case al>jO with a i r - b r e a t h i n g engines now being developed for f l i g h t at high supersonic speeds.
1.2. Engine Comparisons
If/hen making paper comparisons of vehicles fitted with different
types of engine it is customary to take "typical" performance data for each class of engine. This is a convenient but rather dangerous technique, for
the results finally obtained are no more re].iable than the original assxjniptions. Comparisons of total vehicle weight, for example, should be regarded as
having an elastic property. Sometimes apparently relatively small changes in engine performance will have the effect of reversing a previoxisly drawn
conclusion. Having sounded a note of caution, it is nevertheless of interest to attempt such comparisons in the light of engineering data currently
available.
While there are many interesting hybrids and engine ccmbinations possible, it will be assumed for simplicity that only three main engine types will be used, and only one engine type per vehicle stage. The engines considered are the rocket, ramjet and turbojeto
The rocket operates by discharging material initially carried by the vehicle. It is thus independent of any surrounding atmosphere. The ramjet and tiirbojet are atmosphere-breathing engines, the thrust being obtained by discharging atmospheric gas (together with injected fuel in some cases) at a higher velocity than it vras entrained. The principles of these engines are fairly well understood and will not be dealt with at length here. A ramjet for supeirsonic operation is shown in Pig. 1 . The supersonic turbojet would look rather similar to this with a
compressor-2
t u r b i n e set interposed betvreen the intake and exhaust nozzle.
A comparison of different engine types for f l i g h t i n the E a r t h ' s atmosphere has been d e a l t rn.th i n a number of papers. ( l , 2, 3 ) . A i r
-breathing engines can ohl.y be operated within a l i m i t e d range of altitu5.es and speeds. P i g . 2 shows i n g e n e r a l i s e d form t h e " l i m i t a t i o n s " of ramjet operation. The top point of t h e range might be put a t Mach 5 a t 120,000 f t . The t\jrbo-jet ejq^eriences s i m i l a r r e s t r i c t i o n s , the main difference being t h a t i t produces thrijst from r e s t . The rocket as an engine experiences no altitude-xspeed r e s t r i c t i o n , although a t lav<r a l t i t u d e s a rocket v e h i c l e would be subject t o speed l i m i t a t i o n s due t o aerodynamic heating a t high
speeds jxost as any other vehicle would,
For sustained f l i g h t i n t h e atmosphere a t a constant speed. P i g . 3 gives an approximate idea of the f l i g h t time a t Tfhich a ramjet powered v e h i c l e wcjuld be l i g h t e r than a rocket powered v e h i c l e . (Assunrptions and
c a l c u l a t i o n s are given i n Appendix I . ) At low a l t i t u d e s t h i s occ\jrs a f t e r only a few seconds ! At high a l t i t u d e s i t may take a minute or more. The reason for t h i s i s t h a t a t high a.ltitudes a ramjet engine produces l e s s t h r u s t for a given T/eight, and t h e r e f o r e a longer period of time has t o
elapse before the r e l a t i v e l y l a v s p e c i f i c propellant consiornption of the engine r e s u l t s in an advantage. Longer fliglrit times are requii^ed f o r the t u r b o j e t than for the ramjet, as the ttirbojet i s a heavier engine.
I t i s again necessary t o advise caution i n i n t e r p r e t i n g information such as t h a t shewn on Pig. 3 . However, i t does b r i n g out one of the main f e a t u r e s of r o c k e t - a i r - b r e a t h i n g engine comparison.
"Vlhile i t i s obviously d e s i r a b l e t o aim a t a low f i n a l v e h i c l e weight, a great numbe-r of other f a c t o r s have t o be considered i n p r a c t i c e . These include c o s t , r e l i a b i l i t y , ease of development, safety and s i m p l i c i t y . I n t h i s st\3dy we s h a l l consider mainly comparisons of vehicle weight. 1.3. Influence cf f l i g h t programme
For an adequate a n a l y s i s , proper considera.tion must be given t o s u i t a b l e f l i g h t paths and programming. This i s too long a job for a single worker, however e n t h u s i a s t i c , and a b r i e f atteript has been made t o deal with t h i s by considering two a l t e r n a t i v e f l i g h t paths together with r a t h car a r b i t r a r y a c c e l e r a t i o n s . I t i s hoped however t h a t t h i s w i l l serve as an i n t r o d u c t i o n to t h e problem.
The simplest and most obvious path t o consider f i r s t i s t h a t of v e r t i c a l a s c e n t . The onlsr paper Iciovm t o the author dealing with t h i s i s by T s i e n ' 4 } , His analysis has a number of r a t h e r severe l i m i t a t i o n s
in p a r t i c u l a r he allows a ramjet povrered v e h i c l e t o produce t h r u s t from r e s t , -a condition not r e -a l i s -a b l e i n p r -a c t i c e !
J
The method s e l e c t e d by the author v/as t o examine t h e t h r u s t produced by a ramjet engine of given s i z e during ver-fcical f l i g h t assuming i n t u r n the conditions l i s t e d an P i g . 4 . The assijraption of constant a c c e l e r a t i o n was made t o e s t a b l i s h the s p e e d - a l t i t u d e v a r i a t i o n . I t i s c l e a r t h a t the r e s u l t i n g t h r u s t v a r i a t i o n s therefore imply t h a t an a d d i t i o n a l and v a r i a b l e t h r u s t producing system rrould be necessary t o give t h e corxstant a c c e l e r a t i o n assumed. However, an approximation may be irade by taking the engine t h r u s t as s u b s t a n t i a l l y constant over the middle p a r t of the speed rangCf
(Reduction i n fuel weight vd.ll, i n any event, tend t o compensate f o r fallx-ig t h r u s t in the upper p a r t of the range.)
