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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 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Research turbine for high temperature core engine application. 1: Cold-airoverall performance of solid scaled turbine

A solid, half-scale model of a 50.8-cm (20-in) research turbine designed for a high temperature core engine application was investigated over a range of speeds and pressure ratios. The results of this test are presented. The effect of rotor blade twist was also investigated. At the design equivalent speed and specific work output, the total efficiency of the turbine with untwisted rotor blades was 87.1 percent; at the same pressure ratio the efficiency of the turbine with twisted rotor blades was 88.0 percent.

Szanca, E. M.↗
NASA NTRS1975

Cold-air investigation of a 3 1/2-stage fan-drive turbine with a stage loading factor of 4 designed for an integral lift engine. 1: Turbine design and performance of first stage

The design of the 3 1/2-stage turbine is described, and the cold-air performance of the first stage, modified for axial inlet conditions, is presented. The performance of the modified single-stage turbine and of two comtemporary high-stage-loading-factor turbines is compared with that estimated with a reference prediction method.

Whitney, W. J.↗
NASA NTRS1977

Air cooling of disk of a solid integrally cast turbine rotor for an automotive gas turbine

A thermal analysis is made of surface cooling of a solid, integrally cast turbine rotor disk for an automotive gas turbine engine. Air purge and impingement cooling schemes are considered and compared with an uncooled reference case. Substantial reductions in blade temperature are predicted with each of the cooling schemes studied. It is shown that air cooling can result in a substantial gain in the stress-rupture life of the blade. Alternatively, increases in the turbine inlet temperature are possible.

Gladden, H. J.↗
NASA NTRS1978

An overview of aerospace gas turbine technology of relevance to the development of the automotive gas turbine engine

Technology areas related to gas turbine propulsion systems with potential for application to the automotive gas turbine engine are discussed. Areas included are: system steady-state and transient performance prediction techniques, compressor and turbine design and performance prediction programs and effects of geometry, combustor technology and advanced concepts, and ceramic coatings and materials technology.

Evans, D. G.↗
NASA NTRS1978

An overview of aerospace gas turbine technology of relevance to the development of the automotive gas turbine engine

The NASA-Lewis Research Center (LeRC) has conducted, and has sponsored with industry and universities, extensive research into many of the technology areas related to gas turbine propulsion systems. This aerospace-related technology has been developed at both the component and systems level, and may have significant potential for application to the automotive gas turbine engine. This paper summarizes this technology and lists the associated references. The technology areas are system steady-state and transient performance prediction techniques, compressor and turbine design and performance prediction programs and effects of geometry, combustor technology and advanced concepts, and ceramic coatings and materials technology.

Evans, D. G.↗
NASA NTRS1978

Cold-air performance of free power turbine designed for 112-kilowatt automotive gas-turbine engine 3: Effect of stator vane end clearances on performance

An experimental investigation of a free power turbine designed for a 112-kW, automotive, gas turbine engine was made to determine the penalty in performance due to the stator vane end clearances. Tests were made over a range of mean section stator vane angles from 26 deg to 50 deg (as measured from the plane of rotation) with the vane end clearances filled. These results were compared with test results of the same turbine with vane end clearances open. At design equivalent values of rotative speed and pressure ratio and at a vane angle of 35 deg, the mass flow with the vane and clearances filled was about 8 percent lower than mass flow with vane end clearances open. The decrease in mass flow was mitigated by increasing the vane angle. With the vane end clearances filled, there was about a 66 percent reduction in mass flow when the vane angle was decreased from 40 deg to 26 deg. For the same decrease in vane angle the stator throat area decreased by about 50 percent. This result indicates that the rotor losses were increasing with decreasing vane angle.

Kofskey, M. G.↗
NASA NTRS1980

Algorithm for calculating turbine cooling flow and the resulting decrease in turbine efficiency

An algorithm is presented for calculating both the quantity of compressor bleed flow required to cool the turbine and the decrease in turbine efficiency caused by the injection of cooling air into the gas stream. The algorithm, which is intended for an axial flow, air routine in a properly written thermodynamic cycle code. Ten different cooling configurations are available for each row of cooled airfoils in the turbine. Results from the algorithm are substantiated by comparison with flows predicted by major engine manufacturers for given bulk metal temperatures and given cooling configurations. A list of definitions for the terms in the subroutine is presented.

