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Oxidation resistance and high temperature tests of rhenium, tungsten, hafnium, and tantalum matrix composites with iridium in oxygen, fluorine, and boron atmospheres for liquid propellant engines
The space shuttle main engine propellant valve actuators (SSME) were designed to simulate the loads reflected into the SSME by the chamber coolant valve, the fuel preburner, and the oxidizer. The design, and functional description are included along with a list of the drawings. The load fixture control transform, friction torque, and flow calculations are reported.
Results are presented for an investigation intended to provide experimental data that can quantitatively describe the way in which various coaxial injector designs affect the stability of gaseous propellant rocket motors. The response factors of configurations that simulate the flow conditions in a gaseous-fuel injector element and a gaseous-oxidizer injector element are measured by using a modified impedance-tube technique and under cold-flow conditions simulating those observed in rocket motors with axial instability. The measured injector response factor data are presented and discussed. It is shown that there is reasonable agreement between the measured injector response factors and those predicted by the Feiler and Heidmann model (1967), and that the orifice length can be varied to shift the resonant frequency of the injector without any change in the magnitude of the response factor at resonance. A change in the injector open-area ratio is found to have a significant effect on the characteristics of the injector response factor.
Two cylindrical rocket thrust chamber cylinders were coated with a 0.203 mm layer of zirconium oxide using a process that employed electrodeposition of metal to a spray-coated mandrel. The cylinders were cyclically tested using hydrogen-oxygen propellants at a nominal chamber pressure of 4.14 MN/sq m to show the effect of the coating on life. Both cylinders failed prematurely due to causes unrelated to the coatings. Post destructive analysis showed no cooling passage wall deformation. Where erosion of the coating occurred, the coating thickness stabilized at 0.061 mm within 80 cycles and remained well adhered throughout the tests. The hot-gas-side metal wall temperature and the theoretical maximum effective strain were reduced by 80 and 92%, respectively, where the coating was 0.203 mm thick and by 40 and 50%, respectively, where the coating was 0.061 mm thick compared to an uncoated cylinder. The 50 and 92% reductions in maximum effective strain resulted in theoretical lives of 1900 cycles and greater than 10,000 cycles, respectively, compared to 300 cycles for the uncoated cylinder.
The In-Situ production of propellants for Mars missions has been considered to utilize the Carbon dioxide (CO2) in Mars atmosphere to produce Oxygen using a high temperature solid oxide electrolyzer. The oxygen then needs to be cooled, liquefied, and stored to be available for propulsion and other end users. The storage period could be up to two years either in the actual Mars ascent propulsion tanks or in a separate tank. Recent investigations have demonstrated the feasibility of both achieving zero-boil-off and controlling the pressure of oxygen within a tank using high efficiency cryocoolers. A representative configuration of tube on tank liquefaction using cryocooler is shown in Fig. 1.
In-Situ Resource Utilization (ISRU) is key to long term presence at any extraterrestrial destination. Current NASA direction is to achieve a sustainable presence on the lunar surface by 2028. Mission plans currently target the lunar South Pole to leverage the extended periods of solar illumination and allows for potential access to water ice in the permanently shadowed areas around the poles. With water, it is possible to produce both fuel and oxidizer to fully refuel a vehicle. In order to address ISRU infusion into mission planning, a study of an end-to-end ISRU propellant production system was initiated to assess ISRU architectures and obtain mass and power estimates for each. The results of these case studies will be presented.
