The initial Nuclear Thermal Propulsion workshop was intended to identify a range of reactor concepts, select Concept Focal Points to specify essential concept elements, conduct a mission analysis, identify critical initial tests, and provide schedule, milestone, cost and facility requirements estimates.
These reactor concepts were evaluated according to their ability to meet baseline Mars mission requirements, which included use of the 2016 launch window for a less than 600 day piloted trip time with a 30 day crew stay on Mars. The Mars vehicle would be assembled in LEO, with a 334 kN engine requiring three perigee burns for trans-Mars injection with a baseline 124 ton out-bound payload and a 40 ton return payload. In the course of the evaluation process, a range of variations on this baseline emerged.(1)
Additional mission options identified at the Workshop included an early Mars mission in 2006, which would require early propulsion downselect, and a late Mars mission in 2026, which would permit a longer technology development effort.(2) The choice of these options results from a balance of various figures of merit, and may result in a sequence of missions in which the first Mars mission could be slower and less risky, with later Mars missions having shorter trip times, building on earlier experience and infrastructure. A number of potential figures of merit were discussed at the Workshop:
PARAMETER BASELINE VARIATION
Engine availability 2015 2004 - 2017
Thrust per engine kN 334 110 - 1110
Specific impulse m/s 9,065 9,065 - 11,760
Engine thrust/weight 6 6 - 10
Number of engines 1 multiple ( 7 ? )
Reactor power MWt 1,500 500 - 5,000
Electric power - Low kWe 0 1 - 5 dual mode
Electric power - High kWe 0 25 - 50 dual mode
Reactor operating time minutes 250 / mission 250 - 1,000 / mission
Reactor cycles per mission 6 1 - 30
Number of missions 1 1 - 5
Crew Radiation Limit REM/year 5 0 - 5
FIGURE OF MERIT IMPACTS
IMLEO Launch requirements
(Initial Mass to Trip Time
Low Earth Orbit) Cost
Trip Time IMLEO
Crew Considerations (radiation, etc)
Free Return trajectory
Safety/Reliability Complexity
Free Return trajectory
Flight Qualification
Operational & Complexity (eg Mars Rendezvous)
Mission Flexibility IMLEO
Reusability
Technology Cost
Development Complexity
Evolvability
For the most part, system analyses traded on IMLEO and trip time, because of the relative ease of quantifying and calculating these figures of merit. But there was a frequently expressed view at the Feedback Workshop that a more sophisticated approach was needed, and that other figures of merit might prove more informative. It was also recognized that much more work was needed on making the case for which figures of merit were most relevant for mission analysis, since little work had been done on comparative analysis of figures of merit per se.
A total of 17 propulsion concepts were identified, in four general classes: solid core; liquid core; gas core; and other.(3) The panel recommended that high priority near term technology reactor efforts focus on higher temperature fuels, fuel qualification, test facilities, control mechanisms, systems lifetime and coatings and coatings design criteria. Recommendations for near term propulsion efforts included high temperature Hydrogen and neutron compatible materials, nozzle cooling, radiation hardened feed systems, Hydrogen kinetics, and system reliability, redundancy and operability. Each of the reactor classes was thought to offer a 10-15% performance growth potential through technology advances.
Solid core concepts encompass the widest range of systems, with ten configurations identified (nine at the initial workshop, and a tenth, the tungsten-water-moderated, identified subsequent to the initial workshop). These fall into three categories, including homogeneous thermal reactors in which the fuel is mixed with a carbon or beryllium oxide moderator, heterogeneous thermal reactors with separate fuel and moderator, and fast reactors with refractory materials such as tungsten for structures. The size of these reactors is dependent on the efficiency of heat transfer, and performance of these systems is a function of the maximum fuel temperature, which is limited to between 2300 K and 3500 K. Major design considerations include selection of materials to avoid hot corrosion with multiple reactor restarts, and fuel element fabrication to retain fission products.
Two liquid core reactors were identified, with operating temperatures in the range of 3000 K to 9000 K. Major technical uncertainties include the effectiveness of heat transfer from the reactor to the propellant, as well as the risk of significant fuel droplet loss through the nozzle. A third class of reactors includes three gaseous core reactor concepts, which, with core temperatures of 9000 K to 15,000 K or higher offer, the highest performance levels. But major questions remain concerning plasma containment for these concepts. A final category included architectures which called for use of in-situ propellant production, or a dual-mode nuclear-thermal / nuclear-electric system.
1 - NERVA-Derivative Reactor
Westinghouse is the leading proponent of the NERVA Derivative Reactor (NDR), which has been evaluated for SDI(4) and SEI electric propulsion(5) and nuclear thermal rocket applications.
Principal advantages of the NERVA-Derived system are:
+ An established technology base which has been demonstrated over a wide range of thrust capabilities that requires no technical breakthroughs and has modest development requirements and risks.
+ Identified solutions to safety and environment concerns. Loss of coolant accidents result in reactor shut-down, since the hydrogen propellant makes a significant contribution to core reactivity.
+ Proliferation resistant nuclear fuel, since highly dispersed low inventory discourages diversion, and recovery from TRISO-type beads is a difficult process.(6)
Major technical issues include:
- Demonstration of the effectiveness and lifetime of the ZrC "Super Coat" fuel element coating, which received limited testing in the Nuclear Furnace program.
- Safety issues include concerns about fission product exhaust during ground testing, reassurance that intact reentry will not lead to reactor critical or super-critical geometry if the full core is immersed in water, permanent shutdown capabilities using safety rods and control drums, and fuel integrity, as well as public perceptions of these issues.
Near-term Phase I efforts in years 2 through 5 of the program will include:
* Fuel element tests to demonstrate ZrC coating lifetime and effectiveness;
* Demonstration of complete fuel element fabrication;
* Evaluation of reactor safety issues, and demonstration of reactor shutdown.
Mid-term Phase II research in years 4 through 6 of the program include:
* System control and prototype flight system control tests;
* Shielding tests;
* Safety tests;
Far-term Phase III activities beginning in year 8 of the program include:
* Cold flow engine experiments;
* Hot tests demonstrating operating envelopes;
* Demonstration of start-up, shut-down, cool-down and emergency responses;
* Verification of endurance and cycle capabilities.