Conditions to which the engine wovild be subjected during a tjjpical f l i g h t axe shovm i n F i g . 5- The convention of p l o t t i n g a l t i t u d e v e r t i c a l l y , though i t i s the independent v a r i a b l e , has been adopted. Conditions a t a given a l t i t u d e are obtained by reading across to t h e appropriate curve and then down to t h e relevant l i n e . I t T/d-ll be appreciated t h a t mechanical troubles a r i s i n g from high ram tonperatures w i l l increase with a l t i t u d e , although t h i s i s t o some extent a l l e v i a t e d by the f a l l i n ram pressure. The f l i g h t v e l o c i t y p r o g r e s s i v e l y increases, although the f l i g h t Mach No. f a l l s off f o r a while a t t h e upper end of the range ovdng t o the increase in atmospheric temperature above the s t r a t o s p h e r e .
The constant t h r u s t t r a j e c t o r y r e f e r r e d t o on P i g . 4 cannot be obtained i n v e r t i c a l f l i g h t •vvdthout a sharp increase in a c c e l e r a t i o n a.t medium a l t i t u d e s . This may be seen from Pig. 6 viiiere the required
speed-a l t i t u d e v speed-a r i speed-a t i o n i s given. Avoidspeed-ance of speed-a f speed-a l l off i n engine t h r u s t with a l t i t u d e i s c l e a r l y d e s i r a b l e , and i n p r a c t i c e t h i s w i l l c a l l f o r a departure from ve^rtical ascent and some l e v e l l i n g - o f f at medium a l t i t u d e s .
The r e l a t i v e weights of ramjet and rocket powered vehicles over
r e a l i s t i c ramjet speed ranges for constant payloads and for the t r a j e c t o r i e s and acceleraticjn programcies given above are shovm i n F i g . 7. These q u a n t i t i e s have the e l a s t i c p r o p e r t i e s mentioned e a r l i e r , and tend t o penalise the rocket •which i s made t o operate over a l i m i t e d speed range. I t does i n d i c a t e , hcjwever, t h a t ramjet povrered v e h i c l e s , t o be competitive vd.th rocket powered v e h i c l e s ,
(a) need t o be boosted t o speed at low a l t i t \ i d e s , (b) need t o follcrw a t r a j e c t o r y which s u s t a i n s t h e t h r u s t over t h e speed range.
1. if. S a t e l l i t e vehicles
An attempt vra.s made t o canpare vehicles for e s t a b l i s h i n g a fixed
payload i n a c i r c u l a r o r b i t near the E a r t h . (Note - these stiidies were c a r r i e d out p r i o r t o the f i r s t s a t e l l i t e launching.) A conventional t h r e e - s t a g e
rocket was compared id-th a two-stage rcxiket launched from a ramjet powered stage t h a t had been a c c e l e r a t e d t o operational speed by a s o l i d propellant booster. The r e s u l t i n g t o t a l v e h i c l e weights are shovm i n F i g . 8 , The
conventional rocket vehicle i s l i g h t e r and simpler, ha'ving one fev/er stage. One of the reasons f o r t h i s i s t h a t the ramjet stage f i n a l v e l o c i t y i s only about 1 m i l e / s e c , thus forcing upon each of the tv/o stages of i t s "payload" a higher inaremental v e l o c i t y reqioirement than each of the tiiree stages of the
- 2 , - .
ooiriparison rocket v e h i c l e .
A much earlica? i n ' v e s t i ^ t i c n of t h e s a t e l l i t e launching problem isas made by Proell*'-'''. His a n a l y s i s has c e r t a i n wealoiesses i n d e t a i l , but he advocates t h e use of a slow take-off type of space ship as t h e only p r a c t i c a l tjrpe. He examined t h e p o s s i b i l i t y of varying speed and a l t i t u d e using a •winged ramjet ship - t o carry the payload almost t o o r b i t a l v e l o c i t y ! This looks frightening a t present owing t o t h e high ram temperatures t h a t wouM be experienced. Nevertheless the p r i n c i p l e i s sound, and t h e more
the speed range of t h e ramjet can be s t r e t c h e d i n f u t u r e , t h e more i n t e r e s t i n g t h i s proposal becomes.
Possibly one of t h e most i n t e r e s t i n g studies i n recent years was t h a t c a r r i e d out by Sandorff^°J, He considered the problem of e s t a b l i s h i n g a 500 l b , payload i n a c i r c u l a r o r b i t using e i t h e r a conventional t h r e e stage rocket or a two stage rocket launched by an a i r c r a f t powered by turbo j e t s , The speed a l t i t u d e v a r i a t i o n i s shewn i n P i g . 9. The vehicle v/eights are given i n P i g . 10. Aga.in the conventional rocket v e h i c l e i s seen t o be l i g h t e r i n weight. Some important f a c t o r s a r e , hcrtTevcr, emp^iasised by Sandcïrff.
F i r s t l y the weight of t h e expensive rcxjket coinponent i s halved. Secondly, the aeroplane f i r s t stage i s l i k e l y to be more r e l i a b l e than a rocket f i r s t s t a g e , and the o v e r a l l r e l i a b i l i t y of the syston shcTuld be increased. F i n a l l y , t h e aercplane stage i s reco-verable.
