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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 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Hover performance analysis of advanced rotor blades

This is an effort aimed at validating recent hover prediction methods. The experimental basis for this validation work is an extensive set of loads, wake and performance data, which were obtained from a pressure instrumented model UH-60 rotor tested at the Sikorsky hover test facility and at Duits-Nederlandse Windtunnel (DNW). This model was equipped with replaceable tips - including a tapered and a BERP-type tip - which permitted studies of the effects of rotor geometry. The central prediction method studied is a free-wake, vortex embedded, full-potential CFD method - called HELIX-I. It is found that the HELIX-I code produces very good comparisons with the data including wake, surface pressure and performance. Comparisons with the measured radial load distributions have permitted an improved understanding of the wake resolution modelling requirements of CFD methods. Since HELIX-I is a combined Eulerian/Lagrangian method, limited comparisons are also made with a Lagrangian boundary element code (called EHPIC) and an Eulerian Navier-Stokes code (called TURNS). In most cases all methods produce good comparisons with the data. It is found that the HELIX-I code provides a good compromise between the speed of boundary integral methods and the comprehensive nature of Navier-Stokes methods.

Tung, C.↗
NASA NTRS1982

Rotor tip clearance effects on overall and blade-element performance of axial-flow transonic fan stage

The effects of tip clearance on the overall and blade-element performance of an axial-flow transonic fan stage are presented. The 50-centimeter-diameter fan was tested at four tip clearances (nonrotating) from 0.061 to 0.178 centimeter. The calculated radial growth of the blades was 0.040 centimeter at design conditions. The decrease in overall stage performance with increasing clearance is attributed to the loss in rotor performance. For the rotor the effects of clearance on performance parameters extended to about 70 percent of blade span from the tip. The stage still margin based on an assumed operating line decreased from 15.3 to 0 percent as the clearance increased from 0.061 to 0.178 centimeter.

Moore, R. D.↗
NASA NTRS1979

Experimental investigation of effects of blade tip geometry on loads and performance for an articulated rotor system

Wind-tunnel tests of an aeroelastically designed helicopter rotor model were carried out to determine the effects on dynamic response and aerodynamic performance of varying the design of the outboard 8 percent of the blade lengths. Four different blade tip geometries or shapes having different amounts of planform sweep, thickness and chordwise taper, and anhedral angle were studied. Each configuration was tested at several shaft angles of attack for advance ratios of 0.20, 0.30, 0.35, and 0.40. For each combination of shaft angle and advance ratio, rotor lift was varied over a wide range to include high lift conditions.

Weller, W. H.↗
NASA NTRS1990

Status of integrated multidisciplinary rotorcraft optimization research at the Langley Research Center

This paper describes a joint NASA/Army research activity at the Langley Research Center to develop optimization procedures aimed at improving the rotor blade design process by integrating appropriate disciplines and accounting for important interactions among the disciplines. The activity is being guided by a Steering Committee made up of key NASA and Army researchers and managers. The paper describes the optimization formulation in terms of the objective function, design variables, and constraints. The analysis aspects are discussed, and the interdisciplinary interactions are defined in terms of the information that must be transferred among disciplinary analyses as well as the trade-offs between disciplines in determining the details of the design. At this writing, some significant progress has been made. Results given in the paper represent accomplishments in rotor aerodynamic performance optimization for minimum horsepower, rotor dynamic optimization for vibration reduction, approximate analysis of frequencies and mode shapes, rotor structural optimization for minimum weight, and integrated aerodynamic load/dynamics optimization for minimum vibration and weight.

Mantay, Wayne R.↗
NASA NTRS2001

Full-Span Tiltrotor Aeroacoustic Model (TRAM) Overview and 40- by 80-Foot Wind Tunnel Test

