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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 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS2009

Liquid Oxygen Rotating Friction Ignition Testing of Aluminum and Titanium with Monel and Inconel for Rocket Engine Propulsion System Contamination Investigation

Metallic contaminant was found in the liquid oxygen (LOX) pre-valve screen of the shuttle main engine propulsion system on two orbiter vehicles. To investigate the potential for an ignition, NASA Johnson Space Center White Sands Test Facility performed (modified) rotating friction ignition testing in LOX. This testing simulated a contaminant particle in the low-pressure oxygen turbo pump (LPOTP) and the high-pressure oxygen turbo pump (HPOTP) of the shuttle main propulsion system. Monel(R) K-500 and Inconel(R) 718 samples represented the LPOTP and HPOTP materials. Aluminum foil tape and titanium foil represented the contaminant particles. In both the Monel(R) and Inconel(R) material configurations, the aluminum foil tape samples did not ignite after 30 s of rubbing. In contrast, all of the titanium foil samples ignited regardless of the rubbing duration or material configuration. However, the titanium foil ignitions did not propagate to the Monel and Inconel materials.

Peralta, S.↗
NASA NTRS1984

Evaluation of propellent tank insulation concepts for low-thrust chemical propulsion systems: Executive summary

Cryogenic propellant tank insulations or liquid oxygen/liquid hydrogen low-thrust 2224N (500 lbf) propulsion systems (LTPS) were assessed. The insulation studied consisted of combinations of N2-purged foam and multilayer insulation (MLI) as well as He-purged MLI-only. Heat leak and payload performance predictions were made for three shuttle-launched LTPS designed for shuttle bay packaged payload densities of 56 kg cu/m (3.5 lbm/cu ft), 40 kg/cu m (2.5 lbm/cu ft) and 24 kg/cu m (1.5 lbm/cu ft). Foam/MLI insulations were found to increase LTPS payload delivery capability when compared with He-purged MLI-only. An additional benefit of foam/MLI was reduced operational complexity because orbiter cargo bay N2 purge gas could be used for MLI purging. Maximum payload mass benefit occurred when an enhanced convection, rather than natural convection, heat transfer was specified for the insulation purge enclosure. The enhanced convection environment allowed minimum insulation thickness to be used for the foam/MLI interface temperature selected to correspond to the moisture dew point in the N2 purge gas. Experimental verification of foam/MLI benefits was recommended. A conservative program cost estimate for testing a MLI-foam insulated tank was 2.1 million dollars. This cost could be reduced significantly without increasing program risk.

Kramer, T.↗
NASA NTRS1973

Study of advanced techniques for determining the long term performance of components

The application of existing and new technology to the problem of determining the long-term performance capability of liquid rocket propulsion feed systems is discussed. The long term performance of metal to metal valve seats in a liquid propellant fuel system is stressed. The approaches taken in conducting the analysis are: (1) advancing the technology of characterizing components through the development of new or more sensitive techniques and (2) improving the understanding of the physical of degradation.

Source record↗
NASA NTRS1983

Discriminating Between Liquid and Gas Flows

Flow sensor distinguishes flowing gases from liquids. Sensor allows liquids to pass, but would signal valve to turn off flow of gases. Concept developed for propulsion systems in which liquid fuel is forced out of storage tank by high-pressure gas. Controller stops flow of gas from fuel tank after liquid has been depleted.

Berdahl, C. M.↗
NASA NTRS1986

Mission and vehicle sizing sensitivities

Representative interplanetary space vehicle systems are sized to compare and show sensitivity of the initial mass required in low Earth orbit to one mission mode and mission opportunity. Data are presented to show the requirements for Earth-Mars opposition and conjunction class roundtrip flyby and stopover mission opportunities available during the time period from year 1997 to year 2045. The interplanetary space vehicle consists of a spacecraft and a space vehicle acceleration system. Propellant boil-off for the various mission phases is given for the Lox/LH (Liquid Oxygen/Liquid Hydrogen) propulsion systems. Mission abort information is presented for the 1999 Venus outbound swingby trajectory, transfer profile.

