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Results for “combustion efficiency”

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 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Non-grey radiation in a liquid rocket combustion chamber

The effects of radiation on droplets gasification in liquid rocket combustion chambers has been studied. The modeling includes a Legendre pseudo-spectral collocation approximation to solve the Radiative Transfer Equation (RTE). The band model has been utilized to account for non-grey emitting and absorbing gases present in the comustion chamber. The GEMCHIP II code has been utilized to study the fuel and oxidizer droplets combustion and interaction. The submodels within this code are capable of accounting for group combustion and conjugate effects between many droplets. The radiative model has been coupled with the GEMCHIP II code accounting for radial effects only, to provide the tool for studying the combustion-radiation coupling effects in a bipropellant system. The results indicate that the gasification/combustion process will be enhanced upstream of the chamber causing thicker flame sheet and associated higher combustion efficiency.

Kehtarnavaz, H.↗
NASA NTRS1990

Transpiration cooling in the locality of a transverse fuel jet for supersonic combustors

The objective of the current work was to determine the feasibility of transpiration cooling for the relief of the local heating rates in the region of a sonic, perpendicular, fuel jet of gaseous hydrogen. Experiments were conducted to determine the interaction between the cooling required and flameholding limits of a transverse jet in a high-enthalpy, Mach 3 flow in both open-jet and direct-connect test mode. Pulsed shadowgraphs were used to illustrate the flow field. Infrared thermal images indicated the surface temperatures, and the OH(-) emission of the flame was used to visualize the limits of combustion. Wall, static presures indicated the location of the combustion within the duct and were used to calculate the combustion efficiency. The results from both series of tests at facility total temperatures of 1700 K and 2000 K are presented.

Northam, G. Burton↗
NASA NTRS1991

A parametric study of scramjet combustors for Mach 8 to Mach 20 flight

Performance trends for a combustor with ramp and slot fuel injectors are investigated through a three-dimensional parametric study at flight Mach numbers of 8, 10, 12, 16, and 20. Focus is placed on the mixing (combustion) efficiency, combustor loads, and thrust over the Mach number range. It is observed that the mixing efficiency decreases by 0.2 when a flight Mach number is increased from 8 to 12 but changes little for higher Mach numbers; the lower mixing is attributed more to a large decrease in the fuel-to-air velocity ratio and residence time over the Mach number range than to compressibility. It is pointed out that the ramp generates a shock impinging on the cowl, while a combined ramp and slot injection produces better mixing, both because of the action of this shock on the slot fuel and because of better fuel distribution in the duct.

Kamath, Pradeep S.↗
NASA NTRS1991

Mixing enhancement of reacting parallel fuel jets in a supersonic combustor

Pursuant to a NASA-Langley development program for a scramjet HST propulsion system entailing the optimization of the scramjet combustor's fuel-air mixing and reaction characteristics, a numerical study has been conducted of the candidate parallel fuel injectors. Attention is given to a method for flow mixing-process and combustion-efficiency enhancement in which a supersonic circular hydrogen jet coflows with a supersonic air stream. When enhanced by a planar oblique shock, the injector configuration exhibited a substantial degree of induced vorticity in the fuel stream which increased mixing and chemical reaction rates, relative to the unshocked configuration. The resulting heat release was effective in breaking down the stable hydrogen vortex pair that had inhibited more extensive fuel-air mixing.

Drummond, J. P.↗
NASA NTRS1991

Mixing, ignition and combustion studies using discrete orifice injection at hypervelocity flight conditions

Testing at hypervelocity test conditions continues in the Hypulse expansion tube. A newly-calibrated Mach 13 operating point has been achieved, and hydrogen injection studies were performed at this test condition. The experiments were conducted in a 2-in x 1-in combustor model employing three injector configurations: a single flush-wall circular orifice; two such orifices on opposite walls (the 'dual' injector configuration); and two orifices on the upper and the lower walls (the 'quad' injector configuration). The quad injectors were distinctly superior to the other two configurations in promoting mixing, ignition, and combustion. The single injector, however, proved to be superior to the dual injectors, supporting the concept that the aspect ratio of the 'injection cell' plays a large role in promoting the processes required to achieve high combustion efficiency.

