45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit
2 - 5 August 2009, Denver, Colorado
AIAA 2009-5311
Bimodal Nuclear Thermal Rocket Propulsion Systems for
Human Exploration of Mars
Marc N. Wilson1 and Dr. Alan Wilhite.2
Georgia Institute of Technology/National Institute of Aerospace, Hampton, VA, 23666
D.R. Komar3
NASA Langley Research Center, Hampton, VA 23681
The purpose of this paper is to perform a preliminary comparative analysis of four
leading bimodal nuclear thermal propulsion system designs to determine which would result
in the lowest initial mass in low-Earth orbit and how it would affect Mars mission payload
capability and a launch vehicle’s lift requirement. The Commonwealth of Independent
States bimodal system selected in NASA’s 1998 Design Reference Mission is compared to
particle bed reactor, ceramic-metallic, and Nuclear Engine for Rocket Vehicle Applicationtype engines. System trades are performed on the four propulsion systems as they relate to
the piloted mission in NASA’s human Mars exploration architecture. The system masses are
calculated using an in-space vehicle model with an integrated bimodal nuclear thermal
rocket model. The vehicle subsystem masses resulting from each engine type were modeled
using both constant thrust and constant thrust-to-weight constraints, with 0%, 15%, and
30% dry mass contingency applied. In every case, and especially when constrained to a
thrust-to-weight of 0.3 for gravity loss reduction and a practical 30% contingency mass, the
19-element particle bed reactor is the propulsion system requiring the least vehicle gross
mass. Using this propulsion system, a trend is established between the useable payload
capability of a launch vehicle and the corresponding limit on the trans-Mars injection stage’s
payload mass.
Nomenclature
BNTP
=
BNTR
=
CBC
=
CERMET =
CIS
=
DRM
=
ERV
=
EXAMINE =
IMLEO =
Isp
=
=
LEO
MER
=
MLI
=
MMOD =
NDR
=
NERVA =
NTP
=
Bimodal Nuclear Thermal Propulsion
Bimodal Nuclear Thermal Rocket
Closed Brayton Cycle
Ceramic-metallic
Commonwealth of Independent States
Design Reference Mission
Earth Return Vehicle
Exploration Architecture Model for In-Space and Earth to Orbit
Initial Mass in Low Earth Orbit
Specific Impulse, s
Low-Earth Orbit
Mass Estimating Relationship
Mulitilayer Insulation
Micrometeoroid and Orbital Debris
NERVA-derived Reactor
Nuclear Engine for Rocket Vehicle Application
Nuclear Thermal Propulsion
1
Graduate Research Assistant, Georgia Institute of Technology / National Institute of Aerospace, 100 Exploration
Way, Member AIAA.
2
Langley Professor, Georgia Institute of Technology / National Institute of Aerospace, 100 Exploration Way, AIAA
Associate Fellow.
3
Senior Aerospace Engineer, Vehicle Analysis Branch, Mail Stop E401, Member AIAA.
1
American Institute of Aeronautics and Astronautics
Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
NTR
OMS
PBR
PMAD
RCS
SEI
T/W
TCS
TMI
TPS
=
=
=
=
=
=
=
=
=
=
Nuclear Thermal Rocket
Orbital Maneuvering System
Particle Bed Reactor
Power Management and Distribution
Reaction Control System
Space Exploration Initiative
Thrust to Weight Ratio
Thermal Control System
Trans-Mars Injection
Thermal Protection System
I. Introduction
O
N January 14, 2004, President George W. Bush announced a new Vision for Space Exploration Program in
which he advocated for continuing human and robotic exploration of our solar system, beginning with humans
returning to the Moon and ultimately enabling human missions to Mars and beyond. In order to make the journey to
Mars viable, new propulsion technologies would need to replace the current chemical systems. Work on advance
concepts has stalled for years, but the studies that had been done in the past provided a firm foundation for continued
investigation.
