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IN-SITU PROPELLANT INFRASTRUCTURE FOR ALTERNATIVE PROPELLANT
NUCLEAR THERMAL PROPULSION ENGINES
Dennis Nikitaev1 and Dr. L. Dale Thomas1
1301 Sparkman Dr., Huntsville, AL, 35899
Primary Author Contact Information: 702-287-6852 dn0038@uah.edu
Abstract
Recently, NASA has pushed for a return to the Moon with In-Situ resource utilization (ISRU)
being the central focus to sustainably achieve this goal. It so happens that the Moon has an
abundancy of water and ammonia in permanently shadowed regions. These are two potential
alternative propellants, are much more dense than liquid hydrogen, and do not require post
processing, such as electrolysis, to be used directly by Nuclear Thermal Propulsion (NTP) engines.
These Alternative propellant NTP (A-NTP) engines have a lower specific impulse than the
reference engines and the initial vehicle mass will be higher thus requiring more mining efforts if
the architecture is based around A-NTP vehicles. It could be advantageous to augment a reference
architecture with an A-NTP Supplemental Vehicle to increase the total number of missions
performed by the architecture if enough of either water or ammonia is produced to support a
Supplemental Vehicle performing the same mission as the Main Mission Vehicle.
By using an abstracted analysis from a cost per mission perspective, it was determined that if
the resulting cost per mission of the Supplemental Vehicle is lower than the Main Mission Vehicle,
then the overall cost per mission of the architecture will be lower. However, the cost per mission
may not be the only parameter that needs to be considered when selecting an architecture. The
total architecture cost and the total missions performed by the architecture are also important
parameters. Furthermore, the costs of the architecture elements will scale with economies of scale.
Therefore, larger architectures with several vehicles performing multiple missions simultaneously
are expected to result in the lowest costs per mission.
Nomenclature
Symbols
C = Maximum rated reactor power level
= Specific impulse, s
ΔV = Change in velocity, m/s
Subscripts
miss = Mission
E = Engine
2
Infra = Infrastructure
MMV = Main Mission Vehicle
SV = Supplemental Vehicle
V = Vehicle
Acronyms
A-NTP = Alternative propellant Nuclear Thermal Propulsion
AR = Aerojet Rocketdyne
Exp = Expendable
F9 = Falcon 9
FH = Falcon Heavy
H-NTP = Hydrogen Nuclear Thermal Propulsion
IHOP = ISRU water purification and Hydrogen Oxygen Production
ISRU = In-Situ Resource Utilization
LADV = Lunar Ascent/Descent Vehicle
LDHEO = Lunar Distant High Earth Orbit
LDRO = Lunar Distant Retrograde Orbit
LOX = Liquid Oxygen
NAFCOM = NASA Air Force COst Model
NTP = Nuclear Thermal Propulsion
NERVA = Nuclear Engine for Rocket Vehicle Application
Par-Recov = Partially Recoverable
Recov = Recoverable
SLS = Space Launch System
I. INTRODUCTION/BACKGROUND
NASA’s Artemis Program aims to produce propellant In-Situ on the Lunar surface as useful
volatiles such as water, ammonia, methane, and carbon dioxide have been found in permanently
shadowed regions [1,2]. The goal of this program is to gain the experience necessary to
successfully send crewed missions to Mars with In-Situ Resource Utilization (ISRU) being the
central focus. [1] This will require mining Lunar volatiles and processing them through filtration
3
to obtain propellant grade water which will then be fed into electrolyzers to obtain separated
hydrogen and oxygen which could be used by both H2+LOX Chemical and hydrogen nuclear
thermal propulsion (H-NTP) engines.
Nuclear Thermal Propulsion (NTP) is a concept which dates to the late 1950’s through the
early 1970’s with reactor and engines demonstrated during Project Rover and later by the Nuclear
Engine for Rocket Vehicle Application (NERVA) program with a goal of creating a propulsion
system capable of transporting humans to Mars. Unlike chemical propulsion, NTP does not depend
on combustion of an oxidizer and fuel to produce thrust; instead, a propellant is pumped into a
nuclear reactor and heated to high temperatures before being expelled through the nozzle.
