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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 NTRS1995

Metallized Gelled Propellants: Oxygen/RP-1/aluminum Rocket Combustion Experiments

A series of combustion experiments were conducted to measure the specific impulse, Cstar-, and specific-impulse efficiencies of a rocket engine using metallized gelled liquid propellants. These experiments used a small 20- to 40-1bf (89- to 178-N) thrust, modular engine consisting of an injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt% loadings of aluminum and gaseous oxygen was the oxidizer. Ten different injectors were used during the testing: 6 for the baseline 02/RP-1 tests and 4 for the gelled fuel tests which covered a wide range of mixture ratios. At the peak of the Isp versus oxidizer-to-fuel ratio (O/F) data, a range of 93 to 99% Cstar efficiency was reached with ungelled 02/RP-1. A Cstar efficiency range of 75 to 99% was obtained with gelled RP-l (0-wt% RP-1/Al) while the metallized 5-wt% RP-1/Al delivered a Cstar efficiency of 94 to 99% at the peak Isp in the O/F range tested. An 88 to 99% Cstar efficiency was obtained at the peak Isp of the gelled RP1/Al with 55-wt% Al. Specific impulse efficiencies for the 55-wt% RP-1/Al of 67%-83% were obtained at a 2.4:1 expansion ratio. Injector erosion was evident with the 55-wt% testing, while there was little or no erosion seen with the gelled RP-1 with 0- and 5-wt% Al. A protective layer of gelled fuel formed in the firings that minimized the damage to the rocket injector face. This effect may provide a useful technique for engine cooling. These experiments represent a first step in characterizing the performance of and operational issues with gelled RP-1 fuels.

Palaszewski, Bryan↗
DOE OSTI2020

Linking constituent phase properties to ductility and edge stretchability of two DP 980 steels

Two DP 980 steels were characterized with multiple methods to determine its microstructure characteristics, tensile properties and edge stretchability. The first (DP980-B) is a commercial grade DP980 steel produced by AK Steel, while the second (DP980-T) is a modification of the first one by a tempering process. Both contain two individual phases: ferrite and martensite with body centered cubic (BCC) and tetragonal (BCT) crystal structure respectively. ASTM standard tensile tests were performed for both materials, revealing that DP980-B has higher work hardening and larger total elongation than DP980-T. Standard hole piercing and subsequent hole expansion and extrusion tests were also performed. The critical hole expansion ratio (HER) of hole pierced sheet at different nominal clearances, on the other hand, show an opposite trend: the HER of DP980-T almost doubles that of DP980-B. This indicates the tensile ductility and edge stretchability are determined by different mechanisms due to different deformation modes. The ductility of sheet metal during uniaxial tension is determined by deformation instability, i.e. necking, due to less constraint at the region of deformation and fracture will occur quickly after necking. The contribution of post-necking deformation to the total elongation is insignificant for the high strength grades such as DP980 steels. The ductility during hole-expansion, however, is mainly determined by material fracture behavior that is dependent on phase property disparity and material intrinsic fracture parameters. With the individual phase properties for both DP steels determined by in situ tensile tests under high energy X-ray diffraction, the results show that DP980-T steel has lower property disparities between the two phases than DP980-B steel, due to the tempering process. This explains why the DP980-T steel has higher HER than DP980-B steel: lower phase disparity will lead to less local deformation during loading. The higher work hardening rate for DP980-B steel contributes to its higher uniform elongation compared with DP980-T steel based on the maximum load condition of deformation instability. An integrated finite element simulation framework for studying hole expansion is also presented here, based on the calculated individual phase properties from the combined high energy x-ray diffraction (HEXRD) and elastic plastic self-consistent modeling. The simulation results correlate well with the experimental results on the HER difference between the two materials.

