N71 - 31069 NASA TECHNICAL NASA TM X-67940 MEMORANDUM a sh \e xI cs a v)a 7

PERFORMANCE POTENTIAL OF GAS -CORE AND FUSION ROCKETS: A MISS ION APPLICATIONS SURVEY by Laurence H. Fishbach, and Edward A. Willis, Jr. Lewis Research Center C1eveland, Ohio

TECHNICAL PAPER proposed for presentation at the Second Symposium on Uranium Plasmas: Research and Applications sponsored by the American Institute of Aeronautics and Astronautics Atlanta, Georgia, November 15- 17, 1971 PERFORMANCE POTENTIAL OF GAS-CORE AND FUSION ROCKETS: A MISSION APPLICATIONS SURVEY Laurence 11. Fishbach, Aerospace Engineer Edward A. Willis, Jr., Head, rlight Systems Section NASA-Lewis Research Center . Cleveland, Ohio

Abstract rockets (REGEN.GCNR), 1000 to 3000 sec Isp, F/M of 14 to 25 N/kg. This paper reports an evaluation ok the performance potential of five nuclear rocket 3. Light bulb gds-core nuclear rockets engines for four mission classes. These (LBGCNR), 1700 to 2650 sec Isp, F/M of 10 to 20 engines are: the regeneratively cooled gds- N/kg . core nuclear rocket; the light bulb gas-core nuclear rocket; the space-radiator cooled gas- 4. Space radiator cooled gas-core nuclear h core nuclear rocket; the fusion rocket; and an rockets (SRGCNR), 2600 to 6500 sec Isp, F/M of advanced solid-core nuclear rocket which is in- 1 to 3 NAg. cluded for comparison. The missions considered are: Earth-to-orbit launch; near-Earth space 5. Fusion rockets [FUSION) up to 200,000 missions; close-interplanetary missions; and sec Isp, power to engine mass ratio (P/M, or distant interplanetary missions. For each of l/') OF 1 kW/kg. these missions, the capabilities of each rocket engine type are compared in terms of payload The purpose of this paper is to determine ratio for the Earth launch mission or by the which engine offers the best performance poten- initial vehicle mass in Earth orbit for space tial for each of four mission classes: missions (a measure of initial cost). Other factors which might determine the engine choice 1. Earth-to-orbit launch are discussed. It is shown that a 60 day manned round trip to Mars is conceivable. 2. Near-Earth space missions

I. Introduction 3. Close interplanetary missions

The relatively low specific impulse Isp of 4. Distant interplanetary missions. chemical rockets severely restricts mission capability. For missions more difficult than For Earth-to-orbit launch vehicles, payload lunar exploration and one-way planetary probes, ratio- is inversely related to total initial high propulsive velocity requirements or AV's mass (vehicle initial mass times the number of result in extremely large initial vehicle launches) required to put up a given payload. masses at Earth. These high masses in turn Initial mass is presumably a measure of vehicle imply the need for large and presumably expen- initial cost. (Costs per unit of payload, how- sive vehicles and launch facilities, and for ever, depend significantly upon whether the very complex operations such as orbital assem- vehicle in question is reusable or not.) bly, multi-staging, etc. When a more ambitious program of space exploration is considered, For the other three missions, performance is there is an evident need for higher specific measured in terms of initial mass in Earth impulses together with moderate engine weights. orbit (IMEO) required to perform the given mis- sion with the thrust level optimized. In addi- Solid-core nuclear rockets (SCNR) would tion, for reusable vehicles, propellant load- develop relatively high thrusts at about twice ings are presented since propellant and payload file specific impulse of the best chemical would have to be replenished for missions sub- rockets (800 + sec as opposed to 400 + sec). sequent to the first. Even this increase in Isp may be inadequate for very energetic, very high payload missions, 11. Symbols partly because the SCNR is considerably heavier in relation to its thrust Yhan is the chemical F thrust, N rocket. At the opposite end of the advanced- engine spectrum, elcctric rockets can develop g standard valy of gravity, extremely high specific impulse. On the other 9.80665 m/sec hand, their very low thrust implies long pro- pulsion times and in inany cases, undesirably H/U hydrogen-to uranium-flow rate ratio long mission times. ISP Specifir impulse, sec In this study we hJve investigated the per- formance of several nucledr rockets for a Me mdss of enqine, kg variety of potentially intcresting missions. Included in the study drc: 'is interstdge structure mass, kg

1. An advanced solid-corp nucleclr rockrt Mjcttlson mass jettis~ined, such as Mars (SCNR), 930 sec Isp, thrust til c,nqint* m,tss lander, kg r,itin (r/M) ctr lOfl Nwtrms/kq. propellant mass, kg 2. Regpnrrdt ivr,ly rvmled q.l\-r.r,rc. nLlrlrclr %

