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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS2010

Liquid Oxygen/Liquid Methane Component Technology Development at MSFC

The National Aeronautics & Space Administration (NASA) has identified Liquid Oxygen (LOX)/Liquid Methane (LCH4) as a potential propellant combination for future space vehicles based upon exploration studies. The technology is estimated to have higher performance and lower overall systems mass compared to existing hypergolic propulsion systems. Besides existing in-house risk reduction activities, NASA has solicited from industry their participation on component technologies based on the potential application to the lunar ascent main engine (AME). Contracted and NASA efforts have ranged from valve technologies to engine system testbeds. The application for the AME is anticipated to be an expendable, pressure-fed engine for ascent from the moon at completion of its lunar stay. Additionally, the hardware is expected to provide an abort capability prior to landing, in the event that descent systems malfunction. For the past 4 years, MSFC has been working with the Glenn Research Center and the Johnson Space Center on methane technology development. This paper will focus on efforts specific to MSFC in pursuing ignition, injector performance, chamber material assessments and cryogenic valve technologies. Ignition studies have examined characteristics for torch, spark and microwave systems. Injector testing has yielded insight into combustion performance for shear, swirl and impinging type injectors. The majority of chamber testing has been conducted with ablative and radiatively cooled chambers with planned activities for regenerative and transpiration cooled chambers. Lastly, an effort is underway to examine the long duration exposure issues of cryogenic valve internal components. The paper will summarize the status of these efforts.

Robinson, Joel W.↗
NASA NTRS1989

Liquid Oxygen Cooling of Hydrocarbon Fueled Rocket Thrust Chambers

Rocket engines using liquid oxygen (LOX) and hydrocarbon fuel as the propellants are being given serious consideration for future launch vehicle propulsion. Normally, the fuel is used to regeneratively cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability. Another possibility for the coolant is the liquid oxygen. Combustion chambers previously tested with LOX and RP-1 as propellants and LOX as the collant demonstrated the feasibility of using liquid oxygen as a coolant up to a chamber pressure of 13.8 MPa (2000 psia). However, there was concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot gas side wall. In order to study this effect, chambers were fabricated with slots machined upstream of the throat between the cooling passage wall and the hot gas side wall to simulate cracks. The chambers were tested at a nominal chamber pressure of 8.6 MPa (1247 psia) over a range of mixture ratios from 1.9 to 3.1 using liquid oxygen as the coolant. The results of the testing showed that the leaking LOX did not have a deleterious effect on the chambers in the region of the slots. However, there was unexplained melting in the throat region of both chambers, but not in line with the slots.

Armstrong, Elizabeth S.↗
NASA NTRS1989

Liquid oxygen cooling of hydrocarbon fueled rocket thrust chambers

Rocket engines using liquid oxygen (LOX) and hydrocarbon fuel as the propellants are being given serious consideration for future launch vehicle propulsion. Normally, the fuel is used to regeneratively cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability. Another possibility for the coolant is the liquid oxygen. Combustion chambers previously tested with LOX and RP-1 as propellants and LOX as the coolant demonstrated the feasibility of using liquid oxygen as a coolant up to a chamber pressure of 13.8 MPa (2000 psia). However, there was concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot gas side wall. In order to study this effect, chambers were fabricated with slots machined upstream of the throat between the cooling passage wall and the hot gas side wall to simulate cracks. The chambers were tested at a nominal chamber pressure of 8.6 MPa (1247 psia) over a range of mixture ratios from 1.9 to 3.1 using liquid oxygen as the coolant. The results of the testing showed that the leaking LOX did not have a deleterious effect on the chambers in the region of the slots. However, there was unexplained melting in the throat region of both chambers, but not in line with the slots.

