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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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS2011

Small Scale Turbopump Manufacturing Technology and Material Processes

As part of an internal research and development project, NASA Marshall Space Flight Center (MSFC) has been developing a high specific impulse 9,000-lbf LOX/LH2 pump-fed engine testbed with the capability to throttle 10:1. A Fuel Turbopump (FTP) with the ability to operate across a speed range of 30,000-rpm to 100,000-rpm was developed and analyzed. This small size and flight-like Fuel Turbopump has completed the design and analysis phase and is currently in the manufacturing phase. This paper highlights the manufacturing and processes efforts to fabricate an approximately 20-lb turbopump with small flow passages, intricately bladed components and approximately 3-in diameter impellers. As a result of the small scale and tight tolerances of the hardware on this turbopump, several unique manufacturing and material challenges were encountered. Some of the technologies highlighted in this paper include the use of powder metallurgy technology to manufacture small impellers, electron beam welding of a turbine blisk shroud, and casting challenges. The use of risk reduction efforts such as non-destructive testing (NDT) and evaluation (NDE), fractography, material testing, and component spin testing are also discussed in this paper.

Alvarez, Erika↗
NASA NTRS1995

Dual-Fuel Propulsion in Single-Stage Advanced Manned Launch System Vehicle

As part of the United States Advanced Manned Launch System study to determine a follow-on, or complement, to the Space Shuttle, a reusable single-stage-to-orbit concept utilizing dual-fuel rocket propulsion has been examined. Several dual-fuel propulsion concepts were investigated. These include: a separate-engine concept combining Russian RD-170 kerosene-fueled engines with space shuttle main engine-derivative engines: the kerosene- and hydrogen-fueled Russian RD-701 engine; and a dual-fuel, dual-expander engine. Analysis to determine vehicle weight and size characteristics was performed using conceptual-level design techniques. A response-surface methodology for multidisciplinary design was utilized to optimize the dual-fuel vehicles with respect to several important propulsion-system and vehicle design parameters, in order to achieve minimum empty weight. The tools and methods employed in the analysis process are also summarized. In comparison with a reference hydrogen- fueled single-stage vehicle, results showed that the dual-fuel vehicles were from 10 to 30% lower in empty weight for the same payload capability, with the dual-expander engine types showing the greatest potential.

Lepsch, Roger A., Jr.↗
NASA NTRS1991

Thermal analysis of conceptual designs for GPHS/FPSE power systems of 250 We and 500 We

Thermal analyses were performed for two distinct configurations of a proposed space nuclear power system which combines General Purpose Heat Source (GPHS) modules with the state of the art Free-Piston Stirling Engines (FPSEs). The two configurations correspond to systems with power levels of 250 and 500 W(sub e). The 250 W(sub e) GPHS/FPSE power system utilizes four GPHS modules and one FPSE, and the 500 W(sub e) contains eight GPHS modules and two FPSEs. The configurations of the systems and the bases for selecting the configurations are described. Brief introductory sections are included to describe the GPHS modules and free piston Stirling engines. The primary focus of the thermal analyses is on the temperature of the iridium fuel clad within the GPHS modules. A design goal temperature of 1573 K was selected as the upper limit for the fuel clad during normal operating conditions. The basis for selecting this temperature limit is discussed in detail. Results obtained from thermal analysis of the 250 W(sub e) GPHS/FPSE power system indicate fuel clad temperatures which slightly exceed the design goal temperature of 1573 K. The results are considered favorable due to the numerous conservative assumptions used in developing the thermal model and performing the thermal analysis. To demonstrate the effects of the conservatism, a brief sensitivity analysis is performed in which a few of the key system parameters are varied to determine their effect on the fuel clad temperatures. It is concluded that thermal analysis of a more detailed thermal model would be expected to yield fuel clad temperatures below the design foal temperature limiy 1573 K.

Mccomas, Thomas J.↗
NASA NTRS2015

Affordable Development and Demonstration of a Small NTR Engine and Stage: How Small is Big Enough?

