SEARCH · Search NASA

Results for “conventional gas”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 109 records · Page 6

All sources NASA NTRS DOE OSTI

NASA NTRS2001

Simplified Analysis of Pulse Detonation Rocket Engine B1owdown Gasdynamics and Performance

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. The goal of this paper is to improve understanding of PDRE blowdown gasdynamics and performance issues through use of a simplified model that captures the essential features of the unsteady blowdown process, and yet remains computationally inexpensive. The PDRE system studied here is highly idealized, consisting of a constant-area detonation tube with one end closed and the other end open to the environment. The tube is prefilled with a gaseous propellant mixture with no initial velocity or outflow to the environment. The detonation is initiated instantaneously at the closed end of the device. Chapman-Jouguet (C-J) post-detonation gas conditions are calculated using the CET89 version of the NASA thermochemical code. The I-D, unsteady method of characteristics is used to calculate the flowfield following the detonation front. See the compressible flow texts by Thompson and Zucrow and Hoffman for details of this method. Parametric studies of the effect of mixture stoichiometry, fill temperature, and blowdown pressure ratio on performance are reported. A comparison of the performance of an idealized straight-tube PDRE with a conventional steady-state rocket engine is provided. The effect of constant-gamma and equilibrium chemistry assumptions is also examined. Additionally, in order to form an assessment of the accuracy of the model, the flowfield time history is compared to experimental data from Stanford University.

Morris, Christopher I.↗

NASA NTRS2005

Power Requirements Determined for High-Power-Density Electric Motors for Electric Aircraft Propulsion

Future advanced aircraft fueled by hydrogen are being developed to use electric drive systems instead of gas turbine engines for propulsion. Current conventional electric motor power densities cannot match those of today s gas turbine aircraft engines. However, if significant technological advances could be made in high-power-density motor development, the benefits of an electric propulsion system, such as the reduction of harmful emissions, could be realized.

Johnson, Dexter↗

NASA NTRS1997

Comparison of High Aspect Ratio Cooling Channel Designs for a Rocket Combustion Chamber

An analytical investigation on the effect of high aspect ratio (height/width) cooling channels, considering different coolant channel designs, on hot-gas-side wall temperature and coolant pressure drop for a liquid hydrogen cooled rocket combustion chamber, was performed. Coolant channel design elements considered were: length of combustion chamber in which high aspect ratio cooling was applied, number of coolant channels, and coolant channel shape. Seven coolant channel designs were investigated using a coupling of the Rocket Thermal Evaluation code and the Two-Dimensional Kinetics code. Initially, each coolant channel design was developed, without consideration for fabrication, to reduce the hot-gas-side wall temperature from a given conventional cooling channel baseline. These designs produced hot-gas-side wall temperature reductions up to 22 percent, with coolant pressure drop increases as low as 7.5 percent from the baseline. Fabrication constraints for milled channels were applied to the seven designs. These produced hot-gas-side wall temperature reductions of up to 20 percent, with coolant pressure drop increases as low as 2 percent. Using high aspect ratio cooling channels for the entire length of the combustion chamber had no additional benefit on hot-gas-side wall temperature over using high aspect ratio cooling channels only in the throat region, but increased coolant pressure drop 33 percent. Independent of coolant channel shape, high aspect ratio cooling was able to reduce the hot-gas-side wall temperature by at least 8 percent, with as low as a 2 percent increase in coolant pressure drop. The design with the highest overall benefit to hot-gas-side wall temperature and minimal coolant pressure drop cooling can now be done in relatively short periods of time with multiple iterations.

Wadel, Mary F.↗

NASA NTRS1998

Comparison of High Aspect Ratio Cooling Channel Designs for a Rocket Combustion Chamber with Development of an Optimized Design