Armstrong, Elizabeth S.↗
NASA NTRS1979

Structural analysis of cylindrical thrust chambers, volume 1

Life predictions of regeneratively cooled rocket thrust chambers are normally derived from classical material fatigue principles. The failures observed in experimental thrust chambers do not appear to be due entirely to material fatigue. The chamber coolant walls in the failed areas exhibit progressive bulging and thinning during cyclic firings until the wall stress finally exceeds the material rupture stress and failure occurs. A preliminary analysis of an oxygen free high conductivity (OFHC) copper cylindrical thrust chamber demonstrated that the inclusion of cumulative cyclic plastic effects enables the observed coolant wall thinout to be predicted. The thinout curve constructed from the referent analysis of 10 firing cycles was extrapolated from the tenth cycle to the 200th cycle. The preliminary OFHC copper chamber 10-cycle analysis was extended so that the extrapolated thinout curve could be established by performing cyclic analysis of deformed configurations at 100 and 200 cycles. Thus the original range of extrapolation was reduced and the thinout curve was adjusted by using calculated thinout rates at 100 and 100 cycles. An analysis of the same underformed chamber model constructed of half-hard Amzirc to study the effect of material properties on the thinout curve is included.

Armstrong, W. H.↗
NASA NTRS2010

Testing and Results of Vacuum Swing Adsorption Units for Spacesuit Carbon Dioxide and Humidity Control

A principal concern for extravehicular activity (EVA) space suits is the capability to control carbon dioxide (CO2) and humidity (H2O) for the crewmember. The release of CO2 in a confined or unventilated area is dangerous for human health and leads to asphyxiation; therefore, CO2 and H2O become leading factors in the design and development of the spacesuit. An amine-based CO2 and H2O vapor sorbent for use in pressure-swing re-generable beds has been developed by Hamilton Sundstrand. The application of solid-amine materials with vacuum swing adsorption technology has shown the capacity to concurrently manage CO2 and H2O levels through a fully regenerative cycle eliminating mission constraints imposed with non-regenerative technologies. Two prototype solid amine-based systems, known as rapid cycle amine (RCA), were designed to continuously remove CO2 and H2O vapor from a flowing ventilation stream through the use of a two-bed amine based, vacuum-swing adsorption system. The Engineering and Science Contract Group (ESCG) RCA is the first RCA unit implementing radial flow paths, whereas the Hamilton Sundstrand RCA was designed with linear flow paths. Testing was performed in a sea-level pressure environment and a reduced-pressure environment with simulated human metabolic loads in a closed-loop configuration. This paper presents the experimental results of laboratory testing for a full-size and a sub-scale test article. The testing described here characterized and evaluated the performance of each RCA unit at the required Portable Life Support Subsystem (PLSS) operating conditions. The test points simulated a range of crewmember metabolic rates. The experimental results demonstrate the ability of each RCA unit to sufficiently remove CO2 and H2O from a closed loop ambient or subambient atmosphere.

McMillin, Summer↗
NASA NTRS2011

Testing and Results of Vacuum Swing Adsorption Units for Spacesuit Carbon Dioxide and Humidity Control

A principal concern for extravehicular activity (EVA) spacesuits is the capability to control carbon dioxide (CO2) and humidity (H2O) for the crewmember. The release of CO2 in a confined or unventilated area is dangerous for human health and leads to asphyxiation; therefore, CO2 and H2O control become leading factors in the design and development of the spacesuit. An amine-based CO2 and H2O vapor sorbent for use in pressure-swing regenerable beds has been developed by Hamilton Sundstrand. The application of solidamine materials with vacuum swing adsorption technology has shown the capacity to concurrently manage CO2 and H2O levels through a fully regenerative cycle eliminating mission constraints imposed with nonregenerative technologies. Two prototype solid amine-based systems, known as rapid cycle amine (RCA), were designed to continuously remove CO2 and H2O vapor from a flowing ventilation stream through the use of a two-bed amine based, vacuum-swing adsorption system. The Engineering and Science Contract Group (ESCG) RCA implements radial flow paths, whereas the Hamilton Sundstrand RCA was designed with linear flow paths. Testing was performed in a sea-level pressure environment and a reduced-pressure environment with simulated human metabolic loads in a closed-loop configuration. This paper presents the experimental results of laboratory testing for a full-size and a sub-scale test article. The testing described here characterized and evaluated the performance of each RCA unit at the required Portable Life Support Subsystem (PLSS) operating conditions. The test points simulated a range of crewmember metabolic rates. The experimental results demonstrated the ability of each RCA unit to sufficiently remove CO2 and H2O from a closed loop ambient or sub-ambient atmosphere.

