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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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At least 541 records · Page 30

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NASA NTRS1989

Arcjet Cathode Phenomena

Cathode tips made from a number of different materials were tested in a modular arcjet thruster in order to examine cathode phenomena. Periodic disassembly and examination, along with the data collected during testing, indicated that all of the tungsten-based materials behaved similarly despite the fact that in one of these samples the percentage of thorium oxide was doubled and another was 25 percent rhenium. The mass loss rate from a 2 percent thoriated rhenium cathode was found to be an order of magnitude greater than that observed using 2 percent thoriated tungsten. Detailed analysis of one of these cathode tips showed that the molten crater contained pure tungsten to a depth of about 150 microns. Problems with thermal stress cracking were encountered in the testing of a hafnium carbide tip. Post test analysis showed that the active area of the tip had chemically reacted with the propellant. A 100 hour continuous test was run at about 1 kW. Post test analysis revealed no dendrite formation, such as observed in a 30 kW arcjet lifetest, near the cathode crater. The cathodes from both this test and a previously run 1000 hour cycled test displayed nearly identical arc craters. Data and calculations indicate that the mass losses observed in testing can be explained by evaporation.

Curran, Francis M.↗

NASA NTRS1973

Design for pressure regulating components

The design development for Pressure Regulating Components included a regulator component trade-off study with analog computer performance verification to arrive at a final optimized regulator configuration for the Space Storable Propulsion Module, under development for a Jupiter Orbiter mission. This application requires the pressure regulator to be capable of long-term fluorine exposure. In addition, individual but basically identical (for purposes of commonality) units are required for separate oxidizer and fuel pressurization. The need for dual units requires improvement in the regulation accuracy over present designs. An advanced regulator concept was prepared featuring redundant bellows, all metallic/ceramic construction, friction-free guidance of moving parts, gas damping, and the elimination of coil springs normally used for reference forces. The activities included testing of actual size seat/poppet components to determine actual discharge coefficients and flow forces. The resulting data was inserted into the computer model of the regulator. Computer simulation of the propulsion module performance over two mission profiles indicated satisfactory minimization of propellant residual requirements imposed by regulator performance uncertainties.

Wichmann, H.↗

NASA NTRS2001

ISRU Technologies for Mars Life Support

Life support systems can take advantage of elements in the atmosphere of Mars to provide for necessary consumables such as oxygen and buffer gas for makeup of leakage. In situ consumables production (ISCP) can be performed effectively in conjunction with in situ propellant production, in which oxygen and methane are manufactured for rocket fuel. This project considers ways of achieving the optimal system objectives from the two sometimes competing objectives of ISPP and ISCP. In previous years we worked on production of a nitrogen-argon buffer gas as a by- product of the CO2 acquisition and compression system. Recently we have been focusing on combined electrolysis of water vapor and carbon dioxide. Combined electrolysis of water vapor and carbon dioxide is essential for reducin,o the complexity of a combined ISPP/ISCP plant. Using a solid oxide electrolysis cell (SOEC) for this combined process would be most advantageous for it allows mainly gas phase reactions, O2 gas delivered from the electrolyzer is free of any H2O vapor, and SOE is already a proven technology for pure CO2 electrolysis. Combined SOEC testing is conducted at The University of Arizona in the Space Technologies Laboratory (STL) of the Aerospace and Mechanical Engineering Department.

Finn, John E.↗

NASA NTRS1983

A summary of recent developments in transportation hazard classification activities for ammonium perchlorate

The transportation hazard classification of Ammonium Perchlorate is discussed. A test program was completed and data were forwarded to retain a Class 5.1 designation (oxidizer) for AP which is shipped internationally. As a follow-on to the initial team effort to conduct AP tests existing data were examined and a matrix which catalogs test parameters and findings was compiled. A collection of test protocols is developed to standardize test methods for energetic materials of all types. The actions to date are summarized; the participating organizations and their roles as presently understood; specific findings on AP (matrix); and issues, lessons learned, and potential actions of particular interest to the propulsion community which may evolve as a result of future U.N. propellant transportation classification activities.