Bakos, R.↗
NASA NTRS1992

A two-phase restricted equilibrium model for combustion of metalized solid propellants

An Eulerian-Lagrangian two-phase approach was adopted to model the multi-phase reacting internal flow in a solid rocket with a metalized propellant. An Eulerian description was used to analyze the motion of the continuous phase which includes the gas as well as the small (micron-sized) particulates, while a Lagrangian description is used for the analysis of the discrete phase which consists of the larger particulates in the motor chamber. The particulates consist of Al and Al2O3 such that the particulate composition is 100 percent Al at injection from the propellant surface with Al2O3 fraction increasing due to combustion along the particle trajectory. An empirical model is used to compute the combustion rate for agglomerates while the continuous phase chemistry is treated using chemical equilibrium. The computer code was used to simulate the reacting flow in a solid rocket motor with an AP/HTPB/Al propellant. The computed results show the existence of an extended combustion zone in the chamber rather than a thin reaction region. The presence of the extended combustion zone results in the chamber flow field and chemical being far from isothermal (as would be predicted by a surface combustion assumption). The temperature in the chamber increases from about 2600 K at the propellant surface to about 3350 K in the core. Similarly the chemical composition and the density of the propellant gas also show spatially non-uniform distribution in the chamber. The analysis developed under the present effort provides a more sophisticated tool for solid rocket internal flow predictions than is presently available, and can be useful in studying apparent anomalies and improving the simple correlations currently in use. The code can be used in the analysis of combustion efficiency, thermal load in the internal insulation, plume radiation, etc.

Sabnis, J. S.↗
NASA NTRS1991

Characteristics of smoke emissions from biomass fires of the Amazon region - BASE-A experiment

The Biomass Burning Airborne and Spaceborne Experiment-Amazonia was designed for study of both aerosol and gaseous emissions from fires using an airborne sampling platform. The emission factors for combustion products from four fires suggest that the proportion of carbon released in the form of CO2 is higher than for fires of logging which has been burned in the western U.S. Combustion efficiency was of the order of 97 percent for the Amazonian test fire and 86-94 percent for deforestation fires. The inorganic content of particles from tropical fires are noted to be different from those of fires in the U.S.

Ward, Darold E.↗
NASA NTRS1991

Mixing and mixing enhancement in supersonic reacting flowfields

Propulsion systems planned for use late in this century and beyond will require appropriate physical models for describing supersonic combustion and numerical techniques for solving the model governing equations. A computer program to study these flows is reported which considers the multicomponent diffusion and convection of important chemical species, the finite-rate reaction of these species, and the resulting interaction of the field mechanics and the chemistry. The application of the program to a spatially developing and reacting mixing layer, which serves an an excellent physical model for the mixing and reaction processes that take place in a scramjet combustor, is reported. Several techniques to enhance the fuel-air mixing and growth of that layer and improve its overall combustion efficiency are considered.

Drummond, J. P.↗
NASA NTRS1992

Laser absorption flow diagnostics for use in hypersonic ground-based and flight experiments

Consideration is given to laboratory ground-based scale systems for measuring OH and NO and diode laser systems for measuring O2 and H2O. Several diagnostic instruments which are currently being used in a scramjet combustor test rig of the Ames 16-inch shock tunnel are described. Data on mole fraction, temperature, density, pressure, and velocity is obtained using specific data reduction techniques. These measurements provide critical data for evaluation of the propulsion system performance which include mass capture for quantification of fuel equivalence ratio (O2), flow contamination and emission levels (NO), combustion progress (OH), and injector mixing and combustion efficiency (H2O).

Cavolowsky, John A.↗
NASA NTRS1932

Considerations of Air Flow in Combustion Chambers of High-Speed Compression-Ignition Engines

The air flow in combustion chambers is divided into three fundamental classes - induced, forced, and residual. A generalized resume is given of the present status of air flow investigations and of the work done at this and other laboratories to determine the direction and velocity of air movement in auxiliary and integral combustion chambers. The effects of air flow on engine performance are mentioned to show that although air flow improves the combustion efficiency, considerable induction, friction, and thermal losses must be guarded against.

Spanogle, J A↗
NASA NTRS1940

Analysis of Cylinder-pressure-indicator Diagrams Showing Effects of Mixture Strength and Spark Timing

An investigation was made to determine the effect of mixture strength and of normal as well as optimum spark timing on the combustion, on the cylinder temperature, and on the performance characteristics of an engine. A single-cylinder test unit utilizing an air-cooled cylinder and a carburetor and operating with gasoline having an octane rating of 92 was used. The investigation covered a range of fuel-air ratios from 0.053 to 0.118. Indicator diagrams and engine-performance data were taken for each change in engine conditions. Examination of the indicator shows that for fuel-air ratios less than and greater than 0.082 the rate and the amount of effective fuel burned decreased. For a fuel-air ratio of 0.118 the combustion efficiency was only 58 percent. Advancing the spark timing increased the rate of pressure rise. This effect was more pronounced with leaner mixtures.