In-Situ Resource Utilization (ISRU) is key to long term presence at any extraterrestrial destination. Current NASA direction is to achieve a sustainable presence on the lunar surface by 2028. Mission plans currently target the lunar South Pole to leverage the extended periods of solar illumination and allows for potential access to water ice in the permanently shadowed areas around the poles. With water, it is possible to produce both fuel and oxidizer to fully refuel a vehicle. In order to address ISRU infusion into mission planning, a study of an end-to-end ISRU propellant production system was initiated to assess ISRU architectures and obtain mass and power estimates for each. The results of these case studies will be presented. For this study, The ISRU system architecture involved two sites; the mine site in a shadowed crater where water ice is excavated and extracted from the regolith and the propellant production site at an illuminated ridge where the water is processed into liquefied O2 and H2 propellants. Fixed hardware would be emplaced at each site, with two alternating water tankers to transport water between them. Notional lunar sites were identified for this architecture for baseline environmental parameters. Technology solutions for each subsystem were selected based on those with the highest fidelity models or those that have empirical laboratory data to anchor to. While power needs were identified for each location, a power solution was not prescribed, therefore the masses presented do not include surface power systems. The baseline case in this study assumed that 10 mT of oxygen, along with enough hydrogen to support a propulsion mixture ratio of 6, must be produced in 225 days. Therefore 15 mT of water would need to be collected and processed. The baseline solution resulted in a system mass of5 mT and a total required power of 68kW. The majority of the mass was split between the ridge site system and the two water tankers (2.6 mTand 1.8mTrespectively). The majority of the required power was with three subsystems at approximately 20 kW each: hydrogen liquefaction, electrolysis, and the water extractor subsystem. Trades for four key variables are also presented, namely production rate, water concentration, dry overburden depth, and number of water transport trips. This can be compared to a system that targets oxygen from the minerals in the surface regolith material. A carbothermal reduction reactor system was used for this comparison. The use of direct solar thermal energy to process the regolith and the ease of access of the resource resulted in significantly lower values for the oxygen case: approximately 2.7 mT and 11.8 kW. However, the mass trade would favor the water case over successive missions where the hydrogen up-mass of 2 mT per mission will accrue against the oxygen system.1Aerospace Research Engineer, Chemical and Thermal Propulsion Systems Branch, Senior Member.2Mechanical Engineer, Propulsion and Power Division/Energy Conversion Systems, 2101 Nasa Pkwy/Mailcode EP3 Houston, TX 77058, AIAA member.
The IPD rocket engine is the product of a joint Air Force/NASA program to demonstrate the concept of a full-flow staged combustion power cycle. In this type of rocket engine cycle, both the entire fuel flow and the entire oxidizer flow are combusted in one of two pre-bumers and used to drive two turbopumps, thus utilizing the flow work of the total propellant flow. The basis of the IPD engine program relies on integrating newly developed and pre-existing hardware to demonstrate the component and material technologies to make this concept feasible, while simultaneously saving development time and costs. To provide insight to the project team and contractors during engine development and test phases, the engine system was modeled at MSFC using the Rocket Engine Transient Simulation (ROCETS) software to analyze system performance and determine component integration issues. The ROCETS software is used extensively a MSFC to perform steady-state power-balances and transient simulations of thermodynamic power and general fluid systems. The software is favored for its capability to solve large systems of non-linear equations, its librarie of fluid properties and flow devices, its flexibility to modify existing code to improve the physics-derived approximations of real fluid thermodynamic behavior, and the ability to add unique system constraints. The purpose of this paper is to present the methodology used to model the IPD engine system, detail the pitfalls encountered with the software, and explain the approximations made to more accurately represent engine component and fluid behavior. Engine system performance output from the model will be presented and explained in comparison with real fluid and component behavior.
Mining and processing of lunar material could possibly lead to more cost-efficient scenarios for permanent presence of man in space and on the Moon. Production of oxygen for use as propellant seems especially important. Different candidate processes for oxygen-extraction from lunar soil were proposed, of which the reduction of ilmenite by hydrogen was studied most. This process, however, needs the concentration of ilmenite from lunar regolith to a large extent and releases oxygen only with low efficiency. Another possibility - the fluorination method - which works with lunar bulk material as feedstock is discussed. Liberation of oxygen from silicate or oxide materials by fluorination methods has been applied in geoscience since the early sixties. The fact that even at moderate temperatures 98 to 100 percent yields can be attained, suggests that fluorination of lunar regolith could be an effective way of propellant production. Lunar soil contains about 50 percent oxygen by weight which is gained nearly completely through this process as O2 gas. The second-most element Si is liberated as gaseous SiF4. It could be used for production of Si-metal and fluorine-recycling. All other main elements of lunar soil will be converted into solid fluorides which also can be used for metal-production and fluorine-recycling. Preliminary results of small scale experiments with different materials are discussed, giving information on specific oxygen-yields and amounts of by-products as functions of temperature. These experiments were performed with an already existing fluorine extraction and collection device at the University of Bonn, normally used for determination of oxygen-isotopic abundances. Optimum conditions, especially concerning energy consumption, are investigated. Extrapolation of the experimental results to large industrial-type plants on the Moon is tried and seems to be promising at first sight. The recycling of the fluorine is, however, crucial for the process. It might be achieved by means of electrolysis. This needs further investigation. The technical problem of transport and handling of the toxic and corrosive fluorine seems to be solvable and could be done by inert storage vessels.