Most helicopter data trends cannot be extrapolated to tiltrotors because blade geometry and aerodynamic behavior, as well as rotor and fuselage interactions, are significantly different for tiltrotors. A tiltrotor model has been developed to investigate the aeromechanics of tiltrotors, to develop a comprehensive database for validating tiltrotor analyses, and to provide a research platform for supporting future tiltrotor designs. The Full-Span Tiltrotor Aeroacoustic Model (FS TRAM) is a dual-rotor, powered aircraft model with extensive instrumentation for measurement of structural and aerodynamic loads. This paper will present the Full-Span TRAM test capabilities and the first set of data obtained during a 40- by 80-Foot Wind Tunnel test conducted in late 2000 at NASA Ames Research Center. The Full-Span TRAM is a quarter-scale representation of the V-22 Osprey aircraft, and a heavily instrumented NASA and U.S. Army wind tunnel test stand. Rotor structural loads are monitored and recorded for safety-of-flight and for information on blade loads and dynamics. Left and right rotor balance and fuselage balance loads are monitored for safety-of-flight and for measurement of vehicle and rotor aerodynamic performance. Static pressure taps on the left wing are used to determine rotor/wing interactional effects and rotor blade dynamic pressures measure blade airloads. All of these measurement capabilities make the FS TRAM test stand a unique and valuable asset for validation of computational codes and to aid in future tiltrotor designs. The Full-Span TRAM was tested in the NASA Ames Research Center 40- by 80-Foot Wind Tunnel from October through December 2000. Rotor and vehicle performance measurements were acquired in addition to wing pressures, rotor acoustics, and Laser Light Sheet (LLS) flow visualization data. Hover, forward flight, and airframe (rotors off) aerodynamic runs were performed. Helicopter-mode data were acquired during angle of attack and thrust sweeps for a variety of tunnel speeds. Wake geometry images were acquired using LLS photographs and suggest dual tip vortex formation at low thrust conditions. The full paper will include comparisons to isolated-rotor TRAM data acquired at the Duits-Nederlandse Windtunnel (DNW) in 1998. The FS TRAM has been established as a valuable national asset for tiltrotor research. Data reduction and analysis of the 40- by 80-Foot Wind Tunnel test results are underway. Follow-on testing of the FS TRAM is currently being planned for the NASA Ames 80- by 120-Foot Wind Tunnel in late 2001.

McCluer, Megan S.↗
NASA NTRS1989

Optimization of rotor blades for combined structural, performance, and aeroelastic characteristics

The strategies whereby helicopter rotor blades can be optimized for combined structural, inertial, dynamic, aeroelastic, and aerodynamic performance characteristics are outlined. There are three key ingredients in the successful execution of such an interdisciplinary optimization. The first is the definition of a satisfactory performance index that combines all aspects of the problem without too many constraints. The second element is the judicious choice of computationally efficient analysis tools for the various quantitative components in both the cost functional and constraints. The third element is an effective strategy for combining the various disciplines either in parallel or sequential optimizations.

Peters, David A.↗
NASA NTRS1957

Experimental investigation of an axial-flow-compressor inlet stage operating at transonic relative inlet Mach numbers V : rotor blade-element performance at a reduced blade angle

At a corrected speed of 1100 feet per second, the low-blade-angle rotor operated with a relative inlet Mach number of 1.2, a diffusion factor of 0.65, and an axial velocity ratio of 0.71 in the tip region (11 percent of passage height away from the outer wall). The measured minimum-loss coefficient was 0.35, and this value falls above a previous correlation of rotor losses with diffusion factor. Through a comparison with data for three other rotors, the occurrence of high losses was related to a high suction-surface Mach number. These comparisons also indicated that axial velocity ratios between 0.73 and 1.10 have no independent effect on losses.

Schwenk, Francis C↗
NASA NTRS2012

Counteracting Rotor Imbalance in a Bearingless Motor System with Feedforward Control

In standard motor applications, traditional mechanical bearings represent the most economical approach to rotor suspension. However, in certain high performance applications, rotor suspension without bearing contact is either required or highly beneficial. Such applications include very high speed, extreme environment, or limited maintenance access applications. This paper extends upon a novel bearingless motor concept, in which full five-axis levitation and rotation of the rotor is achieved using two motors with opposing conical air-gaps. By leaving the motors' pole-pairs unconnected, different d-axis flux in each pole-pair is created, generating a flux imbalance which creates lateral force. Note this is approach is different than that used in previous bearingless motors, which use separate windings for levitation and rotation. This paper will examine the use of feedforward control to counteract synchronous whirl caused by rotor imbalance. Experimental results will be presented showing the performance of a prototype bearingless system, which was sized for a high speed flywheel energy storage application, with and without feedforward control.

Bearingless Motor↗
NASA NTRS1986

Effects of planform geometry on hover performance of a 2-meter-diameter model of a four-bladed rotor

Hover tests were conducted on three small scale rotors to evaluate the effects of blade planform taper on rotor hover performance. Tests were conducted on a rectangular swept-tip configuration, on a configuration with a 3 to 1 taper over the outboard 20 percent of the span, and on a configuration with a 5 to 1 taper over the outboard 20 percent of the blade span. The investigation covered a range of thrust coefficients from 0 to 0.0075 and a range of tip speeds from 300 to 600 ft/sec. The tests showed that both tapered configurations had better hover performance than the swept-tip rectangular configuration and that the 3 to 1 taper configuration was better than the 5 to 1 taper configuration. The test results were compared with predictions made with a prescribed wake analysis, a momentum analysis, and a simplified free wake analysis.