Koller, A. M., Jr.↗

NASA NTRS1991

Primary helium heater for propellant pressurization systems

The primary helium heater is a unique design that provides direct heating of pressurant gas for large pressure fed propulsion systems. It has been conceptually designed to supply a heated (800-1000 R) pressurization gas to both a liquid oxygen and an RP-1 propellant tank. This pressurization gas is generated within the heater by mixing super critical helium (40-300 R and 3000-1600 psi) with an appropriate amount of combustion products from a 4:1 throttling stoichiometric LO2/LH2 combustor. This simple, low cost and reliable mixer utilizes the large quantity of helium to provide stoichiometric combustor cooling, extend the throttling limits and enhance the combustion stability margin. Preliminary combustion, thermal, and CFD analyses confirm that this low-pressure-drop direct helium heater can provide the constant-temperature pressurant suitable for tank pressurization of both fuel and oxidizer tanks of large pressure fed vehicles.

Reichmuth, D. M.↗

NASA NTRS2020

Corrosion Behavior of Stainless Steel 304 and Nickel 625 Under Iodine Vapor at 300 °C

The chemical, structural, and microstructural behavior of stainless steel 304 (SS 304) and nickel 625 (Ni 625) were probed after exposure to iodine vapor laminar flow at 300 °C. This work was conducted in support of the development of in-space propulsion utilizing iodine as a propellant. The kinetics of corrosion was measured in a custom-built iodine-compatible thermogravimetric analyzer rig at 300 °C in which samples were exposed to an iodine laminar vapor flow of 1 mg/min (carried by 20 mL/min argon gas) for up to 31 days. Samples were characterized after the experiment by X-ray diffraction, electron microscopy, energy dispersive X-ray spectroscopy, and dynamic secondary ion mass spectrometry. All samples formed scales consisting mainly of metal oxyiodides showing different chemistry, microstructure, and crystalline phases. The paralinear rate law best describes the kinetics of oxidation of the samples. SS 304 exhibited slightly slower kinetics of oxidation (106±3·10(exp –5) mg/sq.cm/day) than Ni 625 (122±3·10(exp –5) mg/sq.cm/day) in the low-oxygen and low-water environment.

Chemistry and Materials↗

NASA NTRS2008

Liquid Oxygen/Liquid Methane Ascent Main Engine Technology Development

The National Aeronautics & Space Administration (NASA) has identified Liquid Oxygen (LO2)/Liquid Methane (LCH4) as a potential propellant combination for future space vehicles based upon the Exploration Systems Architecture Study (ESAS). The technology is estimated to have higher performance and lower overall systems mass compared to existing hypergolic propulsion systems. The current application considering this technology is the lunar ascent main engine (AME). AME is anticipated to be an expendable, pressure-fed engine to provide ascent from the moon at the completion of a 210 day lunar stay. The engine is expected to produce 5,500 lbf (24,465 N) thrust with variable inlet temperatures due to the cryogenic nature of the fuel and oxidizer. The primary technology risks include establishing reliable and robust ignition in vacuum conditions, maximizing specific impulse, developing rapid start capability for the descent abort, providing the capability for two starts and producing a total engine bum time over 500 seconds. This paper will highlight the efforts of the Marshall Space Flight Center (MSFC) in addressing risk reduction activities for this technology.

Robinson, Joel W.↗

DOE OSTI2024

Advances in electrosynthesis for a greener chemical industry

As nations unite to curb anthropogenic greenhouse gas emissions, the decarbonization of the chemical industry has been propelled to the forefront of scientific research. Renewable electricity will play a central role in this effort. In addition to powering chemical plants and allowing for the sustainable production of heat to drive thermocatalytic processes, renewable electricity also provides the chemical industry with opportunities to engage in the sustainability revolution and broadly reduce its environmental footprint through breakthrough innovations in direct electrochemical transformations. The electrification of chemical synthesis—electrosynthesis—is a promising route to promote sustainability without compromising economic competitiveness. Electrosynthesis uses electrons both as an energy source to drive reactions and as a green reagent for chemical reductions and oxidations under ambient conditions. Therefore, it holds tremendous potential (pun intended) to increase selectivity to desired products, open green reaction pathways for challenging transformations (e.g., Birch reduction, epoxidations, coupling reactions), and reduce chemical waste.