36 MATERIALS SCIENCE↗
NASA NTRS2004

Hybrid Automotive Engine Using Ethanol-Burning Miller Cycle

A proposed hybrid (internal-combustion/ electric) automotive engine system would include as its internal-combustion subsystem, a modified Miller-cycle engine with regenerative air preheating and with autoignition like that of a Diesel engine. The fuel would be ethanol and would be burned lean to ensure complete combustion. Although the proposed engine would have a relatively low power-to-weight ratio compared to most present engines, this would not be the problem encountered if this engine were used in a non-hybrid system since hybrid systems require significantly lower power and thus smaller engines than purely internal-combustion-engine-driven vehicles. The disadvantage would be offset by the advantages of high fuel efficiency, low emission of nitrogen oxides and particulate pollutants, and the fact that ethanol is a renewable fuel. The original Miller-cycle engine, named after its inventor, was patented in the 1940s and is the basis of engines used in some modern automobiles, but is not widely known. In somewhat oversimplified terms, the main difference between a Miller-cycle engine and a common (Otto-cycle) automobile engine is that the Miller-cycle engine has a longer expansion stroke while retaining the shorter compression stroke. This is accomplished by leaving the intake valve open for part of the compression stroke, whereas in the Otto cycle engine, the intake valve is kept closed during the entire compression stroke. This greater expansion ratio makes it possible to extract more energy from the combustion process without expending more energy for compression. The net result is greater efficiency. In the proposed engine, the regenerative preheating would be effected by running the intake air through a heat exchanger connected to the engine block. The regenerative preheating would offer two advantages: It would ensure reliable autoignition during operation at low ambient temperature and would help to cool the engine, thereby reducing the remainder of the power needed for cooling and thereby further contributing to efficiency. An electrical resistance air preheater might be needed to ensure autoignition at startup and during a short warmup period. Because of the autoignition, the engine could operate without either spark plugs or glow plugs. Ethanol burns relatively cleanly and has been used as a motor fuel since the invention of internal-combustion engines. However, the energy content of ethanol per unit weight of ethanol is less than that of Diesel fuel or gasoline, and ethanol has a higher heat of vaporization. Because the Miller cycle offers an efficiency close to that of the Diesel cycle, burning ethanol in a Miller-cycle engine gives about as much usable output energy per unit volume of fuel as does burning gasoline in a conventional gasoline automotive engine. Because of the combination of preheating, running lean, and the use of ethyl alcohol, the proposed engine would generate less power per unit volume than does a conventional automotive gasoline engine. Consequently, for a given power level, the main body of the proposed engine would be bulkier. However, because little or no exhaust cleanup would be needed, the increase in bulk of the engine could be partially offset by the decrease in bulk of the exhaust system. The regenerative preheating also greatly reduces the external engine cooling requirement, and would translate to reduced engine bulk. It may even be possible to accomplish the remaining cooling of the engine by use of air only, eliminating the bulk and power consumption of a water cooling system. The combination of a Miller-cycle engine with regenerative air preheating, ethyl alcohol fuel, and hybrid operation could result in an automotive engine system that satisfies the need for a low pollution, high efficiency, and simple engine with a totally renewable fuel.

Weinstein, Leonard↗
NASA NTRS2002

Quasi-One-Dimensional Modeling of Pulse Detonation Rocket Engines

Pulsed detonation rocket engines (PDREs) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred a great deal of interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the difficulties in comparing the available experimental measurements with numerical models. In a previous paper by the author, parametric studies of the performance of a single, straight-tube PDRE were reported. A 1-D, unsteady method of characteristics code, employing a constant-gamma assumption behind the detonation front, was developed for that study. Models of this type are computationally inexpensive, and are particularly useful for parametric performance comparisons. For example, a plot showing the specific impulse of various PDRE and steady-state rocket engine (SSRE) configurations as a function of blowdown pressure ratio. The performance curves clearly indicate that a straight-tube PDRE is superior in specific impulse to a SSRE with a sonic nozzle over the entire range of pressure ratios. Note, however, that a straight-tube PDRE in general does not compare favorably to a SSRE fitted with an optimized de Laval supersonic nozzle, particularly at the high pressure ratios typical for boost or in-space rocket applications. However, the calculations also show that if a dynamically optimized, supersonic de Laval nozzle could be could be fitted to a PDRE, then the specific impulse of the device would exceed that of a comparable SSRE. While such a nozzle is a considerable idealization, it is clear that nozzle design and optimization will play a critical role in whether the performance potential of PDREs can be effectively realized in practice. In order to study PDRE nozzle issues with greater accuracy, a quasi-one-dimensional, finite-rate chemistry CFD code has been developed by the author. Comparisons of the code with both the previous MOC model and experimental data from Stanford University are reported. The effect of constant-gamma and finite-rate chemistry assumptions on the flowfield and performance is examined. Parametric studies of the effect of nozzle throat size and expansion ratio, at various blowdown pressure ratios, are reported.

