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AIAA 92-3778

The Application of Nuclear Power and Propulsion for Space Exploration Missions

Robert M. Zubrin and Tal K. Sulmeisters Martin Marietta Astronautics Group Denver, Colorado, USA

AI ANSA E f AS M WAS E E 28th Joint Propulsion

Conference and Exhibit July 6-8, 1992 / Nashville, TN

For permlsslon to Copy or republish. contact the American Institute of Aeronautics and Astronautics 370 L’Entant Promenade. S.W., Washlngton, D.C. 20024

A I A A- 92-3778

The Application of Nuclear Power and Propulsion for Space Exploration Missions

Robert M. Zubrint and Tal K. Sulmeisterst Martin Marietta Astronautics

PO Box 179 Denver, CO 80201

Abstract

Nuclear propulsion has recently attracted attention as a potential means of enabling the human exploration of the Moon and Mars. However there are numerous other missions, including unmanned exploration of the Earth, its Moon, Mars, the asteroid belt and the outer solar system that can be greatly enhanced or enabled through the use of either nuclear thermal propulsion, nuclear electric propulsion, or both in combination. This paper examines the use of nuclear propulsion in such modes. It is found that a small (10 klbf) nuclear thermal rocket engine offers optimal utility for supporting such missions, enabling an otherwise impossible ballistic trajectory Pluto orbiter mission to be flown with a flight time of about 13 years and a payload of 200 kg, direct flight (no gravity assist) missions to all the outer major planets, and a tripling of the payload delivered to orbit around Uranus and Neptune compared to that possible with cryogenic chemical propulsion. It is also found that such engines can be used to enable manned Mars missions with minimal performance penalty compared to that possible through the employment of large (75 klbf class) NTR engines. Since such large engines are not useful at all on the unmanned exploration missions, and require a much more expensive and extended development program than the small NTR engines, it is recommended that the NTR development effort focus on the rapid development and production of small, near term technology NTR engines for general use.

Missions requiring large amounts of electric power to support their payload functions can be enabled through the employment of nuclear electric power reactors, which in some cases can also assist the mission by making possible the employment of high specific impulse electric propulsion as well. However it is found that the practicality and versatility of using a power reactor to provide advanced propulsion is enormously enhanced if the reactor is configured in such a way to allow it to generate a certain amount of direct thrust as well. The use of such a system allows the creation of a common bus upper stage that can t Member A I M

provide both high power and high impulse (with short orbit transfer times). It is shown that such a system, termed an Integral Power and Propulsion Stage (IPAPS), Is optimal for supporting many Earth, Lunar, planetary and asteroidal observation, exploration, and communication support missions, and it is therefore recommended that the nuclear power reactor ultimately selected by the government for development and production be one that can be configured for such a function.

lntroduct ion

Nuclear thermal rocketry has been identified as an enabling technology for manned Mars missions. Yet the requirements that have been proposed for such engines, in terms of thrust, thrustlweight and specific impulse, as well as the technology requirements for the gigantic NTR spaceships envisioned, in terms of required launch capability, on-orbit assembly, checkout and cryogenic fluid management, are so formidable that they relegate Mars missions based upon such concepts to the distant future. Thus, so conceived, NTR becomes not an enabling technology, but a disabling technology for the Mars mission. Furthermore, by actifiy as an expensive and perpetual toll-gate blocking the initiation of the manned Mars exploration, NTR actually sabotages itself, for so long as its chosen application remains several decades in the future there will never be a strong incentive to put the NTR development program on a schedule that will produce real results. A vicious circle is thus set up in which the advanced NTR and the Mars mega-mission application each act to prevent the other from ever coming to be. Clearly what is needed is to break this circle, and the way to do it is to put the NTR program on a fast track to produce a useful, albeit not optimum engine, that can meet a host of mission needs that already exist, and will continue to exist regardless of the political fortunes of the manned Lunar and Mars program. Once such an engine exists, it in-turn will greatly help the manned Lunar and Mars mission by providing real, flight demonstrated hardware that can be incorporated into practical, cost-effective mission plans in which

Copyright 0 1992 by Martin Marietta C o p Published by the American Instilute of Aeronautics and Astronautics, Inc. with oermission.

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payloads and trajectories are chosen to accommodate the limitations of real (although not optimal) launch and space propulsion systems. In this paper we shail examine the application of such near-term NTR and spacecraff technology to Lunar, Mars and outer solar system exploration missions and show that in addition lo being enabling for a host of exciting unmanned missions currently under active consideration, it is also fully capable of meeting all the requirements of the Space Exploration Initiative.

Similarly, space nuclear electric power reactors have been stillborn because they have been designed for very high power applications which to-date are not actively supported by any significant user community. This is because in the lower power range, reactor mass scaling generally favors the use of photovoltaics over nuclear power sources. However, because of the unacceptably long orbit transfer times associated with low power electric propulsion, only high power applications have the capability of offering mission enhancement through the use of electric propulsion. Thus the development of both electric propulsion and nuclear electric power sources have both been prevented to date. However if a low power reactor is configured so as to allow the utilization of its thermal output to produce direct thrust, then advanced propulsion can be employed by even small nuclear power sources to enable an orbit transfer within an acceptable amount of time, and by taking advantage of this capability, such a system offers superior mission performance than photovoltaics for missions with power requirements as low as 5 kWe. Since many missions are currently being planned with power requirements in the 5 to 10 kWe range, the ability to compete in this regime can insure a user base for a space nuclear power technology development program. Once such technology is created to meet the needs of the existing low power requirements of the user community, high power applications will also be enabled and will be called into being by mission planners in an evolutionary way.

"If we build it, they will come," may have worked well for Kevin Kostner in "Field of Dreams," but it is a strategy that is unlikely to produce satisfactory resuits as a means of introducing revolutionary new technologies into general use in the space program. In this paper we offer an allernative approach, one based upon conditioning nuclear technology to be broadly applicable across the existing mission set, while insuring that it is sufficiently evolvable to play a

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w key role in enabling future missions which are not on the books today because the requisite technology to support them does not exist. We can get "there" from "here," but only by starting from "here."

