Vol. 1-467

Pluto Orbiter/Lander/Sample Return Missions Using the MITEE Nuclear Engine

Jamcs Powell, George Maise, and John Paniagua Plus Ultra Technologies, Inc.

PO Box 547 Shoreham, New York 1 1786 PhonelFax: 631 -744-5707

in7 wol6 m t on1 i nc net

Ab.~/i*uc~- Pluto OrbiterAander and sample return missions are not impossible using chemical propulsion, but are possible with nucle*ar thermal propulsion. Using the MITEE nuclear engine, a spacecraft could first orbit Pluto, mapping it, and then land at a selected site, 12 years after the departure from Earth. If surface waterhce is available, fresh H: propellant could be manufactured by electrolysis of melt 60 using power from the bi-modal nuclear engine, enabling multiple hops to new sites for further data collection. A warm water probe could also be deployed to explore the interior of Pluto's ice sheets. After completing exploration, the spacecraft could return samples from Pluto to Earth with a 12-year trip time. Mission architectures and the design of the spacecraft, nuclear propulsion engine, propellant manufacturing unit and warm water probe are described herein.

TABLE OF CONTENTS

1. INTRODUCTION ................................................................... 1 2. OPTIONS FOR A PLUTO MlSSlON .................................... 2

SAMPLE RETURN CONCEPT .............................................. 6 4. THE MITEE NUCLEAR PROPULSION ENGINE ............. 8

IN-SITU RESOURCES ......................................................... 18

MISSION .............................................................................. 20 7. CONCLUSIONS ................................................................... 22 REFERENCES .......................................................................... 22

3. DESCRIPTION OF THE PLUTO ORBlTER/LANDER/-

5. REPLENISHMENT OF H2 PROPELLANT FROM

6. TECHNOLOGY DEVELOPMENT FOR THE PLUTO

1. INTRODUCTION Pluto remains the most mysterious object in the Solar System. No detail is known about its origin, surface topography, composition, geology, internal structure, and atmospheric behavior.

The proposed New Horizons mission would provide a quick snapshot of Pluto, and its moon Charon, as its spacecraft flew by at high speed a decade after its launch

' 0-7803-7651 -X/03/$17.00 0 2003 IEEE ' IEEEAC paper #1088, Updated October 10,2002

from Earth in 2006. A fast flyby is the best that can be accomplished using chemical propulsion. The mission would cost over 500 million dollars. Moreover, if not launched in 2006, the window for the required Jupiter gravity assist will be gone; with the result that the next date will be too late to observe Pluto's expected atmospheric condensation in -2020 AD.

Nuclear propulsion offers the opportunity to obtain much more data about Pluto, because of the much hijgher Isp and AV's that it enables, relative to chemical propellants.

There are 2 basic nuclear propulsion methods for a Pluto mission - nuclear electric and nuclear thermal. [A third hybrid method also exists, involving a combination of the 2 basic methods, which may be advantageous in certain cases. This is discussed later.]

In the nuclear electric propulsion method, a nuclear reactor would generate electricity for an ion thruster, or other type of electric propulsion unit. The spacecraft would slowly accelerate to its maximum velocity over a period of several years, and then slowly decelerate for several more years as it approached Pluto and finally went into orbit.

While nuclear electric propulsion could enable a spacecraft to orbit Pluto, it has a number of sgnificant disadvantages: 1. The propulsion system involves complex equipment

that must be developed to operate reliably for many years. The spacecraft cannot land on Pluto unless it were to use a chemical propulsion system for the landing, because the available thrust liom the electric propulsion unit is far too small. The spacecraft cannot hop to multiple sites on Pluto to obtain a detailed understanding of its geology, composition, etc. The spacecraft cannot return samples from Pluto to Earth.

2.

3.

4.

With regard to the first point, Woodcock [ I ] has analyzed various nuclear electric missions to the outer planets. A nuclear electric orbiter mission to Pluto would take

Vol. 1-468

Launch Payload to LEO

approximately 12 years and require a nuclear electric power system of -100 KW(e), for a total lMLEO (Initial Mass on Low Earth Orbit) of 10,000 kilograms.

Perform 1st NTP Perform Td NTP Burn to Enter I

Earth Orbit Increase * V Pluto Orbit

Developing an integrated reactorlpower conversionl- electric thruster system that could operate reliably for 10 years or more will require a major development and testing effort that will take many years to carry out. The time scale for the development and implementation of a nuclear electric mission to Pluto is not consistent with observing atmospheric condensation on Pluto in 2020 AD.

Explore Site - Deploy Warm Perform

With regard to point 2, including the mass of a chemical propulsion unit to land on Pluto would increase the IMLEO requirement to an impractical level. With regard to points 3 and 4, even if a chemical unit were used for a landing, it would not be possible to replenish the spacecraft propellant for multiple hops or the return to Earth. t

Probe- Manufacture H,

Propellant- Transmit Data

- Burn-Hop 4 to Next Site

The nuclear thermal propulsion method offers much greater capability over nuclear electric for the exploration of Pluto. Its advantages include: a) The nuclear propulsion system operates for relatively

short periods, i.e., hours instead of years. b) The nuclear propulsion system enables the spacecraft

to land on the surface of Pluto within acceptable IMLEO requirements.

c ) The nuclear propulsion system can be replenished with 6 propellant obtained from in-situ wateriice,

V Perform 3'd NTP Map Pluto - J Burn to Land on 4 Transmit Data to

Pluto Surface Earth '

enabling it to hop to multiple sites on Pluto, and return samples to Earth.

Pluto Exploration Complete -

With regard to the point a), the nuclear propulsion system can be validated using a relatively short period of testing, enabling it to be potentially available for arrival at Pluto by 2020 AD.

Perform Final Coast to Earth - Land BumtoDepart b on Board Power b Samples

for Earth Only (-1 0 YTS) on Earth

With egard to point b), the propulsion burn to land on Pluto is no different than the bums to depart from Earth orbit or to enter Pluto orbit, so that no additional technology or components are needed. With regard to points e), if waterlice is available on Pluto, H2 propellant can readily be manufactured by a simple, low weight (30 kg) electrolysisiliquefier unit that is based on existing commercial technology.

For the above reasons, some type of nuclear propulsion appears to be the best choice for exploration of Pluto. The next section describes the various options for using nuclear thermal propulsion for a Pluto mission.

2. OPTIONS FOR A PLUTO MISSION

Figure 1 shows 2 different options for missions employing nuclear thermal propulsion (NTP) to Pluto. Both options use a high thrust NTP burn to depart from Earth orbit, followed by a NTP burn to orbit Pluto, with a 3rd burn to land on the surface.

I

Figure 1 Nuclear Thermal Propulsion Options for Pluto Mission

Vol. 1-469

0 II 2 3 4 5 6 7 8

Time After NTP Burn In LEO 4vcaarsl

Figure 2 Comparison of V Capability of Bi-Modal MITEEB Engine with NTP only MITEE Engine

Both options employ a bi-modal nuclear engine in which the high thrust NTP mode is used when departing from Earth orbit, entering a Pluto orbit, landing on Pluto’s surface, and hopping to multiple sites on the surface. (Mid course corrections can also be made using the nuclear engine.) The principal difference between the two options is that in option A, the spacecraft coasts on a direct trajectory (no gravity assists, which greatly enlarges the launch window opportunities), while in option B, the spacecraft acquires additional AV along its trajectory by using electric power generated from the bi-modal engine to drive a high specific impulse electric propulsion unit, e.g., an ion engine or other device, thereby, reducing trip time. This hybrid, NTP/NEP option has the advantage of shorter trip time, but is offset by the additional equipment and propellant required for the electric thruster, and the need to develop an electric thruster system that can operate reliably for many years in deep space.

The electric propulsion unit probably would not be able to be used for a return trip to Earth, since the spacecraft could not replenish the required propellant from in-situ resources on Pluto and sufficient propellant could not be carried from Earth. As a result, only the MP mode would be used for the return trip.