In FY11, NASA formulated a plan for Nuclear Thermal Propulsion (NTP) development that included Foundational Technology Development followed by system-level Technology Demonstrations The ongoing NTP project, funded by NASAs Advanced Exploration Systems (AES) program, is focused on Foundational Technology Development and includes 5 key task activities:(1) Fuel element fabrication and non-nuclear validation testing of heritage fuel options;(2) Engine conceptual design;(3) Mission analysis and engine requirements definition;(4) Identification of affordable options for ground testing; and(5) Formulation of an affordable and sustainable NTP development program Performance parameters for Point of Departure designs for a small criticality-limited and full size 25 klbf-class engine were developed during FYs 13-14 using heritage fuel element designs for both RoverNERVA Graphite Composite (GC) and Ceramic Metal (Cermet) fuel forms To focus the fuel development effort and maximize use of its resources, the AES program decided, in FY14, that a leader-follower down selection between GC and cermet fuel was required An Independent Review Panel (IRP) was convened by NASA and tasked with reviewing the available fuel data and making a recommendation to NASA. In February 2015, the IRP recommended and the AES program endorsed GC as the leader fuel In FY14, a preliminary development schedule DDTE plan was produced by GRC, DOE industry for the AES program. Assumptions, considerations and key task activities are presented here Two small (7.5 and 16.5 klbf) engine sizes were considered for ground and flight technology demonstration within a 10-year timeframe; their ability to support future human exploration missions was also examined and a recommendation on a preferred size is provided.

spacecraft design↗
NASA NTRS2015

Affordable Development and Demonstration of a Small NTR Engine and Stage: A Preliminary NASA, DOE, and Industry Assessment

In FY11, NASA formulated a plan for Nuclear Thermal Propulsion (NTP) development that included Foundational Technology Development followed by system-level Technology Demonstrations. The ongoing NTP project, funded by NASA's Advanced Exploration Systems (AES) program, is focused on Foundational Technology Development and includes 5 key task activities: (1) Fuel element fabrication and non-nuclear validation testing of heritage fuel options; (2) Engine conceptual design; (3) Mission analysis and engine requirements definition; (4) Identification of affordable options for ground testing; and (5) Formulation of an affordable and sustainable NTP development program. Performance parameters for Point of Departure designs for a small criticality-limited and full size 25 klb (sub f) (kip, or 1000 pound-force)-class engine were developed during FYs 13-14 using heritage fuel element designs for both Rover NERVA (Nuclear Engine for Rocket Vehicle Applications) Graphite Composite (GC) and Ceramic Metal (Cermet) fuel forms. To focus the fuel development effort and maximize use of its resources, the AES program decided, in FY14, that a leader-follower down-selection between GC and cermet fuel was required. An Independent Review Panel (IRP) was convened by NASA and tasked with reviewing the available fuel data and making a recommendation to NASA. In February 2015, the IRP recommended and the AES program endorsed GC as the leader fuel In FY14, a preliminary development schedule DDTE (Design Development Test and Evaluation) plan was produced by GRC (Glenn Research Center), DOE (Department of Energy) and industry for the AES program. Assumptions, considerations and key task activities are presented here. At the direction of NASA HQ (32515), NASA and DOE are to work together to formulate a detailed development plan and schedule allowing the affordable development of a small (7.5-16.5 klbf) GC engine for possible flight technology demonstration (FTD) mission within a 10-year timeframe.

Lunar Flyby Mission↗
NASA NTRS2000

Isentropic Compression of Multicomponent Mixtures of Fuels and Inert Gases

In selected aerospace applications of the fuels hydrazine and monomethythydrazine, there occur conditions which can result in the isentropic compression of a multicomponent mixture of fuel and inert gas. One such example is when a driver gas such as helium comes out of solution and mixes with the fuel vapor, which is being compressed. A second example is when product gas from an energetic device mixes with the fuel vapor which is being compressed. Thermodynamic analysis has shown that under isentropic compression, the fuels hydrazine and monomethylhydrazine must be treated as real fluids using appropriate equations of state. The appropriate equations of state are the Peng-Robinson equation of state for hydrazine and the Redlich-Kwong-Soave equation of state for monomethylhydrazine. The addition of an inert gas of variable quantity and input temperature and pressure to the fuel compounds the problem for safety design or analysis. This work provides the appropriate thermodynamic analysis of isentropic compression of the two examples cited. In addition to an entropy balance describing the change of state, an enthalpy balance is required. The presence of multicomponents in the system requires that appropriate mixing rules are identified and applied to the analysis. This analysis is not currently available.