I t would appear t h a t t h i a scheme i s an a t t r a c t i v e a l t e r n a t i v e t o t h e conventional rocket scheme, Sandorff envisages developments i n which ramjets are used t o increase t h e f i n a l speed of t h e aeroplane s t a g e .
Some proposals from M. Varvarov^''^of Russia are of i n t e r e s t . His s p e e d - a l t i t u d e v a r i a t i o n f o r a combined turbojet-ramjet-rocket v e h i c l e i s shown in P i g . 9. Engines would be j e t t i s o n e d a t t h e end of each phase of operation. As an a l t e r n a t i v e he suggests a s e r i e s of vehicles each with i t s ovm type of power p l a n t . His s e l e c t i o n of a l t i t u d e s a t v/hich a given speed i s t o be reached appeor*s t o be on the high s i d e f o r the development of an adequate t h r u s t from the a i r - b r e a t h i n g engines.
1,5. Conclusion
Air breathing engines appear t o offer a reasona.bly a t t r a c t i v e a l t e r n a t i v e t o t h e rocket engine f o r sane a p p l i c a t i o n s of a s t r o n a u t i c a l i n t e r e s t during f l i g h t i n t h e E a r t h ' s atmosphere.
PART I I
2 . 1 , I n t r o d u c t i o n
Into the foreseeable futiore mass 'will continue t o be a most expensive and d i f f i c u l t property t o project any distance av/ay from the Earth. Since, as f a r as. a v e h i c l e i s concerned, t h e production of t h r u s t i s mass consuming i t •will always be imperative t o produce i t i n the most e f f i c i e n t manlier. To d a t e , a l l p r o j e c t e d studies of inter'planetary voyages appear t o have considered some form of rocket engine as an automatic s e l e c t i o n f o r a l l stages of the journey outside cur ovm atmosphere,, Greeriv/ood, i n the f i r s t p a r t of t h i s paper, and o t h ' r s have shown t h a t a i r breathing engines may be able t o play a useful p a r t i n the i n i t i a l stage of a launching programme fran t h e E a r t h ' s surface. Looking f u r t h e r , i n distance and time, t h i s p a r t
i n v e s t i g a t e s the p o s s i b i l i t y of using t h e atmospheres of other p l a n e t s as a means of producing t h r u s t more e f f i c i e n t l y than t h e rocket engine. I t i s thought t h a t suc^h a p o s s i b i l i t y virill be of i n t e r e s t t o those •who may study the f e a s i b i l i t y of j.ourneys involving e i t h e r circ^umnavigation of, or landing
and subsequent t a k e cff from, any of the b a i i e s of t h e solar system possessing considerable atmospheres. See P i g . 1 1 .
2.2, Atmor-oheres of the Solar System
Because of the phj/sical d i f f i c u l t i e s involved in vievidng and analysing the r e s t of the s o l a r system from the bottom of our own atmosphere, t h e
information a v a i l a b l e upon t h e chemical composition of t h e various pla.netary atmospheres i s vague and contra d i e t ctry. Unf ort'uriately, as f a r as t h i s study i s concerned, opinions c o n f l i c t not only upon the amounts of •\/-arious cons^bituents p r e s e n t , but upon the absence, or presence of a particiiLar gas i n some
considerable proportion. There i s in addition p r a c t i c a l l y no information a t a l l upon the temperature, pressures and d e n s i t i e s a t poin'ks v/ithin atmospheric envelopes.
Obviously the comparative method.s discussed i n the f i r s t p a r t of t h i s paper have only been p o s s i b l e because adequate information i s a-\/ailable upon the chemistry and d i s t r i b u t i o n of the E a r t h ' s a-tmosphere. Such r e l i a b l e comparisons cf t h e r e l a t i v e merits of a^bmospheric b r e a t h i n g and rocket
engines for s p e c i f i c applications on other p l a n e t s vn.ll only become p o s s i b l e vrfien our knowledge of t h e i r a'tmospheres becomes as extensive as our knowledge of our own. These difficrulties have forced t h e author t o define, hypcrthetDi^al atmosptercs, anri •liion t o i n v e s t i g a t e the possri.biJ.i^ty of using engines
i n than.
Each hypothetical atmosphere has, therefore, been defined as being made up e n t i r e l y of one gas. The choice of these gases has been governed by the main atmospheric ingredients of the solar system excluding E a r t h . On t h i s point there i s a remarkable lack of disagreement amongst various a u t h o r i t i e s ; i n consequence, hydi-ogen, methane, ammonia and carbon dioxide have been chosen, The f i r s t t h r e e , a l l capable of producing heat when chemically reacted vidth
~ 6
-Carbon dioxide, on the other hand, i s f a r too s t a b l e t o be used i n such chemical processes and so receives separate consideration
The next problam was the s e l e c t i o n of s u i t a b l e temperatiares and pressures fctr these hypothetical atmospheres. To simplify the i n i t i a l i n v e s t i g a t i o n , and t o make t h e r e s u l t s conrparable, i t v/as decided t o use the same temperature and pressure figures for each atmosphere. Since the l a r g e r p l a n e t s , Saturn and J u p i t e r , have considerable atmospheres, t h e y were used t o give reasonable f i g u r e s for t h i s f i r s t t e n t a t i v e vrork, F i n a l l y , 1 23 JK (-150 C) and 14.7 I b / i n ^ were s e l e c t e d as being p o s s i b l e
and f a i r l y tij-pica.!. Only t h i s one spot point vra,s considered; no work has yet been c a r r i e d out on "alti-tvide e f f e c t s " .
