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At least 307 records · Page 17
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

Surface Temperature Measurements from a Stator Vane Doublet in a Turbine Engine Afterburner Flame Using a YAG:Tm Thermographic Phosphor

Luminescence-based surface temperature measurements were obtained from a YAG:Tm-coated stator vane doublet exposed to the afterburner flame of a J85 test engine at University of Tennessee Space Institute (UTSI). The objective of the testing was to demonstrate that reliable surface temperatures based on luminescence decay of a thermographic phosphor producing short-wavelength emission could be obtained from the surface of an actual engine component in a high gas velocity, highly radiative afterburner flame environment. YAG:Tm was selected as the thermographic phosphor for its blue emission at 456 nm (1D23F4 transition) and UV emission at 365 nm (1D23H6 transition) because background thermal radiation is lower at these wavelengths, which are shorter than those of many previously used thermographic phosphors. Luminescence decay measurements were acquired using a probe designed to operate in the afterburner flame environment. The probe was mounted on the sidewall of a high-pressure turbine vane doublet from a Honeywell TECH7000 turbine engine coated with a standard electron-beam physical vapor deposited (EB-PVD) 200-m-thick TBC composed of yttria-stabilized zirconia (YSZ) onto which a 25-m-thick YAG:Tm thermographic phosphor layer was deposited by solution precursor plasma spray (SPPS). Spot temperature measurements were obtained by measuring luminescence decay times at different afterburner power settings and then converting decay time to temperature via calibration curves. Temperature measurements using the decays of the 456 and 365 nm emissions are compared. While successful afterburner environment measurements were obtained to about 1300C with the 456 nm emission, successful temperature measurements using the 365 nm emission were limited to about 1100C due to interference by autofluorescence of probe optics at short decay times.

thermographic phosphor↗
NASA NTRS2015

Unsteady Loss in the Stator Due to the Incoming Rotor Wake in a Highly-Loaded Transonic Compressor

The present paper reports an investigation of unsteady loss generation in the stator due to the incoming rotor wake in an advanced GE transonic compressor design with a high-fidelity numerical method. This advanced compressor with high reaction and high stage loading has been investigated both experimentally and analytically in the past. The measured efficiency in this advanced compressor is significantly lower than the design intention goal. The general understanding is that the current generation of compressor design analysis tools miss some important flow physics in this modern compressor design. To pinpoint the source of the efficiency miss, an advanced test with a detailed flow traverse was performed for the front one and a half stage at the NASA Glenn Research Center.

Compressor↗
NASA NTRSDate not supplied

Design Study of a Coupled Inner-Stator Magnetically Geared Motor for Electric Aircraft Applications

Electric aircraft require high performance and high reliability electric motor drivetrains. A geared electric motor drivetrain will outperform a direct drive motor drivetrain in most applications. Traditional mechanical gearing, however, has mechanical contact-based wear and failure modes that result in added maintenance costs and require an oil lubrication system. Magnetically geared motor drives are a potential technology for electric aircraft applications capable of enabling the benefits of a geared drive without the maintenance, reliability, and lubrication system cost of mechanical gears. In this paper, a topology of magnetically geared motor, an inner stator magnetically geared motor, is explored to estimate its achievable performance for electric aircraft applications. Optimization results on the topology show that it can achieve greater than 15 Nm/kg and 96% efficiency at 100 kW of power.

Thomas Tallerico↗
NASA NTRSDate not supplied

Design Study of a Coupled Inner-Stator Magnetically Geared Motor for Electric Aircraft Applications

Abstract- Electric aircraft require high performance and high reliability electric motor drivetrains. A geared electric motor drivetrain will outperform a direct drive motor drivetrain in most applications. Traditional mechanical gearing, however, has mechanical contact-based wear and failure modes that result in added maintenance costs and require an oil lubrication system. Magnetically geared motor drives are a potential technology for electric aircraft applications capable of enabling the benefits of a geared drive without the maintenance, reliability, and lubrication system cost of mechanical gears. In this paper, a topology of magnetically geared motor, an inner-stator magnetically geared motor, is explored to estimate its achievable performance for electric aircraft applications. Optimization results on the topology show that it can achieve greater than 15 Nm/kg specific torque and 96% efficiency at 100 kW of power.