Gerrish, Harold C↗
NASA NTRS1949

Altitude-wind-tunnel investigation of AN-F-58 fuel in experimental version of J47 turbojet engine

An altitude-wind-tunnel investigation of the performance of AN-F-58 and AN-F-32 fuels in an experimental turbojet engine was conducted over a range of simulated altitudes and flight Mach numbers. Combustion efficiencies obtained with AN-F-58 and AN-F-32 fuels were approximately equal. The minimum-speed altitude operational limit was essentially the same with either AN-F-58 or AN-F-32 fuel. Starting characteristics of the two fuels were approximately the same at low wind milling speeds. Visual observation showed no apparent differences in the carbon-deposition rates of AN-F-58 and AN-F-32 fuels.

CHARGING AND CONTROL - GAS TURBINE ENGINES↗
NASA NTRS1950

Altitude-test-chamber Investigation of Performance of a 28-inch Ram-jet Engine I : Combustion and Operational Performance of Four Combustion-chamber Configurations

An altitude-test-chamber investigation of a 28-inch-diameter ram-jet engine at a simulated flight Mach number of approximately 2.0 for altitudes of 40,000 to 50,000 feet was conducted at the NACA Lewis laboratory. Three different flame holders, varying in the number and size of the annular gutters, in conjunction with several fuel-injection systems were investigated. The combustion efficiency for the flame-holder fuel-injection system that provided the best over-all operational fuel-air-ratio range (0.03 to 0.075) was over 0.9 at a fuel-air ratio of about 0.065 for the altitude range investigated.

MISSILES, SPECIFIC TYPES↗
NASA NTRS1950

Altitude Performance Characteristics of Tail-pipe Burner with Convergingconical Burner Section on J47 Turbojet Engine

An investigation of turbojet-engine thrust augmentation by means of tail-pipe burning was conducted in the NACA Lewis altitude wind tunnel. Performance data were obtained with a tail-pipe burner having a converging conical burner section installed on an axial-flow-compressor type turbojet engine over a range of simulated flight conditions and tail-pipe fuel-air ratios with a fixed-area exhaust nozzle. A maximum tail-pipe combustion efficiency of 0.86 was obtained at an altitude of 15,000 feet and a flight Mach number of 0.23. Tail-pipe burner operation was possible up to an altitude of 45,000 feet at a flight Mach number of 0.23.

Prince, William R↗
NASA NTRS1950

NACA research on combustors for aircraft gas turbines I : effects of operating variables on steady-state performance

Some of the systematic research conducted by the NACA on aircraft gas-turbine combustors is reviewed. Trends depicting the effect of inlet-air pressure, temperature, and velocity and fuel-air ratio on performance characteristics, such as combustion efficiency, maximum temperature rise attainable, pressure loss, and combustor-outlet temperature distribution are described for a variety of turbojet combustors of the liquid-fuel type. These trends are further discussed as effects significant to the turbojet engine, such as altitude operational limits, specific fuel consumption, thrust, acceleration, and turbine life.

Olson, Walter T↗
NASA NTRS1951

Effect of fuel volatility on performance of tail-pipe burner

Fuels having Reid vapor pressures of 6.3 and 1.0 pounds per square inch were investigated in a tail-pipe burner on an axial-flow-type turbojet engine at a simulated flight Mach number of 0.6 and altitudes from 20,000 to 45,000 feet. With the burner configuration used in this investigation, having a mixing length of only 8 inches between the fuel manifold and the flame holder, the low-vapor-pressure fuel gave lower combustion efficiency at a given tail-pipe fuel-air ratio. Because the exhaust-nozzle area was fixed, the lower efficiency resulted in lower thrust and higher specific fuel consumption. The maximum altitude at which the burner would operate was practically unaffected by the change in fuel volatility.

FUELS - TURBINE ENGINES, RAM JETS, AND PULSE JETS↗
NASA NTRS1956

Performance of a short combustor at high altitudes using hydrogen fuel

Performance characteristics of a 16-inch annular-type combustor installed in a full-scale engine using gaseous-hydrogen fuel were obtained at simulated altitudes from 66,000 to 86,000 feet at a flight Mach number of 0.8. Combustion efficiencies of 86 percent were obtained at 86,000 feet (combustor pressure, 420 lb/sq ft abs). Combustor blowout was not encountered during the investigation.

Sivo, Joseph N↗
NASA NTRS1955

Transient and steady-state performance of a single turbojet combustor with four different fuel nozzles

Acceleration and steady-state performance of a tubular combustor was evaluated at two simulated altitudes with four different fuel nozzles. Temperature response lag was observed with all the nozzles. Except for rich-limit blowout, the only combustion failures observed during acceleration were with a fuel nozzle that gave an interrupted flow delivery during the acceleration. This same nozzle, because of superior fuel atomization, gave the highest steady-state combustion efficiencies.

Mccafferty, Richard J↗