SP2016_AP4

Foundational Methane Propulsion Related Technology Efforts, and Challenges for

Applications to Human Exploration Beyond Earth Orbit
Space Propulsion 2016
Marriott Park Hotel, Rome, Italy / 2-6 May 2016

Thomas Brown (1), Mark Klem (2), Patrick McRight (3)

(1)
National Aeronautics & Space Administration, NASA Engineering and Safety Center, NASA George C.
Marshall Space Flight Center, Huntsville, Alabama 35812. USA thomas.m.brown@nasa.gov
(2)
National Aeronautics & Space Administration, Propulsion Division, NASA Glenn Research Center,
Cleveland, Ohio, 44135. USA mark.d.klem@nasa.gov
(3)
National Aeronautics & Space Administration, Propulsion Department, NASA Marshall Space Flight
Center, Huntsville, AL 35812. USA pat.mcright@nasa.gov

KEYWORD: Methane Propulsion, Cryogenic Fluid studies have identified the potential for commonality
Management, Cryogenic Space Propulsion between interplanetary and descent/ascent
propulsion solutions using liquid methane (LCH4)
ABSTRACT: and liquid oxygen (LOX) propellants. These
Current interest in human exploration beyond earth commonalities may lead to reduced overall
orbit is driving requirements for high performance, development costs and more affordable exploration
long duration space transportation capabilities. architectures.
Continued advancement in photovoltaic power With this increased interest, it is critical to understand
systems and investments in high performance the current state of LOX/LCH4 propulsion technology
electric propulsion promise to enable solar electric and the remaining challenges to its application to
options for cargo delivery and pre-deployment of beyond earth orbit human exploration. This paper
operational architecture elements. However, higher provides a survey of NASA’s past and current
thrust options are required for human in-space methane propulsion related technology efforts,
transportation as well as planetary descent and assesses the accomplishments to date, and
ascent functions. examines the remaining risks associated with full
While high thrust requirements for interplanetary scale development.
transportation may be provided by chemical or 1. INTRODUCTION:
nuclear thermal propulsion systems, planetary
descent and ascent systems are limited to chemical Human, beyond-earth-orbit, exploration architecture
solutions due to their higher thrust to weight and studies have identified LOX/LCH4 as a strong
potential planetary protection concerns. Liquid candidate for both interplanetary and descent/ascent
hydrogen fueled systems provide high specific propulsion solutions. While methane fuel has yet to
impulse, but pose challenges due to low propellant be implemented in such in-space flight systems,
density and the thermal issues of long term significant research efforts have been conducted for
propellant storage. Liquid methane fueled over 50 years, ranging from fundamental combustion
propulsion is a promising compromise with lower and mixing efforts to rocket chamber and system-
specific impulse, higher bulk propellant density and level demonstrations. In addition, over the past 15
compatibility with proposed in-situ propellant years, NASA and its partners have built upon these
production concepts. Additionally, some architecture early research activities, conducting many advanced
development efforts that have demonstrated the The following sections will review/summarize recent
practical components and sub-systems needed to NASA, LOX/LCH4 advanced development efforts,
field future methane space transportation elements consider remaining risks to develop future flight
(e.g. thrusters, main engines, and propellant storage systems, and make some general recommendations
and distribution systems). Relevant advanced for a path forward.
development efforts began with a push to field non-
toxic Orbital Maneuvering System (OMS) and 2. LOX/LCH4 IGNITERS
Reaction Control Systems (RCS) for NASA’s Space Relative to more conventional, hypergolic storable
Shuttle System. Early Non-Toxic RCS efforts did not solutions, one of the largest risks associated with
utilize methane fuel. However, these demonstrations LOX/LCH4 propulsion is reliable ignition. In the 2005
are applicable from the common challenges of – 2010 timeframe, the NASA Propulsion and
cryogenic propellants for on demand systems. Cryogenics Advanced Development Project (PCAD)
Likewise some earlier pump-fed throttleable lander conducted numerous in-house experimental efforts
engine efforts used liquid hydrogen (LH2) fuel, but to examine the issue [1, 2]. The work was completed
are applicable from a cryogenic propellant and at both Reaction Control Engine (RCE) and larger
throttle control/stability perspective. main engine scales. The majority of the work was
These related efforts and a significant number of conducted with spark torch igniters. However, there
direct methane propulsion demonstration activities were also successful demonstrations of microwave
have formed a foundation of LOX/LCH4 (and related) torch ignition, and a combination spark torch/glow
propulsion knowledge that has significantly reduced plug igniter.
the development risks of future methane based Overall there were no significant issues identified
space transportation elements for human exploration that would prohibit the reliable ignition over a range
beyond earth orbit. of conditions with LOX/LCH4. One of the last ignition
While LOX/LCH4 propulsion has been identified as a specific activities completed was the demonstration
potential solution for multiple transportation of 30,000 ignition cycles on a spark torch ignition
functions, some architecture efforts have identified system at vacuum conditions. Completion of this
the potential for commonality between interplanetary activity did not identify any issues with the hardware
and descent/ascent propulsion solutions using or designs for long duration applications. The work
LOX/LCH4 propellants (common approaches could did, however, identify issues with spark plug
reduce development costs). These architecture durability and the reliability of power exciter units.
efforts have generally indicated needs for the Figs. 1 through 4 present examples of the igniters
following propulsion subsystem and components that were demonstrated and evaluated.
capabilities:

o RCS Propulsion: ~ 25-lbf – 100-lbf class

o Pressure-fed main engine: ~ 6000-lbf class
o Pump-fed (throttleable) main engine:
~ 25,000-lbf class
o Long Duration Cryogenic Fluid Management
and Distribution (CFM&D), including:
 High performance pressurization
systems
 Thermal management with high
performance multilayer insulation and
Figure 1. Main Engine Class Spark Torch Igniter
90K class cryocooler systems integrated
(Firing)
with CFM&D
 Management of propellant losses due to
boiloff and component leakage
 alternatives to more conventional hypergolic storable
OMS/RCS were initiated. The primary focus was
reduction in ground processing costs due to
simplified operations. These efforts are applicable to
LOX/LCH4 propulsion due to the common challenges
related to cryogenic propellants, and because some
of the hardware was later transitioned to perform the
early PCAD LOX/LCH4 RCS demonstrations.

Two non-toxic RCS efforts were conducted. Aerojet

Figure 2. RCE Class spark torch igniter (Firing) developed and demonstrated a dual thrust (25-lbf
and 870-lbf) LOX/Ethanol RCE [3]. This thruster was
successfully demonstrated at both thrust levels in
pulsed and steady state modes.

Figure 3. Dual, Diverse Ignition Torch (Firing)