1 payload mass, kg neglcctcd, it is diificult to see how this one Mpay characteristic OF the SCNR could be reconciled Mpstr propell,int-structurc mass, kg with the mplasma ball which is at extremely Subscripts: high temperatures and heats hydrogen primarily by radiation. The plasma is approximately i ith maneuver separated from the hydrogen by hydrodynamic means. A small amount of "seed" material, imax last maneuver e.g., tungsten powder, is added to improve the hydrogen's-- radiative absorption properties. 111. Characteristics and Masses of Prospective Nuclear Engines Basic research has resulted in apparent progress toward a demonstration of feasibility. Before taking up a detailed discussion of Laboratory tests applicable to this concept missions, trajectories, mass fractions, etc., (reference 2) indicate that the necessary types it is appropriate to briefly review the perfor- of flow pattern, uranium plasma confinement and mance and operating characteristics used for separation, etc. can be achieved under condi- the advanced engine concepts considered in tions suggestive of a GCNR reactor. The favor- this study. able engine weight and impulse characteristics predicted by that study (see reference 3) yield Solid-Core Nuclear Rocket [SCNR) good performance for interplanetary missions (reference 4) and also appear promising for This is the only one of the engines to be Earth launch missions. An advantage in this discussed that could be brought to operational application is the fact that the fission status within a decade. It operates on a fa- fragment inventory is not retained in the reac- miliar principle: the heating of hydrogen to tor after shutdown. Thus, it would probably produce thrust. The heating is done by a not present a major radiation hazard after reactor constructed of solid fue1.element.s and shutdown. Unfortunately, it does discharge a structure. To keep these parts from melting, small amount of partly-fissioned uranium (e.g., the reactor temperature, and hence the hydrogm 1 percent by mass of the hydrogen flow rate) at temperature, is limited to values less than all times during engine operation, and this is 3300 K. Correspondingly, the maximum specific evidently an undesirable feature from the impulse attainable by this concept is only a atmospheric-pollution viewpoint if an Earth to little greater than 900 seconds. The NERVA orbit launch vehicle with low-altitude startup program has already demonstrated that a value is envisioned. of at least 800 seconds can be attained in practice. This engine might also be undesirable for near-Earth and lunar missions since it would Unfortunately, the masses estimated for tend.to contaminate the area with radioactive such engines have increased significantly dur- wastes due to the large numbers of such ing the development process. This has resulted, missions that would be contemplated. in pdrt, simply from an increased awareness of the need to shield certain vehicle and system The mass of the regeneratively cooled GCNR elements against nuclear radiation. Neverthe- is based on equations presented in reference 4 less, from edrly estimates as high as L'160 (which in turn updated those OP references 2 (kg mass per Newton thrust), SCNR specific and 3). These equations relate engine mass to masses hdve now increased to values approxi- specific impulse, engine diameter, hydrogen to mating 1/30 or 1/25. Tor the purposes of this uranium flow rate r.itio, critical mass, and study an improvement in technology has been thrust. For the Rl:GI:N.GCNR's studied herein, assumed: A sper*iCic impulse of 930 sec for the the diameter is 3.66 m, the critical mass is SCNR wLis chosen dnd tlie c-ngine mdss was assumed 148 kg, the flow rate ratio is 100 and the tn IIP rcpresentcd by specific impulse is 2500 secs. The engine mass [reference 11) becomes

2 REGEN-GCNR. 'L'lic iihi ]in* diilc3miic.e5 a% tli,l t tile Fus ~IIIIRIII 1ki ts (I'U~~I~II~' light bulb cmploys helium-3 excess heat. It, tliert*lor-c, h.is ,111 thcs bcnc- is injc-ctr d inti^ cl i'edction clldmber. A iew pcr- fits of the RCCCN-CCNR, but does not dischsirgc crnt 01 tlic in jcctiad furl undrrgoes fusion radioactive dchris. Thus, the LBGCNR dppedrs redrtions. The onr'r'gy rttlcwsed heats the un- to be an attractive candidate tor near-Edrth redcted fuel to csxtrcmely high temperatures missions. where it ionizes to lorm phsmd. Magnetic fields hold the p1,ism;l fuel away from the reclc- The mass and specific impulse of the LBGCNR tion chamber walls and divert some of it into ~1 are given in Table I as a €unction of thrust. mixing chamber. Hydrogen propellant is injectcd These values are from computations by United into the mixinf chamber where it is ionized dnd Aircraft Resedrch Laboratories based on a heated. The thermal energy of the propellant is radiator mass of 135 kg/W and a chamber converted into directrd motion by a magnetic pressure of 1000 atmospheres. nozzle to produce thrust. Mixing with a propel- lant is required because the escaplng fusion- TABLE I: LBGCNR DATA reaction products by themselves would have a specific impulse in the range of 200 000 sec, F- n I Me-kg far dbOVe the optimal value for most planetary 133 370 I 14 050 missions. By adjusting the amount of hydrogen 222 370 15 650 1905 added, the specific impulse can be varied as 311 360 17 200 1990 desired throughout the mission. 400 360 19 050 2050 667 100 23 125 2180 The specific powerplant mass,&, has been 1 334 200 34 Lk75 2355 taken to be 1 kg/kW as estimated in reference 7. 1778 YO0 44 000 2425 This reference assumes a radiator specific mass 2 668 YO0 83 000 2530 of about 15 kghW as compared to the 135 kg/ Mw 3 113 100 122 450 2570 used herein for the LBGCNR and SRGCNR. 3 557 900 204 075 2605 385 500 1 2635 IV. Propulsive Velocity Increments

