Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “gas turbine combustors”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

The effect of incomplete fuel-air mixing on the lean blowout limit, lean stability limit and NO(x) emissions in lean premixed gas turbine combustors

Gas turbine engines for both land-based and aircraft propulsion applications are facing regulations on NOx emissions which cannot be met with current combustor technology. A number of alternative combustor strategies are being investigated which have the potential capability of achieving ultra-low NOx emissions, including lean premixed combustors, direct injection combustors, rich burn-quick quench-lean burn combustors and catalytic combustors. The research reported in this paper addresses the effect of incomplete fuel-air mixing on the lean limit performance and the NOx emissions characteristics of lean premixed combustors.

Shih, W.-P.↗
NASA NTRS1990

Application of mixing-controlled combustion models to gas turbine combustors

Gas emissions were studied from a staged Rich Burn/Quick-Quench Mix/Lean Burn combustor were studied under test conditions encountered in High Speed Research engines. The combustor was modeled at conditions corresponding to different engine power settings, and the effect of primary dilution airflow split on emissions, flow field, flame size and shape, and combustion intensity, as well as mixing, was investigated. A mathematical model was developed from a two-equation model of turbulence, a quasi-global kinetics mechanism for the oxidation of propane, and the Zeldovich mechanism for nitric oxide formation. A mixing-controlled combustion model was used to account for turbulent mixing effects on the chemical reaction rate. This model assumes that the chemical reaction rate is much faster than the turbulent mixing rate.

Nguyen, Hung Lee↗
NASA NTRS2011

Gas turbine combustor

A gas turbine engine has a combustor module including an annular combustor having a liner assembly that defines an annular combustion chamber having a length, L. The liner assembly includes a radially inner liner, a radially outer liner that circumscribes the inner liner, and a bulkhead, having a height, H1, which extends between the respective forward ends of the inner liner and the outer liner. The combustor has an exit height, H3, at the respective aft ends of the inner liner and the outer liner interior. The annular combustor has a ratio H1/H3 having a value less than or equal to 1.7. The annular combustor may also have a ration L/H3 having a value less than or equal to 6.0.

Burd, Steven W.↗
NASA NTRS1973

Effect of increased fuel temperature on emissions of oxides of nitrogen from a gas turbine combustor burning natural gas

An annular gas turbine combustor was tested with heated natural gas fuel to determine the effect of increasing fuel temperature on the formation of oxides of nitrogen. Fuel temperatures ranged from ambient to 800 K (980 F). Combustor pressure was 6 atmospheres and the inlet air temperature ranged from 589 to 894 K (600 to 1150 F). The NOx emission index increased with fuel temperature at a rate of 4 to 9 percent per 100 K (180 F), depending on the inlet air temperature. The rate of increase in NOx was lowest at the highest inlet air temperature tested.

Marchionna, N. R.↗
NASA NTRS1979

The effect of fuel sprays on emissions from a gas turbine combustor

A research gas turbine combustor which allows for independent control of drop size, fuel/air ratio, air inlet temperature, pressure, reference velocity, and residence time has been used to study the influence of fuel spray characteristics on primary zone emissions. Local concentrations of oxides of nitrogen, carbon monoxide and unburned hydrocarbons at various locations in the flow, local gas temperature, gas velocity, and the turbulence intensity (measured with laser velocimetry) are obtained. A comprehensive data analysis program to compute the local fuel/air ratio and the sum of the mole fractions of the species present confirms the validity of the emissions data. Increasing drop size is correlated with an increase in unburned hydrocarbons, while increasing residence time is associated with sharp decreases in hydrocarbons and carbon monoxide and an increase in oxides of nitrogen.

Nicholls, J. A.↗
NASA NTRS1999

Comparison of Techniques for Non-Intrusive Fuel Drop Size Measurements in a Subscale Gas Turbine Combustor

In aviation gas turbine combustors, many factors, such as the degree and extent of fuel/air mixing and fuel vaporization achieved prior to combustion, influence the formation of undesirable pollutants. To assist in analyzing the extent of fuel/air mixing, flow visualization techniques have been used to interrogate the fuel distributions during subcomponent tests of lean-burning fuel injectors. Discrimination between liquid and vapor phases of the fuel was determined by comparing planar laser-induced fluorescence (PLIF) images, elastically-scattered light images, and phase/Doppler interferometer measurements. Estimates of Sauter mean diameters are made by ratioing PLIF and Mie scattered intensities for various sprays, and factors affecting the accuracy of these estimates are discussed. Mie calculations of absorption coefficients indicate that the fluorescence intensities of individual droplets are proportional to their surface areas, instead of their volumes, due to the high absorbance of the liquid fuel for the selected excitation wavelengths.