A comparison of the NTP only and the hybrid NTP/NEP options shows virtually no benefits from the latter. This is the case whether one compares total mission time (Earth departure to Earth return) assuming equal IMLEO’s, or the IMLEO requirements, assuming equal

trip times. This results from a combination of several factors. Taking, for example, the case where equal IMLEO’s are assumed, the following factors apply: 1. The NTP only system delivers a considerably greater

AV increment in a very short time (-1 hour) at Earth departure than does the NTP/NEP system. This is due to the reduced amount of Hj propellant for the NTP/NEP system, since a substantial portion of its IMLEO is devoted to the equipment and propellant for the NEP sub-system. This deficiency in AV increment is more than made up for by the NEP engine, but it takes many years to do so, and a very long distance to catch up. Both systems have to make a substantial AV NTP burn to orbit and land on Pluto, which further reduces the amount of propellant available for NEP operation. The return trip to Earth can only be done by an NTP burn, since the NEP propellant is not available on Pluto.

2.

3.

To illustrate this, Figure 2 compares the AV for the NTP only and the NTPiNEP options as a function of time after departure from Earth orbit.

Both missions are assumed to start from LEO, with an initial IMLEO of 10,000 kilograms, on a trajectory out of the Solar System. The NTP only MITEE-B spacecraft has an initial V of 20 kilometers per second from its NTP bum and then coasts outwards, with no further propulsion. The NTP/NEP MITEE-B (Mhature &eacIor - EnginE-Bi-Modal) spacecraft receives an NTP burn of

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10 kndsec, which leaves in addition to the 450 kg payload and 250 kg MITEE-B engine mass, a mass of 2000 kg of high Isp propellant (Isp = 5000 sec or equivalently 5000 kgf sec/kg or propellant), 300 kg of electric power system and thruster mass, and 100 kg of propellant tankage.

After the initial NTP burn, the electric thruster begins operation, with an overall efficiency (input electric power to directed kinetic power of the high Isp propellant jet) of 50%. As the 20 kW(e) electric power/thruster system continues operation, an additional V is imparted to the spacecraft.

At a little over 2 years the V imparted by the NTP/NEP MITEE-B engine reaches 20 km per second, the same as the V initially imparted by the NTP burn of the NTP- only engine. However, since NTP-only started with a greater initial velocity than NTP/NEP-only, it will be far ahead of NTP/NEP.

The NTP only system reaches Jupiter’s orbit in 0.6 years, while the NTP/NEP system takes 1 year to reach the same point. Similarly, the NTP-only engine reaches Saturn’s orbit in 1.1 years, while the NTP/NEP engine takes 2+ years. However, i t continues to impart V to its spacecraft reaching a total - V addition of 60 kmisec in 8 years, ensuring that at some point the N?’P/NEP engine will catch up to the NTP only system and from then on, leave it far behind. The catch-up point occurs in 5 years at Pluto’s orbit.

The NTWNEP engine would reach the 100 AU point in the Kuiper Belt at 9 years, while it would take 12 years for the NTP only engine to get there. Many other mission scenarios could be developed for the NTP/NEP system to show its potential for fast trip times to near interstellar space. This example, however, illustrates the attractive capability of the NTPiNEP system for missions beyond the bounds ofthe Solar System.

Clearly any gains in trip time achieved by the hybrid NTP/NEP option will be marginal at best, even for a pure flyby mission. When the additional requirements of orbiting and landing are included, which will require carrying a larger fraction of the spacecraft IMLEO to be devoted to the necessary burns, it appears doubtful that the hybrid NI’P/NEP option will show any advantage in either trip time or mission IMELO.

The lack of a clear mission advantage, plus the additional requirement for operating an electric thruster at high electric powers for many years, makes the hybrid NTP/NEP option less attractive than the NTP option for the Pluto orbiter/lander/sample return mission.

The landing site would be chosen after a detailed mapping of the surface of Pluto from orbit. If a surface waterhce sheet or sheets are detected, a landing site on an ice sheet would be probably be selected. Data on atmospheric composition and properties (temperature, pressure, movement, etc.) would be obtained and samples of the local terrain collected. A warm water probe would be deployed to melt a channel down through the ice sheet to obtain data on the composition and properties of the ice sheet interior.

The melt probe would use reject heat from the electric generation operation of the bi-modal nuclear engine to melt the channel. With a 20 KW(e) generation capacity, approximately 80 KW(th) of heat at -100” C would be available for the probe. This would permit a melt descent rate of -100 meters per day. The warm water would circulate through a small dianieter line attached t o the probe. Data on the ice sheet structure and composition would be transmitted to the spacecraft in real time through an optical fiber link to the probe. Ultrasonic pulses could be used to map the extended structure of the ice sheet before deploying the probe.

The probe could descend deep into the interior of the ice sheet with a maximum descent capability of several hundred meters. The actual descent depth would depend on the site characteristics. Upon reaching its desired depth, the probe would remelt its way back to the spacecraft, letracing the original descent path. Samples collected during the descent/ascent process would be transferred to the spacecraft.

The maximum depth that the probe can penetrate into the ice sheet is constrained by the amount of flexible tube that would convey the warm water from the spacecraft to the probe. This tube is relatively small, -1 cm in diameter and made of lightweight thin wall material that can be collapsed into a compact package attached to the probe. Once expanded by pressurized hot water that flows down to and back from the probe, the tube will maintain its shape.

Much greater penetration depths are possible using a small, lightweight nuclex-powered melt probe that generates its own heat, and does not require piping warm water from the surface. Figure 3 shows one possible design approach for such a probe, based on a small (40 centimeter OD) water cooled and moderated nuclear reactor. It would use cermet fuel like that employed in present defense reactors, consisting of a metal matrix (e.g., zirconium) in yhich was imbedded tiny (-10 micron) particles of L’’UOz. Such fuel has proved highly reliable and capable of high burnups. The reactor core would consist of an assembly of cermet fuel that was cooled and moderated by water. To minimize weight, the

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water for the reactor would probably be obtained from melting H20 ice from the surface of Pluto's ice sheet.

The cermet fuel can be arranged as an assembly of spaced small diameter (e g , 3 millimeter) rods or thin plates, with axial flow of the water coolant/moderator between them, as is usually done. Alternatively, the thin fuel sheets could be fabricated with perforations and rolled into annular fuel elements, with radial flow of the water coolant/moderator through the moderator, as illustrated in Figure 3 . The axial How geometry would permit power outputs of up to -1 megawatts. Even higher power outputs would be possible using the radial flow geometry. Most of the thermal power generated by the melt probe reactor would be used to heat water for the melt process. A small fraction of the coolant channels would run hot enough to generate steam for a mini-turbine that would

produce several KW(e) electric power for instruments, communication and controls.

The reactor melt probe would move downwards into the ice sheet as illustrated in Figure 4, pumping out hot water to melt the surrounding ice and create a channel through which it descended. A small diameter optical fiber would unreel from the probe, enabling real time 2-way communication between the spacecraft and the probe. The melt chaimel would refreeze to ice a few meters behind the descending probe, with the optical fiber frozen in place. To return to the spacecraft on the surface, the probe would reverse the direction of the water jets, and change its buoyancy from slightly negative to slightly positive, and melt its way back up the original descent channel to the surface (Figure 5).

Figure 3 SUSEE Fuel Element and Reactor Configuration (Space NLklear Steam Bectric'gnergy)

r&tlh'

Figure 4 View of MP Unit as it Descends Through Pluto Ice Sheet

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Figure 5 View of MP Unit as it Ascends Through Pluto Ice Sheet

An ascent/descent rate of several hundred meters per day is achievable with a melt channel diameter of -1 meter at a power output of a megawatt. The maximum descent depth for such a probe would be several kilometers. Auxiliary electric power to operate instruments, communications, pumps, etc. would be generated by a mini-turbo generator utilizing a small fraction of the water coolant flow to generate steam.