Space Radiator Cooled Gas-Core Nuclear The required propellant loading for each Rockets (SRGCNR) stage may be easily computed from the chemical rocket equation, i.e.

(Mp)i = (Moli (1 - exp(-AVi/Ispg) (3)

The propulsive velocity increment or AV ulti- mately is obtained from trajectory simulations. For Earth-to-orbit launch missions, the launch- trajectory code of reference 8 was used. This comprises realistic Earth and atmosphere model% representative vehicle aerodynamic coefficients, and an accurate numerical-integration package. Calculus of variations steering logic is used above the sensible atmosphere to maximize the payload delivered.

Ideal impulsive AV's for outer-planet mis- sions were taken from reference 9, a standard trajectory data compilation. Similar data for Mars missions were obtained from unpublished Lewis data which, like reference 9, was gener- ated by means of the trajectory program de- scribed in reference 10. Precomputed gravity loss data (c.f. reference 11) was tabulated and used by means of a tabular lookup routine to account €or AV penalties for finite thrust. This permits the optimization of thrust level F- n Me- kg; Isn-sec without re-integrating the actual trajectories. 22 240 36 280 3400 44 470 45 590 4150 For the fusion rocket, an approximate low- 88 950 68 210 4850 thrust trajectory computer code based on refer- 133 420 82 830 5200 ence 12 was used. In order to facilitate com- 177 900 101 440 5500 parison with the higher thrust systems, all 222 370 120 060 5700 planet physical rmstants arid circular/coplanar 266 850 136 680 5800 orbit clemrnts were chosen to match those of 311 320 157 300 5900 refrrence 9. 355 800 175 920 6000 400 270 194 530 6000 V. Vehicle Mass 444 750 213 150 6000 Veliii~lr~mass is cLilculLitcstl by suiiuniiig thc masses required for each maneuver. As previ-

3 ously mentioned, pr~~pc~llantiniiss is uiilc.ulii1rtl vt~lii<-lr~siii-[a iiiclioiitc!d I)y tltca h~ii.izoitL.11hlilri -from rqitatinn (3). Tlic~ initial iitiiss L'oi* tliv ,111 tlti, l'iy,iiiv*.) '1'11~. piiyti~irtl is tlicrl thca di1'- next maneuver is then I'~~i~~iict-Iii*twc~.ii tlie str-ucttiriil mass l'r*;lctiwi ;iiicI Llicb Iiiii-riout tract ion shown ori the figure. ,l'a!r c*x;lnlJJIl', ii 1iypothc.ti.cal engine fJf ltllJ0 See rsp iirid spc-cil'ie mass of 0.02 kg/Nc*wton would yield ii 0.30 burnout mass fraction. Assuming il The final maneuver wquires that the init tal 0.25 dead mass fraction, a comfortable 5 per- mass equal cent payload fraction remains.

Unfortunatcly, even the advanced SCNR (030 sec Isp) would rfquir*e a specific mass nf 0.005 (less than Iialf tlie presently estimated value, Mpay + Me (5) c.i. equation (1))before it could deliver any payload at il1.1. l'u deliver a 5 percent payload ratio, it wuuld have to be essentially mass- (7) less. The outlook for the GCNR type of engine is considerably better. Based on the most optimistic combination of present estimates of Isp and engine specific mass, payload ratios of For Mars and major planet missions, 15 percent may be obtainable. Hence, it is concluded provisionally [pending more refined Mjettison = 150 000 kg (9) structure and engine mass analyses) that'the LBGCNR is an attractive candidate for this For the missions studied herein there are mission. usually two payloads, one left at the destina- tion which is called Mjettison and one returned VII. Near-Earth Missions to Earth consisting of a command module, crew, etc. called %ay. Lunar Ferry