An analytical investigation on the effect of high aspect ratio (height/width) cooling channels, considering different coolant channel designs, on hot-gas-side wall temperature and coolant pressure drop for a liquid hydrogen cooled rocket combustion chamber, was performed. Coolant channel design elements considered were: length of combustion chamber in which high aspect ratio cooling was applied, number of coolant channels, and coolant channel shape. Seven coolant channel designs were investigated using a coupling of the Rocket Thermal Evaluation code and the Two-Dimensional Kinetics code. Initially, each coolant channel design was developed, without consideration for fabrication, to reduce the hot-gas-side wall temperature from a given conventional cooling channel baseline. These designs produced hot-gas-side wall temperature reductions up to 22 percent, with coolant pressure drop increases as low as 7.5 percent from the baseline. Fabrication constraints for milled channels were applied to the seven designs. These produced hot-gas-side wall temperature reductions of up to 20 percent, with coolant pressure drop increases as low as 2 percent. Using high aspect ratio cooling channels for the entire length of the combustion chamber had no additional benefit on hot-gas-side wall temperature over using high aspect ratio cooling channels only in the throat region, but increased coolant pressure drop 33 percent. Independent of coolant channel shape, high aspect ratio cooling was able to reduce the hot-gas-side wall temperature by at least 8 percent, with as low as a 2 percent increase in coolant pressure drop. ne design with the highest overall benefit to hot-gas-side wall temperature and minimal coolant pressure drop increase was the design which used bifurcated cooling channels and high aspect ratio cooling in the throat region. An optimized bifurcated high aspect ratio cooling channel design was developed which reduced the hot-gas-side wall temperature by 18 percent and reduced the coolant pressure drop by 4 percent. Reductions of coolant mass flow rate of up to 50 percent were possible before the hot-gas-side wall temperature reached that of the baseline. These mass flow rate reductions produced coolant pressure drops of up to 57 percent.

Wadel, Mary F.↗

NASA NTRS1973

Electric vehicle battery research and development

High energy battery technology for electric vehicles is reviewed. The state-of-the-art in conventional batteries, metal-gas batteries, alkali-metal high temperature batteries, and organic electrolyte batteries is reported.

Schwartz, H. J.↗

NASA NTRS1975

Application of cabin atmosphere monitors to rapid screening of breath samples for the early detection of disease states

Analysis of human breath is a nonintrusive method to monitor both endogenous and exogenous chemicals found in the body. Several technologies were investigated and developed which are applicable to monitoring some organic molecules important in both physiological and pathological states. Two methods were developed for enriching the organic molecules exhaled in the breath of humans. One device is based on a respiratory face mask fitted with a polyethylene foam wafer; while the other device is a cryogenic trap utilizing an organic solvent. Using laboratory workers as controls, two organic molecules which occurred in the enriched breath of all subjects were tentatively identified as lactic acid and contisol. Both of these substances occurred in breath in sufficient amounts that the conventional method of gas-liquid chromatography was adequate for detection and quantification. To detect and quantitate trace amounts of chemicals in breath, another type of technology was developed in which analysis was conducted using high pressure liquid chromatography and mass spectrometry.

Valentine, J. L.↗

NASA NTRS1980

An automotive transmission for automotive gas turbine power plants

A joint government-industry program was initiated to investigate the two-shaft gas turbine concept as an alternative to present-day automotive powerplants. Both were examined, compared and evaluated on the basis of the federal automotive driving cycle in terms of specific fuel/power/speed characteristics of the engine and the efficiency and performance of the transmission. The results showed that an optimum match of vehicle, gas turbine engine, and conventional automatic transmission is capable of a significant improvement in fuel economy. This system offers many advantages that should lead to its wide acceptance in future vehicles.

Polak, J. C.↗

NASA NTRS1980

Study of component technologies for fuel cell on-site integrated energy systems

Heating, ventilation and air conditioning equipment are integrated with three types of fuel cells. System design and computer simulations are developed to utilize the thermal energy discharge of the fuel in the most cost effective manner. The fuel provides all of the electric needs and a loss of load probability analysis is used to ensure adequate power plant reliability. Equipment cost is estimated for each of the systems analyzed. A levelized annual cost reflecting owning and operating costs including the cost of money was used to select the most promising integrated system configurations. Cash flows are presented for the most promising 16 systems. Several systems for the 96 unit apartment complex (a retail store was also studied) were cost competitive with both gas and electric based conventional systems. Thermal storage is shown to be beneficial and the optimum absorption chiller sizing (waste heat recovery) in connection with electric chillers are developed. Battery storage was analyzed since the system is not electric grid connected. Advanced absorption chillers were analyzed as well. Recommendations covering financing, technical development, and policy issues are given to accelerate the commercialization of the fuel cell for on-site power generation in buildings.