Figure 5. Aerojet Non-Toxic LOX/Ethanol RCE

Figure 4. Microwave Ignition Torch

Many of the remaining ignition associated risks are

related to specific requirements and duty cycles that
will be imposed on future systems and conduct of
Figure 6. TRW Non-Toxic LOX/LH2 RCE
final spaceflight qualification. One general area that
still requires investigation is ignition in the cold TRW conducted two non-toxic RCS demonstration
thermal environment of space where both the efforts. One demonstrated a dual thrust (25-lbf and
hardware and propellants have been exposed for a 870-lbf) LOX/Ethanol RCE, while the second effort
significant period of time prior to required operation. focused on a 1000-lbf LOX/LH2 RCS Thruster [4].
Both designs were successfully demonstrated in hot
fire tests.
3. REACTION CONTROL SYSTEM THRUSTERS
Later NASA shifted focus from reusable launch
During NASAs 2nd Generation Reusable Launch technologies to advanced chemical propulsion for
Vehicle/Next Gen Launch Technology Program – space exploration. The PCAD project focused the
Auxiliary Propulsion Project (2000-2004), advanced top three risks identified for RCE technology: 1)
development efforts focused on non-toxic Ignition reliability; 2) Performance (vacuum specific
impulse (Isp)); and 3) Pulse width repeatability. To level to examine engine performance, and the design
address the risks, PCAD undertook a combination of used a spark torch igniter. Ultimately, all key
in-house and contract activities. performance criteria were demonstrated using an
impinging injector design. Aerojet conducted over
In 2006, the PCAD project awarded RCE contracts 1300 engine pulse tests at a variety of duty cycles
to Aerojet and Northrop Grumman (previous TRW and accumulated more than 1900 seconds of
propulsion group). Each contract focused on the operating time during sea-level, engine development
development and delivery of a 100-lbf thrust pre- testing. Aerojet met the 317-sec Isp requirement,
prototype engine subsystem. The key performance calculated based on estimated nozzle losses and
requirements were: 1) 317-second vacuum Isp; exceeded the 80-msec EPW requirement by
2) 4-lbfsec minimum impulse bit (Ibit); 3) 80-ms demonstrating 40-msec EPW. Aerojet provided 5
electrical pulse width (EPW); 4) 25,000 valve cycles engines to NASA that were subsequently tested in a
and 5) ignition and operation over a range of inlet multiple engine configurations on the Auxiliary
conditions including liquid and gaseous propellants. Propulsion System Test Bed (APSTB) and 2 units for
The two suppliers pursued different engine concepts testing at the thruster level in NASA’s Altitude
in response to these requirements. Combustion Stand (ACS).
The Aerojet concept was based on the earlier NASA conducted sea-level and altitude performance
LOX/Ethanol engine development and other testing, including a total of 60 altitude hot-fire tests
internally funded activities. Initial testing was with the Aerojet 100-lbf LOX/LCH4 engine over a
performed with 870-lbf engines that were originally wide range of propellant inlet conditions (pressure
designed to operate on LOX/Ethanol propellants and and temperature), to simulate operation in a variety
were modified to accommodate LOX/LCH4. NASA of space environments. Testing was conducted using
successfully tested these modified units at altitude a 45:1 area ratio columbium radiation cooled nozzle.
with the results influencing the 100-lbf engine design.
The main goal of the testing was to develop Isp
performance curves as a function of mixture ratio.
The engine demonstrated that meeting the required
317-sec performance is feasible for the 80:1 nozzle
based on the results with a 45:1 nozzle.

The Northrop Grumman concept was primarily based

on previous work on hypergolic propellant engines.
The combustion chamber and a portion of the nozzle
were regeneratively cooled with both oxygen and
methane. The full engine area ratio (120:1) was
completed with a columbium radiation-cooled nozzle
Figure 7. Aerojet 100-lbf LOX/LCH4 RCE extension. Propellant flow to both the main chamber
The Aerojet 100-lbf RCE consisted of compact and igniter was controlled by a single set of
integral exciter/spark plug system, a dual coil direct- independent single coil fuel and oxidizer valves.
acting solenoid valve for oxidizer and fuel, an integral Ignition was accomplished with the use of a spark
igniter and injector, and a columbium torch igniter. A series of hardware configurations
chamber/nozzle with an expansion area ratio of 80:1 were tested, starting with workhorse hardware, to
(See Fig. 7). develop the engine cooling circuit. Northrop
Grumman developed a single pre-prototype unit that
Propellant flow to both the main chamber and igniter was tested in vacuum conditions at their Capistrano
were controlled by a single set of dual coil valves. test facility (See Fig. 8).
Over 55,000 cycles were demonstrated at cryogenic
temperatures, exceeding the specified 25,000 cycle
life. A series injector concepts were tested at sea
 injectors with multiple combustion chamber lengths.
Testing demonstrated C* efficiencies over 98%. A
water cooled combustion chamber was used to
collect heat transfer data. Different length chambers
were used to obtain performance level correlations to
chamber length.

The chambers were also instrumented to collect

combustion stability data (both for direct injector
design evaluation, and anchoring analytical models).

Figure 8. Northrop Grumman Pre-Prototype 100-lbf 5. PRESSURE-FED MAIN ENGINE EFFORTS