VI. Earth-to-Orbit Launch The Lunar Ferry mission and its associated vehicle are illustrated schematically in fig- Launch costs have been an item of major ure 4. In this mission, the reusable nuclear concern since the earliest days of the space rocket stage,which is initially in a parking program. Cost reduction principles yield only orbit about Earth, follows a minimum energy incremental improvements in cost effectiveness transfer trajectory to deliver various amounts as long as the launch vehicle is discarded payload into a lunar orbit. The vehicle after each use. In order to reduce launch of then returns with a 50 000 kg payload (crew, costs to really attractive levels; e-g., less command module, etc.) on a minimum energy than $400 per kilogram in orbit, a fully reus- transfer to Earth and into the original park- able system, "the space shuttle" concept, is ing orbit. There it would be refueled, pick under intensive study at this time. As pres- up another lunar payload, and depart for lunar ently conceived, it consists of two chemical orbit. The capability to deliver large pay- rocket-propelled airplane-like stages. loads to the moon will be a necessity if and when permanent lunar bases or colonies are Unfortunately, the low of chemical Isp formed. engines combined with the considerable mass of reentry structure result in an uncomfortably The performance of four of the nuclear small payload ratio, e.g., about 1 percent of rocket concepts was measured for this mission. inili.il mas. An alternative approach which The fusion rocket was eliminated because its could be considered for later operational a low thrust levels result in excessive mission date involve the use of advanced nuclear bnuld times. The results are shown in figure 5. The ~ngi11c.s. By tdhing advantage o€ the presumed Itigli specific impulse and other, favorable char- SCNR requires the highest IMEO for this mission artc,ristics of (e.g.) the LBGCNR, it is possi- The CBGCNR and REGEN-GCNR each offer a poten- ble in principle to achieve excellent payload tial 20 percent reduction in IMEO over that of idtins and to do it with a single st=tge vehicle. the SCNR. The reader may recall that these two engines are very similar in performance; the Niib latter concept, illustrated in figure 2, ha5 apparent advantages in terms of operational mdjor difference betwccn them is the method of simplicity and a presumably lower initial cost. separating the hydrogen from the uranium plasma. Their specific impulscs mc about the same7 as A parametric study of this possibility is .ii-e their m.t

'I that of the LIIGCNR ,ind l

The second qear-Earth mission studied is the By increasing travel time, IMEO can be re- "slingshot", which is essentially an advanced duced dramatically. At 80 days, for example, verqion oE the "space tug". This mission and an IMEO of only 1 million kg is required. For its associated vehicle are illustrated schemat- the rest of this discussion on Mars trips it is ically in figure 8. The vehicle is initially assumed that the 80 day fast trip would be of in an Earth parking orbit, then boosts out of prime interest since it has only half the IMEO orbit to a given hyperbolic excess velocity,. of the 60 day trip yet does not assume any Va, and separates, with a 500 000 kg payload appreciable extension of space time over the continuing along the initial path. The nuclear Skylab program. rocket then retrofires, returning to low Earth orbit where an additional impulse places the The emission of radioactive waste of the vehicle with a 50 000 kg payload back into a SRGCNR is not thought to be significant for a circular parking orbit. This mission is analo- mission of this nature since there would not gous to the reusable launch vehicle; the be as many repetitions. The high I,p of the "slingshot" is ready to boost another payload SRGCNR at the same time as having high thrust as soon as it has been refueled. capability (compared to electric or even fusion rockets) make it a clear choice for this mis- The same four engines are studied for this sion. The shortest travel time that either of mission as for the lunar ferry. The results the other GCNR's can do the mission in is about are shown in figure 9. The SCNR again requires 100 days and even then require 6 times as much the highest IMEO. The LBCCNR and REGEN.GCNR IMEO as the SRGCNR. performance curves were indistinguishable and have been plotted together. The SRGCNR agdin Conventional MdrS Missions. Most studies of shows the best performance. At a V, riF 5.5 Mars round trip missions have considered trip kdsec (about the requirement for a Earth de- times of 1 year and longer (e.g., reference 13). parture maneuver for a 300 day Mars round trip) These missions have much lower AV the SRGCNR requires 93 percent as much IFEO os requirements than the fast ones; thus, they can the LBGCNR and 70 percent as much as the SCNR. be more ambitious in trrms of large payloads, The savings of 7 and 30 percent, respectivcly, reusable vehicles, etc., and still result in increase with increasing V, since propellant reasonable IMEO's. loading increases and the higher Isp of the SRGCNR proves more bcncficial. As in the lunar These "Scicnce/Cxploration" missions also ferry mission, the diflerences in terms of pro- stmt in low Earth orbit dnd proceed to a 0.9 pellant loading requirements is considerably eccentricity parking orbit at Mars. The tr,ins- larger with the gd5 cores having a 3 or 4 to 1 fer times md angles are much longer, however, advantage over the solid core rocket at high than those rrir the courirr missions. The energies. vehiclc rcmdins in Marticin orbit Cor 40 days. A pdylodd ti1 150 000 kg is leEt which might VIII. M'inned Intrrplanetarv Missions hove been used to go down to the surfac?, build ,in rivbiting tibservatory, vtc. The vehicle then A typicLilrecoverable manned interpldnetary returns tu Cmth with a p,iylodd of 100 000 kg

