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NASA NTRS1988

Assessment, development and application of combustor aerothermal models

The gas turbine combustion system design and development effort is an engineering exercise to obtain an acceptable solution to the conflicting design trade-offs between: combustion efficiency, gaseous emissions, smoke, ignition, restart, lean blowout, burner exit temperature quality, structural durability, and life cycle cost. For many years, these combustor design trade-offs have been carried out with the help of fundamental reasoning and extensive component and bench testing, backed by empirical and experience correlations. Recent advances in the capability of computational fluid dynamics (CFD) codes have led to their application to complex three-dimensional flows such as those in the gas turbine combustor. A number of US. Government and industry sponsored programs have made significant contributions to the formulation, development, and verification of an analytical combustor design methodology which will better define the aerothermal loads in a combustor, and be a valuable tool for design of future combustion systems. The contributions made by NASA Hot Section Technology (HOST) sponsored Aerothermal Modeling and supporting programs are described in this paper.

Aerothermal modeling↗

NASA NTRS2015

Effects of Burning Alternative Fuel in a 5-Cup Combustor Sector

A goal of NASA's Environmentally Responsible Aviation (ERA) program is to develop a combustor that will reduce the NOx emissions and that can burn both standard and alternative fuels. To meet this goal, NASA partnered with General Electric Aviation to develop a 5-cup combustor sector; this sector was tested in NASA Glenn's Advanced Subsonic Combustion Rig (ASCR). To verify that the combustor sector was fuel-flexible, it was tested with a 50-50 blend of JP-8 and a biofuel made from the camelina sativa plant. Results from this test were compared to results from tests where the fuel was neat JP-8. Testing was done at three combustor inlet conditions: cruise, 30% power, and 7% power. When compared to burning JP-8, burning the 50-50 blend did not significantly affect emissions of NOx, CO, or total hydrocarbons. Furthermore, it did not significantly affect the magnitude and frequency of the dynamic pressure fluctuations.

alternative fuel↗

NASA NTRS2018

Semi-Empirical Modeling and Prediction of Direct Combustor Noise

Due to turbofan design trends, advances in fan-noise-mitigation techniques, and expected aircraft configuration changes, the impact on overall airport community noise from certain aircraft-propulsion noise sources, such as fan and jet noise, will be reduced at all certification points in the future. The situation with regard to turbofan-core noise is less clear however. Combustor noise in particularâ€”which presently is an important core-noise component, but not dominant from an overall noise perspectiveâ€”could become more prominent in upcoming high-power-density core designs. Its detailed source structure in the combustor, including the indirect combustor-noise mechanism, and the effects of the propagation path through the engine and exhaust nozzle need to be better understood. This report reviews the current status of modeling and prediction of direct combustor noise. This work is carried out under the NASA Advanced Air Vehicles Program, Advanced Air Transport Technology Project, Aircraft Noise Reduction Subproject.

Turbomachinery Noise↗

NASA NTRS2022

Impact of Future Low-Emissions Combustor Technology on Acoustic Scaling Laws

The overall goal of a recent research project at the Raytheon Technologies Research Center, under National Aeronautics and Space Administration sponsorship, was to develop a first-ofits-kind database of detailed unsteady measurements characterizing noise sources of far-term advanced low-emissions aero-combustors. The program addressed the need for fundamental combustion-noise experiments which, in the near term, enable improvements to reduced-order models for use in system level noise assessments at the preliminary design stage for advanced air transport vehicles. In addition, optical measurement techniques were refined and validated for usage at the higher pressures and temperatures relevant to future combustor designs. In the long term, results from this program can be utilized to validate high-fidelity prediction methods suited for detailed multi-disciplinary acoustics/emissions combustor design. This paper concentrates on one aspect of the overall project, namely an examination of the legacy scaling laws for broadband combustor noise.