In 1991 the Space Exploration Initiative (SEI) Synthesis Group, a collection of experts from NASA and the
Departments of Defense, Transportation, and Energy, released the report “America at the Threshold: America’s
Space Exploration Initiative”.1,10 The report outlined a variety of architectures and technical strategies to ensure safe,
cost efficient means of Moon and Mars exploration. Fourteen enabling technologies were identified to accomplish
this goal, one of the top two being nuclear thermal rocket propulsion.1
Panels of experts convened at workshops in July 1990 to evaluate the various technologies and concepts related
to nuclear thermal propulsion. The progress initiated by the workshops led to the formation of six interagency
technical panels. The Nuclear Thermal Propulsion Technology Panel identified four categories of nuclear thermal
propulsion: solid core reactors, liquid core reactors, gas core reactors, and other configurations (including nuclear
rocket using indigenous rocket fuel, hybrid systems, and dual mode systems). The panel’s goal was to have a nuclear
thermal propulsion (NTP) engine reach a technology readiness level of 6 (TRL-6, system ground test complete) by
the year 2006. The solid core reactor was selected as the concept that would most likely meet this goal. In particular,
the four solid reactor types that were identified were Nuclear Engine for Rocket Vehicle Application (NERVA)
derived, particle bed, wire core, and ceramic-metallic (CERMET) reactors.5
The Performance Comparisons Subpanel, a subgroup of the Nuclear Thermal Propulsion Technology Panel, was
tasked with comparing the merits of the solid core reactors. The subpanel decided to evaluate the concepts in terms
of specific impulse, thrust level, thrust-to-weight ratio (T/W), characteristic temperature differential, and the
resultant initial mass in low Earth orbit (IMLEO). Gary F. Polansky, one of the subpanel members, performed a
preliminary assessment of some of the concepts, including the Enabler (an updated NERVA design), a particle bed
reactor, a wire core, and a CERMET design. Polansky found that for a given trip time, IMLEO decreases with an
increase in T/W or specific impulse. In addition, performance parameters presented by the workshop focal points
listed that the PBR core had the highest T/W (20+), the highest maximum chamber temperature (3500 K), and
among the highest specific impulses (1060 s). With the caveat that his was only a preliminary study, Polansky
concluded that the PBR best satisfied the evaluation criteria.6
NASA has released three Mars design reference missions (DRM’s)11,12,1 that identify nuclear thermal propulsion
as the primary option for the inbound Mars transfer. The first DRM, released in 1997, relied on a trans-Mars
injection (TMI) stage equipped with four 66,700 N thrust, 900-s specific impulse engines to propel the crew and a
combination habitat/ascent stage along the trans-Mars trajectory.11 The cargo mission would use the same TMI
stage, except that there would be no radiation shield and one less engine.7,11 DRM 3.0, the addendum to the original
DRM, differed in that a common three-engine TMI stage was used for all cargo and crewed missions.12 But it was
the next version that saw the most drastic change in the TMI stage design. DRM 4.0 incorporated bimodal nuclear
thermal rocket (BNTR) propulsion (using the aforementioned CIS engine), in which the BNTR system produced up
to 50 kWe of power in addition to the required thrust.1 It is clear that bimodal nuclear thermal propulsion (BNTP) is
a necessary component of any human Mars architecture. However, system trades must be performed to determine
which type of BNTP would best satisfy the mass and performance requirements imposed by the Mars missions. In
addition, since the number of launches and type of launch vehicles required can be major prohibitive factors, the
2
American Institute of Aeronautics and Astronautics
relationship between launch vehicle payload capability and the corresponding TMI payload capability must be
quantified.
II.
Background and History
Nuclear thermal rockets (NTRs) differ from chemical rocket propulsion primarily in the manner that heat is
added to the propellant. In a chemical rocket, the propellant is heated by means of combustion. Nuclear thermal
rockets pass the propellant through a nuclear reactor where nuclear fission is the source of heat. While the chemical
propulsion results in specific impulse (Isp) values less than 500 s, nuclear thermal propulsion holds the potential for
Isp’s greater than 1000 s. This disparity in engine efficiency is due to the fact that molecular weight is inversely
proportional to Isp. With NTR’s ability to use a single low molecular weight propellant (such as hydrogen), the
available Isp can be double what it would otherwise be in a chemical system. Reaching a maximum Isp would then
be limited by heat transfer to the propellant and the upper limit of the core fuel temperature.3,16
Four NTR engine reactor types will be examined in this study. They are the particle bed reactor (PBR), ceramicmetallic (CERMET), NERVA–derived reactor (NDR), and Commonwealth of Independent States (CIS) “twisted
ribbon” concepts. Both the NDR and CERMET reactors used solid cores; whereas the NDR had graphite/uranium
carbide composite fuel elements, CERMET used fuel composed of a tungsten or tungsten/rhenium matrix imbedded
with uranium-dioxide particles.3,16 The CERMET fuel materials had a greater compatibility with hot hydrogen (the