5 dnd reenters the initi.11 p.ii*king whit ihiii:; til iiis :iOoii~. 11'1 I,:.: 111' iii,liiiiiiiit. '1111(, :,li;il>c~(11 prupulsivc brdking. tli<.sr*I.III'V~*S is 11ii. slini(' iis tlirisc: l'or 1i%0 iriid tin' i~c~1.1Livi. iit,igriiLtrd(*s dr'v itbout the same A11 five 01- thc nwl(~.irivicl\ct Lylys iii ;I IS,!. '1'11(* I~i~ir~J~~ll;iiiti'c.quirtcmt.nts can be tliis pdpei wercj stuilic.cl lo? Lhis missioi1. As c;ilcuLa tcd by mu1 tiplying the uranium requirc- previously mentioned, c-ven thr 930 sec IslJ mcnts IJY tlic. HA! ratin. This propellant SCNR m.iy not be sufficiciit lcii- Iiigli eii('i'gy, ' reqiii.rcvwii ts is important since thv Scienc:c:/ high payload missions. This was the mse for Kxplorcltirin veliiclcs are reusable. The 80 day this mission. In order to bving tlic IMEO Tor SRGCNR Courier reyu'res about $35~106worth SCNR trips into the 1 to 2 million I,g rdngc, ol' 1~235(3500 kg x $lUit/kg) when 11/11 = 200. it was necessary to reduce the Earth return Although il cirnsideroble sum, this in percc'ritogt' payload to 50 000 kg and reenter atmospher- terms, tloc~s not reprcscnt il major cost incre- ically (not enter into the initial parking ment when coin Jrod (c*.g.) tu il dircsct cost CJ~' orbit). Thus, this vehicle is not recovered. about $220x10!' Lor 1ilunr.h operation alone (lo6 kg IMEO x about $220/kg in orbit). The results of the study are also shjwn in figure 11. Even with an easier missinn profile The optimum tlirust levels for LBGCNR's ant1 thc SCNR still requires the highest IMEO. The SRGCNR's are shown in figure, 13. Thc, LBGCNR LBGCNR and R€GEN.GCNR again yield almost iden- thrust levels are about 5 times thnsr of tlie tical results. A 500 day mission requires SRGCNR. .Since the I, level is abrlut 1/2 tiiilt about 1 million kg LMEO. The trip time cm be of the SRGCNR, the LBECNR reactor ~J~WCT is 2.5 reduced to 1 year for an increase in IMEO to times that of the SRGCNR. It appears that one about 1.6 million kg. SRGCNR reactor could be used to perform a large portion of the missions whereas the The SRGCNR appears to offer significant sav- optimum thrust level changes more rapidly with ings in terms of IMEO or trip time and, in mission time for the LBGCNR. addition, a flattening of the curve. At 500 days the SRGCNR requires only 60 perrent as This point is demonstrated more clearly in much IMEO as the LBGCNR. On the other hand, figure 14 where it can be seen that for the 80 for a 1 million kg IMEO, the SRGCNR can do the day Courier and 400 day Science/Exploration mission in 280 days or a reduction of 45 per- missions, IMEO is fairly insensitive to the cent in mission time. thrust level of the SRGCNR over quite a long range. A 150 000 Newton engine would perform The fusion rocket appears promising at long both missions with essentially the same IMEO trip times, excelling the other four engines. as the optimum thrust engine (180 000/80 days, A rocket of this nature propels fora signifi- 130 000/400 days). Thus it is conceivable cant portion of the trip and, thus, can follow that a standard-design SRGCNR might be usable more optimum flight paths. for a large variety of space missions.