Smal l NTR Fna lnes fa r E x D l o r a t l o n Missions

Two similar, but different technology baselines were used in the analysis presented in this paper. One, based upon estimates made by the Idaho National Engineering Lab of potential performance of near term small NTR engines assumes a specific impulse of 850 s and a thrust to weight ratio (TMI) of 3 for a 10 klbf NTR engine. The other, based upon recent NASA Lewis studies1 assumes NTR engines with a specific impulse of 870 s, a thrust of 10 klbf and a thrust to weight (T/W) of 2, and 25 klbf engines with a T/W of 3. The rationale for such baselines is that the 2500 to 2600 K propellant temperatures required to produce 850 to 870 s Isp is achievable with materials proven in the NERVA program, so that an extensive program of fuel development will be unnecessary. The reason for the modest thrust levels is to enable the engine to be tested within existing ground L/ facilities at the Idaho National Engineering Lab. The modest TMI of 2 to 3 is chosen so as not to force the pushing of unproven technology or design w concepts, as well as to allow mass for sufficient subsystem (turbo-pump, for example) redundancy to assure adequate engine reliability. While the INEL and NASA LeRC performance estimates differ somewhat, they are not entirely inconsistent, as the thrust of an NTR engine can be increased at the cost of a certain decrease in specific impulse by raising the chamber pressure.

There are many other potential applications for NTR propulsion besides the Space Exploration Initiative (SEI) manned Lunar and Mars missions. Such missions include DoD and civil missions to near Earth space (for example geosynchronous orbit. or GEO), and unmanned science missions to Mars and the outer solar system. Hundreds of such missions are currently planned for the next two decades, and more are certain to be planned for the years following, regardless of the fate of any one particular program. Because there are so many such missions it is very worthwhile to examine how effective an NTR engine is in accomplishing them. Indeed. in the long run, the performance of an engine on these missions may be much more relevant in quantifying its potential benefits to the nation's overall space program than its utility on SEI missions alone,

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W TABLE 1. Payload Delivery Capability for Unmanned Missions with 20 Tonne IMLEO -

won A V v ( k m / s ) s S ) 10klbTMI 3 75klbTMI 5 l O l d ! ~ f f M I 6 75klbTMI 1Q

Trans-Mars 3.9 7.35 tonnes 10.27 tonnes 4.47 tonnes 11.72 9.07 tonnes GEO

- - - - - - - - . .

4.3 6.56 9.63 4.33 11.11 8.46 6.6 3.17 6.46 1.16 8.08 5.43

In Table 1, we show the benefits of various NTR engines for 3 unmanned missions; Trans-Mars Injection (TMI), GEO, and Trans-Jupiter injection (TJI.) The importance of the GEO mission is self evident, as over half of all civil space missions and numerous military missions are flown to this destination. Mars is certain to be a continuing target of unmanned probes, regardless of the fate of SEI, and extensive planning is ongoing worldwide for such missions as Mars Rover, Mars Rover Sample Return, MESUR, and Mars Aeronomy Observer, among others, all of which could benefit by enhanced payload capabilities. Finally TJI i s important since nearly all missions to the outer solar system would use a gravity assist at Jupiter, and so the performance on this mission is generic to all

v exploration probes to Jupiter and beyond.

Trip time is not a driver on such unmanned missions, and so all missions analyzed here are flown on minimum energy trajectories. The initial mass in LEO of all missions is set at 20 tonnes, since this is the launch capability of both the Space Shuttle and the Tian IV.

In Table 1. cry0 stage and NTR tank dry mass

Table 2. Orbiter Missions to Uranus and Neptune

fractions were taken as 0.1. It can be seen that the 75 klb/850 s NTR is inferior to cryogenic chemical propulsion on these missions, and while the 950 s175 klb NTR is superior to chemical propulsion, it is inferior to the 10 klb1850 s NTR in every case, despite its vastly more advanced technology. The little 10 klb1850 s NTR significantly enhances all missions considered, increasing deliverable payload by 50 to 100% over the cryogenic chemical state of the aft. The 950 s l l0 klb NTR is certainly best, but whether the 14 to 25% improvement in capability it offers over the 850 s/lO klb model is worth the more diflicult and risky development effort it entails is questionable.

Summarizing, then, it can be seen that a small, near term technology NTR engine offers substantial benefit for the most common unmanned exploration missions, while a large near term technology engine offers no benefit at all.

Now let us consider the more specific case of orbiter missions sent to Uranus or Neptune. Such missions are currently being planned by NASA Code S for early in the 21st Century. A comparison of options for these missions is given in Table 2.

Phnet Trajectoly

Uranus JGA Uranus JGA Uranus JGA Uranus VEJGA Uranus VEJGA

Neptune JGA NeMune JGA Neptune JGA

v Neptune JGA Neptune JGA Neptune VEJGA Neptune VEJGA

v

Launch Arrive Tme(yrs)

2/11/07 2/11/15 8.0 2/10/07 2/10/16 9.0 2/10/07 2/10/17 10.0

2/12/06 2/12/17 11.0 2/12/06 2/12/18 12.0

2/15/07 2/15/17 10.0 2/15/07 2/15/18 11.0 2/15/07 2/15/19 12.0 2/15/07 9/11/20 13.6 1/10/06 1/09/21 15.0

7/26/02 7/26/19 17.0 7/26/02 7/26/20 18.0

(ka) ~ l ! u . r ! m ~ ~ 169 409 296 139 435 1020 745 365 677 1587 1160 568

756 1143 952 689 1214 1835 1529 1106

84 322 203 39 1 79 628 406 99 339 1100 728 21 1 437 1378 91 9 285 737 1813 1305 602

1648 2255 1927 1472 2067 2829 241 7 1847

3

In the analysis shown in Table 2, cryogenic chemical propulsion was assumed to have a specific impulse of 460 s, and a single stage storable bipropellant system with an Isp of 320 s was used for orbit capture. Tank inert mass fractions of 15% was assumed for all systems, and gravity losses of 5% were assumed for Earth escape and 10% for Uranus or Neptune orbital capture. A reserve AV capability of 500 m/s was added to the storable stage to allow for both mid course corrections and in-system maneuvers at the destination planet. The NASA LeRc baseline was used for NTR performance, with an Isp of 870 s and a T/W of 2,3, and 4 for engines of 10 klbf, 25 klbf, and 50 klbf thrust, respectively. trajectory data was provided by SAIC2.