uranium carbide proved to be unstable in hydrogen) and had a higher melting point.13 In addition, NDRs had to slow
neutrons down using a moderator in order for them to operate, while the fast fissioning spectrum employed in
CERMET reactors made moderators unnecessary (thereby saving mass).3 A particle bed reactor differs in that it
“consists of a number of fuel particles packed in a bed and surrounded by hexagonal moderator blocks arrayed in a
cylindrical assembly. Its distinguishing feature is that the hydrogen propellant directly cools small, coated,
particulate fuel spheres.”3 The CIS reactor is a heterogeneous design that uses ternary carbide fuel material and a
hydrogen-cooled zirconium-hydride moderator. This reactor is unique in that it “allows for optimization of the
power density across the core by changing the spacing of the fuel elements in both the radial and circumferential
directions”, providing “a more uniform fuel and exit gas temperature for each element.”17 Each reactor type is
currently at different stages of development. Only the CIS and NDR types have reached proof-of-concept validation
and been through thorough testing. Thus far, the PBR and CERMET reactors have been limited to conceptual
studies.17,18
The United States government became involved in the research and development of nuclear thermal rocket
propulsion technologies in 1955, with tests being conducted from 1959 to 1972. In 1955, the Atomic Energy
Commission and the United States Air Force began the main phases of the ROVER program with the Los Alamos
Scientific Laboratory, eventually spawning the solid-core KIWI, Phoebus, Peewee-1, and Nuclear Furnace-1
reactors. The KIWI reactors purpose was for the initial ground testing, the Phoebus for Mars mission propulsion
design, and the Peewee-1 and Nuclear Furnace-1 reactors for evaluation of advanced fuel elements.16
A joint effort of the Atomic Energy Commission and NASA, the Nuclear Engine for Rocket Vehicle Application
(NERVA) program started in 1960 under the auspices of the combined Space Nuclear Propulsion Office. A flightrated NERVA design was supposed to meet several requirements, including at least 75,000 lbf (333,617 N) thrust,
minimum chamber pressure and temperature of 450 psia and 4250 R, respectively, and a 600-minute runtime
capability with up to 60 cycles. The tested NERVA designs included the NRX reactors and the XE-PRIME engine.
The ROVER and NERVA programs led to further understanding and delineation of the performance, restartability,
and lifetime requirements of Mars mission-focused NTR engines.5,16
The most recent evaluation of an SEI PBR system was a conceptual study performed by an Aerojet/Babcock
&Wilcox consortium in 1992. They found that a 75,000 lbf (333,617 N) engine with a 2770 K reactor exit gas
temperature would produce a specific impulse of 915 seconds and a T/W of 7.2 (including shielding).4 As for the
most recent NERVA-derived model, Rocketdyne and Westinghouse found in 1992 that their reactor design could
attain a maximum T/W of 6.0 at a reactor exit temperature of 2500 K.13 In 2004, Pratt & Whitney designed a
CERMET BNTR with a specific impulse of 905 seconds, a reactor exit temperature of 2,700 K, and a propulsion
T/W of 3.6.14 Another NTR engine type, the Commonwealth of Independent States (CIS) engine, was developed by
an Aerojet, Energopool, and Babcock & Wilcox team in 1993. The CIS engine was capable of a specific impulse of
up to 955 seconds, engine T/W of 3.06, and an exhaust temperature of up to 3,075 K.1
3
American Institute of Aeronautics and Astronautics
III. Trans-Mars Injection Stage Mass Modeling
A. Vehicle and Subsystem Modeling
The system trades in this study are performed using the EXAMINE (Exploration Architecture Model for InSpace and Earth to Orbit) segment model. EXAMINE is a space vehicle model developed by D. R. Komar at NASA
Langley. It provides conceptual mass, geometry, and performance estimates for all of the subsystems (e.g., structure,
propulsion, and power). The tool’s overall purpose is to perform complete architecture modeling involving Earth-toorbit, in-space, and landing vehicles. The segment model used in this study sizes only the trans-Mars injection
(TMI) stage. For future studies, the vehicle attribute estimates created for the TMI stage could be fed into the larger
architecture model, along with data for the other pertinent vehicles.19
In addition to calculating the propellant necessary to complete the mission with the TMI stage, EXAMINE
outputs masses for the structure, protection, propulsion, power, and avionics subsystems; there are other subsystems
that can be modeled that were not necessary for this study. In this segment model, the top level requirements are the
delta-v’s (changes in velocity) associated with propulsive events, the duration of the events, and the payload to be
transported. Each subsystem has additional levels of detailed inputs for the user to specify. For example, the user can
define which type of fuel cell, if any, is to
be used in the power subsystem as well as
the power and duration requirements
imposed on the fuel cells.19
To model a BNTP subsystem (model
shown in Fig. 1), a tool was developed at
SpaceWorks Engineering, Inc. by Jon
Wallace and integrated into EXAMINE.
The BNTP model uses particle bed reactor
(PBR), NERVA-type (Nuclear Engine for
Rocket Vehicle Application), and ceramicmetallic (CERMET) reactor core types.
