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At least 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1957

Performance of an Experimental Annular Turbojet Combustor with Methane and Propane

Combustion efficiencies obtained with gaseous methane were compared with reported data obtained with gaseous propane for the same experimental combustor configuration. The combustion efficiencies obtained with methane were 98, 91, and 77 percent at simulated flight altitudes of 56,000, 70,000 and 80,000 feet, corresponding to combustor inlet-air pressures from 15 to 5 inches of mercury absolute. Combustion efficiencies with propane were equivalent to those with methane up to a simulated altitude of 70,000 feet. At the most severe conditions investigated propane operated with a higher efficiency and over a wider range of fuel-air ratio than methane.

Norgren, Carl T↗
NASA NTRS1959

Application of Gas Analysis to Combustor Research

The performance of turbine-engine combustors usually is given in terms of operating limits and combustion efficiency. The latter property is determined most often by measuring the increase in enthalpy across the combustor through the use of thermocouples. This investigation was conducted to determine the ability of gas-analytical techniques to provide additional information about combustor performance. Gas samples were taken at the outlet and two upstream stations and their compositions determined. In addition to over-all combustion efficiency, estimates of local fuel-air ratios, local combustion efficiencies, and heat-release rates can be made. Conclusions can be drawn concerning the causes of combustion inefficiency and may permit corrective design changes to be made more intelligently. The purpose of this investigation was not to present data for a given combustor but rather to show the types and value of additional information that can be gained from gas-analytical data.

Hibbard, R. R.↗
NASA NTRS1978

Correlations of catalytic combustor performance parameters

Correlations for combustion efficiency percentage drop and the minimum required adiabatic reaction temperature necessary to meet emissions goals of 13.6 g CO/kg fuel and 1.64 g HC/kg fuel are presented. Combustion efficiency was found to be a function of the cell density, cell circumference, reactor length, reference velocity, and adiabatic reaction temperature. The percentage pressure drop at an adiabatic reaction temperature of 1450 K was found to be proportional to the reference velocity to the 1.5 power and to the reactor length. It is inversely proportional to the pressure, cell hydraulic diameter, and fractional open area. The minimum required adiabatic reaction temperature was found to increase with reference velocity and decrease with cell circumference, cell density and reactor length. A catalyst factor was introduced into the correlations to account for differences between catalysts. Combustion efficiency, the percentage pressure drop, and the minimum required adiabatic reaction temperature were found to be a function of the catalyst factor. The data was from a 12 cm-diameter test rig with noble metal reactors using propane fuel at an inlet temperature of 800 K.

Bulzan, D. L.↗
NASA NTRS1976

Swirl-can combustor performance to near-stoichiometric fuel-air ratio

Emissions and performance characteristics were determined for full-annulus swirl-can modular combustors operated to near stoichiometric fuel air ratios. The purposes of the tests were to obtain stoichiometric data at inlet air temperatures up to 894 K and to determine the effect of module number by investigating 120 and 72 module swirl-can combustors. The maximum average exit temperature obtained with the 120-module swirl-can combustor was 2465 K with a combustion efficiency of 95 percent at an inlet-air temperature of 894 K. The 72-module swirl-can combustor reached a maximum average exit temperature of 2306 K with a combustion efficiency of 92 percent at an inlet air temperature of 894 K. At a constant inlet air temperature, maximum oxides of nitrogen emission index values occurred at a fuel-air ratio of 0.037 for the 72-module design and 0.044 for the 120-module design. The combustor average exit temperature and combustion efficiency were calculated from emissions measurements. The measured emissions included carbon monoxide, unburned hydrocarbons, oxides of nitrogen, and smoke.

Diehl, L. A.↗
NASA NTRS1976

Swirl-can combustor performance to near-stoichiometric fuel-air-ratio

Emissions and performance characteristics were determined for two full-annulus swirl-can modular combustors operated to near-stoichiometric fuel air ratios. The purposes of the tests were to obtain stoichiometric data at inlet-air temperatures up to 894 K and to determine the effect of module number by investigating 120 and 72 module swirl-can combustors. The maximum average exit temperature obtained with the 120-module swirl-can combustor was 2465 K with a combustion efficiency of 95 percent at an inlet-air temperature of 894 K. The 72-module swirl-can combustor reached a maximum average exit temperature of 2306 K with a combustion efficiency of 92 percent at an inlet-air temperature of 894 K. At a constant inlet air temperature, maximum oxides of nitrogen emission index values occurred at a fuel-air ratio of 0.037 for the 72-module design and 0.044 for the 120-module design. The combustor average exit temperature and combustion efficiency were calculated from emissions measurements. The measured emissions included carbon monoxide, unburned hydrocarbons, oxides of nitrogen, and smoke.

