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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 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1962

A Survey and Evaluation of High Energy Liquid Chemical Propulsion Systems: Propellant Selection Criteria for Space Missions - Part 1

This report presents the results of a study to develop a procedure for evaluating liquid propellants in order (a) to select the most appropriate propellant (from among those under development) for each of several applications on each of the various missions in the NASA program, or (b) to select new propellants (from among those being proposed) for initiation or continuation of research and development. The analysis begins with a consideration of requirements--either for the specific application or for the various classes of applications. The known characteristics of the propellant or propellants to be evaluated are then put into a convenient form for evaluation. The next step is to determine whether or not there are requirements that simply cannot be met by the propellant. If the propellant passes this test, an optimum vehicle configuration using the propellant (and meeting all requirements) is estimated. (The configuration should be optimized with respect to the total resource consumption for all aspects of the mission, including R&D, production, logistics, and operation.) The total resource consumption for this configuration is then compared with that for similar configurations using other propellants (and meeting all requirements equally well). If all factors have been properly taken into account, this comparison of resource consumption will complete the evaluation. Such an evaluation may be performed several times, in increasing detail and with correspondingly increasing accuracy, as an R&D program proceeds, and the accuracy of the data as well as the cost of the next step in the program increase. The procedure is superior to those in common use in that it minimizes both the amount of analytical work and the number of points at which subjective value judgments are made.

LIQUID PROPELLANT↗
NASA NTRS1982

Molecular absorption cryogenic cooler for liquid hydrogen propulsion systems

A light weight, long life molecular absorption cryogenic cooler (MACC) system is described which can use low temperature waste heat to provide cooling for liquid hydrogen propellant tanks for interplanetary spacecraft. Detailed tradeoff studies were made to evaluate the refrigeration system component interactions in order to minimize the mass of the spacecraft cooler system. Based on this analysis a refrigerator system mass of 31 kg is required to provide the .48 watts of cooling required by a 2.3 meter diameter liquid hydrogen tank.

Klein, G. A.↗
NASA NTRS2015

NASA Propulsion Concept Studies and Risk Reduction Activities for Resource Prospector Lander

The Resource Prospector mission is to investigate the Moon's polar regions in search of volatiles. The government-version lander concept for the mission is composed of a braking stage and a liquid-propulsion lander stage. A propulsion trade study concluded with a solid rocket motor for the braking stage while using the 4th-stage Peacekeeper (PK) propulsion components for the lander stage. The mechanical design of the liquid propulsion system was conducted in concert with the lander structure design. A propulsion cold-flow test article was fabricated and integrated into a lander development structure, and a series of cold flow tests were conducted to characterize the fluid transient behavior and to collect data for validating analytical models. In parallel, RS-34 PK thrusters to be used on the lander stage were hot-fire tested in vacuum conditions as part of risk reduction activities.

Trinh, Huu P.↗
NASA NTRS2014

Liquid and Solid Propulsion Systems Attributes - Unique, Common and Complementary

In this study, attributes are described for solid and liquid propulsion systems based on historical data. This study is not intended to compare liquid and solid propulsion system attributes, rather to present options for their use in various mission scenarios. US launch vehicle data from 1970 to 2008 was analyzed to assess solid and liquid propulsion development cost and schedule characteristics, performance features, and safety and mission success attributes. The study assessed historical trends for liquid and solid systems, and investigated implications of those trends. It was found that the two propulsion technologies have unique, common and complementary attributes that can be leveraged to meet mission requirements.

