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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 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1979

Aerodynamic performance of axial-flow fan stage operated at nine inlet guide vane angles

The overall performance of a fan stage with nine inlet guide vane angle settings is presented. These data were obtained over the stable flow range at speeds from 60 to 120 percent of design for vane setting angles from -25 to 42.5 degrees. At design speed and design inlet guide vane angle, the stage has a peak efficiency of 0.892 at a pressure ratio of 1.322 and a flow of 25.31 kg/s. The stall margin based on peak efficiency and stall was 20 percent. Based on an operating line passing through the peak efficiency point at the design setting angle, the useful operating range of the stage at design speed is limited by stall at the positive setting angles and by choke at the negative angles. At design the calculated static thrust along the operating line varied from 68 to 114 percent of that obtained at design setting angle.

Moore, R. D.↗
NASA NTRS1979

Effect of outboard vertical-fin position and orientation on the low-speed aerodynamic performance of highly swept wings

A theoretical study was conducted to determine the potential low-speed performance improvements which can be achieved by altering the position and orientation of the outboard vertical fins of low-aspect-ratio highly swept wings. Results show that the magnitude of the performance improvements is solely a function of the span-load distribution. Both the vertical-fin-chordwise position and toe angle provided effective means for adjusting the overall span-load distribution.

Johnson, V. S.↗
NASA NTRS1979

Effect of lip and centerbody geometry on aerodynamic performance of inlets for tilting-nacelle VTOL aircraft

Inlets for tilt-nacelle VTOL aircraft must operate over a wide range of incidence angles and engine weight flows without internal flow separation. Wind tunnel tests of scale model inlets were conducted to evaluate the effectiveness of three geometric variables to provide this capability. Increasing the lip contraction ratio increased the separation angle at all engine weight flows. The optimum axial location of the centerbody occurred when its leading edge was located just downstream of the inlet lip. Compared with a short centerbody, the optimum location of the centerbody resulted in an increase in separation angle at all engine weight flows. Decreasing the lip major-to-minor-axis ratio increased the separation angle at the lower engine weight flows.

Burley, R. R.↗
NASA NTRS1980

Aerodynamic performances of three fan stator designs operating with rotor having tip speed of 337 meters per second and pressure ratio of 1.54. Relation of analytical code calculations to experimental performance

A hub-to-shroud and a blade-to-blade internal-flow analysis code, both inviscid and basically subsonic, were used to calculate the flow parameters within four stator-blade rows. The produced ratios of maximum suction-surface velocity to trailing-edge velocity correlated well in the midspan region, with the measured total-parameters over the minimum-loss to near stall operating range for all stators and speeds studied. The potential benefits of a blade designed with the aid of these flow analysis codes are illustrated by a proposed redesign of one of the four stators studied. An overall efficiency improvement of 1.6 points above the peak measured for that stator is predicted for the redesign.

Gelder, T. F.↗
NASA NTRS1978

Low-speed aerodynamic performance of 50.8-centimeter-diameter noise-suppressing inlets for the Quiet, Clean, Short-haul Experimental Engine (QCSEE)

Two basic inlet concepts, a high throat Mach number (0.79) design and a low throat Mach number (0.60) design, were tested with four diffuser acoustical treatment designs that had face sheet porosity ranging from 0 to 24 percent for the high Mach number inlet and 0 to 28 percent for the low Mach number inlet. The tests were conducted in a low speed wind tunnel at free stream velocities of 0, 41, and 62 m/sec and angles of attack to 50 deg. Inlet throat Mach number was varied about the design value. Increasing the inlet diffuser face sheet porosity resulted in an increase in total pressure loss in the boundary layer for both the high and low Mach number inlet designs, however, the overall effect on inlet total pressure recovery of 0.991 at the design throat Mach number, a free stream velocity of 41 m/sec, and an angle of attack of 50 deg; Inlet flow separation at an angle of attack of 50 deg was encountered with only one inlet configuration the high Mach number design with the highest diffuser face sheet porosity (24 percent).

Abbott, J. M.↗
NASA NTRS1983

Wing Subsonic Aerodynamic Performance Estimates

Solution by iteration estimates performance of twisted and cambered wings of arbitrary planform. SUBAER based on linearized theory lifting-surface solution provides spanwise distribution of theoretical leading edge thrust, in addition to surface distribution of perturbation velocities. Program written in FORTRAN IV.

