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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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1948

Simulated Altitude Performance of Combustors for the Westinghouse 24C Jet Engine I-24C-2 Combustor

A Westinghouse 24C-2 combustor was investigated at conditions simulating operation of the 24C Jet engine at zero ram over ranges of altitude and engine speed. The investigation was conducted to determine the altitude operational limits, that is, the maximum altitude for various engine speeds at which an average combustor-outlet gas temperature sufficient for operation of the jet engine could be obtained. Information was also obtained regarding the character of the flames, the combustion efficiency, the combustor-outlet gas temperature and velocity distributions, the extent of afterburning, the flow characteristics of the fuel manifolds, the combustor inlet-to-outlet total-pressure drop, and the durability of the combustor basket. The results of the investigation indicated that the altitude operational limits for zero ram decreased from 12,000 feet at an engine speed of 4000 rpm to a minimum of 9000 feet at 6000 rpm, and thence increased to 49,000 feet at 12,000 rpm.. At altitudes below the operational limits, flames were essentially steady, but, at altitudes above the operational limits, flames were often cycling and either blew out or caused violent explosions and vibrations. At conditions on the altitude operational limits the type of combustion varied from steady to cycling with increasing fuel-air ratio and the reverse occurred with decreasing fuel-air ratio. In the region of operation investigated, the combustion efficiency ranged from 75 to 95 percent at altitudes below the operational limits and dropped to 55 percent or less at some altitudes above the operational limits. The deviations in the local combustor-outlet gas temperatures were within +20 to -30 percent of the mean combustor temperature rise for inlet-air temperatures at the low end of the range investigated, but became more uneven (up to +/-100 percent) with increasing inlet-air temperatures. The distribution of the combustor-outlet gas velocity followed a similar trend. Practically no afterburning downstream of the combustor outlet occurred. At conditions of high inlet-air temperature several factors indicated that fuel vapor or air formed in the fuel manifolds and adversely affected combustion. The combustor inlet-to-outlet total-pressure drop can be correlated as a function of the ratio of the combustion-air inlet density to outlet density and of the inlet dynamic pressure. The walls of the combustor basket were warped and burned during 29 hours of operation.

Manganiello, Eugene J.↗
DOE OSTI2025

A Comparison of Novel Woodstove Catalytic Combustors to Improve Low Temperature Light off and Catalyst Stability at High Temperatures and Corrosive Environments, Including Stability of Catalytic Combustors used in Coastal Environments

Wood heaters with catalytic combustors often emit higher levels of particulate matter (PM) and volatile organic compounds (VOCs) at startup before the catalytic combustor is at operational temperature. This project aims to advance wood heater catalytic combustor technology by studying degradation mechanisms in failed field catalysts and using those insights to develop an improved catalytic combustor. Findings will provide critical insights into the degradation mechanisms of wood heater catalytic combustors and inform the design of catalysts that deliver improved efficiency, durability, and performance compared to current commercial products. The outcomes of this project will not only improve catalytic combustor technology but also contribute to operational guidelines to extend catalyst lifetimes.

99 GENERAL AND MISCELLANEOUS↗
NASA NTRS1947

Effect of combustor-inlet conditions on performance of an annular turbojet combustor

The combustion performance, and particularly the phenomenon of altitude operational limits, was studied by operating the annular combustor of a turbojet engine over a range of conditions of air flow, inlet pressure, inlet temperature, and fuel flow. Information was obtained on the combustion efficiencies, the effect on combustion of inlet variables, the altitude operational limits with two different fuels, the pressure losses in the combustor, the temperature and velocity profiles at the combustor outlet, the extent of afterburning, the fuel-injection characteristics, and the condition of the combustor basket.

