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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 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
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.↗
NASA NTRS1974

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 7: Data and performance for stage E

An axial flow compressor stage, having tandem 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 has 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 single-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.

Cheatham, J. G.↗
NASA NTRS2013

A CFD Database for Airfoils and Wings at Post-Stall Angles of Attack

This paper presents selected results from an ongoing effort to develop an aerodynamic database from Reynolds-Averaged Navier-Stokes (RANS) computational analysis of airfoils and wings at stall and post-stall angles of attack. The data obtained from this effort will be used for validation and refinement of a low-order post-stall prediction method developed at NCSU, and to fill existing gaps in high angle of attack data in the literature. Such data could have potential applications in post-stall flight dynamics, helicopter aerodynamics and wind turbine aerodynamics. An overview of the NASA TetrUSS CFD package used for the RANS computational approach is presented. Detailed results for three airfoils are presented to compare their stall and post-stall behavior. The results for finite wings at stall and post-stall conditions focus on the effects of taper-ratio and sweep angle, with particular attention to whether the sectional flows can be approximated using two-dimensional flow over a stalled airfoil. While this approximation seems reasonable for unswept wings even at post-stall conditions, significant spanwise flow on stalled swept wings preclude the use of two-dimensional data to model sectional flows on swept wings. Thus, further effort is needed in low-order aerodynamic modeling of swept wings at stalled conditions.

Petrilli, Justin↗
NASA NTRS1980

Experimental determination of unsteady blade element aerodynamics in cascades. Volume 2: Translation mode cascade

A two dimensional cascade of harmonically oscillating airfoils was designed to model a near tip section from a rotor which was known to have experienced supersonic translational model flutter. This five bladed cascade had a solidity of 1.52 and a setting angle of 0.90 rad. Unique graphite epoxy airfoils were fabricated to achieve the realistic high reduced frequency level of 0.15. The cascade was tested over a range of static pressure ratios approximating the blade element operating conditions of the rotor along a constant speed line which penetrated the flutter boundary. The time steady and time unsteady flow field surrounding the center cascade airfoil were investigated.

Riffel, R. E.↗
NASA NTRS1979

Computation of subsonic and transonic flow about lifting rotor blades

An inviscid, nonconservative, three-dimensional potential flow code has been developed for computing the quasi-steady flow about an isolated lifting rotor blade. Calculations from the code were compared with chordwise pressure measurements obtained in a wind tunnel on a nonlifting rotor at transonic tip speeds at advance ratios from 0.40 to 0.55. The overall agreement between theoretical calculations and experiment was good. To illustrate the early capability of the program, the flow about a hypothetical lifting rotor blade having twist, airfoil thickness taper, and a 20 deg sweptback tip was analyzed at azimuthal positions of 60, 90, and 120 deg for an advance ratio of 0.342. A typical run on a CDC 7600 computer required about 5 min for one rotor position at transonic tip speeds.

Arieli, R.↗
NASA NTRS1980

Experimental determination of unsteady blade element aerodynamics in cascades. Volume 1: Torsion mode cascade

A two dimensional cascade of harmonically oscillating airfoils was designed to model a near tip section from a rotor which was known to have experienced supersonic torsional flutter. This five bladed cascade had a solidity of 1.17 and a setting angle of 1.07 rad. Graphite epoxy airfoils were fabricated to achieve the realistically high reduced frequency level of 0.44. The cascade was tested over a range of static pressure ratios approximating the blade element operating conditions of the rotor along a constant speed line which penetrated the flutter boundary. The time-steady and time-unsteady flow field surrounding the center cascade airfoil were investigated. The effects of reduced solidity and decreased setting angle on the flow field were also evaluated.

Riffel, R. E.↗
NASA NTRS1991

Aerodynamic characteristics of the HL-20 and HL-20A lifting-body configurations

The data show that the HL-20 is longitudinally and laterally stable over the test range from Mach 10 to 0.2. At hypersonic speeds it has a trimmed lift/drag ratio of 1.4. This values gives the vehicle a cross range capability similar to that of the Space Shuttle. At subsonic speeds, the HL-20 has a trimmed lift/drag ratio of about 3.6. Replacing the flat plate outboard fins with fins having an airfoil shape, increased the maximum trimmed L/D to 4.3. Preliminary evaluation of configuration modifications (the HL-20A series), indicates that trim at higher values of lift at hypersonic speeds could be achieved with an L/D of about 1.0. In the supersonic range, the lift and directional stability characteristics were improved. The untrimmed subsonic L/D was increased to 5.8 with airfoil fins.