Zaller, Michelle↗
NASA NTRS1977

Coherence between internal and external noise generated by gas turbine combustors

Experiments and analysis on a gas turbine combustor unit are reported with a view in mind to separate propagated acoustic power from non-propagating 'pseudo-sound'. Analytically, it is suggested that a transition frequency will exist below which the interior pressure fluctuations are non-propagating, whereas above this frequency, of the order of 100 Hz, the noise is dominated by propagating acoustic waves. Coherence measurements are reported which show this concept to be borne out experimentally. Coherence between interior and exterior microphones is measured over a wide range of experimental conditions for a gas turbine combustor exhausting directly to the atmosphere. The purpose is to show that below a certain frequency, measurements of interior noise are not indicative of combustion noise ultimately propagating from an engine.

Strahle, W. C.↗
NASA NTRS1973

Effect of water injection on nitric oxide emissions of a gas turbine combustor burning natural gas fuel

The effect of direct water injection on the exhaust gas emissions of a turbojet combustor burning natural gas fuel was investigated. The results are compared with the results from similar tests using ASTM Jet-A fuel. Increasing water injection decreased the emissions of oxides of nitrogen (NOX) and increased the emissions of carbon monoxide and unburned hydrocarbons. The greatest percentage decrease in NOX with increasing water injection was at the lowest inlet-air temperature tested. The effect of increasing inlet-air temperature was to decrease the effect of the water injection. The reduction in NOX due to water injection was almost identical to the results obtained with Jet-A fuel. However, the emission indices of unburned hydrocarbons, carbon monoxide, and percentage nitric oxide in NOX were not.

Marchionna, N. R.↗
NASA NTRS1995

Turbulence models for gas turbine combustors

This viewgraph presentation covers gas turbine combustor flow physics, turbulence model investigations, turbulent combustion modeling, and present status and future needs.

Brankovic, Andreja↗
NASA NTRS1996

Multi-Dimensional Measurements of Combustion Species in Flame Tube and Sector Gas Turbine Combustors

The higher temperature and pressure cycles of future aviation gas turbine combustors challenge designers to produce combustors that minimize their environmental impact while maintaining high operation efficiency. The development of low emissions combustors includes the reduction of unburned hydrocarbons, smoke, and particulates, as well as the reduction of oxides of nitrogen (NO(x)). In order to better understand and control the mechanisms that produce emissions, tools are needed to aid the development of combustor hardware. Current methods of measuring species within gas turbine combustors use extractive sampling of combustion gases to determine major species concentrations and to infer the bulk flame temperature. These methods cannot be used to measure unstable combustion products and have poor spatial and temporal resolution. The intrusive nature of gas sampling may also disturb the flow structure within a combustor. Planar laser-induced fluorescence (PLIF) is an optical technique for the measurement of combustion species. In addition to its non-intrusive nature, PLIF offers these advantages over gas sampling: high spatial resolution, high temporal resolution, the ability to measure unstable species, and the potential to measure combustion temperature. This thesis considers PLIF for in-situ visualization of combustion species as a tool for the design and evaluation of gas turbine combustor subcomponents. This work constitutes the first application of PLIF to the severe environment found in liquid-fueled, aviation gas turbine combustors. Technical and applied challenges are discussed. PLIF of OH was used to observe the flame structure within the post flame zone of a flame tube combustor, and within the flame zone of a sector combustor, for a variety of fuel injector configurations. OH was selected for measurement because it is a major combustion intermediate, playing a key role in the chemistry of combustion, and because its presence within the flame zone can serve as a qualitative marker of flame temperature. All images were taken in the environment of actual engines during flight, using actual jet fuel. The results of the PLIF study led directly to the modification of a fuel injector.

Hicks, Yolanda Royce↗
NASA NTRS1999

Real-Time Optical Fuel-to-Air Ratio Sensor for Gas Turbine Combustors

The measurement of the temporal distribution of fuel in gas turbine combustors is important in considering pollution, combustion efficiency and combustor dynamics and acoustics. Much of the previous work in measuring fuel distributions in gas turbine combustors has focused on the spatial aspect of the distribution. The temporal aspect however, has often been overlooked, even though it is just as important. In part, this is due to the challenges of applying real-time diagnostic techniques in a high pressure and high temperature environment. A simple and low-cost instrument that non-intrusively measures the real-time fuel-to-air ratio (FAR) in a gas turbine combustor has been developed. The device uses a dual wavelength laser absorption technique to measure the concentration of most hydrocarbon fuels such as jet fuel, methane, propane, etc. The device can be configured to use fiber optics to measure the local FAR inside a high pressure test rig without the need for windows. Alternatively, the device can readily be used in test rigs that have existing windows without modifications. An initial application of this instrument was to obtain time-resolved measurements of the FAR in the premixer of a lean premixed prevaporized (LPP) combustor at inlet air pressures and temperatures as high as 17 atm at 800 K, with liquid JP-8 as the fuel. Results will be presented that quantitatively show the transient nature of the local FAR inside a LPP gas turbine combustor at actual operating conditions. The high speed (kHz) time resolution of this device, combined with a rugged fiber optic delivery system, should enable the realization of a flight capable active-feedback and control system for the abatement of noise and pollutant emissions in the future. Other applications that require an in-situ and time-resolved measurement of fuel vapor concentrations should also find this device to be of use.