A space simulation procedure was developed for studying rocket plume contamination effects using a 5-pound bipropellant reaction control system thruster. Vacuum chamber pressures of 3 x 10 to the minus 5 torr (70 miles altitude) were achieved with the thruster firing in pulse trains consisting of eight pulses (50 msec on, 100 msec off, and seven minutes between pulse trains). The final vacuum was achieved by cooling all vacuum chamber surfaces to liquid helium temperature and by introducing a continuous argon leak of 48 std. cc/sec into the test chamber. An effort was made to simulate propellant system flow dynamics corresponding to actual spacecraft mission use. Fast time response liquid flow rate measurements showed that large variations occurred in the ratio of oxidizer to fuel flow for pulse-on times up to 120 msec. These variations could lead to poor combustion efficiency and the production of contamination.
The combustion of metals is a field with important practical applications in rocket propellants, high-temperature flames, and material synthesis. Also, the safe operation of metal containers in high-pressure oxygen systems and with cryogenic fuels and oxidizers remains an important concern in industry. The increasing use of metallic components in spacecraft and space structures has also raised concerns about their flammability properties and fire suppression mechanisms. In addition, recent efforts to embark on unmanned and manned planetary exploration, such as on Mars, have also renewed the interest in metal/carbon-dioxide combustion as an effective in situ resource utilization technology. In spite of these practical applications, the understanding of the combustion properties of metals remains far behind that of the most commonly used fuels such as hydrocarbons. The lack of understanding is due to the many problems unique to metal- oxidizer reactions such as: low-temperature surface oxidation prior to ignition, heterogeneous reactions, very high combustion temperatures, product condensation, high emissivity of products, and multi-phase interactions. Very few analytical models (all neglecting the influence of gravity) have been developed to predict the burning characteristics and the flame structure details. Several experimental studies attempting to validate these models have used small metal particles to recreate gravity-free conditions. The high emissivity of the flames, rapid reaction, and intermittent explosions experienced by these particles have made the gathering of any useful information on burning rates and flame structure very difficult. The use of a reduced gravity environment is needed to clarify some of the complex interactions among the phenomena described above. First, the elimination of the intrusive buoyant flows that plague all combustion phenomena is of paramount importance in metal reactions due to the much higher temperatures reached during combustion. Second, a low-gravity environment is absolutely essential to remove the destructive effect of gravity on the shape of a molten metal droplet in order to study a spherically symmetric condition with large bulk samples. The larger size of the spherical metal droplet and the longer burning times available in reduced gravity extend the spatial and temporal dimensions to permit careful probing of the flame structure and dynamics. Third, the influence of the radiative heat transfer from the solid oxides can be studied more carefully by generating a stagnant spherical shell of condensed products undisturbed by buoyancy.