Phelps, A. E., III↗
NASA NTRS2015

Far-Field Acoustic Power Level and Performance Analyses of F31/A31 Open Rotor Model at Simulated Scaled Takeoff, Nominal Takeoff, and Approach Conditions: Technical Report I

Far-field acoustic power level and performance analyses of open rotor model F31/A31 have been performed to determine its noise characteristics at simulated scaled takeoff, nominal takeoff, and approach flight conditions. The nonproprietary parts of the data obtained from experiments in 9- by 15-Foot Low-Speed Wind Tunnel (915 LSWT) tests were provided by NASA Glenn Research Center to perform the analyses. The tone and broadband noise components have been separated from raw test data by using a new data analysis tool. Results in terms of sound pressure levels, acoustic power levels, and their variations with rotor speed, angle of attack, thrust, and input shaft power have been presented and discussed. The effect of an upstream pylon on the noise levels of the model has been addressed. Empirical equations relating model's acoustic power level, thrust, and input shaft power have been developed. The far-field acoustic efficiency of the model is also determined for various simulated flight conditions. It is intended that the results presented in this work will serve as a database for comparison and improvement of other open rotor blade designs and also for validating open rotor noise prediction codes.

Prop-Fan Technology↗
NASA NTRS1992

Numerical investigation of performance degradation of wings and rotors due to icing

The aerodynamic load characteristics and the performance degradation of moderate aspect ratio wings and rotors with simulated glaze leading-edge ice have been studied using a three-dimensional, compressible Navier-Stokes solver. The effect of a splitter plate at the wing root on both clean and iced wing configurations has been studied and the results are compared with the experiment. A significant difference has been observed with and without splitter plates in the magnitude of flow separation and aerodynamic loading at the inboard stations for the iced wing at 8-deg angle of attack. Inviscid calculations were performed and compared with viscous calculations to investigate whether the performance of iced swept wings can be inexpensively predicted using Euler methods. It is shown that inviscid calculations predict higher aerodynamic loading than viscous calculations, and cannot model separation effects. A typical nonlifting helicopter rotor in forward flight condition is also studied, and the penalty due to the leading-edge ice formation on the required torque is numerically demonstrated.

Kwon, Oh J.↗
NASA NTRS1974

Performance of a transonic compressor rotor with an aspect ratio of 6.5

The overall and blade-element performances and the aerodynamic design parameters of a transonic rotor with an aspect ratio of 6.5, designed to investigate the effects of aspect ratio on range and performance are presented. The rotor was designed for a total-pressure ratio of 1.53, an efficiency of 0.898, and a weight flow of 67.76 pounds per second.

Reid, L.↗
NASA NTRS1951

Investigation of Off-design Performance of Shock-in-rotor Type Supersonic Blading

An analysis of the off-design performance of shock-in rotor-type supersonic blading is presented. The over-all losses are assumed divided into two groups;normal-shock losses, and external-wave losses. Throughout the range of relative entrance Mach numbers from 1.35 to starting, the largest losses can be attributed to the normal shock. An improvement in performance would be expected with a reduction of the normal-shock losses by the introduction of external compression waves.

CASCADES, EXPERIMENT↗
NASA NTRS1979

An off-design correlation of part span damper losses through transonic axial fan rotors

The experimental performance of 10 transonic fan rotors was used to correlate losses caused by midchord part-span dampers (PSD) during off-design operation between 50 and 100 percent of design speed. The design tip speed for the rotors used varied from 419 to 425 m/s and the design pressure ratios from 1.6 to 2.0. The loss attributable to the damper and the region influenced along the blade height was correlated with relevant aerodynamic and geometric parameters. The losses at the design point were estimated by a previously reported correlation (Roberts, 1978). Using this as a base, the off-design losses were correlated with variation in blade suction surface incidence. A check with independent data showed that the prediction of damper losses and region of influence was fair to good for most of the off-design data examined.

Roberts, W. B.↗
NASA NTRS1977

Cold-air performance of a 12.766-centimeter-tip-diameter axial-flow cooled turbine. 3: Effect of rotor tip clearance on overall performance of a solid blade configuration

Two tip clearance configurations, one with a recess in the casing and the other with a reduced rotor blade height, were investigated at design equivalent speed over a range of tip clearance from about 2.0 to 5.0 percent of the stator blade height. The optimum configuration with a recess in the casing was the one where the rotor tip diameter was equal to the stator tip diameter (zero blade extension). For this configuration there was an approximate 1.5 percent decrease in total efficiency for an increase in tip clearance of 1 percent of stator blade height. For the reduced blade height configurations there was an approximate 2.0 percent decrease in total efficiency for an increase in tip clearance of 1 percent of stator blade height.

Haas, J. E.↗
NASA NTRS1971

Effect of casing treatment on overall and blade element performance of a compressor rotor

An axial flow compressor rotor was tested at design speed with six different casing treatments across the rotor tip. Radial surveys of pressure, temperature, and flow angle were taken at the rotor inlet and outlet. Surveys were taken at several weight flows for each treatment. All the casings treatments decreased the weight flow at stall over that for the solid casing. Radial surveys indicate that the performance over the entire radial span of the blade is affected by the treatment across the rotor tip.

Moore, R. D.↗