Essentially, the NTP based engine is a monopropellant system and can use any propellant provided
that significant core degradation does not occur. [3]
Aerojet Rocketdyne (AR) has developed low enriched uranium NTP engine power balance
models that use hydrogen as the propellant due to this species’ potential to yield high efficiency
of propellant usage, also known as specific impulse (Isp). The Isp of H-NTP engines is around 900
seconds which is twice as much as H2+LOX Chemical propulsion [4,5]. A higher Isp results in
lower required propellant mass to accelerate a set dry mass by a set ΔV resulting in lower initial
wetted mass of the vehicle. In general, lower mass for space vehicles results in cost savings on the
missions which these vehicles must perform according to the NASA Air Force COst Model
(NAFCOM) [6]. Therefore, vehicle mass is the prime limitation for space missions according to
the Ideal Rocket Equation.
Since NTP can theoretically use any fluid as a propellant, it is necessary to understand the
impact to missions if the infrastructure were to be simplified by eliminating electrolyzers and using
water or ammonia directly as propellants. The availability of propellant in space fundamentally
changes the approach that must be taken for selecting a propulsion system for a mission. This is
because a portion of the vehicle mass does not need to be brought from Earth and can be mined
directly on the Lunar surface [2]. Therefore, engine and vehicle reusability become the central foci
and not the efficiency of propellant usage [7].
I.A. In-Situ Propellant Production
The Commercial Lunar Propellant Architecture (CLPA) study outlined how laboratory tested
hardware can be used to effectively mine Lunar volatiles in permanently shadowed regions and
produce propellant grade water, filtered ammonia, and hydrogen and oxygen via water electrolysis
by using the ISRU water purification and Hydrogen Oxygen Production (IHOP) System with the
volatile flow diagram shown in Figure 1. Here, sublimation mining tents transfer heat into the
Lunar regolith and, due to near vacuum pressures, the volatiles sublimate and are trapped by the
tents. Most of the volatiles are filtered through a cold trap where the condensed raw steam (and
ammonia vapor) is fed into an ammonia scrubber where the water and ammonia are separated. The
water is then polished to remove any residual impurities such as dust and is fed into an electrolyzer
unit. An important characteristic of the IHOP is that it scales linearly and the components have a
life of 10 years and each mining tent can yield 1875 mton of water and 113 mton of ammonia per
year. [2] The entire system is proposed to be powered by Kilopower Reactor Using Stirling
Technology (KRUSTY) nuclear electrical power units which also have an estimated expected life
of 12 to 15 years [8]. Therefore, it is important to analyze infrastructure requirements that will
support a series of missions for up to 10 years.
4
Figure 1: ISRU water purification and Hydrogen Oxygen Production (IHOP) System Flow
Diagram [2]
I.B. Reference Missions
There are two types of missions that can benefit from In-Situ produced propellant: (1) Lunar
surface to Lunar orbit such as ferrying propellant to a space transfer mission and (2) space transfer
missions such as those originating from a Lunar orbit and transporting a payload to Mars. The
NASA Artemis Program considers the Lunar Distant Retrograde Orbit (LDRO) to be the parking
orbit for space transfer vehicles [1,9–11]. If round trip missions are considered, then the vehicles
performing these types of missions will become inherently reusable once the advantages of ISRU
are realized.
An example of a Lunar surface to Lunar orbit mission is that which a Lunar Ascent/Descent
Vehicle (LADV) will perform with a total ΔV of 6856 m/s. This particular mission architecture
considers a total mission time of 24 hours with H2+LOX Chemical propulsion provided by RL-
10 engines for which propellant is obtained completely In-Situ. The purpose of this mission is to
ferry propellant as payload to a customer vehicle parked in any orbit that has a total orbital energy
equal to or less than LDRO. [11] Based on a 10-year infrastructure life, a single LADV will be
able to perform a maximum of 3650 missions when no downtime between missions is considered.
Furthermore, engine blocks are considered to be replaceable when the RL-10 engines reach the
end of their operational life [11] of 1.25 hours [12].