Lee, W. D.↗

NASA NTRS1981

Low thrust chemical propulsion for orbit transfer of large space structures

For transporting Large Space Structures (LSS) from Low Earth Orbit (LEO) to Geosynchronous Equatorial Orbit (GEO), a variety of low thrust engine cycles have been evaluated for oxygen/hydrogen and oxygen/hydrocarbon propellants. The engine cycles included conventional propellant turbine drives (gas generator, expander, and staged-combustion cycles), turboalternator/electric motor pump drive, and fuel-cell/electric motor pump drive, as well as pressure-fed engines. The thrust chamber cooling limits and the engine cycle limits were established for a range of thrust levels. The candidate engine cycles were analyzed, screened, rated, and two engine cycle/configurations were selected for preliminary engine design. Preliminary engine designs for these two engines were formulated and engine design layouts prepared and parametric engine data generated.

Shoji, J. M.↗

NASA NTRS1984

Progress in the utilization of an oxide-dispersion-strengthened alloy for small engine turbine blades

The conventional means of improving gas turbine engine performance typically involves increasing the turbine inlet temperature; however, at these higher operational temperatures the high pressure turbine blades require air-cooling to maintain durability. Air-cooling imposes design, material, and economic constraints not only on the turbine blades but also on engine performance. The use of uncooled turbine blades at increased operating temperatures can offer significantly improved performance in small gas turbine engines. A program to demonstrate uncooled MA6000 high pressure turbine blades in a GTEC TFE731 turbofan engine is being conducted. The project goals include demonstration of the advantages of using uncooled MA6000 turbine blades as compared with cast directionally solidified MAR-M 247 blades.

Beatty, T. G.↗

NASA NTRS1991

Lightweight Valve Closes Duct Quickly

Expanding balloon serves as lightweight emergency valve to close wide duct. Uninflated balloon stored in housing of duct. Pad resting on burst diaphragm protects balloon from hot gases in duct. Once control system triggers valve, balloon inflates rapidly to block duct. Weighs much less than does conventional butterfly, hot-gas, or poppet valve capable of closing duct of equal diameter.

Fournier, Walter L.↗

NASA NTRS1993

In-stream measurements of combustion during Mach 5 to 7 tests of the Hypersonic Research Engine (HRE)

Results of in-stream combustion measurements taken during Mach 5 to 7 true simulation testing of the Hypersonic Research Engine/Aerothermodynamic Integration Model (HRE/AIM) are presented. These results, the instrumentation techniques, and configuration changes to the engine installation that were required to test this model are described. In test runs at facility Mach numbers of 5 to 7, an exhaust instrumentation ring which formed an extension of the engine exhaust nozzle shroud provided diagnostic measurements at 10 circumferential locations in the HRE combustor exit plane. The measurements included static and pitot pressures using conventional conical probes, combustion gas temperatures from cooled-gas pyrometer probes, and species concentration from analysis of combustion gas samples. Results showed considerable circumferential variation, indicating that efficiency losses were due to nonuniform fuel distribution or incomplete mixing. Results using the Mach 7 facility nozzle but with Mach 6 temperature simulation, 1590 to 1670 K, showed indications of incomplete combustion. Nitric oxide measurements at the combustor exit peaked at 2000 ppmv for stoichiometric combustion at Mach 6.

Lezberg, Erwin A.↗

NASA NTRS1996

Stable isotope analysis of molecular oxygen from silicates and oxides using CO2 laser extraction

A laser-excited system for determination of the oxygen isotope composition of small quantities of silicate and oxide minerals was constructed and tested at JSC. This device is the first reported to use a commercially available helium cryostat to transfer and purify oxygen gas quantitatively within the system. The system uses oxygen gas instead of the conventional CO2 for mass spectrometer analyses. This modification of technique permits determination of all three stable oxygen isotopes, an essential requirement for oxygen isotope analysis of meteoritic material. Tests of the system included analysis of standard silicate materials NBS 28 and UWMG2 garnet, six SNC meteorites, and inclusions and chondrules from the Allende meteorite. Calibration with terrestrial standards was excellent. Meteorite values are close to published values and show no evidence of terrestrial oxygen contamination. The one limitation observed is that, in some runs on fine-grained SNC matrix material, sample results were affected by other samples in the sample holder within the reaction chamber. This reemphasizes the need for special precautions in dealing with fine-grained, reactive samples. Performance of the JSC instrument compares favorably with that of any other instrument currently producing published oxygen isotope data.