Parameters for NERVA, the current NERVA-Derived Reactor (NDR), as well as more advanced NERVA-derived systems include:(7)
Parameter Unit NRX XE' NERVA NDR Binary Ternary
1972 Carbide Carbide
Year 1970 1972 1990
Fuel Bead ZrC coated UC-ZrC-C ZrC-UC UC-ZrC-NbC
Moderator Graphite Zr Hx/C Zr Hx/C Zr Hx/C Zr Hx/C
Reactor Vessel Aluminum Steel Steel Steel Steel
Temperature K 2,270 2,500 2,700 3,100 3,300
Expansion ratio 100:1 500:1 500:1 500:1 500:1
Isp secs 710 890 925 1020 1080
Thrust kN 245 334 334 334 334
Power MWt 1,120 1,520 1,613 1,787 1,877
Propellant flow kg/sec 33
Thrust/Weight (no shield) 3.9 4.2 4.0 3.7
Thrust/Weight (w/ shield) 2.4 2.3 2.2
Options for integrated test facilities include exhausting hydrogen into a cleanup-scrubber system, or into an underground tunnel, or testing in space. Existing containment facilities are argued to be adequate, with modification of the hydrogen cleanup system.
Total costs for the program are estimated at $755 million, including:
Reactor design & development $ 350 million
Engine design & development $ 150 million
Full test procurement & assembly $ 100 million
Facility preparation $ 125 million
Tests costs $ 30 million
Rockwell was responsible for the non-nuclear engine components of the NERVA program in the 1960s. Rejoining the NERVA industrial team of the 1960s, Westinghouse and Rockwell recently signed a memorandum of understanding concerning their joint work on NERVA-Derived Reactors for the Space Exploration Initiative -- as with the initial program, Westinghouse will concentrate on nuclear systems, with Rockwell focusing on rocket engineering.(8)
Rockwell mission analysis studies have focused on NERVA applications to both the Moon and Mars missions.(9) For the reference 2015 mission, which envisions a total trip time of 565 days, four alternatives were considered: Chemical + Aerobrake; Nuclear Thermal with engine-out capability; Nuclear thermal with propulsive capture: and Nuclear Thermal + Aerobrake. The resulting IMLEO requirements clearly favor nuclear thermal:
SYSTEM IMLEO (tons)
Chemical + Aerobrake 720
Nuclear Thermal with engine-out 680
Nuclear Thermal with propulsive capture 580
Nuclear Thermal + Aerobrake. 340
A range of mission benefits were identified, including:
+ Significant trip-time reductions, and broadening of launch and stay-time windows, including the potential to conduct single window missions.
+ Higher specific impulse significantly reduces IMLEO.
+ Shorter trip-times reduce mass of expendable and shielding, further reducing IMLEO.
Future studies include consideration of hot engines in Mars orbit, as well as safeing of reactors returned to Earth orbit.
Rockwell has also given consideration to improvements in rocket engine technology applicable to new NERVA-Derived systems.(10) One of the primary innovations is the use of an extendible rocket nozzle skirt, to significantly increase the nozzle expansion ratio. The basic hydrogen-cooled section of the nozzle might extend to an expansion ratio of from 65 if carbon-carbon is used for the extendible skirt, or to 150 if a refractory metal is used. Significant fabrication problems could limit the use of carbon-carbon, which would require a skirt nearly 13 meters long and over 7.5 meters in diameter. In addition, the work of Rockwell's Rocketdyne Division on the Space Shuttle Main Engine in the 1970s significantly advanced the state of the art:
Parameter Unit 1960's NERVA 1990 SSME
Turbopump flow kg/sec 28 72
Turbopump pressure MPa 11.0 48.3
Nozzle pressure MPa 4.3 21.7
Nozzle temperature K 2,500 3,116
NASA Lewis has been a leading actor in recovering the engineering heritage of the NERVA reactor program. Applications studied include both Lunar Base and Mars Expedition missions, with emphasis on modular systems that will support evolution from the Lunar to Mars applications.
Consideration has also been given to NERVA testing requirements, leading to recommendations for construction of a
"... Nuclear Test Facility (NTF) where the integrated system level tests will be conducted. Candidate DOE sites include the Nuclear Rocket Development Station (NRDS) at Jackass Flats, Nevada, or the Idaho National Engineering Laboratory (INEL)...Tests to be conducted include Cold Flow Tests, Startup Transients, Ramps to Intermediate Hold Points, Full Power Operation, Shutdown, and Cooldown. Engine Exhaust is contained and processed within an effluent treatment system which directs hydrogen away from the engine system, removes fission products and disposes of the hydrogen in a safe manner...
"A containment option for consideration is to exhaust the engine into a large underground tunnel. Such tunnels are routinely constructed at the Nevada Test Site for containment of nuclear weapons test (several tunnels already exist within a mile or two from NRDS). Tunnels can be evacuated and used to collect the engine effluent. (Benefits include) flexible effluent scrubbing time (cleanup of exhaust gases can proceed at slower rates (lower mass flows) than the engine exhaust mass flow rate); no environmental contamination in the event of operational accident; and test approval not function of weather conditions."(11)
NASA Lewis envisions a three-phase development effort for the NERVA-Derived Reactor:
Phase Activity Period Cost
1 System Development & Planning 1991-1995 $ 350 M
2 Electric & Nuclear Furnace tests 1996-2000 $1,450 M
3 Reactor & Engine System tests 2001-2006 $1,200 M
TOTAL 15 years $3,000 M
According to Stanley Borowski of NASA Lewis, much of the engineering heritage of the NERVA program could be recaptured within 8 years, leading to an NTR flight test as soon as 2002:
"The punch line is that nuclear propulsion is an enabling technology for the moon and Mars. Missions can be seen that demand nuclear electric as well as nuclear thermal, but budgets will obviously be the dictating force."(12)
2 - Particle Bed Reactors
In response to SEI propulsion requirements, two design philosophies have been suggested,(13) one which maximizes the reactor's thrust-to-weight ratio and requires a high power density, pressure and temperature, as well as small size and high thrust, and another which maximizes specific impulse requires low power density and pressure, with very high temperatures, large size and low thrust.
High High Point
Thrust/Weight Specific Impulse Design
Power MW 1000 - 5000 500 - 2000
Power Density MW/L 20 - 80 5
Chamber Temperature K 2500 - 3500 3000 - 3750 3200
Chamber Pressure MPA 7 - 14 0.5 6.9
Specific Impulse Sec 850 - 1060 1000 - 1300 971
Thrust kN 20 - 1,000 60 - 200 333
Engine Mass kg 650 - 5500 2800 - 6000 1,700
Thrust/Weight (W/O shield) 20.0 - 35.0 4.0 - 7.5
Shield Mass kg 1300 - 6400 3700 - 7900
Thrust/Weight (W/ shield) 8.6 - 14.0 2.0 - 3.2
Maximum Fuel Temperature K 2500 - 3650 3200 - 3900
To resolve the range of development issues posed by the Particle Bed Reactor, total developmental cost for this concept was estimated at about $1.5 billion from 1990 to 2012. This included $320 million through 2000 for an Element Test Reactor, $500 million from 1995 to 2008 for a Ground Test Engine (which would require an altitude chamber for start-up testing), and $600 million from 2007 through 2012 for space qualification testing.