There are two examples of Mars missions which would originate and end in LDRO which are
Mars Conjunction and Mars Opposition missions. The Mars Conjunction mission aims to
minimize the required ΔV resulting in 4223 m/s and maximize the time the crew will spend on the
5
Martian surface resulting in 620 days. The round-trip transit time will be 356 days yielding a total
mission time of 976 days. If a Mars mission opportunity window every 26 months is considered,
then a single vehicle can perform 3 Mars Conjunction missions within the 10-year life of the
infrastructure.
The Mars Opposition mission aims to minimize the total time the crew spends in space by
minimizing the time on the Martian surface to 50 days. This results in a total round-trip transit
time of 679 days and a total mission time of about 2 years. When considering a Mars mission
window of every 26 months, then a total of 4 Mars Opposition missions can be performed by a
single vehicle within the 10-year life of the infrastructure.
II. Alternative Propellant NTP Engines and Their Applications
Alternative propellant NTP (A-NTP) engines were analyzed in previous work to understand if
it is feasible to use the separated and filtered Lunar resources directly without electrolysis and thus
simplify the Lunar propellant infrastructure [13]. It was determined that A-NTP engines using
ammonia as a propellant are viable due to the absence of oxygen as a catalyst for corrosion [14–
16] under a maximum fuel temperature limitation of 2850 K according to BWXT, the reactor
manufacturer [4]. Water A-NTP engines are also viable provided that Silicon Carbide cladding is
used and the maximum cladding surface temperatures do not exceed 2400 K to provide a total
engine life of 5 hours [17]. The parameter that sets this engine life limit is the recession rate of the
cladding. For longer engine life, a maximum cladding surface temperature of 1400 K could be
considered [18] which would then yield an engine life equal to the time it takes to utilize 2% of
the total 235U mass [19]. However, no study has been done to identify the recession rates of SiC
cladding at temperatures between 1400 K and 2400 K [18,20–23].
Bleed cycle A-NTP engines will not be viable with water or ammonia due to the lack of a
gaseous phase prior to the nozzle chamber as the propellant is considered to boil inside the reactor
[13,24]. Therefore, instead of mixing the hot bleed flow with a cold bleed flow to power the
turbines as was done in the NERVA Xe-Prime engine [25,26], all of the bleed flow to the turbines
will be at the chamber temperature exceeding the current turbine material temperature limitation
of 1150 K [27–29] which will result in infeasible engines. Therefore, only expander cycle A-NTP
engines will be considered which provide specific impulse values of up to 371 seconds for
ammonia, 318 seconds for water with a 2400 K maximum cladding surface temperature, and 228
seconds for water with a 1400 K maximum cladding surface temperature. Water A-NTP engines
can provide double the thrust of the H-NTP engines at the same reactor power level and ammonia
A-NTP engines can provide 70% higher thrust than H-NTP engines at the same reactor power
level. [13]
Since the Isp of A-NTP engines is lower than both hydrogen NTP (H-NTP) (900 seconds) and
H2+LOX Chemical engines (465 seconds), it is expected that the initial mass of the vehicle will
be higher when A-NTP engines are utilized due to increased propellant mass based on the Ideal
Rocket Equation. Although no propellant transportation occurs and all the propellant produced is
directly available for the LADV, the wetted vehicle mass is still a limitation to the thrust to weight
ratio for the LADV to be able to complete its mission. The limiting factors on the number of flights
performed per engine block for the LADV are the thrust and burn times. If the thrust increase of
A-NTP engines makes up for the increased initial wetted vehicle mass to provide the same or
higher thrust to weight ratio than the reference RL-10 engines, then A-NTP engines for the LADV
6
could be utilized and thus require a simpler Lunar infrastructure without electrolyzers and only
require mining tents and separation/purification units.
The Mars missions, unlike the LADV, are not limited by the thrust to weight ratio to complete
their mission. Instead, they are strictly limited by the engine life which must exceed the burn time
of at least a single mission to be viable. The burn time is impacted by both the increased initial
vehicle wetted mass from a lower Isp and the increased thrust of the A-NTP engines. According to
the Ideal Rocket Equation, the increase in initial vehicle wetted mass scales exponentially when
the ΔV is increased. Therefore, the Mars Opposition mission will require much more propellant
mass than the Mars Conjunction mission when a raw Lunar propellant is used in an A-NTP engine.
This will result in fewer missions available per engine block for the Mars Opposition mission than
the Mars Conjunction mission since the burn time will be longer due to increased ΔV and increased
vehicle mass.