With decreasing trip time, the long propul- Since the IMEO appears to be relatively sion times of the fusion rocket force it to insensitive to the thrust level, and since propessively higher thrust and lower specific there is only a 2 to 1 variation in optimum F impulse. This inrreoses propellant mass and for the LBGCNR (fig. 13), perhaps a fixed-size powt-rpl dn t fractirm resulting in a crossover LBGCNR would also have multi-mission capability. ktween til(- SRGCNR dnd fusion rocket curves at about 1 yedr- mission time. The SRGCNF;discussed so far operated at a chamber pressure of 1000 atm. If the chamber H.rsed rin the results illustrated herein, it pressure is increased to 2000 atm, the would dppedr that SRGCNR would be the best can- achievable specific impulse increases by dida te for these Science/Explora tion missiuns approximately 1000 sec at a thrust level of to Mas. Trip timcBs are relatively short and 50 000 Newtons to 500 sec at 500 000 Newtons. potential mass wvings of the fusion rochet do If the chamber pressure is decreased to 500 atm not appear to be decisively large. The story the Isp decreases by 1000 sec at 50 000 Newtons could be changed, however, if the fusion rocket to 600 sec at 500 000 Newtons. The engine mass had a specific powerplant mass 01 0.5 l\q/kW does not change significantly at 50 000 Newtons instead of 1.0. This would shift the Lusion and is 2 10 000 kg at 500 000 Newtons. The rocket curve down and to the lc,ft, resulting in effect of achievable specific impulse is shown higher mass savings and reduced mission times. in figure 15. High Isp is for a chambep pres- It should be pointed out that the equivalent sure of 2000 atm, Nominal for 1000 atm. (table specific powerplant mass or& of the SRGCNR is 11) and Low for 500 atm. about .01 to .1 kg/kW, but it suffers somewhat in not being able to achieve the high Isp of The ef€ect of Isp is small for the 400 day the Eusion rocket. Science/Cxploration mission and strong for the much higher energy 80 day Courier mission. For The uranium requirements for the courier and the 500 iitm engine 80 day mission a 40 percent Science/Exploration missions me shown in increase ill IMEO is shown while -the 2000 atm figure 12 for hydrogen to uranium flow rdte enginc decreases thc IMCO by 20 percent. ratios of 100 and 200 for the SRCCNR and for a ur.inium mass to tlirust rdtio 01 .ftft8 hg/N lcn. Md ior Pliinet Missions the SCMt. The LBGCNR fuel requirements are very small since no fuel is lost and is The last class of missions studied are trips determined by burnup rate and reuses. As an to the m.ijor planets .Jupitc~.Saturn, and example, a 400 000 Newton thrust cngiiiv cun- U1,anus.

b A peculiirrity of ina-jor pl,iiic>t round ti*ips, s iiic'i- i I t1tti.s IIIJL i.iiii t riirliriactive wdstcs. as discussrd in ~(-I'vI~(-II~T CI, is that fcvisibl(* trips €or re1;it i.vt*Ly Iiigh thi-list roclicts su(.li A sLgiii.l'Lv:iiit (.l,iss 811' IIOW idst trips to its tllrJSC ilisr:iissc-tl Iic'~.(.iil (i~sclusi.vi, or LIIk' Pl,it~h liiis !Jl'1.11 lii,iilio*l-LII, t1111y at tliscrctc. Liiiic.s iis 1.r~iih HI1 (I.rys, ii1i.innc.d rtJunrl trips intc,rvals (it' 12 tcj I3 inoiitlis. Tlius, ct!~, ciiii I>(* ~1~winiplislic*rl ior, unly 1 million kg these ruclu~ts, inissirins du not exist ;It timias Tl\n:O. TII uddition, should trip times be of inte~*mi,ili.~tc,to tliv dii til poi.tits shown in priinc? i.iiipurtmcc, 611 day duration round trips figui-c~16 wli ich tiiivi' bivri cwnnected with to Mar's appear feclsib1.1. for a 2 million kg str;iiglit lines to idcntify the rocket type tMI:O. ;lnd indicate trends. AlJ oI tlic. gas-core and fusion concepts Consider, first the ,Tupiter missions shown c,xamincd promise appreciably better perfor- in figure 16(a). For these missions there are mance than the solid-core engine. However, ,lgCiin two modes: Courier (no payload to selection of a preferred concept must await planet, 50 OLIO kg biick to Earth and atmospher- continued work to establish the feasibility ic reentry) and Science/Exploration (150 000 of the concepts and better definition of kg to planet, 100 000 kg to Earth, 200 day engine characteristics, followed by considera- stay time. in Q.5 eccentricity parking orbit tion of development and operational costs. with pericipsis

The Couri(,r. mode requires 1.67 years and an 1. Clark, M. R., Sagerman, G. D., and Lahti, IMEO of 350 000 kg for the SRGCNR and 840 000 G. P., "Comparison of Small .Water- kg for the LBGCNR. The fusion rocket which is Graphite Nuclear Rocket Stages with sliown by the long-short-short dashed curve Chemical Upper Stages for Unmanned gives continuous performance but cannot per- Missions," TN D-4827, 1968, NASA, form il 1.67 year trip to Jupiter. Cleveland, Ohio.