During the period that the melt probe was exploring the interior of the ice sheet, the bi-modal nuclear engine would be generating electric power for the electrolysis unit. At an power level of 15 KW(e), 7 kg per day of liquid H2 could be generated, enabling the spacecraft to make multiple hops to new sites on Pluto. 5 KW(e) of the 20 KW(e) total electric output from the bi-modal engine would be used to operate the controls, communications, and instrumentation on the spacecraft, after completing its exploration mission on Pluto, the spacecraft would accumulate sufficient liquid E$ to return to Earth with collected samples.

3. DESCRIPTION OF THE PLUTO ORBITERhWDER/SAMPLE RETURN CONCEPT

Figure 6 shows the IMLEO requirements as a function of trip time to Pluto, using the MITEE-B engine: 1. Simple flyby of Pluto

2. Orbit only mission 3. Lander mission

The curves are based on a total spacecraft dry mass of 800 kg, including engine, structure, and payload, plus a propellant tank mass equal to 5% of the propellant mass carried. The MULIMP code was used to predict the IMLEO values. Direct trajectory flights were assumed with no planetary gravity assists at Jupiter or Saturn. The specific impulse of the MITEE-B engine was 950 seconds.

The flyby-only mission has a very low IMLEO requirement, only 3300 kg for a 7 year trip time - a much smaller IMLEO than for the New Horizons flyby, with 110 need for a Jupiter gravity assist. Even shorter flight times to Pluto, i.e., 5 or 6 years, would have attractively low IMELO's.

The IMLEO values for the orbiter only and the lander missions are considerably larger, because of the requirements to carry additional propellant for the bum at Pluto to orbit or land. The IMLEO values rapidly increase for flight times less than 12 years. Flight times of 10 years appear practical, through IMLEO values are large, i.e., 22 metric tons for the lander mission. Flight times that are substantially less than 10 years are probably not practical

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MISSION

PLUTO FLY-BY

Pluto Orbiter

Pluto Lander

Pluto Sample Return (OutBound)

Pluto Sample Return (Return With

Indigenous Refueling, Earth Aero-brake Re-

En*)

TOF

Delta V Launch Date/ (km/sec) Trip Time

11.93 Feb. 26,2010

20.43 Feb. 18,2010

20.64 Feb. 17,2010

20.64 Feb. 17,2010

12.22 Dec. 12,2023

7 Years

1 1 Years

12 Years

12 Years

12 Years

IMLEO Values (Kiloqrams) IMLEO IMLEO IMLEO

(years) Pluto Fly-By Pluto Orbiter Pluto Lander 7 3300

10 2550 16750 22420 11 2450 1 1900 15240 12 2400 9450 12400

IMLEO vs. Time-of-Flight (TOF)

25000

20000

Y 15000 0

A En

3 10000 z 5000

0 6 7 8 9 10 11 12 13

TOF (Years)

+Pluto Fly-By --C Pluto Orbiter

-A- Pluto Lander

Figure 6 IMLEO vs Time of Flight for Potential P l u t o Missions Using the MITEEB Engine

TABLE 1 POTENTIAL BASELINE MISSIONS TO PLUTO QVABLED BY THE MITEEB ENGINE

IMLEO (kg) 3300

1 Stage/No Restart 11900

1 Stage/Restart 12400

1 Stage/Restart 12400

1 Stage/Restart 1950

1 Stage/Restart

Payload (kg) 500

500

500

500

150

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1 Vol. 1-474

Based on these results, a baseline flight time of 12 years for the orbiter/lander/sample return mission was selected. The AV, IMLEO, and launch date for the mission are summarized in Table 1. The AV is 20.6 km per second - a value far beyond the capabilities of chemical propulsion. Total IMLEO is 12.2 metric tons, with a launch date in February 2010. Also shown for comparison are a pure orbiter mission ( I 1 year flight time) and a pure flyby mission (7 year flight time).

The baseline mission would land on Pluto 12 years after departing Earth orbit, and then spend a year hopping to multiple sites on Pluto (assuming that surhce water ice is available For the manufacture of H2 propellant). It would then take off for Earth with collected samples. Again assuming that surface water ice was available for hydrogen production, with a 12-year return ffight time. The AV required For the return trip is much less than the outbound trip, i.e., 12.2 km per second compared to 20.6 km per second. The takeoff mass from Pluto is correspondingly less, only 2.2 metric tons, compared to 12.4 metric tons for the outbound flight. The return flight time is constrained by the need to keep the entry velocity for aerocapture at Earth within acceptable limits rather than by AV and takeoff mass factors when the spacecraft departs from Pluto.

Table 2 gives the spacecraft mass budget for both the outbound portion of the orbiter/lander/sample return mission, and the Earth return portion.

Figure 7 shows the distance that the spacecraft can hop from site to site on Pluto, as a function of the burnout velocity, b o as the spacecraft takes off for its hop across Pluto. At a production rate of 7.3 kg per day, for example, 100 kg of liquid H2 propellant would be accumulated after a 2-week stay at the site. This would permit a burnout velocity of 670 meterskec, sufficient to hop for a distance of 1200 kilometers. For a small size planet like Pluto, a maximum hop distaance of 1000 kilometers appears satisfactory, with an average hop distance of -200 kilometers. An autonomous guidance control system would be required for the hop maneuvers.

Returning to the development of the MITEE-B engine, there exists a strong technology base for cermet nuclear fuel. In particular, in the 710 reactor program [2], the tungsten/U02 cermet has demonstrated that it is rugged and has the capability to satisfactory withstand very high temperature, e.g., up to 3000 K hydrogen for many hours, and to survive many thermal cycles - in some tests, at temperature rates of change of 10,000 K per second - without damage. In the DOD/SNTP (&ace Nuclear - Thermal Propulsion) program [3], thermal hydraulic tests of radial flow annular packed bed fuel elements have demonstrated power densities of 30 megawatts per liter in

the fuel region, using blowdown tests with & propellant (1000 psi) through high temperature (2500 K ) non- nuclear heated elements. The MITEE-B engine has substantially lower power densities in its fuel elements, i.e., I O MW per liter instead of 30 MW/liter, and should also perform satisfactorily.

Low power nuclear critical reactors on the PBR reactor were operated as part of the DOD/SNTP program. They confirmed with a exceptional degree of accuracy the ability of 3 dimensional Monte Carlo codes to predict Kef6

power distribution, and various nuclear performance factors, includins moderator ’and temperature coefficients, etc. These same computer codes have been used to analyze and predict the neutronic performance of the MITEE-B reactor.

Based on this technology base, it appears possible to develop the MITEE-B engine in 9 years, given a well funded development prograni. With an aggressive, goal oriented program, the development period could probably be shortened to 7 years, making the MITEE-B engine available for a Pluto mission starting in 2010 AD, if started now.

4. THE MITEE NUCLEAR PROPULSION ENGINE

To be able to carry out a practical Pluto OrbiterlLanderiSampIe Return Mission, its nuclear propulsion engine should provide the following performance goals: 1.

2.

3.

4.

5.

6.

7.

8.

High thrust, i.e., >10,000 Newtons, when it is operating in the NTP (Nuclear Therma 1 Propulsion) mode. High specific impulse, i.e., 2900 seconds, in the NTP mode. Be capable of multiple restarts after long shutdown periods. After departing Earth, it will have to bum for orbit insertion at Pluto, then for landing on Pluto, then for multiple hops across Pluto’s surface, and finally, lift off for return to Earth. On-board electric power supply, i.e., -1 KW(e) for controls, instruments, etc. for the 12 year trip to Pluto and the 12 year return to Earth, using a mini turbo- generator. Electric power, i.e., -20 KW(e) on Pluto for manufacture of fresh liquid H2 propellant. Be compact, e.g., 5 1 meter diameter, and lightweight, e.g., 5 500 kilograms. Be reliable and utilize existing technology where possible. Be able to observe the expected condensation of Pluto’s atmosphere in -2020 AD.

