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NASA NTRS2020

Novel Environmental Barrier Coatings for the Protection of SiC Components

Greater gas turbine engine efficiency is a major goal in aeronautics research often pursued through increased engine operating temperatures. However, it is necessary to replace the current hot-stage alloy components with more thermally robust parts, such as Silicon-based ceramics and composites. Unfortunately, these materials are still susceptible to the effects of oxidation, water vapor, and (Calcium-Magnesium-Alumino-Silicate) CMAS interaction, among other issues at high temperature. To mitigate these effects, environmental barrier coating (EBC) materials are employed to help control the rate of degradation to the underlying composite, but must also survive the corrosive environment. The current study explores new design space in Mg-based EBC materials with respect to microstructure, high temperature phase stability, volatilization in water vapor, as well as structure-property relationships at high temperature.

Kowalski, Benjamin A.↗

NASA NTRS2020

A Review of Distributed Electric Propulsion Concepts for Air Vehicle Technology

The continuous, yet incremental performance gains in high-efficiency gas turbine engines over the last eight decades has enabled passengers to travel long distances at high speeds. Modern commercial jet transports are so common that most passengers traveling by air today take the innovative developments of jet engine technologies for granted. However, the ever-increasing demands for travel in the 21st century has also brought an increased awareness of the energy and environmental concerns associated with aviation. Due to the forecasted limited supply of petroleum-based fuel sources and the global warming potential associated with emissions of traditional aircraft propulsion systems, the need for environmentally-responsible solutions in aircraft technology has been recognized as a priority in the development of future aircraft configurations.

Hyun Dae Kim↗

NASA NTRS1982

Cold-air performance of a 15.41-cm-tip-diameter axial-flow power turbine with variable-area stator designed for a 75-kW automotive gas turbine engine

An experimental evaluation of the aerodynamic performance of the axial flow, variable area stator power turbine stage for the Department of Energy upgraded automotive gas turbine engine was conducted in cold air. The interstage transition duct, the variable area stator, the rotor, and the exit diffuser were included in the evaluation of the turbine stage. The measured total blading efficiency was 0.096 less than the design value of 0.85. Large radial gradients in flow conditions were found at the exit of the interstage duct that adversely affected power turbine performance. Although power turbine efficiency was less than design, the turbine operating line corresponding to the steady state road load power curve was within 0.02 of the maximum available stage efficiency at any given speed.

Mclallin, K. L.↗

NASA NTRS2013

Multifunctional Low-Pressure Turbine for Core Noise Reduction, Improved Efficiency, and Nitrogen Oxide (NOx) Reduction

This work studied the feasibility of using Helmholtz resonator cavities embedded in low-pressure-turbine (LPT) airfoils to (1) reduce core noise by damping acoustic modes; (2) use the synthetic jets produced by the liner hole acoustic oscillations to improve engine efficiency by maintaining turbulent attached flow in the LPT at low-Reynolds-number cruise conditions; and (3) reduce engine nitrogen oxide emissions by lining the internal cavities with materials capable of catalytic conversion. Flat plates with embedded Helmholtz resonators, designed to resonate at either 3000 or at 400 Hz, were simulated using computational fluid dynamics. The simulations were conducted for two inlet Mach numbers, 0.25 and 0.5, corresponding to Reynolds numbers of 90 000 and 164 000 based on the effective chordwise distance to the resonator orifice. The results of this study are (1) the region of acoustic treatment may be large enough to have a benefit; (2) the jets may not possess sufficient strength to reduce flow separation (based on prior work by researchers in the flow control area); and (3) the additional catalytic surface area is not exposed to a high velocity, so it probably does not have any benefit.

Miller, Christopher J.↗

NASA NTRS2011

Computational Assessment of the Aerodynamic Performance of a Variable-Speed Power Turbine for Large Civil Tilt-Rotor Application

