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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1951

Coolant-Flow Calibrations of Three Simulated Porous Gas-Turbine Blades

An investigation was conducted at the NACA Lewis laboratory to determine whether simulated porous gas-turbine blades fabricated by the Eaton Manufacturing Company of Cleveland, Ohio would be satisfactory with respect to coolant flow for application in gas-turbine engines. These blades simulated porous turbine blades by forcing the cooling air onto the blade surface through a large number of chordwise openings or slits between laminations of sheet metal or wire. This type of surface has a finite number of openings, whereas a porous surface has an almost infinite number of smaller openings for the coolant flow. The investigation showed that a blade made of sheet-metal laminations stacked on a support member that passed up through the coolant passage was completely unsatisfactory because of extremely poor coolant flow distribution over the blade surface. The flow distribution for two wire-wound blades was more uniform, but the pressure drop between the coolant supply pressure and the local pressure on the outside of the blades was too low by a factor ranging from 3 to 3.5 for the required coolant flow rates. The pressure drop could be increased by forcing the wires closer together during blade fabrication.

Esger, Jack B.↗
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 NTRS1981

Review of NASA progress in thermal barrier coatings for stationary gas turbines

Ceramic thermal barrier coatings for industrial/utility gas turbines were investigated. In burner rig tests of a zirconia yttria/nickel chromium aluminum yttrium ZrO2-12w/0Y2O3/NiCrAlY coating system on air cooled superalloy specimens, ceramic coating life (spallation) was sensitive to Na and V concentration in the fuel. The locations of coating spallation correspond to areas where combustion products were predicted to condense. Three new thermal barrier coating systems were identified. These are based on calcium silicate, ZrO2-8w/0Y2O3, and a MgO-NiCrAlY cermet. The spall resistance can be increased by reducing the ceramic layer thickness from 0.038 to 0.013 cm and by the use of more oxidation/corrosion resistant bond coats.

Hodge, P. E.↗
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.↗
NASA NTRS1985

Advanced liner-cooling techniques for gas turbine combustors

Component research for advanced small gas turbine engines is currently underway at the NASA Lewis Research Center. As part of this program, a basic reverse-flow combustor geometry was being maintained while different advanced liner wall cooling techniques were investigated. Performance and liner cooling effectiveness of the experimental combustor configuration featuring counter-flow film-cooled panels is presented and compared with two previously reported combustors featuring: splash film-cooled liner walls; and transpiration cooled liner walls (Lamilloy).

Norgren, C. T.↗
NASA NTRS1985

Advanced liner-cooling techniques for gas turbine combustors

Component research for advanced small gas turbine engines is currently underway at the NASA Lewis Research Center. As part of this program, a basic reverse-flow combustor geometry was being maintained while different advanced liner wall cooling techniques were investigated. Performance and liner cooling effectiveness of the experimental combustor configuration featuring counter-flow film-cooled panels is presented and compared with two previously reported combustors featuring: splash film-cooled liner walls; and transpiration cooled liner walls (Lamilloy).

Norgren, C. T.↗
NASA NTRS1990

Low NO(x) potential of gas turbine engines

Nitrogen oxide levels of low gas turbine engine emissions are correlated in this paper. The predictions of NO(x) emissions are based on a review of the literature of previous low NO(x) combustor programs and analytical chemical kinetic calculations. Concepts included in the literature review consisted of Lean-Premixed-Prevaporized (LPP), Rich Burn/Quick Quench/Lean/Burn (RQL), and Direct Injection. The NO(x) emissions were found to be an exponential function of adiabatic combustion temperature over a wide range of inlet temperatures, pressures and (lean) fuel-air ratios. It is stated that the low NO(X) potential of LPP is offset by the operational disadvantages of its narrow stability limits and its susceptibility to autoignition/ flashback. The RQL and the Direct Injection concepts have the advantage of wider stability limits comparable to conventional combustors. Results were obtained primarily with gaseous fuels and it is noted that the challenge will be to produce the same low levels of NO(x) with liquid fuels.

Tacina, Robert R.↗
NASA NTRS1995

Object-oriented approach for gas turbine engine simulation

An object-oriented gas turbine engine simulation program was developed. This program is a prototype for a more complete, commercial grade engine performance program now being proposed as part of the Numerical Propulsion System Simulator (NPSS). This report discusses architectural issues of this complex software system and the lessons learned from developing the prototype code. The prototype code is a fully functional, general purpose engine simulation program, however, only the component models necessary to model a transient compressor test rig have been written. The production system will be capable of steady state and transient modeling of almost any turbine engine configuration. Chief among the architectural considerations for this code was the framework in which the various software modules will interact. These modules include the equation solver, simulation code, data model, event handler, and user interface. Also documented in this report is the component based design of the simulation module and the inter-component communication paradigm. Object class hierarchies for some of the code modules are given.

Curlett, Brian P.↗
NASA NTRS2015

Mixer Assembly for a Gas Turbine Engine

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.