The East Test Area at Marshall Space Flight Center has five major test stands, each of which has two or more test positions, not counting the SSME and RD-180 engine test facilities in the West Test Area. These research and development facilities are capable of testing high pressure pumps, both fuel and oxidizer, injectors, chambers and sea-level engine assemblies, as well as simulating deep space environments in the 12, 15 and 20 foot vacuum chambers. Liquid propellant capabilities are high pressure hydrogen (liquid and gas), methane (liquid and gas), and RP-1 and high pressure LOX. Solid propellant capability includes thrust measurement and firing capability up to 1/6 scale Shuttle SRB segment. In the past six months MSFC supported multiple space access and exploration programs in the previous six months. Major programs were Space Exploration, Shuttle External Tank research, Reusable Solid Rocket Motor (RSRM) development, as well as research programs for NASA and other customers. At Test Stand 115 monopropellant ignition testing was conducted on one position. At the second position multiple ignition/variable burn time cycles were conducted on Vacuum Plasma Spatter (VPS) coated injectors. Each injector received fifty cycles; the propellants were LOX Hydrogen and the ignition source was TEA. Following completion of the monopropellant test series the stand was reconfigured to support ignition testing on a LOX Methane injector system. At TS 116 a thrust stand used to test Booster Separation Motors from the Shuttle SRB system was disassembled and moved from Chemical Systems Division s Coyote Canyon plant to MSFC. The stand was reassembled and readied for BSM testing. Also, a series of tests was run on a Pratt & Whitney Rocketdyne Low Element Density (LED) injector engine. The propellants for this engine are LOX and LH2. At TS 300 the 20 foot vacuum chamber was configured to support hydrogen testing in the Multipurpose Hydrogen Test Bed (MHTB) test article. This testing, which went 24/7 for fourteen consecutive days, demonstrated long duration storage methods intended to minimize losses of propellant in support of the Space Exploration Initiative. The facility is being converted to support similar research using liquid methane. The 12 foot chamber at TS 300 was used to create ascent profiles (both heat and altitude effects) for foam panel testing in support of the Shuttle External Tank program. At TS 500, one position was in build-up to support ATK Thiokol research into the gas dynamics associated with high pressure flow across the propellant joint in segmented solid rocket motors. The testing involves flowing high pressure gas through a 24 motor case. Initial tests will be conducted with simulated aluminum grain, followed by tests using actual propellant. The second position at TS 500 has been in build-up for testing a LOX methane thruster manufactured by KT Engineering. At the Solid Propulsion Test Area (SPTA), the first dual segment 24 solid rocket motor was fired for ATK Thiokol in support of the RSRM program. A new axial thrust measurement stand was designed and fabricated for this testing. Real Time Radiography (RTR) will be deployed to examine nozzle erosion on the next dual segment motor.
The Europa mission and spacecraft design presented unique challenges for selection of valve seat materials that met the fluid compatibility requirements, and combined fluid compatibility and high radiation exposure level requirements. The Europa spacecraft pressurization system valves will be exposed to fully saturated propellant vapor for the duration of the mission. The effects of Nitrogen Tetroxide (NTO) and Monomethylhydrazine (MMH) propellant vapors on heritage valve seat materials, such as Vespel SP-1 and Polychlorotrifluoroethylene (PCTFE), were evaluated to determine if an alternate material is required. In liquid system applications, Teflon is the only available compatible valve seat material. Radiation exposure data for Teflon in an air or vacuum environment has been previously documented. Radiation exposure data for Teflon in an oxidizer environment such as NTO, was not available, and it was unknown whether the effects would be similar to those on air-exposed samples. Material testing was conducted by Marshall Space Flight Center (MSFC) and White Sands Test Facility (WSTF) to determine the effects of propellant vapor on heritage seat materials for pressurization valve applications, and the effects of combined radiation and NTO propellant exposure on Teflon. The results indicated that changes in heritage pressurization valve seat materials' properties rendered them unsuitable for the Europa application. The combined radiation and NTO exposure testing of Teflon produced results equivalent to combined radiation and air exposure results.
The ram accelerator, a chemically-propelled mass driver, is presented as a new approach for directly launching acceleration-insensitive pay-loads into LEO. The cargo vehicle resembles the centerbody of a conventional ramjet and travels through a launch tube filled with a premixed gaseous fuel and oxidizer mixture. The tube acts as the outer cowling of the ramjet and the combustion process travels with the vehicle. Two modes of ram accelerator drive are described, which when used in sequence, are capable of accelerating the cargo vehicle to 10 km/sec. The requirements for placing a 2000 kg vehicle with 50 percent payload fraction into a 400 km orbit, with a minimum of on-board rocket propellant for circularization maneuvers, are examined. It is shown that aerodynamic heating during atmospheric transit results in very little ablation of the nose. Both direct and indirect orbital insertion scenarios are investigated, and a three-step maneuver consisting of two burns and aerobraking is found to minimize the on-board propellant mass. A scenario involving a parking orbit below the desired final orbit is suggested as a means to increase the flexibility of the mass launch concept.