The main rotors of the NASA Large Civil Tilt-Rotor notional vehicle operate over a wide speed-range, from 100% at take-off to 54% at cruise. The variable-speed power turbine offers one approach by which to effect this speed variation. Key aero-challenges include high work factors at cruise and wide (40 to 60 deg.) incidence variations in blade and vane rows over the speed range. The turbine design approach must optimize cruise efficiency and minimize off-design penalties at take-off. The accuracy of the off-design incidence loss model is therefore critical to the turbine design. In this effort, 3-D computational analyses are used to assess the variation of turbine efficiency with speed change. The conceptual design of a 4-stage variable-speed power turbine for the Large Civil Tilt-Rotor application is first established at the meanline level. The design of 2-D airfoil sections and resulting 3-D blade and vane rows is documented. Three-dimensional Reynolds Averaged Navier-Stokes computations are used to assess the design and off-design performance of an embedded 1.5-stage portion-Rotor 1, Stator 2, and Rotor 2-of the turbine. The 3-D computational results yield the same efficiency versus speed trends predicted by meanline analyses, supporting the design choice to execute the turbine design at the cruise operating speed.

Welch, Gerard E.↗

NASA NTRS1976

Two-phase turbine engines

A description is given of a two-phase turbine which utilizes a uniform mixture of gas and liquid accelerated in nozzles of the types reported by Elliott and Weinberg (1968). The mixture acts directly on an axial flow or tangential impulse turbine or is separated into gas and liquid streams which operate separately on a gas turbine and a hydraulic turbine. The basic two-phase cycles are examined, taking into account working fluids, aspects of nozzle expansion, details of turbine cycle operation, and the effect of mixture ratio variation. Attention is also given to two-phase nozzle efficiency, two-phase turbine operating characteristics and efficiencies, separator turbines, and impulse turbine experiments.

Elliott, D. G.↗

NASA NTRS1994

Combustor technology for future small gas turbine aircraft

To enhance fuel efficiency, future advanced small gas turbine engines will utilize engine cycles calling for overall engine pressure ratios, leading to higher combustor inlet pressures and temperatures. Further, the temperature rise through the combustor and the corresponding exit temperature are also expected to increase. This report describes future combustor technology needs for small gas turbine engines. New fuel injectors with large turndown ratios which produce uniform circumferential and radial temperature patterns will be required. Uniform burning will be of greater importance because hot gas temperatures will approach turbine material limits. The higher combustion temperatures and increased radiation at high pressures will put a greater heat load on the combustor liners. At the same time, less cooling air will be available as more of the air will be used for combustion. Thus, improved cooling concepts and/or materials requiring little or no direct cooling will be required. Although presently there are no requirements for emissions levels from small gas turbine engines, regulation is anticipated in the near future. This will require the development of low emission combustors. In particular, nitrogen oxides will increase substantially if new technologies limiting their formation are not evolved and implemented. For example, staged combustion employing lean, premixed/prevaporized, lean direct injection, or rich burn-quick quench-lean burn concepts could replace conventional single stage combustors. Due to combustor size considerations, staged combustion is more easily accommodated in large engines. The inclusion of staged combustion in small engines will pose greater combustor design challenges.

Lyons, Valerie J.↗

NASA NTRSDate not supplied

Numerical Investigation of Combustor-Turbine Interactions with a Two-Stage High-Pressure Turbine Including Cooling Airflows

Combustor–turbine interactions (CTIs) are investigated by performing three-dimensional unsteady simulations using a realistic combustor and high-pressure turbine (HPT) configurations from the Energy Efficient Engine (E3) program. To understand CTIs, we compare the numerically predicted flow fields from single-component simulations (Step 1: the combustor + the first-stage stator of turbine; Step 2: the two-stage HPT imposing the time-averaged flow solution from Step 1 as the inflow boundary condition) and a fully coupled combustor–turbine simulation (Step 3) at the sea-level takeoff (SLTO) condition. In addition to three previous simulations where the cooling airflows inside the HPT had been neglected for all Step 1, Step 2, and Step 3, two new simulations of Step 2 and Step 3 take into account the cooling airflows using the source team approach. In this approach, to mimic the cooling airflows, we impose the source term at a specific area of a cooling airflow hole at the solid surfaces without making a mesh of each hole. The objective of this study is twofold. One is to investigate the effect of the presence of the cooling airflows on the aerodynamics of the combustor and HPT as well as the HPT performance. The second is to perform a detailed comparison among the calculated flowfields by two different numerical schemes, the central-difference with the standard Jameson–Schmidt–Turkel (CD-JST) scheme and the AUSM scheme. There is a noticeable difference in the hot-streak distributions at the first-stage stator. In addition, depending on the choice of the numerical scheme and the presence of the cooling airflows, an occurrence of shock waves appearing at the first-stage stator is greatly influenced. Thus, this has a noticeable impact on the HPT performance. It is shown that the estimated turbine efficiencies from Step 3 are about 7 % less than the ones from Step 2.