Ware, George M.↗
NASA NTRS1943

The Effect of Compressibility on the Growth of the Laminar Boundary Layer on Low-Drag Wings and Bodies

The development of the laminar boundary layer in a compressible fluid is considered. Formulas are given for determining the boundary-layer thickness and the ratio of the boundary-layer Reynolds number to the body Reynolds number for airfoils and bodies of revolution. It i s shown that the effect of compressibility will profoundly alter the Reynolds number corresponding to the upper limit of the range of the low-drag coefficients . The available data indicate that for low-drag and high critical compressibility speed airfoils and bodies of revolution, this effect is favorable.

Allen, H. Julian↗
NASA NTRS1988

The effects of leading edge and downstream film cooling on turbine vane heat transfer

The progress under contract NAS3-24619 toward the goal of establishing a relevant data base for use in improving the predictive design capabilities for external heat transfer to turbine vanes, including the effect of downstream film cooling with and without leading edge showerhead film cooling. Experimental measurements were made in a two-dimensional cascade previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils under contract NAS3-22761 and leading edge showerhead film cooled airfoils under contract NAS3-23695. The principal independent parameters (Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio) were maintained over ranges consistent with actual engine conditions and the test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. Data provide a data base for downstream film cooled turbine vanes and extends the data bases generated in the two previous studies. The vane external heat transfer obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The data obtained and presented illustrate the interaction of the variables and should provide the airfoil designer and computational analyst the information required to improve heat transfer design capabilities for film cooled turbine airfoils.

Hylton, L. D.↗
NASA NTRS2003

Calculation of Rotor Performance and Loads Under Stalled Conditions

Rotor behavior in stalled conditions is investigated using wind tunnel test data of a l/l0-scale CH-47B/C type rotor, which provides a set of test conditions extending from unstalled to light stall to some deep stall conditions over a wide range of advance ratios. The rotor performance measured in the wind tunnel is similar to the main rotor performance measured during the NASA/Army UH-60A Airloads Program, although the two rotors are quite different. The analysis CAMRAD II has been used to predict the rotor performance and loads. Full-scale airfoil test data are corrected for Reynolds number effects for comparison with the model-scale rotor test. The calculated power coefficient shows good correlation with the measurements below stall with the Reynolds number-corrected airfoil table. Various dynamic stall models are used in the calculations. The Boeing model shows the lift augmentation at low advance ratios and the Leishman-Beddoes model shows better correlation of torsion moment than the other models at mu = 0.2. However, the dynamic stall models, in general, show only a small influence on the rotor power and torsion moment predictions especially at higher advance ratios.

Yeo, Hyeonsoo↗
NASA NTRS2000

The Surface Pressure Response of a NACA 0015 Airfoil Immersed in Grid Turbulence: Characteristics of the Turbulence - Volume 1

Two grids have been developed for the Virginia Tech 6 ft x 6 ft Stability wind tunnel for the purpose of generating homogeneous isotropic turbulent flows for the study of unsteady airfoil response. The first, a square bi-planar grid with a 12" mesh size and an open area ratio of 69.4%, was mounted in the wind tunnel contraction. The second grid, a metal weave with a 1.2 in. mesh size and an open area ratio of 68.2% was mounted in the tunnel test section. Detailed statistical and spectral measurements of the turbulence generated by the two grids are presented for wind tunnel free stream speeds of 10, 20, 30 and 40 m/s. These measurements show the flows to be closely homogeneous and isotropic. Both grids produce flows with a turbulence intensity of about 4% at the location planned for the airfoil leading edge. Turbulence produced by the large grid has an integral scale of some 3.2 inches here. Turbulence produced by the small grid is an order of magnitude smaller. For wavenumbers below the upper limit of the inertial subrange, the spectra and correlations measured with both grids at all speeds can be represented using the von Karman interpolation formula with a single velocity and length scale. The spectra maybe accurately represented over the entire wavenumber range by a modification of the von Karman interpolation formula that includes the effects of dissipation. These models are most accurate at the higher speeds (30 and 40 m/s).

Bereketab, Semere↗
NASA NTRS1996

Aerodynamic Limitations of the UH-60A Rotor

High quality airloads data have been obtained on an instrumented UH-60A in flight and these data provide insight into the aerodynamic limiting behavior of the rotor. At moderate weight coefficients and high advance ratio limiting performance is largely caused by high drag near the blade tip on the advancing side of the rotor as supercritical flow develops on the rotor with moderate to strong, shocks on both surfaces of the blade. Drag divergence data from two-dimensional airfoil tests show good agreement with the development of the supercritical flow regions. Large aerodynamic pitching moments are observed at high advance ratio, as well, and these pitching moments are the source of high torsional moments on the blade and control system loads. These loads occur on the advancing side of the disk and are not related to blade stall which does not occur for these weight coefficients. At high weight coefficients aerodynamic and structural limits are related to dynamic stall cycles that begin on the retreating side of the blade and, for the most severe conditions, carry around to the advancing side of the blade at the presumed first frequency of the blade/control system.