It can be seen that the use of a 10 klb NTR for these missions offers massive advantages over cryogenic chemical propulsion, with about 2.5 times as much payload being delivered to the destination planet on trajectories that avoid a Venus swingby. If a 25 klb NTR is employed instead, some of this advantage evaporates, but deliverable payload is still 60% greater than that offered by chemical propulsion. A 50 klb NTR engine, however, is too heavy, and its performance i s outclassed by the chemical alternative.

Another mission of significant interest is the placing of an unmanned exploration spacecraft info orbit around Pluto. This mission is much more difficult than a Uranus of Neptune orbiler, because Pluto's weak gravity makes it necessary for nearly all of the hyperbolic excess velocity of the approaching spacecraft to be taken out by propulsive AV. While trajectories can be found which minimize the hyperbolic excess velocity approaching Pluto, such trajectories necessarily involve very long flight times and are unacceptable. A definitive requirement for the Pluto orbiter mission is that the spacecraft arrive

Table 3. Pluto Orbiter Missions Using Small NTR Engines

Flight Time C3 LEO AV LEO Interplanetary Vinf Pluto &asL &E&!& uvnlsl SicMass( kql &D!sL

10 117 8.03 3450 13.72 11 114 7.95 3553 12.10 12 111 7.83 3681 10.79 13 108 7.74 3778 9.69 14 106 7.69 3833 8.77 15 105 7.63 3898 7.98 16 104 7.59 3943 7.30

U

v on station in orbit around Pluto prior to the year 2020, when it is anticipated that Pluto's increasing distance from the Sun will cause its atmosphere to freeze out or "collapse," an event whose observation is of prime interest to the scientific community. Meeting these objectives is very difficult, and a recent study by the Jet Propulsion Lab3, which included only cryogenic chemical rockets for Earth escape propulsion concluded that, with a Titan IV launch vehicle, the maximum mass the spacecraft delivered into orbit around Pluto could have would be 35 kg. This is too small to allow inclusion of a useful set of scientific instruments (in addition to the standard spacecraft functions of communication, attitude control, power, etc.) and so the JPL study concluded that the Pluto Orbiter mission was unfeasible. However if a 10 klb NTR engine is used to drive an expendable stage off of a Titan IV in place of the cryogenic chemical Centaur, the Pluto Orbiter mission can be enabled. The data is presented in Table 3.

The trajectories employed in the analysis shown in Table 3 are based upon JPL data4. and all use Jupiter gravity assists, with the closest approach to Jupiter being greater than 15 Jupiter radii. This is considerably farther from the planet than the Voyager spacecraft traveled, and would insure that the spacecraft is not damaged by Jupiter's strong radiation bels. A December 2004 Titan IV launch is assumed, with the launch window remaining open until 2006. The capture maneuver at Pluto is done by a 2 stage system, with a specific impulse of 375 s assumed for the space storable LOWCH4 option. Gravity losses of 5% were assumed for Earth escape, but none for capture at Pluto, where capture is made into an elliptical orbit 100 m/s into the weak gravity well. A budget of 200 m/s extra AV is also provided by the chemical orbit capture system, to allow for mid- course corrections. The NASA LeRC

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AV SIC AUYa

12.90 11.30 9.98 8.89 7.98 7.20 6.52

Q ~ i e r Mass (kpl Pluto Capture by

NTO/MMHLOX/CH4 6 28

24 €6 57 126

105 197 166 282 242 382 329 491

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I I NTRs Jable 4. D irect rmsslnns to the 0- with Sma . .

Saturn 2.3 13.93 3.62 Uranus 6.6 10.43 3.24

Pluto 19.3 7.57 6.57 Neptune 12.6 8.63 2.10

baseline (870 s Isp) was used for NTR performance. It can be seen that spacecraft masses between 100 and 200 kg can be inserted into orbit around Pluto on missions with a flight time of 13 years. This is sufficient mass for a viable mission, and the flight time is short enough that, even if the launch date were to slip to 2006, the spacecraft would still be on station prior to the 2020 deadline. It may also be observed that the use of space storable LOWCH4 for orbit capture offers about a factor of 2 advantage in delivered payload on this mission compared to Earth storable NTO/MMH, and is also more convenient, since the LOX/CH4 would not be prone to freezing in the outer solar system.

An even larger spacecraft mass (1000 kg or so) can be delivered to Pluto if a Titan IV Centaur is used with a 50 kWe version of SP-100 on an electric propulsion trajectorys. Such a mission, however, requires full power reactor and ion drive operation for a period of about 8 years prior to arrival at Pluto (slightly longer than the SP-100's design life and much longer than the proven lifetime of ion thNSterS), and thus poses severe reliability concerns that make such an option unacceptable to many in the planetary mission planning community. In consideration of such an option it may be fairly pointed out that, if it worked, it would represent a very capable mission, since the large power source on the spacecraft would allow for active sensing of lhe planet.

Direct trajectories (without gravity assists) to the outer planets are not practical with chemical propellants because of the very large Earth escape delta-V required, but are entirely feasible if the small NTR engine is employed. such trajectories are of considerable interest because Jupiter is only in the correct position for an outer planet trajectory assist for a certain period of time every 12 years. If funding or other delays were to force a mission relying upon a Jupiter gravity assist to postpone launch past this window, the mission would be have to be put off for

l 3 a X - m 61 7 670 890 729 774 loo0

1157 1180 1411 88 204 273

at least 12 years. Windows for direct trajectories, on the other hand, open up once a year for every outer planet, making access possible as needed. In Table 4 we show some sample results for direct trajectory missions to each of the outer planets using small NTRs for Earth escape.