Combustor-turbine interaction↗

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 NTRSDate not supplied

Numerical Investigation of the Effect of Cooling Airflow inside High-Pressure Turbine using the Source Term Approach

In this study, the Open National Combustion code (OpenNCC) is applied to simulate the airflow inside the high-pressure turbine (HPT) of the energy efficient engine (EEE). The main objective of this study is to compare the characteristics of the HPT performance obtained by using three different approaches (Case 1: the realistic combustion products are considered at the HPT inlet, but the cooling airflows are not included, Case 2: the working fluid is assumed to be a single gas (i.e., air), and the cooling airflows are considered, Case 3: the cooling airflow and the combustion products at the HPT inlet are included). To mimic the cooling airflows at the HPT surface, we use the surface source team approach, in which we impose the source term at a specific area of a cooling airflow hole at the solid surfaces by specifying an injection angle, temperature, turbulent intensity, and mass flowrate of each cooling airflow. For the validation of the model, the data for the EEE model was obtained from the General Electric (GE) test campaign (RDG10), and the full scale warm-air rig test condition representing a test point close to the integrated core/low spool design condition was considered. This validation test was done using the stationary uniform inflow condition (i.e., the reported total temperature and total pressure). Secondly, we considered the Sea-Level Take-off (SLTO) engine condition and separately performed the combustor simulation to obtain a more realistic time-averaged, spatially-nonuniform HPT inflow condition. This work provides a better understanding the effect of the cooling airflow model with different assumptions on the HPT performance. It is found that using different working gas have a minor effect on the performance metric, however the presence of the cooling airflows significantly affects the HPT performance such as the turbine efficiency.

CFD Hot-streaks↗

NASA NTRS1974

Highly loaded multi-stage fan drive turbine - performance of initial seven configurations

Experimental results of a three-stage highly loaded fan drive turbine test program are presented. A plain blade turbine, a tandem blade turbine, and a tangentially leaned stator turbine were designed for the same velocity diagram and flowpath. Seven combinations of bladerows were tested to evaluate stage performances and effects of the tandem blading and leaned stator. The plain blade turbine design point total-to-total efficiency was 0.886. The turbine with the stage three leaned stator had the same efficiency with an improved exit swirl profile and increased hub reaction. Two-stage group tests showed that the two-stage turbine with tandem stage two stator had an efficiency of 0.880 compared to 0.868 for the plain blade two-stage turbine.

Wolfmeyer, G. W.↗

NASA NTRS1947

Preliminary Results of an Altitude-Wind-Tunnel Investigation of a TG-100A Gas Turbine-Propeller Engine: Compressor and Turbine Performance Characteristics - 4

As part of an investigation of the performance and operational characteristics of the TG-100A gas turbine-propeller engine, conducted in the Cleveland altitude wind tunnel, the performance characteristics of the compressor and the turbine were obtained. The data presented were obtained at a compressor-inlet ram-pressure ratio of 1.00 for altitudes from 5000 to 35,000 feet, engine speeds from 8000 to 13,000 rpm, and turbine-inlet temperatures from 1400 to 2100R. The highest compressor pressure ratio was 6.15 at a corrected air flow of 23.7 pounds per second and a corrected turbine-inlet temperature of 2475R. Peak adiabatic compressor efficiencies of about 77 percent were obtained near the value of corrected air flow corresponding to a corrected engine speed of 13,000 rpm. This maximum efficiency may be somewhat low, however, because of dirt accumulations on the compressor blades. A maximum adiabatic turbine efficiency of 81.5 percent was obtained at rated engine speed for all altitudes and turbine-inlet temperatures investigated.