III. INFRASTRUCTURE ANALYSIS
There are three Lunar infrastructure variations that are made possible once A-NTP engines are
considered as shown in Figure 2. The Reference H2+LOX Production infrastructure aims to
electrolyze water to produce propellant. Since the RL-10 engine uses a hydrogen rich mixture [30],
excess oxygen resulting from the stoichiometric hydrogen and oxygen production from water will
be used for life support systems [2]. Similarly, H-NTP engines only use hydrogen, therefore, all
the produced oxygen will go to life support systems. All other mined volatiles are considered to
be vented and lost.
If the utilization of the scrubbed ammonia is considered while keeping the reference
infrastructure the same, then the H2+LOX Production + Ammonia infrastructure results as shown
in Figure 2. Here, two vehicles can be supported by the same infrastructure at the same time. The
Main Mission Vehicle is the one around which the reference infrastructure is built while the
Supplemental Vehicle simply uses the left-over ammonia. An issue with this infrastructure is that
ammonia constitutes only 4.52% of the total available volatiles on a molar basis while water
constitutes 74.91% as shown in Table 1 [2]. The molecular weights of water and ammonia are
similar with ammonia having 94.44% the mass of water per mole [31]. This results in 0.057 kg of
ammonia for every 1 kg of water mined. However, due to the utilization of hydrogen rich mixtures,
more water mass will need to be mined than usable propellant mass produced. An extreme case of
this would be missions that utilize H-NTP where hydrogen constitutes only 11.11% of the total
water mined [32]. This will result in 0.51 kg of ammonia per 1 kg of hydrogen produced. Once the
significantly lower Isp of ammonia is also considered, this infrastructure will not produce enough
ammonia propellant mass per Main Mission Vehicle for the Supplemental Vehicle to complete the
same mission. Therefore, to utilize the ammonia alternative propellant with the reference Lunar
infrastructure, several retankings of the Main Mission Vehicle must occur to produce enough
ammonia to retank a Supplemental Vehicle once which performs the same mission. Otherwise, a
different mission architecture with a lower ΔV and lower payload and dry mass must be considered
for the Supplemental Vehicle.
The Separated Water and Ammonia infrastructure shown in Figure 2 produces either water or
ammonia for use by an A-NTP based Main Mission Vehicle and the Supplemental Vehicle uses
either ammonia or water, respectively. It is important to note that this infrastructure does not
require electrolysis units. Due to the scarcity of ammonia in the Lunar regolith as shown in Table
7
1, an infrastructure based around an ammonia A-NTP Main Mission Vehicle will produce enough
water for several Supplemental Vehicles to perform the same mission. However, a similar situation
to the H2+LOX Production + Ammonia infrastructure will occur when water is produced for the
Main Mission Vehicle resulting in the necessity of retanking the Main Mission Vehicle several
times before a Supplemental Vehicle could be retanked once.
8
Figure 2: Lunar Infrastructure Variations
9
Table 1: Molar Composition of Lunar Volatiles [2]
Compound
Concentration (%)
74.91197843
12.54775639
4.517192299
2.389692112
2.337253727
1.625589932
1.161135666
0.48692786
0.022473594
IV. ARCHITECTURE ANALYSIS
The Lunar infrastructure for propellant production is a small part of the total architecture
necessary to support reusable vehicles for crewed Mars missions as shown in Figure 3 which
depicts the full architecture. Here, various launch vehicles are considered for different elements of
the architecture with the details on how the elements are allocated among the different launch
vehicles along with associated launch costs shown in Table 2. All launch vehicles are considered
to send their payload into a highly elliptical Lunar Distant High Earth Orbit (LDHEO) with the
payload providing the transfer into both LDRO and the Lunar Surface (L-Surf). The mining tents
are absent from this table as their total unit and launch costs together were taken from the CLPA
study and total to $103M [2]. Elements that are to be sent to L-Surf already have the Lunar Cargo
Lander (LCL) mass built into the payload fairing carrying them with parameters calculated from
the Ideal Rocket Equation using the landing ΔV of the LADV of 3668 m/s and the same Isp as the
upper stage engines of the launch vehicles. Elements that need to be sent into LDRO are considered
to have the required 140 m/s ΔV provided for them by the Reaction Control System (RCS). This
system was also modeled according to the Ideal Rocket Equation with a Isp of 320 seconds
consistent with the reference Mars mission architectures [33].