McMillin, Summer D.↗
NASA NTRS2018

CDKN1a/p21 Plays a Critical Role in Suppressing Stem Cell Regenerative Potential During Aging

Unloading during spaceflight is known to adversely affect mammalian physiology. Mechanical stimulation is required for repair and regeneration by stem cell lineages to maintain tissue health and mass. CDKN1a/p21 functions as a potent cell cycle arrest molecule and we previously found that CDKN1a/p21 was overexpressed in mouse bone during 15-days of spaceflight on STS-131 and localized to osteoprecursor cells in the femur. Therefore, we hypothesized that altered expression of CDKN1a/p21 leads to an arrest of bone formation during spaceflight in response to altered load. To study CDKN1a/p21 and its role in stem cell-based tissue regeneration, we use a CDKN1a/p21 knockout (KO) mouse to investigate the impact on bone structure, osteoprogenitor proliferation, and mineralized nodule formation. We have shown that bone marrow stem cells isolated from juvenile (11-week-old) and skeletally mature (18-week-old) KO mice have an increased bone formation potential as evidenced by increased proliferation and mineralization rates. In addition, we have shown that juvenile KO mice display significantly increased bone volume fraction (BV/TV) relative to wildtype (WT) mice, but not in skeletally mature KO mice, indicating increased resorption and bone turnover in adult mice. To more closely examine age differences in the KO mouse, we will study a wider spectrum of mice ranging from 4 weeks to 12 months in age. To do this, we will analyze differences in bone morphometric parameters using MicroCT and osteoblastogenesis assays. The pelvis, femur, and tibia are key in distributing weight and we expect to see altered remodeling and stem cell potential with age. In combination with histomorphometry, these results will help elucidate the complex mechanisms underlying bone tissue maintenance and stem cell regeneration.

stem cells↗
NASA NTRS2012

Development and Testing of a Sorbent-Based Atmosphere Revitalization System 2010/2011

Spacecraft being developed for future exploration missions incorporate Environmental Control and Life Support Systems (ECLSS) that limit weight, power, and volume thus requiring systems with higher levels of efficiency while maintaining high dependability and robustness. For air revitalization, an approach that meets those goals utilizes a regenerative Vacuum-Swing Adsorption (VSA) system that removes 100% of the CO2 from the cabin atmosphere as well as 100% of the water. A Sorbent Based Atmosphere Revitalization (SBAR) system is a VSA system that utilizes standard commercial adsorbents that have been proven effective and safe in spacecraft including Skylab and the International Space Station. The SBAR system is the subject of a development, test, and evaluation program that is being conducted at NASA s Marshall Space Flight Center. While previous testing had validated that the technology is a viable option, potential improvements to system design and operation were identified. Modifications of the full-scale SBAR test articles and adsorption cycles have been implemented and have shown significant performance gains resulting in a decrease in the consumables required for a mission as well as improved mission safety. Previous testing had utilized single bed test articles, during this period the test facility was enhanced to allow testing on the full 2-bed SBAR system. The test facility simulates a spacecraft ECLSS and allows testing of the SBAR system over the full range of operational conditions using mission simulations that assess the real-time performance of the SBAR system during scenarios that include the metabolic transients associated with extravehicular activity. Although future manned missions are currently being redefined, the atmosphere revitalization requirements for the spacecraft are expected to be quite similar to the Orion and the Altair vehicles and the SBAR test program addressed validation to the defined mission requirements as well as operation in other potential vehicle architectures. The development program, including test articles, the test facility, and tests and results through early 2011 is discussed.

Miller, Lee A.↗
NASA NTRS2008

Lunar Outpost Life Support Trade Studies

Engineering trade-off studies of life support system architecture and technology options were conducted for potential lunar surface mission scenarios within NASA's Constellation Program. The scenarios investigated are based largely on results of the NASA Lunar Architecture Team (LAT) Phase II study. In particular, the possibility of Hosted Sortie missions, the high cost of power during eclipse periods, and the potential to reduce life support consumables through scavenging, in-situ resources, and alternative EVA technologies were all examined. These trade studies were performed within the Systems Integration, Modeling and Analysis (SIMA) element of NASA's Exploration Life Support (ELS) technology development project. The tools and methodology used in the study are described briefly, followed by a discussion of mission scenarios, life support technology options and results presented in terms of equivalent system mass for various regenerative life support technologies and architectures. Three classes of repeated or extended lunar surface missions were investigated in this study along with several life support resource scenarios for each mission class. Individual mission durations of 14 days, 90 days and 180 days were considered with 10 missions assumed for each at a rate of 2 missions per year. The 14-day missions represent a class of Hosted Sortie missions where a pre-deployed and potentially mobile habitat provides life support for multiple crews at one or more locations. The 90-day and 180-day missions represent lunar outpost expeditions with a larger fixed habitat. The 180-day missions assume continuous human presence and must provide life support through eclipse periods of up to 122 hours while the 90-day missions are planned for best-case periods of nearly continuous sunlight. This paper investigates system optimization within the assumptions of each scenario and addresses how the scenario selected drives the life support system to different designs. Subsequently, these analysis results can be used to determine which technologies may be good choices throughout a broad range of architectures.