2 , 3 . Choice of Engine f o r S^tutüy
I t i s a very i n t e r e s t i n g f a c t ttiat a l l forms of a i r b r e a t h i n g i n t e r n a l combustion engine appear t o be capable of redesign t o enable them t o operate •within any of t h e fuel atmospheres. The main modification recjuired would be t o convert the existing f u e l system so t h a t i t vro'uld meter and inject a s e l e c t e d oxidant, A f u l l study of the thermodynamics a.nd mechanism of such engines f o r e i t h e r ground or "airborne" use v»-ill need much careful consideration a t some time, but i s outside t h e l i m i t s of t h i s paper. I t would appear t h a t the f i r s t \ase t o which atmosphere breathing engines may be put be as p a r t of a "probe" programme. This may involve a-tmospheric
entry, circumnavigation and subsequent e x i t , most probably vd.thout landing. Of the p o s s i b l e types of engine a'vailable, t h e ramjet appears t o b e an
obvious choice for s.uch a scheme. I t was, t h e r e f o r e , decided t o i n v e s t i g a t e t h e p o s s i b i l i t y of using such an engine in t h e hypothetical f u e l atmospheres •with one or two oxi^Sants about which adequate information vra.s a v a i l a b l e . Those studied so far include l i q u i d oxygen, l i q u i d ozone, n i t r i c acid, and dinitrogen t e t r a oxijde.
Engine operation •within a carbon diax:i.de atmosphere requires even more study than the operation of engines v/ithin chemically r e a c t i v e atmospheres. I t i s p o s s i b l e , once again, t o envisage design changes of a l l a i r b r e a t h i n g
engines t o make them vrork under those conditionr. This v/ould involve t h e use of both tanked oxidant and fuel burning in a bi-chemlcal process with t h e carbon dioxide atmosphere used only as a heat c a r r i e r or d i l u e n t t o lower the r e s u l t i n g temperature. YiTiilst t h i s appear-s t o make such engines an engineering p o s s i b i l i t y , the question of themiodynamic efficiency i s complicated and cannot be considered here.
The ramjet "was s e l e c t e d f o r the only i n v e s t i g a t i o n i n t o t h i s form of power u n i t f o.r the same reasons as those i n t h e c h a n i c a l l y r e a c t i v e atmos-pheres above. I t i s a point of i n c i d e n t a l i n t e r e s t t h a t stich a system i s very s i m i l a r t o an a i r breathing ducted rocket operating without secondary atmospheric combustion.
To simplify t h i s f i r s t i n v e s t i g a t i o n i n t o engines operating i n other atmospheres, and t o make the r e s u l t s d i r e c t l y comparable, i t was decided t o
. "7 .
-standardise t h e engine operating conditions i n a d d i t i o n t o standardising atmospheric temperatures and p r e s s u r e . Since ramjet engines are s u s c e p t i b l e t o Mach N'umber r a t h e r than v e l o c i t y , Mach 3 was taken as a reasonable
operating condition. An i n t e r e s t i n g point a r i s e s out of t h i s . The speed giving Mach 3 Ifi "the various hypothetical atmospheres v a r i e s considerably, being about 8000 f t . / s e c . i n the case of a hydrogen atmosphere, 3200, 3000 and 2000 f t . / s e c , r e s p e c t i v e l y in methane, ammonia and carbon dioxide. Nevertheless, such v/idely divergent speeds produce the same increases i n p r e s s u r e and t o n p e r a t u r e vyithin t h e various engine i n t a k e s .
I n t e r n a l l y , due allowance vra.Spmade for l o s s e s . This lovrered i n t e r n a l pressure frcm a p o s s i b l e 550 l b . / i n dovm t o l e s s than 375 I b . / i n ,
F i n a l l y , as a c o n t r o l on the amount of combustion permitted, 2200 K was taken as the combustion temperature i n a l l cases. This i s qixite a r e a l ramjet l i m i t a t i o n . The decision v/as influenced to a s l i g h t extent by t h e fac^t t h a t i t sitnplified the sums involved !
2.4. Performance in other Atmospheres
Calculations were f i r s t performed t o determine the amount of various tanked oxidant t h a t would be r e q u i r e d t o produce a temperature cf 2200 K a t constant press^ure. As a comparison, ona pound of a i r r e q u i r e s about .06 l b . of f u e l . Results i n d i c a t e tha't one pound of hydrogen vri.ll require something l i k e tiivo and a q u a r t e r pounds of l i q u i d oxygen • ! This a^we-i n s p j j a^we-i n g f a^we-i g u r e a^we-i s beaten by a consa^we-iderable marga^we-in a^we-i f e a^we-i t h e r n a^we-i t r a^we-i c aca^we-id (2fg- l b / s e c . ) or dinitrogen t e t r a oxide (3-2 l b / s e c . ) a r e used. Ozone r e q u i r e s l e s s than 2 l b . per pound of hydrogen, but i n a c t u a l fact i s a most itnprobable propellant because of other reasons. Recpairanent of liquid-oxygen per pound
of the other two f u e l atmospheres shov/s inipravement over the hydrogen case, but i s s t i l l t e n times greater than tlie Earth comparison figure ! I n t h e carbon dioxide case, using licjuid oxygen and kerosine as a heat source, a f i g u r e of a l i t t l e l e s s than 0.4 l b . of t h e tvro propellants per pound of atmosphere i s obtained.