Once the Mars Transfer Vehicles (MTV), LADV, and Lunar surface infrastructure are in place,
the architecture is maintained by the launch vehicles providing engine replacements as needed
once the usage time on the engines reaches their operational life. The considered launch vehicles
are SpaceX Falcon 9 (F9), SpaceX Falcon Heavy (FH), Space Launch System (SLS) Block 2, and
Blue Origin New Glenn. Furthermore, expendable (Exp), recoverable (Recov), and partially
recoverable (Par-Recov) versions of the SpaceX launch vehicles are considered. The In-Situ
produced propellant is transferred to the LADV for both propulsion usage and to ferry as payload
to the MTV parked in LDRO. As presented in Figure 3, this architecture can not only support
multiple reuses of both the LADV and MTV, but it could also be scaled to support multiple
instances of these vehicles to perform parallel missions simultaneously. The utilization of parallel
missions can unlock the possibility of using ammonia A-NTP versions of these vehicles
simultaneously along with the reference vehicles if the cumulative amount of ammonia produced
is enough to support these missions. The number of required missions performed by the Main
Mission Vehicle to yield enough ammonia to support an ammonia A-NTP Supplemental Vehicle
10
is obtained by dividing the required ammonia mass for the Supplemental Vehicle by the ammonia
mass produced per each tanking of the Main Mission Vehicle and rounding up to an integer value.
Figure 3: Full Architecture
11
Table 2: Unit Allocation per Launch Vehicle [34–36]
Unit
Mass
(kg)
Launch
Vehicle
Payload
Capability
Units
per
Launch
Launch
Cost
($M)
Electrolyzer
1290
FH-Par-Recov
4735
3
105
KRUSTY
2258
FH-Par-Recov
4735
2
105
NTP-530 (L-Surf)
5500
SLS Block 2
16601
3
1600
NTP-330 (L-Surf)
3700
FH-Par-Recov
4735
1
105
Chemical (L-Surf)
400
F9-Recov
1013
2
50
NTP-530 (LDHEO)
5500
New Glenn
8600
1
130
NTP-330 (LDHEO)
3700
F9-Recov
4000
1
50
Chemical (LDHEO)
400
F9-Recov
4000
2
50
Stages (LDHEO)
46800
SLS Block 2
46800
1
1600
####
Infrastructure Elements
####
Engine Replacements
####
Vehicle Assembly
V. COST ANALYSIS
Although the CLPA study considered the mining tents and purification/separation units
linearly scalable with a linear production cost after the developmental costs have been incurred
[2], other elements are assumed to follow the cost power functions of NAFCOM [6]. These
elements include engines, vehicles, KRUSTY, and electrolyzers. Therefore, when the
infrastructure grows and/or multiple vehicles are considered to perform simultaneous missions,
economies of scale can be realized and the total cost per mission can be expected to decrease with
an increasing number of missions.
The cost per mission for the Main Mission Vehicle is shown in Eq. 1. Here, is
the total cost of the infrastructure that supports the propellant production for the Main Mission
Vehicle only. This cost includes launch vehicles that ferry the infrastructure elements to the Lunar
surface. and are the total costs of the Main Mission Vehicle and engines including
the associated launch vehicle costs. is only applicable to vehicles parked in LDRO and
is the total cost to launch all the LADV vehicles to retank the parked Main Mission Vehicles
throughout the life of the architecture including all costs associated with the infrastructure required
to support the LADV vehicles. is the number of missions that the Main Mission Vehicle
provides throughout the life of the architecture.
Eq. 1
12
Similarly, Eq. 2 provides the cost per mission for the Supplemental Vehicle where the
same cost variables are used as in Eq. 1, but with the subscript MMV replaced by SV indicating
that these costs are now associated with the Supplemental Vehicle only. A key distinction between
Eq. 1 and Eq. 2 is that is absent from Eq. 2 since the Supplemental Vehicle uses only the
excess resources left-over from In-Situ propellant production for the Main Mission Vehicle.
