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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 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Effect of fuel zoning and fuel nozzle design on pollution emissions at ground idle conditions for a double-annular ram-induction combustor

An exhaust emission survey was conducted on a double-annular ram induction combustor at simulated ground idle conditions. The combustor was designed for a large augmented turbofan engine capable of sustained flight speeds up to Mach 3.0. The emission levels of total hydrocarbon (THC), carbon monoxide, carbon dioxide, and nitric oxide were measured. The effects of fuel zoning, fuel nozzle design, and operating conditions (inlet temperature and reference Mach number) on the level of these emissions were determined. At an overall combustor fuel/air ratio of 0.007, fuel zoning reduced THC emissions by a factor of 5 to 1. The reduction in THC emissions is attributed to the increase in local fuel/air ratio provided by the fuel zoning. An alternative method of increasing fuel/air ratio would be to operate with larger-than-normal compressor overboard bleed; however, analysis on this method indicated an increase in idle fuel consumption of 20 percent. The use of air-atomizing nozzles reduced the THC emissions by 2 to 1.

Clements, T. R.↗
NASA NTRS1973

Evaluation of an experimental short-length annular combustor: One-side-entry dilution airflow concept

A test program was conducted to evaluate an experimental short-length annular combustor that uses a one-side-entry dilution airflow concept. The combustor design features scoops on the outer liner for controlling the primary- and secondary-zone airflow distribution. Combustor inlet total pressures were limited to 62 N/sq cm (90 psia) with inlet-air temperatures from 590 K (600 F) to 890 K (1150 F). At a diffuser inlet Mach number of 0.25, the exit temperature pattern factor was 0.44 with an average exit temperature of 1436 K (2124 F) and a total pressure loss of 4.3 percent. At a diffuser inlet Mach number of 0.31, the exit temperature pattern factor was reduced to 0.29 with an average exit temperature of 1450 K (2151 F) and a total pressure loss of 6.1 percent. Nominal combustion efficiencies of 100 percent were obtained with the ASTM A-1 fuel. Exhaust gas emissions, smoke, and altitude relight data are included with exit-temperature profiles and distribution patterns.

Humenik, F. M.↗
NASA NTRS1974

Emissions of nitrogen oxides from an experimental hydrogen-fueled gas turbine combustor

The effect of operating variables of a hydrogen fueled combustor on exhaust concentrations of total oxides of nitrogen was determined at inlet-air temperature levels up to 810 K, pressure of 414,000N/sa m, and reference velocity of 21.3 m/sec. The combustor, which was originally designed for hydrocarbon fuel produced a NO(x) concentration of 380 ppm with hydrogen at 810 K inlet-air temperature. A reduction in NO(x) of about 30 % was obtained by modification to a lean or rich primary zone. The lowest NO(x) levels obtained with hydrogen were equivalent to those of the reference combustor burning hydrocarbon fuels.

Norgren, C. T.↗
NASA NTRS1974

Stability of partially lined combustors with distributed combustion

An integral method for theoretical evaluation of the stability of confined flows with mass sources is presented. This method is applied to evaluate the effect of partial length acoustic liners on liquid propellant combustors with distributed combustion sources. The combustor is modeled as a right circular cylinder terminated by a multiorifice (constant Mach number) nozzle. The spatial spread of combustion in the chamber is represented by an arbitrary number of pressure sensitive planar mass sources. On the cylindrical periphery of the chamber is an acoustic absorber of arbitrary length, position, and damping capability. Integral equations are developed from the partial differential equations describing the nonsteady flow and are solved iteratively. Results are given in terms of linear neutral stability plots using a pressure sensitive combustion response. Calculations indicate that increasing the axial spread of combustion increases the stability of the combustor and the effectiveness of the absorber. The most effective linear placement appears at the location of greatest combustion concentration.

Baer, M. R.↗
NASA NTRS1975

Pollution emissions from single swirl-can combustor modules at parametric test conditions

Exhaust pollutant emissions were measured from single swirl-can combustor modules operating over a pressure range of 69 to 276 N/sq cm (100 to 400 psia), over a fuel-air ratio range of 0.01 to 0.04, at an inlet air temperature of 733 K (860 F), and at a constant reference velocity of 23.2 m/sec). Many swirl-can module designs were evaluated; the 11 most promising designs exhibited oxides of nitrogen emission levels lower than that from conventional gas-turbine combustors. Although these single module test results are not necessarily indicative of the performance characteristics of a large array of modules, the results are very promixing and offer a number of module designs that should be tested in a full combustor.

Mularz, E. J.↗
NASA NTRS1975

Performance and Pollution Measurements of Two-Row Swirl-Can Combustor Having 72 Modules

A test program was conducted to evaluate the performance and gaseous-pollutant levels of an experimental full-annulus 72-module swirl-can combustor. A comparison of data with those for a 120-module swirl-can combustor showed no significant difference in performance or levels of gaseous pollutants. Oxides of nitrogen were correlated for the 72- and 120-swirl-can combustors by using a previously developed parameter.