Archie C Young↗
NASA NTRS1991

Design issues for propulsion systems using metallized propellants

Metallized propellants are liquid propellants that contain metal particles. These particles are suspended in a gelled fuel or oxidizer. Aluminum is used as the metal additive. The addition of metal to conventional propellants can increase their specific impulse and their density over conventional propellants, and consequently, the payload delivered on Mars and lunar transportation vehicles, Earth-to-Orbit vehicles and upper stages for robotic planetary missions. Gelled fuels also provide increased safety during accidental propellant leakage or spills. To take full advantage of these potential performance increases, there are changes that must be made to the vehicle design. The differences are discussed between metallized propellant and traditional liquid propellants and their effect on the propulsion system design. These differences include the propellant density, mixture ratio, engine performance, and propellant rheology. Missions related to the Space Exploration Initiative are considered as design examples to illustrate these issues.

Palaszewski, Bryan↗
NASA NTRS1991

Design issues for propulsion systems using metallized propellants

Metallized propellants are liquid propellants that contain metal particles. These particles are suspended in a gelled fuel or oxidizer. Aluminum is used as the metal additive. The addition of metal to conventional propellants can increase their specific impulse and their density over conventional propellants, and consequently, the payload delivered on Mars and lunar transportation vehicles, Earth-to-Orbit vehicles and upper stages for robotic planetary missions. Gelled fuels also provide increased safety during accidental propellant leakage or spills. To take full advantage of these potential performance increases, there are changes that must be made to the vehicle design. The differences are discussed between metallized propellant and traditional liquid propellants and their effect on the propulsion system design. These differences include the propellant density, mixture ratio, engine performance, and propellant rheology. Missions related to the Space Exploration Initiative are considered as design examples to illustrate these issues.

Palaszewski, Bryan↗
NASA NTRS2011

Liquid Oxygen/Liquid Methane Propulsion and Cryogenic Advanced Development

Exploration Systems Architecture Study conducted by NASA in 2005 identified the liquid oxygen (LOx)/liquid methane (LCH4) propellant combination as a prime candidate for the Crew Exploration Vehicle Service Module propulsion and for later use for ascent stage propulsion of the lunar lander. Both the Crew Exploration Vehicle and Lunar Lander were part the Constellation architecture, which had the objective to provide global sustained lunar human exploration capability. From late 2005 through the end of 2010, NASA and industry matured advanced development designs for many components that could be employed in relatively high thrust, high delta velocity, pressure fed propulsion systems for these two applications. The major investments were in main engines, reaction control engines, and the devices needed for cryogenic fluid management such as screens, propellant management devices, thermodynamic vents, and mass gauges. Engine and thruster developments also included advanced high reliability low mass igniters. Extensive tests were successfully conducted for all of these elements. For the thrusters and engines, testing included sea level and altitude conditions. This advanced development provides a mature technology base for future liquid oxygen/liquid methane pressure fed space propulsion systems. This paper documents the design and test efforts along with resulting hardware and test results.

Klem, Mark D.↗
NASA NTRS1998

Preliminary Assessment of Using Gelled and Hybrid Propellant Propulsion for VTOL/SSTO Launch Systems

A novel, reusable, Vertical-Takeoff-and-Vertical-Takeoff-and-Landing, Single-Stage-to-Orbit (VTOL/SSTO) launch system concept, named AUGMENT-SSTO, is presented in this paper to help quantify the advantages of employing gelled and hybrid propellant propulsion system options for such applications. The launch vehicle system concept considered uses a highly coupled, main high performance liquid oxygen/liquid hydrogen (LO2/LH2) propulsion system, that is used only for launch, while a gelled or hybrid propellant propulsion system auxiliary propulsion system is used during final orbit insertion, major orbit maneuvering, and landing propulsive burn phases of flight. Using a gelled or hybrid propellant propulsion system for major orbit maneuver burns and landing has many advantages over conventional VTOL/SSTO concepts that use LO2/LH2 propulsion system(s) burns for all phases of flight. The applicability of three gelled propellant systems, O2/H2/Al, O2/RP-1/Al, and NTO/MMH/Al, and a state-of-the-art (SOA) hybrid propulsion system are examined in this study. Additionally, this paper addresses the applicability of a high performance gelled O2/H2 propulsion system to perform the primary, as well as the auxiliary propulsion system functions of the vehicle.