Barragan, Michelle↗
NASA NTRSDate not supplied

Current Test Capabilities of NASA Langley’s Arc-Heated Scramjet Test Facility

The Arc-Heated Scramjet Test Facility at NASA Langley Research Center has demonstrated itself repeatedly as a key contributor to fundamental hypersonic research and scramjet engine technology development in the Mach 4.7 to 8 range. Since its inception in the 1970’s, the facility has hosted numerous scramjet engine configurations that explored the performance and operability of airframe-integrated hydrogen-fueled scramjet flow path concepts, successfully maturing both analysis tools and design technique and methodology for hypersonic propulsion. Contributions to major programs include extensive tests of the hydrogen-fueled NASP and Hyper-X scramjet engines, and the hydrocarbon-fueled HIFiRE Flight 2 scramjet flow path, in addition to fundamental in-house research. The facility capabilities have expanded to include direct-connect flow path support, gaseous hydrocarbon fuels, dynamic trajectory simulation and extension of flight simulation down to Mach 2.2. The low-Mach upgrades include increased air flow rate and controls capability and two new low-Mach nozzles (Mach 2.2 and 3.5). The facility recently completed verification of the new operational range through successful envelope expansion tests with the Mach 2.2 nozzle. With these new modifications, the operating envelope for the facility is now Mach 2.2 – 8.0(+), with a demonstrated dynamic pressure simulation of 2000 psf at the low end, and margin available for even more. Lastly, a continuous run time of over 5 minutes has been demonstrated, indicating that run time is limited only by test hardware or back-pressure requirements. This paper will provide a detailed description of these previously unreported upgrades and expanded simulation capabilities.

Hypersonic↗
NASA NTRS1978

Design and analysis of a plate-fin sandwich actively cooled structural panel

The skin structure of hydrogen fueled hypersonic transport vehicles traveling at Mach 6 and above must be designed to withstand, for relatively long periods of time, the aerodynamic heating effects which are far more severe than those encountered by the supersonic aircraft of today. The use of conventional aircraft materials such as aluminum in combination with forced convection active cooling to accommodate aerodynamic heating is addressed. The basic active cooling concept consists of a stringer stiffened plate-fin sandwich. The sandwich surface is subjected to the aerodynamic heat flux which is transferred, via convection, to a coolant that is forced through the sandwich under pressure. The coolant, in turn, circulates in a closed loop through a hydrogen heat exchanger and back through the skin panel.

Smith, L. M.↗
NASA NTRS1980

Initial characterization of an Experimental Referee Broadened-Specification (ERBS) aviation turbine fuel

Characterization data and a hydrocarbon compositional analysis are presented for a research test fuel designated as an experimental referee broadened-specification aviation turbine fuel. This research fuel, which is a special blend of kerosene and hydrotreated catalytic gas oil, is a hypothetical representation of a future fuel should it become necessary to broaden current kerojet specifications. It is used as a reference fuel in research investigations into the effects of fuel property variations on the performance and durability of jet aircraft components, including combustors and fuel systems.

Prok, G. M.↗
NASA NTRS1983

Experimental design of an interlaboratory study for trace metal analysis of liquid fluids

The accurate determination of trace metals and fuels is an important requirement in much of the research into and development of alternative fuels for aerospace applications. Recognizing the detrimental effects of certain metals on fuel performance and fuel systems at the part per million and in some cases part per billion levels requires improved accuracy in determining these low concentration elements. Accurate analyses are also required to ensure interchangeability of analysis results between vendor, researcher, and end use for purposes of quality control. Previous interlaboratory studies have demonstrated the inability of different laboratories to agree on the results of metal analysis, particularly at low concentration levels, yet typically good precisions are reported within a laboratory. An interlaboratory study was designed to gain statistical information about the sources of variation in the reported concentrations. Five participant laboratories were used on a fee basis and were not informed of the purpose of the analyses. The effects of laboratory, analytical technique, concentration level, and ashing additive were studied in four fuel types for 20 elements of interest. The prescribed sample preparation schemes (variations of dry ashing) were used by all of the laboratories. The analytical data were statistically evaluated using a computer program for the analysis of variance technique.