Due to the timeless nature of the direct trajectory missions, a circular orbit approximation was used for the planets orbits, instead of exact trajectory data for a specific launch opportunity. This is fairly accurate for Saturn, Uranus, and Neptune, whose orbits are close to circular, but is somewhat inaccurate for dealing with Pluto's more eccentric orbit. In fact, since Pluto is currently closer to the sun than its average distance, the real flight time to Pluto on the trajectory chosen in this example would be about 15 years (assuming a 2000 launch), instead of the 19.3 years cited in Table 4. The trajectory chosen, somewhat arbitrarily, was a parabolic trajectory departing LEO with a C3 relative to the Earth of 152 km2/s2, and then leaving the Earth with a C3 relative to the Sun of 0. Elliptical capture orbits were employed at the major planets, with periapsis located 2 planetary radii from the planet's center. A periapsis of 1.3 planetary radii was employed at Pluto, since it seems unlikely that Pluto has dangerous radiation belts. The orbiter masses cited include any propellant required for in-system maneuvers. All missions are flown with a NASA LeRC baseline 10 klb NTR used to drive an expendable upper stage for Earth escape off of a Titan IV launch vehicle. It can be seen that the small NTR can enable direct missions to any of the outer planets, offering continuous access to the outer solar system. Remarkably, flight times to Saturn can be as short as 2 years.

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Small g M J4isslonS

In the sections above we have shown the utility of small NTR engines for unmanned exploration missions. It may be objected, however, that such engines are not useful for the manned Mars mission which is the primary objective of the NTR program. Such objectives are unfounded. As we shall show, properly designed manned Mars missions can be carried out with the same small NTR engines that are the basis of the unmanned program.

While it is certainly possible to design manned Mars missions involving huge masses lifted to orbit in multiple launches of heavy lift boosters, in the course of which a vast assemblage of cryogenic tankage, trusswork, plumbing, and so forth is assembled, such missions are very undesirable from the point of view of cost, complexity, risk, and basic feasibility. Much more attractive are Mars missions which minimize launch mass and which can be launched by direct throw of an upper stage lifted by a heavy lift booster, without any on orbit assembly and with minimum requirements for in space cryogenic propellant storage.

Vehicle configurations for accomplishing such more- attractive manned Mars missions with two alternative strategies, Mars orbit rendezvous (MOR) and direct return (DR), are depicted in fig.1. Both mission plans have an initial mass in LEO of 300 tonnes and are launch able all-up with two launches of a 150 tonne to LEO class heavy lift booster.

In Figure l a we show a conventional MOR conjunction mission. In this case two payloads are each launched to Mars on medium energy Type 1 trajectories using expendable stages powered by small NTRs off of separate boosters. The NTR stages are expended after trans Mars injection (TMI.) One of the payloads, which flies out unmanned, is a Mars surface habitat plus equipment and a wet Mars Ascent Vehicle (MAV) all positioned on a lander. The other payload, which flies out manned, is a Mars Transfer Vehicle (MN) habitat, a Earth Crew Capture Vehicle (ECCV), and a cryogenic hydrogenloxygen trans Earth injection (TEI) stage. Both payloads perform a low energy aerocapture into a 250 km x 1 sol elliptical orbit about Mars, after which they rendezvous and the crew of the M N transfer to the

< < * v MAV for descent to the surface. After a 1.5 year surface stay, the crew ascends in the MAV and ' rendezvous with the MTV. The hydrogen/oxygen stage is then used to send the MTV onto trans-Earth injection. Upon arrival at Earth, the crew bails out in the ECCV to perform an Apollo type direct entry, and the remainder of the spacecraft is expended.

In figure l b we show a version of the Mars Direct6 mission concept using a NTR stage and low energy aerocapture. (Aerocapture is highly advantageous for Mars missions incorporating large surface payloads, since all payloads destined for the Mars surface must carry an aeroshield in any case.) In this mission a NTR stage is used to throw a payload lifted to LEO by a 150 tonne ET0 class HLV on direct trans-Mars injection, after which the payload aerocaptures and lands on Mars, while the NTR stage is expended into interplanetary space. The payload consists of an unfueled two-stage Earth Return Vehicle (ERV) driven by methane/oxygen engines, a propellant processing plant (PPP), and about 10 tonnes of liquid hydrogen which subsequent to landing is reacted with the C02 martian atmosphere to produce the methane and oxygen required by the ERV. (184 tonnes of CH4/LOX are produced; 166 for the ERV and 18 to enable long range ground operations high-powered internal combustion engine driven rovers.) Power to drive this process can be provided by a 100-kWe power reactor landed with payload. Two years (the next Earth-Mars launch window) after the ERV has landed and fueled itself, a second HLV is launched sending to Mars a crew of 6 astronauts in a Mars Transfer and Surface Hab (MTASH), which lands in the immediate vicinity of the ERV. The crew will remain on the Martian surface for 1.5 years, exploring widely with the aid of combustion engine driven ground vehicles, and then execute a direct return to Earth in the ERV. Simultaneous with the MTASH launch, a second ERV payload is also sent out to Mars where it aerocaptures into orbit. If anything is wrong with the ERV launched earlier, this second ERV can be landed near the MTASH and used as a backup, otherwise it can be landed elsewhere to open up a new potential landing site for the next MTASH which will be launched 2 years later. Thus every two years two HLVs are launched, an average of one per year, to sustain a continuous and very robust program of manned Mars exploration.

