Content uploaded by James Horton
Author content
All content in this area was uploaded by James Horton on Oct 15, 2020
Content may be subject to copyright.
Nuclear and Emerging Technologies for Space
Knoxville, TN, April 6 – April 9, 2020, available online at https://nets2020.ornl.gov
LEU NTP FLIGHT DEMONSTRATION VEHICLE AND APPLICATIONS TO OPERATIONAL MISSIONS
Timothy Kokan1, James Horton2, C. Russell Joyner II3, Daniel J.H. Levack2, Brian J. Muzek1, and Christopher B. Reynolds1
1Aerojet Rocketdyne, Huntsville, Alabama 35806, USA,
2Aerojet Rocketdyne, Canoga Park, California 91309, USA,
3Aerojet Rocketdyne, West Palm Beach, Florida 33410, USA,
256-922-2579; timothy.kokan@rocket.com
Nuclear thermal propulsion (NTP) has been
extensively researched as a potential main propulsion
option for human Mars missions. NTP’s combination of
high thrust and high fuel efficiency makes it an ideal main
propulsion candidate for these types of missions,
providing architectural benefits including smaller
transportation system masses, reduced trip times,
increased abort capabilities, and the potential for
transportation infrastructure reuse.
Since 2016, AR has been working with NASA and
members of industry as part of the NASA Space
Technology Mission Directorate Game Changing
Development Nuclear Thermal Propulsion Project. The
overall goal of this project is to determine the feasibility
and affordability of a low enriched uranium (LEU)-based
NTP engine with solid cost and schedule confidence.
Having shown feasibility and affordability, program
planning has been underway for follow-on activities to
continue to mature the LEU NTP engine technology.
These activities include program planning for reactor
fuels testing, reactor component design, engine
component technology development, test facility design
and demonstration, and a demonstration engine available
for ground test and potentially flight test. These follow-on
activities would set the stage for full scale development of
a human rated NTP flight engine for use in human
exploration missions.
This paper presents details of a potential LEU NTP
prototype flight test and corresponding first flight vehicle
along with potential applications of an evolved vehicle for
subsequent operational missions.
NOMENCLATURE
AR = Aerojet Rocketdyne
CFM = Cryogenic Fluid Management
CLV = Commercial Launch Vehicle
DoD = Department of Defense
E-M = Earth-Moon
L1 = First Lagrange Point
GCD = Game Changing Development
Isp = Specific Impulse
LEO = Low Earth Orbit
LEU = Low Enriched Uranium
LH2 = Liquid Hydrogen
MEO = Medium Earth Orbit
MLI = Multilayer Insulation
MMOD= Micro-meteoroid Orbital Debris
MSFC = Marshall Space Flight Center
NASA = National Aeronautics and Space Administration
NTP = Nuclear Thermal Propulsion
RAAN = Right Ascension of the Ascending Node
RCS = Reaction Control System
SLS = Space Launch System
SOFI = Spray-on Foam Insulation
STMD = Space Technology Mission Directorate
TDRS = Tracking and Data Relay Satellite
ULA = United Launch Alliance
I. INTRODUCTION
Since 2016, AR has been working with NASA, the
Department of Energy, and members of industry as part of
the NASA Space Technology Mission Directorate
(STMD) Game Changing Development (GCD) Nuclear
Thermal Propulsion (NTP) Project. The overall goal of
this project is to determine the feasibility and affordability
of a low enriched uranium (LEU)-based NTP engine with
solid cost and schedule confidence.
Having shown feasibility and affordability, program
planning has been underway for follow-on activities to
continue to mature the LEU NTP engine technology.
These activities include program planning for:
1. Initial NTP engine system technology
development including reactor fuels testing,
reactor component design, engine component
technology development, test facility design and
demonstration;
2. Prototype NTP engine development including
potential testing either on the ground or in flight;
3. Human rated NTP flight engine full scale
development for the full scale flight engine for
human exploration mission.
As seen in Figure 1 below, the prototype NTP engine
development and testing provides a path, along with the
initial NTP engine system technology development
activities, to a human rated NTP flight engine system.