Lange, Kevin E.↗
NASA NTRS1992

Pressurized Lunar Rover (PLR)

The objective of this project was to design a manned pressurized lunar rover (PLR) for long-range transportation and for exploration of the lunar surface. The vehicle must be capable of operating on a 14-day mission, traveling within a radius of 500 km during a lunar day or within a 50-km radius during a lunar night. The vehicle must accommodate a nominal crew of four, support two 28-hour EVA's, and in case of emergency, support a crew of six when near the lunar base. A nominal speed of ten km/hr and capability of towing a trailer with a mass of two mt are required. Two preliminary designs have been developed by two independent student teams. The PLR 1 design proposes a seven meter long cylindrical main vehicle and a trailer which houses the power and heat rejection systems. The main vehicle carries the astronauts, life support systems, navigation and communication systems, lighting, robotic arms, tools, and equipment for exploratory experiments. The rover uses a simple mobility system with six wheels on the main vehicle and two on the trailer. The nonpressurized trailer contains a modular radioisotope thermoelectric generator (RTG) supplying 6.5 kW continuous power. A secondary energy storage for short-term peak power needs is provided by a bank of lithium-sulfur dioxide batteries. The life support system is partly a regenerative system with air and hygiene water being recycled. A layer of water inside the composite shell surrounds the command center allowing the center to be used as a safe haven during solar flares. The PLR 1 has a total mass of 6197 kg. It has a top speed of 18 km/hr and is capable of towing three metric tons, in addition to the RTG trailer. The PLR 2 configuration consists of two four-meter diameter, cylindrical hulls which are passively connected by a flexible passageway, resulting in the overall vehicle length of 11 m. The vehicle is driven by eight independently suspended wheels. The dual-cylinder concept allows articulated as well as double Ackermann steering. The primary power of 8 kW is supplied by a dynamic isotope system using a closed Brayton cycle with a xenon-hydrogen mixture as the working fluid. A sodium-sulfur battery serves as the secondary power source. Excess heat produced by the primary power system and other rover systems is rejected by radiators located on the top of the rear cylinder. The total mass of the PLR 2 is 7015 kg. Simplicity and low total weight have been the driving principles behind the design of PLR 1. The overall configuration consists of a 7-m-long, 3-m-diameter cylindrical main vehicle and a two-wheeled trailer. The cylinder of the main body is capped by eight-section, faceted, semi-hemispherical ends. The trailer contains the RTG power source and is not pressurized. The shell of the main body is constructed of a layered carbon fiber/foam/Kevlar sandwich structure. Included in the shell is a layer of water for radiation protection. The layer of water extends from the front of the rover over the crew compartment and creates a safe haven for the crew during a solar flare-up. The carbon fiber provides the majority of the strength and stiffness and the Kevlar provides protection from micrometeoroids. The Kevlar is covered with a gold foil and multi-layer insulation (MLI) to reduce radiation degradation and heat transfer through the wall. A thin thermoplastic layer seals the fiber and provides additional strength.