Perry, Eugene↗

NASA NTRS1997

Mixing of Pure Air Jets with a Reacting Fuel-Rich Crossflow

Jets in a crossflow play an integral role in practical combustion systems such as can and annular gas turbine combustors in conventional systems, and the Rich-burn/Quick-mix/Lean-burn (RQL) combustor utilized in stationary applications and proposed for advanced subsonic and supersonic transports. The success of the RQL combustor rests with the performance of the quick-mixing section that bridges the rich and lean zones. The mixing of jet air with a rich crossflow to bring the reaction to completion in the lean zone must be performed rapidly and thoroughly in order to decrease the extent of near-stoichiometric fluid pocket formation. Fluid pockets at near-stoichiometric equivalence ratios are undesirable because the high temperatures attained accelerate pollutant formation kinetics associated with nitric oxide (NO). The present study develops a model experiment designed to reveal the processes that occur when jet air is introduced into hot effluent emanating from a fuel-rich reaction zone.

Leong, M. Y.↗

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.↗

NASA NTRS2005

Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

Pulse detonation rocket engines (PDREs) have generated 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 considerable 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 progress that has been made in comparing the available experimental measurements with analytical and numerical models. In recent work by the author, a quasi-one-dimensional, finite rate chemistry CFD model was utilized to study the gasdynamics and performance characteristics of PDREs over a range of blowdown pressure ratios from 1-1000. Models of this type are computationally inexpensive, and enable first-order parametric studies of the effect of several nozzle and extension geometries on PDRE performance over a wide range of conditions. However, the quasi-one-dimensional approach is limited in that it cannot properly capture the multidimensional blast wave and flow expansion downstream of the PDRE, nor can it resolve nozzle flow separation if present. Moreover, the previous work was limited to single-pulse calculations. In this paper, an axisymmetric finite rate chemistry model is described and utilized to study these issues in greater detail. Example Mach number contour plots showing the multidimensional blast wave and nozzle exhaust plume are shown. The performance results are compared with the quasi-one-dimensional results from the previous paper. Both Euler and Navier-Stokes solutions are calculated in order to determine the effect of viscous effects in the nozzle flowfield. Additionally, comparisons of the model results to performance data from CalTech, as well as experimental flowfield measurements from Stanford University, are also reported.

Morris, Christopher I.↗

NASA NTRS1996

Optimization Study of the Ames 0.5 Two-Stage Light Gas Gun

There is a need for more faithful simulation of space debris impacts on various space vehicles. Space debris impact velocities can range up to 14 km/sec and conventional two-stage light gas guns with moderately heavy saboted projectiles are limited to launch velocities of 7-8 km/sec. Any increases obtained in the launch velocities will result in more faithful simulations of debris impacts. It would also be valuable to reduce the maximum gun and projectile base pressures and the gun barrel erosion rate. In this paper, the results of a computational fluid dynamics (CFD) study designed to optimize the performance of the NASA Ames 0.5' gun by systematically varying seven gun operating parameters are reported. Particularly beneficial effects were predicted to occur if (1) the piston mass was decreased together with the powder mass and the hydrogen fill pressure and (2) the pump tube length was decreased. The optimum set of changes in gun operating conditions were predicted to produce an increase in muzzle velocity of 0.7-1.0 km/sec, simultaneously with a substantial decrease in gun erosion. Preliminary experimental data have validated the code predictions. Velocities of up to 8.2 km/sec with a 0.475 cm diameter saboted aluminum sphere have been obtained, along with large reductions in gun erosion rates.

Bogdanoff, D. W.↗

NASA NTRS1991

Characteristics of a high pressure gas proportional counter filled with xenon

The characteristics of a conventional cylindrical geometry proportional counter filled with high pressure xenon gas up to 10 atm. were fundamentally investigated for use as a detector in hard X-ray astronomy. With a 2 percent methane gas mixture the energy resolutions at 10 atm. were 9.8 percent and 7.3 percent for 22 keV and 60 keV X-rays, respectively. From calculations of the Townsend ionization coefficient, it is shown that proportional counters at high pressure operate at weaker reduced electric field than low pressure counters. The characteristics of a parallel grid proportional counter at low pressure showed similar pressure dependence. It is suggested that this is the fundamental reason for the degradation of resolution observed with increasing pressure.

Sakurai, H.↗