2
NASA STMD
GCD NTP
Project
Initial NTP
Engine System
Technology
Development
Prototype NTP
Engine
Development
and System
Testing
Human Rated
NTP Flight
Engine Full Scale
Development
Fig. 1. Prototype NTP engine development and system
testing, either on the ground or in flight, provides a path
from STMD GCD NTP feasibility assessment to NTP
flight engine full scale development
AR is currently performing preliminary definitions of
potential prototype NTP engine flight test options that can
reduce technical risk for the larger human rated NTP
flight engine full scale development. The following
sections will discuss these flight test options along with
potential options for evolved vehicles based on the flight
test vehicle to perform operational missions.
II. PROTOTYPE ENGINE FLIGHT TEST OPTIONS
In 2019, AR started examining various approaches
for prototype NTP engine flight test vehicles that would
have lineage to a NTP-based human Mars architecture.
An initial screening of flight test mission concepts
examined missions that provide information on NTP
operational verification, demonstration of integrated
cryogenic systems versus non-cryogenic systems, NTP
integration with a cryogenic stage similar to the Mars
vehicle, packaging capability for launch on a commercial
launch vehicle (CLV), and many other attributes.
Based on these initial screening activities, the best
NTP and stage flight test approach appears to be one that
achieves the following goals:
1. Have drop-off orbit that provides safety -
independent of prototype NTP engine operation;
2. Demonstrate operation of a NTP engine (reactor)
in space: Perform multiple burn sequences (start-
up, main stage, shutdown, cooldown) with burn
times to demonstrate NTP capability;
3. Demonstrate processes for a safe launch and
operation of a nuclear reactor into space via
commercial launch similar to Department of
Defense launches;
4. Demonstrate passive cryogenic fluid
management (CFM) for an extended period of
time applicable to Lunar and Mars missions and
apply data to design of robust passive/active
CFM technologies;
5. Demonstrate launch of a cryogenic stage in the
payload fairing of a launch vehicle (Figure 2).
4-5 meter
Diameter
NTP Stage
Launched
on a CLV
Fig. 2. A prototype NTP flight test vehicle can be sized to
fit on existing or near-term future CLVs
Many potential flight test mission options can satisfy
these goals, including low Earth orbit (LEO) plane
changes (either changes in inclination or right ascension
of the ascending node (RAAN)), LEO-to-medium Earth
orbit (MEO) altitude changes, and LEO-to-Earth-Moon
(E-M) L1 transfers.
The LEO plane change demonstration mission
(Figure 3) was selected for further study because it has
several operational advantages:
1. Flexibility in the final orbit allowing for shorter
or longer than nominal NTP burn times;
2. Continuous nuclear-safe orbit throughout the
mission;
3. Altitude is kept below the global continuous
coverage provided by the Tracking and Data
Relay Satellite (TDRS) system.
Fig. 3. LEO plane change flight test mission
Inclination Change or Shift in RAAN
3
III. LEO PLANE CHANGE MISSION
PARAMETRIC TRADES
Parametric trades of different prototype flight vehicle
sizes and different prototype NTP engine thrust levels are
provided in Section III.
A primary flight test mission goal is to demonstrate
NTP operability over several main engine burns. In order
to achieve this goal, the LEO plane change mission is
envisioned to consist of two burns, each with a minimum
burn time of six minutes (30 second startup, minimum of
5 minute main stage, 30 second shutdown). This results in
the need for a stage large enough to permit up to 10
minutes of NTP main stage burn time.
A nuclear safe LEO starting orbit of 2,000 km x
2,000 km x 25° is selected. This orbit is advantageous for
several reasons, including:
1. Low orbital debris spatial density;
2. Negligible atmospheric drag;
3. Continuous tracking and data relay coverage
provided by TDRS.
Figure 4 provides a sensitivity of burn time to stage
gross mass for two different prototype NTP engine thrust
levels (12.5 klbf and 15.0 klbf) with different launch
vehicle class capabilities called out. A prototype NTP
engine flight test vehicle with a 12.5 klbf NTP engine and
a 20 mT vehicle gross mass is highlighted with the blue
star as it provides sufficient total NTP main stage burn
time and can be launched on the Delta IV Heavy or a
future medium CLV such as Vulcan or Omega.