RCE In 2006 NASA funded ATK and KT Engineering
Test results indicated that the engine concept was (KTE) to conduct LOX/LCH4 main engine workhorse
able to meet the performance goals, including demonstration efforts. Each contract was focused on
exceeding the Isp requirement. The measured Isp the development and delivery of a 7,500-lbf thrust
was approximately 331 sec, which exceeded the pre-prototype engine. The key performance targets
demonstration requirement of 317 sec. for the activity were: 1) 7,500-lbf thrust, 355-sec
vacuum Isp; 2) 90% rated thrust within 0.5 seconds;
3) total of 24 restarts; and 5) operation over a range
4. MAIN ENGINE INJECTOR PARAMETRICS of inlet conditions from gas to liquid for start. The
companies design solutions varied significantly with
In parallel to the contracted efforts, NASA conducted one pursuing a regenerative cooling approach and
in-house development of larger scale LOX/LCH4 the other implementing an ablative design. ATK
injectors [5]. Tests were conducted on both 2-inch teamed with XCOR to develop a pressure-fed engine
diameter and 6-inch diameter chambers at NASA. concept that was regeneratively cooled by the
methane fuel. Sea-level testing was conducted with
both water and methane cooled combustion
chambers at XCOR facilities in Mojave, CA [6].

Figure 9. Impinging Injector Tests

Figure 11. ATK/XCOR Engine Testing at Mojave,

CA.

KT Engineering pursued an ablative combustion

chamber design. A number of sea-level tests were
conducted at NASA on this workhorse design as
Figure 10. Coaxial and Swirl Coaxial Injector Tests well. Unfortunately, shifting technology demon-
stration requirements (toward Lunar Lander
This effort investigated performance and stability applications) resulted in the ATK/XCOR, and KT
characteristics of impinging, coaxial and swirl coaxial Engineering contract options not being exercised.
In response to the evolving technology of 150:1, a specific impulse of approximately 348 sec
demonstration requirements, NASA funded Aerojet could be achieved, which was within 2% of the
to develop a vacuum workhorse engine performance goal.
demonstrator [7]. This effort focused on demon-
strating the following requirements 1) 5,500-lbf thrust, More recent in-house efforts at NASA are currently
355-sec vacuum Isp; 2) 90% rated thrust within 0.5 pursuing additively manufactured (3D printed)
seconds; 3) total of 24 restarts; and 5) operation over regenerative cooled LOX/LCH4 engine concepts [5].
a range of inlet conditions from gas to liquid for start. Demonstration hardware includes a 3D printed
The Aerojet design included an ablative chamber Inconel injector, with a separate porous faceplate.
and liquid oxygen/liquid methane injection system. The injector also includes variable fuel film cooling
The overall activity was broken into two phases. The and a center igniter port. The regeneratively cooled
first phase involved Aerojet fabrication and sea-level chamber includes fully printed coolant channels,
testing of multiple injector designs. In the second eliminating the need for a separate liner/jacket joint.
phase, NASA took delivery of the engines and The chamber design also includes printed
conducted altitude performance testing. Testing at thermocouple ports along one coolant channel
NASA proceeded with the first injector produced (Fig. 13). Work is also underway to evaluate a
under the Aerojet contract. While sea-level GRCop-84 (Copper) printed unit.
performance was lower than desired, altitude testing
was conducted to correlate the sea-level and altitude
results and to validate nozzle performance analysis.

Figure 13. Additively Manufactured LOX/LCH4

Pressure-Fed Main Engine

Initial hot fire testing has verified injector stability, and

has successfully demonstrated the 3D printed
concept (Fig. 13). Testing also provided detailed
regenerative cooling data for a 2-phase thermal
model (critical for future pressure-fed, regenerative
engine development.

6. PUMP FED MAIN ENGINE EFFORTS

Figure 12. Aerojet 5500-lbf LOX/LCH4 Main Engine
The NASA Propulsion Cryogenics & Advanced
Demonstration Testing
Development (PCAD) Project also conducted both
Testing was conducted with an 8-inch long ablative contracted and in-house efforts related to deeply
combustion chamber and a radiation cooled throttleable pump-fed main engines. These efforts
columbium Space Shuttle OMS engine nozzle (conducted between 2005 and 2010) were focused
extension, which provided an area ratio of 129:1. on demonstrating technologies for lunar lander
Design area ratio for the prototype engine design descent stage applications, and all efforts utilized
was 150:1. A total of 187 seconds of run time was LOX/LH2 propellant combinations. Future Mars
accumulated on the engine including seven 20- transfer stage and Mars lander/ascent vehicle
second tests and one 40-second test. The injector, applications require LOX/Methane propellant
chamber and nozzle were all in good physical combinations. However, the PCAD efforts are
condition after the testing. The average vacuum Isp relevant due to lessons learned related to deep
calculated for the test program was 344 sec and the throttle injector stability, pump performance and
maximum was 345 sec. Extrapolating to an area ratio system response of cryogenic engine systems.
PCAD funded Pratt and Whitney Rocketdyne to
demonstrate the Common Extensible Cryogenic
Engine (CECE) [1]. This demonstrator utilized a
modified RL10 engine. Design changes included
injector modifications, valve modifications, and
system trim adjustments.