The next opportunity for high thrust 2. Ragsdale, R. G., "Some Fuel Loss Rate and rockets occurs at 2.8 years. The IMEO Weight Estimates of an Open-Cycle Gas- incredses over that at 1.67 years as a result Core Nuclear Rocket Engine," Paper 70- of switching Erom Courier to Science/Explora- 690, 1970, AIAA, New York, N.Y. tion type trips. At this trip time, fusion ririd LXCNR IMEO's are at about 1.4 million kg 3. Ragsdale, R. G., "Relationship Between while the SRGCNR requires 750 000 kg or about -Engine Parameters and the Fuel Mass Con- 54 percent as much. tained in an Open-Cycle Gas-Core Reactor," Research on Uranium Plasmas and their Going to longer trip time, the IMEO for the Technological Applications. SP-236, 1971, fusion rockets drops rapidly and for trips NASA, Washington, DC, pp. 13-22. beyond 4.3 years outperforms the other two rockets. As missions become difficult and 4. Fishbach, L. H., "Mission Performance trip times longer, the practically unlimited Potential of Regeneratively Cooled Gas- Isp capability of fusion rockets makes them Core Nuclear Rockets," TM X-2256, 1971, increasingly attractive. NASA, Cleveland, Ohio.

Thus in figure 16(b) it can be seen that, 5. McLafferty, G. H. and Baur, H. E., "Studies fusion rockets always outperform LBGCNR's for of Specific Nuclear Light-Bulb and Open- Saturn missions and offers slight improvement Cycle Vortex-Stabilized Gaseous Nuclear over SRGCNR's beyond 4.8 years. Engine," CR-1030, 1968, NASA, Washington, DC . For missions to Uranus (fig.l6(c)) the fusion rocket outperforms both GCNR's for 6. Ragsdale, R. G. and Willis, E. A., Jr., Science/Exploration missions. Even so, the "Gas-Core Rocket Reactors - A New Look," performance gains of the FUSION rocket are TM X-67823, 1971, NASA, Cleveland, Ohio not decisively large for the missions con- and Paper 71-641, 1971, AIAA, New York, sidered here. By extrapolating the trends NY . shown in figures 16(a), (b), and (c), the fusion rocket appears to be the best candidate, 7. Reinmann, J. J., "Fusion Rocket Concepts, for very liigh AV, very long time missions such TM X-67826, 1971, NASA, Cleveland, Ohio. BS Neptune and Pluto round trips and Solar System escape. These missions were not con- 8. Spurlock, 0. F. and Teren, F., "A Trajec- sidered here. tory Codr for Maximizing the Payload of Multistage Launch Vehicles," TN D-4729, IX. Concluding Remarks 1968, NASA, Clwcland, Ohio.

For the five nuclear rockets studied here- 9. Fishbiicli, L. H., Giventer, L. L., and in the space-radiator-cooled gas-core rocket Willis, E. A., Jr., "Appriixirnotc Tra- appears to always require thr lcast IMI:O fnr ,ic,ctory Data for. Missions tn the Major the missions studied if excessive trip times Planets. TN D-6141, 1971, NASA, are rukd out. Other ecol%ical restructions Clcweland, Ohio. may make the light-bulb gas-core nuclc,ar rocket the choice Tor the near-Earth missions

7 Z C. L., "A Flc~tli~rdrjl Apprrrximiiting Pi*(11 rr 9 11 ; in t I<(*el u i. i,cmcn t s a t I' Lt IW- 1'1 ir,ust '~i~;i,ic~~~tiii~i(~s," 1'N 11-3400, 19 66, NASA, Cltw.L:tncl, Ohin.

T,uiclc~iis,. I<. W., Ilur~lc~y,R. I<., Liscnberg, J. I)., l(ti]~para~I-,J. M., Miller, n. A., Sliiivlin, M. D., :irid Willis, E. A., Jr., "l'l;iniied Elms Landing Flissions by Means oL' lli;$t-'l'lirust I

83x185 KM

SHORT VERTICAL RISE / / / / / / / / / / / /// / / /////ISURFACE

/ rAERODYNAMlC STRUCTURE

PRESSURE SHEL ,-FUEL AND PROPELLANT

EXCHANGER LOOP ,ADVANCED ENGINE

C D-11087-22

Figure 1. - Open Cycle gas-core nuclear rocket engine - schematic diagram. Figure 2. - Earth launch vehicle and trajectory. NUCLEAR ROCKn VACUUM SPECIFIC