Nevertheless, these figures give no indication of t h r u s t producing e f f i c i e n c y vrfiich must be compared an a s p e c i f i c t h r u s t b a s i s . Present day rockets cjperating i n an atmosphere a t 14.7 l b / i n prod'uce betvreen 220 and 250 pounds of t h r u s t for a t o t a l propellant flew of one pound a second. The s p e c i f i c t h r u s t s of ramjets operating in f u e l atmospheres are a l l b e t t e r than t h i s . Using l i q u i d oxygen, a specific t h r u s t of 3^0 Lb per l b / s e c . i s obtained in a hydrogen atmosphere vAiilst methane and ammonia atmospheres produce f i g u r e s of 430 and h80 L b . s e c / l b . r e s p e c t i v e l y . These r e s u l t s i n d i c a t e t h a t a strong case can be made for the use of a ramjet i n t h e fuel atmospheres considered, but i t i s worth underlining the point t h a t these comparisons have been c a r r i e d out on a s p e c i f i c t h r u s t b a s i s f o r jxist one condition i n each atmosphere and much more vrork remains t o be done before r e a l i s t i c ansv/ers can be obtained.
I n t h e carbon dioxide case the r e s u l t s are much l e s s promising. Using t h e b i p r c p e l l a n t s l i q u i d oxygen and kerosine, a s p e c i f i c thrijst f i g u r e cf
8
-260 Lb, s e c / l b i s obtained v/hich i s very close t o p o s s i b l e rocket engine performance vri.th t h e same p r o p e l l a n t s . Nevertheless, fvirther vrork must be
c a r r i e d out before the ducted rocket form of ramjet for t h i s a p p l i c a t i o n i s dismissed as of no i n t e r e s t .
One or two points t h a t a r i s e fran the figures so f a r obtained are
worthy of seme connent. I n the f i r s t p l a c e , in a l l the a^tmospheres considered, the mass of propellant added t o t h e atmosphere i s of q u i t e considerable
proportion. This added mass c o n t r i b u t e s l a r g e l y t o t h e f i n a l t h r u s t
obtained. P r e c i s e engine geometries liave not been worked out, but i t •would appear t h a t intake areas w i l l be very small compared with those a t e x i t , t h i s being p a r t i c u l a r l y exaggerated i n t h e hydrogen c a s e . F i n a l l y i t must be pointed out t h a t a ccmplcjte comparison be-tween ramjet and rocket can
only be c a r r i e d out by taking engine s p e c i f i c masses i n t o account. No vrork i n t h i s d i r e c t i o n has yet been undertaken. Method of c a l c u l a t i o n and d e t a i l s of assunrptions made are contained i n Appendix I I .
2 . 5 . ConcJ.usionq
Whilst i t has been shown t h a t t h e r e i s good reason t o continue with t h i s i n v e s t i g a t i o n , much care should be taken i n using e x i s t i n g r e s u l t s as the apparent improvements i n perforinance of atmosphere b r e a t h i n g engines over the rocket are a t present based on a very l i m i t e d i n v e s t i g a t i o n .
Nevertheless, a t the conditions l a i d down, i n hj'pothetical fuel atmospheres, very considerable improvements on a tanlced l i q u i d consimption b a s i s have been demonstrated. The investigatican of the carbon dioxide atmosphere i s
at present inconclusive, Acknowle dgement
The author of P a r t I I v/iahes t o place on record t h e c o n t r i b u t i o n made by S.¥. Greenvrood, v?ho s t a r t e d a formal stutïy on t h e use of a ramjet in a hydrogen atmosphere before t h e author, I n p a r t i c u l a r , h i s suggestions were accepted vriith r e ^ r d t o the atmospheric conditions of pressure and temperature and a l s o the external ramjet operating conditions and intake losses used
9 -REFERENCES 1 . M o u l t , E . S . 2. Gardner, G.Vf.H. 3 , C l e a v e r , A.V. 4 , T s i e n , H . S , 5 . P r o e l l , Yf, 6, S a n d o r f f , P . E . 7. A r t i c l e 8, R e i c h e r t , H, 9. P e n n e r , S . S . 1 0 . R o s s i n i , F . D , and o t h e r s Pc3v/er P l a n t s f o r S u p e r s o n i c P l i g l i t , The de H a v i l l a n d G a z e t t e , August, 1955. Guided M i s s i l e s . The C h a r t e r e d Mechanical E n g i n e e r , J a n , 1955« S u p e r s o n i c F l i g h t - P r o p u l s i o n . J o u r n a l of t h e Royal A e r o n a u t i c a l S o c i e t y , November, 1956.
A Method f o r Comparing t h e Pe.rformance of Power P l a n t s f o r V e r t i c a l P l i g h t ,
J o u r n a l of t h e American Rocket S o c i e t y , J u l y - A u g u s t , 1952.
The Design of a P r a c t i c a l Space S h i p .
The J o u r n a l of Space P l i g h t , September, 1949 ( P a r t l ) , October. 1949 ( P a r t 2 ) , November, 1949 ( P a r t 3 ) .
Some New Thoughts on Space P l i g h t .
Paper p r e s e n t e d a t t h e American Rocket S o c i e t y ' s Semi-Ann'ual Meeting, B o s t o n , June 1955.
J e t P r o p u l s i o n News, J o u r n a l of t h e American Rocket S o c i e t y , May 1957. E n t h a l p y C h a r t s . A . R . C . T e c h . R e p o r t , R. & M. No.3015. H . M . S t a t i o n e r y O f f i c e , 1957. C h o n i s t r y Problems i n J e t P r o p u l s i o n . Pergamon P r e s s , 1957. S e l e c t e d V a l u e s of Chemical Thertaodynamic P r o p e r t i e s , N a t i o n a l Bureau of S t a n d a r d s C i r c u l a r 500, 1952.
