SEARCH · Search NASA
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.
A mixer assembly for a gas turbine engine is provided, including a main mixer with fuel injection holes located between at least one radial swirler and at least one axial swirler, wherein the fuel injected into the main mixer is atomized and dispersed by the air flowing through the radial swirler and the axial swirler.
A mixer assembly for a gas turbine engine is provided, including a main mixer, and a pilot mixer having an annular housing. The pilot mixer can further include features to cool portions of the annular housing, such as a radial edge of the annular housing.
A seal system for a gas turbine vane includes a first seal layer and an optional second seal layer. Each first seal layer includes multiple seal segments, each of which includes a first leg, a second leg, and an intermediate portion between the first leg and the second leg. Each adjacent pair of seal segments of the first seal is separated by a gap. The seal segments define a substantially complete perimeter of a seal slot in which the first seal is installed. The second seal, which is also segmented, may or may not define the substantially complete perimeter of the seal slot. If present, the second seal segments are circumferentially offset from the first seal segments to block the gaps between the first seal segments. A turbine vane including the present seal system is also disclosed.
A new era is drawing in the ability to predict convection heat transfer in the turbine gas path. We feel that the technical community now has the capability to mount a major assault on this problem, which has eluded significant progress for a long time. We hope to make a case for this bold statement by reviewing the state of the art in three major heat transfer, configuration-specific experiments, whose data have provided the big picture and guided both the fundamental modeling research and the code development. Following that, we review progress and directions in the development of computer codes to predict turbine gas path heat transfer. Finally, we cite examples and make observations on the more recent efforts to do all this work in a simultaneous, interactive, and more synergistic manner. We conclude with an assessment of progress, suggestions for how to use the current state of the art, and recommendations for the future.
A fuel-air mixer for use in a combustion chamber of a gas turbine engine is provided. The fuel air mixing apparatus comprises an annular fuel injector having a plurality of discrete plain jet orifices, a first swirler wherein the first swirler is located upstream from the fuel injector and a second swirler wherein the second swirler is located downstream from the fuel injector. The plurality of discrete plain jet orifices are situated between the highly swirling airstreams generated by the two radial swirlers. The distributed injection of the fuel between two highly swirling airstreams results in rapid and effective mixing to the desired fuel-air ratio and prevents the formation of local hot spots in the combustor primary zone. A combustor and a gas turbine engine comprising the fuel-air mixer of the present invention are also provided as well as a method using the fuel-air mixer of the present invention.
This presentation provides a high-level review of the potential failure modes and hazards to which electrified aircraft propulsion (EAP) systems are susceptible, along with potential gas turbine control-based strategies to assist in the mitigation of those failures. To further illustrate the role of gas turbine controls in mitigating EAP failure modes, an example based on a simulated EAP concept aircraft proposed by NASA is given. The effects of failures are discussed, along with turbomachinery control strategies, including reversionary control modes, and control limit logic.
Estimates of the heat transfer from the gas to stationary (vanes) or rotating blades poses a major uncertainty due to the complexity of the heat transfer processes. The gas flow through these blade rows is three dimensional with complex secondary viscous flow patterns that interact with the endwalls and blade surfaces. In addition, upstream disturbances, stagnation flow, curvature effects, and flow acceleration complicate the thermal transport mechanisms in the boundary layers. Some of these fundamental heat transfer effects are discussed. The chief purpose of the discussion is to acquaint those in the heat transfer community, not directly involved in gas turbines, of the seriousness of the problem and to recommend some basic research that would improve the capability for predicting gas-side heat transfer on turbine blades and vanes.
In response to the National Research Council (NRC) recommendations, the Workshop on Aerosols and Particulates from Aircraft Gas Turbine Engines was organized by the NASA Lewis Research Center and held on July 29-30, 1997 at the Ohio Aerospace Institute in Cleveland, Ohio. The objective is to develop consensus among experts in the field of aerosols from gas turbine combustors and engines as to important issues and venues to be considered. Workshop participants' expertise included engine and aircraft design, combustion processes and kinetics, atmospheric science, fuels, and flight operations and instrumentation.
Programs initiated by NASA to develop and demonstrate low emission advanced technology combustors for reducing aircraft gas turbine engine pollution are reviewed. Program goals are consistent with urban emission level requirements as specified by the U. S. Environmental Protection Agency and with upper atmosphere cruise emission levels as recommended by the U. S. Climatic Impact Assessment Program and National Research Council. Preliminary tests of advanced technology combustors indicate that significant reductions in all major pollutant emissions should be attainable in present generation aircraft gas turbine engines without adverse effects on fuel consumption. Preliminary test results from fundamental studies indicate that extremely low emission combustion systems may be possible for future generation jet aircraft. The emission reduction techniques currently being evaluated in these programs are described along with the results and a qualitative assessment of development difficulty.
Technical work on the design and effort leading to the testing of a 74.5 kW (100 hp) automotive gas turbine engine is reviewed. Development of the engine compressor, gasifier turbine, power turbine, combustor, regenerator, and secondary system is discussed. Ceramic materials development and the application of such materials in the gas turbine engine components is described.