Therefore, no additional infrastructure is required to support the Supplemental Vehicle except for
the additional costs incurred by using the LADV to retank the parked Supplemental Vehicle in
LDRO.
Eq. 2
The lower Isp of A-NTP engines results in more LADV launches for supporting the
Supplemental Vehicles than required for supporting the Main Mission Vehicles. However, this
does not necessarily mean that it is never advantageous to use A-NTP Supplemental Vehicles. The
breakeven point occurs when . Therefore, if , it is preferrable to include
a Supplemental Vehicle in the architecture from a cost standpoint, otherwise, it is more cost
effective to vent the excess resources and only use the Main Mission Vehicle to perform the
missions. If the Supplemental Vehicle is included in the architecture, then the final cost per mission
of the entire architecture is given by Eq. 3.
Eq. 3
It is important to note that although estimated values in Eq. 3 could be used, it is best to
incorporate economies of scale into these values as they change with respect to the total size of the
architecture. Furthermore, the cost per mission is not the only parameter to consider when deciding
on an architecture as other factors such as the total cost of the architecture and the total number of
missions could play a large role. If there is enough propellant to support a Supplemental Vehicle
that will perform the same mission as the Main Mission Vehicle, then the total number of missions
performed by the architecture will always increase. Determining an advantageous architecture will
require detailed architecture modeling and will be explored in future work.
VI. CONCLUSIONS
An abstracted analysis on the use of A-NTP engines on board vehicles within the NASA
Artemis Program context was performed. Previous work determined approximate performance
parameters that could be expected from A-NTP engines using water or ammonia as a propellant.
Based on the difference in Isp values, it was estimated that A-NTP vehicles will have larger initial
masses than the reference vehicles utilizing the reference engines with higher Isp which include the
RL-10 and H-NTP. Basing an infrastructure on A-NTP vehicles alone will result in more mining
tents due to more required propellant mass but will save on the requirement of having electrolyzers
produce hydrogen and oxygen for higher performance propulsion systems. Another option is to
augment an architecture without changing the Lunar surface infrastructure by introducing
13
Supplemental Vehicles which will utilize some of the excess resources such as ammonia left-over
from producing the propellant for the Main Mission Vehicle. The Supplemental Vehicle will
conduct missions alongside the Main Mission Vehicle to increase the total number of missions
performed by the architecture. However, the addition of the Supplemental Vehicle will only be
advantageous from a cost per mission perspective if the cost per mission of the Supplemental
Vehicle is less than the cost per mission of the Main Mission Vehicle. However, if there is enough
water or ammonia to serve as a propellant to support an A-NTP Supplemental Vehicle, then the
total number of missions will always increase. To determine if the inclusion of a Supplemental
Vehicle is overall advantageous, the cost per mission may not be the only parameter to consider
as the total cost of the infrastructure as well as the total number of missions performed by the
architecture could be equally important. Economies of scale could also have significant impacts
on the costs of the individual elements of the architecture. Future work will analyze the augmented
architectures in detail to determine which missions benefit from A-NTP Supplemental Vehicles
when economies of scale are considered.
ACKNOWLEDGMENTS
This work was supported by the National Aeronautics and Space Administration under
contract number MTS_UAH_021717.
REFERENCES
[1] Artemis Plan: NASA’s Lunar Exploration Program Overview. Publication NP-2020-05-
2853-HQ. NASA, 2020.
[2] Kornuta, David, Abbud-Madrid, Angel, Atkinson, Jared, Barr, Jonathan, Barnhard, Gary,
Bienhoff, Dallas, Blair, Brad, Clark, Vanessa, Cyrus, Justin, DeWitt, Blair, Dreyer, Chris,
Finger, Barry, Goff, Jonathan, Ho, Koki, Kelsey, Laura, Keravala, Jim, Kutter, Bernard,
Metzger, Philip, Montgomery, Laura, Morrison, Phillip, Neal, Clive, Otto, Erica, Roesler,
Gordon, Schier, Jim, Seifert, Brandon, Sowers, George, Spudis, Paul, Sandahl, Mark,
Zacny, Kris, and Zhu, Guangdong. “Commercial Lunar Propellant Architecture: A
Collaborative Study of Lunar Propellant Production.” REACH - Reviews in Human Space
Exploration, Vol. 13, No. 1, 2019, pp. 1–77. https://doi.org/10.1016/j.reach.2019.100026.