TECHNISCHE HOGESCHOOL
VLIEGTUIGBOUWKUNDE _ -] O - Kanaalstraat 10 - DELFT
APPENDIX I
Sunmary of a s s u m p t i o n s made, and o u t l i n e of methods of c a l c u l a t i o n u s e d .
Note: Assumptions a r e b a s e d on " t y p i c a l " d a t a . S i n i p l i f i e d methods of a n a l j ^ i s a r e employed, e . g . v e h i c l e a i r d r a g i s n e g l e c t e d . 1 , B r e a k - e v e n t i m e s ( P i g . 3) P l i g h t Mach No. = 3 Ramjet Rocket T h r u s t ( S.L. 28 ifi Weight ( 60,000 t t . 2^.5 40 S p e c i f i c p r o p e l l a n t ( S.L. 2 . 6 ^6 consuniption l b / h r , L b . ( 60,000 f t , 2 . 2 1 3 . 5 e . g . f o r same t h r u s t , and f o r b r e a k - e v e n t i m e a t S . L . , engine weigJit +
p r o p e l l a n t v/eight i s t h e sane i n each c a s e .
T h r u s t 2 . 6 _,, J. J. • / N — + -rr--^ X T h r u s t X time ( s e e s ; 28 •*• 3600 i r u s t 16 40 "•• 3600~ T h r u s t l 6 mu J. J.- /" \ + ^dnn ^ T h r u s t X t i m e ( s e e s ) Time = 2.9 s e e s . 2» C o n d i t i o n s durin.g t r a j e c t o r y ( P i g s . 4 , 5, ^ ^ ^ 6) N.A.C.A, S t a n d a r d Atmosphere C o n d i t i o n s .
As an example, some d e t a i l s a r e g i v e n below f o r c a s e s ( c ) and ( d ) , The s p e c i f i c h e a t r a t i o f o r t h e a i r i n t h e i n t a k e s y s t o n has b e e n t a k e n a s y = 1,4 f o r t h e s e p r e l i m i n a r y i n v e s t i g a t i o n s ,
The i n t a k e t o t a l p r e s s u r e r e c o v e r y i s assumed t o v a r y l i n e a r l y from 1.0 .at Mach 1 t o 0 . 6 a t Mach 5 ,
- 11 "
Case (c) :
15 g boost to 2000 f t , / s e o o followed by 3g constant acceleratic^n v e r t i c a l t r a j e c t o r y , a i - n 4 - - j - ;3 2 0 0 0 x 2 0 0 0 , ^ i A .P4. S e p a r a t i o n a l t i t u d e = ———--=——^^-^ = 4,140 f t . •\rRLOCITï f t . / s e c . 2000 3000 4000 5000 6000 Case (d) ; ALTITUDE f t . 4 , 1 4 0 3 0 , 0 4 0 66,340 113,140 170,140 MA.CH NO. 1 . 8 2 3.05 4 . 1 2 ^ 9 5 4 . 8 7 RIM TEMP. 190 380 675 1100 1725 RAM PRESajRE S.L, Atmospheres lu5k 9 . 2 0 6.57 1.74 0 . 2 0 4 Approx, I 5 g v e r t i c a l b o o s t t o Mach 2 a t 4 , 1 4 0 f t . f o l l o w e d by c o n s t a n t t h r u s t t r a j e c t o r y .
The f o l l o w i n g v a r i a t i o n of t h r u s t c o e f f i c i e n t w i t h Mach No. i s assumed : -MA.CH NO. : 2 3 4 5 THHJST COEPPICXEMT : 1 1.5 1.2 1 The a l t i t u d e n e c e s s a r y f o r c o n s t a n t t h r u s t i s t h e n gi-ven b y MACH NO. : 2 3 4 5 AKEITUDE f t . : 4,140 3 3 , 3 0 0 40,700 4 6 , 0 0 0 3 . Comparison of v e h i c l e Vireigh"ts ( F i g . 7)
As a n example, c o n s i d e r c a s e ( c ) vri.th Mach No. r a n g e 2 - 3 . 5 . ( P i g . 4 i n d i c a t e s t h r i j s t r o u g h l y c o n s t a n t over t h i s r e g i o n - p e r m i t t i n g
12 Thrust Engine weight Ramjet Rocket lé 40 Structure weight Prcpellant weight "'^ ^'^
Specific propellant consumption Ib./hr.Lb. Mass ratio For payload of 10,000 Lb. Lb. Ram,iefc 56,200 3,510 605 6i 2.5 1.04!f Rocket 53,000 1,320 3,190 319 14 1.274 Thrust Lb, Engine weight Lb. Propellant v/eight Lb, Structvire •'-'eight Lb, Total weight Lb, 4,176 4,829 (excluding payload)
Note: Structure weight has been taken as f r a c t i o n of propellant weight. S t r u c t u r e weight influenced by engine weight i s assumed t o be incorporated i n engine Thrust/weight assurnption.