Palaszewski, Bryan↗
NASA NTRS2005

Electromagnetic Pumps for Conductive-Propellant Feed Systems

Prototype electromagnetic pumps for use with lithium and bismuth propellants were constructed and tested. Such pumps may be used to pressurize future electric propulsion liquid metal feed systems, with the primary advantages being the compactness and simplicity versus alternative pressurization technologies. Design details for two different pumps are described: the first was designed to withstand (highly corrosive) lithium propellant, and t he second was designed to tolerate the high temperature required to pump liquid bismuth. Both qualitative and quantitative test results are presented. Open-loop tests demonstrated the capability of each device to electromagnetically pump its design propellant (lithium or bismuth). A second set of tests accurately quantified the pump pressure developed as a function of current. These experiments, which utilized a more easily handled material (gallium), demonstrated continuously-adjustable pump pressure levels ranging from 0-100 Torr for corresponding input current levels of 0-75 A. While the analysis and testing in this study specifically targeted lithium and bismuth propellants, the underlying design principles should be useful in implementing liquid metal pumps in any conductive-propellant feed system.

Markusic, Thomas E.↗
NASA NTRS2014

Advanced Compatibility Characterization Of AF-M315E With Spacecraft Propulsion System Materials Project

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.

McClure, Mark B.↗
NASA NTRSDate not supplied

On the Equivalence of Gas and Liquid Leak Flow Rates for Application to Propulsion and Propellant Transfer Systems

The equivalence of leak rates for gas flow and liquid flow through an arbitrary small leak size in pressurized systems has been investigated theoretically and experimentally. Measurements were taken for gaseous helium (GHe) rates and water flow issuing from the same leak configuration. The predictions for the correlation of GHe leak rates to water leak rates have been satisfactorily compared with present test data and those of other investigators over a wide range leak flow rates (many orders of magnitude) corresponding to a wide range of leak sizes (0.1 m to 0.1 mm). The results show that at extremely small leak sizes (sub-micron range), the effects of Knudsen number in the free molecular regime (rarefied gases) become important. By a proper scaling procedure, the measured flow rate for water could be applied to equivalent propellant leaks in flight operations. In the present tests the micro-channel range of the leak size is satisfactorily simulated by repeatable precision set needle position valve. Test results show that in general the theoretical model predicts the data reasonably well for GHe flow rates in the range of 0.001 to 1 sccs. Outside this range of GHe flow rates some adjustments (improvements) to the analytical model are required along with further testing.

Max Kandula↗
NASA NTRS2012

System Engineering and Technical Challenges Overcome in the J-2X Rocket Engine Development Project

Beginning in 2006, NASA initiated the J-2X engine development effort to develop an upper stage propulsion system to enable the achievement of the primary objectives of the Constellation program (CxP): provide continued access to the International Space Station following the retirement of the Space Station and return humans to the moon. The J-2X system requirements identified to accomplish this were very challenging and the time expended over the five years following the beginning of the J- 2X effort have been noteworthy in the development of innovations in both the fields for liquid rocket propulsion and system engineering.

Ballard, Richard O.↗
NASA NTRS1981

Analysis of low-g heat transfer for a cryogenic oxygen pressurization heater

After the external tank separates from the Orbiter about 2000 pounds of residual liquid oxygen remain in the main propulsion system lines. The pressurization of liquid oxygen from a subcritical to a supercritical state by the use of the heaters of the PRSA tanks while in a low-g environment is investigated. The performance of the heaters while bringing the state of the substance from the subcritical state to the supercritical one is studied, with particular emphasis on the time the pressurization process takes, and the temperature of the heater as the process proceeds.

Rios, J.↗
NASA NTRS2010

Liquid Rocket Engine Testing - Historical Lecture: Simulated Altitude Testing at AEDC

The span of history covered is from 1958 to the present. The outline of this lecture draws from historical examples of liquid propulsion testing done at AEDC primarily for NASA's Marshall Space Flight Center (NASA/MSFC) in the Saturn/Apollo Program and for USAF Space and Missile Systems dual-use customers. NASA has made dual use of Air Force launch vehicles, Test Ranges and Tracking Systems, and liquid rocket altitude test chambers / facilities. Examples are drawn from the Apollo/ Saturn vehicles and the testing of their liquid propulsion systems. Other examples are given to extend to the family of the current ELVs and Evolved ELVs (EELVs), in this case, primarily to their Upper Stages. The outline begins with tests of the XLR 99 Engine for the X-15 aircraft, tests for vehicle / engine induced environments during flight in the atmosphere and in Space, and vehicle staging at high altitude. The discussion is from the author's perspective and background in developmental testing.