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Table 2 Spacecraft Mass Breakdown for Baseline Pluto Orbiter/Lander/Sample Return Mission

Spucecraji Baseline Design Pluto Landing MITEE-B Engine Shield: MP Unit: Electrolyzer/Liquefier: MP Deployment Mechanism: Landing Gear: Telecommunications: Ion Thrusters (Navigation): Sample Return Tank: Sample Return Tank Aeroshield: Propellant Tank Insulation: Miscellaneous & Contingency:

Spacecraft Payload Mass:

MITEE-B Bi-Modal Engine Systeitt :

Spacecruj? Total Dry Mass:

Earth Departure and Pluto Landing AV Propellant: Spacecraft Propellant Tank Mass:

Spacecraft Total Mass (lMLE0):

Spacecraj? Baseline Design Earth Return MITEE-B Engine Shield: Telecommunications: Ion Thrusters (Navigation): Sample Return Tank & Samples: Sample Return Tank Aeroshield: Propellant Tank Insulation: Miscellaneous & Contingency:

Spacecruft Puyload Muss:

MITEE-B Bi-Modal Engine System:

Spacecrap Total Dty Mass:

Pluto Departure AV Propellant Loading: Spacecraft Propellant Tank Mass:

300 kg

90 kg 65 kg 30 kg 40 kg 80 kg 45 kg 25 kg 20 kg 25 kg 50 kg 30 kg 500 kg

300 kg

800 kg

I 1050 kg (includes cooldown propellant) 550 kg

12400 kg

50 kg 45 kg 20 kg 30 kg 25 kg 20 kg 10 kg 200 kg

500 kg

1570 kg (includes cooldown propellant) 80 kg

Spacecraj? Total Mass (Pluto Surface): 21 50 kg

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HOD Ranae (Kilometers) Burnout Velocitv Elevation Anqle

(rnlsec) 15dea 30dea 45dea 50 deo 400 233 380 408 332 500 4 52 679 689 537 600 a02 1130 1069 796 700 1611 1865 1588 1121 800 3526 2955 2248 I 508

Range as Function of Vbo 8 Elevation Angle

4000

3500 - 3000 -45deg

5 2500

& 2000 5 1500

1000 500

0 : I : , : I : I : I I

300 400 500 600 700 800 900

Vbo (mlsec)

Figure 7 Spacecraft Hop Range on Pluto as a Function of Burnout Velocity

The above are very challenging requirements. In particular, the goal of reaching Pluto by 2020 AD would be difficult unless a strong, committed development program were to be undertaken very soon.

What is the status of nuclear propulsion engines, and what is the best route that could meet the above perfomiance goals ‘?

NTP systems have undergone extensive development and ground testing in the US and the Former Soviet Union (FSU). However, no NTP systems have yet operated in space. 3 1 low output power thermoelectric nuclear reactors have been operated in space by the FSU, as well as two thermionic Topaz reactors. The US has tested one thermoelectric reactor, SNAP- 1 OA in Space in 1965. The US later carried out development work on the 100 kW(e) SP- 100 space nuclear reactor, but no ground or flight tests were carried out. To date, for space missions requiring a nuclear power source, the US has relied on low power [< 1 KW(e)] RTG’s (Radisotope Thermoelectric Generator).

Most of the US and FSU development work on NTP systems carried out during the period of the 1950’s through the 1970’s focused on NERVA type reactors. These used neutronically inefficient graphite moderator, making the resultant nuclear propulsion engine excessively large, heavy, and with a low thrust to weight ratio. The US carried out successful ground tests of the NERVA system, but no flight tests. US work on NERVA stopped in the early 1970’s, due to the lack of a defined mission. Development on NERVA type engines

continued in the FSU until the 1980’s, and successful tests of fuel element assemblies were carried hut, though no full-up engine systems were tested.

In the late 1980’s in the DoD/SNTP program, the US undertook the development of a very compact and lightweight NTP engine. The SNTP engine was based on the Particle Bed Reactor (PBR) design, in which lithium7 hydride was used as the moderator, enabling a much smaller and lighter reactor [ 3 , 41. The PBR fuel elements consist of annular packed beds of small (-400 micron diameter) HTGR (High ’Emperature Graphite Reactor) type nuclear fuel particles, held in place between two porous frits. Hydrogen propellant flows radially inwards through the outer cold frit into the annular packed bed of fuel particles, where it is heated to 3000 K, and then out through the inner hot frit to a central axial flow channel, along which it travels to the exit nozzle.

The SNTP program developed and successfully tested the various components for the SNTP engine. While a full-up engine was not built and tested, low power critical assemblies were operated. Thermal hydraulic tests of prototype fuel elements demonstrated the capability to operate at very high power densities, of 30 megawatts per liter. The PBWSNTP engine design had a power of 1000 megawatts and a thrust-to-weight ratio of 3011, comparable to that of high performance chemical propulsion engines.

The SNTP program stopped in the early 90’s at the close of the Cold War. More recently, design studies of a smaller, lighter, high performance NTP engine, based in

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part on the PRR technology for robotic planetary science missions have been carried out. This engine concept, termed MITEE Una tu re ReacTor &ginE], adopts the basic radial flow geometry of the PBR file1 elements [5, 6, 71. However, it has two-hndaniental improvements from the PRR. First, instead of having the fuel elements arranged as a core assembly inside a common pressure vessel, each fuel element is positioned inside its own individual pressure tube, which contains besides the fuel element, an annular outer shell of lithiurn7 hydride moderdtor and a small nozzle at the exit end of the pressure tube. This arrangeniciit reduces the weight, and simplifies the construction of the engine. Moreover, it simplifies the nuclear test program and greatly reduces its time and cost. Instead of testing a full up assembly of 37 or 61 elements, validation tests can be carried out on a single pressure tubeifuel element assembly.

The second fundmiental improvement is the use of a multi-layer assembly of perforated tungsten-U02 cermet fuel sheets instead of a packed bed of nuclear fuel particles. The tungsten-UO2 fuel performs excellently in 3000 K hydrogen for many hours. Moreover, the local voidage and propellant flow geometry in the sheet assembly can be controlled more precisely than in a bed of randomly packed particles, reducing the effects of hot channel factors and variation in voidage. Also, the possibility of mechanical distortion due to shifts in

particle position caused by the thennal cycling effects during multi-burn operation, is eliminated.

The tungsten-U02 cermet consists of a tungsten matrix that incorporates up to 50% by volume of micron size U02 particles Similar cemiet fuels have operated successfully in nuclear reactors for many years with no problems. Tests of the tungsten-U02 type cermet fuel were carried out in hot hydrogen at temperatures up t o 3000 K in the 710 nuclear engine program [ 2 ] , and demonstrated the capability to operate for many hours without significant fuel loss.

Test pieces of the tungsten-U02 fuel were subjected to dozens of thermal cycles and temperature rise rdte of 10,000 K per second, without significant problems or degradation. To reduce overall reactor weight, molybdenumUO2 cermet fuel is used in the lower temperature zone (below 2000 K) of the fuel element in place of the higher density tungsten-U02 fuel, which is retained in the high temperature zone.

Figure 8 shows a cross section of a bi-modal MITEE-B reactor core assembly with 61 pressure tube/fuel elements. Designs having fewer, e.g., 37, pressure tubeifuel elements are also possible. The core assembly is surrounded by a ring of pressure tubes containing only lithium7 hydride, which acts as a neutron reflector for the core.

Figure 8 Cross Section Through MITEE-B Reactor

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F ptwa

FIGURE 9 Cross Section of MITEEB Pressure Tube/Fuel Element

Figure 9 shows a detailed cross section of an individual pressure tube. To achieve bi-modal capability, 8 small diameter beryllium tubes are bonded to the cold frit of the fuel element. These 8 tubes carry a closed cycle coolant, e.g., helium, when the reactor is operating in the electric power mode. Heat from the tungsten-U02 fuel sheets is transferred to the closed cycle coolant through the cold frit and its beryllium tubes, both by thermal radiation and conduction. This thermal energy is then carried to an external electric power generation system by the flowing coolant inside the 8 tubes. During the electric generation mode, there is no €J propellant flow through the fuel region. The nozzle at the end of the pressure tube is open to vacuum, so all of the region inside the pressure tube is also under vacuum conditions, except for that helium cycle portion inside the closed 8 coolant tubes. A description of the electric generation system is given later.