Creel, Kenneth↗
NASA NTRS2002

Lightweight, Actively Cooled Ceramic Matrix Composite Thrustcells Successfully Tested in Rocket Combustion Lab

In a joint effort between the NASA Glenn Research Center and the NASA Marshall Space Flight Center, regeneratively cooled ceramic matrix composite (CMC) thrustcells were developed and successfully tested in Glenn's Rocket Combustion Lab. Cooled CMC's offer the potential for substantial weight savings over more traditional metallic parts. Two CMC concepts were investigated. In the first of these concepts, an innovative processing approach utilized by Hyper-Therm, Inc., allowed woven CMC coolant containment tubes to be incorporated into the complex thruster design. In this unique design, the coolant passages had varying cross-sectional shapes but maintained a constant cross-sectional area along the length of the thruster. These thrusters were silicon carbide matrix composites reinforced with silicon carbide fibers. The second concept, which was supplied by Ceramic Composites, Inc., utilized copper cooling coils surrounding a carbon-fiber-reinforced carbon matrix composite. In this design, a protective gradient coating was applied to the inner thruster wall. Ceramic Composites, Inc.'s, method of incorporating the coating into the fiber and matrix eliminated the spallation problem often observed with thermal barrier coatings during hotfire testing. The focus of the testing effort was on screening the CMC material's capabilities as well as evaluating the performance of the thermal barrier or fiber-matrix interfacial coatings. Both concepts were hot-fire tested in gaseous O2/H2 environments. The test matrix included oxygen-to-fuel ratios ranging from 1.5 to 7 with chamber pressures to 400 psi. Steady-state internal wall temperatures in excess of 4300 F were measured in situ for successful 30-sec test runs. Photograph of actively cooled composite thrustcell fabricated by Hyper-Therm is shown. The thrustcell is a silicon-carbide-fiber-reinforced silicon carbide matrix composite with woven cooling channels. The matrix is formed via chemical vapor infiltration. Photograph of hot-fire test of an actively cooled carbon-fiber-reinforced carbon matrix composite thrustcell is also shown. This composite thrustcell, which was fabricated by CCI, Inc., was wound with copper cooling coils to contain the water coolant. The tests were run with oxygen fuel ratios up to seven with chamber pressures of 200 psia.

Jaskowiak, Martha H.↗
NASA NTRS2004

Emerging Fuel Cell Technology Being Developed: Offers Many Benefits to Air Vehicles

Fuel cells, which have recently received considerable attention for terrestrial applications ranging from automobiles to stationary power generation, may enable new aerospace missions as well as offer fuel savings, quiet operations, and reduced emissions for current and future aircraft. NASA has extensive experience with fuel cells, having used them on manned space flight systems over four decades. Consequently, the NASA Glenn Research Center has initiated an effort to investigate and develop fuel cell technologies for multiple aerospace applications. Two promising fuel cell types are the proton exchange membrane (PEM) and solid oxide fuel cell (SOFC). PEM technology, first used on the Gemini spacecraft in the sixties, remained unutilized thereafter until the automotive industry recently recognized the potential. PEM fuel cells are low-temperature devices offering quick startup time but requiring relatively pure hydrogen fuel. In contrast, SOFCs operate at high temperatures and tolerate higher levels of impurities. This flexibility allows SOFCs to use hydrocarbon fuels, which is an important factor considering our current liquid petroleum infrastructure. However, depending on the specific application, either PEM or SOFC can be attractive. As only NASA can, the Agency is pursuing fuel cell technology for civil uninhabited aerial vehicles (UAVs) because it offers enhanced scientific capabilities, including enabling highaltitude, long-endurance missions. The NASA Helios aircraft demonstrated altitudes approaching 100,000 ft using solar power in 2001, and future plans include the development of a regenerative PEM fuel cell to provide nighttime power. Unique to NASA's mission, the high-altitude aircraft application requires the PEM fuel cell to operate on pure oxygen, instead of the air typical of terrestrial applications.