aeroacoustics, turbomachinery noise, combustor noi↗

NASA NTRSDate not supplied

Impact of Future Low-Emissions Combustor Technology on Acoustic Scaling Laws

The work is a summary and further analysis of unlimited-rights acoustic data acquired under NRA Contract NNC16CA39C (September 2016 - November2019) with United Technologies (now Raytheon Technologies) Research Center, East Hartford, Connecticut. NASA/CR-20205011099/VOL1, https://ntrs.nasa.gov/citations/20205011099. The overall goal of the Contract was to develop a first-of-its- kind database of detailed unsteady measurements characterizing noise sources of far-term advanced low-emissions aero-combustors. The program addressed the need for fundamental combustion-noise experiments which, in the near term, enable improvements to reduced-order models for use in system level noise assessments at the preliminary design stage for advanced air transport vehicles. In addition, optical measurement techniques were refined and validated for usage at the higher pressures and temperatures relevant to future combustor designs. In the long term, results from this program can be utilized to validate high-fidelity prediction methods suited for detailed multi-disciplinary acoustics/emissions combustor design. This presentation concentrates on one aspect of the overall project, namely an examination of the legacy scaling laws for broadband combustor noise.

aeroacoustics, turbomachinery noise, combustor noi↗

NASA NTRS1969

Development of an Advanced Annular Combustor

The objective of the effort described in this report was to determine the structural durability of a full-scale advanced annular turbojet combustor using ASTM A-1 type fuel and operating at conditions typical of advanced supersonic aircraft. A full-scale annular combustor of the ram-induction type was fabricated and subjected to a 325-hour cyclic endurance test at conditions representative of operation in a Mach 3.0 aircraft. The combustor exhibited extensive cracking and scoop burning at the end of the test program. But these defects had no appreciable effect on combustor performance, as performance remained at a high level throughout the endurance program. Most performance goals were achieved with pressure loss values near 6% and 8%, and temperature rise variation ratio (deltaTVR) values near 1.25 and l.22 at takeoff and cruise conditions, respectively. Combustion efficiencies approached l004 and the exit radial temperature profiles were approximately as desired.

Rusnak, J. P.↗

NASA NTRS1972

Preliminary results of large supersonic burning combustor testing.

A hydrogen fueled supersonic-burning combustor 18 in. in diameter, which is equivalent to that of an engine about 6 ft in diameter, was tested as a direct-connected duct at inlet conditions which simulated Mach 8 flight at 115,000 ft alt. A synthetic air consisting of oxygen with 39% nitrogen and 38% water vapor at a total temperature of 4500 deg R and a total pressure of 300 psia was supplied to the combustor inlet by a hydrazine-nitrogen tetroxide hot gas generator which maintained a uniform inlet flow Mach number of 2.8. The large combustor size required a new approach to fuel injector design. Some hydrogen was injected through flush-wall injectors, but most was injected from two rows of swept and tapered struts immersed in the flow stream. Supersonic combustion was obtained at hydrogen equivalence ratios of 0.94 without encountering thermal cho king. Wall static pressures, and the radial distribution of hydrogen, Pitot pressure, and Mach number were determined at the combustor exit.

Metzler, A. J.↗

NASA NTRS1973

The effect of water injection on nitric oxide emissions of a gas turbine combustor burning ASTM Jet-A fuel

Tests were conducted to determine the effect of water injection on oxides of nitrogen (NOx) emissions of a full annular, ram induction gas turbine combustor burning ASTM Jet-A fuel. The combustor was operated at conditions simulating sea-level takeoff and cruise conditions. Water at ambient temperature was injected into the combustor primary zone at water-fuel ratios up to 2. At an inlet-air temperature of 589 K (600 F) water injection decreased the NOx emission index at a constant exponential rate: NOx = NOx (o) e to the -15 W/F power (where W/F is the water-fuel ratio and NOx(o) indicates the value with no injection). The effect of increasing combustor inlet-air temperature was to decrease the effect of the water injection. Other operating variables such as pressure and reference Mach number did not appear to significantly affect the percent reduction in NOx. Smoke emissions were found to decrease with increasing water injection.