0
2
4
6
8
10
12
14
16
18
20
22
15 16 17 18 19 20 21 22 23 24 25
NTP Total Burn Time (min)
Vehicle Gross Mass (mT)
NTP Demonstrator Vehicle
Gross Mass Sensitivity
12.5k NTP Engine
15k NTP Engine
Current
Atlas V
Capability
Future
Medium CLV
Capability
Future
Heavy
CLV
Capability
Fig. 4. Several launch vehicle options are capable of
launching a prototype NTP engine flight test vehicle sized
to achieve the prototype engine total burn time goal of
>10 min
Figure 5 shows the orbital changes envisioned for the
prototype NTP engine flight test mission. The mission
details are for an example 12.5 klbf prototype NTP engine
thrust and an initial test flight vehicle gross mass of 20
mT. This engine and vehicle size combination allows for
over 12 minutes of main stage burn time.
Burn #1 is envisioned to operate at a lower reactor
temperature, providing additional reactor temperature
margin for the first use of the reactor in space, resulting in
an initial Isp of 800 seconds. Burn #2 is then envisioned
to operate at the 2700 K nominal reactor operating
temperature, resulting in an Isp of 900 seconds.
Initial
Orbit
Inter-
mediate
Orbit
Final
Orbit
Burn #1
Thrust: 12,500 lbf
Isp: 800 sec
Main Stage Burn: 6.1 min
RAAN Shift: 24.5°
Burn #2
Thrust: 12,500 lbf
Isp: 900 sec
Main Stage Burn: 6.1 min
RAAN Shift: 30.1°
Fig. 5. Two-burn prototype NTP engine flight test
mission achieves key demonstration goals (multiple burn
sequences, >10 minutes of NTP main stage burn time)
while staying within a nuclear safe orbit
4
Figure 6 provides the envisioned concept of
operations for the example 12.5 klbf / 20 mT prototype
NTP flight test vehicle mission shown in Figure 5. The
near-24-hour mission consists of an initial checkout of
approximately 6 hrs, a first 6-minute main stage burn, a 6-
hr coast / cooldown, a second 6-minute main stage burn, a
second 6-hr coast / cooldown, and a final approximate 6-
hr for mission closeout and stage safing and monitoring.
12.5k NTP Engine / 20 mT Gross Mass Vehicle
TInitial
(hr)
T
(hr)
TFinal
(hr)
Launch to 2,000 km
circ @ 25 deg
0.0
0.5
0.5
Spacecraft Checkout
(3 orbits)
0.5
6.0
6.5
First Burn (Startup /
Main Stage /
Shutdown)
6.5
0.1
6.6
Coast / Engine
Cooldown (3 orbits)
6.6
6
12.6
Second Burn (Startup /
Main Stage /
Shutdown)
12.6
0.1
12.7
Engine Cooldown /
Checkout (3 orbits)
12.7
6
18.7
Mission Closeout /
Monitoring
18.7
6
24.7
Fig. 6. Near-24-hour prototype NTP engine flight test
vehicle mission duration is sufficient to achieve mission
goals
Figure 7 shows the example 12.5 klbf / 20 mT
prototype NTP flight test vehicle within the United
Launch Alliance (ULA) Vulcan launch vehicle 5.4m
diameter payload fairing. The NTP test vehicle can be
sufficiently sized with enough liquid hydrogen (LH2) to
permit at least 10 minutes of NTP main stage burn time
while still fitting within the dynamic envelope of the
Vulcan payload fairing.
14.7 m
Fig. 7. Prototype NTP flight test vehicle sized to fit in the
ULA Vulcan 5.4m diameter payload fairing
Figure 8 provides a summary mass roll-up for the
12.5 klbf / 20 mT NTP flight test vehicle. The flight test
vehicle is envisioned to leverage CLV upper stage
subsystems to the greatest extent possible, including: LH2
tank and primary structures, storable reaction control
systems (RCS), batteries, command and data handling,
guidance, navigation, and control, communications,
passive CFM (spray-on foam insulation (SOFI),
multilayer insulation (MLI)), and micro-meteoroid orbital
debris (MMOD) shielding.