Morris, Christopher I.↗
DOE OSTI2026

Strong gradient neoclassical transport in the plateau regime

Strong gradient regions in tokamaks such as the pedestal or internal transport barriers are regions of reduced turbulence where neoclassical transport can play a dominant role. In pedestals, gradient lengths comparable to the ion poloidal gyroradius have been measured. Standard neoclassical theory can miss important strong gradient effects in these regions because it assumes that the gradient length scales of density, temperature and potential are larger than the ion poloidal gyroradius. We extend plateau regime neoclassical theory into regions of gradients of the order of the ion poloidal gyroradius to capture strong gradient effects on transport processes in the pedestal and internal transport barriers. The fundamental idea behind our new framework is to keep a scale separation between the orbit widths and the gradient length scales by performing a large aspect ratio expansion. In the plateau regime, strong gradients cause poloidal variation that is in–out as well as up–down asymmetric. We study two different test cases assuming either radial force balance or the absence of turbulence and show that strong gradient effects can enhance or reduce standard neoclassical theory predictions in the plateau regime in strong gradient regions.

fusion plasma↗
DOE OSTI2020

Hydrodynamic Characterization of a Shallow Rectangular Spouted Bed Using Bed Pressure Signals and Image Processing Methods

To support the design of spouted bed reactors, experiments are performed in a rectangular spouted bed to measure the peak pressure drop (Δ P max ) and the internal and external spouting velocities ( U is and U es ). Four particle types spanning under Geldart D and B are used. U is and U es are measured from bed pressure drop and expansion ratio profiles, respectively, with the U es / U is ratio of each particle type obtained. Additionally, correlations predicting the Δ P max of Geldart D and B particles are developed by reviewing published correlations and investigating the nozzle effect. Good prediction is achieved with a slope of 0.997 and an adjusted R 2 of 0.998 in a 95% confidence interval. U es is predicted through correlations developed by adding effects of d p / D i and H 0 in the Mathur–Gishler equation. The calculated U es agrees well with the experimental data, achieving a slope of 0.998 and an adjusted R 2 of 0.998 in a 95% confidence interval.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗
NASA NTRS1962

Measured Steady-State Performance of Water Vapor Jets for Use in Space Vehicle Attitude Control Systems

Measurements have been made in a vacuum environment to determine the steady-state performance of several nozzles having thrusts up to 1000 dynes for use in space vehicle attitude control systems. Water vapor was used as a propellant. The results indicate that the trend of the variation of specific impulse and thrust coefficient with expansion ratio is predicted by calculations based on one-dimensional isentropic flow. The level of these quantities, however, is dependent upon the nozzle diameter. The specific impulse, for example, varies from about 30 percent to 80 percent of the theoretical value as the nozzle thrust is increased from about 10 to 1000 dynes.

JET↗
NASA NTRS1972

Space shuttle orbit maneuvering engine reusable thrust chamber

A data dump is presented containing space shuttle orbiter maneuvering engine performance, weight, envelope, and interface pressure requirements for candidate propellant combinations (NTO/MMH, NTO50-50, LOX/MMH, LOX/50-50, LOX/N2H4, LOX/C3H8, and LOX/RP-1) and cooling concepts (regenerative and dump/film). These data are presented parametrically for the thrust, chamber pressure, nozzle expansion ratio, and engine mixture ratio ranges of interest. Also included is information describing sensitivity to system changes; reliability, maintainability and safety; development programs and associated critical technology areas; engine cost comparisons during development and operation; and ecological effects.

Source record↗
NASA NTRS1973

Molecular flux measurements in the back flow region of a nozzle plume

A series of tests were conducted to measure the mass flux in the far field of a nozzle plume in a high vacuum with emphasis on the back flow region. The measurements presented provided fairly accurate data for off-axis angles as large as 140 deg (i.e., in the back flow region). This region, since it is well behind the exit plane, is of paticular interest to those concerned with instrument contamination. Usually sensitive spacecraft surfaces are located in the region affected by the back flow. Parameters such as expansion ratio, throat diameter, nozzle lip shape, and plenum (chamber) pressure were varied, carbon dioxide and nitrogen gases were flowed and mass flux measurements were taken using quartz crystal microbalances in as many as nine different locations relative to the tests nozzle. Several conclusions with respect to the effect of nozzle and gas parameters on the amount of back flow mass flux are offered, and it was demonstrated that gaseous mass fluxes, which are not predictable by present theories, are encountered in the region behind the nozzle exit plane. This knowledge is significant if materials incompatible with the gaseous exhaust products are used in this region.