An extensive set of Monte Carlo neutronic analysis of MITEE reactors has been carried out using the MCNP computer code [7]. These analyses utilized full 3 dimensional geometric representation of the reactor geometry, and pointwise neutron cross sections, and provide the most accurate method available for analyzing small, high leakage, highly heterogeneous reactors. MCNP analyzes were able to very accurately predict the experimental criticality constant, Keff, of small PBR experimental reactors to within 0.5 percent, and could accurately model the other reactor parameters, including moderator coefficient, power distribution, etc.

The criticality of MITEE reactors is strongly affected by the pitch/diameter (P/D) ratio of its fuel elements. As the distance between fuel elements increases, P/D also increases the diameter and height of the reactor, and the volume of its lithium7 moderator increases. The value of Keff initially increases with increasing P/D value, because neutron leakage decreases. However, as the P/D ot io continues to increase, Keff decreases because of the increasing absorption of neutrons in the hydrogen atoms of the moderator. However, the reactor mass, which is composed of its U-235 fuel, the tungsten and molybdenum matrices, lithium7 hydride moderator and reflector, and beryllium structure, monotonically increases with the P/D ratio.

Figure 10 shows the effect of P/D ratio on the hl element core, plus the effect of having the beryllium closed cycle coolant tubes on the cold frits. Assuming that all of the 61 pressure tubeshe1 elements are functioning with a P/D ratio of 2.5/1 and there are 8 beryllium tubes on the cold frit, the value of Keff is 1.09, which is quite adequate. The criticality condition for reactor operation is a value of Keff = 1.00. With a P/D ratio of 2/1, Keff drops to 1.04, which is not adequate to accommodate for the range of control margin normally used for operating reactors. The effect of having the 8 beryllium tubes on the cold frit is modest, about a 0.05 drop in Keff, but still significant. Accordingly, the P/D ratio of 2.5/1 appears optimum, and necessary, for MITEE-B.

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Figure 10 MITEEB Criticality as a Function of Position of Non-Operational Pressure Tube Fuel Element

One of the unique attributes of the MITEE pressure tuhe/fuel element concept is its ability to continue operation even if one of the pressure tubeifuel elements were to fail. Propellant flow to the failed element would be simply cut-off by a shut-off valve, and the remainder of the reactor assembly would continue to operate. This capability is not possible for reactors that are situated inside a common pressure vessel.

As shown in Figure 10, shutting down one of the 61 pressure tube/fuel elements in the MITEE-B core has only a minor effect on the overall Keff, in the 0.02 to 0.03 range. The reactor could continue to operate with 1 of its elements shutdown, and probably 2. The impact on Keff depends on the position of the fueled element in the core - elements in the outer region (rings 4 or 5 ) have less impact on Keff than elements at the center or in the rd or 3rd row.

The use of the pressure tube configuration increases system reliability, and reduces the need to establish very high levels of reliability during development and testing. Being able to continue to operate with a failed element, a situation not possible with a single pressure vessel, significantly decreases the required level of reliability for each fuel element.

The criticality of MITEE reactors is strongly affected by the nature of the moderator used, and also the type of fissile fuel employed. F ip re 11 shows the Keff and reactor mass of MITEE for two different moderators,

7 . LiH and BeH2, and 3 different fissile fuels, U-235, U-233, and Am242m. Using BeH2 moderator instead of LiH reduces reactor mass from 100 kg down to 70 kg, at

a Keff of 1.07. This occurs because the atomic density of hydrogen in BeH2 is substantially greater than in 7LiH, making neutron leakage less for a given dianieter reactor.

7 .

Using U-233 fissile fuel further reduces reactor mass from 70 kg down to 40 kg, because of the higher net number of neutrons released per absorption in U-233 nuclei, as compared to absorptions in U-235 nuclei. Using Ain242m fuel reduces reactor mass still further down to 25 kg - a result of an even greater number of net neutrons per absorption and a greater absorption cross section.

This present study assumes 7LiH moderator and U235 fissile fuel; however, it should be realized that the MITEE-B reactor mass could be reduced substantially by use of a more efficient moderator and/or fissile f'uel, which would significantly reduce spacecraft dry mass and mission IMLEO. Table 3 summarizes the reactor design parameters used for the MITEE-B neutronic analyses. A 61 element core was chosen, in order to maximize its heat transfer capability when it operates in the closed cycle electric power generation mode.

The molybdenurnU02 and tungsten-U02 cermet fuel sheets are perforated with a large number of small diameter holes, through which the radially inflowing hydrogen propellant passes as it goes from a given cermet sheet to the next one at a smaller radius.

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Figure 11 Criticality Constant (Keff) vs MTIEE Reactor Mass for Different Fissile Fuels and Core Configuration

The diameter of the perforation holes is 0.018 centimeters, with a total hole area of 0.25 cm2 per cm' of sheet surface area. A conservative heat transfer analysis approach was used, in which convective heat transfer was assumed to only occur on the cylindrical inner surface of the perforating holes. This assumption neglects the additional convective heat transfer that occurs on the surfaces of the cermet sheets outside the holes. Since this external heat transfer area is several times greater than that inside the holes, the actual film drop will be substantially smaller than the calculated value.

Using a heat transfer correlation given by Rohsenhow [8], the radial temperature distribution of the fuel sheets and the hydrogen propellant across the fuel region is shown in Figure 12. The local temperature difference between the fuel sheet and the bulk temperature of the propellant at that point is the film drop required to transfer the heat from the fuel to the propellant. The film drop is relatively large at the cold side of the fuel region and relatively small at the hot side, because the thermal conductivity of hydrogen monotonically decreases with temperature. As discussed earlier, the actual film drop will be substantially smaller because of the additional heat transfer area.

During the electric power generation mode, the reactor operates at low thermal power. The heat generated in the fuel sheets transfers by a combination of radiative and conductive transport between the sheets (the fuel sheets are in close mechanical contact) in the fuel region to the beryllium cold frit, where it then transfers to the closed coolant circuit through the walls of the 8 tubes bonded to the frit.

The total surface area for heat trapfer into the cold frits of the 61-element core is 1.74 nf. Assuming a thermal cycle efficiency of 2 9 4 , operating at 1 KW(e) MITEE-B has a thermal power of 4kW(th). This correyonds to an average input heat flux of 0.23 watts per cm'. Operating at 20 KW(e), the average input flux is 4.6 watts per cm'. Because the radial thickness of the fuel region is small (0.6 cm) and the fuel sheets are in intimate mechanical contact, the temperature difference between the hottest fuel sheet and the cold frit will be small, on the order of 10 K under this condition.

Mission electric power requirements are 1 KW(e) continuous electric on-board power during the12 year trip to Pluto and the 12 year return to Earth, plus 20 KW(e) of continuous electric power for propellant manufacture while the spacecraft is on the surface of Pluto.