Diehl, L. A.↗
NASA NTRS1975

Idle efficiency and pollution results for two-row swirl-can combustors having 72 modules

Two 72-swirl-can-module combustors were investigated in a full annular combustor test facility at engine idle conditions typical of a 30:1 pressure-ratio engine. The effects of radial and circumferential fuel scheduling on combustion efficiency and gaseous pollutants levels were determined. Test conditions were inlet-air temperature, 452 K; inlet total pressure, 34.45 newtons per square centimeter; and reference velocity, 19.5 meters per second. A maximum combustion efficiency of 98.1 percent was achieved by radial scheduling of fuel to the inner row of swirl-can modules. Emission index values were 6.9 for unburned hydrocarbons and 50.6 for carbon monoxide at a fuel-air ratio of 0.0119. Circumferential fuel scheduling of two 90 degree sectors of the swirl-can arrays produced a maximum combustion efficiency of 97.3 percent. The emission index values were 12.0 for unburned hydrocarbons and 69.2 for carbon monoxide at a fuel-air ratio of 0.0130.

Biaglow, J. A.↗
NASA NTRS1981

Reynolds number effects in combustion noise

Acoustic emission spectra have been obtained for non-premixed turbulent combustion from two small diameter laboratory gas burners, two commercial gas burners and a large gas burner in the firebox of a Babcock-Wilcox Boiler (50,000 lb steam/hr). The changes in burner size and firing rate represent changes in Reynolds number and changes in air/fuel ratio represent departure from stoichiometric proportions. The combustion efficiency was measured independently through gas analysis. The acoustic spectra obtained from the various burners exhibit a persistent shape over the Reynolds number range of 8200-82,000. The spectra were analyzed for identification of a predictable frequency domain that is most responsive to, and readily correlated with, combustion efficiency. A simple parameter (consisting of the ratio of the average acoustic power output in the most responsive frequency bandwidth to the acoustic power level of the loudest frequency) is proposed whose value increases significantly and unmistakably as combustion efficiency approaches 100%. The dependence of the most responsive frequency domain on the various Reynolds numbers associated with turbulent jets is discussed.

Seshan, P. K.↗
NASA NTRS1989

Hydrocarbon-fueled scramjet combustor component development tests

Technology is being developed for a hydrocarbon-fueled engine operating as a scramjet over the flight Mach number range from 5.6 to 7. A series of connected-pipe tests were performed to define scramjet combustor design criteria applicable to the United Technologies Research Center (UTRC) engine concept which comprises a pair of modular axisymmetric combustor underslung on a supersonic/hypersonic missile. The development of key pilot and fuel injector components of the combustor is pursued in a variable-geometry two-dimensional test section over a range of combustor entrance conditions simulating the intended flight regime. The applicability of the two-dimensional test results to the axisymmetric engine is ensured by maintaining a proper simulation of combustor entrance conditions and preserving the actual length scale in the two-dimensional test configuration. An air-breathing pilot was developed and tested to evaluate flame stabilization and flame propagation characteristics. A pilot configuration was developed that operated stably, with minimal flow spillage, at exhaust stagnation temperatures as high as 4500 R with ethylene fuel. It was demonstrated that the pilot promotes efficient combustion of either gaseous ethylene or preheated liquid Jet-A (JP-5) when they are injected into the supersonic mainstream flow as primary fuels. For the tests with Jet-A fuel, the fuel was heated in an array of internal cooling passages within the pilot walls to a thermodynamic condition such that it would flash-vaporize upon injection into the mainstream flow. The idea of using the air-breathing pilot and distributed primary and secondary fuel injectors to achieve efficient supersonic combustion over a wide range of equivalence ratios was also experimentally demonstrated during the program. During staged fuel injection tests with gaseous ethylene fuel, high secondary fuel combustion efficiencies were achieved and smooth transitions from fully supersonic to dual mode (supersonic/subsonic) operation were demonstrated at high overall equivalence ratios.