The power generating component is
modeled along with each of these reactor
Figure 1. Bimodal nuclear thermal rocket process.
types. Top-level performance requirements
are input (thrust level required, electrical
power required, etc.) along with nuclear
reactor, rocket propulsion, closed Brayton cycle (CBC), and dry mass margin inputs. These inputs feed into
calculations to determine the gas properties, system pressure levels, and CBC parameters. The reactor, thrust
chamber, and radiation shield are subsequently sized. The final output is a mass statement for the BNTP system
which (in addition to the reactor, thrust chamber, and radiation shield) also lists the rocket turbomachinery, NTR
support structure, CBC power conversion system, dry mass margin, and Brayton cycle working fluid mass values.
Dimensions of the thrust chamber, reactor, and radiation shield are included in the calculations and can be used for
volume estimates. For sizing the thermal control system within EXAMINE, the BNTP model also reports a radiator
heat output value.
The NERVA, PBR, and CERMET NTR mass estimating equations in this model are from the “Space Propulsion
Analysis and Design” text.3 For NERVA NTR modeling, simplified assumptions were made for the diffusion theory
calculations. Including factors such as Doppler feedback and control rod effects into the analysis goes into detailed,
complicated design. For preliminary dimensional estimates, a “lumped parameter approach” was used, wherein the
reactor is assumed to be homogeneous with constant performance through the core. These dimensions can then be
used to approximate the core’s mass. The equations used to size the PBR core resulted from a parametric analysis
that made use of the design code for the Los Alamos reactor. The 7, 19, and 37 fuel elements configurations were
chosen for the equations as to cover a wide range of possible reactor power requirements. The CERMET reactor
dimension equations were created using a least-squares curve fit of available data. As the CIS engine option was not
included in “Space Propulsion Analysis and Design”, a thrust-based mass estimating relationship (MER) was
developed using empirical data provided by Stanley Borowski (one of the architects of DRM 4.0) of NASA Glenn
Research Center.
Modeling of the CBC power generation component of the BNTP model was described by the designer, Jon
Wallace, as follows:
Performance of the closed Brayton cycle is estimated using basic thermodynamic equations for
each of the components of the system. Equations relating enthalpy, temperature, and pressure
4
American Institute of Aeronautics and Astronautics
have been created from large data sets of thermodynamic properties for the Helium-Xenon
working fluid mixture. A variety of efficiencies and pressure drop deltas are applied throughout the cycle to reduce the
overall efficiency to a more realistic level.
The NTR and CBC power conversion system operate as a BNTP system as shown in Fig. 1 above.
B. Mass Breakdown Statement
Table 1 lists the system model mass breakdown statement as it is output within EXAMINE. Primary Body
Structure mass consists of the stage forward skirt, aft skirt, and thrust structure. The skirts are cylindrical structures
7.4 m in diameter (diameter of the tank) and 4.2 m in length. The thrust structure model uses the tank diameter (7.4
m), engine thrust, and number of engines as inputs into a mass estimating relationship (MER) used to determine the
structure’s mass. Micrometeoroid and Orbital Debris (MMOD) Protection is calculated based on inputs of unit
weight (1.52 kg/m2), thickness (0.0005 m), standoff distance (0.0065 m), and surface area (27 m2).
While the Reaction Control System (RCS) Engine & Installation is not broken down into lower component
levels, NTR Engine & Installation includes the reactor and thrust chamber, radiation shield, rocket turbomachinery,
NTR support structure, and the closed Brayton cycle power conversion system. These masses were calculated using
the inputs from Table 3 below; the resulting mass breakdown of the components is also included in Table 3. Orbital
Maneuvering System (OMS) Tanks & Feed/Fill/Drain System includes the pressure vessel, multi-layer insulation,
cryocooler system, vacuum jacket, supports, penetration, and ancillary components (feed, fill, and drain system).
The 2-stage cryoocooler employed in EXAMINE is a zero boil-off cryogenic storage system.19 Inputs used in
calculating all of the aforementioned masses are listed in Table 2.