Cannon, James L.↗
NASA NTRS2024

Welding of Crack Sensitive Aluminum Alloys for Liquid Rocket Propulsion Applications

In this welding test campaign, electron beam welding (EBW) and manual gas tungsten arc welding (GTAW) processes were utilized to join an additively manufactured aluminum alloy nozzle to a wrought aluminum flange. Both aluminum alloys were 6061, which in their conventional form are susceptible to solidification cracking issues. The AM nozzle utilized a novel high-performance modified version of 6061 which exhibits crack resistant properties during the AM process and during welding. When manually conducting GTAWs a filler wire with the same modified version of 6061 was utilized to weld different combinations of 6061 coupons (either modified or unmodified). Due to the grain refinement technology in the modified 6061, either as filler material or base material, there was complete elimination of weld solidification cracking. During EBW development a combination of wrought 6061 with a modified and additively manufactured 6061 was investigated. The EBWs developed exhibited no crack like defects. Each weld development program utilized non-destructive evaluation, metallography, and mechanical testing to determine the final weld acceptance. Mechanical testing involved hardness evaluation and tension testing on welded coupons. Both welding techniques were successfully developed and exhibited adequate properties for the final test articles and met all requirements of the development campaign. These welds were ultimately developed to be utilized on a set of all-aluminum liquid rocket nozzles to demonstrate additively manufactured aluminum alloys for propulsion applications. These aluminum nozzles were tested at NASA Marshall Space Flight Center’s hot fire test facility and subsequently underwent a test campaign involving a variety of chamber pressures and liquid propellant combinations. Following the hot fire test campaign each weldment was investigated for post operation performance and any defects that may have evolved during testing.

Welding↗
NASA NTRS1991

A subscale facility for liquid rocket propulsion diagnostics at Stennis Space Center

The Diagnostics Testbed Facility (DTF) at NASA's John C. Stennis Space Center in Mississippi was designed to provide a testbed for the development of rocket engine exhaust plume diagnostics instrumentation. A 1200-lb thrust liquid oxygen/gaseous hydrogen thruster is used as the plume source for experimentation and instrument development. Theoretical comparative studies have been performed with aerothermodynamic codes to ensure that the DTF thruster (DTFT) has been optimized to produce a plume with pressure and temperature conditions as much like the plume of the Space Shuttle Main Engine as possible. Operation of the DTFT is controlled by an icon-driven software program using a series of soft switches. Data acquisition is performed using the same software program. A number of plume diagnostics experiments have utilized the unique capabilities of the DTF.

Raines, N. G.↗
NASA NTRS2008

Summary of Propulsion System Needs in Support of Project Constellation

In January 2004, the President of the United States established the Vision for Space Exploration (VSE) to return man to the moon and ultimately to extend manned space travel to Mars. This paper will summarize the manned space flight liquid propulsion system needs in support of Project Constellation over the next 10 years. It will include all engine needs to return man to the moon. An overview of engines currently under contract, those baselined but not yet under contract, and those engine needs that hav.e yet to be initiated. Project Constellation includes the components as shown Figure 1. Liquid propulsion systems supporting the manned portion of these elements include the following: the Crew Exploration Vehicle named Orion (crew module reaction control system (CMRCS), service module Orion Main Engine (OME), service module auxiliary RCS, and service module reaction control system (SMRCS)), the Crew Launch Vehicle named Ares 1 (J2X upper stage, first stage roll control system, second stage reaction control system, and the Ares I-X roll control system), the Heavy Lift Launch Vehicle named Ares V (RS68B first stage booster, J-2X upper stage, roll control systems, and the Earth Departure Stage (EDS) (powered by the same Ares V Upper Stage J-2X), and the Lunar Lander named Altair with both descent and ascent stages (lunar orbit insertion and descent main engine, ascent main engine, and attitude control systems for both stages). In addition, there may be additional engine needs for early demonstrators, but those will not be speculated on as part of this paper. Also, other portions of the VSE architecture, including the planned Orion abort demonstrations and the Lunar Precursor Robotic Program, are not addressed here as they either use solid motors or are focused on unmanned precursor missions.

Sumrall, Phil↗
NASA NTRS2004

Liquid Rocket Propulsion for Atmospheric Flight in the Proposed ARES Mars Scout Mission

Flying above the Mars Southern Highlands, an airplane will traverse over the terrain of Mars while conducting unique science measurements of the atmosphere, surface, and interior. This paper describes an overview of the ARES (Aerial Regional-scale Environmental Survey) mission with an emphasis on airplane propulsion needs. The process for selecting a propulsion system for the ARES airplane is also included. Details of the propulsion system, including system schematics, hardware and performance are provided. The airplane has a 6.25 m wingspan with a total mass of 149 kg and is propelled by a bi-propellant liquid rocket system capable of carrying roughly 48 kg of MMH/MON3 propellant.

