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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Parameters controlling nitric oxide emissions from gas turbine combustors.

Nitric oxide forms in the primary zone of gas turbine combustors where the burst gas composition is close to stoichiometric and gas temperatures are highest. It has been found that combustor air inlet conditions, mean primary zone fuel-air ratio, residence time, and the uniformity of the primary zone are the most important variables affecting nitric oxide emissions. Relatively simple models of the flow in a gas turbine combustor, coupled with a rate equation for nitric oxide formation via the Zeldovich mechanism are shown to correlate the variation in measured NO sub x emissions. Data from a number of different combustor concepts are analyzed and shown to be in reasonable agreement with predictions. The NO sub x formation model is used to assess the extent to which an advanced combustor concept, the NASA swirl can, has produced a lean well-mixed primary zone generally believed to be the best low NO sub x emissions burner type.

Heywood, J. B.↗
NASA NTRS1973

Parametric test results of a swirl-can combustor.

Pollutant levels of oxides of nitrogen, unburned hydrocarbons, and carbon monoxide were measured for three models of an experimental, annular swirl can combustor. The combustor was 1.067 meters in outer diameter, incorporated 120 modules, and was specifically designed for elevated exit temperature performance. Test conditions included combustor inlet temperatures of 589, 756, and 839 K, inlet pressures of 3 to 6.4 atmospheres, reference velocities of 21 to 38 meters per second, and combustor equivalence ratios, based on total combustor flows, of 0.206 to 1.028. Maximum oxides of nitrogen emission index values occurred at an equivalence ratio of 0.7 with lower values measured for both higher and lower equivalence ratios.

Niedzwiecki, R. W.↗
NASA NTRS1973

Parameters controlling nitric oxide emissions from gas turbine combustors

Nitric oxide forms in the primary zone of gas turbine combustors where the burnt gas composition is close to stoichiometric and gas temperatures are highest. It was found that combustor air inlet conditions, mean primary zone fuel-air ratio, residence time, and the uniformity of the primary zone are the most important variables affecting nitric oxide emissions. Relatively simple models of the flow in a gas turbine combustor, coupled with a rate equation for nitric oxide formation via the Zeldovich mechanism are shown to correlate the variation in measured NOx emissions. Data from a number of different combustor concepts are analyzed and shown to be in reasonable agreement with predictions. The NOx formation model is used to assess the extent to which an advanced combustor concept, the NASA swirl can, has produced a lean well-mixed primary zone generally believed to be the best low NOx emissions burner type.

Heywood, J. B.↗
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 NTRS1976

The pollution reduction technology program for can-annular combustor engines - Description and results

Pollutant reduction and performance characteristics were determined for three successively more advanced combustor concepts. Program Element I consisted of minor modifications to the current production JT8D combustor and fuel system to evaluate means of improved fuel preparation and changes to the basic airflow distribution. Element II addressed versions of the two-staged Vorbix (vortex burning and mixing) combustor and represented a moderate increase in hardware complexity and difficulty of development. The concept selected for Element III employed vaporized fuel as a means of achieving minimum emission levels and represented the greatest difficulty of development and adaptation to the JT8D engine. Test results indicate that the Element I single-stage combustors were capable of dramatic improvement in idle pollutants. The multistage combustors evaluated in Program Elements II and III simultaneously reduced CO, THC and NOx emissions, but were unable to satisfy the current 1979 EPA standards.

Roberts, R.↗
NASA NTRS1977

Experimental clean combustor program, phase 3

A two-stage vortex burning and mixing combustor and associated fuel system components were successfully tested at steady state and transient operating conditions. The combustor exceeded the program goals for all three emissions species, with oxides of nitrogen 10 percent below the goal, carbon monoxide 26 percent below the goal, and total unburned hydrocarbons 75 percent below the goal. Relative to the JT9D-7 combustor, the oxides of nitrogen were reduced by 58 percent, carbon monoxide emissions were reduced by 69 percent, and total unburned hydrocarbons were reduced by 9 percent. The combustor efficiency and exit temperature profiles were comparable to those of production combustor. Acceleration and starting characteristics were deficient relative to the production engine.

Roberts, R.↗
NASA NTRS1977

NASA/Pratt and Whitney experimental clean combustor program: Engine test results

A two-stage vorbix (vortex burning and mixing) combustor and associated fuel system components were successfully tested in an experimental JT9D engine at steady-state and transient operating conditions, using ASTM Jet-A fuel. Full-scale JT9D experimental engine tests were conducted in a phase three aircraft experimental clean combustor program. The low-pollution combustor, fuel system, and fuel control concepts were derived from phase one and phase two programs in which several combustor concepts were evaluated, refined, and optimized in a component test rig. Significant pollution reductions were achieved with the combustor which meets the performance, operating, and installation requirements of the engine.

