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Liquid Oxygen / Liquid Methane Test Results of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility
Abstract
Tests were conducted with the RS-18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA's Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to ∼122,000 ft (∼37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. LO2 flow ranged from 5.9 - 9.5 lbm/sec (2.7 - 4.3 kg/sec), and LCH4 flow varied from 3.0 - 4.4 lbm/sec (1.4 - 2.0 kg/sec) during the RS-18 hot-fire test series. Propellant flow rate was measured using a coriolis mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup due to two-phase flow effects. Subsequent cold-flow testing demonstrated that the propellant manifolds must be adequately flushed in order for the coriolis flow meters to give accurate data. The coriolis flow meters were later shown to provide accurate steady-state data, but the turbine flow meter data should be used in transient phases of operation. Thrust was measured using three load cells in parallel, which also provides the capability to calculate thrust vector alignment. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes. All of these objectives were met with the RS-18 data and additional testing data from subsequent LO2/methane test programs in 2009 which included the first simulated-altitude pyrotechnic ignition demonstration of LO2/ methane.
... Because of the envisioned booster engine application, the combustion chamber conditions were at relatively high chamber pressures (at least larger than 100 bar). Starting from the 2000s, NASA renewed the interest on this propellant combination for in-space rather than booster applications [6][7][8][9]. Oxygen-methane propulsion for planetary landing/ascending or orbit maneuvers would replace the toxic hypergolic propellants (like nitrogen tetroxide-hydrazine) that are typically used for such applications. Besides the higher specific impulse, the advantages over hypergolic propellants for in-space propulsion are the safer environment for the ground crews and astronauts when manipulating oxygen and methane, the possibility to produce these propellants in situ, thus reducing the need to carry them to extraterrestrial surfaces, and the ability to use common tanks for other subsystems (e.g., oxygen for life-support systems, methane for solid-oxide fuel cell power systems). ...
... Based on the above discussion, it is not surprising that studies on the possible occurrence of pseudo-phase change in rocket cooling channels have flourished recently, when the attention to the use of methane as a rocket fuel has more and more increased. The interest towards methane comes from its possible use as a denser and cheaper replacement of hydrogen in launch vehicles [5,6] and as a cheaper replacement of toxic storable propellants for space propulsion [7,8]. ...
Heating of liquid propellants used as the coolant in rocket engines may lead to undesired phenomena such as pseudo-boiling or heat transfer deterioration under specific conditions. This can be an issue for propellants characterized at the same time by relatively low critical pressure and temperature. Light hydrocarbons, as for instance methane, belong to this family. In the present paper, a critical review is made of the main results obtained by Authors and their coworkers for the present application. Focus is on the correlations and trends inferred by their numerical simulations mainly carried out considering methane as the coolant, perhaps the most challenging one.
... Several studies on the use of Hydrogen Peroxide/Kerosene (H 2 O 2 -RP1) as a green propellant combination suited for propulsion applications have been realized [7][8][9][10][11][12]. Yet, the overall performance can be limited by coking which remains a major problem for current Hydrocarbon cooled rocket engines; but currently and from a different aspect, because of the possibility of harvesting Methane (CH 4 ) from Mars, the interest of its use for interplanetary missions has increased [13][14]. Another alternative is hybrid systems, as they are known for their simplicity, safety and environmental advantages over conventional ones. ...
The magnitude of the exhaust velocity is dependent on molecular and chemical properties of the propellant, and the expansion ratio of the engine. The main requirement for all propellant combinations is to maximize the energy release per kilogram; in other words, the lower the mass for a given energy release, the higher is the ultimate velocity of the vehicle. This also implies that the mass flow rate depends on the density, while the exhaust velocity depends on the energy contained in the hot gas. The main performance parameters can be defined as the thrust coefficient related to the nozzle performance, and the characteristic velocity related to the propellant and combustion performance. The first parameter reaches its optimum when the exhaust pressure is equal to the ambient pressure while the second parameter is defined by the choice of propellant combination. The objective of this work is to investigate, numerically and thermodynamically, different combinations of oxidizers with conventional and alternative fuels, for different rocket propulsion and energy generation systems. Based on the state-of-the-art and the analysis results on the most promising fuel-oxidizer combinations for the future of aerospace propulsion and space transportation, in a technically and environmentally sound manner, conclusions are made.
... The study of heat transfer to near-critical fluids, that is to supercritical pressure fluids whose temperature is close to the pseudo-critical value a , has recently captured the interest of liquid rocket engine designers because of the possible use of methane as a denser and cheaper replacement of hydrogen in launch vehicles 2 and as a cheaper replacement of toxic storable propellants for space propulsion. 3,4 In case of liquid-oxygen/liquid-methane engines with chamber pressures larger than about 50 bar, a peculiar behavior of methane as a coolant is found: it enters channels at supercritical pressure and subcritical temperature and then, under heating from hot gas, its temperature increases up to reach and exceed the pseudo-critical value. Moreover, because the typical pressure and temperature of methane in cooling channels are relatively close to the critical point, the coolant flow behavior is strongly affected by the large property variations that occur in the near-critical region. ...