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Table 3 MITEEB Design Parameters (NTP Mode)

Fuel Element Descriptions

= 0.3873 cm (hot frit ID) * I = 0.6971 cni (ID ofW/UO: cermet section) 0 : = 1.0069 cm (ID of Mo/UO* cermet section) 0 3 = 2.3 cm (ID of pressure tube) 04 = 2.5 cm (OD ofpressure tube) Fuel element pitch P = 5.0345 cin Core OD/Height = 9 x pitch 45.4 cm Reactor OD = 17 x pitch 55.4 cm Upper axial reflector and grid plate structure thickness. 5 cm (Be) No lower retlector included Multiplication factor ( k , ) = 1.09 8 Be tubes attached to outside of fuel region for Brayton cycle Tube OD = 0.195 cm TubeID- 0.117 cm WIUO;! density = 8.85776 gmkc Mo/UO? density = 6.1468 gm/cc Moderator/reflector material density = 0.87541 5 gm/cc (LiH/Be)

Reactor Mass Description (Masses in kg)

Fuel Mass W/UO2 = 26.0

Moderator Mass = 69.44 (includes LiH/Be, Be tubes, and Be spacers between elements) Radial reflector mass = 27.41 (includes LiH/Be tubes and spacers between elements) Upper reflector = 12.81 Inlet plenum = 4.15 Reactor total = 168.1

Fuel Mass M d U 0 2 = 28.25

Engine Estimated Masses (kg)

75 MW

Thrust Reactor Control TPA Nozzle PMS TVC Plumbing Total

1456 kg, 14.268 NI 3200 lbc 168.1 10 5 25 6.5 7.5 10 232.1

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b

0 0,3873 1 .m9

Radius Icm’l

SUSEE Reactor SteamNVater Superheated Superheated

Mini Steam Electric Power Turbine 8

Fuel t Zone Tl, Pl Generator Element In Fuel To, Po Element

Zone In Steam b 4 Steam

Figure 12 Fuel and Propellant Distribution Through MITEEB Fuel Element

SteamNVater Turbine Exhaust

Figure 13 SUSEE Power Cycle

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- 5

n a - 4

L (U P

- 3 2

2 - 2 8

z

d 0

i- Q .-

- 1 E

The optimum choice for the system to supply this power appears to be the conventional steam cycle. The technology is very well established. Figure 13 shows the flowsheet for a space power system that utilizes the conventional steam cycle. During the in-space portions of the mission, reject heat from the steam condenser could be radiated to space. When on the surf'ace of Pluto, the reject heat would be dissipated in a sinal1 melt pool located in the ice sheet adjacent to the point where the spacecraft had landed.

The condenser pressure for a space steam cycle is considerably higher than in standard steam cycles on Earth, e.g., several atmospheres compared to -0.1 atmosphere. The higher condenser pressure allows a higher radiator temperature (i.e., -400 K), which in turn reduces radiator area and mass per KW(e) of generated power. Even with high condenser pressure, however, the thermal cycle efficiency (electric output divided by thermal input) is very good, i.e., in the range of 20 to 25%.

Figure 14 shows the thermal cycle efficiency and radiator area for a steam cycle as a function of condenser pressure, assuming standard steam cycle conditions, e.g., a turbine inlet temperature of 810 K (1000 F ) and a turbine inlet pressure of 68 atni (1000 psi). A conservative turbine efficiency of 80% is assumed, together with a radiator

efficiency of 0.9, and thermal radiation from only one side of the radiator. Also, a conservative mean temperature difference of 17 K is assumed between the turbine outlet and the radiator. The optimuni operating point appears to be a condenser pressure of 2atmospheres. It enables a thermal cycle efficiency of 23 pycent, which is attractive, and a one-sided radiator of 3 nf per kW(,,. With a two-sided radiator, the specific area per KW(e) is reduced t o 1.5 n?/KW(e).

Figure 15 shows a possible radiator configuration for the space steam cycle. The radiator consists of a series of flat metal strips that have internal grooved flow channels that carry the condensing steani/water mixture. The two ends of the flat strips are connected to inlet and outlet headers. The inlet header carries essentially pure steam, with a small fraction of liquid water droplets, while the outlet header carries a fully condensed liquid water stream.

The flat strips are connected along their length by simple flat fins ofmetal. Heat is conducted fkom the flat strips to the fins, where it radiates to space. Heat also radiates to space from the surfaces of the flat strips. 'lhe fins and strips are fabricated from a high thermal conductivity, low density metal such as aluminum or beryllium. Beryllium is probably a better choice than aluminum, since it has a lower density ( 1.8 g/cm3 vs 2.7 fdr aluminum), and higher thermal conductivity (2.2 wicm K vs 1.7 for aluminum).

Conditions: - Turbine inlet Temperature 810 K (1000 F)

Turbine inlet Pressure = 68 Atm (1000 psi) Turbine Efficiency = 80?& One Sided Radiator Radiator Emissivrty = 0.9

-

Figure 14 Radiator Area and Cycle Efficiency for Space Steam Cycle as a Function of Condenser Pressure

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Vol. 1-484

Cross Section of Space Radiator Plan View of Space Radiator

Wateristeam

From Ta s t q m Condenser nttbrne Purnr,

Figure 15 Space Radiator Design for Steam Cycle for MITEEB Electric Power Generator

The temperature difference hetween the roots of the fins and their midpoints is very small, even for very thin fins. At 1 kW(th) per square meter (one sided) for example, the

T for a 2centimeter wide, 0.025 Centimeter (10 ml) thick beryllium fin is only 1.8 K.

The nominal overall thickness of the flat strips is 0.1 centimeter (40 mils) with 0.025 centimeter (10 mil) deep grooved internal channels for the flowing steandwater mixture. As the steam condenses, because of surface tension effects, the mix will form into a linear sequence of individual water and steam slugs that move along the grooved channels. As the steamiwater mixture continues to condense, the distance between the water slugs will diminish, until a stream of pure water is discharged into the outlet header.

The flow velocity in the grooved channels is very low. For a radiator strip 2 meters in length, and 4centimeters in width (including a 2 centimeter wide fin), the water flow velocity out of the flat strip is only 0.5 centimeter per second, while the steam flow velocity into the strip is 4.5 meters per second. Pressure drop is small, less than a psi, and readily compensated by the suction pump that returns the condensed water back to the reactor. The radiator would be initially coiled into a rolled package, and then extended to form a flat panel after it was launched into space, using expandable truss structural elements. Assuming a two sided radiator panel, a radiator of only 1.2 x 1.2 meter in extent would be sufficient for

1 KW(e) generation. For the 20 KW(e) generation phase on Pluto, the waste heat would be dissipated into the melt pool via a compact heat exchanger that would be deployed from the spacecraft. Parameters and a mass breakdown for the electric portion ofthe bi-modal MITEE engine are given in Table 4.

5. REPLENISHMENT OF H2 PROPELLANT FROM IN-S ITU RESOURCES

Ice sheets on Pluto are ideally suited for the production of H2 propellant for the spacecraft because they provide: 1. A simple, easily utilized heat sink for the waste heat

from the MITEE-B electric power generation cycle. 2. A ready source of melt water for electrolytic

production of hydrogen. 3. A very low temperature heat sink for the efficient

liquefaction of hydrogen at low energy input.

Figure 16 shows the overall flow sheet for the hydrogen propellant production system on Pluto. The electric power output of the MITEE-€3 engine when it is sitting on the surface of Pluto is nominally 20 KW(e). Of this, 13 KW(e) is utilized in to electrolyze melt water to 6 and 9. The H2 gas is collected, liquefied, and stored in the spacecraft propellant tank, and the 02 gas is vented to space.

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Table 4 Electric Power System Parameters for MITEEB Engine

Power output in space Power output on Pluto surface Therinal cycle efficiency Turbine inlet teiiiperature Turbine inlet pressure Condenser pressure Radiator area in space Power system masses

Space radiator Mini-turbine Generator Plunxbing Heat exchangx on Pluto

Total bi-modal engine mdSS

(NTP mode plus electric mode)

1 KW(e) 20 KW(e) 23% 810 K (1000F) 68 Atm ( 1000 psi)

2 Atm 1.5 rd (2 sided)

6 kg 20 kg 15 kg 15 kg 15 kg

Total 71 kg 303 kg

m 1 20 kW(e) Water Electric Electrolpei ~1 LI~UCIICI

Power (80% Efficient) (20% Efficient) systm

5c

I L 2 kW(e) I Control

omrnunications FI Propellant

Figure 16 Overall Flow Sheet for & Production on Pluto

The electrolyzer uses conventional Solid Polymer Electrolyzers are generally built as multilayer stacks of Electrolyze (SPE) technology. The H2 production rate is cells instead of a single cell with a large electrode area. 7.3 kg of hydrogen per day. An 80% efficiency for the An 80:ell stack would then have an electrode area of electrolyzer appears conservative, and should be readily 0.01 nf per cell, with dimensions of 10 cm x 10 cm, for achievable based on electrolyzer experience on Earth. example. Optimizing the electrolyzer is outside the scope The required electrode area is 0.8 rn2, based on a of this study. However, it appears that it will be compact conservative current density of 1 O4A/m2. and light in weight.