Wallner, Lewis E.↗

NASA NTRS2001

Mechanical Components Branch Overview

Mr. James Zakrajsek, chief of the Mechanical Components Branch, presented an overview of research conducted by the branch. Branch members perform basic research on mechanical components and systems, including gears and bearings, turbine seals, structural and thermal barrier seals, and space mechanisms. The research is focused on propulsion systems for present and advanced aerospace vehicles. For rotorcraft and conventional aircraft, we conduct research to develop technology needed to enable the design of low noise, ultra safe geared drive systems. We develop and validate analytical models for gear crack propagation, gear dynamics and noise, gear diagnostics, bearing dynamics, and thermal analyses of gear systems using experimental data from various component test rigs. In seal research we develop and test advanced turbine seal concepts to increase efficiency and durability of turbine engines. We perform experimental and analytical research to develop advanced thermal barrier seals and structural seals for current and next generation space vehicles. In space mechanisms, we conduct fundamental research on lubricants, materials, components and mechanisms subjected to deep space and planetary environments.

Source record↗

NASA NTRS2002

Overview of Glenn Mechanical Components Branch Research

Mr. James Zakrajsek, chief of the Mechanical Components Branch, gave an overview of research conducted by the branch. Branch members perform basic research on mechanical components and systems, including gears and bearings, turbine seals, structural and thermal barrier seals, and space mechanisms. The research is focused on propulsion systems for present and advanced aerospace vehicles. For rotorcraft and conventional aircraft, we conduct research to develop technology needed to enable the design of low noise, ultra safe geared drive systems. We develop and validate analytical models for gear crack propagation, gear dynamics and noise, gear diagnostics, bearing dynamics, and thermal analyses of gear systems using experimental data from various component test rigs. In seal research we develop and test advanced turbine seal concepts to increase efficiency and durability of turbine engines. We perform experimental and analytical research to develop advanced thermal barrier seals and structural seals for current and next generation space vehicles. Our space mechanisms research involves fundamental investigation of lubricants, materials, components and mechanisms for deep space and planetary environments.

Zakrajsek, James↗

NASA NTRS1987

Liner cooling research at NASA Lewis Research Center

Described are recently completed and current advanced liner research applicable to advanced small gas turbine engines. Research relating to the evolution of fuel efficient small gas turbine engines capable of meeting future commercial and military aviation needs is currently under way at NASA Lewis Research Center. As part of this research, a reverse-flow combustor geometry was maintained while different advanced liner wall cooling techniques were investigated and compared to a baseline combustor. The performance of the combustors featuring counterflow film-cooled (CFFC) panels, transpiration cooled liner walls (TRANS), and compliant metal/ceramic (CMC) walls was obtained over a range of simulated flight conditions of a 16:1 pressure ratio gas turbine engine and fuel/air ratios up to 0.034. All the combustors featured an identical fuel injection system, identical geometric configuration outline, and similar designed internal aerothermodynamics.

Acosta, Waldo A.↗

NASA NTRS1987

Liner cooling research at NASA Lewis Research Center

Acosta, Waldo A.↗

NASA NTRS1957

Component Performance Investigation of J71 Experimental Turbine VII: Effect of First-stator Adjustment; Over-all Performance of J71-97 Turbine with 87-percent-design Stator Area

The performance results, as well as a compressor-turbine match point study based on a particular mode of engine operation, are compared with those previously obtained with the same turbine having first-stator areas 70, 95.6, and 132 percent of design. Maximum efficiencies for all turbine configurations were high, on the order of 87 to 89 percent. The efficiencies obtained at the match points for the 87-, 95.6-, and 132-percent turbines were slightly lower. No match point for the 70-percent turbine was obtained because, with first-stator areas less than 75.5 percent of design, the turbine would develop insufficient work to drive the compressor at the match point.

Schum, Harold J↗

NASA NTRS1949

Experimental study of flow past turbine blades

The requirements on gas turbines for aircraft power units, namely, adequate efficiency, operation at high gas temperatures, low weight, and small dimensions, must be taken into consideration during the design of the blading. To secure good efficiency, it is necessary that the gas flow past the blades as smoothly as possible without separation. This is relatively easily obtainable in the accelerated flow of turbine blading, if the blade spacing is chosen small enough. A small blade spacing, however, is detrimental to the other requirements outlined above. Operation at high gas temperatures usually calls for blade cooling. This cooling is associated with a power input that lowers the turbine efficiency. Since the amount of heat that must be carried off for coding a blade can be influenced rather little, the gross power input for a turbine stage can be reduced by keeping the number of blades to a minimum, that is, with blades of high spacing ratio. But here also a limit is imposed, the exceeding of which is followed by separation of flow. Hence the requirement of finding blade forms on which the flow separates at rather high spacing ratios .

Eckert, E.↗