Liquid propellants, which are typically used for regenerative cooling of rocket thrust chambers, can
ow in channels at supercritical pressures and in the neighborhood of pseudo-critical temperature (near-critical fluid). This could be for instance the case for the envisioned liquid-oxygen/liquid-methane engines with chamber pressures larger than about 50 bar. When the fluid is in such a near-critical condition, deterioration in heat transfer can occur if the heat transfer level is higher than a threshold value. In this study detailed three-dimensional numerical analyses are performed to study the coupled wall/coolant environment of an electrically-heated test article designed in order to inves-tigate the thermo-uid dynamic behavior of methane inside a rectangular cooling channel which can be representative of a regenerative system. Di�fferent coolant pressure and surface roughness levels are considered in order to understand their infuence on the heat transfer capability of the cooling system. Results evidence that the heat transfer deterioration can be mitigated either by increasing the coolant pressure or by increasing the surface roughness. In the latter case, a penalty in terms of coolant pressure drop is expected.
... The study of heat transfer to near-critical fluids, that is to supercritical pressure fluids whose temperature is close to the pseudo-critical value, 1 has recently captured the interest of liquid rocket engine designers because of the possible use of methane as a denser and cheaper replacement of hydrogen in launch vehicles [5,6] and as a cheaper replacement of toxic storable propellants for space propulsion [7,8]. The interest in such flows for liquid rocket engines is driven by the fact that thrust chamber is cooled by one of the available propellants, which flows in suitable narrow channels. ...
The knowledge of the flow behavior inside cooling channels is of great importance to improve design and performance of regeneratively cooled rocket engines. The modeling of the coolant flow is a challenging task because of its particular features, such as the high wall temperature gradient, the high Reynolds number and the three-dimensional geometry of the passages. In case of methane as coolant, a further complication is the transcritical operating condition of the fluid. In this thermodynamic regime large changes of the fluid properties can greatly influence the coolant flow-field and the heat transfer. Numerical simulations of transcritical methane flow-field in asymmetrically heated rectangular channel with high aspect ratio and strong wall temperature differences are carried out by a suitable CFD solver. Results are discussed in detail and compared with supercritical methane flow-fields. Finally the aspect ratio effect on methane flow is analyzed by comparison of four different rectangular cooling channel geometries with fixed hydraulic diameter and coolant flows with the same mass flow-rate per unit area. Emphasis is given to the comparison of fluid cooling performance and pressure loss.
The NASA Space Launch Systems (SLS) outstanding capabilities for launching heavy, large diameter payloads will enable robust lunar architectures where the Near Rectilinear Halo Orbit (NRHO) is used as an aggregation node for lander and Orion elements. SLS capabilities and production status is discussed, as is the new large Exploration Upper Stage, currently in development, which will optimize the SLS Core and Booster Stages. A simplified lunar architecture is presented that takes advantage of the SLS launch capabilities. An overview of lander engine options is given and a propulsion trade study is performed and conclusions are drawn.
In 2008 the Boeing company conducted a six month design study for the NASA Johnson Space Center as part of the of the Altair Lunar lander concept desi gn effort. Information presented here is taken from Boeings final report presented to NASA in October 2008 in Houston to the Altair program office. This report contains a summary of one aspect of that stu dy; its objective is to present propulsion system t rade studies and configuration options for the Altair As cent Module. Vehicle subsystem concept design sought to maximize crew safety and vehicle reliability whi le staying within the mass and volume envelope provided by the Ares-V launch vehicle. Tradable ele ments in the Ascent module design included main engine location and number (single vs several, exte rnal vs embedded, bottom vs side placement); propellant type (hypergolic N2O4/MMH vs LO2/LCH4); Ascent crew cabin layout (horizontal or vertical cabin alignment) and several other items.
In recent years, methane has attracted attention as a propellant for liquid rocket engines because of its various advantages compared to typical propellants such as hydrogen. When methane is used as a coolant for a regenerative cooling system, its near-critical thermodynamic and transport properties experience large variations because its critical pressure is higher than that of typical propellants; this significantly influences the flowfield and heat transfer characteristics. Therefore, adequate understanding of the flowfield and heat transfer characteristics of methane in regenerative cooling channels is a prerequisite for future engine development. In this study, conjugated coolant and heat transfer simulations were performed to investigate the flowfield and heat transfer characteristics of transcritical methane flows in a sub-scale methane-cooled thrust chamber. The computed results were validated against experimental data measured in hot firing tests. They compared well with the measured pressures and temperatures in cooling channels, and wall temperatures were within the permitted levels. Detailed flow analysis revealed peculiar flow structures in the cooling channel: a strong secondary flow induced in the concave-heated part in the channel throat section and the coexistence of two different gas phases-ideal and real-in a single cross-section in the cylindrical region. A high wall temperature appeared in the cylindrical region of the thrust chamber under the considered conditions; this was due to the heat transfer deterioration induced by an M-shaped velocity profile and a turbulent heat flux reduction. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
The use of non-toxic propellants in future exploration vehicles would enable safer, more cost effective mission scenarios. One promising "green" alternative to existing hypergols is liquid methane/liquid oxygen. To demonstrate performance and prove feasibility of this propellant combination, a 100lbf LO2/LCH4 engine was developed and tested under the NASA Propulsion and Cryogenic Advanced Development (PCAD) project. Since high ignition energy is a perceived drawback of this propellant combination, a test program was performed to explore ignition performance and reliability versus delivered spark energy. The sensitivity of ignition to spark timing and repetition rate was also examined. Three different exciter units were used with the engine's augmented (torch) igniter. Propellant temperature was also varied within the liquid range. Captured waveforms indicated spark behavior in hot fire conditions was inconsistent compared to the well-behaved dry sparks (in quiescent, room air). The escalating pressure and flow environment increases spark impedance and may at some point compromise an exciter's ability to deliver a spark. Reduced spark energies of these sparks result in more erratic ignitions and adversely affect ignition probability. The timing of the sparks relative to the pressure/flow conditions also impacted the probability of ignition. Sparks occurring early in the flow could trigger ignition with energies as low as 1-6mJ, though multiple, similarly timed sparks of 55-75mJ were required for reliable ignition. An optimum time interval for spark application and ignition coincided with propellant introduction to the igniter and engine. Shifts of ignition timing were manifested by changes in the characteristics of the resulting ignition.