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6

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v

1. Manned Mars missions using Practical NTR Stages. No on-orbit assembly is required.

While the Mars Direct option requires the in-situ manufacture of propellant, the processes required (a u . simple Sabatier reactor to react the hydrogen with Martian C02 to produce CH4 and H20, after which

U the CH4 is stored as fuel, while the H20 is electrolysed to produce oxygen propellant and hydrogen feedstock for the Sabatier reaction) are 19th century technology. On the other hand, the Mars Direct approach allows the use the the preferable DR mission architecture, much greater surface mobility by enabling the use of combustion engine driven ground vehicles, and a much greater useful surface payload. In the case of the conventional mission plan shown in figure l a the useful surface payload is 36 tonnes, about the same as that delivered by the reference mission in the NASA 90 Day Report. The small NTR Mars Direct mission described above has, between the ERV and MTASH landings, a useful surface payload of 11 1 tonnes, over 3 times as much as the conventional MOR mission.

By reducing mission mass to reasonable levels and eliminating on-orbit assembly or servicing, these two mission plans provide two alternative routes to a manned Mars mission using near-term technology.

V T h e question arises, however, how much could these missions be enhanced if the NTR technology level were raised beyond the modest level assumed for the small NTRs. Since the initial mass in LEO of v

these missions is fixed at 300 tonnes by the lift capacity of the launch vehicle, the relevant measure of transportation technology effectiveness is the useful payload delivered to the Martian surface. In Table 5. we show a comparison of useful payload delivered to the Martian surface for both mission plans assuming either cryogenic hydrogenloxygen, small near-term NTRs, or various levels of advanced NTRs .

In Table 5 the entries marked with an asterisk represent mission plans using 3 perigee kicks for trans Mars injection; all other entries represent mission plans where TMI is accomplished with a single burn. The thrust levels chosen have been optimized for each option. The thrust levels of 40 or 20 kibf selected for the near-term NTR (850 s Isp) missions represent pods of either four or two 10 klbf engines, depending upon whether a single burn or 3 perigee kicks are employed for TMI.

It can be seen in Table 5 that the largest mission benefit, a tripling of capability, results from dropping the conventional MOR plan in favor of the Mars Direct plan; and this should clearly be done since the in-situ propellant production technology required for the Mars Direct plan can be developed for a cost about 2 orders of magnitude lower than any of the NTR options. Having adopted the Mars Direct plan, a 48% improvement in mission capability can be achieved

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TABLE 5. Useful Payload Delivered to Martian Surface in 300 tonne IMLEO Manned Mars Missions.

. . TMI P r o u i o n isson Plan ISD Tnnl ThNst Conventional MOR irect

465 50 75 klbf 24 tonnes 75 tonnes 850 3 401 2V 33/ 36' 1021 111' 900 6 50/ 20' 36 / 37̂ 1 1 1 / 116. 950 10 65 38 117 3m 40 €6 40 124

by shifting from chemical propulsion to small near- term NTR technology for TMI. Advancing from small near-term NTR technology to the best medium-term technology NTR option, (950 s Isp, T/W=lO) however, results in a mere 5% further improvement, while even invoking the performance levels of a super advanced NTR (1000 s Isp, T/W=40) only increases the mission capability by 12% over that of the small NTR. The fact that the benefits offered by advanced NTR over near-term NTR in an optimized Mars mission plan are so marginal makes the wisdom of stretching the NTR development program to achieve such difficult levels of engine performance highly questionable.

The same point is made parametrically in the results

shown in Fig. 2., below. Here we show a comparison of the amount of mass that can be thrown on TMI for a given IMLEO by either cryogenic propulsion, clusters of small 10 klbf NTR engines with an Isp of 850 s and a T/W of 3, and larger NTR engines with a thrust of 75 klbf, an Isp of 850 s and a T/W of 5. It can be seen that even for very large scale manned Mars missions with initial masses in LEO of 400 tonnes, the lion's share of the benefit of going from cryogenic chemical propulsion to NTR is already manifesterd by the introduction of small NTRs. The larger engines, which are useless for supporting the unmanned exploration program and much more expensive to develop, offer only marginal benefits on this, their mission of choice, overthe cheaper and more versatile small NTR technology.

W

W

Trans-Mars Injection C3=15 krn2/s2 180 day transit time

600 I

500 U

400 - --1- - 300 - 200 -

100

0 0 100 200

TMI Payload

cry0 110kNTR 2 10k NTR 6 10k NTR 1 75k NTR 2 75k NTR

Fig. 2. Trans-Mars Injection Payload Capability of Chemical and NTR Engines

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IPAPS Conc- v

It has been observed on numerous occasions that space missions requiring significant amounts of electric power could potentially be enhanced further by utilizing the on-board electric power source to drive advanced propulsion to deliver the spacecraft to its destination orbit, as well as to provide the power for the mission's payload. The most obvious way to attempt to use the electric power source in such a dual role is to drive an electric propulsion (EP) system, as the high specific impulse offered by such devices frequently makes possible a significant mission mass benefit. Unfortunately, the extreme low thrust limitations inherent in electric propulsion frequently causes the transfer times it produces to be excessive, especially for precisely those large AV missions for which the EP system offers a significant mass saving relative to chemical propulsion. However, if the electrical power source is nuclear, it is also a source of thermal power at levels an order of magnitude greater than its electrical output. This makes possible a different form of use for the power source, with direct thermal power generating rocket thrust without the intermediate step of electrical conversion.

A thermionic reactor concept has recently been proposed that is particularly amenable to such dual- mode utilization. Known as the SEHPTR (Small Ex- Core Heat Pipe Thermionic Reactor), this concept incorporates tungsten clad U02 fuel elements separated by small gaps from thermionic heat pipe modules, which convert about 10% of the core's thermal power to electricity and then transport the remaining heat to a radiator for disposal. Heat is transferred across the gaps from the fuel to the thermionic heat pipes via radiation. The gaps were