Figure 15. Northrop Grumman Throttling LOX/LH2

Pintle injector

A second variable geometry injector was designed

in-house at NASA [1]. This two-stage injector utilized
separate injector manifolds to enable a transition
between two fixed injector geometries. Unlike the
Northrop Grumman pintle design, the two stage
injector is not continuously variable, but is able to
shift between a lower flow resistance, high power
geometry and more resistive low power geometry.
Like the pintle the two stage design also enables
greater system-level performance by reducing
injector pressure drop at high power. The two-stage
injector was also successfully demonstrated
Figure 14. Pratt and Whitney Rocketdyne Common
throughout its throttle range, in injector/chamber,
Extensible Cryogenic Engine (CECE)
sea-level testing at NASA (Fig. 16).
The CECE effort successfully demonstrated stable
throttling (> 10:1), and met overall performance goals
(448 sec at 100% Power, 436 sec at low power).
Testing also demonstrated reliable ignition over 20
engine starts.

While the CECE effort utilized a fixed injector (with

increased pressure drop, enabling deep throttling),
variable geometry injector concepts were also
investigated. Variable injector geometry concepts Figure 16. Two Stage Throttleable Injector Testing
maintain stability margins at low power levels,
without large increases in injector pressure drop at More recent in-house pump-fed engine efforts (2012
high power levels, resulting in higher overall system – Current) at NASA have focused on an Additive
performance. PCAD funded Northrop Grumman’s Manufacturing Demonstration (AMD) Engine [8].
efforts to develop a throttling LOX/LH2 pintle injector This effort demonstrated the ability to utilize additive
[1]. The throttling pintle injector (Fig. 15) is manufacturing to greatly reduce development time
continuously adjustable throughout the throttle and production costs of a 30-Klbf-class LOX/LH2
range, and was successfully demonstrated in engine. This activity produced the majority of the
injector/chamber sea-level testing at NASA. engine system components (including rotating turbo-
machinery parts) through additive/3D printing
techniques. The projects’ Integrated AMD
breadboard system testing demonstrated multiple
components simultaneously, in relevant environ-
ments for relatively low costs
 disposal (Fig. 19). However, development of other
LOX/LCH4 engine system components are under-
way, and system-level test-bed demonstrations are
planned for the near future.

7. CRYOGENIC FLUID MANAGEMENT AND

DISTRIBUTION

Since the primary application for cryogenic

propulsion systems has been launch vehicle
boosters and upper stages, LOX/LCH4 propulsion
systems for missions beyond earth orbit must
overcome challenges that traditional cryogenic
propulsion systems have not yet encountered. For
example, while pre-launch thermal loads and ascent
Figure 17. Additive Manufacturing Demonstration heating have dominated the design requirements for
(AMD) Engine at NASA traditional cryogenic propulsion systems, farther-
reaching missions employing cryogenics must also
Due to the technology pull from future Mars withstand post-ascent thermal loads during venting
exploration missions, the AMD test bed concept is and equilibration of multi-layer insulation (MLI) and
being transitioned to demonstrate LOX/LCH4 engine deal with subsequent effects that will impact system
components and systems operation. In March, 2016 performance, such as radiative heating from
the NASA team successfully demonstrated an planetary bodies, solar heating, and microgravity
additively manufactured LOX/LCH4 Turbo-Pump fluid behavior and its impacts on thermal
stratification, heat transfer, and pressures within the
propellant tanks. Tank mounting schemes and tank
penetrations will also play a much more dominant
role in system performance, as total heat leak into
cryogenic tanks becomes far more impactful for
longer mission durations.