Roberts, R.↗
NASA NTRS1979

Experimental Clean Combustor Program (ECCP), phase 3

A double annular advanced technology combustor with low pollutant emission levels was evaluated in a series of CF6-50 engine tests. Engine lightoff was readily obtained and no difficulties were encountered with combustor staging. Engine acceleration and deceleration were smooth, responsive and essentially the same as those obtainable with the CF6-50 combustor. The emission reductions obtained in carbon monoxide, hydrocarbons, and nitrogen oxide levels were 55, 95, and 30 percent, respectively, at an idle power setting of 3.3 percent of takeoff power on an EPA parameter basis. Acceptable smoke levels were also obtained. The exit temperature distribution of the combustor was found to be its major performance deficiency. In all other important combustion system performance aspects, the combustor was found to be generally satisfactory.

Gleason, C. C.↗
NASA NTRS1979

Analytical evaluation of the impact of broad specification fuels on high bypass turbofan engine combustors

Six conceptual combustor designs for the CF6-50 high bypass turbofan engine and six conceptual combustor designs for the NASA/GE E3 high bypass turbofan engine were analyzed to provide an assessment of the major problems anticipated in using broad specification fuels in these aircraft engine combustion systems. Each of the conceptual combustor designs, which are representative of both state-of-the-art and advanced state-of-the-art combustion systems, was analyzed to estimate combustor performance, durability, and pollutant emissions when using commercial Jet A aviation fuel and when using experimental referee board specification fuel. Results indicate that lean burning, low emissions double annular combustor concepts can accommodate a wide range of fuel properties without a serious deterioration of performance or durability. However, rich burning, single annular concepts would be less tolerant to a relaxation of fuel properties. As the fuel specifications are relaxed, autoignition delay time becomes much smaller which presents a serious design and development problem for premixing-prevaporizing combustion system concepts.

Taylor, J. R.↗
NASA NTRS1980

Study of research and development requirements of small gas-turbine combustors

A survey is presented of the major small-engine manufacturers and governmental users. A consensus was undertaken regarding small-combustor requirements. The results presented are based on an evaluation of the information obtained in the course of the study. The current status of small-combustor technology is reviewed. The principal problems lie in liner cooling, fuel injection, part-power performance, and ignition. Projections of future engine requirements and their effect on the combustor are discussed. The major changes anticipated are significant increases in operating pressure and temperature levels and greater capability of using heavier alternative fuels. All aspects of combustor design are affected, but the principal impact is on liner durability. An R&D plan which addresses the critical combustor needs is described. The plan consists of 15 recommended programs for achieving necessary advances in the areas of liner thermal design, primary-zone performance, fuel injection, dilution, analytical modeling, and alternative-fuel utilization.

Demetri, E. P.↗
NASA NTRS1980

Experimental evaluation of combustor concepts for burning broad property fuels

A baseline CF6-50 combustor and three advanced combustor designs were evaluated to determine the effects of combustor design on operational characteristics using broad property fuels. Three fuels were used in each test: Jet A, a broad property 13% hydrogen fuel, and a 12% hydrogen fuel blend. Testing was performed in a sector rig at true cruise and simulated takeoff conditions for the CF6-50 engine cycle. The advanced combustors (all double annular, lean dome designs) generally exhibited lower metal temperatures, exhaust emissions, and carbon buildup than the baseline CF6-50 combustor. The sensitivities of emissions and metal temperatures to fuel hydrogen content were also generally lower for the advanced designs. The most promising advanced design used premixing tubes in the main stage. This design was chosen for additional testing in which fuel/air ratio, reference velocity, and fuel flow split were varied.

Kasper, J. M.↗
NASA NTRS1982

Techniques for enhancing durability and equivalence ratio control in a rich-lean, three-stage ground power gas turbine combustor

Rig tests of a can-type combustor were performed to demonstrate two advanced ground power engine combustor concepts: steam cooled rich-burn combustor primary zones for enhanced durability; and variable combustor geometry for three stage combustion equivalence ratio control. Both concepts proved to be highly successful in achieving their desired objectives. The steam cooling reduced peak liner temperatures to less than 800 K. This offers the potential of both long life and reduced use of strategic materials for liner fabrication. Three degrees of variable geometry were successfully implemented to control airflow distribution within the combustor. One was a variable blade angle axial flow air swirler to control primary airflow while the other two consisted of rotating bands to control secondary and tertiary or dilution air flow.