Dougherty, N. S.↗
NASA NTRS2016

Additive Manufacturing a Liquid Hydrogen Rocket Engine

Space Propulsion is a 5 day event being held from 2nd May to the 6th May 2016 at the Rome Marriott Park Hotel in Rome, Italy. This event showcases products like Propulsion sub-systems and components, Production and manufacturing issues, Liquid, Solid, Hybrid and Air-breathing Propulsion Systems for Launcher and Upper Stages, Overview of current programmes, AIV issues and tools, Flight testing and experience, Technology building blocks for Future Space Transportation Propulsion Systems : Launchers, Exploration platforms & Space Tourism, Green Propulsion for Space Transportation, New propellants, Rocket propulsion & global environment, Cost related aspects of Space Transportation propulsion, Modelling, Pressure-Thrust oscillations issues, Impact of new requirements and regulations on design etc. in the Automotive, Manufacturing, Fabrication, Repair & Maintenance industries.

Jones, Carl P.↗
NASA NTRS2019

Extreme-Temperature Carbon- and Ceramic-Matrix Composite Nozzle Extensions for Liquid Rocket Engines

The United States (US) National Aeronautics and Space Administration (NASA) and its US industry partners are developing extreme-temperature composite nozzle extensions for a variety of cryogenic liquid rocket engine propulsion systems. Applications under consideration at the NASA Marshall Space Flight Center (MSFC) include launch vehicle (upper stage), in-space, and lunar lander descent/ascent propulsion systems. Composite material systems of interest include carbon-carbon (C-C) and carbon/silicon-carbide (C-SiC) composites, as well as modified versions of such composites (coatings, mixed matrices, oxidation inhibitors, etc.). Missions addressed include access to low Earth orbit, the lunar surface, and more distant destinations. Cryogenic upper-stage and in-space liquid rocket engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures, increasing exhaust velocities. As a result of the large size of such nozzles, and the related engine performance requirements, composite nozzle extensions are being considered to reduce mass impacts. Currently, metallic and foreign composite nozzle extensions used on the Atlas V, Delta IV, Falcon 9, and Ariane 5 launch vehicles represent the state-of-the-art. Such extensions are limited to use at or slightly above 2000°F (1093°C). Materials under development by MSFC and its industry partners have the potential to operate at temperatures up to (or above) 4250°F (2343°C). Marshall Space Flight Center efforts are aimed at (a) further developing the technology and databases needed to enable the use of composite nozzle extensions on cryogenic liquid rocket engines, (b) developing and demonstrating low-cost capabilities for testing and qualifying such nozzle extensions, and (c) advancing the US domestic supply chain for nozzle extensions. Through the Small Business Innovation Research (SBIR) and Small Business Technology Transfer (STTR) Programs, low-level internal MSFC research projects, and teaming arrangements with domestic industry partners, composite material design, development, hot-fire test, and evaluation efforts are progressing. Hot- fire engine tests of subscale hardware have been conducted using oxygen/hydrogen (LOX/H2), oxygen/methane (LOX/CH4), and oxygen/kerosene (LOX/RP-1) propellants. These engine tests at MSFC have enabled the evaluation of both heritage and state-of-the-art material systems, demonstrating the initial capabilities of the extreme temperature materials and their fabrication methods. Recent, ongoing, and potential future work supporting cryogenic propulsion systems development will be presented. The composite nozzle extension technology and test capabilities being developed are intended to support both NASA and US Department of Defense requirements, as well as those of the broader Commercial Space industry.

Valentine, Peter G.↗
NASA NTRS2014

Green Mono Propulsion Activities at MSFC

In 2012, the National Aeronautics & Space Administration (NASA) Space Technology Mission Directorate (STMD) began the process of building an integrated technology roadmap, including both technology pull and technology push strategies. Technology Area 1 (TA-01) for Launch Propulsion Systems and TA-02 In-Space Propulsion are two of the fourteen TAs that provide recommendations for the overall technology investment strategy and prioritization of NASA's space technology activities. Identified within these documents are future needs of green propellant use. Green ionic liquid monopropellants and propulsion systems are beginning to be demonstrated in space flight environments. Starting in 2010 with the flight of Prisma, a 1-N thruster system began on-orbit demonstrations operating on ammonium dinitramide based propellant. The NASA Green Propellant Infusion Mission (GPIM) plans to demonstrate both 1-N, and 22-N hydroxyl ammonium nitrate (HAN)-based thrusters in a 2015 flight demonstration. In addition, engineers at MSFC have been evaluating green propellant alternatives for both thrusters and auxiliary power units (APUs). This paper summarizes the status of these development/demonstration activities and investigates the potential for evolution of green propellants from small spacecraft and satellites to larger spacecraft systems, human exploration, and launch system auxiliary propulsion applications.

Robinson, Joel W.↗