08 HYDROGEN↗

NASA NTRS1996

Augmentation of Rocket Propulsion: Physical Limits

Rocket propulsion is not ideal when the propellant is not ejected at a unique velocity in an inertial frame. An ideal velocity distribution requires that the exhaust velocity vary linearly with the velocity of the vehicle in an inertial frame. It also requires that the velocity distribution variance as a thermodynamic quantity be minimized. A rocket vehicle with an inert propellant is not optimal, because it does not take advantage of the propellant mass for energy storage. Nor is it logical to provide another energy storage device in order to realize variable exhaust velocity, because it would have to be partly unfilled at the beginning of the mission. Performance is enhanced by pushing on the surrounding because it increases the reaction mass and decreases the reaction jet velocity. This decreases the fraction of the energy taken away by the propellant and increases the share taken by the payload. For an optimal model with the propellant used as fuel, the augmentation realized by pushing on air is greatest for vehicles with a low initial/final mass ratio. For a typical vehicle in the Earth's atmosphere, the augmentation is seen mainly at altitudes below about 80 km. When drag is taken into account, there is a well-defined optimum size for the air intake. Pushing on air has the potential to increase the performance of rockets which pass through the atmosphere. This is apart from benefits derived from "air breathing", or using the oxygen in the atmosphere to reduce the mass of an on-board oxidizer. Because of the potential of these measures, it is vital to model these effects more carefully and explore technology that may realize their advantages.

Taylor, Charles R.↗

NASA NTRS2018

Study of a Tricarbide Grooved Ring Fuel Element for Nuclear Thermal Propulsion

Deep space exploration, especially that of Mars, is on the horizon as the next big challenge for space exploration. Nuclear propulsion, through which high thrust and efficiency can be achieved, is a promising option for decreasing the cost and logistics of such a mission. Work on nu- clear thermal engines goes back to the days of the NERVA program. Currently, nuclear thermal propulsion is under development again in various forms to provide a superior propulsion system for deep space exploration. The authors have been working to develop a concept nuclear thermal engine that uses a grooved ring fuel element as an alternative to the traditional hexagonal rod design. The authors are also studying the use of carbide fuels. The concept was developed in order to increase surface area and heat transfer to the propellant. The use of carbides would also raise the operating temperature of the reactor. It is hoped that this could lead to a higher thrust to weight nuclear thermal engine. This paper describes the modeling of neutronics, heat transfer, and fluid dynamics of this alternative nuclear fuel element geometry. Fabrication experiments of grooved rings from carbide refractory metals are also presented along with material characterization and interactions with a hot hydrogen environment. Results of experiments and associated analysis are desired densities with some success in material distribution and reaching a solid solution. Future work is needed to improve distribution of material, minimize oxidation during the milling process, and de ne a fabrication process that will serve for constructing grooved ring fuel rods for large system tests.

Taylor, Brian↗

NASA NTRS1974

The 15 cm diameter ion thruster research

The startup reliability of a 15 cm diameter mercury bombardment ion thruster which employs a pulsed high voltage tickler electrode on the main and neutralizer cathodes is examined. Startup of the thruster is achieved 100% of the time on the main cathode and 98.7% of the time on the neutralizer cathode over a 3640 cycle test. The thruster was started from a 20 C initial condition and operated for an hour at a 600 mA beam current. An energy efficiency of 75% and a propellant utilization efficiency of 77% was achieved over the complete cycle. The effect of a single cusp magnetic field thruster length on its performance is discussed. Guidelines are formulated for the shaping of magnetic field lines in thrusters. A model describing double ion production in mercury discharges is presented. The production route is shown to occur through the single ionic ground state. Photographs of the interior of an operating-hollow cathode are presented. A cathode spot is shown to be present if the cathode is free of low work-function surfaces. The spot is observed if a low work-function oxide coating is applied to the cathode insert. Results show that low work-function oxide coatings tend to migrate during thruster operation.