Biaglow, James A.↗
NASA NTRS1974

Exhaust emissions of a double annular combustor: Parametric study

A full scale double-annular ram-induction combustor designed for Mach 3.0 cruise operation was tested. Emissions of oxides of nitrogen, carbon monoxide, unburned hydrocarbons, and smoke were measured over a range of combustor operating variables including reference velocity, inlet air temperature and pressure, and exit average temperature. ASTM Jet-A fuel was used for these tests. An equation is provided relating oxides of nitrogen emissions as a function of the combustor, operating variables. A small effect of radial fuel staging on reducing exhaust emissions (which were originally quite low) is demonstrated.

Schultz, D. F.↗
NASA NTRS1975

Experimental clean combustor program noise measurement addendum, phase 1

The test results of combustor noise measurements taken with waveguide probes are presented. Waveguide probes were shown to be a viable measurement technique for determining high sound pressure level broadband noise. A total of six full-scale annular combustors were tested and included the three advanced combustor designs: swirl-can, radial/axial, and double annular.

Emmerling, J. J.↗
NASA NTRS1975

Gas turbine combustor stabilization by heat recirculation

The feasibility of heat recirculation for stabilization of lean mixtures and emission reduction has been studied in detail for a typical aircraft gas turbine combustor. Thermodynamic calculations have indicated temperature and heat recirculation rates for operation of the combustor over a range of combustion zone equivalence ratios and for varying modes of desired engine operation. Calculations indicate the feasibility of stabilizing the combustion zone at equivalence ratios as low as 0.2 with achievable heat recirculation rates. Detailed chemical kinetic calculations suggest that combustor heat release is maintained with reaction completion substantially before the NO forming reactions, even though CO is rapidly oxidized in this same region.

Ganji, A.↗
NASA NTRS1975

The experimental clean combustor program: Description and status to November 1975

The generation of technology was studied for the development of advanced commercial CTOL aircraft engines with lower exhaust emissions than current aircraft. The program is in three phases. Phase 1, already completed, consisted of screening tests of low pollution combustor concepts. Phase 2, currently in progress, consists of test rig refinement of the most promising combustor concepts. Phase 2 test results are reported. Phase 3, also currently in progress, consists of incorporating and evaluating the best combustors as part of a complete engine. Engine test plans and pollution sampling techniques are described in this report. Program pollution goals, specified at engine idle and take-off conditions, are idle emission index value of 20 and 4 for carbon monoxide (CO) and total unburned hydrocarbons (THC), respectively, and at take-off are an oxides of nitrogen (NOx) emission index level of 10 and a smoke number of 15. Pollution data were obtained at all engine operating conditions. Results are presented in terms of emission index and also in terms of the Environmental Protection Agency's 1979 Standards Parameter.

Niedzwiecki, R. W.↗
NASA NTRS1976

Effect of premixing quality on oxides of nitrogen in gas turbine combustors foi HC

Experiments were conducted to determine the effectiveness of several premixing prevaporizing gas turbine combustor designs in reducing formation of oxides of nitrogen at the supersonic cruise condition. An atomized spray from a single injector mounted on the axis of the mixer tube produced a high initial concentration of fuel near the axis and only moderate premixed conditions entering the combustor. A fuel spray produced by 12 flush-mounted normal injection orifices in the mixer tube wall produced a good initial despersion of fuel and resulted in nearly complete premixing. Oxides of nitrogen emission levels of the order of 0.2 g NO2/kg fuel were obtained at 99 percent combustion efficiency at an equivalence ratio of 0.4. Overall total pressure drop was less than 3 percent through the 1-meter combustor module.

Roffe, G.↗
NASA NTRS1976

Effect of inlet temperature and pressure on emissions from a premixing gas turbine primary zone combustor

Experiments were conducted to determine the performance of a premixing prevaporizing gas turbine primary zone combustor design over a range of combustor inlet temperatures from 700 to 1000 K and a range of inlet pressures from 40 to 240 N/sq cm. The 1 meter long combustor could be operated at pressures up to and including 120 N/sq cm without autoignition in the premixing duct or flashback from the stabilized combustion zone. Autoignition occurred in the mixer tube at the 240 N/sq cm pressure level with an entrance temperature of 830 K and a mixer residence time of 4 msec. Measured NOx level, combustion inefficiency, and hydrocarbon emission index correlated well with adiabatic flame temperature. The NOx levels varied from approximately 0.2 to 2.0 g NO2/kg fuel at combustion inefficiencies from 4 to 0.04 percent, depending upon adiabatic flame temperature and pressure. Measured NOx levels were sensitive to pressure. Tests were made at equivalence ratios ranging from 0.35 to 0.65. The overall total pressure drop for the configuration varied slightly with reference velocity and equivalence ratio, but never exceeded 3 percent.

Roffe, G.↗
NASA NTRS1976

Design study of a gas turbine combustor with heat recirculation

A means of avoiding stoichiometric combustion, reducing emissions, and yet providing stable burning for lean mixtures is based on the use of heat recirculation rather than flow recirculation. This paper is concerned with the calculations of the design parameters of a gas turbine combustor with heat exchanger to produce the desired preheat temperature. The combustor inlet temperature, maximum temperature, equivalence ratio and recirculated heat are determined by thermodynamic analysis. The heat transfer analysis then provides the dimensions of the system to produce the predetermined boundary conditions. It is indicated that practical combustor design may be feasible for reactant mixtures as low as equivalence ratio 0.2.