Table 1. System Model Mass Breakdown Statement
1.0 Structure
Primary Body Structure
2.0 Protection
MMOD Protection
3.0 Propulsion
NTR Engine & Installation
RCS Engines & Installation
OMS Fuel Tanks & Feed/Fill/Drain System
RCS Fuel Tanks & Feed/Fill/Drain System
RCS Oxidizer Tanks & Feed/Fill/Drain System
Pressurization System
4.0 Power
PMAD
5.0 Avionics
Command, Control, and Data Handling
Communications
6.0 Environment
Thermal Control System
Heat Acquisition
Heat Transport
Heat Rejection
Miscellaneous Passive & Active TCS
7.0 Non-Cargo
Reserve, Residual Fluids, and Gases
Pressurant
OMS/RCS Press
OMS Fuel Autogeneous Pressurization
Unused Fuel
Residual OMS
8.0 Propellant
Usable OMS Fuel
Usable RCS Fuel
Usable RCS Oxidizer
The Power Management and Distribution (PMAD) system is sized based on the vehicle’s 50 kWe peak power
requirement, as identified in DRM 4.0 and as listed in Table 2. The Avionics modeled on this vehicle consist of the
communications components and the command, control, and data handling components. These avionics are all
linearly scaled based on heritage data.19 In the Environment subsystem, the Thermal Control System is sized to
manage the heat produced by the NTRs (an output of the BNTP model). The Miscellaneous Passive & Active TCS
component includes the total heat pump, shell heater, and insulation masses. Non-Cargo components include
Pressurant and Residual Orbital Maneuvering System (OMS) Fuel. Helium is used as the RCS pressurant, while the
main fuel tank uses a portion of the LH2 stored within it for autogenous pressurization. As shown in Table 2, the
Residual Fuel is calculated as 3% of the usable fuel. The Usable Propellant for the OMS and the RCS is based on the
delta-v requirements, the Isp of the propulsion systems, and the masses of the other components of the vehicle.
IV. Mission Analysis for Bimodal Nuclear Thermal Propulsion Designs
A. Concept of Operations
5
American Institute of Aeronautics and Astronautics
The portion of NASA’s 1998 DRM modeled in this paper is the 2014 Piloted Lander Mission. In this Mars
architecture, two cargo missions are completed before the Piloted Mission can begin. The Cargo Lander and Earth
Return Vehicle transports are launched on a low-energy trajectory towards Mars. Once both vehicles arrive at Mars
and complete the necessary ground and on-orbit operations, the piloted mission would be cleared to launch three
years later. All three mission types are sent on conjunction class trajectories and use the same core stage, sized for
the energetically demanding 2016 Cargo ERV Mission.
For the piloted mission, the core stage will be launched into low-Earth orbit (LEO) first. The crewed component
would be launched and would dock with the core stage within a ~32 day window. After a ~two day vehicle checkout
period, the TMI engines would perform its single ~39 minute burn and send the crew on their 180-day trajectory.
Unlike in previous DRMs, the TMI stage will not be jettisoned immediately after its burn. The core stage would
remain attached to provide electrical power and midcourse correction propulsion to the crew vehicle. In the DRM
scenario used for this study, the core stage would not bring the crew into Mars orbit but rather would be jettisoned
prior, being disposed along the interplanetary path.1
B. Ground Rules and Assumptions
Accurately modeling the alternative BNTP systems for comparing to the DRM CIS engine masses requires that
they share the same performance requirements. The ground rules and assumptions for this study come directly from
DRM 4.0 and are shown in Table 2 below.1 The active control system for zero boil-off assumes the necessary
technological advances will occur before the Mars missions transpire.
Table 2. Ground Rules and Assumptions1
Mission
Mission Type
Earth Departure Orbit
Total delta-v
TMI Burn
296 km to 407 km Hohmann Transfer
Gravity Loss
1% Flight Reserve
Tank Sizing
Material (for tank and all other structures)
Acceleration Load
Internal Tank Pressure
Residuals
Ullage Factor
Safety Factor
Cooldown
Shape
Inner Diameter
MMOD Protection
Contingency/Margin
Dry Mass
Piloted Conjunction Class (long-stay)
407 km Circular
4.157 km/s
3.672 km/s
0.064 km/s
0.380 km/s
0.042 km/s
Composites
5g
241316 Pa
3%
3%
1.5
3%
Cylindrical
√2/2 Domes
7.4 m
0.55 mm Shielding
Boil-off Control
Multi-layer Insulation
Heat Flux
Sink Temperature
Active Boiloff Control
80 Layers
2.1 inch total thickness
0.161 W/m2
230 K
Zero Boil-off Refrigeration
BNTP System
Thrust per Engine (#)
Power Generation
Propellant
65656 N (3)
50 kWe
LH2
RCS
Delta-v
Propellant
0.1 km/s
MMH / N2O4
Payload Mass (t)
Total Payload Mass
15%
56.38
Crew (6) & Suits
Surface Payload
Descent Stage
Aerobrake/Descent Shell
Parachutes
Propellant
1.44
26.81
4.20
13.24
0.70
9.99
C. Engine Model Inputs
DRM 4.0 uses a core stage propulsion system with three 65,656 N CIS engines operating with an Isp of 955 s.