A quantitative evaluation methodology for utilizing in-situ resources on Mars for the production of useful substances. The emphasis is on the chemical processes. Various options considering different feedstock (mostly, carbon dioxide, water, and iron oxides) are carefully examined for the product mix and the energy needs. Oxygen, carbon monoxide, alcohols, and other chemicals are the end products. The chemical processes involve electrolysis, methanation, and variations. It is shown that maximizing the product utility is more important than the production of oxygen, methane, or alcohols. An important factor is the storage of the chemicals produced. The product utility is dependent, to some extent, upon the mission. A combination of the stability, the enthalpy of formation, and the mass fraction of the products is seen to yield a fairly good quantitative feel for the overall utility and maximum mission impact.
Oxygen was identified as the most important product of initial lunar materials processing efforts. A source of oxygen on the Moon provides an alternative to the costly transport of propellant to the Moon or to low earth orbit. Pyrolysis, or vapor-phase reduction, involves heating a feedstock to temperatures sufficient to decompose the constituent metal oxides and release oxygen. The process relies on the vaporization of metal oxides in the form of reduced suboxides or atomic species. The reduced species must then be condensed without re-oxidizing, yielding oxygen in the gas phase. The feasibility of obtaining oxygen from common lunar minerals was demonstrated using solar furnace experiments. These results are discussed together with chemical equilibrium models which were extended to include the multicomponent oxides used in experiments. For the first time, both experiments and theoretical models dealt with the complex oxides that make up potential lunar feedstocks. Two major conclusions are drawn from this preliminary work. First, unbeneficiated regolith is a suitable feedstock for pyrolysis. Second, the process can operate at moderate temperatures, circa 2000 K, which could be supplied by direct solar or electrical energy. In addition to these advantages in choice of feedstock and energy source, the pyrolysis process requires no chemicals or reagents, making it an attractive process for lunar oxygen production.
All spacecraft require propulsion systems for thrust and maneuvering. Propulsion systems can be chemical, nuclear, electrical, cold gas or combinations thereof. Chemical propulsion has proven to be the most reliable technology since the deployment of launch vehicles. Performance, storability, and handling are three important aspects of liquid chemical propulsion. Bipropellant systems require a fuel and an oxidizer for propulsion, but monopropellants only require a fuel and a catalyst for propulsion and are therefore simpler and lighter. Hydrazine is the state of the art propellant for monopropellant systems, but has drawbacks because it is highly hazardous to human health, which requires extensive care in handling, complex ground ops due to safety and environmental considerations, and lengthy turnaround times for reusable spacecraft. All users of hydrazine monopropellant must contend with these issues and their associated costs. The development of a new monopropellant, intended to replace hydrazine, has been in progress for years. This project will apply advanced techniques to characterize the engineering properties of materials used in AF-M315E propulsion systems after propellant exposure. AF-M315E monopropellant has been selected HQ's Green Propellant Infusion Mission (GPIM) to replace toxic hydrazine for improved performance and reduce safety and health issues that will shorten reusable spacecraft turn-around time. In addition, this project will fundamentally strengthen JSC's core competency to evaluate, use and infuse liquid propellant systems.
NASA Marshall Space Flight Center’s Materials and Processes Department, with support from the Propulsion Systems Department has developed a test bed for exposing ablative thermal protection materials to an environment similar to that of solid rocket motors (SRMs) by using a hybrid-propellant combustor. The Solid Fuel Torch (SFT) is a pressurized. small-scale hybrid rocket motor which utilizes an aluminized solid fuel and a gaseous oxygen (GOx) oxidizer to vary internal environments which can be tailored to specific program needs. The baseline design for the SFT was developed to support the Space Shuttle Program and was redesigned to support Space Launch System (SLS) Booster program requirements. There are four configurations of the SFT testbed: 1.) Insulation configuration with the option of a “dome” or “converging cone” insulation test section, 2.) Super Sonic Blast Tube configuration to simulate nozzle exit cone environments, 3.) Flow Straightened configuration which includes an additional instrumented length between the fuel grain and the insulation, and 4.) Constant Diameter Throat configuration which includes a cylindrical phenolic test section to test samples with different ply angles under the same environment simultaneously. Motor pressures for each configuration can range from 150 to 1500 psi with heat fluxes up to 1200 BTU/sq.ft/s. This paper discusses the development of each configuration and the results obtained during the initial test series.