4. Comparison of s a t e l l i t e vehicle weights (Pig, 8) 3 stage rocket •
V e r t i c a l t r a j e c t o r y assumed t o simplify mass r a t i o c a l c u l a t i o n s ,
Specific propellant consumption 14 I b . / h r . L b . and t h r u s t / e n g i n e weigJit = 40 throughout. 3g mean a c c e l e r a t i o n . Each stage operates over a v e l o c i t y i n t e r v a l of I'f- m i l e s / s e c ,
13 -Stage Payload Lb, Engine weight Lb, Propellant weight Lb. S t r u c t u r e weight Lb. 3 500 237 3,355 336 2 4,428 2,090 29,750 2,975 1 39,243 18,500 263,500 26,350 Total w e i ^ t Lb. 4,428 39,243 347,593 (excluding payload)
2 stage rocket with ramjet and booster rocket stages :
V e r t i c a l boost by s o l i d propellant booster rocket at 15g a c c e l e r a t i o n to separation a t 2000 f t . / s e c ,
Constant t h r u s t ramjet s t a g e with 3g a c c e l e r a t i o n to s e p a r a t i o n a t 1 m i l e / s e c , (Trajectory no longer v e r t i c a l , but assumed v e r t i c a l for mass r a t i o c a l c u l a t i o n ) .
2 stage rocket vd.th a v e l o c i t y increment of 2 m i l e s / s e c . for each s t a g e . ( t r a j e c t o r y assumed v e r t i c a l for mass r a t i o c a l c u l a t i o n s ) .
2 stage rocket - assumptions as for 3 stage rocket.
Ramjet s t a g e . Specific propellant consumption = 2,5 l b . / h r , L b . and t h r u s t / e n g i n e weight = 16,
Booster r o o k r t . Specific propellant consumption 17.2 l b . / h r , L b , Total s t r u c t u r e weight/propellant weight = -r-r
2 Stage Stage 2 500 453 7,700 770 Rocket Stage 1 9,423 8,525 145,000 14,500 Ramjet 177,448 64,650 24,400 2,2,40 Booster Rocket 268,938 -143,500 14,350 Payload Lb. Engine vreight Lb. Propellant W e i ^ t Lb, Structure Weight Lb. Total Weight Lb. 9,423 177,448 268,938 426,788 (including payload)
14
-APPENDIX II
Methods and assumptions Symbols, units and values
a = acoustic velocity (— ) ft./sec, m
f = thrust pdls. P = thrust Lbf,
A = mass flow rate lb,/sec.
ill = molecular mass V M = Mach number —
a T = temperature IC, V = v e l o c i t y f t , / s e c ,
Thermochemistry and gas d.^\namics H = enthalpy chu/lb.mol. H = enthalpy c b v ' l b . h = enthalpy f t . p d l , / l b . m o l , h = enthalpy f t . p d l , / l b , s i m i l a r l y I = r e a c t i o n enthalpy chu/lb.mol. e t c . H^ = enthalpy of formation c b y i b . m o l . e t c . R = u n i v e r s a l gas constant chu/lb.mol. A. e t c ,
= 1.98
C = heat capacity ( p r e s s u r e constant) chu/lb.mol, K. e t c .
y = V ( C p - R )
y-1
Subscripts
O = free stream 3 = intal® exit 5 = nozzle entry e = nozzle exit a = atmosphere p = tanked propellant
To convert chu to ft.pdl. multiply by 4.5 x 10 T.o convert pdl. to Lbf. divide by 32,147.
16
-1 . Intake and Pressure Lo.'r'.ses M 3.0
V
^•^•
•^3/?. o - - " - t 37 0.7^t3^^
V o
25.9 25These values have been used f o r a l l atmospheres and are based on y = 1,45. 2. Engine Specific Performance
Assuming complete expansion i n t h e nozzle t o P = P
p d l . ( l )
For isentrcroic f l a v i n t h e nozzle
° t : T c + 2 V ? = c T + i v ^ f t . p d l , / l b . p5 5 5 p e e ' ^ e ^ ' f = ( A + ] ! i ) v - m v ^ a p^ e a o
Rev/riting i n terms of exit v e l o c i t y
= [
2(o J- T_ - c T ) + Vcp5 5 pe e' ^ 5 - ft,/see.
as
Prom this v^. can reasonably be ignored as small. Taking a value c
approximately equal to c ^ and c , i. e. assuming frozen chemical equilibrium
2 c T^ P 5
• / P \ i ^ i
-1 -
-S^
^
V5)
ft,/sec.From equation (l) the specific thrust in pdl. per lb/sec. of tanked propellant is then obtained
•
i4'
N.B. c T^ = he and P 5 5 = V A A a ^ ^ . 1 ^ ( 2 h 77)2 -A / ^ -A V P V p d l , s e c , / l b , (2) and m 32,15 A^ L b , s e c , / l b .- 17 "
3. TherTnophemis t r y
X i s defined as t h e number of mols. of oxidant recjuired t o r a i s e the t o t a l products of combustion t o 2200"^.
Chemical composition balance.
1 l b . mol. atmosphere + x l b . mol. oxidant •• Products of combustion and excess atmosphere as functions of x.
I f the water gas reaxition, and d i s s o c i a t i o n are ignored, the functions of X for each simple component can be determined. This process incjurs an „rror of unknown magnitude but i s not considered too u n r e a l i s t i c a t the comparatively lev/ tonperature and high p r e s s u r e occurring i n the chamber.
Enthalpy balance
I « +. X I * 298 _ y j ^ 2200
atmosphere oxidant ~ products where I»^ ^98 ^ ^
and I « 2200 ^ j ^ ^ ^^ 22CX)
Since a l l I values are a v a i l a b l e from t h e references, x can now be determined and t h e r e f o r e •—• ,
m P
Then H^ = g "'^products ~ 2 K^ chu/lb.mol, atmosphere.