Payload placed 50 000-kg Earth return payload 7, in lunar orbit

,/Propellant and tankage

EarthReturned orbit to Nuclear /r LBGCNR PARAM- {i:- rocket engine -4k \ \ ,/ ETER RANGE /e---- r EXPECTED DEAD MASS \ // \\ .3 \ \ \ Lunar .2 escape

L './ 6)Lunar .1 capture / / / 0 0 .01 .02 .03 .04 .05 .06 ,-- ,' NUCLEAR ROCKET ENGINE MASSlTHRUST, KGlNEWTON Figure 3. - Performance of single stage nuclear rocket Figure 4. - Lunar ferry mission. launch vehicle. Final orbit: 83x185 km, 460 mlsec AV reserve. E-5547

500x 103 ----- ADVANCED SCNR --- LBGCNR ----- ADVANCED SCNR SRGCNR 2 (HYDROGEN ONLY) REGEN GCNR vi 400 --- LBG CNR tn / ----- / REGEN GCNR 0 / / 2 SRGCNR 0 E 2 2 300 I

G 0 1 2 3 4 5x lo5 PAYLOAD MASS, KG PAYLOAD MASS, KG Figure 5. - Effect of payload mass on lunar ferry Figure 6. - Propellant requirements for lunar ferry. mission. mission. 50 000-kg EARTH PAYLOAD

130x lo4 \ R€rURN PAYLOAD '\ TO V, PROPELLANT AND TANKAGE RETURNED TO EARTH ORBIT . NUCLEAR !':;I: i!':;I: ROCKET ENGINE

EARTH ESCAPE/

CAPTURE STAGE ONLY---,, /4---- \\ \ DESIGN / /' \ I I / / / 80 /

1000 2000 3000 AM- SPECIFIC IMPULSE, SEC --/- PAYLOADTO V, EARTH ESCAPElCAPTURE / Figure 7. - Effect of specific impulse on lunar ferry STAGE AND PAYLOAD iRETRO mission. Light bulb gas core nuclear engine, STAGE ONLY 500 000 kg payload. Figure 8. - Slingshot mission. 4 16ox 10 I ---- SCNR 140 - --- LBGCNR/REGEN GCNR / SRGCNR / / c3 / Ld 120- / / 0 0- 0 w - .’ 100 I# -4- ---/- --- 80------60 I I I I 1 I I

r PLANET-CAPTURE ‘, HYDROGEN TANKAGE \ CORE VEHICLE RETURNED TO 600 Km PAYLOAD CIRCULAR EARTH PARKING ORBIT \, TO PLANET7

’$1I TANKAGE (TYPICAL)

i m

I “-COMMAND MODULE LURANIUM STORAGE AND SUPPLY SYSTEM

Figure 10. - Nuclear rocket vehicle schematic manned interplanetary missions. 0 S CI ENCElEX PLORATION (40 DAYS STAY1 0 COURIER SRGCNR --- LBGCNR ----- REGEN GCNR --m- FUSION a = 1 KGlKW ----- OPTIMISTIC SCNR (NOT RECOVERED AT EARTH RETURN)

5x lo6 I

MISSION TIME, DAYS Figure 11. - Effect of Mars round trip mission time, for various nuclear rocket engines. E-5547

0 COURIERS 0 S ClENCElEXPLORATION COURIERS SRGCNR 0 SClENCElEXPLORATlON

lo Oo0 r m ooot g -

3s4 z 2 =2 SRGCNR I I ? 0 100 200 300 400 500 600 0 100 200 300 400 500 600 MISSION TIME, DAYS MISSION TIME, DAYS Figure 12. - Total uranium propellant requirements Figure 13. - Optimum thrust levels for various for SCNR and for SRGCNR rocket engines as a nuclear rocket engines. Mars round trips. function of propellant to fuel flow rate ratios. Mars round trips. 1.

1. 4-

1. c3= E- 1. 5 8- 400 DAY S CIEN CElEX PLORATI ON

*6* *6* .450 100 150 BOX103

THRUST, NEWTONS Figure 14. - Effect of thrust level on JMEO for Mars round-trip missions. Space radiator cooled GCNR.

1. 6x106

80 DAY COURIER =c3 E- -8 - HE 400 DAY SClENCElEXPLORATlON

0' 0' LOW , NOMINAL HIGH SPECIFIC IMPULSE LEVEL Figure 15. - Effect of 'attainable specific impulse of radiator-cooled GCNR - thrust levels reoptimized. I irl 0 1 2 3 4 5 6 7 (A) JUPITER.

KW

O21 3 4 5 6 7 8 (B) SATURN. 12106 I

4- f"-'i",", I KW \ '\4 -\LBGCNR

NASA-Lewis-Com'l