Childs, J Howard↗
NASA NTRS1947

Simulated Altitude Performance of Combustors for the 24C Jet Engine: 24C-4 Combustor - 2

The performance of a 24C-4 combustor was investigated with three different combustor baskets and five modifications of these baskets at conditions simulating static (zero-ram) operation of the 24C jet engine over ranges of altitude and engine speed to determine and improve the altitude operational limits of the 24C combustor. Information was also obtained regarding combustion characteristics, the fuel-flow characteristics of the fuel manifolds, and the combustor total-pressure drop. NACA modifications, which consisted of blocking rows of holes on the baskets, increased the minimum point on the altitude-operational-limit curve, which occurs at low engine speeds, for a narrow-upstream-end basket by 8000 feet (from 23, 000 to 31,000 ft_ and for a wide-upstream-end basket by 21,000 feet (from 12, 000 to 34,000 ft). These improvements were approximately maintained over the entire range of engine speeds investigated.

Bernardo, Everett↗
NASA NTRS1975

Variable combustor geometry for improving the altitude relight capability of a double annular combustor

A test program was conducted to evaluate several ways of improving the altitude relight capability of a double annular ram-induction combustor designed for Mach 3.0 cruise operation. Using various techniques including two modes of simulated combustor variable geometry, altitude relights were obtained down to a pressure of 3.6 newtons per square centimeter against 7.05 newtons per square centimeter for the unmodified combustor. This was at a test condition of 0.05 reference Mach number using ambient temperature inlet air and fuel.

Schultz, D. F.↗
NASA NTRS1975

Ground idle performance improvement of a double-annular combustor by using simulated variable combustor geometry

A test program was undertaken to determine if variable combustor geometry could be used to reduce exhaust emissions of a low-pressure-ratio jet engine operating at ground idle conditions. Three techniques for varying combustor geometry were simulated. Other techniques evaluated were radial fuel staging and the use of preheated fuel. When simulated variable combustor geometry was employed with radial fuel staging, combustion efficiency at a fuel-air ratio of 0.01 was increased from 77 to 95 percent, and exhaust emissions of unburned hydrocarbons and carbon monoxide were significantly reduced.

Schultz, D. F.↗
NASA NTRS1976

Low pollution combustor designs for CTOL engines - Results of the Experimental Clean Combustor Program

The NASA/Pratt & Whitney Aircraft Experimental Clean Combustor Program is a multi-year, major contract effort. Primary program objectives are the generation of combustor technology for development of advanced commercial CTOL engines with lower exhaust emissions than current aircraft and demonstration of this technology in a full-scale JT9D engine in 1976. This paper describes the pollution and performance goals, Phase I and II test results, and the Phase III combustor hardware, pollution sampling techniques, and test plans. Best results were obtained with the Vorbix concept which employs multiple burning zones and improved fuel preparation and distribution. Substantial reductions were achieved in all pollutant categories, meeting the 1979 EPA standards for NOx, THC, and smoke when extrapolated to JT9D cycle conditions. The Vorbix concept additionally demonstrated the capability for acceptable altitude relight and did not appear to have unsolvable durability or exit temperature distribution problems.

Roberts, R.↗
NASA NTRS1976

Effect of ceramic coating of JT8D combustor liner on maximum liner temperatures and other combustor performance parameters

The effect of ceramic coating of a JT8D combustor liner was investigated at simulated cruise and takeoff conditions with two fuels of widely different aromatic contents. Substantial decreases in maximum liner temperatures and flame radiation values were obtained with the ceramic-coated liner. Small reductions in exhaust gas smoke concentrations were observed with the ceramic-coated liner. Other performance parameters such as combustion efficiency and emissions of unburned hydrocarbons, CO, and NOx were not affected significantly. No deterioration of the ceramic coating was observed after about 6 hours of cyclic operation including several startups and shutdowns.

Butze, H. F.↗
NASA NTRS1948

Simulated Altitude Performance of Combustor of Westinghouse 19XB-1 Jet-Propulsion Engine