Marchionna, N. R.↗

NASA NTRS1974

The experimental clean combustor program: Description and status

The objectives, program plan, schedule, pollution and performance goals, program approaches to pollution reduction, and status of the project are presented. Advanced supersonic transport and combustion noise programs are also discussed. The phases of the program consist of the following: (1) screening of low pollutant combustors, (2) refinement of the best combustors, and (3) engine demonstration of the best combustors. It is stated that all combustor designs investigated appear capable of achieving the idle pollutant goals with additional development effort.

Niedzwiecki, R. W.↗

NASA NTRS1974

Effect of diffuser bleed on performance of an annular swirl can combustor

A full-scale annular swirl can combustor equipped with diffuser bleed capability was tested at low fuel-air ratio, burning natural gas fuel. Test results show that, by drawing off a small amount of air on the inner or outer diffuser wall, the radial profile of combustor exit temperature could be changed from a strongly tip biased, to a strongly hub biased shape. Furthermore, combustor total pressure loss could be reduced by about one third by using a total bleed rate equal to 6 percent of diffuser inlet flow. At simulated engine idle conditions combustion efficiency was increased from 64 percent to nearly 100 percent by using 3 percent outer wall bleed. Combustor blowout performance was also improved significantly by using this diffuser bleed scheme.

Juhasz, A. J.↗

NASA NTRS1975

Exhaust pollutant emissions from swirl-can combustor module arrays at parametric test conditions

Improved designs of swirl-can combustor modules were tested using seven-module arrays in a combustor. The combustor was operated over a pressure range of 69 to 207 N/sq cm, a fuel-air ratio range of 0.015 to 0.046, at a constant inlet air temperature of 733 K, and at reference velocities of 23.9 and 30.6 m/sec. The three designs tested performed with high combustion efficiency at all conditions tested and exhibited oxides of nitrogen emissions substantially lower than that of conventional gas turbine combustors. A correlating parameter used to extrapolate oxides of nitrogen emissions to full power or takeoff conditions for large commercial turbofan engines predicts oxides of nitrogen emissions somewhat higher than those specified in the 1979 government emissions standards.

Mularz, E. J.↗

NASA NTRS1975

Effect of design features on performance of a double-annular ram-induction combustor

An extensive test program was undertaken to determine the effect of many design features such as the size and number of air scoops, and the type of diffuser airflow distribution to use to optimize performance of a double-annular ram-induction combustor of 94 cm outer diameter. Six combustor configurations were tested. It was found that a snouted double annular combustor built with 256 ram-induction air scoops with a combustor open area giving a total pressure loss of 5.0 percent at a diffuser inlet Mach number of 0.25 gave the best overall performance of the configurations tested.

Schultz, D. F.↗

NASA NTRS1975

Experimental clean combustor program, phase 1

A program of screening three low emission combustors for conventional takeoff and landing, by testing and analyzing thirty-two configurations is presented. Configurations were tested that met the emission goals at idle operating conditions for carbon monoxide and for unburned hydrocarbons (emission index values of 20 and 4, respectively). Configurations were also tested that met a smoke number goal of 15 at sea-level take-off conditions. None of the configurations met the goal for oxides of nitrogen emissions at sea-level take-off conditions. The best configurations demonstrated oxide of nitrogen emission levels that were approximately 61 percent lower than those produced by the JT9D-7 engine, but these levels were still approximately 24 percent above the goal of an emission index level of 10. Additional combustor performance characteristics, including lean blowout, exit temperature pattern factor and radial profile, pressure loss, altitude stability, and altitude relight characteristics were documented. The results indicate the need for significant improvement in the altitude stability and relight characteristics. In addition to the basic program for current aircraft engine combustors, seventeen combustor configurations were evaluated for advanced supersonic technology applications. The configurations were tested at cruise conditions, and a conceptual design was evolved.