In aircraft propulsion as well as stationary power generation, gas turbine engines remain a key energy conversion technology due to their high thermal efficiencies and low emissions. However, as emission requirements become increasingly stringent, engine manufacturers have sought to design combustion systems that operate near the fuel-lean limit of flammability. In this study, superadiabatic matrix-stabilized combustion, also known as porous media combustion, is evaluated as an advanced combustion concept for extending the lean flammability limit to achieve improved efficiency and emissions. To this end, a Brayton cycle analysis is developed and key parameters of the porous matrix are identified for maximizing the extension of the lean flammability limit. It is shown that stabilization of combustion below the nominal lean flammability limit allows for the design of engines with significantly higher pressure ratios and lower dilution ratios without increasing turbine inlet temperatures, thus improving cycle thermal efficiency. Combustor flammability limits were shown to be extendable by up to 32% when employing matrix-stabilized combustion, resulting in thermal efficiency gains of up to 11% compared to a nominal design.
Overview: Solid state energy harvesting using waste heat available in gas turbine engine offers potential for power generation to meet growing power needs of aircraft; Thermoelectric material advances offer new opportunities; Weight-optimized integrated turbine engine structure incorporating energy conversion devices.
The paper describes and discusses the results from some of the research and development programs for reducing aircraft gas turbine engine emissions. Although the paper concentrates on NASA programs only, work supported by other U.S. government agencies and industry has provided considerable data on low emission advanced technology for aircraft gas turbine engine combustors. The results from the two major NASA technology development programs, the ECCP (Experimental Clean Combustor Program) and the PRTP (Pollution Reduction Technology Program), are presented and compared with the requirements of the 1979 U.S. EPA standards. Emission reduction techniques currently being evaluated in these programs are described along with the results and a qualitative assessment of development difficulty.
The use of ammonia as a fuel source in gas turbine engine power cycles represents an attractive means to decarbonize the energy sector due to higher energy density and achieving liquid state at far lower pressures compared to pure hydrogen. However, due to low flammability and a propensity for high NOx emissions, its use is not without challenge. Here, a number of 0D and 1D modeling tools were utilized to study the combustion characteristics of ammonia and ammonia/hydrogen mixtures, examining basic fundamental properties such as laminar flame speed, variability among existing chemical kinetic mechanisms, and considering the use of two-stage rich-lean combustion strategies to achieve low NOx emissions.
This report is the final in a series of Technical Summary Reports for the Advanced Gas Turbine (AGT) Technology Development Project, authorizrd under NASA Contract DEN3-167 and sponsored by the DOE. The project was administered by NASA-Lewis Research Center of Cleveland, Ohio. Plans and progress are summarized for the period October 1979 through June 1987. This program aims to provide the US automotive industry the high risk, long range technology necessary to produce gas turbine engines for automobiles that will reduce fuel consumption and reduce environmental impact. The intent is that this technology will reach the marketplace by the 1990s. The Garrett/Ford automotive AGT was designated AGT101. The AGT101 is a 74.5 kW (100 shp) engine, capable of speeds to 100,000 rpm, and operates at turbine inlet temperatures to 1370 C (2500 F) with a specific fuel consumption level of 0.18 kg/kW-hr (0.3 lbs/hp-hr) over most of the operating range. This final report summarizes the powertrain design, power section development and component/ceramic technology development.
The potential benefits of solar/fossil hybrid gas turbine power systems were assessed. Both retrofit and new systems were considered from the aspects of; cost of electricity, fuel conservation, operational mode, technology requirements, and fuels flexibility. Hybrid retrofit (repowering) of existing combustion (simple Brayton cycle) turbines can provide near-term fuel savings and solar experience, while new and advanced recuperated or combined cycle systems may be an attractive fuel saving and economically competitive vehicle to transition from today's gas and oil-fired powerplants to other more abundant fuels.
The potential benefits of solar/fossil hybrid gas turbine power systems were assessed. Both retrofit and new systems were considered from the aspects of cost of electricity, fuel conservation, operational mode, technology requirements, and fuels flexibility. Hybrid retrofit (repowering) of existing combustion (simple Brayton cycle) turbines can provide near-term fuel savings and solar experience, while new and advanced recuperated or combined cycle systems may be an attractive fuel saving and economically competitive vehicle to transition from today's gas and oil-fired powerplants to other more abundant fuels.
The present treatise reports on theoretical investigations and test-stand measurements which were carried out in the BMW Flugmotoren GMbH in developing the hollow blade for exhaust gas turbines. As an introduction the temperature variation and the stress on a turbine blade for a gas temperature of 900 degrees and circumferential velocities of 600 meters per second are discussed. The assumptions onthe heat transfer coefficients at the blade profile are supported by tests on an electrically heated blade model. The temperature distribution in the cross section of a blade Is thoroughly investigated and the temperature field determined for a special case. A method for calculation of the thermal stresses in turbine blades for a given temperature distribution is indicated. The effect of the heat radiation on the blade temperature also is dealt with. Test-stand experiments on turbine blades are evaluated, particularly with respect to temperature distribution in the cross section; maximum and minimum temperature in the cross section are ascertained. Finally, the application of the hollow blade for a stationary gas turbine is investigated. Starting from a setup for 550 C gas temperature the improvement of the thermal efficiency and the fuel consumption are considered as well as the increase of the useful power by use of high temperatures. The power required for blade cooling is taken into account.