3 - Pellet Bed Reactor
Science Applications International (SAIC) is a leading proponent of the Pellet-Bed Reactor, which has been evaluated for SDI(14) and SEI applications. The University of New Mexico
Institute for Space Nuclear Power Studies (with the support of McDonnell-Douglas) is the Concept Focal Point for Pellet Bed Reactors for SEI nuclear thermal propulsion.(15)
A baseline nuclear thermal reactor concept for SEI propulsion includes the following parameters:
Parameter Value Baseline
Rated Power MWt 1,000
Core Diameter m 0.8
Core Height m 1.3
Core Power Density kW/cm3 3.0
Maximum fuel temperature K 3,100
Maximum Propellant Exit Temperature K 3,000
Propellant Flow Rate kg/sec 32
Reactor Specific Mass (W/O shield) kg/kWt 1.0
Resolution of these issues is estimated to require over $2 billion by 2006, including;
Phase Time (years) Cost ($M)
Conceptual Design & Technology Issue Resolution 3 - 5 $ 100
Preliminary Design & Component Development 5 - 7 $ 800
System Ground Demonstration 1 - 2 $1,000
Flight Qualification 1 $1,200
TOTAL 10 - 16 $3,100
4 - PLUTO-Derivative Reactor
The PLUTO-derivative reactor, prevoiusly discussed in the chapter on SDI power systems, has also been considered for SEI propulsion applications. The applicability of the PLUTO reactor to nuclear thermal propulsion is limited by a UO2 BeO fuel phase transition near 1900 K and a low-temperature eutectic slightly above 2000 K. These effectively eliminate the PLUTO reactor from consideration for operations at these outlet temperatures.
5 - PLUTO/NERVA Derivative Reactor
The PLUTO/NERVA derivative reactor, prevoiusly discussed in the chapter on SDI power systems, has also been considered for SEI propulsion applications.
6 - Cermet Reactor
General Electric (GE) is the leading proponent for the Cermet Reactor, which has been evaluated for SDI(16) and SEI applications.(17) A range of advantages of the cermet reactor have been suggested:
+ Cermet reactors offer improved thermal conductivity compared to metal oxide fuel elements.
+ An extensive fuel test engineering heritage exists from the ANP and 710 programs.
+ A principle advantage of the Cermet reactor, which was the basis for interest in the technology for aircraft propulsion, is the high retention of fission products in the fuel matrix. This was experimentally demonstrated in the 710 Program, in which most test fuel elements demonstrated fission gas fraction release of less than 10-9, while some fuel elements released fission fragment fractions in the range of 10-5 to 10-4. This should produce less stringent containment and confinement restrictions on Ground Test Facilities.
+ It has also been suggested that the cermet fuel may offer improved swelling behavior, but this remains uncertain.
Several issues remain open:
- Due to their lower fuel density relative to metal fuel elements, a potential disadvantage of cermet fuels is large core size, and thus greater core and shielding mass.
- In order to improve weldability, Rhenium is a potential Cermet cladding material.
Relevant system parameters include:
Parameter Unit 710 SEI/ANL Baseline
Study period years 1962-68 1988+
Power MWt 2,000 2,000
Thrust kN 133 - 890 445
Specific Impulse Seconds 850 - 870 832
Propellant flow rate kg/sec 55
8 - Advanced Dumbo
Los Alamos is the Concept Focal Point for the Advanced Dumbo Reactor, based on the Dumbo Reactor concept originally developed in the late 1950s.(18) The Dumbo configuration consists of a number of vertical columns, composed of washer-shaped fuel elements separated by spiders to facilitate propellant flow. Hydrogen propellant is injected from the exterior of the column, and flows axially down the hollow center of the column. Work on Dumbo was cancelled in September 1959, based on the conclusion that the design offered no performance advantages relative to the graphite core KIWI reactor, and the Dumbo involved a more complex design which required a significant number of precision parts.
The Advanced Dumbo configuration incorporates several innovations. One of the major changes from the original 1959 concept is replacing the Cermet fuel with UC-ZrC fuel elements. In addition, several variations on the precise fuel configuration are under consideration.
Advantages of this configuration include:
+ An alternative, radially self-supporting, geometry for UC-ZrC fuel which offers reduced thermal stress, and potentially eased fuel fabrication.
+ Improved efficiency due to increased fuel surface area (intermediate between Rover and Particle Bed reactors) with greater fuel flow area (much greater than Rover, and greater than Particle Bed).
Drawbacks include engineering complexity, fuel fabrication and thermal stress, and axial structural support.
9 - Foil Reactor
Sandia is the Concept Focal Point for the Fission Fragment Assisted Foil Reactor,(19) and the Direct Heating Reactor. In the Foil Reactor, a thin 2 micron Uranium layer is supported on a Be or Al2O3 substrate, and escaping fission fragments directly heat the hydrogen propellant. Concentric fuel plates are separated by alternating layers of coolant channels for substrate cooling and propellant exhaust.
The concept was first developed by Bussard and DeLauer in 1958, and was used in nuclear pumped laser tests in the FALCON program in the 1980s. The FALCON tests verified the relationship between UO2 foil thickness and fission fragment energy, demonstrated metallic and ceramic substrates, and resulted in gas heating significantly in excess of the 1,500 K substrate temperature. About 100 of fuel modules would be clustered into a reactor with the following characteristics:
Parameter Value Baseline
Reactor Diameter m 3.75
Reactor Height m 4.00
Reflector Thickness m 0.70
Module Diameter m 0.358
Fuel Mass kg 18 - 30
Total Mass (no shield) kg 42,000
Reactor Power MWt 13,300
Foil Temperature K 2,700
Exhaust Temperature K 3,400
Thrust kN 2,669
Specific Impulse Seconds 990
Design excursions range over the following values
Temperature of Temperature of Specific Thrust
Substrate (K) Propellant (K) Impulse (Sec) (kN)
2,000 2,700 836 3,050
2,300 3,100 898 2,815
2,500 3,370 937 2,687
2,700 3,630 975 2,571
3,000 4,040 1030 2,424
The principle advantages of the Foil reactor concept are:
+ High thrust and power levels with attractive specific impulse.