Roberts, R.↗

NASA NTRS1976

Effect of flame stabilizer design on performance and exhaust pollutants of a two-row 72-module swirl-can combustor

A test program was conducted to evaluate the effects of four flame stabilizer designs on the performance and gaseous pollutant levels of an experimental full-annular swirl-can combustor. Combustor operating parameters, including inlet-air temperature, reference velocity, and fuel-air ratio, were set to simulate conditions in a 30:1 pressure ratio engine. Combustor inlet total pressure was held constant at 6 atm due to the facility limit. Combustor performance and gaseous pollutant levels were strongly affected by the geometry and resulting total pressure loss of the four flame stabilizer designs investigated. The addition of shrouds to two designs produced an 18 to 22% decrease in the combustion chamber pressure loss and thus resulted in doubling the exit temperature pattern factor and up to 42% higher levels of oxides of nitrogen. A previously developed oxides of nitrogen correlating parameter agreed with each model within an emission index of plus or minus 1 but was not capable of correlating all models together.

Biaglow, J. A.↗

NASA NTRS1976

Experimental clean combustor program, phase 2

The alternate fuels investigation objective was to experimentally determine the impacts, if any, on exhaust emissions, performance, and durability characteristics of the hybrid and vorbix low pollution combustor concepts when operated on test fuels which simulate composition and property changes which might result from future broadened aviation turbine fuel specifications or use of synthetically derived crude feedstocks. Results of the program indicate a significant increase in CO and small NOX increase in emissions at idle for both combustor concepts, and an increase in THC for the vorbix concept. Minimal impact was observed on gaseous emissions at high power. The vorbix concept exhibited significant increase in exhaust smoke with increasing fuel aromatic content. Altitude stability was not affected for the vorbix combustor, but was substantially reduced for the hybrid concept. Severe carbon deposition was observed in both combustors following limited endurance testing with No. 2 home heat fuel. Liner temperature levels were insensitive to variations in aromatic content over the range of conditions investigated.

Roberts, R.↗

NASA NTRS1976

Experimental clean combustor program: Noise study

Under a Noise Addendum to the NASA Experimental Clean Combustor Program (ECCP) internal pressure fluctuations were measured during tests of JT9D combustor designs conducted in a burner test rig. Measurements were correlated with burner operating parameters using an expression relating farfield noise to these parameters. For a given combustor, variation of internal noise with operating parameters was reasonably well predicted by this expression but the levels were higher than farfield predictions and differed significantly among several combustors. For two burners, discharge stream temperature fluctuations were obtained with fast-response thermocouples to allow calculation of indirect combustion noise which would be generated by passage of the temperature inhomogeneities through the high pressure turbine stages of a JT9D turbofan engine. Using a previously developed analysis, the computed indirect combustion noise was significantly lower than total low frequency core noise observed on this and several other engines.

Sofrin, T. G.↗

NASA NTRS1976

Exhaust gas emissions of a vortex breakdown stabilized combustor

Exhaust gas emission data are described for a swirl stabilized continuous combustor. The combustor consists of confined concentric jets with premixed fuel and air in the inner jet and air in the outer jet. Swirl may be induced in both inner and outer jets with the sense of rotation in the same or opposite directions (co-swirl and counter-swirl). The combustor limits NO emissions by lean operation without sacrificing CO and unburned hydrocarbon emission performance, when commercial-grade methane and air fired at one atmosphere without preheat are used. Relative swirl direction and magnitude are found to have significant effects on exhaust gas concentrations, exit temperatures, and combustor efficiencies. Counter-swirl gives a large recirculation zone, a short luminous combustion zone, and large slip velocities in the interjet shear layer. For maximum counter-swirl conditions, the efficiency is low.

Yetter, R. A.↗

NASA NTRS1977

Direct-connect tests of hydrogen-fueled supersonic combustors

Direct-connect tests of hydrogen-fueled supersonic combustors were performed using arc-heated air at combustor inlet Mach numbers of 2.9 to 3.2. Various axisymmetric combustor geometries of 5.89 and 6.96 cm (inner diameter) inlet were investigated; the fuel was injected from the wall either from a ring of equally spaced holes normal to the air stream, or from a circumferential slot oriented 45 deg downstream. The hole-type injectors consistently gave better results. The effects of various parameters are examined, and the performance comparison procedure is described. A theoretical model of the supersonic combustion process which includes a precombustion shock-compression is used to explain the character of the observed pressure distributions and to assess the effects of the measured heat transfer rates, deduced wall shear, and combustor geometry on performance.

Waltrup, P. J.↗