Up to 50 kWe would be created by any two of the engines, each using a 25 kWe Brayton rotating unit. This CBC
power conversion system operates at ~2/3 of its rated capacity, thereby ensuring an engine-out capability.1 The 7element PBR, 19-element PBR, and CERMET designs were able to operate at three engines at 65,656 N each, but
running the 37-element PBR and NERVA-derived reactors at that setting violated their minimum reactor power
limits. In order to maintain the NERVA-derived design as an option, two alternative settings with one and two
engines were included; the two-engine option was used only to ensure at least one NERVA-derived design with an
engine out capability. For the NERVA-derived options, the closest amount of thrust possible to match the combined
196,968 N of the other BNTR types (three engines operating at 65,656 N each) was 220,209 N. This value was
brought up to 220,630 N so as not to be too close to the outer limits of the design space. For power generation, the
BNTP model can only divide the total power requirement equally between each of the engines. Therefore, the
configurations with three engines operated at 16.67 kWe per engine, the two-engine configuration at 25 kWe per
engine, and the one-engine configuration at 50 kWe.
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American Institute of Aeronautics and Astronautics
Expansion ratios for the three alternative BNTP systems were set at the highest tested or analyzed values.3,16 The
300:1 expansion ratio of the CIS engine is the setting used in DRM 4.0.1 The chamber temperatures for each type
were set at the maximum values tested amongst the designs used for the parametric equations. Table 3 also lists the
Isp, propellant flow rate, and required reactor power that resulted from the inputs to the model. For the CIS and
alternative engines, H2 was used as the propellant. The propellant tank pressure values for the model were set at the
241,317 Pa used in the DRM.1
Table 3. Engine Specifications and Calculated Values
Reactor Type
Engine Thrust (#)
Expansion Ratio
Chamber Temperature
Chamber Pressure
Specific Impulse
Propellant Flow Rate
T/W System
Required Reactor Power
NTR Gross Mass
Reactor and Thrust Chamber
Radiation Shield
Rocket Turbomachinery
NTR Support Structure
Closed Brayton Cycle Power Conversion System
Brayton Cycle Working Fluid
kN
--K
MPa
s
kg/s
--MW
kg
kg
kg
kg
kg
kg
kg
CIS
65.66 (3)
300:1
2889
13.79
955
7.01
0.14
335
6,670
-------------
PBR 7-element PBR 19-element
65.66 (3)
65.66 (3)
125:1
125:1
3200
3200
6.89
6.89
1015
1015
6.60
6.60
0.16
0.16
343.73
343.73
8,375
7,874
3,924
2,408
2,563
3,624
85
85
657
612
996
996
149.42
149.42
CERMET
NERVA 1-engine NERVA 2-engine
65.66 (3)
220.63 (1)
220.63 (2)
120:1
100:1
100:1
2507
2361
2361
4.14
3.10
3.10
885
850
850
7.56
26.48
26.48
0.15
0.15
0.29
299.37
977.99
977.87
5,149
5,271
10,050
2,398
2,196
4,391
1,153
1,849
3,697
88
87
174
364
413
826
996
632
836
149.42
94.75
125.38
For piloted missions, a shield is required to mitigate the crew’s exposure to the radiation from the reactor. In the
BNTP model, radiation shielding is calculated using an areal density (kg/m2) constant supplied by the user. It
assumes a cylindrical reactor shape and calculates the area from the reactor radius. The areal density used in
determining each BNTP system’s shield mass was calculated based on the 0.325 m CIS engine vessel radius and its
3.24 t radiation shield.1
V. Validation
EXAMINE was used to model the 2007 Cargo ERV Mission and the 2014 Piloted Mission from DRM 4.0. A
mass comparison of the DRM 4.0 and EXAMINE mass breakdowns for these two vehicle types is shown in Table 4
below. The characteristics of the two TMI stage vehicles designed for the propulsion requirements of these two
Table 4. 2014 Piloted Mission Bimodal Nuclear Thermal Propulsion Vehicle Mass Comparison
BNTR Core Stage Elements
Structure
Avionics and Power
Reaction Control System (RCS)
Propellant Tank
Passive TPS/Micrometeor Shield
LH2 Refrigeration System (@ ~75 Wt)
Brayton Power System (@50 kWe)
NTR Assemblies
NTRs
External Radiation Shields (3)
Propellant Feed, etc.
Contingency (15%)
"Dry" Bimodal Core Stage
LH2 Propellant (max LH2 Capacity)
RCS Propellant
Fuel Cell Reactants (O2)
"Wet" Bimodal Core Stage
Mass (t)
2007 Cargo ERV Mission
2014 Piloted Mission
56 t DRM Core EXAMINE 51 t DRM Core EXAMINE
2.45
2.30
2.50
2.29
1.20
0.74
1.47
0.10
0.42
1.04
0.45
1.03
6.66
7.83
5.98
6.67
1.39
1.00
1.29
0.89
0.30
0.01
1.35
1.35
6.67
6.67
0.56
2.90
25.59
56.00
0.77
0.43
82.79
0.33
2.99
22.89
56.60
1.56
0.48
82.31
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American Institute of Aeronautics and Astronautics
6.67
2.82
0.47
3.50
26.80
50.19
1.83
6.67
2.82
0.30
3.32
25.44
50.73
1.31
78.82
77.48
energetically demanding missions are shown in Fig. 2 and Fig. 3.1 The ERV Cargo Mission uses propulsion-only
NTRs, while the Piloted Mission uses BNTRs.