Wilbur, P. J.↗

NASA NTRS1988

High/variable mixture ratio O2/H2 engine

Vehicle/engine analysis studies have identified the High/Dual Mixture Ratio O2/H2 Engine cycle as a leading candidate for an advanced Single Stage to Orbit (SSTO) propulsion system. This cycle is designed to allow operation at a higher than normal O/F ratio of 12 during liftoff and then transition to a more optimum O/F ratio of 6 at altitude. While operation at high mixture ratios lowers specific impulse, the resultant high propellant bulk density and high power density combine to minimize the influence of atmospheric drag and low altitude gravitational forces. Transition to a lower mixture ratio at altitude then provides improved specific impulse relative to a single mixture ratio engine that must select a mixture ratio that is balanced for both low and high altitude operation. This combination of increased altitude specific impulse and high propellant bulk density more than offsets the compromised low altitude performance and results in an overall mission benefit. Two areas of technical concern relative to the execution of this dual mixture ratio cycle concept are addressed. First, actions required to transition from high to low mixture ratio are examined, including an assessment of the main chamber environment as the main chamber mixture ratio passes through stoichiometric. Secondly, two approaches to meet a requirement for high turbine power at high mixture ratio condition are examined. One approach uses high turbine temperature to produce the power and requires cooled turbines. The other approach incorporates an oxidizer-rich preburner to increase turbine work capability via increased turbine mass flow.

Adams, A.↗

NASA NTRS1993

Design of power electronics for TVC EMA systems

The Composite Development Division of the Propulsion Laboratory at Marshall Space Flight Center (MSFC) is currently developing a class of electromechanical actuators (EMA's) for use in space transportation applications such as thrust vector control (TVC) and propellant control valves (PCV). These high power servomechanisms will require rugged, reliable, and compact power electronic modules capable of modulating several hundred amperes of current at up to 270 volts. MSFC has selected the brushless dc motor for implementation in EMA's. This report presents the results of an investigation into the applicability of two new technologies, MOS-controlled thyristors (MCT's) and pulse density modulation (PDM), to the control of brushless dc motors in EMA systems. MCT's are new power semiconductor devices, which combine the high voltage and current capabilities of conventional thyristors and the low gate drive requirements of metal oxide semiconductor field effect transistors (MOSFET's). The commanded signals in a PDM system are synthesized using a series of sinusoidal pulses instead of a series of square pulses as in a pulse width modulation (PWM) system. A resonant dc link inverter is employed to generate the sinusoidal pulses in the PDM system. This inverter permits zero-voltage switching of all semiconductors which reduces switching losses and switching stresses. The objectives of this project are to develop and validate an analytical model of the MCT device when used in high power motor control applications and to design, fabricate, and test a prototype electronic circuit employing both MCT and PDM technology for controlling a brushless dc motor.

Nelms, R. Mark↗

NASA NTRS1987

Analysis of flow decay potential on Galileo

The risks posed to the NASA's Galileo spacecraft by the oxidizer flow decay during its extended mission to Jupiter is discussed. The Galileo spacecraft will use nitrogen tetroxide (NTO)/monomethyl hydrazine bipropellant system with one large engine thrust-rated at a nominal 400 N, and 12 smaller engines each thrust-rated at a nominal 10 N. These smaller thrusters, because of their small valve inlet filters and small injector ports, are especially vulnerable to clogging by iron nitrate precipitates formed by NTO-wetted stainless steel components. To quantify the corrosion rates and solubility levels which will be seen during the Galileo mission, corrosion and solubility testing experiments were performed with simulated Galileo materials, propellants, and environments. The results show the potential benefits of propellant sieving in terms of iron and water impurity reduction.

Cole, T. W.↗

NASA NTRS2006

Development and Implementation of Non-Newtonian Rheology Into the Generalized Fluid System Simulation Program (GFSSP)