Chirivella, J. E.↗
NASA NTRS1973

Mass flux measurements and correlations in the back flow region of a nozzle plume.

A series of tests were conducted to measure the mass flux in the far field of a nozzle plume in a high vacuum with emphasis on the back flow region. The measurements presented in this report provide fairly accurate data for off-axis angles as large as 140 deg (i.e., the back flow region). The tests utilized five different nozzles. Parameters such as expansion ratio, throat diameter, nozzle lip shape, and plenum (chamber) pressure were varied. Carbon dioxide and nitrogen gases were passed through, and mass flux measurements were taken using quartz crystal microbalances in as many as nine different locations relative to the test nozzle. The tests resulted in a large matrix of data that were correlated and compared to existing prediction theories. Several conclusions with respect to the effect of nozzle and gas parameters on the amount of back flow mass flux are offered, and it was demonstrated that gaseous mass fluxes, which are not predictable by present theories, are encountered in the region behind the nozzle exit plane.

Chirivella, J. E.↗
NASA NTRS1973

Advanced space engine preliminary design

A preliminary design was completed for an O2/H2, 89 kN (20,000 lb) thrust staged combustion rocket engine that has a single-bell nozzle with an overall expansion ratio of 400:1. The engine has a best estimate vacuum specific impulse of 4623.8 N-s/kg (471.5 sec) at full thrust and mixture ratio = 6.0. The engine employs gear-driven, low pressure pumps to provide low NPSH capability while individual turbine-driven, high-speed main pumps provide the system pressures required for high-chamber pressure operation. The engine design dry weight for the fixed-nozzle configuration is 206.9 kg (456.3 lb). Engine overall length is 234 cm (92.1 in.). The extendible nozzle version has a stowed length of 141.5 cm (55.7 in.). Critical technology items in the development of the engine were defined. Development program plans and their costs for development, production, operation, and flight support of the ASE were established for minimum cost and minimum time programs.

Cuffe, J. P. B.↗
NASA NTRS1974

Flight-test investigation of the aerodynamic characteristics and flow interference effects about the aft fuselage of the F-111A airplane

Static pressure measurements were made on the aft fuselage of an F-111A airplane to determine local flow characteristics and engine/airframe interaction effects. Data were obtained over the Mach number range from 0.5 to 2.0. Aspiration effects associated with low ejector nozzle expansion ratios reduced the local pressure coefficients particularly on the interfairing but also extending to the trailing edge of the nacelle. The presence of afterbodies also affected the behavior of the air flowing into and about the ejector nozzle. Pressures about the aft fuselage were improved by an increase in primary nozzle area at a supersonic speed. A comparison of wind-tunnel and flight-test results showed generally good agreement, although there was a large disparity in pressure level about the ejector nozzle. However, the shape of the data curves and the local flow behavior were basically similar.

Taillon, N. V.↗
NASA NTRS1976

Discrete estimation of continuous angle-modulated signals over multipath channels for aeronautical communication

The multipath channel for communication between an aerospace vehicle and a ground terminal is modeled by a multiplicative first-order Markov process. The multiplicative process is treated as a component of the message model and the discrete-time demodulation algorithms using the extended Kalman nonlinear estimation technique are developed for continuous-time angle-modulated signals. The equivalent baseband form of the demodulator structure is derived. Two examples of the message process are discussed for an FM system. The simulation results are presented for various values of the bandwidth expansion ratio and the additive SNR. The performance of the baseband algorithms is discussed.

