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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

Forward acoustic performance of a model turbofan designed for a high specific flow (QF-14)

Forward noise and overall aerodynamic performance are presented for a high-tip-speed fan having an exceptionally high average axial Mach number at the rotor inlet. This high Mach number is intended to attenuate forward noise at both the design-speed takeoff point, and at the unconventional low-pressure-ratio, design-speed approach point. As speed was increased near design, all forward noise components were reduced, and rear noise in the discharge duct was increased, indicating that the high Mach number flow at the rotor face is attenuating forward noise at takeoff. The fan at takeoff is some 5.5 to 11 dB quieter than several reference fans. Data at the point closest to approach indicated tentatively that the design-speed approach mode was 3 dB quieter than the conventional mode.

Lucas, J. G.↗
NASA NTRS1964

Status of v/stol research and development in the united states

It is the purpose of this paper to review the status of V/-STOL research and development in the United States with particular emphasis on significant research results obtained since the last Anglo-American Conference in 1961. Research information will be presented dealing with the helicopter, with propeller, ducted-fan, and jet V/STOL aircraft, and with the general area of V/STOL handling-qualities requirements.

V/STOL aircraft↗
NASA NTRS1985

Feasibility of Mapping Velocity Flow Fields in SSME Powerhead by Laser Anemometry Techniques

Because of the flow environment associated with the SSME powerhead pressure (3000 psia), temperature (1800 R), and mechanical complexity and the high vibration test stand environment, detailed flow measurements are difficult to make. The feasibility of using laser anemometry techniques to map velocity flow fields in an SSME powerhead is studied. In the study three engine powerhead component flow environments: (1) the high pressure fuel turbopump preburner, (2) the fuel turbopump turbine rotor and stator region, and (3) the 180 deg turnaround duct - are being considered. Flow parameters measured by the anemometry techniques are time averaged values of the velocity magnitude and flow direction, turbulence intensity, velocity component correlation, integral time scale, and turbulence spectrum.

Pelaccio, D. G.↗
NASA NTRS2010

An Experimental Study of Fan Inflow Distortion Tone Noise

The tone noise generated when a fan ingests circumferentially distorted flow was studied by an experiment conducted with the Advanced Noise Control Fan at the NASA Glenn Research Center. The inflow was distorted by inserting cylindrical rods radially into the duct. The rods were arranged in circumferentially irregular patterns in three of the five configurations tested. Rods were held in place using a mounting ring with 30 equally spaced holes placed at an axial location one rotor chordlength upstream of the fan. Acoustic pressure was measured in the inlet and exhaust duct of the fan using the Rotating Rake fan tone measurement system. Sound power levels, calculated from the measured data, were plotted as a function of circumferential mode. An analytic description of the unsteady pressure distribution at the interaction plane between the stationary rods and the fan rotor is presented in a form suitable for representing the circumferentially irregularly placed rods. Terms in the analytical description for sound power were proven to be useful in determining the dominant circumferential modes measured in the experiment and the differences in mode power level between the configurations tested. Insight gained through this work will be useful in the development of tools to compute fan inflow distortion tone noise.

Koch, L. Danielle↗
NASA NTRS1978

Apparatus and method for reducing thermal stress in a turbine rotor

A gas turbine is described wherein the thermal stresses in the turbine rotor are reduced. The rotor includes a central disc with a peripheral rim, and a plurality of blades extending radially outward from the rim. To reduce thermal stresses, a duct arrangement is provided which selectively directs hot gases from the turbine combustor to the rim during the turbine start up. The hot gases from the combustor serve to heat the rim, and decrease the start up period necessary to bring the temperature profile of the rotor into the operating temperature range. After the start up period, the duct arrangement is then used to direct cool gases from the turbine compressor to the rim of the rotor in order to maintain a lower rotor equilibrium temperature.

Heller, J. A.↗
NASA NTRS1981

Sound generated in a cascade by three-dimensional disturbances convected in a subsonic flow

Discrete tone sound generation in a subsonic fan subject to three-dimensional disturbances is investigated. The analytical model used treats the fan rotor and stator as linear cascades of thin airfoils in a rectangular duct subject to a three-dimensional gust for which a complete aerodynamic theory already exists. The sound pressure can then be cast as the sum of a finite number of discrete sound waves (modes) the magnitude of which depends on an unknown function satisfying a singular integral equation. Similarity rules are derived to reduce the problem to that of a two-dimensional gust. Three-dimensional effects on the cut-off condition, the sound pressure, and the acoustic power are first investigated for each mode. The theory is then applied to noise generated by typical rotor-wake-defect and rotor-tip-vortex disturbances interacting with a stator.