7.1. Passive CFM capabilities

Previous NASA work has shown that cooling LOX

and LCH4 below their boiling points prior to launch
Figure 18. Additively Manufactured LOX/LCH4 and fielding a passive cryogenic fluid management
Turbo-pump Testing (CFM) system could be sufficient for missions to
polar regions on the surface of the moon. This
mission would require an insulation system designed
for the pre-launch, ascent, and vacuum environ-
ments, low-conductivity tank mounting, and
directional sun shielding. A thermodynamic vent
system (TVS) would provide mixing and recirculation
to counteract thermal stratification and control
pressure rise within the tanks. NASA modeling tools
indicated that an LOX/LCH4 system integrating these
features could enable a polar-region lunar surface
Figure 19. Additively Manufactured LOX/LCH4
dwell time as long as 240 days without active
Turbo-pump Testing
refrigeration [2]. NASA has demonstrated 13-day
Early LOX/LCH4 turbo-pump level demonstrations storage of LCH4 with helium pressurization using
sent the entire fuel flow to an external burn stack for passive CFM techniques in the Multipurpose
Hydrogen Test Bed (MHTB) in 2006 (Fig. 20a) [9]. overall heat leak into the tank. While CPST focused
Also, NASA’s Methane Lunar Surface Thermal on liquid hydrogen stored at 20 K, the strut concept
Control (MLSTC) Test (Fig. 20b) validated control is directly relevant to LOX/LCH4 systems, whose
predictions for the tanks of a lunar ascent vehicle warmer storage temperatures are less challenging
concept [2]. than the hydrogen application in which the strut has
already succeeded.