Gauntner, J. W.↗
NASA NTRS1982

Energy efficient engine: Turbine intermediate case and low-pressure turbine component test hardware detailed design report

A four stage, low pressure turbine component has been designed to power the fan and low pressure compressor system in the Energy Efficient Engine. Designs for a turbine intermediate case and an exit guide vane assembly also have been established. The components incorporate numerous technology features to enhance efficiency, durability, and performance retention. These designs reflect a positive step towards improving engine fuel efficiency on a component level. The aerodynamic and thermal/mechanical designs of the intermediate case and low pressure turbine components are presented and described. An overview of the predicted performance of the various component designs is given.

Leach, K.↗
NASA NTRS1956

Investigation of Turbines Suitable for Use in a Turbojet Engine with High Compressor Pressure Ratio and Low Compressor-tip Speed VI: Experimental Performance of Two-stage Turbine

The brake internal efficiency of the highly loaded two-stage turbine was 0.79 equivalent design work and speed. The maximum brake internal efficiency was 0.84. A radial survey revealed these major defects: (1) the first-rotor throat area was too large, and a large area of underturned flow existed near the tip;(2) considerable underturning existed at the second-stator outlet; and (3) tangential components of velocity at the turbine outlet amounted to 2.5 points in turbine efficiency.

Davison, Elmer H↗
NASA NTRS1947

Fundamentals of the Control of Gas-Turbine Power Plants for Aircraft: Standardization of the Computations Relating to the Control of Gas-Turbine Power Plants for Aircraft by the Employment of the Laws of Similarity - Part 1

It will be shown that by the use of the concept of similarity a simple representation of the characteristic curves of a compressor operating in combination with a turbine may be obtained with correct allowance for the effect of temperature. Furthermore, it bec~mes possible to simplify considerably the rather tedious investigations of the behavior of gas-turbine power plants under different operating conditions. Characteristic values will be derived for the most important elements of operating behavior of the power plant, which will be independent of the absolute valu:s of pressure and temperature. At the same time, the investigations provide the basis for scale-model tests on compressors and turbines.

Luehl, H.↗
NASA NTRS2000

Turbine Blade and Endwall Heat Transfer Measured in NASA Glenn's Transonic Turbine Blade Cascade

Higher operating temperatures increase the efficiency of aircraft gas turbine engines, but can also degrade internal components. High-pressure turbine blades just downstream of the combustor are particularly susceptible to overheating. Computational fluid dynamics (CFD) computer programs can predict the flow around the blades so that potential hot spots can be identified and appropriate cooling schemes can be designed. Various blade and cooling schemes can be examined computationally before any hardware is built, thus saving time and effort. Often though, the accuracy of these programs has been found to be inadequate for predicting heat transfer. Code and model developers need highly detailed aerodynamic and heat transfer data to validate and improve their analyses. The Transonic Turbine Blade Cascade was built at the NASA Glenn Research Center at Lewis Field to help satisfy the need for this type of data.

Giel, Paul W.↗
NASA NTRS2013

Aerodynamic Measurements of a Variable-Speed Power-Turbine Blade Section in a Transonic Turbine Cascade at Low Inlet Turbulence

Aerodynamic measurements obtained in a transonic linear cascade were used to assess the impact of large incidence angle and Reynolds number variations on the 3-D flow field and midspan loss and turning of a 2-D section of a variable-speed power-turbine (VSPT) rotor blade. Steady-state data were obtained for ten incidence angles ranging from +15.8 deg to −51.0 deg. At each angle, data were acquired at five flow conditions with the exit Reynolds number (based on axial chord) varying over an order-of-magnitude from 2.12×10(exp 5) to 2.12×10(exp 6). Data were obtained at the design exit Mach number of 0.72 and at a reduced exit Mach number of 0.35 as required to achieve the lowest Reynolds number. Midspan total-pressure and exit flow angle data were acquired using a five-hole pitch/yaw probe surveyed on a plane located 7.0 percent axial chord downstream of the blade trailing edge plane. The survey spanned three blade passages. Additionally, three-dimensional half-span flow fields were examined with additional probe survey data acquired at 26 span locations for two key incidence angles of +5.8 deg and −36.7 deg. Survey data near the endwall were acquired with a three-hole boundary-layer probe. The data were integrated to determine average exit total-pressure and flow angle as functions of incidence and flow conditions. The data set also includes blade static pressures measured on four spanwise planes and endwall static pressures. Tests were conducted in the NASA Glenn Transonic Turbine Blade Cascade Facility. The measurements reflect strong secondary flows associated with the high aerodynamic loading levels at large positive incidence angles and an increase in loss levels with decreasing Reynolds number. The secondary flows decrease with negative incidence as the blade becomes unloaded. Transitional flow is admitted in this low inlet turbulence dataset, making it a challenging CFD test case. The dataset will be used to advance understanding of the aerodynamic challenges associated with maintaining efficient power turbine operation over a wide shaft-speed range. deg