Figure 2. TMI Stage Sized for the ERV Cargo Mission1
Figure 3. TMI Stage Sized for the Piloted Mission1
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VI. Results
A. Analysis Based on Design Reference Mission 4.0 Thrust Requirement
The Piloted Mission vehicle mass breakdown for each of the reactor types as calculated in the model is displayed
in Table 5. For the three-65,656 N engine propulsion system option, the 19-element PBR and the CERMET reactor
had lower engine masses than the CIS variant, whereas the 7-element PBR is slightly higher. The radiation shield
mass, a function of the radius of the reactor core, is lower for the CERMET case. At this thrust and power level (the
PBRs have a calculated 343.73 MW power requirement compared to 299.37 for the CERMET reactor, as shown in
Table 3 above), the CERMET engine is both smaller and less massive. As expected, the one and two-NERVA
engine types, with their 220,630 N and 441,260 N thrusts, respectively, have the largest engine masses. With their
1015 s Isp’s, the PBR engines have the lowest performance requirements. The less efficient 885 s Isp CERMET
engine needs more propellant for the mission than the 955 s Isp CIS option.
Table 5. Piloted Mission Vehicle Mass Comparison Using EXAMINE Thrust-based Sizing
Mass (t)
Reactor Type
Structure
Avionics
Reaction Control System (RCS)
Propellant Tank
Passive TPS/Micrometeor Shield
LH2 Refrigeration System
Brayton Power System (@50 kWe)
NTR Assemblies
NTRs
External Radiation Shields
Propellant Feed, etc.
Contingency (15%)
"Dry" Bimodal Core Stage
LH2 Propellant
RCS Propellant
"Wet" Bimodal Core Stage
IMLEO (with 56.38 t payload)
CIS
PBR7
PBR19
CERMET
2.29
0.10
1.03
6.87
0.91
0.30
1.35
2.29
0.10
0.98
6.22
0.84
0.01
1.27
2.29
0.10
0.98
6.16
0.83
0.01
1.27
2.29
0.10
1.09
7.36
0.96
0.01
1.27
6.67
2.82
0.31
3.40
26.04
52.09
1.33
79.46
135.84
7.20
2.56
0.29
3.27
25.03
47.68
1.28
73.99
130.37
5.64
3.62
0.29
3.18
24.39
47.31
1.27
72.97
129.35
5.39
1.15
0.33
2.99
22.95
55.31
1.33
79.60
135.98
NERVA NERVA
1-engine 2-engine
2.30
2.30
0.10
0.10
1.23
2.22
8.61
9.39
1.08
1.15
0.01
0.02
0.91
1.11
9.30
1.85
0.36
3.86
29.61
63.28
1.48
94.36
150.74
8.94
3.70
0.53
4.42
33.87
68.03
1.57
103.48
159.86
B. Analysis Based on Minimum Thrust-to-Weight Requirement and 30% Contingency
Unfortunately, all of the TMI stages detailed in Table 5 are based on DRM 4.0 thrust requirements that result in
system T/W’s of approximately 0.15 (see Table 3 for specific values). In order to counteract the effects of gravity
loss incurred during the TMI burn, an initial vehicle T/W of at least 0.3 is necessary (the NERVA 2-engine option
has a relatively high T/W of 0.29 because its thrust is at least twice that of the other propulsion systems with a less
dramatic mass increase).3 To this end, EXAMINE was used to calculate the vehicle masses based on a 0.3 T/W
constraint.
In addition to the T/W adjustment, another modification in the modeling was necessary to make the results more
realistic. A 15% contingency is optimistic and does not reflect the mass growth experienced by past aerospace
projects. To more accurately capture the mass that can be expected when the final vehicle is delivered, a
contingency of ~30% is more appropriate.20 Table 6 shows vehicle masses that result when using a T/W of 0.3 and a
historically accurate contingency value of 30%.
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American Institute of Aeronautics and Astronautics
Table 6. Piloted Mission Vehicle Mass Comparison Using EXAMINE T/W = 0.3 and 30% Contingency
Mass (t)
Reactor Type
Structure
Avionics
Reaction Control System (RCS)
Propellant Tank
Passive TPS (@2" MLI)/Micrometeor Shield
LH2 Refrigeration System
Brayton Power System (@50 kWe)
NTR Assemblies
NTRs
External Radiation Shields
Propellant Feed, etc.