A study was conducted for the performance characteristics of methane taking recently the limelight in the world as a next-generation propellant, with the survey for state of the art in the development of methane/oxygen rocket engine being accompanied. Liquid methane as a rocket fuel has the favorable characteristics such as non-toxic, low cost, regenerative cooling capability, and potential for in-situ resource utilization (ISRU). The combination of liquid methane and liquid oxygen also provides the excellent performance including high specific impulse and low system mass. For these reasons, many researches have been actively carried out on the methane/oxygen engine, nevertheless, its technology readiness level is not that high enough just yet. Therefore, it is judged that it is the time to mitigate the technical gap with the space technology of advanced countries through a swift onset of the development of methane rocket engine.
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A ground test article was prepared for RS-18 engine testing using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA Glenn Research Center's Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propellant systems in lunar vehicle ascent engine applications. LO2/LCH4 testing capability at altitude conditions did not previously exist at WSTF for this size engine, and modifications were made to the Auxiliary Propulsion Systems Test Bed (APSTB) article. Altitude simulation is achieved using the WSTF Large Altitude Simulation System, which provides altitude conditions equivalent to ∼90,000 ft (∼27 km). For specific impulse calculations, engine thrust and propellant mass flow rates are measured. Propellant flow rate is measured using a coriolis-style mass-flow meter, and accuracy is compared with a serial turbine-style flow meter. Thrust is measured using three load cells in parallel. Igniter system capability is being developed to demonstrate two methods, a gaseous oxygen/methane spark torch igniter and solid propellant pyrotechnic igniter. Design, procurement and assembly are complete for the test article and test readiness is expected for hot-fires to begin pending completion of manifold buildup and system checkout. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/LCH4 lunar ascent engines, 2) obtain nozzle kinetics data to anchor two-dimensional kinetics codes, and 3) obtain vacuum ignition data for the torch and pyrotechnic igniters.
When United States President George W. Bush announced the Vision for Space Exploration in January 2004, twelve propulsion and launch system projects were being pursued in the Next Generation Launch Technology (NGLT) Program. These projects underwent a review for near-term relevance to the Vision. Subsequently, five projects were chosen as advanced development projects by NASA's Exploration Systems Mission Directorate (ESMD). These five projects were Auxiliary Propulsion, Integrated Powerhead Demonstrator, Propulsion Technology and Integration, Vehicle Subsystems, and Constellation University Institutes. Recently, an NGLT effort in Vehicle Structures was identified as a gap technology that was executed via the Advanced Development Projects Office within ESMD. For all of these advanced development projects, there is an emphasis on producing specific, near-term technical deliverables related to space transportation that constitute a subset of the promised NGLT capabilities. The purpose of this paper is to provide a brief description of the relevancy review process and provide a status of the aforementioned projects. For each project, the background, objectives, significant technical accomplishments, and future plans will be discussed. In contrast to many of the current ESMD activities, these areas are providing hardware and testing to further develop relevant technologies in support of the Vision for Space Exploration.
A non-toxic, dual thrust reaction control engine (RCE) was successfully tested over a broad range of operating conditions at the Aerojet Sacramento facility. The RCE utilized LOX/Ethanol propellants; and was tested in steady state and pulsing modes at 25-lbf thrust (vernier) and at 870-lbf thrust (primary). Steady state vernier tests vaned chamber pressure (Pc) from 0.78 to 5.96 psia, and mixture ratio (MR) from 0.73 to 1.82, while primary steady state tests vaned Pc from 103 to 179 psia and MR from 1.33 to 1.76. Pulsing tests explored EPW from 0.080 to 10 seconds and DC from 5 to 50 percent at both thrust levels. Vernier testing accumulated a total of 6,670 seconds of firing time, and 7,215 pulses, and primary testing accumulated a total of 2,060 seconds of firing time and 3,646 pulses.
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