Carlson, H. W.↗
NASA NTRS1982

The aerodynamic performance of several flow control devices for internal flow systems

An experimental reseach and development program was undertaken to develop and document new flow-control devices for use in the major modifications to the 40 by 80 Foot wind tunnel at Ames Research Center. These devices, which are applicable to other facilities as well, included grid-type and quasi-two-dimensional flow straighteners, louver panels for valving, and turning-vane cascades with net turning angles from 0 deg to 90 deg. The tests were conducted at model scale over a Reynolds number range from 2 x 100,000 to 17 x 100,000, based on chord. The results showed quantitatively the performance benefits of faired, low-blockage, smooth-surface straightener systems, and the advantages of curved turning-vanes with hinge-line gaps sealed and a preferred chord-to-gap ratio between 2.5 and 3.0 for 45 deg or 90 deg turns.

Eckert, W. T.↗
NASA NTRS1985

A note on blade wake interaction influence on compressor stator row aerodynamic performance

Attention is given to the effect of blade wake interactions on the performance of an axial flow compressor's stator row, for the case of compressor and fan stator flows without shocks. The measured midspan loss of total pressure can be related to stator surface boundary layers, the chopped rotor wakes passing through the stator row, and the interaction between these flows. The interaction between rotor blade wake segments and stator blade surface boundary layers generates much higher losses than expected in both the stator wake and in the region of flow between stator boundary layers/wakes, if interaction effects are ignored.

Okiishi, T. H.↗
NASA NTRS1985

Aerodynamic performance of a 27-percent-scale AH-64 wind-tunnel model with baseline/advanced rotor blades

Performance of a 27-percent scale model of an advanced rotor designed for the AH-64 helicopter was measured in hover and at forward speeds of 50 to 130 knots. A baseline rotor, modeled after the current AH-64 rotor, was also tested to provide data for comparison. The investigation was conducted to validate procedures used at Langley to design rotors with increased performance potential and to provide a database for evaluation of current and future rotor systems. Both rotors were operated at full-scale tip speeds. Rotor thrust, forward speed, and ground height were varied for each rotor.

Kelley, H. L.↗
NASA NTRS1986

A new, improved method for separating turbulent boundary layer for aerodynamic performance prediction of trailing edge stall airfoils

A method based on integral techniques is presented to determine the turbulent separation location and separation pressure for trailing edge stall airfoils exhibiting extensive separated flow regions during prestall and poststall conditions. The method and experimental laser velocimetry data in the turbulence amplification region are used to establish a relationship between fluctuating velocity components and mean velocity in the boundary layer. The method is demonstrated for symmetrical, cambered, conventional and supercritical airfoils in the 5-to-21-percent thickness ratio range, and over a Reynolds number range of 2 to 22 million. Good agreement with experimental data is found, and the method is used in conjunction with the NASA Multi-Component Airfoil Computer Program to predict maximum L/D ratios and maximum section/lift coefficients.

Goradia, S.↗
NASA NTRS1985

Aerodynamic performance investigation of advanced mechanical suppressor and ejector nozzle concepts for jet noise reduction

Advanced Supersonic Transport jet noise may be reduced to Federal Air Regulation limits if recommended refinements to a recently developed ejector shroud exhaust system are successfully carried out. A two-part program consisting of a design study and a subscale model wind tunnel test effort conducted to define an acoustically treated ejector shroud exhaust system for supersonic transport application is described. Coannular, 20-chute, and ejector shroud exhaust systems were evaluated. Program results were used in a mission analysis study to determine aircraft takeoff gross weight to perform a nominal design mission, under Federal Aviation Regulation (1969), Part 36, Stage 3 noise constraints. Mission trade study results confirmed that the ejector shroud was the best of the three exhaust systems studied with a significant takeoff gross weight advantage over the 20-chute suppressor nozzle which was the second best.