Contingency (30%)
"Dry" Bimodal Core Stage
LH2 Propellant
RCS Propellant
"Wet" Bimodal Core Stage
IMLEO (with 56.38 t payload)
CIS
PBR7
PBR19
CERMET
2.30
0.10
2.02
8.10
1.03
0.30
1.35
2.30
0.10
1.84
7.26
0.95
0.01
1.27
2.30
0.10
1.75
6.82
0.90
0.01
1.27
2.30
0.10
2.09
8.36
1.05
0.01
1.27
9.50
2.82
0.48
8.40
36.39
60.08
1.52
97.98
154.36
10.10
2.56
0.44
8.05
34.87
54.66
1.45
90.98
147.36
6.43
3.04
0.41
6.91
29.95
51.71
1.37
83.03
139.41
6.13
1.15
0.50
6.89
29.87
61.71
1.47
93.04
149.42
NERVA NERVA
1-engine 2-engine
2.31
2.31
0.10
0.10
2.69
2.54
11.53
10.65
1.34
1.26
0.02
0.02
0.91
1.11
13.68
5.45
0.64
11.60
50.26
80.57
1.85
132.69
189.07
9.90
5.01
0.61
10.05
43.55
75.50
1.74
120.79
177.17
According to the system analysis with the 0.3 T/W constraint and realistic 30% contingency, the PBR 19element BNTR results in the lowest IMLEO. Therefore the capabilities of this BNTP system should be further
explored to understand how it can be used effectively.
C. Trans-Mars Injection Stage Payload Capability and Launch Vehicle Loading
Since the Magnum 80 t-payload-capable launch vehicle was employed in the DRM 4.0 study, there have been
further developments in heavy-lift launch vehicle design. Most notably, NASA’s Exploration Systems Architecture
Study identified the Ares V as the best candidate for satisfying both lunar and Mars mission ETO transportation
needs.15 The Ares V can transport approximately 128.8 t of payload to low-Earth orbit, at which point the TMI stage
would propel the cargo to Mars.
As launch vehicle useable payload capability changes, so does the payload capability of the TMI stage.
EXAMINE was used to analyze this relationship between launch vehicle and the 19-element PBR TMI stage and
formulate a trend. Figure 4 shows this trend for the PBR propulsion system, with the corresponding Ares V useable
capability. This approach assumes one launch that includes both the TMI stage and its payload.
In the aforementioned NASA study, the Earth Departure Stage (EDS) is used as the trans-Mars injection stage.
In the concept of operations, the EDS fires after the Ares V had completed its burn. Using a LOX/LH2 chemical
propulsion system, the EDS can achieve an isp of 454 seconds with a T/W of 43. A trend was formed relating the
EDS’s payload capability to the launch vehicle useable payload capability. A comparison between the EDS and PBR
TMI stages is pictured in Figure 5. Using the Ares V (denoted in the figure by the dotted line), the PBR TMI stage
can transport nearly twice as much payload as the chemical EDS.
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American Institute of Aeronautics and Astronautics
Figure 4. 19-element Particle Bed Reactor Bimodal Nuclear Thermal Rocket Trans-Mars Injection
Payload Capability
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American Institute of Aeronautics and Astronautics
Figure 5. 19-element Particle Bed Reactor Bimodal Nuclear Thermal Rocket Trans-Mars Injection
Payload Capability Compared with Current Chemical Earth Departure Stage
VII. Conclusion
This preliminary analysis can serve as a starting point for a more in-depth study. PBR BNTP-powered core
stages have been identified as the least massive of the four options presented in this paper (in particular the 19element design). Besides IMLEO, there are other considerations that must factor into making the final selection. The
fact that NERVA and CIS are farther along in development and testing than the other options is significant.
However, the combination of the facts that NERVA-type engines have not been tested at the low thrust levels
required for this mission and that they have appreciably higher resulting IMLEOs may eliminate the propulsion
system as a candidate. In the case of the CERMET core, its superior endurance, restart ability, and compatibility
with the hydrogen propellant may outweigh its higher IMLEO.
The next step in making comparisons between the BNTP options would be to perform a more comprehensive
study that trades cost, schedule, and feasibility. In addition, once the less attractive propulsion system options are
eliminated, a higher level mass estimate can be performed to solidify the relative rankings. The degree to which the
superior propulsion system would progress would then be a factor of the current policies and will of the decision
makers.
Acknowledgments
The authors would like to thank Jon Wallace of SpaceWorks Engineering, Inc. for creating the bimodal nuclear
thermal propulsion model. Our appreciation also goes out to Douglas Stanley of the Georgia Institute of
Technology/National Institute of Aerospace for his help in determining the modeling approach for this paper.
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American Institute of Aeronautics and Astronautics
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