Walker, James F.↗
NASA NTRS1990

A solar power system for an early Mars expedition

As NASA looks at missions that will expand human presence in the solar system, the power requirements for such missions need to be defined, developed and analyzed. One mission under consideration consists of a 40 day manned Mars surface expedition to perform science experiments. The mission time was centered around an aerocentric longitude (Ls) of 90 deg to lessen the probability of an occurrence of a local or planetary dust storm. The mission site was arbitrarily located at the Martian equator. The power requirements were assumed to be 40 kWe for life support and experiment power during the Martian day and 20 kWe for life support during the Martian night. A solar energy system consisting of roll-out amorphous silicon arrays and a hydrogen-oxygen regenerative fuel cell energy storage system was chosen for the study. The power available from a roll-out array, when plotted against time, approaches a cosine-like curve and depends on both array area and the amount of solar irradiance impinging on its horizontal surface. The array is sized to provide at least 20 KWe when the sun is 12.5 deg above the horizon and ramp up to 140 kWe peak power at Martian noon. In this configuration, the array is capable of supplying 40 KWe continuously to the user for the majority of the Martian day while supplying the excess energy to the electrolyzer portion of the energy storage system. A roll-out, pumped loop radiator system is used to dissipate the waste heat produced by the fuel cell. The power management and distribution system inverts the power from the individual solar array sub-modules and the fuel cell stacks and connects them to a 440 VAC single phase 20 kHz main bus. The total power system is comprised of 80 individual solar array modules with an integral bus and three energy storage modules consisting of fuel cell and electrolyzer stacks, reactant storage tanks, and a roll-out radiator. Power system mass, stowed volume, and deployed area were determined. Day/night power splits of 40/10 kWe, 40/30 kWe, and 40/40 kWe were also considered to determine the impact of a range of nighttime power requirements on the baseline system.

Mckissock, Barbara I.↗
NASA NTRS2006

Closure of Regenerative Life Support Systems: Results of the Lunar-Mars Life Support Test Project

Future long duration human exploration missions away from Earth will require closed-loop regenerative life support systems to reduce launch mass, reduce dependency on resupply and increase the level of mission self sufficiency. Such systems may be based on the integration of biological and physiocochemical processes to produce potable water, breathable atmosphere and nutritious food from metabolic and other mission wastes. Over the period 1995 to 1998 a series of ground-based tests were conducted at the National Aeronautics and Space Administration, Johnson Space Center, to evaluate the performance of advanced closed-loop life support technologies with real human metabolic and hygiene loads. Named the Lunar-Mars Life Support Test Project (LMLSTP), four integrated human tests were conducted with increasing duration, complexity and closure. The first test, LMLSTP Phase I, was designed to demonstrate the ability of higher plants to revitalize cabin atmosphere. A single crew member spent 15 days within an atmospherically closed chamber containing 11.2 square meters of actively growing wheat. Atmospheric carbon dioxide and oxygen levels were maintained by control of the rate of photosynthesis through manipulation of light intensity or the availability of carbon dioxide and included integrated physicochemical systems. During the second and third tests, LMLSTP Phases II & IIa, four crew members spent 30 days and 60 days, respectively, in a larger sealed chamber. Advanced physicochemical life support hardware was used to regenerate the atmosphere and produce potable water from wastewater. Air revitalization was accomplished by using a molecular sieve and a Sabatier processor for carbon dioxide absorption and reduction, respectively, with oxygen generation performed by water hydrolysis. Production of potable water from wastewater included urine treatment (vapor compression distillation), primary treatment (ultrafiltration/reverse osmosis and multi-filtration) and post processing. For the Phase II test the water recovery rate ranged from 95 to 98%, depending on the processor. LMLSTP Phase III, the fourth test of the series, had a duration of 91 days and included four crew members. The test demonstrated an integration of physicochemical and biological technologies for air revitalization, water recovery and waste processing. Wheat supplemented the physicochemical air revitalization systems by providing approximately 25% of the oxygen required for the 4-person crew. The water recovery system included immobilized cell and trickling filter bioreactors for primary water treatment, reverse osmosis and air evaporation systems for secondary water treatment, followed by post processing. The 8 day initial supply of water was recycled through the chamber and crew 10 times over the course of the test. Grain from the wheat together with fresh lettuce from a small growth chamber within the crew chamber provided supplementation to the stored food system, but at a level less than 5% of the crew s caloric requirement. An incinerator was used to demonstrate mineralization of the crew s solid waste, with the combustion products (mainly carbon dioxide) returned to the wheat for conversion to oxygen.