[3] Robbins, W. H., and Finger, H. B. A Historical Perspective of the NERVA Nuclear Rocket
Engine Technology Program. Publication 187154. North Olmsted, OH, 1991.
[4] Joyner, C. R., Eades, M., Horton, J., Jennings, T., Kokan, T., Levack, D. J. H., Muzek, B.
J., and Reynolds, C. LEU NTP Engine System Trades and Mission Options. Presented at
the Proceedings of the Nuclear and Emerging Technologies for Space, Richland, WA,
2019.
[5] Levack, D. J. H., Horton, J. F., Jennings, T. R., Joyner, C. R. I., Kokan, T., Mandel, J. L.,
Muzek, B. J., Reynolds, C., and Widman, F. W. J. Evolution of Low Enriched Uranium
Nuclear Thermal Propulsion Vehicle and Engine Design. Indianapolis, IN, 2019.
[6] McAfee, Julie, Colver, George, and Naderi, Mahmoud. NASA Air Force Cost Model
(NAFCOM): Capabilities and Results. Huntsville, AL.
[7] Zuppero, Anthony. To Inhabit the Solar System Using the Water in Space. 2009.
[8] Poston, D. I., Gibson, M. A., Godfroy, T., and McClure, P. R. “KRUSTY Reactor Design.”
Nuclear Technology, Vol. 206, No. sup1, 2020, pp. S13–S30.
https://doi.org/10.1080/00295450.2020.1725382.
14
[9] Joyner, C. R., Eades, M., Hanks, D., Horton, J., Jennings, T., Kokan, T., and Levack, D. J.
H. NTP Design Derivatives and Enhancements for Lunar and Mars Missions. Presented at
the AIAA Propulsion and Energy 2019 Forum, Indianapolis, IN, 2019.
[10] Reynolds, C. B., Joyner, C. R. I., Kokan, T., Levack, D. J. H., and Muzek, B. J. NTP
Robustness for Mars Conjunction and Opposition Class Missions. Presented at the
ASCEND, Virtual Event, 2020.
[11] Carabellese, D., Pino, P., Barbero, D., Rabagliati, L., Karnal, M., and Nambiar, S.
Preliminary Design of a Reusable Lunar Lander/Ascender for on-Orbit Refueling.
Presented at the AIAA Propulsion and Energy 2020 Forum, Virtual Event, 2020.
[12] Askins, Bruce, Childress-Thompson, Rhonda, and Thomas, L. Dale. Methodology for
Identifying Key Parameters Affecting Reusability for Liquid Rocket Engines. JANNAF
Propulsion Meeting, , 2019.
[13] Nikitaev, D., and Thomas, L. In-Situ Alternative Propellants for Nuclear Thermal
Propulsion. Presented at the AIAA Propulsion and Energy 2021 Forum, VIRTUAL
EVENT, 2021.
[14] Hinshelwood, Cyril Norman and Burk, Robert Emmett. “The Thermal Decomposition of
Ammonia Upon Various Surfaces.” Journal of the Chemical Society, Vol. 127, No. 1, 1925,
pp. 1105–1117. https://doi.org/10.1039/CT9252701105.
[15] Kunsman, C. H. “The Thermal Decomposition of Ammonia on Tungsten, Molybdenum,
and Nickel.” Journal of the American Chemical Society, Vol. 50, No. 8, 1928, pp. 2100–
2113. https://doi-org.elib.uah.edu/10.1021/ja01395a006.
[16] Shindo, Hitoshi, Egawa, Chikashi, Onishi, Takaharu, and Tamaru, Kenzi. “Reaction
Mechanism of Ammonia Decomposition on Tungsten.” Journal of the Chemical Society
Faraday Transactions 1, Vol. 76, No. 1, 1980, pp. 280–290.
https://doi.org/10.1039/f19807600280.
[17] Rashkeev, Sergey N., Glazoff, Michael V., and Tokuhiro Akira. “Ultra-High Temperature
Steam Corrosion of Complex Silicates for Nuclear Applications: A Computational Study.”