Atassi, H.↗
NASA NTRS1989

The effects of Reynolds number, rotor incidence angle, and surface roughness on the heat transfer distribution in a large-scale turbine rotor passage

A combined experimental and computational program was conducted to examine the heat transfer distribution in a turbine rotor passage geometrically similiar to the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP). Heat transfer was measured and computed for both the full-span suction and pressure surfaces of the rotor airfoil as well as for the hub endwall surface. The primary objective of the program was to provide a benchmark-quality data base for the assessment of rotor passage heat transfer computational procedures. The experimental portion of the study was conducted in a large-scale, ambient temperature, rotating turbine model. Heat transfer data were obtained using thermocouple and liquid-crystal techniques to measure temperature distributions on the thin, electrically-heated skin of the rotor passage model. Test data were obtained for various combinations of Reynolds number, rotor incidence angle and model surface roughness. The data are reported in the form of contour maps of Stanton number. These heat distribution maps revealed numerous local effects produced by the three-dimensional flows within the rotor passage. Of particular importance were regions of local enhancement produced on the airfoil suction surface by the main-passage and tip-leakage vortices and on the hub endwall by the leading-edge horseshoe vortex system. The computational portion consisted of the application of a well-posed parabolized Navier-Stokes analysis to the calculation of the three-dimensional viscous flow through ducts simulating the a gas turbine passage. These cases include a 90 deg turning duct, a gas turbine cascade simulating a stator passage, and a gas turbine rotor passage including Coriolis forces. The calculated results were evaluated using experimental data of the three-dimensional velocity fields, wall static pressures, and wall heat transfer on the suction surface of the turbine airfoil and on the end wall. Particular attention was paid to an accurate modeling of the passage vortex and to the development of the wall boundary layers including crossflow.

Blair, Michael F.↗
NASA NTRS1980

Study of blade aspect ratio on a compressor front stage

A single stage, low aspect ratio, compressor with a 442.0 m/sec (1450 ft/sec) tip speed and a 0.597 hub/tip ratio typical of an advanced core compressor front stage was tested. The test stage incorporated an inlet duct which was representative of an engine transition duct between fan and high pressure compressors. At design speed, the rotor stator stage achieved a peak adiabatic efficiency of 86.6 percent at a flow of 44.35 kg/sec (97.8 lbm/sec) and a pressure ratio of 1.8. Surge margin was 12.5 percent from the peak stage efficiency point.

Behlke, R. F.↗
NASA NTRS1985

Forced response analysis of an aerodynamically detuned supersonic turbomachine rotor

High performance aircraft-engine fan and compressor blades are vulnerable to aerodynamically forced vibrations generated by inlet flow distortions due to wakes from upstream blade and vane rows, atmospheric gusts, and maldistributions in inlet ducts. In this report, an analysis is developed to predict the flow-induced forced response of an aerodynamically detuned rotor operating in a supersonic flow with a subsonic axial component. The aerodynamic detuning is achieved by alternating the circumferential spacing of adjacent rotor blades. The total unsteady aerodynamic loading acting on the blading, as a result of the convection of the transverse gust past the airfoil cascade and the resulting motion of the cascade, is developed in terms of influence coefficients. This analysis is used to investigate the effect of aerodynamic detuning on the forced response of a 12-blade rotor, with Verdon's Cascade B flow geometry as a uniformly spaced baseline configuration. The results of this study indicate that, for forward traveling wave gust excitations, aerodynamic detuning is very beneficial, resulting in significantly decreased maximum-amplitude blade responses for many interblade phase angles.

Hoyniak, D.↗
NASA NTRS1975

Prediction of Externally Blown Flap Noise and Turbomachinery Strut Noise

Methods were developed for predicting externally blown flap (EBF) noise and turbomachinery strut noise. Noise radiated by under-the-wing and upper-surface-blowing EBF configurations is calculated as a sum of lift dipole noise, trailing edge noise, and jet quadrupole noise. Resulting predictions of amplitudes and spectra generally were in good agreement with data from small-scale models. These data cover a range of exhaust velocity, flap deflection, exhaust nozzle position, exhaust nozzle shape, and ratio of exhaust nozzle diameter to wing chord. A semi-empirical method for predicting dipole noise radiation from a strut with incident turbulence was in good agreement with data. Leading-edge regions made of perforated plate backed by a bulk acoustic absorber achieved up to 7 db reduction of strut noise caused by incident turbulence at high frequencies. Radial turbulence in a turbofan exit duct was found to have a relatively high level associated with the mean velocity defect in the rotor blade wakes. Use of these turbulence spectra and a dipole noise radiation equation gave general prediction of measured aft-radiated sound power caused by a splitter ring in a full-scale fan exit duct.