+ Structural elements are much cooler than the propellant, and elevated temperatures are confined to small areas. This permits a significant range of choice of moderator (D2O, Be, D2 liquid or gas, CD4)
+ Safety is enhanced through redundancy of fuel modules, as well as the low power density which leads to graceful failure modes and reduces the rate at which a core accident progresses.
+ Fission fragments discharged in propellant flow and short burn times (less than 30 minutes) reduce residual radiation hazards.
Unresolved issues include:
- Impact of large dilute configuration on criticality and structural design. This results in large size and significant reflector cooling mass penalty.
- Materials issues, including fuel integrity in the face of hydrogen erosion and thermal gradients, reflector cooling, coating techniques, and frit and porous ceramic fabrication.
- Discharge of fission fragment in propellant flow could significantly complicate ground testing (although this could be mitigated by overcoating all but one of the modules in the ground test to prevent fission fragment escape, and using an open loop scrubber for the one uncoated module).
A program to resolve these issue would include:
Category Activity Time Cost ($M)
Physics H2 Excitation Reaction 2 years $ 5
Scoping Studies Criticality, Cooling, Structures 2 years $ 5
Technology Development Substrates, Coatings, Integrity 5 years $ 60
Component Testing Channel Tests (FALCON etc) 4 years $ 40
System Integration Tests Site Preparation 5 years $ 500
Engineering & Fabrication 15 years $1,500
Facility Operation 5 years $ 300
10 - Tungsten Water Moderated Reactor
NASA Lewis is the Concept Focal Point for the Tungsten Water Moderated Reactor, although few details of this concept have been published to date.
11 - Foam-Fuel Reactor
The Foam Fuel reactor, prevoiusly discussed in the chapter on SDI power systems, has also been considered for SEI propulsion applications.
12 - Wire Core Reactor
The Wire Core reactor, prevoiusly discussed in the chapter on SDI power systems, has also been considered for SEI propulsion applications. The concept proposed for SEI applications weighs 5,080 kg, is 1.85 m long and 1.08 m in diameter, with a thrust of 910 kN based on a reactor thermal power of 4,400 Mw.(20) The system provides a relatively high 930 second specific impulse.
13 - Low Pressure Nuclear Thermal Rocket
INEL is the leading proponent of the Low Pressure Nuclear Thermal Rocket concept, which was originated by Carl Leyse.(21) Pressure-fed nuclear reactors were considered in the early 1960s, but the conventional graphite core reactors of that period operated at such high core pressures as to render the concept impractical. But in the late 1960s a preliminary study by Leyse of a low-pressure ( 345 kPa ) NERVA-derived reactor identified a number of operational advantages, including improved performance and reliability.
More recently, studies at INEL have focused on using more advanced high-temperature ZrC/UC particle or pebble fuel elements, and carbide (such as TaC/ZrC/UC or TaC/ZrC) reactor materials. These materials are anticipated to permit total operating endurance of over ten hours at 3,000 K, and over two hours at 3,200 K, in contrast to the 2,300 K to 2,500 K operating endurance range for conventional Graphite matrix fuels. In initial inward-radial-flow configurations, the reactor pressure vessel is at the cold end of the propellant flow, placing the core in compression, permitting higher chamber temperatures and thus specific impulse. More recent designs have focused on outward-radial-flow configurations.
The Low Pressure reactor takes advantage of the disassociation of molecular H2 to monatomic hydrogen, with significant increases in specific impulse being achieved with chamber pressures below 1.0 MPa (10 atmospheres), and chamber temperatures above 3,000 K. The Low Pressure Reactor relies on conventional propellant tanks pressure, in the range of 0.2 to 0.3 MPa (2-3 atmospheres) to expel the LH2 propellant into the reactor without the need for turbopump machinery.
Some typical reactor characteristics include:
Thrust kN 22.2 111.2 111.2 48.9 48.9
Chamber Pressure kPa 69 69 69 103 103
Chamber Temp K 3500 3200 3600 3200 3600
Specific Impulse Sec 1325 1075 1250 1050 1210
Thrust/Weight Ratio 0.5-1.0 6 6 6 6
Vessel Diameter m 1.4 1.5 1.5
Nozzle Area ratio 300 60 60 40 40
Reactor Power Mwt 500 500 260 260
Fuel loading U-235 Kg 70 80 50 70
Proponents claim a number of advantages for the Low Pressure Reactor:
+ the LP reactor, currently at Technology Readiness Level 2, has a development status intermediate between that of the NERVA and more exotic concepts such as the gas core;
+ the LP reactor has a theoretical specific impulse up to 50% higher than NERVA, ranging from 1050 to 1350 seconds;
+ the lower operating pressure can reduce reactor thermal problems, with improved core heat transfer and lower nozzle heat flux;
+ the pressure-fed reactor eliminates the control drums, engine gimbal and complicated turbopump machinery of other solid-core reactor concepts, as well as reducing the number of valves and reactor parts;
+ this greatly simplified propellant feed system of the LP reactor could provide a significantly more reliable propulsion system than either the NERVA or gas core reactors, since the turbopump accounted for about half of the failure budget of NERVA;
+ the simpler propellant feed system also could prove much easier to qualify for manned missions, given the absence of complex redundancies typical of NERVA systems;
+ the much lower thrust levels typical of LP reactors (one to two orders of magnitude lower than typical of NERVA systems) could result in substantially smaller, and thus less expensive, ground test facilities;
+ clusters of engines (such as seven) would permit two-engine-out abort modes;
However, several issues remain unresolved:
- the extent to which theoretical molecular hydrogen disassociation and recombination specific impulse gains can contribute to actual engine performance requires experimental verification, as well as the linearity of the relationship between chamber pressure and specific impulse - some analyses suggest major gains in specific impulse can be achieved at 100 kPa chamber pressures, while others suggest that relatively modest improvements might be achieved even with significantly lower chamber pressures;
- the high temperature HfC and TaC materials needed to realize the higher range of specific impulses predicted for this concept are beyond the current state of the art, and pose significant development challenges;
- the ground test facility will require a long (30 meter) low pressure blow-down chamber to simulate the space environment for reactor testing, with duct diameters 2 to 5 times as large as for conventional NERVA reactors, as well as very low pressure drop effluent handling systems with multi-stage ejectors, which may demonstrate very poor (5% -10%) efficiencies;
- the relatively low thrust of this concept would make it difficult to avoid using multiple periapsis trans-Mars injection burns, which poses radiation and operational safety hazards that can be avoided by higher thrust single-burn systems.