Journal of Nuclear Materials, Vol. 444, No. 1, 2014, pp. 56–64.
https://doi.org/10.1016/j.jnucmat.2013.09.021.
[18] Markham, G., Hall, R., and Feinroth, H. Recession of Silicon Carbide in Steam under
Nuclear Plant LOCA Conditions up to 1400 °C. In Ceramic Engineering and Science
Proceedings (K. Fox, Y. Katoh, H.-T. Lin, I. Belharouak, M. Halbig, and S. Mathur, eds.),
John Wiley & Sons, Inc., Hoboken, NJ, USA, 2012, pp. 111–120.
[19] Gustafson, Jeremy L., Witter, Jonathan K., and Houts, Michael. Discuss UAH NTP Fuel
Question. Apr 21, 2021.
[20] Koyanagi, T., Katoh, Y., Singh, G., and Snead, M. SiC Cladding Materials Properties
Handbook. https://info.ornl.gov/sites/publications/Files/Pub100714.pdf.
[21] Avincola, V. Angelici, Grosse, M., Stegmaier, U., Steinbrueck, M., and Seifert, H. J.
“Oxidation at High Temperatures in Steam Atmosphere and Quench Ofsilicon Carbide
Composites for Nuclear Application.” Nuclear Engineering and Design, Vol. 295, No. 1,
2015, pp. 468–478. https://doi.org/10.1016/j.nucengdes.2015.10.002.
[22] Cheng, T., and Tortorelli, P. F. “Silicon Carbide Oxidation in High-Pressure Steam.”
Journal of the American Ceramic Society, Vol. 96, No. 7, 2013, pp. 2330–2337.
https://doi.org/10.1111/jace.12328.
[23] Lee, Y., McKrell, T. J., Montecot, A., Pantano, M., Song, Y., and Kazimi, M. S. “Oxidation
Behavior of Sintered Tubular Silicon Carbide in Pure Steam Ι: Experiments.” Ceramics
15
International, Vol. 42, No. 1, 2016, pp. 1916–1925.
https://doi.org/10.1016/j.ceramint.2015.09.161.
[24] Zuppero, Anthony and Richins, William. Pump and Pressure Vessel Considerations for
Nuclear Heated Steam Rocket. Presented at the American Nuclear Society 1998 Annual
Meeting, Boston, MA.
[25] Wetmore, W. O. NERVA Engine Reference Data. Publication S130-CP090290-AF1.
Aerojet Nuclear Systems Company, 1970.
[26] Wetmore, W. O. Detail Specification Part 1 Performance/Design and Qualifications
Requirements for Engine, NERVA, 75K, Full Flow. Publication CP-90290A. Aerojet
Nuclear Systems Company, 1970.
[27] Special Metals. Inconel Alloy 625. https://www.specialmetals.com/documents/technical-
bulletins/inconel/inconel-alloy-625.pdf.
[28] Special Metals. Inconel Alloy 718. https://www.specialmetals.com/documents/technical-
bulletins/inconel/inconel-alloy-718.pdf.
[29] Special Metals. Nimonic Alloy 90. https://www.specialmetals.com/documents/technical-
bulletins/nimonic-alloy-90.pdf.
[30] Binder, Michael P. A Transient Model of the RL10A-3-3A Rocket Engine. Publication
195478. NYMA, Inc., Brook Park, OH, 1995.
[31] Lemmon, E. W., Bell, I. H., Huber, M. L., and McLinden, M. O. 10.0. NIST Standard
Reference Database 23: Reference Fluid Thermodynamic and Transport Properties-
REFPROP. National Institute of Standards and Technology.
[32] WebBook, N. I. S. T. C. NIST Standard Reference Database Number 69. National Institute
of Standards and Technology, 2018.
[33] Kokan, Timothy S. Mars Architecture Ground Rules and Assumptions.
[34] SpaceX. SpaceX Falcon 9 Data Sheet. SpaceX, 2017.
[35] New Glenn Payload User’s Guide.
https://yellowdragonblogdotcom.files.wordpress.com/2019/01/new_glenn_payload_users_g
uide_rev_c.pdf.
[36] Jackman, A. L. Space Launch System: SLS Transportation Beyond LEO. Washington D.C.,
Mar 01, 2017.