originally conceived for the purpose of allowing non- nuclear testing of a complete system (by replacing the tungsten clad fuel with a tungsten clad electrical heating element) and to give the system a much longer operating life than is possible in a conventional thermionic reactor (which bind thermionic conversion systems around the fuel in the form of thermionic fuel elements, or TFEs) by allowing for an order of magnitude greater tolerances for fuel swelling. Due to the SEHPTR reactor's unique construction, tungsten clad flow channels (Le. the gaps) are available through the core which provide a means of directly heating hydrogen propellant to a temperature of 2200 K. While the 700 to 750 s specific impulse offered by such direct thermal thrusters is well below that offered by E.P., it is still much greater than that possible with chemical propulsion, and for missions of interest offers the lion's share of the mass benefits available with EP. Moreover, the level of thrust such systems make possible may be 2 orders of magnitude greater than that of the same reactor's EP capability, with a corresponding drastic reduction in the required trip time. Furthermore, since the thermal thruster uses the same propellant as one of the main EP options (the hydrogen arcjet), both systems can be integrated on the same stage, so that the EP option can be taken advantage of for those operations where time is not of the essence. Thus an electric power reactor design, such as the SEHPTR, which is amenable lo the production of direct thermal thrust as well, allows the creation of an integrated stage incorporating the symbiotic utilization of electric power, direct thrust, and hydrogen arcjets. Such a stage, termed the Integrated Power and Propulsion Stage (IPAPS, Fig. 3) offers a superior option for accomplishing many space missions.

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/swR7 SEHPTR Outer Satellite Thrust Structure 4.6 m Dla. ,

LH2 Tank 4.27 m Dla. 7.0 m Cylinder Length

f i Avionics ’ - w o r m

h 8x Active RCS - A / Radiators

\ RCS Mounting Ring & LH2 Tank Support Rlnc

LTotal Length 15.1 m I Figure. 3. An Integrated Power and Propulsion Stage (IPAPS) Designed for Titan IV Launch

The Use o f IPAPS for Earth O r b l t a l Missfons

In Figs. 4,5,6,and 7 we show a comparison of the payload delivery capability of the IPAPS, a thermionic power source with electric propulsion (hydrogen arcjets) only capability, lightweight and survivable solar photovoltaic systems employing electric propulsion, and a cryogenic chemical propulsion stage. In Fig. 4 a we see a study of a GEO mission launched off of a Atlas 2AS. It can be seen that a 40 kWe thermionic reactor such as the SEHPTR can use its electric propulsion capability to deliver itself and its arcjet system plus 1500 kg of instruments from a 300 km LEO obit to GEO in about 150 days. By contrast, if the dual mode capability is employed, the system can deliver about 250 kg of instruments plus the 40 kWe reactor in about 10 days. Such a large difference in delivered payload would probably indicate the arcjets as the SEHPTR’s preferred mode of transfer for the mission thus defined. However,

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the large majority of space mission require a much lower power level than 40 kWe, and if we consider Fig. 4 with a 10 kWe power level, we see that the dual mode IPAPS can deliver 1900 kg of instruments in 40 days, which compares quite favorably with the 500 kg deliverable by chemical propulsion, or the unacceptable 600 day transfer time required by all the electric propulsion options.

The GEO mission, while important, is only one of a number of interest to NASA and the US. DOD. Also of importance is payload delivery to highly elliptic orbits, which are less kind than GEO to the orbit transfer characteristics of electric propulsion. In Fig. 5, we show payload delivery of an Atlas 2AS from LEO to a 24 hour period highly elliptical orbit. We see that for all power levels between 5 and 40 kWe, the dual mode IPAPS delivers about 50% more payload than a pure nuclear electric system, and does so with about 5% the transit time.

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GEO Mission Launched bv Atlas 7000 ka Initial Mass In LEO

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Power ievels of 5, 10, 20 and 40 kWe

- IPAPS - NEP - Lightweight Solar - Survivable Solar * Clyo

O A I - I I I

0 2 0 0 400 6 0 0 8 0 0

Trip Time (days)

Fig. 4 Comparison of Upper Stages for GEO Mission Flown Off of Atlas 2AS

HEEO Mission Launched bv Atlas

4000 1

1000

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7000 kg Initial Mass In LEO Power levels of 5, 10, 20 and 40 kWe

IPAPS - NEP - Lightweight Solar - Survivable Solar * Glyo o ! I / I I I

0 2 0 0 400 600 800 Trip Time (days)

Fig. 5 Comparison of Upper Stages for High Elliptic Orbit Mission Flown Off of Atlas 2AS

GEO Mission Launched bv Titan IV 20

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8000

- 0)

5000 kg lnltlal Mass in LEO Power levels of 5, 10, 20, and 40 kWe

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20 - . 40

- NEP - Lightweight Solar - Survivable Solar * cry0

4000

3000 - a m L c

-

0 1 I I I 0 200 4 0 0 6 0 0 800

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Fig. 6 Comparison of Upper Stages for GEO Mission Flown oft of Titan IV

- NEP - Lighfweight Solar - Survivable Solar

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would have to use a chemical transfer stage to deliver it to its destination orbit Since the SEHPTR/IPAPS would not have to do this, the result is that a 5 kWe SEHPTFVIPAPS would be able to deliver 4 times the payload to orbit as a 5 kWe solar electric system. Thus, the dual mode capability allows the SEHPTR to out-compete photovoltaic power in the low power range, where the majority of current missions are planned. By contrast, non-dual mode nuclear electric systems cannot compete with photovoltaics in this range, because the unfavorable mass scaling of low-power nuclear systems cause then to be somewhat heavier.

The dual mode SEHPTR offers many advantages to the user beyond the pure performance benefits associated with greater payloads and much shorter transfer times as such. Reducing the transfer time also reduces the amount of reactor usage prior to attaining the payloads duty station. This increases the mission reliability since for the same payload operational lifetime, the required reactor lifetime is reduced by the amount of time saved in transit, about 6 months for the baseline mission case. Mission reliability is also increased by the fact that the SEHPTR/IPAPS has two different modes of propulsion, thus a failure of the arcjet system need not abort the satellite transfer and end the mission. The total radiation dose delivered to the payload from the reactor prior to the commencement of mission operation is reduced by more than an order of magnitude, as is the dose received by the spacecraft systems from the Van Allen Belts. The dual mode SEHPTR also reduces cost and increases mission reliability for those missions which cannot accept the 6 to 12 month transit times associated with electric propulsion. Such missions ordinarily would require a dedicated transfer stage, either chemical or nuclear thermal. The use of such a stage increases the expense of the mission, and also risk of mission failure, since if eitherthe transfer stage or the reactor power source fails the mission is lost.