Greenbauer-Seng, L. A.↗
NASA NTRS2002

Three Dimensional CFD Analysis of the GTX Combustor

The annular combustor geometry of a combined-cycle engine has been analyzed with three-dimensional computational fluid dynamics. Both subsonic combustion and supersonic combustion flowfields have been simulated. The subsonic combustion analysis was executed in conjunction with a direct-connect test rig. Two cold-flow and one hot-flow results are presented. The simulations compare favorably with the test data for the two cold flow calculations; the hot-flow data was not yet available. The hot-flow simulation indicates that the conventional ejector-ramjet cycle would not provide adequate mixing at the conditions tested. The supersonic combustion ramjet flowfield was simulated with frozen chemistry model. A five-parameter test matrix was specified, according to statistical design-of-experiments theory. Twenty-seven separate simulations were used to assemble surrogate models for combustor mixing efficiency and total pressure recovery. ScramJet injector design parameters (injector angle, location, and fuel split) as well as mission variables (total fuel massflow and freestream Mach number) were included in the analysis. A promising injector design has been identified that provides good mixing characteristics with low total pressure losses. The surrogate models can be used to develop performance maps of different injector designs. Several complex three-way variable interactions appear within the dataset that are not adequately resolved with the current statistical analysis.

Steffen, C. J., Jr.↗
NASA NTRS1989

Meteoroid and orbital debris shielding on the Orbital Maneuvering Vehicle

NASA's Orbital Maneuvering Vehicle (OMV) is being designed to withstand a 10-year lifetime in polar and low earth orbits. A large percentage of OMV's lifetime will be spent operating in the vicinity of the Space Shuttle and Space Station or in storage at these manned locations. An extensive analysis has been performed to determine the effects of the meteoroid and orbital debris environments on OMV's external fuel tanks. A finite element model of OMV was constructed using NASTRAN and analyzed with the meteoroid and debris design analysis code BUMPER. The results show that the long design lifetime, and the ever increasing man-made orbital debris environment, will require the use of shielding over the external fuel tanks.

Kirkpatrick, Marc E.↗
NASA NTRS2008

Forebody and Inlet Design for the HIFiRE 2 Flight Test

A forebody and inlet have been designed for the HIFiRE 2 scramjet flight test. The test will explore the operating, performance, and stability characteristics of a simple hydrocarbon-fueled scramjet combustor as it transitions from dual-mode to scramjet-mode operation and during supersonic combustion at Mach 8+ flight conditions. Requirements for the compression system were derived from inlet starting and combustor inflow requirements as well as physical size constraints. The design process is described. A planar, fixed geometry, mixed compression concept was used to produce laterally uniform flow at the inlet entrance and a conservative amount of internal contraction with respect to inlet starting. A grid sensitivity study was performed so that important flow physics caused by three-dimensional shock boundary layer interactions could be captured with confidence. Results from low Mach number operability studies, nominal trajectory cases, and high dynamic pressure heat load cases are discussed. The forebody and inlet solutions provide information for on-going combustor calculations, mass capture across the trajectory for fuel system design, and surface heating rates for thermal/structural analysis. The design has a one freestream Mach number margin for inlet starting, exceeds the high Mach number combustor entrance pressure requirement, produces high quality flow at the inlet exit for all Mach numbers and vehicle attitudes in the design space, and fits inside the booster shroud.