A 19XB-1 combustor was operated under conditions simulating zero-ram operation of the 19XB-1 turbojet engine at various altitudes and engine speeds. The combustion efficiencies and the altitude operational limits were determined; data were also obtained on the character of the combustion, the pressure drop through the combustor, and the combustor-outlet temperature and velocity profiles. At altitudes about 10,000 feet below the operational limits, the flames were yellow and steady and the temperature rise through the combustor increased with fuel-air ratio throughout the range of fuel-air ratios investigated. At altitudes near the operational limits, the flames were blue and flickering and the combustor was sluggish in its response to changes in fuel flow. At these high altitudes, the temperature rise through the combustor increased very slowly as the fuel flow was increased and attained a maximum at a fuel-air ratio much leaner than the over-all stoichiometric; further increases in fuel flow resulted in decreased values of combustor temperature rise and increased resonance until a rich-limit blow-out occurred. The approximate operational ceiling of the engine as determined by the combustor, using AN-F-28, Amendment-3, fuel, was 30,400 feet at a simulated engine speed of 7500 rpm and increased as the engine speed was increased. At an engine speed of 16,000 rpm, the operational ceiling was approximately 48,000 feet. Throughout the range of simulated altitudes and engine speeds investigated, the combustion efficiency increased with increasing engine speed and with decreasing altitude. The combustion efficiency varied from over 99 percent at operating conditions simulating high engine speed and low altitude operation to less than 50 percent at conditions simulating operation at altitudes near the operational limits. The isothermal total pressure drop through the combustor was 1.82 times as great as the inlet dynamic pressure. As expected from theoretical considerations, a straight-line correlation was obtained when the ratio of the combustor total pressure drop to the combustor-inlet dynamic pressure was plotted as a function of the ratio of the combustor-inlet air density to the combustor-outlet gas density. The combustor-outlet temperature profiles were, in general, more uniform for runs in which the temperature rise was low and the combustion efficiency was high. Inspection of the combustor basket after 36 hours of operation showed very little deterioration and no appreciable carbon deposits.

Childs, J. Howard↗
NASA NTRS1958

A Preliminary Investigation of the Performance of a Short-length Turbojet Combustor Using Vaporized Hydrocarbon Fuels

Two short turbojet combustors designed for use with vaporized hydrocarbon fuels were tested in a one-quarter annular duct. The experimental combustors consisted of many small "swirl-can" combustor elements manifolded together. This design approach allowed the secondary mixing zone to be considerably reduced over that of conventional combustors. The over-all combustion lengths, for the two configurations were 13.5 and 11.0 inches, approximately one-half the length of the shortest conventional combustors. These short combustors did not provide combustion efficiencies as high as those for conventional combustors at low pressures. However, over the range of combustor-inlet total-pressures expected in aircraft capable of flight at Mach numbers of 2.5 and above, these short combustors gave very high efficiencies. A combustion efficiency of 97 percent was obtained at a combustor-inlet total-pressure of 25.0 inches of mercury absolute, reference velocity of 120 feet per second, and inlet-air total temperature of 1160 deg R. By proportioning the fuel flow between the manifold rows of can combustor elements, control of the combustor-outlet radial total-temperature profile was demonstrated. Combustor totalpressure loss varied from 0.75 percent of the inlet total pressure at isothermal conditions and a reference velocity of 75 feet per second to 5.5 percent at a total-temperature ratio of 1.8 and a reference velocity of 180 feet per second.

Jones, R. E.↗
NASA NTRS1976

Experimental clean combustor program, phase 2

Combustor pollution reduction technology for commercial CTOL engines was generated and this technology was demonstrated in a full-scale JT9D engine in 1976. Component rig refinement of the two best combustor concepts were tested. These concepts are the vorbix combustor, and a hybrid combustor which combines the pilot zone of the staged premix combustor and the main zone of the swirl-can combustor. Both concepts significantly reduced all pollutant emissions relative to the JT9D-7 engine combustor. However, neither concept met all program goals. The hybrid combustor met pollution goals for unburned hydrocarbons and carbon monoxide but did not achieve the oxides of nitrogen goal. This combustor had significant performance deficiencies. The Vorbix combustor met goals for unburned hydrocarbons and oxides of nitrogen but did not achieve the carbon monoxide goal. Performance of the vorbix combustor approached the engine requirements. On the basis of these results, the vorbix combustor was selected for the engine demonstration program. A control study was conducted to establish fuel control requirements imposed by the low-emission combustor concepts and to identify conceptual control system designs. Concurrent efforts were also completed on two addendums: an alternate fuels addendum and a combustion noise addendum.

Roberts, R.↗