Search NASA⌕ Search

SEARCH · Search NASA

Results for “TURBINES, GAS - PERFORMANCE”

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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1981

Overview: DOE/NASA automotive gas turbine and stirling projects

An overview on the progress of the automotive gas turbine and automotive Stirling engine technology projects is presented. The following items are reported: (1) formulation and execution of projects in accordance with the Auto Propulsion Research and Development Act of 1978; (2) substantive technology accomplishments; and (3) future path options of the programs.

Beremand, D. G.↗
NASA NTRS1981

Thermal barrier coatings for thermal insulation and corrosion resistance in industrial gas turbine engines

Four thermal barrier coatings were subjected to a 500-hour gas turbine engine test. The coatings were two yttria stabilized zirconias, calcium ortho silicate and calcium meta titanate. The calcium silicate coating exhibited significant spalling. Yttria stabilized zirconia and calcium titanate coatings showed little degradation except in blade leading edge areas. Post-test examination showed variations in the coating due to manual application techniques. Improved process control is required if engineering quality coatings are to be developed. The results indicate that some leading edge loss of the coating can be expected near the tip.

Vogan, J. W.↗
NASA NTRS1982

Combustor development for automotive gas turbines

The development of a combustion system for the AGT 100 automotive gas turbine engine is described. A maximum turbine inlet temperature of 1288 C is reached during the regenerative cycle, and air up to 1024 C is supplied to the combustor inlet. A premix/prevaporization ceramic combustor employing variable geometry to control burning zone temperature was developed and tested. Tests on both metal and ceramic combustors showed that emissions were a function of burner inlet temperature (BIT). At 999 C BIT, NO(x) emissions were two orders of magnitude below program goals, and at the same temperature but at a different variable geometry position, the CO was 30 times below program goal. Tests to evaluate the durability of the ceramic materials showed no failures during steady-state operation; however, some cracks developed in the dome during extended transient operation.

Ross, P. T.↗
NASA NTRS1986

Ceramic thermal barrier coatings for electric utility gas turbine engines

Research and development into thermal barrier coatings for electric utility gas turbine engines is reviewed critically. The type of coating systems developed for aircraft applications are found to be preferred for clear fuel electric utility applications. These coating systems consists of a layer of plasma sprayed zirconia-yttria ceramic over a layer of MCrAly bond coat. They are not recommended for use when molten salts are presented. Efforts to understand coating degradation in dirty environments and to develop corrosion resistant thermal barrier coatings are discussed.

Miller, R. A.↗
NASA NTRS1987

Soot loading in a generic gas turbine combustor

Variation in soot loading along the centerline of a generic gas turbine combustor was experimentally investigated. The 12.7-cm dia burner consisted of six sheet-metal louvers. Soot loading along the burner length was quantified by acquiring measurements first at the exit of the full-length combustor and then at upstream stations by sequential removal of liner louvers to shorten the burner length. Alteration of the flow field approaching removed louvers, maintaining a constant liner pressure drop. Burner exhaust flow was sampled at the burner centerline to determine soot mass concentration and smoke number. Characteristic particle size and number density, transmissivity of the exhaust flow, and local radiation from luminous soot particles in the exhaust flow were determined by optical techniques. Four test fuels were burned at three fuel-air ratios to determine fuel chemical property and flow temperature influences. Data were acquired at two combustor pressures. Particulate concentration data indicated a strong oxidation mechanism in the combustor secondary zone, though the oxidation was significantly affected by flow temperature. Soot production was directly related to fuel smoke point. Less soot production and lower secondary-zone oxidation rates were observed at reduced combustor pressure.

Eckerle, W. A.↗
NASA NTRS1991

Gas turbine engine control system

A control system and method of controlling a gas turbine engine. The control system receives an error signal and processes the error signal to form a primary fuel control signal. The control system also receives at least one anticipatory demand signal and processes the signal to form an anticipatory fuel control signal. The control system adjusts the value of the anticipatory fuel control signal based on the value of the error signal to form an adjusted anticipatory signal and then the adjusted anticipatory fuel control signal and the primary fuel control signal are combined to form a fuel command signal.

Idelchik, Michael S.↗
NASA NTRS2017

Method for Making Measurements of the Post-Combustion Residence Time in a Gas Turbine Engine

A method of measuring a residence time in a gas-turbine engine is disclosed that includes measuring a combustor pressure signal at a combustor entrance and a turbine exit pressure signal at a turbine exit. The method further includes computing a cross-spectrum function between the combustor pressure signal and the turbine exit pressure signal, calculating a slope of the cross-spectrum function, shifting the turbine exit pressure signal an amount corresponding to a time delay between the measurement of the combustor pressure signal and the turbine exit pressure signal, and recalculating the slope of the cross-spectrum function until the slope reaches zero.