Assuming program initiation in 1991, INEL estimates that a LP reactor development program would include the following major milestones:
1991 - 1994 Environmental Impact Statement
1991 - 1996 Laboratory development of fuel elements
1996 - 2000 Nuclear Furnace testing of fuel elements
1994 - 1998 Test reactor studies
1996 - 2002 Engine design and component tests
1999 - 2004 Engine tests
2001 - 2003 Inert engine flight tests
2003 - 2006 Live engine flight tests
Total program costs in the initial four years of the program were estimated at nearly $200 million, peaking at $125 million in 1994, including $7 million for the EIS, $57 for a test reactor, and $78 million for Nuclear Furnace testing. Total program cost from 1991 through 2006 was estimated at slightly more than $4 billion, with typical annual budget requirements from 1996 to 2005 of $300 million, peaking at nearly $500 million in 2000. Nuclear furnace testing would require $565 million, twelve engines for ground tests $2,255 million, and three flight engines $150 million.
14 - Liquid Annulus Reactor
Brookhaven is also the Concept Focal Point for the Liquid Annulus Reactor System (LARS).(22) In this concept, hydrogen propellant passes through the center of the reactor core and is heated as it passes over a surface molten Uranium, which is maintained in a liquid annulus on a substrate of solid uranium supported by a rotating Be moderator, which is in turn supported by a stationary moderator. Notional reactor parameters include:
Parameter Value f= 0.4 f = 0.8
Total Power MWt 200 200
Outlet Temperature K 6,000 6,000
Outlet Pressure atm 10 10
Fuel Bed Radius cm 9.4 8.1
Rotating Drum Radius cm 12.4 11.1
Reactor Diameter cm 110.0 110.0
Reactor Height cm 192.7 135.5
Reactor mass kg 3,000 3000
Mass of UC2 kg 30.0 30.0
Several advantages have been asserted for LARS:
+ A high specific impulse, in the range of 1500 to 2000 seconds, since the operating temperature is not constrained by the melting point of the fuel, which permits additional impulse gain due to molecular hydrogen disassociation.
+ Enhanced reliability due to operation of moderator, control and structural components at propellant inlet temperature.
+ Fission products evaporate directly into the exhaust stream, reducing fission product inventory, which reduces post-operation radiation hazard and after heat coolant requirements.
However, a number of limitations have also been identified:
- Longer burn-times and multiple engines will be required due to low thrust levels.
- Significant development effort will be required due to immature technology base. Since this effort would be comparable to that required for higher performance gas-core reactors, this concept has not been recommended for further work.
Although little developmental work has been conducted on the LARS concept, a proof-of-principle experiment conducted in 1963 demonstrated rotational containment of liquid refractories by a cooled solid outer layer using an aluminum cylinder. Overall costs of a development program are estimated at $1.6 billion, requiring work through the year 2008. The stages of this program include:
Phase Activity Period (years) Cost ($M)
1 Design & Analysis 1990 - 2006 $ 30
2 Technology Development 1990 - 1994 $ 50
3 Element Test Reactor 1992 - 2000 $ 350
4 Engine Development & Ground Test 1994 - 2008 $ 570
5 Space Qualification 2007 - $ 600
TOTAL $1,600
15 - Droplet Core Reactor
University of Florida Innovative Nuclear Space Power Institute (INSPI) is also the Concept Focal Point for the Droplet Core Reactor.(23) This concept is the culmination of an evolutionary process that began with the Colloid Core concept initially identified in 1970. The Colloid Core used a vortex flow cavity to confine U-C-Zr particles in a compact core with a temperature of 3,700 K providing a specific impulse of 1100 seconds. Although the vortex properties of the core were experimentally simulated in 1972, this concept demonstrated very high Uranium particle loss rates. The Rotating Liquid Core reactor concept, identified in 1972, used a liquid UC-ZrC fuel contained by 7,000 RPM centrifugal force, with pressurized hydrogen bubbling through the fuel to produce a specific impulse of 1,500 seconds at a temperature of 4,800 K. Deficiencies of this concept included high fuel loss rates, as well as the absence of a reliable rotation mechanism for the liquid fuel.
The Droplet Core reactor concept, first identified in 1988, is applicable to both nuclear thermal propulsion and space electric power generation.(24) This reactor recirculates a liquid Uranium fuel maintained in a stable liquid phase, which at a pressure of 500 atmospheres runs from 1,400 K to 9,500 K. Uranium droplets vary in size from 2 to 20 microns, depending on hydrogen flow velocity, which can range from 200 to 2,500 m/sec. Tangential injection of hydrogen propellant maintains the vortex flow of the fuel and propellant mixture, protecting the reactor wall from excess heating. Entrainment of fuel droplets could minimize fuel loss to as little as 50 kg per mission. This configuration can provide thrust in the range from 11 kN to 1,800 kN, with a specific impulse ranging from 1,500 to 3,000 seconds.
Parameter Unit Baseline Variation
Thrust per engine kN 333 111 - 1,111
Reactor Power MWt 1,500 500 - 5,000
Fuel exit temperature K 6,100 5,000 - 6,600
Propellant exit temperature K 6,000 4,000 - 6,500
Thrust/Weight ratio (shielded) 1.6 1 - 5
Specific Impulse seconds 2,000 1,500 - 3,000
U2 recirculation rate kg/sec 70 70 - 1,000
A number of advantages are associated with this concept:
+ The excellent mixing of fuel and propellant provides extremely high surface area density (over 106 M2/M3), a thousand-fold improvement over solid-core concepts. This results in the deposition of about 50% of the fission energy into the propellant flow, which is heated to 3,000 K to 7,000 K,
+ Specific impulse is improved by the high level of non-equilibrium disassociation of hydrogen molecules by fission fragments, which can approach 20% at 6,000 K.
+ These high operating temperatures produce a specific impulse greater than 2,000 seconds (at 6,000 K), and high thrust-to-weight ratios (unshielded T/W ratio of 5 for a 333 kN rocket, 1.6 with shielding). This can reduce Mars mission trip-times to less than 200 days.
+ Engine thrust is limited by nozzle cooling capability, and can be throttled by varying the propellant flow rate (a reactor producing 111 kN at 7 kg/sec will produce 666 kN at 33 kg/sec).
+ Safety features include: in-orbit fueling to avoid launch accident criticality hazards; low residual radioactivity due to expulsion of all gaseous and some non-gaseous fission products in the propellant stream during operations; and low Uranium U-233 loading (20 kg in core and 100-150 kg total).
+ Mechanical simplicity improves reliability.