Fink, M. R.↗
NASA NTRS2004

Fan Noise Prediction

Aircraft noise emission level restrictions in and around airports continue to grow more stringent every few years. Thus, it is important to predict noise emissions from aircraft accurately. Predicting noise from the engine(s) is an integral part of the efforts to characterize the noise signature of an aircraft. An important source of engine noise is the rotor-stator interaction noise produced as a result of impingement of fan rotor wakes on the fan exit guide vanes. Interaction noise propagates through the inlet and exhaust ducts of the engine and radiates to the far field. noise levels for a range of model fans stages that represent current aircraft engine designs. Eversman's radiation codes calculate both the inlet and exhaust noise radiation by propagating the internally measured rotor-stator interaction noise to the far field. Predicted far field sound pressure levels are then compared to the measured levels from wind tunnel tests. This effort's objective is to prove that the predicted levels actually describe the measured levels.

France, Joshua I.↗
NASA NTRS1985

Effect of Steady-State Temperature Distortion on Inlet Flow to a High-Bypass-Ratio Turbofan Engine

The effects of circumferential inlet temperature distortion on the flow characteristics between a distortion generator and a high bypass ratio turbofan engine and through its compression system were evaluated to support the effort to generate analytical models. The flow characteristics are defined by the inlet duct, the flow angles, and the total temperature, total pressure, and static pressure profiles in the inlet duct and through the fan and compressor. The effects of Reynolds number, rotor speed, and distortion extent are also considered.

Soeder, R. H.↗
NASA NTRS1983

Turbofan aft duct suppressor study

Suppressions due to acoustic treatment in the annular exhaust duct of a model fan were theoretically predicted and compared with measured suppressions. The predictions are based on the modal analysis of sound propagation in a straight annular flow duct with segmented treatment. Modal distributions of the fan noise source (fan-stator interaction only) were measured using in-duct modal probes. The flow profiles were also measured in the vicinity of the modal probes. The acoustic impedance of the single degree of freedom treatment was measured in the presence of grazing flow. The measured values of mode distribution of the fan noise source, the flow velocity profile and the acoustic impedance of the treatment in the duct were used as input to the prediction program. The predicted suppressions, under the assumption of uniform flow in the duct, compared well with the suppressions measured in the duct for all test conditions. The interaction modes generated by the rotor-stator interaction spanned a cut-off ratio range from nearly 1 to 7.

Syed, A. A.↗
NASA NTRS1977

Acoustic and aerodynamic performance of a 1.5-pressure-ratio, 1.83-meter (6 ft) diameter fan stage for turbofan engines (QF-2)

The fan was externally driven by an electric motor. Design features for low-noise generation included the elimination of inlet guide vanes, long axial spacing between the rotor and stator blade rows, and the selection of blade-vane numbers to achieve duct-mode cutoff. The fan QF-2 results were compared with those of another full-scale fan having essentially identical aerodynamic design except for nozzle geometry and the direction of rotation. The fan QF-2 aerodynamic results were also compared with those obtained from a 50.8 cm rotor-tip-diameter model of the reverse rotation fan QF-2 design. Differences in nozzle geometry other than exit area significantly affected the comparison of the results of the full-scale fans.

Woodward, R. P.↗
NASA NTRS1981

Rotor wake characteristics relevant to rotor-stator interaction noise generation

Mean and turbulent wake properties at three axial locations behind the rotor of an aerodynamically loaded 1.2 pressure ratio fan were measured using a stationary cross film anemometer in an anechoic wind tunnel. Wake characteristics at four radial immersions across the duct at four different fan speeds were determined utilizing a signal enhancement technique. The shapes of the waveforms of the mean rotor relative and mean upwash velocities were shown to change significantly across the span of the blades. In addition, an increase in fan rotational speed caused an increase in the maximum wake turbulence intensity levels near the hub and tip. Spectral analysis was used to described the complex nature of the rotor wake.

Shaw, L. M.↗
NASA NTRS1981

Rotor wake characteristics relevant to rotor-stator interaction noise generation

Mean and turbulence wake properties at three axial locations behind the rotor of an aerodynamically loaded 1.2 pressure ratio fan were measured using a stationary cross film anemometer in an anechoic wind tunnel. Wake characteristics at four radial immersions across the duct at four different fan speeds were determined utilizing a signal enhancement technique. The shapes of the waveforms of the mean rotor relative and mean upwash velocities were shown to change significantly across the span of the blades. In addition, an increase in fan rotational speed caused an increase in the maximum wake turbulence intensity levels near the hub and tip. Spectral analysis was used to describe the complex nature of the rotor wake.

Shaw, L. M.↗