Thomas Tallerico↗
NASA NTRS1972

Effect of trailing edge geometry and thickness on the performance of certain turbine stator blading

The experimental and analytical investigation included solid blades with five different trailing-edge thicknesses and four different trailing-edge geometries. One of the geometries was round, one was square, one was tapered from the suction surface, and the other tapered from the pressure surface. One of the trailing-edge thicknesses was sharp edged; the other four thicknesses were equivalent to about 5, 11, 16, and 20 percent of the blade throat width. The experimental results show increased efficiency loss for increased trailing-edge thickness for all trailing-edge geometries. The blade with round trailing edge, equal to about 11 percent of the blade throat width, had 60 percent more loss than the sharp-edged blade. For the same trailing-edge thickness, square trailing edges caused more loss than round trailing edges, and the tapered trailing edges caused about the same loss as the round trailing edges.

Prust, H. W., Jr.↗
NASA NTRS1972

Nonuniform flow field generation for supersonic compressor stator development: Design and preliminary results

Design and preliminary results are given for a technique which can generate a nonuniform flow similar to the discharge of an impulse supersonic compressor rotor. The technique utilizes a carefully contoured, two-dimensional blunt body in a conventional hypersonic wind tunnel to generate the required flow field. To show the effects of the flow vorticity, a cascade of impulse-type blades was tested in this test stream. Some preliminary performance data are given along with comparison with previously determined uniform flow results.

Lipfert, F. W.↗
NASA NTRS1972

Flow conditions around the exit and downstream of certain stator blading with various trailing edge thicknesses and geometries

The blading investigated was curved back with thick profiles. The variations in flow conditions considered were flow angle and isentropic energy. Experimental data were obtained in a two-dimensional cascade from surveys with a combined angle, total-, and static-pressure probe and from an array of end-wall static pressure taps. Analytical results were obtained from ideal flow theory. The results showed large variations in flow conditions close to the plane of the trailing edge that were largely attenuated at a distance a little greater than one blade pitch downstream of the trailing edge in the direction of flow. The results were affected by the geometry and thickness of the trailing edge. The agreement between experimental and analytical results is generally fair to excellent.

Prust, H. W., Jr.↗
NASA NTRS1973

Cold-air aerodynamic study in a two-dimensional cascade of a turbine stator blade with suction-surface film cooling

The effect on aerodynamic performance of a single row of spanwise-spaced coolant holes located at four positions chordwise along the suction surface was investigated. In addition, multiple-row data were obtained. The data are presented in terms of primary efficiency, coolant-to-primary-flow percentage, and coolant pressure ratio. Primary efficiencies of multirow blade configurations compare satisfactorily with that calculated from the single-row efficiency increments. At any given coolant flow percentage, the efficiency was about the same for all row locations. For a given coolant pressure, the efficiency varied by as much as 1 percent, depending on the row location.

Brown, D. B.↗
NASA NTRS1973

Single-stage experimental evaluation of tandem-airfoil rotor stator blading for compressors. Part 6: Data and performance for stage D

An axial flow compressor stage, having single-airfoil blading, was designed for zero rotor prewhirl, constant rotor work across the span, and axial discharge flow. The stage was designed to produce a pressure ratio of 1.265 at a rotor tip velocity of 757 ft/sec. The rotor had an inlet hub/tip ratio of 0.8. The design procedure accounted for the rotor inlet boundary layer and included the effects of axial velocity ratio and secondary flow on blade row performance. The objectives of this experimental program were: (1) to obtain performance with uniform and distorted inlet flow for comparison with the performance of a stage consisting of tandem-airfoil blading designed for the same vector diagrams; and (2) to evaluate the effectiveness of accounting for the inlet boundary layer, axial velocity ratio, and secondary flows in the stage design. With uniform inlet flow, the rotor achieved a maximum adiabatic efficiency of 90.1% at design equivalent rotor speed and a pressure ratio of 1.281. The stage maximum adiabatic efficiency at design equivalent rotor speed with uniform inlet flow was 86.1% at a pressure ratio of 1.266. Hub radial, tip radial, and circumferential distortion of the inlet flow caused reductions in surge pressure ratio of approximately 2, 10 and 5%, respectively, at design rotor speed.

Clemmons, D. R.↗