Several problems have been identified with this concept:
- The Liquid Droplet Core is highly dependent on the development of new materials with acceptable mechanical and thermal properties at ultra-high temperatures. These materials include single crystal tantalum and tungsten, low-porosity glossy carbon, W-Re-HfC alloys, and E-ThO2 alloys.
- Neutron activation of W and Ta structural materials will produce high levels of radiation.
- Significant uncertainties associated with the performance of the liquid fuel recirculation system.
- Liquid fuel reactors appear to offer inferior performance relative to gas-core concepts, while requiring similar development efforts.
16 - Gas Core Nuclear Light Bulb
United Technologies Research Laboratory/Center (UTRC) was the lead NASA industrial contractor on gas core reactors in a program that had a budget of $25 million and included about 400 work years of effort. This project, which included LANL as the lead DOE facility, ran from 1959 to 1978.
The NASA program included successful critical assembly of UF6.(25) More recently, UTC has become the Concept Focal Point for the Nuclear Light Bulb Gas Core Reactor configuration.(26)
The Nuclear Light Bulb marks a departure from other Gas Core reactor concepts, in which the Uranium vapor is in physical contact with the Hydrogen propellant. Instead, in the Light Bulb, the Uranium gas is confined in a cylinder of fused silica by a tangential vortex of buffer gas, which like the fused silica is transparent to radiation. The hydrogen propellant flows outside this cylinder, and is heated by injection of micron-size tungsten particles which absorb the thermal radiation from the hot fuel.
Principle advantages of this system are:
+ Significant reductions in triptime and IMLEO, with round trips to Mars typically requiring less than one year.
+ Fuel is contained, avoiding fuel loss of other Gas Core concepts.
+ In contrast to open core cycles, which analysis has suggested have significant stability control problems, Nuclear Light Bulb core stability can be maintained by varying fuel injection rates. Power oscillations either quickly damp, or oscillate with long periods within acceptable limits.
+ The reactor is designed for reuse, since the fuel is stored separately.
Outstanding technical issues include:
- Reactor and system stability over the full range of operating conditions.
- Failure modes and safety impacts, and overall system reliability.
- Fuel and buffer gas separation, and recirculation system performance.
- UF4 deposition on confinement wall interiors can reduce transparency, although this can be mitigated by addition of small amounts of fluorine to the buffer gas.
- A significant fraction (over 10%) of the gas core heat output is at ultraviolet wavelengths at which the fused silica confinement wall is opaque. Unacceptable wall thermal loading would result, unless the buffer gas is seeded with silicon.
- Environmental impacts of Earth Development Facility.
- The complexity of the system precludes early assessment of many safety and reliability issues.
Up to 1961 Air Force support of physics analysis
1961 Program transferred to NASA / AEC
1961 - 1973 Propulsion studies - Open Cycle NASA Lewis & UTC
- Nuclear Light Bulb UTC
- Cavity Reactor Experiments GE & INEL
1969 - 1973 Gas Reactors studied for power, breeder reactor and laser generation.
1970 - 1972 Cylindrical Cavity Test Reactor low power tests by Aerojet at INEL.
Spherical Test Reactor low power (0.5 kW) tests by NASA Lewis at INEL.
1973 Propulsion program terminated, NASA to study terrestrial power applications
1974 Nuclear pumped laser demonstrated (NASA JPL, NASA Langley, University of Florida, University of Illinois, Los Alamos and Sandia).
1975 - 1979 Los Alamos Critical Cavity Reactor Assembly with confined UF6 in seven tests achieved 20 kW for about 100 seconds.
UTRC Argon seeded with UF6 demonstrated reconversion and energy coupling in RF heated plasma confined in water-cooled fused silica walls, with equivalent Isp of 1,350 seconds.
1977 - 1979 Gas Core Reactor studied for Sustainer Breeder Reactor for Nonproliferation Alternative Systems Assessment Program (NASAP) at Los Alamos, UTRC and Science Applications.
1979 Program terminated (with over 160 reports issued)
Major design parameters proposed for the Nuclear Light Bulb concept include:
Parameter Unit Baseline Radiator Demonstrated
Reactor power MWt 4,600
Fuel Radiating Temperature K 8,300 6,000
Propellant Exit Temperature K 6,700 4,500
Engine mass Kg 32,000
Specific Impulse seconds 1,870 3,200
Buffer Gas Hydrogen Argon
Propellant flow kg/sec 22.2
Thrust kN 409
Thrust-to-Weight ratio 1.3
Use of a space radiator for management of moderator and exhaust nozzle heat could significantly increase the specific impulse.
A multi-phase development program is contemplated to qualify this engine concept:
Phase Activity Dates Cost ($M)
Low High
1 Next-step studies 1991 - 1995 $ 4 $ 6
2 Non-nuclear Component Development 1992 - 2003 $ 15 $ 25
3 Nuclear Component Development 1995 - 2005 $ 150 $ 250
4 Ground Engine Development 2000 - 2012 $1,500 $2,500
17 - Open Cycle Gas Core
Sverdrup is the Concept Focal Point for Open Cycle Gas Core Reactors,(27) based in large measure on the work of Robert Ragsdale, who was a leading proponent of these concepts at NASA Lewis prior to his arrival at Sverdrup.(28) Initial studies in the 1960's concluded that Open Cycle Gas Core Reactors with 8,500 MWt power and a thrust of 222 kN had the potential to conduct piloted Mars expeditions carrying payloads of 50 tons with round trip times of 60 days with vehicle IMLEO of 2,050 tons, and round trip missions of 80 days could be accommodated with vehicle IMLEO of 1,000 tons. Mars expeditions carrying payloads of 150 tons with round trip times of 150 days could be conducted with vehicle IMLEO of 1,200 tons, and round trip missions of 250 days could be accommodated with vehicle IMLEO of 600 tons. Open Cycle Gas Core Reactors consists of a central Uranium plasma, which is supported in the center of the reactor chamber by the flow of hydrogen propellant, which is directly heated.
Reactor parameters include:
Parameter Unit Baseline
Reactor Power MWt 7,500
Specific Impulse seconds 5,200
Thrust kN 222
Engine mass (with radiator) kg 114,000
Engine diameter m 4.25
The principal advantage of this concept is the potential for Mars missions with round trips taking 60 to 150 days.
Major technical issues include:
- Adequate containment with acceptable fuel loss rates and reactor pressure.
- Nozzle cooling will pose a major challenge, and satisfactory testing of a high specific impulse 5-10 MW nozzle prototype will be a critical near term technical milestone.