Figure 20. a. Multipurpose Hydrogen Test Bed

(MHTB) tank (left) and b. Methane Lunar Surface
Thermal Control (MLSTC) tank (center) and c. LOX Figure 21. NASA’s CPST Composite Strut Design,
zero-boiloff test (right). as installed.
7.2. Active CFM (Cryocoolers) 7.4. Propellant Quantity Gauging
Longer-duration missions such as near-earth loiters Since most existing cryogenic propulsion systems
or missions to Mars would, however, heighten the have found use in launch environments or
Need for active refrigeration and would increase atmospheric flight environments, historical cryogenic
demand for on-board electrical power. Therefore, in propellant quantity gauging methods have benefitted
recent years, NASA has invested in 90 K flight- from propellant settling, which avoids the
weight cryocooler technology to address this need complexities associated with microgravity fluid
and has taken delivery of an operational prototype 90 behavior. As a result, relatively robust technologies
K reverse Brayton prototype developed by Creare. exist for gauging oxygen and methane under settled
This unit has also been used to demonstrate Initial conditions. These technologies include discrete
subscale testing with LOX indicated its robust sensor “rakes”, capacitance probes, and derivatives
capability for maintaining LOX storage with zero boil- of these basic concepts. Settling, of course,
off (see Fig. 20c) [10]. Further work remains to assumes that either a cryogenic reaction control
assess cryocooler performance in an integrated system (RCS) incorporates functional LADs to allow
LCH4 system, and to reduce active CFM risks for a RCS start-up from unsettled conditions in order to
full-scale LOX/LCH4 flight system. settle propellants or the spacecraft initially relies on
a more bulky gas-fed system or a conventional
7.3. Composite Struts
storable-liquid system.
All foreseeable cryogenic missions will rely heavily
Some more advanced missions, of course, may be
on thermal isolation of the propellant tanks. NASA’s
less tolerant of the impacts of settling propellants
recent work with the now-cancelled Cryogenic
whenever a propellant quantity measurement is
Propellant Storage and Transfer (CPST) project
needed. Hence, NASA has continued to invest in
included analysis, design, manufacturing, and test of
gauging methods for unsettled cryogenic
low-conductivity light-weight composite tank struts
propellants. Previous work within the Exploration
for cryogenic propellant tanks. Load testing (both
Technology Development Program CFM Project
compression and tension) under liquid-hydrogen-to-
(CFMP) investigated the application of pressure-
ambient thermal gradient conditions showed very
volume-temperature (PVT) methods, which are
positive results for this strut design, indicating robust
routinely used in storable-propellant systems, for
mechanical properties and significant reductions in
application in cryogenic systems. This methodology
proved feasible [2], although intuitively with greater developed and matured thermal and fluid physics
uncertainties than are typical for storable modeling tools for the design of future cryogenic
applications. Seeking greater accuracy for cryogenic LADs. Remaining risks for LOX/LCH4 LADs can be
gauging in microgravity, NASA has continued to retired with ground testing and detailed modeling of
invest in the promising Radio-Frequency Mass flight-representative LADs, followed by demon-
Gauge (RFMG) concept, through the CFMP, CPST, stration in a microgravity environment before fully
and Evolvable Cryogenics (eCRYO) projects. This relying on LAD performance in a high-risk mission.
concept uses the propellant’s dielectric properties Hence, initial missions may need to rely operationally
and the electromagnetic Eigenmodes (natural on propellant settling while carrying LADs as a flight
resonant frequencies) of the tank and propellant. demonstration objective.
This approach involves injecting a radio frequency
signal into the tank and pattern-matching the 8. INTEGRATED SYSTEMS DEMONSTRATION
reflected power spectrum to a database of simulated As shown above, much work has been completed in
Eigenmode frequencies to determine propellant the development and maturation of technologies
mass. This concept has been validated for oxygen required for an in-space LOX/LCH4 propulsion
and methane under settled conditions during ground system capable of performing basic functions. More
testing and for a simulant fluid under unsettled recent efforts at NASA have sought to evaluate these
conditions during parabolic aircraft flights [2]. At technologies within a system framework with the goal
present, NASA’s investment continues, as eCRYO is of identifying system interactions, investigating
developing an RFMG for demonstration aboard the integrated system timelines, and evaluating
International Space Station (ISS) as part of Robotic integrated system performance. Even within an
Refueling Mission 3. This continued progress (as environment of fluctuating budgets and shifting
well as ongoing low-level investments in alternate priorities, NASA has continued to take steps toward
backup concepts) bodes well for the availability of this goal of demonstrating the integration of CFM
unsettled propellant gauging capabilities within the technologies and integrated operations within
foreseeable future. storage tanks and feed lines, although without yet
7.5. Liquid Acquisition Devices accomplishing the goal of a fully integrated system-
level ground test incorporating all requisite
To enable the full range of missions without the technologies.
burden of separate propulsion systems for cryogenic
propellant settling, NASA has also continued to
invest in the development of cryogenic liquid
acquisition devices (LADs) which exploit surface
tension properties to separate liquid from gas and to
assure expulsion of gas-free liquid from the
propellant tanks in microgravity. One notable
achievement within the CFMP included
measurement of bubble point pressures (i.e. the
differential pressure across the LAD screen at which
gas pressure overcomes surface tension on the
Figure 22. a. NASA’s Auxiliary Propulsion System
wetted screen surface) for both LOX and LCH4.
Test Bed (left) and b. NASA’s Morpheus Flight Test
CFMP also investigated heat entrapment effects and
Vehicle (right)
helium evolution effects and commissioned an
independent LAD concept development through a NASA’s partially integrated system demonstrations
competitive procurement. Subsequent work under include work performed within several projects,
the LH2-focused CPST project conquered weld and including PCAD, the CFM project, Morpheus, CPST,
manufacturing challenges and produced prototype and the ongoing eCRYO project. The PCAD project
LAD designs that were successfully manufactured demonstrated LOX/LCH4 conditioning and distri-
and tested under gravity conditions. CPST also bution with an integrated flight-weight feed system
and thrusters in the APSTB (Fig. 22a) [11]. This test 9. CHALLENGES FOR FUTURE HUMAN
demonstrated use of a thermodynamic vent to chill EXPLORATION
the propellant manifold. By demonstrating feed line
thermal performance that met or exceeded thruster Considering the advanced development efforts
inlet condition requirements, this test showed that conducted by NASA (and industry partners) over the
distributed feed lines can be successfully designed last 15 years, the overall development risk for
to deliver gas-free liquid cryogenic propellants to LOX/LCH4 in-space propulsion has been significantly
thruster inlets in a spacecraft or vehicle application. reduced. While these efforts have provided a strong
foundation for the pursuit of an initial flight capability,
Although not focused on long-duration cryogenic some challenges still exist, requiring additional
storage, the well-publicized Morpheus vehicle flight investigations/risk reduction testing. These
tests (Fig. 22b) have successfully provided short- remaining challenges include the following:
duration atmospheric flight demonstrations,
investigating control algorithms and response times  Integrated Storage testing with 90-Kelvin
of an LOX/LCH4 propulsion system during time- cryocoolers
critical ascent and descent operations [12].  Reaction control thruster design maturation
 Design maturation for regeneratively cooled
Finally, the hydrogen-focused CPST project main engines
successfully performed vibroacoustic tests of a  Design of low-leakage, long-duration
cryogenic tank with integrated foam and MLI cryogenic valves
insulation (Fig. 23a) and later demonstrated (more
difficult) LH2 storage in the Engineering Development More advanced in-space capabilities (landers,
Unit (EDU) tank (Fig. 23b) incorporating an ascent stages, depots, etc.) require additional
integrated passive CFM approach as well as the technology maturation for:
composite struts, prototype LADs, and RFMG
 Pump-fed LOX/LCH4 engines with deep
described in the preceding section.
throttle capability
 Leak detection
 Zero-G mass gauging technology
maturation
 Automated fluid couplings for space
cryogenic systems
 Zero-G demonstration of cryogenic liquid
acquisition devices