Miles, Jeffrey H.↗
NASA NTRS1991

Optical fiber sensor for temperature measurement from 600 to 1900 C in gas turbine engines

A temperature sensor system has been fabricated specifically for the harsh environment encountered in temperature measurement on gas turbine engines. Four components comprised the system: a thermally emissive source, a high temperature lightguide, a flexible optical cable and an electro-optic signal processor. The emissive source was located inside a sapphire rod so that the sapphire serves as both a lightguide and as a protective shroud. As the probe was heated, the thermal radiation from the emissive source increased with increasing temperature. The flexible optical cable was constructed with 200 micron core fiber and ruggedized for turbine engine applications. The electro-optic signal processor used the ratio of intensity in two wavelength intervals to determine a digital value of the temperature. The probe tip was operated above 1900 C in a low velocity propane flame and above 1500 C at Mach .37. Probe housings, optical cables, and signal processors were constructed and environmentally tested for the temperature and vibration experienced by turbine engine sensors. This technology was used to build an optical exhaust gas sensor for a General Electric Aircraft Engines F404 turbine. The four optical probes and optical cable were a functional replacement for four thermocouple probes. The system was ground tested for 50 hours with an excess of 1000 thermal cycles. This optical temperature sensor system measured gas temperature up to the operational limit of the turbine engine.

Tregay, G. W.↗
NASA NTRS1947

Cooling of Gas Turbines: Effectiveness of Rim Cooling of Blades - 2

An analysis is presented of rim cooling of gas-turbine blades; that is, reducing the temperature at the base of the blade (wheel rim), which cools the blade by conduction alone. Formulas for temperature and stress distributions along the blade are derived and, by the use of experimental stress-rupture data for a typical blade alloy, a relation is established between blade life (time for rupture), operating speed, and amount of rim cooling for several gas temperatures. The effect of blade parameter combining the effects of blade dimensions, blade thermal conductivity, and heat-transfer coefficient is determined. The effect of radiation on the results is approximated. The gas temperatures ranged from 1300F to 1900F and the rim temperature, from 0F to 1000F below the gas temperature. This report is concerned only with blades of uniform cross section, but the conclusions drawn are generally applicable to most modern turbine blades. For a typical rim-cooled blade, gas temperature increases are limited to about 200F for 500F of cooling of the blade base below gas temperature, and additional cooling brings progressively smaller increases. In order to obtain large increases in thermal conductivity or very large decreases in heat-transfer coefficient or blade length or necessary. The increases in gas temperature allowable with rim cooling are particularly small for turbines of large dimensions and high specific mass flows. For a given effective gas temperature, substantial increases in blade life, however, are possible with relatively small amounts of rim cooling.

Wolfenstein, Lincoln↗
NASA NTRS1977

A method to estimate weight and dimensions of aircraft gas turbine engines. Volume 1: Method of analysis

Weight and envelope dimensions of aircraft gas turbine engines are estimated within plus or minus 5% to 10% using a computer method based on correlations of component weight and design features of 29 data base engines. Rotating components are estimated by a preliminary design procedure where blade geometry, operating conditions, material properties, shaft speed, hub-tip ratio, etc., are the primary independent variables used. The development and justification of the method selected, the various methods of analysis, the use of the program, and a description of the input/output data are discussed.

Pera, R. J.↗
NASA NTRS1978

Gas turbine engine with recirculating bleed

Carbon monoxide and unburned hydrocarbon emissions in a gas turbine engine are reduced by bleeding hot air from the engine cycle and introducing it back into the engine upstream of the bleed location and upstream of the combustor inlet. As this hot inlet air is recycled, the combustor inlet temperature rises rapidly at a constant engine thrust level. In most combustors, this will reduce carbon monoxide and unburned hydrocarbon emissions significantly. The preferred locations for hot air extraction are at the compressor discharge or from within the turbine, whereas the preferred reentry location is at the compressor inlet.

Adamson, A. P.↗
NASA NTRS1978

Liquid-cooling technology for gas turbines review and status

A review of research related to liquid cooling of gas turbines was conducted and an assessment of the state of the art was made. Various methods of liquid cooling turbines were reviewed. Examples and results with test and demonstrator turbines utilizing these methods along with the advantages and disadvantages of the various methods are discussed.

Vanfossen, G. J., Jr.↗
NASA NTRS1947

The Application of High-temperature Strain Gauges to the Measurements of Vibratory Stresses in Gas-turbine Buckets

The feasibility of measuring the vibration in the buckets of a gas turbine under service conditions of speed and temperature was determined by use of a high temperature wire strain gauge cemented to a modified supercharger turbine bucket. A high-temperature wire strain gauge and the auxiliary mechanical and electrical equipment developed for the investigation are described.

Kemp, R H↗