- Fuel loss rates will pose major challenges to ground testing with current environmental standards with acceptable cost and risk, although review of the 1972 Preliminary Engineering Report (PER) facility studies did not disclose insurmountable obstacles. An Engine Test Facility would include a scrubber channel 60 meters long and 15 meters in diameter, fed by 11.5 million liters of water, with additional filtration systems.
A system development schedule includes the following stages:
Phase Activity Dates Cost ($M)
1 Start-up Studies 1990 -1993 $ 2
2 Focused Technology (eg Nozzle) 1992 - 1998 $ 60
3 Ground Engine Tests 1997 -2009 $1,000
4 Space Engine Tests 2008 - 2014 $ 600?
18 - Vapor (Gas) Core
The University of Florida Innovative Nuclear Space Power Institute (INSPI) is the Concept Focal Point for the Vapor Core reactor.(29)
Two gas core reactor concepts have been developed, the Nuclear Vapor Core Rocket, and the Vortex-Confined Nuclear Rocket.
In the Nuclear Vapor Core Rocket, a UF4 vapor helium gas mixture is contained in hexagonal cooled carbon-carbon blocks. The axial flow of Hydrogen propellant is circulated in HfC-lined carbon-carbon tubes that penetrate these blocks.
In the Vortex Confined Nuclear Rocket, an inner stream of liquid Uranium at a temperature of 3,000 K is injected into the core, along with an outer stream of hydrogen propellant at a temperature of 1,000 K. The Uranium vapor is maintained in an annulus in the center of the core by tangential injection of hydrogen propellant from the walls of the core, with the central region of the Uranium vapor annulus dominated by hydrogen propellant, which exits the rear of the reactor at temperatures of 6,000 K to 8,000 K.
A rotary fluid separator collects the Uranium, which is subsequently cooled in a heat exchanger and recirculated for injection at the forward end of the reactor. The power levels of these engines scale directly with propellant mass flow rates. Reactor control is maintained by fuel density variation rather than with control rods or drums.
Parameter Unit Vapor Vortex
Engine Availability year 2015 2020
Thrust per engine kN 333 333
Number of engines 1 - 7 1
Reactor Power MWt 1250 1650
Dual Mode Electrical Power kWe 50 25 - 3000
Core Mass - Unshielded kg 15,000 14,000
Shield (0.75 M) mass kg 10,000 10,000
Thrust/Weight ratio 1 - 2 1 - 2
Specific Impulse seconds 1280 1810
Nozzle expansion ratio 50:1 50:1
In addition to those advantages of the Vapor Core reactor previously noted in the chapter on SDI power applications, attractive features specifically related to SEI include:
+ Very high fuel and propellant operating temperatures, limited by wall cooling capabilities rather than fuel temperatures, and thus high specific impulse. This translates reductions of more than 30% in Mars mission trip times, as well as lower IMLEO.
+ Efficient energy transfer from fuel to propellant by direct molecular collision, radiative heat, and direct fission fragment deposition.
+ Since the reactors are launched unfueled, there is no possibility of criticality due to launch accident.
As discussed in the chapter on SDI power applications, however, a number of very serious technical challenges are associated with the Vapor Core reactor.
Major test facilities to resolve these issues include:
1991-93 - Non-critical UF4 fueled mini-cavities tested at high temperatures without a neutron flux, and at lower temperatures with neutron flux. Tests could be conducted at the Fast Flux Test Reactor, as well as the UTREC (Ultrahigh Temperature Reactor and Energy Conversion), EBR II and HFIR (High Flux Irradiation Reactor) facilities. These tests will require a total of $8 million.
1993-95 - Low power (20-100 kwt) low temperature (500K), UF6 fueled flowing critical facility without coolant, at the LANL Plasma Core Cavity Reactor.
1994-98 - A new Facility to conduct mid-power (0.1-1.0 Mwt) mid-temperature (1000 K) criticality and reactivity feedback and power density tests, with Helium cooled UF4 core with cavity and internal moderator. The total cost of this and related efforts to determine engineering feasibility is estimated to be about $90 million.
1998-200? - A new Facility to conduct an integral demonstration experiment at high power (5 Mwt) high-temperature (2500 K) criticality and reactivity feedback and power density tests, with H2 cooled slip-stream recirculation UF4 core with internal moderation. This will require annual funding of about $40 million per year.
Work in the Nuclear Engineering Program at the University of Missouri has include investigation of the stability of Gas Core Reactors.(30) In response to a NASA Lewis workshop in the summer of 1988, a study was undertaken of the stability of Gas Core Reactors. In the past,
"the relatively long prompt neutron lifetime (6 milliseconds) of this reactor system, with its heavy water reflector, led to the belief that the system could always be controlled by fast acting control system actuators."
The subject of this analysis was a notional reactor with the following parameters:
Parameter Unit value
Reactor Power MWt 36,577
Engine Thrust kN 1,880
Cavity pressure atmospheres 300
Cavity internal radius cm 183
U-235 fuel ball radius cm 74
Hydrogen flow rate kg/sec 71.2
Specific Impulse seconds 2,697
Fuel edge temperature K 17,760
Fuel center temperature K 131,600
The study noted that:
"The fuel mass is the most difficult parameter to control. Small, but rather rapid, changes in fuel mass can be expected because of the turbulent mixing of the propellant and fuel. The flowing gas containment experiments had shown that volume flow rate ratios of the outer to inner gas in the range of 100 to 1 can be achieved, with a central fuel ball size 40% of the cavity volume. This gives an effective fuel mass replacement in approximately 250 seconds... Smaller fuel volumes gave even better containment rates."
The results of the study include the observation that:
"In the case of oscillation in the fuel mass, such as what one might expect from the turbulence causing mixing of the hydrogen and uranium, the response was found to quickly become exponentially unstable... It is apparent that the response is prompt and rapid, to even slight variations in propellant flow (which is easily controlled by the operator) or fuel mass (which the operator has virtually no control over, or even a method of observing). These effects are well into the "prompt critical" range of the system, in which the delayed neutrons have little influence. The response to the asymmetrical oscillation of fuel mass was somewhat surprising... The unstable positive effect on reactivity and power over two cycles indicates the inherent instability of the system. It is also interesting to note that, as the negative swing is made just slightly greater in magnitude, the power oscillations change from increasing instability to appear more stable, or unstable in the negative direction, i.e. toward shutting down the reactor. This leads one to the conclusion that with fast response control actuators, conditions of quasi-inherent stability should be attainable."
The study also noted that:
"These results are a qualitative evaluation of the system's stability. Additional details and refinements may reduce the magnitude of the inherent instability, bur one would not expect that such refinements would reverse the system stability (i.e. make it inherently stable)."