Subsystem
Predicted
Mass (kg)
1.0 Structures 3,018
2.0 Propulsion 5,625
MPS 5,511
RCS/OMS 114
3.0 Power 252
4.0 Avionics 405
5.0 Thermal (SOFI, MLI, MMOD) 685
Dry Mass 9,986
6.0 Non-Propellant Fluids 450
Inert Mass 10,436
7.1 MPS Usable Propellant 5,935
7.2 RCS Usable Propellant 180
Gross Mass 16,550
Payload 1,000
LV Payload Attach Fitting 500
LV Payload Margin 1,950
LV Payload System Mass 20,000
Fig. 8. Summary mass roll-up of example NTP flight test
vehicle sized to launch on a ULA Vulcan launch vehicle
5
IV. EVOLVED VEHICLE OPTIONS FOR
OPERATIONAL MISSIONS
In addition to providing risk reduction for a full scale
human rated NTP flight vehicle, the prototype NTP flight
test vehicle can also provide an initial starting point for an
evolved operational stage (Figure 9).
Prototype NTP
Flight Test Vehicle
NTP Operational Stage
Prototype Flight Test
Vehicle Leading to
Operational Stage
Fig. 9. A prototype NTP flight test vehicle can be evolved
for use on future operational missions1
Missions such as outer planetary science, Cislunar
cargo delivery, and Earth orbit altitude / plane changes
could potentially benefit from an operational NTP in-
space propulsive stage.
Example outer planetary science mission trade results
are provided in Figures 10 and 11 for Jupiter and Uranus
deep space science missions2. This NTP operational stage
is sized to fit, along with a payload, within the Long SLS
8.4m diameter payload fairing. This stage has an
estimated gross mass of 31 mT, approximately 50% larger
than the NTP flight demonstrator vehicle discussed in
Section III.
Unlike the NTP flight test vehicle, the NTP
operational stage would require active CFM utilizing
cryocoolers to maintain the LH2 in a liquid state for the
duration of the potentially multi-year mission.
Fig. 10. Deep Space NTP Operational Stage for Jupiter Orbiter Missions using NASA SLS Block 2 Launch Vehicle
6
Fig. 11. Deep Space NTP Operational Stage for Uranus Orbiter Missions using NASA SLS Block 2 Launch Vehicle
V. CONCLUSIONS
Prototype NTP flight test vehicle options with
applications to operational missions were defined. Flight
test missions were identified that allow for the safe testing
of the NTP flight test vehicle in space, demonstrate NTP
engine operation in space with multiple engine burns with
sufficient burn times to demonstrate main stage
capability, provide risk reduction on NTP and stage CFM
systems, and launch on CLV’s. Furthermore, operational
missions were identified that utilize an evolved NTP in-
space propulsion stage to provide significant mission
benefit. Examples were provided for outer planetary
science missions to both Jupiter and Uranus.
ACKNOWLEDGMENTS
The authors would like to acknowledge the support
of NASA Marshall Space Flight Center and Glenn
Research Center, NASA Space Technology Mission
Directorate (STMD) Game Changing Development
(GCD) and the Department of Energy engineers that
continue working NTP for Mars crewed missions and
other missions that can benefit from NTP.
REFERENCES
1. Levack, D.J.H., Horton, J.F., Jennings, T.R., Joyner,
C.R., Kokan, T., Mandel, J.L., Muzek, B.J.,
Reynolds, C., and Widman, F.W., “Evolution of Low
Enriched Uranium Nuclear Thermal Propulsion
Vehicle and Engine Design”, AIAA 2019-3943,
AIAA Propulsion and Energy Forum and Exposition,
Indianapolis, Indiana, August 19-22, 2019.
2. Joyner, C.R., Eades, M., Horton, J., Jennings, T.,
Kokan, T., Levack, D.J.H., Muzek, B.J., and
Reynolds, C.B., “LEU NTP Engine System Trades
and Missions”, Nuclear and Emerging Technologies
for Space, Richland, Washington, February 25-28,
2019.