Takhar, G. S.↗
NASA NTRS1977

Scramjet exhaust simulation technique for hypersonic aircraft nozzle design and aerodynamic tests

Current design philosophy for scramjet-powered hypersonic aircraft results in configurations with the entire lower fuselage surface utilized as part of the propulsion system. The lower aft-end of the vehicle acts as a high expansion ratio nozzle. Not only must the external nozzle be designed to extract the maximum possible thrust force from the high energy flow at the combustor exit, but the forces produced by the nozzle must be aligned such that they do not unduly affect aerodynamic balance. The strong coupling between the propulsion system and aerodynamics of the aircraft makes imperative at least a partial simulation of the inlet, exhaust, and external flows of the hydrogen-burning scramjet in conventional facilities for both nozzle formulation and aerodynamic-force data acquisition. Aerodynamic testing methods offer no contemporary approach for such vehicle design requirements. NASA-Langley has pursued an extensive scramjet/airframe integration R&D program for several years and has recently developed a promising technique for simulation of the scramjet exhaust flow for hypersonic aircraft. Current results of the research program to develop a scramjet flow simulation technique through the use of substitute gas blends are described in this paper.

Hunt, J. L.↗
NASA NTRS1977

Update on development of the Space Shuttle main engine /as of 5 May 1977/

The development and testing of the Space Shuttle main engine are described. The preburners, main injector, main combustion chamber, and 35:1 expansion ratio test nozzle were successfully run at full power level (109 percent of rated power level). Integral combustion stability aids damped induced instability oscillations within 6 milliseconds. Three significant turbomachinery problems were identified and solved during the past year. These problems involved high-pressure fuel turbopump subsynchronous whirl, severe overheating of the turbine bearing and components and, in the high-pressure oxidizer turbopump, burning that first occurred in the drain cavity downstream of the primary oxidizer seal.

Johnson, J. R.↗
NASA NTRS1980

Technology status of a liquid fluorine-hydrazine rocket engine for a planetary spacecraft propulsion system

This paper discusses the current status of a fluorine-hydrazine rocket engine development program. Incorporation of a thin rhenium inner liner successfully eliminated corrosion of the carbon/carbon composite thrust chamber wall experienced during a previous test program. The results of hot-fire tests utilizing reworked and new injectors which provide increased fuel film cooling showed that thrust chamber head-end temperatures could be maintained at an acceptable level. As expected, the accompanying specific impulse performance loss requires optimizing the amount of film cooling to minimize the loss. The efforts to refine the rhenium liner vapor deposition process culminated in a carbon/carbon composite thrust chamber total test duration of 1008 seconds. Tasks presently in process include: (1) fabrication of two carbon/carbon composite thrust chambers incorporating 60:1 expansion ratio nozzles; (2) injector tests to optimize performance and cooling; (3) additional refinements to the rhenium lining process; and (4) fabrication and test of a freestanding rhenium thrust chamber.

Appel, M. A.↗
NASA NTRS1981

Jay Carter Enterprises, Incorporated steam engine

The Small Community Solar Thermal Power Experiment (SCSE) selected an organic rankine cycle (ORC) engine driving a high speed permanent magnet alternator (PMA) as the baseline power conversion subsystem (PCS) design. The back-up conceptual PCS design is a steam engine driving an induction alternator delivering power directly to the grid. The development of the automotive reciprocating simple rankine cycle steam engine and how an engine of similar design might be incorporated into the SCSE is discussed. A description of the third generation automotive engine is included along with some preliminary test data. Tests were conducted with the third generation engine driving an induction alternator delivering power directly to the grid. The purpose of these tests is to further verify the effects of expander inlet temperature, input thermal power level, expansion ratio, and other parameters affecting engine performance to aid in the development of an SCSE PCS.

Source record↗
NASA NTRS1981

Shuttle solid rocket motor nozzle alternate ablative evaluation

A series of subscale tests are shown to suggest that a lower-cost ablative material than the rayon-based carbon ablative currently used in the Space Shuttle Solid Rocket Motor (SRM) may be used as a substitute. Six such ablatives with outstanding performance characteristics, using spun PAN and continuous pitch and PAN fibers instead of the present, continuous rayon, were identified in the course of tests with HTPB/AL/AP solid propellant grains with a burn time of 12 sec. The test nozzle features an initial throat diameter of 2.2 in. and a 6.1 expansion ratio. In addition to nozzle structural feature drawings, extensive test data tables and propellant formulation and properties tables are provided.

Powers, L. B.↗