One of the goals of this program was to develop the experimental and analytical/computational tools required to predict the flow of non-Newtonian fluids through the various system components of a propulsion system: pipes, valves, pumps etc. To achieve this goal we selected to augment the capabilities of NASA's Generalized Fluid System Simulation Program (GFSSP) software. GFSSP is a general-purpose computer program designed to calculate steady state and transient pressure and flow distributions in a complex fluid network. While the current version of the GFSSP code is able to handle various systems components the implicit assumption in the code is that the fluids in the system are Newtonian. To extend the capability of the code to non-Newtonian fluids, such as silica gelled fuels and oxidizers, modifications to the momentum equations of the code have been performed. We have successfully implemented in GFSSP flow equations for fluids with power law behavior. The implementation of the power law fluid behavior into the GFSSP code depends on knowledge of the two fluid coefficients, n and K. The determination of these parameters for the silica gels used in this program was performed experimentally. The n and K parameters for silica water gels were determined experimentally at CFDRC's Special Projects Laboratory, with a constant shear rate capillary viscometer. Batches of 8:1 (by weight) water-silica gel were mixed using CFDRC s 10-gallon gelled propellant mixer. Prior to testing the gel was allowed to rest in the rheometer tank for at least twelve hours to ensure that the delicate structure of the gel had sufficient time to reform. During the tests silica gel was pressure fed and discharged through stainless steel pipes ranging from 1", to 36", in length and three diameters; 0.0237", 0.032", and 0.047". The data collected in these tests included pressure at tube entrance and volumetric flowrate. From these data the uncorrected shear rate, shear stress, residence time, and viscosity were evaluated using formulae for non-Newtonian, power law fluids. The maximum shear rates (corrected for entrance effects) obtained in the rheometer with the current setup were in the 150,000 to 170,000sec- range. GFSSP simulations were performed with a flow circuit simulating the capillary rheometer and using Power Law gel viscosity coefficients from the experimental data. The agreement between the experimental data and the simulated flow curves was within +/-4% given quality entrance effect data.

DiSalvo, Roberto↗

NASA NTRS2018

Study of a Tricarbide Grooved Ring Fuel Element for Nuclear Thermal Propulsion

Deep space exploration, especially that of Mars, is on the horizon as the next big challenge for space exploration. Nuclear propulsion, through which high thrust and efficiency can be achieved, is a promising option for decreasing the cost and logistics of such a mission. Work on nuclear thermal engines goes back to the days of the NERVA program. Currently, nuclear thermal propulsion is under development again in various forms to provide a superior propulsion system for deep space exploration. The authors have been working to develop a concept nuclear thermal engine that uses a grooved ring fuel element as an alternative to the traditional hexagonal rod design. The authors are also studying the use of carbide fuels. The concept was developed in order to increase surface area and heat transfer to the propellant. The use of carbides would also raise the operating temperature of the reactor. It is hoped that this could lead to a higher thrust to weight nuclear thermal engine. This paper describes the modeling of neutronics, heat transfer, and fluid dynamics of this alternative nuclear fuel element geometry. Fabrication experiments of grooved rings from carbide refractory metals are also presented along with material characterization and interactions with a hot hydrogen environment. Results of experiments and associated analysis are discussed. The authors demonstrated success in reaching desired densities with some success in material distribution and reaching a solid solution. Future work is needed to improve distribution of material, minimize oxidation during the milling process, and define a fabrication process that will serve for constructing grooved ring fuel rods for large system tests.

Taylor, Brian↗

NASA NTRSDate not supplied

Advanced Diagnostics for Megahertz Imaging of Mixing, Fuel Spray, and Combustion Processes for Rotating Detonation Combustors

Recent advancements in megahertz rate, high-power, burst-mode laser technology are leveraged to perform and explore imaging measurements that spatially and temporally resolve the mixing, combustion, and detonation flow field in two laboratory-scale rotating detonation combustors (RDCs). In a non-premixed annular RDC, multiple imaging diagnostics are explored to investigate gaseous and liquid injector behavior, the detonation wave structure, and the propellant refill. In one instance, OH planar laser-induced fluorescence (OH-PLIF) imaging is performed up to a 2 MHz repetition rate to track the combustion products and reaction zone locations. In the same annular RDC, a single liquid fuel jet is injected, and laser-based 355-nm imaging of the fuel spray is performed up to a 1 MHz repetition rate. For this configuration, the annular RDC is used as a detonation driver to impose periodic detonation waves to interact with the fuel spray. Moreover, in this annular RDC, a range of tracer-based laser imaging measurements are explored to time-resolve the unsteady oxidizer air recovery and refill process. In a non-premixed linear RDC, planar imaging measurements of the fuel mixing are performed up to a 200 kHz repetition rate using PLIF of a tracer in the fuel supply. The fuel mixing imaging helps explain the origin of the observed pre and post wave burning, detonation structure, and enables quantifying injector recovery timescales. This paper will provide a high-level broad survey of diagnostics applied in these RDCs, lessons learned, and interesting observations.

Combustion↗

NASA NTRSDate not supplied

Advanced Diagnostics For Megahertz Imaging Of Mixing, Fuel Spray, And Combustion Processes For Rotating Detonation Combustors

Combustion↗