19 - Dual Mode
One of Sandia's major contributions to the nuclear propulsion field was a wide-ranging review of gas-cooled reactor concepts for SDI multi-megawatt burst-mode and steady state power production applications.(31) This review was focused on power production rather than nuclear thermal propulsion, and the study noted that the "conclusions reached in this study are limited to the applications discussed... (and) the rank order of the basic concepts may be radically different for other power ranges, operating times and missions." Nonetheless, many of the reactor concepts considered and the associated technical issues are common to both applications.
The growing interest in dual mode nuclear propulsion concepts (in which the reactor is initially operated for thermal propulsion for injection into and braking from trans-Mars and trans-Earth trajectories, then operates at lower power to produce electrical power following injection and prior to capture) increases the relevance of Sandia's work. However, the study did not directly address bi-modal reactors "because the merits of bi-modal reactors are not conclusive, and because many combinations of power level and operating times could be considered." The study also noted "the complexity associated with mode switching and the potential for fuel damage during mode switching..." for such reactors.
General Electric has also performed studies of hybrid or dual mode propulsion systems, which combine low-thrust high specific impulse system such as Nuclear Electric Propulsion (NEP) with higher-thrust less efficient systems, such as Nuclear Thermal.(32) Several combinations of Nuclear Thermal Propulsion (NTP) with other systems have been suggested:
NTP + Chemical Although this combination provides only limited performance improvement, it may be needed for short-burn orbital capture maneuvers.
NTP + NEP or High thrust NTP would be used for Earth escape to avoid the Van Allen radiation belts, and low-thrust NEP would be used for the rest of the mission.
NTP + NEP High thrust NTP would be used for Earth escape, and Mars capture and escape, with low-thrust NEP would be used the rest of the mission to reduce trip-time.
Battelle Pacific Northwest Lab is the Concept Focal Point for Dual-Mode NEP/NTP hybrid propulsion concepts, which have also been examined by General Electric.(33) The PNL concept uses a NERVA reactor for Trans-Mars and Trans-Earth injection, which is supplemented by an electric propulsion system. The propulsion system for a 124 ton payload has a mass of 16 tons, which results in a total IMLEO of 524 tons.
Parameter Unit NTP NEP
Type NERVA MPD
Propellant Hydrogen Argon
Reactor Power MWt 1,500 35
Reactor Power MWe 8
Specific Impulse seconds 850 5,000
Thrust kN 334 0.066
Operating Lifetime 10 hours 2 years
Advantages of hybrid systems include:
+ Power requirements for comparable trip-times can be significantly reduced relative to pure NEP systems.
+ For equal power levels trip-times can be significantly reduced.
+ Improved mission reliability due to redundant systems, including electric power production for house-keeping and surface exploration (through power beaming), as well as the availability of alternate propulsion means.
+ The availability of high-thrust at any point in the mission significantly increases mission resiliency.
+ High thrust during Earth escape avoid the Van Allen belt radiation hazard (although Solar flare shielding requirements may obviate this benefit).
The greatest drawback is the need to develop two different propulsion systems (NEP and NTP), which significantly increases development cost and risk.
Total development cost is estimated at $950 million from 1991 through 2003.
Sverdrup has considered the benefit of hybrid propulsion, under contract to NASA Lewis.(34)
"Minimizing trip time exacts a penalty in higher IMEO, conversely, a minimum IMEO results in longer trip time. These two conflicting figures of merit, trip time and IMEO, correspond to two types of propulsion that are able to minimize one or the other, but not both... An alternate solution to the paradox of minimizing trip time and IMEO is the concept of a "hybrid" propulsion systems, using both high thrust, low Isp and low thrust, high Isp systems to gain the advantages of both... The nuclear thermal (NTR) and NEP hybrid also have the option of separate propulsion and power reactors or a single reactor used to provide energy for both forms of propulsion. By utilizing the two forms of propulsion in the proper regimes, reasonable trip times can be achieved with less IMEO than either of the two systems separately."
This analysis also reported conclusions derived from an earlier study of hybrid systems, conducted in 1968. This prior study assumed a Mars Mission Module mass of 72 tons, and a Mars Excursion Module mass of 179 tons.
Some representative vehicle figures of merit include:
Year Trip IMEO
Time tons
Days
Chemical (Split Sprint) 2003 440 1754
NTR Nuclear Thermal 1980 400 2160
NTR + Nuclear Electric (NEP) 1980 400 1584
Chemical + NEP 1980 400 2520
Chemical + NEP 1980 500 1080
Rockwell has analyzed Nuclear Thermal / Nuclear Electric propulsion systems, of the type for which Pacific Northwest Lab is the Concept Focal Point.(35) A major attraction of such systems is their potential to reduce trip time, and thus crew exposure to radiation and weightlessness.
This analysis considered a system based on:
Parameter Value NTR NEP
Type Phoebus MPD
Power MW 4,000 MWt 1,320 MWe
Thrust kN 1,125 22
Specific Impulse seconds 1,000 5,000
Propellant flow rate kg/sec 113.4 0.454
This analysis noted that one-way trip times of less than 50 days can be implemented with hybrid propulsion systems for total IMLEO of less than 1,200 tons and payloads of 175 tons.
20 - NIMF - Nuclear Indigenous Mars Propellent
Martin Marietta is the Concept Focal Point for the Nuclear Indigenous Mars Propellent mission scheme.(36) NIMF proposes to drastically reduce IMLEO requirements by refueling a nuclear powered Mars Excursion Module on the surface of Mars with propellants derived from the Martian environment. Candidate propellants include Carbon Dioxide condensed from the atmosphere, Water melted from ice or permafrost, and methane synthesized from carbon dioxide using nuclear reactor heating. Used with a nuclear reactor, the specific impulse of these propellants range from 280 seconds for carbon dioxide, which would only be adequate for reaching Mars orbit, to 350 seconds for water, which would be adequate for direct ascent to Trans-Earth injection, to 560 seconds for methane, which would permit direct insertion into a high-energy Trans-Earth injection orbit (methane production would require catalytic reaction with additional hydrogen reactant from Earth).
Successful operation of this system would require reactor fuel materials or coatings capable of resisting corrosion by high temperature carbon dioxide (above 2,200 K). The BeO fuel elements developed in the PLUTO program may enable operations at temperatures up to 2,400 K. Illustrative mission architectures include a scheme under which a single Shuttle-C derived heavy lift booster would launch a single 47 ton automated propellant production plant to Mars. Once the 100 kWe power plant had produced adequate propellant, a subsequent mission would emplace the crew on the Martian surface, using the propellant from the first mission for the return flight.
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