Dai, Zhongtao↗
NASA NTRS2016

Gas Turbine Engine Having Fan Rotor Driven by Turbine Exhaust and with a Bypass

A gas turbine engine has a core engine incorporating a core engine turbine. A fan rotor is driven by a fan rotor turbine. The fan rotor turbine is in the path of gases downstream from the core engine turbine. A bypass door is moveable from a closed position at which the gases from the core engine turbine pass over the fan rotor turbine, and moveable to a bypass position at which the gases are directed away from the fan rotor turbine. An aircraft is also disclosed.

Suciu, Gabriel L.↗
NASA NTRS2018

Mixer Assembly for a Gas Turbine Engine

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.

Dai, Zhongtao↗
NASA NTRS1983

IGCC performance comparison for variations in gasifier type and gas turbine firing temperature

Performance estimates were made for a series of integrated coal gasification combined cycle (IGCC) power systems using three generic types of coal gasification subsystems. The objectives of this study were (1) to provide a self consistent comparison of IGCC systems using different types of gasifiers and different oxidants and (2) to use this framework of cases to evaluate the effect of a gas turbine firing temperature and cooling approach an overall system efficiency. The basic IGCC systems considered included both air and oxygen blown versions of a fluidized bed gasifier, represented by the Westinghouse design, and an entrained bed gasifier, represented by the Texaco design. Also considered were systems using an oxygen blown, fixed bed gasifier, represented by the British Gas Corporation (BGC) slagging gasifier. All of these gasifiers were integrated with a combined cycle using a gas turbine firing temperature of 1700 K (2600 F) and a compressor pressure ratio of 16:1. Steam turbine throttle conditions were chosen to be 16.6 MPa/811 K (2400 psia/1000 F) with a single reheat to 810 K (1000 F). Some of these cases were modified to allow the evaluation of the effect of gas turbine firing temperature. Turbine firing temperatures from state of the art 1365 K (2000 F) to an advanced technology 1920 K (3000 F) were analyzed. A turbine cooling technology that maintains metal temperatures below acceptable limits was assumed for each level of firing temperature. System performance comparisons were made using three advanced turbine cooling technologies for the 1920 K (3000 F) firing temperature. The results indicate that the IGCC using the BGC gasifier had the highest net system efficiency (42.1 percent) of the five gasification cases considered.

Stochl, R. J.↗
NASA NTRS2002

Combustion Sensors: Gas Turbine Applications

This report documents efforts to survey the current research directions in sensor technology for gas turbine systems. The work is driven by the current and future requirements on system performance and optimization. Accurate real time measurements of velocities, pressure, temperatures, and species concentrations will be required for objectives such as combustion instability attenuation, pollutant reduction, engine health management, exhaust profile control via active control, etc. Changing combustor conditions - engine aging, flow path slagging, or rapid maneuvering - will require adaptive responses; the effectiveness of such will be only as good as the dynamic information available for processing. All of these issues point toward the importance of continued sensor development. For adequate control of the combustion process, sensor data must include information about the above mentioned quantities along with equivalence ratios and radical concentrations, and also include both temporal and spatial velocity resolution. Ultimately these devices must transfer from the laboratory to field installations, and thus must become low weight and cost, reliable and maintainable. A primary conclusion from this study is that the optics-based sensor science will be the primary diagnostic in future gas turbine technologies.

Human, Mel↗
NASA NTRS1980

An automotive transmission for automotive gas turbine power plants

A joint government-industry program was initiated to investigate the two-shaft gas turbine concept as an alternative to present-day automotive powerplants. Both were examined, compared and evaluated on the basis of the federal automotive driving cycle in terms of specific fuel/power/speed characteristics of the engine and the efficiency and performance of the transmission. The results showed that an optimum match of vehicle, gas turbine engine, and conventional automatic transmission is capable of a significant improvement in fuel economy. This system offers many advantages that should lead to its wide acceptance in future vehicles.

Polak, J. C.↗
NASA NTRS2018

Pressure-Gain Combustion for Gas Turbines

Pressure Gain Combustion (PGC) is under investigation as a means to improve the thermal efficiency of gas turbines. PGC is a fundamentally unsteady combustion process which, through some means of confinement, raises the total pressure of the working fluid relative to the initial process state. When implemented in a combustor it yields a total pressure gain across the device instead of the typical loss seen in a conventional combustor. This pressure gain can significantly improve gas turbine performance. The tutorial will provide an introduction to the concept of PGC. It will begin with a process description in fundamental thermodynamic terms, and will quantify potential benefits. The majority of the presentation will describe the various approaches to implementation that are under investigation by the PGC.

Paxson, Dan↗
NASA NTRS1992

Progress towards understanding and predicting convection heat transfer in the turbine gas path

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.

Simoneau, Robert J.↗
NASA NTRS2006

Fuel-air mixing apparatus for reducing gas turbine combustor exhaust emissions

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.

Zupanc, Frank J.↗