Barta, Daniel↗
NASA NTRS2013

Cu-Cr-Nb-Zr Alloy for Rocket Engines and Other High-Heat- Flux Applications

Rocket-engine main combustion chamber liners are used to contain the burning of fuel and oxidizer and provide a stream of high-velocity gas for propulsion. The liners in engines such as the Space Shuttle Main Engine are regeneratively cooled by flowing fuel, e.g., cryogenic hydrogen, through cooling channels in the back side of the liner. The heat gained by the liner from the flame and compression of the gas in the throat section is transferred to the fuel by the liner. As a result, the liner must either have a very high thermal conductivity or a very high operating temperature. In addition to the large heat flux (>10 MW/sq m), the liners experience a very large thermal gradient, typically more than 500 C over 1 mm. The gradient produces thermally induced stresses and strains that cause low cycle fatigue (LCF). Typically, a liner will experience a strain differential in excess of 1% between the cooling channel and the hot wall. Each time the engine is fired, the liner undergoes an LCF cycle. The number of cycles can be as few as one for an expendable booster engine, to as many as several thousand for a reusable launch vehicle or reaction control system. Finally, the liners undergo creep and a form of mechanical degradation called thermal ratcheting that results in the bowing out of the cooling channel into the combustion chamber, and eventual failure of the liner. GRCop-84, a Cu-Cr-Nb alloy, is generally recognized as the best liner material available at the time of this reporting. The alloy consists of 14% Cr2Nb precipitates in a pure copper matrix. Through experimental work, it has been established that the Zr will not participate in the formation of Laves phase precipitates with Cr and Nb, but will instead react with Cu to form the desired Cu-Zr compounds. It is believed that significant improvements in the mechanical properties of GRCop-84 will be realized by adding Zr. The innovation is a Cu-Cr-Nb-Zr alloy covering the composition range of 0.8 to 8.1 weight percent Cr, 0.7 to 7.2 weight percent Nb, 0.1 to 1.5 weight percent Zr, and balance Cu. The alloy combines two known strengthening mechanisms - dispersion strengthening by Cr2Nb precipitates (GRCop-84), and precipitation strengthening by CuxZr (AMZIRC) - to produce a synergistic increase in the capabilities of the alloy with the goal of achieving properties greater than either of the methods could achieve alone. The anticipated advantages of the alloy are higher strength at temperatures up to 700 C, improved creep strength, and significantly higher LCF lives relative to GRCop-84. The thermal expansion, thermal conductivity, and processing of the alloy are anticipated to remain largely unchanged relative to GRCop-84.

Ellis, David L.↗
NASA NTRS1993

Test results from a simple, low-cost, pressure-fed liquid hydrogen/liquid oxygen rocket combustor

A simple, low-cost rocket engine was designed, fabricated, and successfully hot fire tested over a wide range of interface conditions and operating parameters. The engine used low enthalpy hydrogen (45 to 70 R, 200 to 390 psia) and oxygen (139 to 163 R, 210 to 480 psia) propellants pressure-fed directly from facility cryogenic tanks. The engine demonstrated excellent performance, with 97% average combustion efficiency, and absence of combustion instabilities. Engine design chamber pressure was 300 psia, yielding about 16,500 pounds thrust at sea level with a 3:1 expansion ration test nozzle. The engine used a fixed-element injector based on TRW's unique coaxial pintle design, but was operated at 60%, 80%, and 100% thrust levels by throttling facility propellant valves. The engine was tested at propellant mixture ratios (O/F) from 5.8 to 8.4; design O/F was 6.6. To document combustion stability, in five tests RDX explosive pulse guns were detonated in radial and tangential directions across the combustion chamber during steady-state operation. The largest disturbance consisted of simultaneous detonation of a 20-grain radial gun and a 40-grain tangential gun. In no case was an instability, either feed system mode or chamber acoustic mode, excited. High-frequency piezoelectric pressure transducers documented stable recovery from disturbance overpressures within 40 milliseconds of peak pressure. A total of 67 firing tests, accumulating 149 seconds of firing time above 10% P(sub c), were performed. Since parametric testing required run durations of only 2 to 3 seconds, a heat sink combustion chamber was employed for most runs. To evaluate the feasibility of a low-cost ablative system for a flight engine design, one 20-second continuous firing was conducted with a silicone rubber chamber/throat/nozzle liner cast in one piece directly into the engine. The ablative engine operated at the equivalent of 309 seconds sea level specific impulse, when adjusted to a 98% efficient 6:1 expansion ration nozzle, and 430 seconds vacuum specific impulse, when adjusted to a 98% efficient 50:1 expansion ratio nozzle. This engine and test series represent an initial subscale demonstration of a new booster-class engine that eliminates the cost and complexity associated with regenerative cooling and typical engine cycles. This paper presents a description of the engine design and discussion and summary data plots of the performance measured during the parametric testing.

Dressler, G. A.↗