Nguyen, Quang-Viet↗
NASA NTRS1984

Ceramic composite liner material for gas turbine combustors

Advanced commercial and military gas turbine engines may operate at combustor outlet temperatures in excess of 1920 K (3000 F). At these temperatures combustors liners experience extreme convective and radiative heat fluxes. The ability of a plasma sprayed ceramic coating to reduce liner metal temperature has been recognized. However, the brittleness of the ceramic layer and the difference in thermal expansion with the metal substrate has caused cracking, spalling and some separation of the ceramic coating. Research directed at turbine tip seals (or shrouds) has shown the advantage of applying the ceramic to a compliant metal pad. This paper discusses recent studies of applying ceramics to combustor liners in which yttria stabilized zirconia plasma sprayed on compliant metal substrates which were exposed to near stoichiometric combustion, presents performance and durability results, and describes a conceptual design for an advanced, small gas turbine combustor. Test specimens were convectively cooled or convective-transpiration cooled and were evaluated in a 10 cm square flame tube combustor at inlet air temperatures of 533 K (500 F) and at a pressure of 0.5 MPa (75 psia). The ceramics were exposed to flame temperatures in excess of 2000 K (3320 F). Results appear very promising with all 30 specimens surviving a screening test and one of two specimens surviving a cyclic durability test.

Ercegovic, D. B.↗
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 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 model 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 formulation 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 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

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 NTRS1977

The effect of ambient conditions on carbon monoxide emissions from an idling gas turbine combustor

A test program employing a gas turbine combustor is outlined; the results of which quantize the effects of changes in ambient temperature and humidity on carbon monoxide emissions at simulated idle operating conditions. A comparison of the experimental results with analytical results generated by a kinetic model of the combustion process, and reflecting changing ambient conditions, is given. It is demonstrated that for a complete range of possible ambient variations, significant changes do occur in the amount of carbon monoxide emitted by a gas turbine at idle, and that the analytical model is reasonably successful in predicting changes.

Subramanian, A. K.↗
NASA NTRS1975

Predicted exhaust emissions from a methanol and jet fueled gas turbine combustor

A computer model of a gas turbine combustor has been used to predict the kinetic combustion and pollutant formation processes for methanol and simulated jet fuel. Use of the kinetic reaction mechanisms has also allowed a study of ignition delay and flammability limit of these two fuels. The NOX emissions for methanol were predicted to be from 69 to 92% lower than those for jet fuel at the same equivalence ratio which is in agreement with experimentally observed results. The high heat of vaporization of methanol lowers both the combustor inlet mixture temperatures and the final combustion temperatures. The lower combustion temperatures lead to low NOX emissions while the lower inlet mixture temperatures increase methanol's ignition delay. This increase in ignition delay dictates the lean flammability limit of methanol to be 0.8, while jet fuel is shown to combust at 0.4.

Adelman, H. G.↗
NASA NTRS1984

Ceramic composite liner material for gas turbine combustors

The application of ceramics to gas turbine combustor liners to reduce liner metal temperature was studied in an experiment in which yttria-stabilized zirconia plasma was sprayed on compliant metal substrates exposed to near stoichiometric combustion. The strain isolation pad materials chosen were Hoskins Alloy 875 and BRUNSLLOY 534 Fiber Metal of 0.25 and 0.38 cm thicknesses and 35 and 45 percent density levels. Combustor screening tests of all specimens showed no evidence of deterioration or failure. Specimens exposed to flame temperatures in excess of 2100 K were convectively or convective-transpiration cooled and were evaluated in a 10 cm sq flame tube at inlet air temperature of 533 K and pressure of 0.5 MPa. The results suggest the superiority of a system composed of the Hoskins Alloy 875 compliant pad with 0.25 cm thickness and 35 percent density coupled with a NiCrAlY bond coat and a 8 percent Y2O3-ZrO2 ceramic top coat of 0.19 cm thickness.

Ercegovic, D. B.↗