Kay, Ira W.↗
NASA NTRS1947

Performance Charts for a Turbojet System

Convenient charts are presented for computing the thrust, fuel consumption, and other performance values of a turbojet system. These charts take into account the effects of ram pressure, compressor pressure ratio, ratio of combustion-chamber-outlet temperature to atmospheric temperature, compressor efficiency, turbine efficiency, combustion efficiency, discharge-nozzle coefficient, losses in total pressure in the inlet to the jet-propulsion unit and in the combustion chamber, and variation in specific heats with temperature. The principal performance charts show clearly the effects of the primary variables and correction charts provide the effects of the secondary variables. The performance of illustrative cases of turbojet systems is given. It is shown that maximum thrust per unit mass rate of air flow occurs at a lower compressor pressure ratio than minimum specific fuel consumption. The thrust per unit mass rate of air flow increases as the combustion-chamber discharge temperature increases. For minimum specific fuel consumption, however, an optimum combustion-chamber discharge temperature exists, which in some cases may be less than the limiting temperature imposed by the strength temperature characteristics of present materials.

CHARTS-PERFORMANCE-ENGINES, TURBOJET↗
NASA NTRS1992

Smoke and fire characteristics for cerrado and deforestation burns in Brazil - BASE-B experiment

Five test fires were performed during August and September 1990 in the cerrado (savannalike region) in central Brazil (three fires) and tropical moist forest (two fires) in the eastern Amazon. This paper details the gases released, the ratios of the gases to each other and to particulate matter, fuel loads, and the fraction consumed (combustion factors), and the fire behavior associated with biomass consumption. Models are presented for evaluating emission factors for CH4, CO2, CO, H2, and particles less than 2.5 micron diam (PM2.5) as a function of combustion efficiency. The ratio of carbon released as CO2 (combustion efficiency) for the cerrado fires averaged 0.94 and for the deforestation fires it decreased from 0.88 for the flaming phase to less than 0.80 during the smoldering phase of combustion. For tropical ecosystems, emissions of most products of incomplete combustion are projected to be lower than previous estimates for savanna ecosystems and somewhat higher for fires used for deforestation purposes.

Ward, D. E.↗
NASA NTRS1983

Atomization and combustion performance of antimisting kerosene and jet fuel

Combustion performance of antimisting kerosene (AMK) containing FM-9 polymer was investigated at various levels of degradation (restoration of AMK for normal use in a gas turbine engine). To establish the relationship of degradation and atomization to performance in an aircraft gas turbine combustor, sprays formed by the nozzle of a JT8-D combustor with Jet A and AMK at 1 atmosphere (atm) (14.1 lb/square in absolute) pressure and 22 C at several degradation levels were analyzed. A new spray characterization technique based on digital image analysis of high resolution, wide field spray images formed under pulsed ruby laser sheet illumination was developed. Combustion tests were performed for these fuels in a JT8-D single can combustor facility to measure combustion efficiency and the lean extinction limit. Correlation of combustion performance under simulated engine operating conditions with nozzle spray Sauter mean diameter (SMD) measured at 1 atm and 22 C were observed. Fuel spray SMD and hence the combustion efficiency are strongly influenced by fuel degradation level. Use of even the most highly degraded AMK tested (filter ratio = 1.2) resulted in an increase in fuel consumption of 0.08% to 0.20% at engine cruise conditions.

Fleeter, R.↗
NASA NTRS1960

A Study of the Combustion of Aluminum Borohydride in a Small Supersonic Wind Tunnel

The combustion of aluminum borohydride in the Mach 2 airstream of a 3.84- by 10-inch wind tunnel was studied by analyzing gas samples taken from the airstream. Gas mixture composition was determined with the aid of a vacuum apparatus and gas chromatograph. The overall combustion efficiency of the fuel was assumed to be good because the combustion efficiency of its hydrogen component was found to be high. The increase of the lateral surface area of the combustion region was dependent upon the degree of mixing of fuel and its combustion products with air. In the flame zone where there was insufficient oxygen for complete combustion, the metal components of the fuel appeared to burn preferentially to the hydrogen component.

Allen, Harrison, Jr.↗
NASA NTRS1972

Effects of fuel nozzle design on performance of an experimental annular combustor using natural gas fuel

Tests of various fuel nozzles were conducted with natural gas fuel in a full-annulus combustor. The nozzles were designed to provide either axial, angled, or radial fuel injection. Each fuel nozzle was evaluated by measuring combustion efficiency at relatively severe combustor operating conditions. Combustor blowout and altitude ignition tests were also used to evaluate nozzle designs. Results indicate that angled injection gave higher combustion efficiency, less tendency toward combustion instability, and altitude relight characteristics equal to or superior to those of the other fuel nozzles that were tested.