The reliability of the dual mode SEHPTR reactor is also improved, relative to other concepts, by the ease with which complete and intact units may be ground tested. By using hot hydrogen gas produced by an outside source to heat the reactor by flowing it through the hydrogen channels, the entire system may be brought to an elevated temperature and the thermionic power conversion system verified prior to launch. This is not possible with competing concepts, which cannot be brought in an intact state to their operating temperature

without being made critical and thus producing an unacceptable radiological launch hazard.

Some missions of interest to NASA and the DOD may require the ability to reposition the satellite, either rapidly or slowly. Electric propulsion systems are ideally suited to accomplish slow transfers, however to accomplish fast maneuvers pure electric systems must carry a contingency mass of chemical bipropellant. The dual mode SEHPTR/IPAPS can accomplish all required maneuvers quickly using its thermal thruster, or slowly using its hydrogen arcjets, Since these two systems use a common propellant, no preallocation of propellant between high and low thrust maneuvering systems is needed, and thus on- orbit maneuvering capability is maximized.

The environmental risk posed by the dual mode SEHPTR reactor is much less than that associated with pure electric thermionic systems, due to the fact that it generates a factor of 50 less radiological inventory for a given amount of rocket impulse, as shown in Fig. 8 Thus if the reactor operates a certain amount of time and then fails, the dual mode SEHPTR/IPAPS will represent a much lower hazard to the public. For the same burn time it will have attained a much higher orbit with a much longer orbit decay time than a pure nuclear electric system, or alternatively, for the same orbit transfer it will contain only about 2% the radiological inventory as a pure electric concept. A comparison of the radiological inventory in the system at the time of re-entry following system failure during climb from a 300 km LEO orbit is given in Fig. 9.

The possibility of a reactor reentry is also reduced for the dual mode SEHPTR, since if the reactor has to be shut off early in the mission, a significant amount of orbit raising can be accomplished using decay heat alone. The dual mode flow channels also add to system safety, by providing a secondary method of cooling the core should there be a complete failure of the radiator cooling system. Such a backup cooling system, analogous to the reserves of water held in cooling ponds near commercial nuclear power plants, is not available to alternative pure electric thermionic concepts. These feature may make the dual mode SEHPTRAPAPS more politically acceptable than alternative concepts, and may also enable it to operate from lower orbits, thus expanding the payload capability of the launch vehicles supporting mission operations.

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Radiological Hazard to Environment

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c 0

- IPAPS - Arcjets - Ion

rn 0 -2 3 -I - 2 - 1 0 1 2

Log (10) of T h e After Start of Climb (days)

Fig. 9 Radioactive Inventory upon Re-Entry After Failure During Climb of Nuclear Systems

The dual mode SEHPTR/IPAPS can also be used as an open cycle electric power system by venting the hydrogen flow through a turbine instead of a rocket nozzle. This allows it to function as a source of burst electric power in the 1 to 5 MWe range. This capability is not shared by akernative pure electric systems.

The Use o f IPAPS f o r ExDlOratIOn MlssionS

A host of potential interplanetary exploration applications have been identified for space nuclear

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power and nuclear electric propulsion, including unmanned probes to the asteroids and outer planets, comet rendezvous missions, and surface power to support human exploration and settlement of the Moon and Mars. The SEHPTR is capable of supporting all of these missions, as also are various alternative pure electric concepts such as SP-100. Because, however, it is also useful for Supporting the far more frequent (by more than an order of magnitude) Earth orbital applications, while pure nuclear electric systems generally are not, the rationale supporting the development of SEHPTR is

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much stronger. Thus, while to-date the programs for - developing space nuclear power have drifted in an entropic manner for lack of a sufficiently large defined user community, a SEHPTR program could be launched on a fast and secure track, supported by a large and varied customer base. Put simply, in today’s funding environment it is unrealistic to imagine that space nuclear power can be developed solely for the purpose of supporting interplanetary missions. Thus i f interplanetary missions incorporating space nuclear power are ever to occur, the nuclear technology proposed for development must be one whose capabilities are synergistic with the requirements of Earth orbital applications. As we have seen in the section above, only by incorporating a dual mode capability can space nuclear power make itself truly competitive with photovoltaics for these applications.

Dual mode capability is thus in fact enabling for any interplanetary mission plan utilizing nuclear electric power. However, beyond being essential from a programmatic viewpoint, the direct thrust capability of a dual mode system also enables or strongly enhances a host of lunar and interplanetary exploration mission applications. For example, the AV required to transfer from LEO to low lunar orbit

w (LLO) by high thrust maneuvers (or repeated perigee kicks) is 4.2 krnfs, essentially identical to the high thrust AV from LEO to GEO. The electric orbit transfer AV, however, is about 9 kmls, due to the large gravity losses associated with outward and inward spirals. The result is that a 40 kWe IPAPS flying off of a Titan IV (20 tonne to LEO capability) can deliver itself plus 6.2 tonnes of payload from LEO lo LLO in about 30 days, while a 40 kWe pure nuclear electric system employing hydrogen arcjets can deliver about 5.9 tonnes of payload, but require about 690 days for the transfer. Since this transfer time is too long (and would be even longer if ion engines were employed), the pure electric system would have to be delivered to station by a chemical stage. If this were the cryogenic Centaur, the deliverable payload in addition to the reactor would be about 2 tonnes, less than a third of the IPAPS capability, and the cost of the Centaur would be added to the mission. The large payload and high power available to the IPAPS would allow a very capable system to be placed in LLO, and many

w exciting applications would be possible. One such application would be to orbii the Moon with a platform capable of active sensing, using the IPAPS’s power to generate laser or particle beams that could ” vaporize small amounts of targeted Lunar surface

materials, which could then be analyzed spectroscopically from orbit. This would allow for detailed, high resolution, chemical mapping of the Moon. Similar mission strategies could be pursued to explore Mercury, Mars, and the moons of Jupiter and Saturn, all of which the IPAPS could access with little more propellant expenditure than that required for the Lunar mission by injecting itself from LEO onto a trans-Venus or trans-Mars injection trajectory, and then proceeding from there via a Venus or Mars swingby gravity assist. Radar mapping of Venus and Titan, and active electromagnetic probing of the atmospheres of all of the major planets would also be possible with a similar flight plan.