Flegel-McVetta, Ashlie B.↗
DOE OSTI2020

Ceramic matrix composite (CMC) turbine blades and methods of forming CMC turbine blades

A ceramic matrix composite (CMC) turbine blade includes a root region, a narrowed neck region extending from the root region, and a hub region extending from the narrowed neck region. At least one central CMC ply extending from the root region is interwoven with at least one insert extending toward the narrowed neck region from the hub region. A method of forming a CMC turbine blade includes interweaving at least one central CMC ply extending from the root region with at least one insert extending toward the narrowed neck region from the hub region. The CMC turbine blade includes a root region, a narrowed neck region extending from the root region, and a hub region extending from the narrowed neck region.

36 MATERIALS SCIENCE↗
DOE OSTI2022

Turbine component and methods of making and cooling a turbine component

A turbine component includes a root and an airfoil extending from the root to a tip opposite the root. The airfoil forms a leading edge and a trailing edge portion extending to a trailing edge. Radial cooling channels in the trailing edge portion of the airfoil permit radial flow of a cooling fluid through the trailing edge portion. Each radial cooling channel has a first end at a lower surface at a root edge of the trailing edge portion or at an upper surface at a tip edge of the trailing edge portion and a second end opposite the first end at the lower surface or the upper surface. A method of making a turbine component and a method of cooling a turbine component are also disclosed.

Dutta, Sandip↗
DOE OSTI2026

Wind Turbine Materials Recycling Prize Phase 2 (Commercialization of Wind Turbine Blade Waste (WTBW)-Based, Lightweight, Cementitious Composite Materials): Cooperative Research and Development Final Report, CRADA Number CRD-24-31305

The National Laboratory of the Rockies (NLR) and AltiSora, LLC., will develop new lightweight cementitious composite material technologies that will utilize wind turbine blade waste as a raw material to; (1) allow a high value addition (2) at low cost, enabling (3) a significant waste consumption volume; while (4) consuming the entire wind turbine blade, without (5) creating any waste or emissions and to also (6) offer specific benefits to communities involved.

17 WIND ENERGY↗
NASA NTRS1952

Dynamic Response of Turbine-blade Temperature to Exhaust-gas Temperature for Gas-turbine Engines

The frequency of blade temperature to exhaust gas temperature is presented for two locations in the blade and at several operating conditions. The frequency response was determined by Fourier analysis of transient data. Two analytical methods are presented, and results are compared with experimental data. Dynamic response of turbine-blade temperature to exhaust-gas temperature exhibited the form of an approximate first-order lag. The results are used to predict blade temperature for a typical controlled and uncontrolled gas-turbine engine.

HEAT TRANSFER THEORY AND EXPERIMENT↗
DOE OSTI2022

Thermo-economic analysis of a solid oxide fuel cell-gas turbine hybrid with commercial off-the-shelf gas turbine

The thermodynamic and economic performance of a natural gas-fueled solid oxide fuel cell (SOFC)-gas turbine (GT) hybrid system with a commercial off-the-shelf GT is investigated. Until today, commercial GTs are primarily engineered for the direct use of energy dense fuels, such as natural gas, and the integration of commercial GTs into a SOFC-GT hybrid remains challenging. In this work technically feasible and economically viable GT operating modes are identified to gauge the technology’s competitiveness on the free market. Steady state, full load operating conditions were established for various GT operating modes: I) constant spool speed operation, II) variable spool speed operation and III) partially closed compressor inlet guide vanes. For each GT operating mode, the performance was investigated over a range of SOFC fuel utilization factors, while considering physical constraints inside the SOFC, such as local temperature gradients in flow direction as well as overall cell temperature differences. The results of the thermodynamic evaluation served as inputs for the economic analysis of the SOFC-GT hybrid power plant. The results show that the integration of an off-the-shelf GT not only results in a significant derating of the GT, but also substantially impacts the SOFC operation, the main power producer in this SOFC-GT hybrid. To maximize the SOFC power output it is desirable operate the GT in a region of high air mass flow rates and low pressure ratios, which increases the number of stacks that can be accommodated and reduces the SOFC cooling requirement. The lowest costs of electricity are obtained at constant spool speed operation, while the highest efficiencies are reached at variable spool speed operation. Critical for the integration of off-the-shelf GTs, which historically have been designed for natural gas, is the surge margin. The largest surge margins are obtained by closing the compressor inlet guide vanes. Furthermore, higher SOFC fuel utilization factors are shown to increase the surge margin as the turbine firing temperature is decreased.

25 ENERGY STORAGE↗