Wagenknecht, C. D.↗
NASA NTRS1988

Experimental aerodynamic performance of advanced 40 deg-swept 10-blade propeller model at Mach 0.6 to 0.85

A propeller designated as SR-6, designed with 40 deg of sweep and 10 blades to cruise at Mach 0.8 at an altitude of 10.7 km (35,000 ft), was tested in the NASA Lewis Research Center's 8- by 6-Foot Wind Tunnel. This propeller was one of a series of advanced single rotation propeller models designed and tested as part of the NASA Advanced Turboprop Project. Design-point net efficiency was almost constant to Mach 0.75 but fell above this speed more rapidly than that of any previously tested advanced propeller. Alternative spinners that further reduced the near-hub interblade Mach numbers and relieved the observed hub choking improved performance above Mach 0.75. One spinner attained estimated SR-6 Design-point net deficiencies of 80.6 percent at Mach 0.75 and 79.2 percent at Mach 0.8, higher than the measured performance of any previously tested advanced single-rotation propeller at these speeds.

Mitchell, Glenn A.↗
NASA NTRS1987

Nacelle aerodynamic performance

The boundary layer transition location was measured on a nacelle shape using the sublimating chemical flow visualization technique. This technique involves coating the surface with a thin film of volatile chemical solid, which, during exposure to a free stream airflow, rapidly sublimates in the turbulent boundary layer as a result of high shear stress and high mass transfer near the surface. Transition is indicated because the chemical coating remains relatively unaffected in the laminar region due to lower shear and low mass transfer. The slow response time of the chemical in a laminar boundary allowed for two test conditions during the same flight. The aircraft was first flown at the desired airspeed and altitude with the noise source off. Once a pattern had developed, the noise source was turned on to the desired setting and a new chemical pattern was sought. In this fashion a direct comparison of the effect of the noise could be determined.

Obara, Clifford J.↗
NASA NTRS1990

Aerodynamic Performance of a 0.27-Scale Model of an AH-64 Helicopter with Baseline and Alternate Rotor Blade Sets

Performance of a 27 percent scale model rotor designed for the AH-64 helicopter (alternate rotor) was measured in hover and forward flight and compared against and AH-64 baseline rotor model. Thrust, rotor tip Mach number, advance ratio, and ground proximity were varied. In hover, at a nominal thrust coefficient of 0.0064, the power savings was about 6.4 percent for the alternate rotor compared to the baseline. The corresponding thrust increase at this condition was approx. 4.5 percent which represents an equivalent full scale increase in lift capability of about 660 lbs. Comparable results were noted in forward flight except for the high thrust, high speed cases investigated where the baseline rotor was slightly superior. Reduced performance at the higher thrusts and speeds was likely due to Reynolds number effects and blade elasticity differences.

Kelley, Henry L.↗
NASA NTRS1992

Effect of a simulated glaze ice shape on the aerodynamic performance of a rectangular wing

The effect of a simulated glaze-ice accretion on the flowfield of a 3D wing is studied experimentally. The model used for these tests was a semispan wing of effective aspect ratio five, mounted from the sidewall of a subsonic wind tunnel. The model has a NACA 0012 airfoil section on a rectangular untwisted planform with interchangeable leading edges to allow for testing both the baseline and the iced-wing geometry. A four-beam two-color fiberoptic laser Doppler velocimeter (LDV) was used to map the flowfield along three spanwise cuts on the model. Measurements on the centerline of the clean model compared favorably with theory and centerline measurements on the iced model compared well with measurements on a similar 2D model. The flow has the largest separation bubble at the model midspan with the smallest separation bubble occurring near the root and the wing tip.

Khodadoust, Abdi↗
NASA NTRS1993

Aerodynamic performance of a supersonic throughflow fan stage

The NASA LeRC has begun a program to design, analyze, build, and test a fan stage that can operate with supersonic axial velocities throughout the stage. This report presents the results from the initial experimental evaluation of the supersonic throughflow fan stage. The stage was tested with flows at subsonic and supersonic inlet axial Mach numbers and with rotor speeds at 40 to 100 percent of design speed. Overall performance as well as detailed performance obtained from radially and circumferentially traversing probes downstream of the stator vanes is presented. At the design rotor tip speed of 1500 ft/s and the design Mach number of 2.0, the measured overall pressure ratio was 2.22, and the corresponding adiabatic efficiency was 0.694. Transition from subsonic to supersonic flow was successfully demonstrated at 75, 90, and 100 percent of design speed. In that speed range, the stage 'unstart' (transition from supersonic to subsonic inlet axial conditions) occurred at an inlet axial Mach number of about 1.36, which corresponds to the Mach number at which the stage started.

Royce D Moore↗