In order to attain the required trans-Venus or trans Mars injection velocities (C3’s of about 16 km2/s2) to enable such gravity assists, the IPAPS can use 2 tactics. One is to do all of the propulsion itseit, which would involve massive gravity losses since after a AV of 3.2 km/s to reach Earth escape from LEO during a series of perigee kicks, another 4 kmls would be required to attain the required hyperbolic excess velocity using the IPAPS‘s modest (50 Ibs) direct thrust capability. The propellant mass penalty associated with this extra AV can be mitigated by using the IPAPS’s arcjets for the post-escape burn, with a post-escape burn time of about 180 days required for a 40 kWe system. However if instead, the IPAPS employs its direct thrust engine in a series of perigee kicks only to bring the spacecraft into a highly elliptic orbit (just short of escape energy) and then a small high thrust chemical engine is kicked in for the escape burn, gravity losses can be eliminated. The required AV for interplanetary injection than must be accomplished by the chemical engine is then only about 0.8 kmls. Since the IPAPS itself only had to perform a 3.1 kmls AV, the total propellant expenditure is similar to that required by the IPAPS for the LEO to GEO or LEO to LLO missions described above, while the time required for Earth escape is even less. This latter orbit transfer scheme has been shown to be advantageous in prior studies for use by solar thermal propulsion on interplanetary missions7, and is clearly ideal for adoption by IPAPS.

Another exciting exploration application that has been suggested for the IPAPS is a main-belt asteroid exploration cruise8. Current asteroid exploration plans involve using chemical propulsion to send a small spacecraft to rendezvous with a main- belt asteroid, and then stationkeep with it for several months in order to be able to assess its composition

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via passive sensing techniques. If conditions allow, another asteroid may be visited afterwards. but probably not more than one, since the Hohmann transfer from one asteroid to the next (for low AV rendezvous) will take about 3 years. Since there are about 3000 currently known main-belt asteroids, sampling only one or two per mission leaves much to be desired. In contrast to this concept, an IPAPS asteroid exploration mission offers many advantages. The IPAPS would be launched into an elliptical orbit with an aphelion of about 3 AU, thus causing it to constantly cross the main-belt. Using active sensing capabilities, the composition of numerous asteroids could be rapidly assessed during large numbers of close fly-bys, while the IPAPS's dual mode capability and large injected payload (which would include a large amount of reserve propellant) would give the system an enormous reserve A V to use to provide for numerous mid-course correction maneuvers to assure that the fly-bys were sufficiently close. A pure electric system could not do this mission anywhere near as well, because to avoid a 600 day Earth escape burn, chemical propulsion would have to be used for interplanetary injection, which would drastically cut the available reserve propellant load. Furthermore, the low thrust available from an electric propulsion only system would limit the versatility of the repertoire of course change maneuvers available to the spacecraft, making many asteroids inaccessible.

Nuclear electric propulsion has been advanced as a possible means of enabling Pluto and other outer planet orbiter missions with large payloads and comparatively short transfer limes. For all such missions, the IPAPS outclasses a pure electric system, as the system's direct thrust mode allows it to deliver triple the payload of spacecraft and propellant (Le. beyond the reactor mass) as would be possible by a Titan IV-Centaur onto an initial interplanetary trajectory, thus enhancing the mission capability by more than a factor of 3. A pure electric system could equal or beat this performance in payload delivered by using its NEP capability to drive all the way from LEO to the destination planet, but only at the cost of facing the reliability problems associated with extraordinary amounts of reactor burnup, thruster use, and trip time prior to arrival at the destination planet.

Conclusion

The key to the successful development of nuclear

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w power and propulsion technology is versatility, or the ability of the technology selected for development to find synergistic uses across a broad range of potential mission applications. In the case of high thrust NTR propulsion, this consideration leads forcibly to the conclusion that the engine selected for development be small, i.e. in the 10 l o 20 klbf range. Such an engine is required for all unmanned NTR mission applications, and is fully capable of supporting the manned Mars and Lunar mission application as well. By contrast. a large (75 klbf range) engine designed soley for use in supporting manned missions is likely to prove a dinosaur. Similarly, if nuclear electric power is desired for interplanetary exploration applications, necessity dictates that the technology chosen for development must be useful for supporting the much more common Earth-orbital mission application. A detailed analysis of Earth orbital power applications shows that only a dual-mode nuclear electric system offers significant advantages over current photovoltaics, and thus reason dictates a strong preference for a dual-mode system if interplanetary nuclear electric missions are hoped for.

References

1. S. Borowski, Private communication, April 1992.

2. A. Friedlander, Private communication. May 1992.

3. R. Staehle, P. Henry, C. Salvo, R. Wallace, and S. Weinstein. "Pluto Mission Development Status Report." Presented at the Solar System Exploration Division Advanced Studies Branch Quarterly Review, JPL, Pasadena, CA, May 20,1992.

4. S. Weinstein. Private Communication, may 1992

5. J. H. Kelley and C.L. Yen, "Planetary Mission Opportunities with Nuclear Electric Propulsion," AlAA 92-1560, AlAA Space Programs and Tech. Conference. Huntsville, AL, March 24-27, 1992.

6. R. Zubrin, D. Baker, and 0. Gwynne, "Mars Direct: A Simple, Robust, and Cost-Effective Architecture for the Space Exploration Initiative," AlAA 91 -0326, 29th Aerospace Sciences Meeting, Reno NV Jan. 1991.

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7. A. Bruckner, Private Communication

8. G. Briggs, Private Communication