Ganji, A.↗
NASA NTRS1976

Results of the NASA/General Electric Experimental Clean Combustor Program

The NASA/General Electric Experimental Clean Combustor Program is a multi-year, major contract effort. Primary program objectives are the generation of technology for development of advanced commercial CTOL engines with lower exhaust emissions than current aircraft and, demonstrations of this technology in a full-scale CF6-50C engine in 1976. This paper describes pollution and performance goals, Phase I and II test results and Phase III hardware, pollution sampling techniques and test plans. Pollution results are presented in emission index and Environmental Protection Agency 1979 Standard Parameters (EPAP). Best results were obtained with a double annular combustor concept. This concept, which incorporates multistage burning, produced EPAP values extrapolated to CF6-50C engine conditions for CO, HC, and NOx of 3.3, 0.3 and 4.5, respectively. These represent respective CO, HC and NOx percentage reductions of 69, 93 and 42%, compared to current CF6-50 engine values. The combustor also met development engine performance requirements.

Gleason, C. C.↗
NASA NTRS1976

Effect of exhaust gas recirculation on emissions from a flame-tube combustor using Liquid Jet A fuel

The effects of uncooled exhaust gas recirculation as an inert diluent on emissions of oxides of nitrogen (NO + NO2) and on combustion efficiency were investigated. Ratios of recirculated combustion products to inlet airflow were varied from 10 to 80 percent by using an inlet air ejector nozzle. Liquid Jet A fuel was used. The flame-tube combustor was 10.2 cm in diameter. It was operated with and without a flameholder present. The combustor pressure was maintained constant at 0.5 MPa. The equivalence ratio was varied from 0.3 to 1.0. The inlet air temperature was varied from 590 to 800 K, and the reference velocity from 10 to 30 m/sec. Increasing the percent recirculation from 10 to 25 had the following effects: (1) the peak NOx emission was decreased by 37 percent, from 8 to 5 g NO2/kg fuel, at an inlet air temperature of 590 K and a reference velocity of 15 m/sec; (2) the combustion efficiency was increased, particularly at the higher equivalence ratios; and (3) for a high combustion efficiency of greater than 99.5 percent, the range of operation of the combustor was nearly doubled in terms of equivalence ratio. Increasing the recirculation from 25 to 50 percent did not change the emissions significantly.

Marek, C. J.↗
NASA NTRS1976

Evaluation of a staged fuel combustor for turboprop engines

Proposed EPA emission regulations require emission reduction by 1979 for various gas turbine engine classes. Extensive combustion technology advancements are required to meet the proposed regulations. The T56 turboprop engine requires CO, UHC, and smoke reduction. A staged fuel combustor design was tested on a combustion rig to evaluate emission reduction potential in turboprop engines from fuel zoning. The can-type combustor has separately fueled-pilot and main combustion zones in series. The main zone fueling system was arranged for potential incorporation into the T56 with minor or no modifications to the basic engine. Three combustor variable geometry systems were incorporated to evaluate various airflow distributions. Emission results with fixed geometry operation met all proposed EPA regulations over the EPA LTO cycle. CO reduction was 82 percent, UHC reduction was 96 percent, and smoke reduction was 84 percent. NOx increased 14 percent over the LTO cycle. At high power, NOx reduction was 40 to 55 percent. This NOx reduction has potential application to stationary gas turbine powerplants which have different EPA regulations.

Verdouw, A. J.↗
NASA NTRS1976

The impact of emission standards on the design of aircraft gas turbine engine combustors

The advent of environmental standards for controlling aircraft gas turbine engine emissions has led to a reevaluation of combustor design techniques. Effective emission control techniques have been identified and a wide spectrum of potential applications for these techniques to existing and advanced engines are being considered. Results from advanced combustor concept evaluations and from fundamental experiments are presented and discussed and comparisons are made with existing EPA emission standards and recommended levels for high altitude cruise. The impact that the advanced low emission concepts may impose on future aircraft engine combustor designs and related engine components is discussed.

Rudey, R. A.↗
NASA NTRS1977

Experimental and analytical separation of hydrodynamic, entropy and combustion noise in a gas turbine combustor

This paper deals with noise sources which are central to the problem of core engine noise in turbopropulsion systems. The sources dealt with are entropy noise and direct combustion noise, as well as a non-propagating psuedosound which is hydrodynamic noise. It is shown analytically and experimentally that a transition can occur from a combustion noise dominant situation to an entropy noise dominant case if the contraction of a terminating nozzle to the combustor is high enough. In the combustor tested, entropy noise is the dominant source for propagational noise if the combustor is choked at the exit. Analysis techniques include spectral, cross spectral, cross correlation, and ordinary and partial coherence analysis. Measurements include exterior and interior fluctuating and mean pressures and temperatures.

Muthukrishnan, M.↗