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The low surface temperature of the Pluto ice sheet, on the order cf 70 K, greatly simplifies the liquefaction and refrigeration of the €$ propellant. For a refrigeration temperature TI, with an ideal Carnot engine carrying heat at TI and rejecting it at a higher sink temperature T2, the reversible work of refrigeration for liquid 6 at 20 K on is Earth 14 watts(e) per watt(th). On Pluto, bccause of the much lower surfice temperature, the reversible work of refrigeration is only 2.5 watts(e) per walt(th). Actual refrigerators are substantially less efficient than an ideal Carnot engine. For propellant production and storage a refrigerator on Pluto will require on the order of 5 times the ideal Cariiot work, or about 12.5 watts(e) per watt(th). Additional refrigeration power is required to cool the incoming H? gas from 70 K down to 20 K for storage as liquid in the spacecraft's propellant tank.

The refrigeration fiactor to cool the Hj gas will be less than 12.5 watts(e) per watt(th), since the average temperature of the fi gas as it is being cooled down is well above 20 K. However, to be conservative, the same refrigeration factor will be used. The electric power to refrigerate incoming H? gas is then, for a production rate of 7 kg per day, 770 watts. Adding in the 460 watts to liquefy the 6, the total refrigeration electric power is then 1230 watts. Adding in controls and refrigeration of the propellant tank, the total power for the refrigeration portion of the €$ propellant system is taken as 2000 watts(e). Accordingly, as illustrated in Figure 16, of the total 20 KW(e) provided by the MITEE-B engine, 13 KW(e) goes to the H2 gas production system by electrolysis ofmelt water and 2 KW(e) to liquefaction and refrigeration of the Hz propellant. The remaining 5 KW(e) is used for a variety of tasks, including communication with the melt probe/ASV unit, transmission to Earth, spacecraft and power system controls, etc.

As discussed earlier, the 6 electrolysis unit would use existing SPE technology. A simple compressoriexpander unit, either a turbine or piston device, would be used to liquefy the H? propellant. The H2 would be compressed to 180 psi (since the electrolyzer would be pressured, the Hz input from the electrolyzer would not require the full range of compression) cooled to 33 K in a recuperative heat exchanger by the returning H? gas from the expander, and then expanded by the turbine or piston portion of the compression/expansion cycle. At the end point (1 atm pressure) of the expansion phase, 35% of the 6 steam would be in liquid form. This would then Eturn to the intake of the compression device, where it would be recompressed to 180 psi. The returning cold H? gas would cool the compressed gas down to 3 3 K. A combined compressodexpander efficiency of 70?4 is assumed.

The waste heat rejection unit for operation on Pluto is similar to the space radiator shown in Figure 15, except that the waste heat would be transferred to a pool of melt water formed in the Pluto ice sheet rather than radiated to space. Steam from the turbine exhaust would be distributed by an input header to a set of flat aluminum strips inside of which were multiple small channels through which the steam would flow and condensed to liquid water as heat was transferred to the surrounding pool of melt water from the ice sheet. The liquid water exiting the outlet ends of the flat strips would flow into and be collected by an outlet header, which would return i t to the power generation unit where it would be pumped up to the appropriate pressure, turned back to steam, arid sent to the turbine inlet.

The wastc heat rejection unit would be configured as a flexible coil of the flat beryllium strips and headers. While in space it would be coiled up in a compact roll. After landing on the Pluto ice sheet, the coil would be lowered from the spacecraft and unrolled to form a flat panel on the surface of the ice sheet. A variety of possible mechanisms could be used to unroll the waste heat unit, including a memory-shape metal to tension the panel (the shape is controlled by the temperature of the memory metal), a pressurized pneumatic structure to form the backbone of the panel, a mechanical, extendable structure, etc. When the exploration of the particular site on the ice sheet was complete, the waste heat rejection unit would be rolled back into its normal coil shape, lifted off the ice sheet surface, and reattached to the spacecraft. The spacecraft would then hop to a new site on Pluto and repeat the process of deploying the waste heat unit. The total surface area required to reject the -80 KW(th) waste heat load from the 20 KW(e) power generation cycle is re1at;vely small. The panel having a projected area 1 rt? ( 2 nf counting both sides) that rejected heat at a flux of 4 watts per square Centimeter, for example, could handle 80 KW(th) of reject waste heat. Since the temperature difference between the condensing steam and melt pool water would be large, on the order of 100" C, the 4 watts per square centimeter is quite conservative. Considerably higher heat fluxes are normally found in conventional heat exchangers. The time required to create the lx lx 0.5 m melt pool is approximately 0.5 hour.

6 . TECHNOLOGY DEVELOPMENT FOR THE PLUTO MISSION

The six principal sub-systems for the Pluto orbiter/lander/sample return mission are: 1. MITEE bi-modal nuclear engine 2. 3. Melt probe 4. Control and communications 5. Scientific instrumentation package

Electric power system and Hz electrolyzer/liquefier

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I

SyStemS ~

Test 1 and 20 KW(e) Power

6. Spacecraft and Tankage

Flight Test MITEE-B Engine

in Orbit

Assembleand Test Full-Up

MITEE-B Engine

Development of the MITEE -B engine subsystem will probably take the most effort and longest time. Its development is discussed below. The power system utilizes conventional steam cyc!e technology, and its heat rejection equipment (a -1 m’ condensive radiator for 1 KW(e) on-board power during the flight to Pluto, and a heat exchanger for the heat rejected to the melt pools in Pluto’s ice sheets, can be developed and validated in a couple of years.

The 1% elcctrolyze/liquefier also would use established technology, i.e., a SPE (Solid Polymer Electrolyzer) unit, with a small turboicompressor to liquefy the Hz. It could be developed and fully tested in 3 to 4 years.

The nuclear reactor technology for the melt probe is very well established in an ongoing, large commercial industry. The cermet fuel has been successfully used for many years in hundreds of water-cooled operating reactors World wide, and is capable of very high burnups. The particular reactor geometry proposed for the melt probe is not commercially used, but could be tested and ready for implementation well before 20 10 AD.

A m r e detailed study is needed to determine what kind of control and communication system would be used for the spacecraft that would carry out the Pluto mission. Much of the required hardware and software probably could he adapted/modified from that already being used or planned for other missions. Similarly, much of the scientific instrumentation could be adapted/modified from the New Horizons and other missions. The New Horizons mission [9] envisions a power supply of only 0.22 KW(e) using an RTG. The proposed orbiter/lander/sample return mission would have an in-space power capability of 1 KW(e), and a surface power supply of 5 KW(e) for spacecraft control, communications, and instrumentations.

Figure 17 shows an illustrative road map for the development of the MITEE-B hi-modal engine, assuming a 9-year program. In Phase I , the various components - electric thruster, radiator, electric power generator, mink turbine, etc. - would be separately developed and tested. Full size single pressure tubelfuel elements would be also tested, first in an electrically heated version and then in an existing nuclear reactor such as TREAT. The cermet nuclear fuel would also undergo in reactor testing, as well as tests in high temperature hydrogen.

, Phase 1 : Component Phase 2: Integration and I Phase3: Engine development and validation ~ shakedown testing demonstration and application

Figure 17 MITEEB Development Road Map

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Phase 1 would take approximately 4 years. It would also include tests of low power, room temperature critical assemblies of the MITEE-B design, in order to verify the neutronic analyses. These would not involve H2 flow or power generation.