Schultz, D. F.↗
NASA NTRS1982

Computations of soot and and NO sub x emissions from gas turbine combustors

An analytical program was conducted to compute the soot and NOx emissions from a combustor and the radiation heat transfer to the combustor walls. The program involved the formulation of an emission and radiation model and the incorporation of this model into the Garrett 3-D Combustor Perfomance Computer Program. Computations were performed for the idle, cruise, and take-off conditions of a JT8D can combustor. The predicted soot and NOx emissions and the radiation heat transfer to the combustor walls agree reasonably well with the limited experimental data available.

Srivatsa, S. K.↗
NASA NTRS1982

Flow aerodynamics modeling of an MHD swirl combustor - Calculations and experimental verification

This paper describes a computer code for calculating the flow dynamics of constant density flow in the second stage trumpet shaped nozzle section of a two stage MHD swirl combustor for application to a disk generator. The primitive pressure-velocity variable, finite difference computer code has been developed to allow the computation of inert nonreacting turbulent swirling flows in an axisymmetric MHD model swirl combustor. The method and program involve a staggered grid system for axial and radial velocities, and a line relaxation technique for efficient solution of the equations. Turbulence simulation is by way of a two-equation Kappa-epsilon model. The code produces as output the flowfield map of the nondimensional stream function, axial, and swirl velocity. Good agreement was obtained between the theoretical predictions and the qualitative experimental results. The best seed injector location for uniform seed distribution at combustor exit is with injector located centrally on the combustor axis at entrance to the second stage combustor.

Gupta, A. K.↗
NASA NTRS1981

Energy efficient engine sector combustor rig test program

Under the NASA-sponsored Energy Efficient Engine program, Pratt & Whitney Aircraft has successfully completed a comprehensive combustor rig test using a 90-degree sector of an advanced two-stage combustor with a segmented liner. Initial testing utilized a combustor with a conventional louvered liner and demonstrated that the Energy Efficient Engine two-stage combustor configuration is a viable system for controlling exhaust emissions, with the capability to meet all aerothermal performance goals. Goals for both carbon monoxide and unburned hydrocarbons were surpassed and the goal for oxides of nitrogen was closely approached. In another series of tests, an advanced segmented liner configuration with a unique counter-parallel FINWALL cooling system was evaluated at engine sea level takeoff pressure and temperature levels. These tests verified the structural integrity of this liner design. Overall, the results from the program have provided a high level of confidence to proceed with the scheduled Combustor Component Rig Test Program.

Dubiel, D. J.↗
NASA NTRS1983

Testing of felt-ceramic materials for combustor applications

The feasibility of using composite felt ceramic materials as combustor liners was experimentally studied. The material consists of a porous felt pad sandwiched between a layer of ceramic and one of solid metal. Flat, rectangular test panels, which encompassed several design variations of the basic composite material, were tested, two at a time, in a premixed gas turbine combustor as sections of the combustor wall. Tests were conducted at combustor inlet conditions of 0.5 MPa and 533 K with a reference velocity of 25 m/s. The panels were subjected to a hot gas temperature of 2170 K with 1% of the total airflow used to film cool the ceramic surface of the test panel. In general, thin ceramic layers yield low ceramic stress levels with high felt ceramic interface temperatures. On the other hand, thick ceramic layers result in low felt ceramic interface temperatures but high ceramic stress levels. Extensive thermal cycling appears to cause material degradation, but for a limited number of cycles, the survivability of felt ceramic materials, even under extremely severe combustor operating conditions, was conclusively demonstrated.

Venkat, R. S.↗
NASA NTRS1983

Small Gas Turbine Combustor Primary Zone Study

The combustion research program, small gas turbine combustor primary zone study is summarized. The basic elements of a design methodology program to obtain the maximum performance potential of small reverse-flow annular combustors is described. Three preferred combustion design approaches for internal flame stabilization patterns were selected. Design features are incorporated in the combustors to address the performance limiting problem areas associated with smaller annular combustors. Performance is predicted by using a 3-D aerodynamic/chemical kinetic elliptic flow analysis, initially developed by Garrett Corporation for the USARTL. It is shown that the analytical flow field predictive models provide a useful design tool for understanding the combustion performance of a small reverse flow annular combustor.

Sullivan, R. E.↗