Ferlemann, Paul G.↗
NASA NTRS2008

Lunar Surface Access Module Descent Engine Turbopump Technology: Detailed Design

The need for a high specific impulse LOX/LH2 pump-fed lunar lander engine has been established by NASA for the new lunar exploration architecture. Studies indicate that a 4-engine cluster in the thrust range of 9,000-lbf each is a candidate configuration for the main propulsion of the manned lunar lander vehicle. The lander descent engine will be required to perform minor mid-course corrections, a Lunar Orbit Insertion (LOI) burn, a de-orbit burn, and the powered descent onto the lunar surface. In order to achieve the wide range of thrust required, the engines must be capable of throttling approximately 10:1. Working under internal research and development funding, NASA Marshall Space Flight Center (MSFC) has been conducting the development of a 9,000-lbf LOX/LH2 lunar lander descent engine testbed. This paper highlights the detailed design and analysis efforts to develop the lander engine Fuel Turbopump (FTP) whose operating speeds range from 30,000-rpm to 100,000-rpm. The capability of the FTP to operate across this wide range of speeds imposes several structural and dynamic challenges, and the small size of the FTP creates scaling and manufacturing challenges that are also addressed in this paper.

Alarez, Erika↗
NASA NTRS2010

Lunar Surface Access Module Descent Engine Turbopump Technology: Detailed Design

The need for a high specific impulse LOX/LH2 pump-fed lunar lander engine has been established by NASA for the new lunar exploration architecture. Studies indicate that a 4-engine cluster in the thrust range of 9,000-lbf each is a candidate configuration for the main propulsion of the manned lunar lander vehicle. The lander descent engine will be required to perform multiple burns including the powered descent onto the lunar surface. In order to achieve the wide range of thrust required, the engines must be capable of throttling approximately 10:1. Working under internal research and development funding, NASA Marshall Space Flight Center (MSFC) has been conducting the development of a 9,000-lbf LOX/LH2 lunar lander descent engine technology testbed. This paper highlights the detailed design and analysis efforts to develop the lander engine Fuel Turbopump (FTP) whose operating speeds range from 30,000-rpm to 100,000-rpm. The capability of the FTP to operate across this wide range of speeds imposes several structural and dynamic challenges, and the small size of the FTP creates scaling and manufacturing challenges that are also addressed in this paper.

Alvarez, Erika↗
NASA NTRS2002

Three Dimensional CFD Analysis of the GTX Combustor

The annular combustor geometry of a combined-cycle engine has been analyzed with three-dimensional computational fluid dynamics. Both subsonic combustion and supersonic combustion flowfields have been simulated. The subsonic combustion analysis was executed in conjunction with a direct-connect test rig. Two cold-flow and one hot-flow results are presented. The simulations compare favorably with the test data for the two cold flow calculations; the hot-flow data was not yet available. The hot-flow simulation Indicates that the conventional ejector-ramjet cycle would not provide adequate mixing at the conditions tested. The supersonic combustion ramjet flowfield was simulated with frozen chemistry model. A five-parameter test matrix was specified, according to statistical design-of-experiments theory. Twenty-seven separate simulations were used to assemble surrogate models for combustor mixing efficiency and total pressure recovery. Scramjet injector design parameters (injector angle, location, and fuel split) as well as mission variables (total fuel mass flow and freestream Mach number) were included in the analysis. A promising injector design has been identified that provides good mixing characteristics with low total pressure losses. The surrogate models can be used to develop performance maps of different injector designs. Several complex three-way variable interactions appear within the dataset that are not adequately resolved with the current statistical analysis.

Steffen, C. J., Jr.↗
NASA NTRS1972

Scramjet fuel injector design parameters and considerations: Development of a two-dimensional tangential fuel injector with constant pressure at the flame

The factors affecting a tangential fuel injector design for scramjet operation are reviewed and their effect on the efficiency of the supersonic combustion process is evaluated using both experimental data and theoretical predictions. A description of the physical problem of supersonic combustion and method of analysis is followed by a presentation and evaluation of some standard and exotic types of fuel injectors. Engineering fuel injector design criteria and hydrogen ignition schemes are presented along with a cursory review of available experimental data. A two-dimensional tangential fuel injector design is developed using analyses as a guide in evaluating the effects on the combustion process of various initial and boundary conditions including splitter plate thickness, injector wall temperature, pressure gradients, etc. The fuel injector wall geometry is shaped so as to maintain approximately constant pressure at the flame as required by a cycle analysis. A viscous characteristics program which accounts for lateral as well as axial pressure variations due to the mixing and combustion process is used in determining the wall geometry.

Agnone, A. M.↗