In Phase 2, the various components would be integrated into a hll-up MITEE-B engine that would be tested in both operating modes, i.e., the NTP and electric generation mode. In contrast to an NTP only MITEE engine, which could conceivably not have to be ground tested, the MITEE-B engine would rcquire ground testing, to ensure that the reactor and electric power generationhhruster system could operate reliably for years. Accordingly, a minimum length of timc for Phase 2 would appear to be 3years Following the ground testing, the program would move into Phase 3 , which would flight test MITEE-B in orbit. A ayear period appears necessary for Phase 3.

Developing the bi-modal MITEE-B would then take on the order of 9 years before it could be applied for operational missions. As discussed above, with an aggressive development program this could probably be shortened to 7 years, by carrying out certain development tasks in parallel rather than in sequence. Such a program would permit application in 201 0 AD.

7. CONCLUSIONS

Present propulsion systems will only yield a brief glimpse of Pluto as the spacecraft flies past the planet at high speed. To really explore this mysterious planet, it will be necessary not only to orbit it, but to land on its surface at multiple locations, and perhaps, even to return samples of Pluto back to Earth.

Such a mission is completely impossible using chemical propulsion. It is possible with nuclear thernial propulsion. A bi-modal nuclear engine that generates high propulsive thrust for insertion into orbit and landinghakeoff from the surface of Pluto would enable a spacecraft to map Pluto from orbit, and then land on its surface. If, as seems likely, there are surface deposits of water ice, electric power gnerated by the bi-modal nuclear engine could electrolyze melt water to replenish liquid 6 propellant for the spacecraft, enabling it to hop to many widely separated sites on Pluto, before taking off for Earth with collected samples. In addition, the spacecraft would deploy a small melt probe, powered by a compact ultra lightweight nuclear reactor that could melt a channel deep into Pluto’s ice sheets, to obtain data on their composition and structure.

There is a strong technology base for the components needed for such a mission. The cermet nuclear fuel has been successfully tested for many hours on high temperature (22750 K ) hydrogen, while the required neutronic and thermal hydraulic performance has been achieved in progranis on similar nuclear thermal propulsion reactors. The proposed steam cycle power system uses technology that has been in comniercial use for many years.

With a vigorous development program, thc Pluto mission could be ready for launch by 2010, to arrive at Pluto shortly after 2020 AD, when the expected atmospheric condensation would probably still be in progress.

REFERENCES

[ l ] G. Woodcock, et al., “Benefits of Nuclear Electric Propulsion for Outer Planet Exploration”, 38”’ Joint Propulsion Conference, Indianapolis, Indiana,

[2] Argonne National Laboratory, “Nuclear Rocket Program Technical Report”, ANL 236, 1966.

[ 3 ] H. Ludwig, et al., “Design of Particle Bed Reactors for the Space Nuclear Thermal Propulsion”, Brookhaven National Laboratory, BNL 52408, 1993.

[4] H. Ludwig, et al., “Design of Particle Bed Reactor for the Space Thermal Propulsion Program”, Prog. In Nuc. Eng., 3, No. I , 1996.

[ 5 ] J. Powell, G. Maise, and J. Paniagua, “MITEE: An Ultra-Lightweight Nuclear Engine for New and Unique Planetary Science and Exploration Missions”, Paper IAF-98-r.l.O1, 4gth Int. Astro. Cong., Sept. 28, - Oct. 2, 1998, Melbourne, Australia.

[6] J. Powell, G. Maise, and J. Paniagua, “Phase 1 - Final Report. Lightweight High Specific Impulse (1 000 sec) Space Propulsion System”, NASA Contract 8-9903 (STTR Contract), Plus Ultra Technologies and Stony Brook University, Oct. 1999.

[7] J. Powell, G. Maise, J. Paniagua, and S. Borowski, “Compact MITEE-B: Bi-Modal Nuclear Engine for Unique New Planetary Science Missions”, Paper AIAA 2002-3652, AIAA Joint Propulsion Conference, Indianapolis, Indiana, July 10-1 3 , 2002.

[ X I W.M. Rosenhow, et al., Handbook of Heat Transfer Fundamentals, McGraw Hill, New York, 1985.

[9] Y. Guo and R. Farguhar, “New Horizons Pluto- Kuiper Belt Mission: Design and Simulation of the Plut o-Charon Encournter”, Paper IAC-0 2-Q. 2 .0 7, 53d International Astronautical Congress, Oct. 10- 19, 2002, Houston, Texas.

July 7-10,2002.

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Jaines R. Powell - S6.D. iir Nrrcleus Engrrieesiirg, MIT. Fsoitr 19% to 1996, DI: Powell wosked ut Biaokhuveii Nutioirul Lubosutoiv (BNL) 011 U vusietv of feclriiologies

oii uird firsiori seucloi:\,

reirewuble erreigv s wteitrs, sir~~ei~contlirctiirg upplicutioirJ, uirtl envrsoiinreiitul arid sufetv strrtlies Dirsiirg tlre period 1 9S7-93, he MWS the Muirugei.,fos tlre BNL portioii of the DOlj/&)u~e Nrrcleus Tlreiwrul Pi*oprilsioii Psogsuiii oir tlre Pustide Betl Ntrcleur Rocket (PBR) He is czisseirtly woskiirg us U coirsrrltuirt oil

irriinvutive teclrirologies fos iirii.u.~ri.Licttii.e, disposul of iiiicleui. waste, arid space e.yploiutioir. 111 udditiori to his iiiventiorr ol the Yai?i~le Betl Nircleui. Rocket, Ire is the co- iirverrtw of Maglev (niupireticullv levitated tsuirspst). uird is disecfiiig the cleveloptient of U Muglev Progi*ui~r betweeir Orlunrlo uirrl Poi? C'uiiuverul, wliich ~1ould be rlirfissI Muglev roufe 111 rlie U.S.

George Muise - Ph. D. in dei*ospuce uird Mechuiricul Scieirces, Psineeton U Mujor useus oj eyertise use'jluiil nrechunirs uird Ireut fsuirTfes, jjusticirlui.lv us relutecl to the ais-bseutliiiig uiicl rocket ~~r*opirlsron, rui-qfiecl gas clviiaiiiirs. re-eiitiv lreutiirg. nrrcleui* seuctos th erit I U U'1ij)di-u ii lies, air ti elect so.\ tu tic pso bes , fos plusit I u diugirostics. Di.. Muue worked jas se~~esul uesospuce coitipurrie.\ psior to joirriiig BNL iir 1974. Aitroirp his ussigiiiireirts ut BNL MU .five veuia qf R&D 011 the theinrul/~i~~c~ruzrlic uspeels o/ the Pusticle Bed Reurtos irtrcleui* socket He left BNL iii I997 urrtl is curseiitlv ownes of the coirsirltiiig jiriir G. Muise Aei*orruirtiruI Assocrutes.

Joliti C. Patiiagua - Ph. D. in Mecliuiricul Engiiieesing, Stute Uiiiversit-y of New Yosk (SUNY) ut StonJ9 Bsook. Dr. Puniugiru 's areus qf exyes use roit~ututiaiiul.fliri~l iiviiurii ics, heu t tsuirs fes, it1 irlti-phase flow systems, orbitul itieclruirics, spaceci-uji vehicle riesigii arid iiiissioii u nu lvs is. His doctosu 1 dim ertu t ioii e m nr ined Sin rpl@etl Boiliirg Wuter Reucfor (SB WR) two-phuse flow

es clui*ing stustiip llwirsieirts. Di: Yuiiiuguu. prior to joining the teuclring stulf ut SUNY ut Stoiiv Brook, worked for terr veuss ut Gi~ritritrun Aesospucr Coiporutron. Aiiroirg his ass igrriirent.\ wese tlri-ee yeair\ of R&D 011 the Pustide Bed Reucfos irzicleus socket.

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