Kuhl, Christopher A.↗
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 NTRS2010

Non-Pyrotechnic Zero-Leak Normally Closed Valve

This valve is designed to create a zero-leak seal in a liquid propulsion system that is a functional replacement for the normally closed pyrovalve. Unlike pyrovalves, Nitinol is actuated by simply heating the material to a certain temperature, called the transition temperature. Like a pyrovalve, before actuation, the upstream and downstream sections are separated from one another and from the external environment by closed welded seals. Also like pyrovalves, after actuation, the propellant or pressurant gas can flow without a significant pressure drop but are still separated from the external environment by a closed welded seal. During manufacture, a Nitinol bar is compressed to 93 percent of its original length and fitted tightly into the valve. During operation, the valve is heated until the Nitinol reaches the transition temperature of 95 C; the Nitinol "remembers" its previous longer shape with a very large recovery force causing it to expand and break the titanium parent metal seal to allow flow. Once open, the valve forever remains open. The first prototype valve was designed for high pressure [5,000 psi (=34.5 MPa)] and low flow, typical requirements for pressurant gas valves in liquid propulsion systems. It is possible to modify the dimensions to make low-pressure models or high-flow models, for use downstream of the propellant tanks. This design is simpler, lower risk, and less expensive than the pyrovalve. Although the valve must be in a thermally controlled state (kept below 80 C) to prevent premature actuation, the pyrovalves and electrically actuated initiators have far more taxing handling requirements.

Gillespie, Rebecca↗
NASA NTRS1993

Liquid rocket propulsion impeller CFD modeling

Steady-state impeller geometric modeling and typical Navier-Stokes CFD algorithm analysis procedures are assessed using two benchmark quality impeller data sets. Two geometric modeling and grid generation software packages, ICEM-CFD and PATRAN, are considered. Results show that a significant advantage of PATRAN's open-ended architecture is the potential interaction between CFD and structural/thermal analysts inside the mechanical computer-aided engineering environment. However the time required to construct the inducer grid would be unacceptable in a design and engineering environment. The ICEM-CFD package is considered to be more appropriate for structural grid generation but lacks the mature link to structural/thermal analysis arena as compared to PATRAN.

Ratcliff, Mark L.↗
NASA NTRS1993

Liquid rocket propulsion dynamic flow modeling using the ROCETS engineering modules in the EASY5x environment

A report on the progress of porting the ROCETS (ROCket Engine Transient Simulator) into the EASY5x simulation environment is presented. Brief descriptions of each of the software systems, information regarding the actual port process, and examples comparing the results of the two systems are given. It is shown that EASY5x is a suitable environment for utilization of the ROCETS engineering modules, and that, for the example systems shown, EASY5x actually seems to give more accurate solutions than the straight ROCETS code.

Follett, Randolph F.↗
NASA NTRS1999

Liquid Rocket Propulsion: RS-68 and Linear Aerospike

This paper presents viewgraphs of the RS-68 and Linear Aerospike engines. The topics include: 1) The mission; 2) Design and Development Strategies; 3) Cost Driven Design Trades; 4) Moderate Pressure Requirements Enable Reduced Complexity; 5) Experience Based Design; 6) Development Costs Dominated by Cost of Eliminating Failure Modes; and 7) Reducing Development Costs. This paper also includes the design of the RS-68 Engine and the engine performance.

Murphy, Terry↗
NASA NTRS2005

Fifth International Symposium on Liquid Space Propulsion

Contents include the fiollowing: Theme: Life-life Combustion Devices Technology. Technical Sessions: International Perspectives. System Level Effects. Component Level Processes. Material Considerations. Design Environments -- Predictions. Injector Design Technology. Design Environments -- Measurements. Panel Discussion: Views on future research and development needs and Symposium observations. Aquarium Welcome and Southern Belle Riverboat Recognition Banquet evening events.

Garcia, R.↗