H = ^ ' ^
5 -^ chfi/lb, exhaust, m ,
atmos
h^ = 4.5 3c10^H f t , p d l . / l b , exhaust.
The only value on t h e r , h , s. of equation 2 nov/ r e q u i r i n g consideration i s TJ, Since t h e expansion r a t i o i s loncwn, only a "value of y must be assigned. This can be obtained frcm C t a b l e s since proportions of various components are known, and t h e i r temperature.
C I N T R 6 BODY
INCLINED SHOCK
C O M B U S T I O N
FIG. I DIAGRAMATIC LAYOUT OF TYPICAL SUPERSONIC RAMJET.
I N T E R N A L P R E S S U R E T O O LOW TEMPERATURE R A M T O O HIGH OPERATING T I M E ^v. MINS
FIG. 2 OPERATING REGION OF RAMJET.
FLtGHT MACH N«. • 3
ALTTTUDE
FIG. 3 BREAK-EVEN TIMES AT WHICH ENGINE + PROPELLANT WEIGHT IS THE SAME FOR RAMJET AND ROCKET VEHICLES.
3 G CONSTANT ACCELERATrON VERTICAL TRAJECTORY 5G
I5G 6 0 0 S T T O 2 O O 0 FT/SEC FOLLOWED BY 3 G CONSTANT ACCELERATION VERTICAL TRAJECTORY
FIG. 4 VARIATION OF THRUST WITH MACH No FOR THE DIFFERENT TRAJECTORIES CONSIDERED.
H S BO Z X
^ -l
. 0 ol j RAM PRESSURE 1 / /V
Y \ \ \ \1
/ /L
ji^k
RAM TEMP. > / / /y
A
\ //
1 / / / / / / i //
/ 1
1
VELOCITY TN. MACH l< It /f // // tVI
i
^_ / ' "// / r/I
/ // '
/ / / 2 p O O l . O O O 10 apoo ipoo dpoo VELOCITY F T / S E C MACH No. ^ O O O FIG.S 15 G RAM T E M P , ^ ; RAM PRESSURE-ATMOSPHERESCONSTANT ACCELERATION VERTKIAL TRAJECTORY TO SEPARATION AT 2 . 0 0 0 FT/SEC FOLLOWED BY 3 G CONSTANT ACCELERATION VERTICAL TRAJECTORY.
4 , 0 0 0 SpOO 6 , 0 0 0 VELOCITY FT/SEC 4 5 6 MACH No 2 , 0 0 0 RAM TEMP C RAM PRESSURE-ATMOSPHERES
FIG. 6 CONDITIONS EXPERIENCED DURING A CONSTANT THRUST TRAJECTORY.
VERTICAL TRAJECTORIES
CONSTANT THRUST TRAJECTORY
<J
a
FIG. 7 RELATIVE WEIGHTS FOR DIFFERENT TRAJECTORIES FOR CONSTANT PAYLOAD.
WEIGHT THOUSANDS OF LBS. lOO . 3 STAGE ROCKET STRUCTURE ETC. PROPELLANT FUEL
•V'--C-r 1 I - . . , •cV, .' -^
•A-•^
[2
•
•
••'c •.-; N' • ^, *• .'' :• ' i ' 1 ' ! ' ' S'c ii 2 STAGE ROCKET ' RAMJET 1 BOOSTER ROCKET 'COMPARISON OF VEHICLES TO PLACE 5 0 0 LB PAYLOAD IN CIRCULAR ORBIT AROUND EARTH.
FIG. 8
SANDORFF , CVELOCITY INCLUDES EARTHS ROTATION)
I 2 3
VELOCITY MILES/SEC.
WEIGHT THOUSANDS OF LBS. 4 0 0 |
r
3 0 O 2 0 0 ( O O 3 STAGE ROCKET1
STRUCTURE E T C . PROPELLANT FUEL V.o\\ \ ^'^U-^i
1 • •> > 1 ; A 1 • %^'. 1 J • 1 •^.•^'" 1 2 STAGE ROCKET AIR PLANE STEPCOMPARISON OF VEHICLES TO PLACE 5 0 0 L B . PAYLOAD IN CIRCULAR ORBIT AROUND EARTH CSANDORFF)
FIG. lO Body Mercury Venus Earth Moon Mars Phobos Deinio.s Jupitor lo Europa Ganymede Callisto 7 others Saturn Titan Rhea — 7 others Uranus -r5 moons Neptune Trition + 1 other Pluto Distance from Sun (lO-kms.) .••,8 io8 1 5 0 2 2 B 779 '.430 2.870 4.500 5.900 Radius —(kms.) 2,400 6,100 6 . 3 7 0 ' . 7 4 0 3 , 4 0 0 10 7 0 , 0 0 0 1,700 1,500 2 , 6 0 0 2 , 5 0 0 6 0 , 0 0 0 2 , 8 0 0 9 0 0 2 5 , 0 0 0 3 6 , 0 0 0 2 , 5 0 0 3 , 0 0 0 0.14 0.91 I 0.07 0 . 2 8 120 0.07 0.06 0.17 0 . 1 5 84 0.2 '5 «7 0 . 1 5 0 . 2 0.26 0.90 1 0.16 0.38 2.65 0.1 0.1 0.2 0.2 1.14 0.2 I 0 . 2 ? Day 88d J • 24h 27d 25h loh ? ? •? ? I oh i i h i6h ?