Wear, J. D.↗
NASA NTRS1954

Investigation of Three Low-temperature-ratio Combustor Configurations in a 48-inch-diameter Ram-jet Engine

A preliminary evaluation was made of three types of combustor configurations in a 48-inch-diameter ram-jet engine to determine the one with the greatest promise of ultimately achieving efficient and stable combustion at low fuel-air rations and low combustor-inlet pressure. The combustor configurations evaluated include (1) a can-piloted configuration using a gutter-type flame holder with can-type pilot burners for stability, (2) an annular-piloted configuration using a gutter-type flame holder with an annular pilot burner to provide stability and to serve as a flow divider for confined fuel-air mixing, and (3) a can-type configuration using a flow divider for confined fuel-air mixing. No modifications were made to optimize the performance of any configuration. On the basis of this preliminary evaluation at an inlet-air temperature of about 990 degrees R, the annular-piloted combustor configuration appeared to have the greatest promise of permitting combustion efficiencies in excess of 0.90 at low fuel-air ratio with combustor-inlet pressures as low as 700 pounds per square foot absolute.

COMBUSTION - RAM-JET ENGINES↗
NASA NTRS1990

Trace gas emissions from burning Florida wetlands

Measurements of biomass burn-produced trace gases were obtained using a helicopter at low altitudes above burning Florida wetlands on November 9, 1987, and from both helicopter and light-aircraft samplings on November 7, 1988. Carbon dioxide normalized emission ratios for carbon monoxide, hydrogen, methane, total nonmethane hydrocarbons, and nitrous oxide were obtained over burning graminoid wetlands consisting primarily of Spartina bakeri and Juncus roemerianus. Some interspersed scrub oak and saw palmetto were also burned. No significant differences were observed in the emission ratios determined for these gases from samples collected over flaming, mixed, and smoldering phases of combustion during the 1987 fire. Combustion-categorized differences in emission ratios were small for the 1988 fire. Combustion efficiency was relatively good (low emission ratios for reduced gases) for both fires. It is believed that the consistently low emission ratios were a unique result of graminoid wetlands fires, in which the grasses and rushes burned rapidly down to standing water and were quickly extinguished. Consequently, the efficiency of the combustion was good and the amount and duration of smoldering combustion was greatly deminished.

Cofer, Wesley R., III↗
NASA NTRS1953

Effect of Diffuser Design, Diffuser-exit Velocity Profile and Fuel Distribution on Altitude Performance of Several Afterburner Configurations

An investigation was conducted in the NACA Lewis altitude wind tunnel to improve the altitude performance and operational characteristics of an afterburner primarily by modifying the diffuser-exit velocity profile by changes in diffuser design and by changing the fuel distribution and the flame holder. Twenty configurations, consisting of combinations of six diffuser geometries, six flame-holder types, and twelve fuel systems, were investigated. Data were obtained over a range of afterburner fuel-air ratios at diffuser-inlet total pressures from 2750 to 620 pounds per square foot. Changes in fuel distribution affected the fuel-air ratio at which peak combustion efficiency occurred as well as the efficiency level. Screeching combustion, which was most prevalent at low altitudes and medium-to-high fuel-air ratios, imposed a restriction on the operable range of a number of configurations.

Diffusers, Subsonic↗
NASA NTRS1980

Evaluation of a bulk calorimeter and heat balance for determination of supersonic combustor efficiency

Results are presented from the shakedown and evaluation test of a bulk calorimeter. The calorimeter is designed to quench the combustion at the exit of a direct-connect, hydrogen fueled, scramjet combustor model, and to provide the measurements necessary to perform an analysis of combustion efficiency. Results indicate that the calorimeter quenches reaction, that reasonable response times are obtained, and that the calculated combustion efficiency is repeatable within + or -3 percent and varies in a regular way with combustor model parameters such as injected fuel equivalence ratio.

Mcclinton, C. R.↗
NASA NTRS1955

Injection Principles from Combustion Studies in a 200-Pound-Thrust Rocket Engine Using Liquid Oxygen and Heptane

The importance of atomizing and mixing liquid oxygen and heptane was studied in a 200-pound-thrust rocket engine. Ten injector elements were used with both steel and transparent chambers. Characteristic velocity was measured over a range of mixture ratios. Combustion gas-flow and luminosity patterns within the chamber were obtained by photographic methods. The results show that, for efficient combustion, the propellants should be both atomized and mixed. Heptane atomization controlled the combustion rate to a much larger extent than oxygen atomization. Induced mixing, however, was required to complete combustion in the smallest volume. For stable, high-efficiency combustion and smooth engine starts, mixing after atomization was most promising.

Heidmann, M. F.↗