Due to currently evolving architecture requirements

a flexible test-bed approach to risk reduction testing
is recommended. A system-level ground test bed
capable of parametric adjustment of operating and
Figure 23. a. The CPST Vibroacoustic Test Article test conditions could evolve as the architecture
(VATA) with integrated foam and MLI (left) and requirements solidify, and could ultimately lead to a
b. The CPST Engineering Development Unit (EDU) potential risk reduction flight demonstration.
in its eventual, fully outfitted configuration (right)
10. SUMMARY AND CONCLUSIONS
Considered in total, these partial system-level
demonstrations combine to increase confidence that Building on years of foundational R&D activities
the infusion of LOX/LCH4 technologies into initial NASA has conducted multiple LOX Methane
mission capabilities is nearly within reach, with only advanced development efforts and hardware
a short list of challenges remaining. demonstrations over the last 15 years. While, over
the years, these efforts were focused on different
ultimate applications (e.g. non-toxic propulsion for
RLVs, crew module and lunar lander propulsion, AIAA Aerodynamic Measurement Technology
human space exploration) these efforts combine to and Ground Testing Conference
significantly reduce development risks associated 7. Judd, D. D., Buccella, S., Alkema, M., Hewitt, R.,
with future methane propulsion systems for human McLaughlin, B., Hart, G., Veith, E. (2006).
exploration. Building on these foundational risk Development Testing of a LOX/Methane
reduction efforts, we are well positioned to pursue an Engine for In-Space Propulsion. 42nd
initial operational capability. A system-level ground AIAA/ASMESAE/ASEE Joint Propulsion
test bed capable of parametric operating and test Conference & Exhibit.
conditions is a logical next step. This test bed would 8. Jones, C. P., Robertson, E. H., Koelbl, M.B., and
evolve as the architecture requirements solidify, and Singer, C. (2016) Additive Manufacturing a
would ultimately lead to a potential risk reduction Liquid Hydrogen Rocket Engine. Space
flight demonstration. Propulsion 2016. 20160129.
While development risks still exist (requiring some 9. Hastings, L.J. & Bolshinskiy, L.G. & Hedayat, A.,
advanced development efforts), the majority are Flachbart, R.H., Sisco, J.D., Schnell, A.R.
related to engineering challenges rather than the (2014) Liquid Methane Testing With a Large-
development of entirely new technologies. Scale Spray Bar Thermodynamic Vent System.
NASA Technical Report. NASA/TP-2014-
Sufficient investments have been made to enable a 218197.
path toward an initial LOX/LCH4 propulsion 10. Plachta, D.W., Johnson, W.L., Feller, J.R.
capability. (2014) Cryogenic Boil-Off Reduction System
Testing. 50th AIAA/ASME/SAE/ASEE Joint
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