Coleman, Colin P.↗
NASA NTRS1978

Analytic investigation of advancing blade drag reduction by tip modifications

Analytic techniques were applied to study the effect on the performance of the nonlifting advancing blade when the outboard 5% of the blade is modified to reduce drag. The tip modifications studied consisted of reducing airfoil thickness, sweepback, and planform taper. The reductions in instantaneous drag and torque were calculated for tip speed ratios from about 0.19 to 0.30, corresponding to advancing blade tip Mach numbers of 0.855 to 0.936, respectively. Approximations required in the analysis introduce uncertainties into the computed absolute values of drag and torque; however, the differences in the quantities should be a fairly reliable measure of the effect of changing tip geometry. For example, at the highest tip speed, instantaneous drag, and torque were reduced by 20% and 24%, respectively, for tip sweep of 40 deg on a blade using an NACA 0010 airfoil and by comparable amounts for 30-deg sweep on a blade having an NACA 0012 airfoil section. The present method should prove to be a useful, inexpensive technique for identifying promising configurations for additional study and testing.

Tauber, M. E.↗
NASA NTRS1950

Extension of boundary-layer heat-transfer theory to cooled turbine blades

An equation for average heat transfer of a surface was derived when the boundary layer changed from laminar to turbulent. Influences on the heat transfer through a laminar boundary layer of Mach number, temperature ratio (gas temperature divided by wall temperature), and exponents of gas-property temperature relations were shown to be relatively small for air with Mach numbers less than 2 and temperature ratios between 1 and 4. Good agreement was obtained with experimental results from cylinders, an airfoil, and turbine blades.

HEAT TRANSFER, AERODYNAMIC↗
NASA NTRS1937

An analysis of the factors that determine the periodic twist of an autogiro rotor blade, with a comparison of predicted and measured results

Report presents an analysis of the factors that determine the periodic twist of a rotor blade under the action of the air forces on it. The results of the analysis show that the Fourier coefficients of the twist are linear expressions involving only the tip-speed ratio, the pitch setting, the inflow coefficient, the pitching-moment coefficient of the blade airfoil section, and the physical characteristics of the rotor blade and machine. The validity of the analysis was examined by using it to predict the twist of a rotor whose twist characteristics had previously been measured in flight. The agreement between the calculated and experimental results was satisfactory.

Wheatley, John B↗
NASA NTRS2011

Measurement of the Noise Resulting from the Interaction of Turbulence with a Lifting Surface

An experimental study of the noise resulting from the interaction of an airfoil with incident turbulence is presented. The test models include NACA0015 airfoils of different chord lengths, a flat plate with a sharp leading edge, and an airfoil of same section as a reference Fowler flap. The airfoils are immersed in nearly isotropic turbulence. Two approaches for performing the noise measurements are used and compared. The effects that turbulence intensity and scales, airfoil geometry, velocity and angle of attack have on the incident turbulence interaction noise are examined. Detailed directivity measurements are presented. It is found that noise spectral levels beyond the peak frequency decrease more with decreasing airfoil leading edge sharpness, and that spectral peak level (at 0 deg. angle of attack) appears to be mostly controlled by the airfoil fs thickness and chord. Increase in turbulence scale and intensity are observed to lead to a uniform increase of the noise spectral levels with an LI(sup 2) dependence (where L is the turbulence longitudinal integral scale and I is the turbulence intensity). Noise levels are found to scale with the 6th power of velocity and the 2nd power of the airfoil chord. Sensitivity to changes in angle of attack appears to have a turbulence longitudinal integral scale to chord (C) ratio dependence, with large effects on noise for L/C greater than or equal to 1 and decreased effects as L/C becomes smaller than 1. For all L/C values, the directivity pattern of the noise resulting from the incident turbulence is seen to remain symmetric with respect to the direction of the mean flow until stall, at which point, the directivity becomes symmetric with respect to the airfoil chord. It is also observed that sensitivity to angle of attack changes is more pronounced on the model suction side than on the model pressure side, and in the higher frequency range of the spectra for the largest airfoils tested (L/C less than 0.24).

Hutcheson, Florence V.↗
NASA NTRS2012

A Computational Study for the Utilization of Jet Pulsations in Gas Turbine Film Cooling and Flow Control

This report is the second part of a three-part final report of research performed under an NRA cooperative Agreement contract. The first part is NASA/CR-2012-217415. The third part is NASA/CR-2012-217417. Jets have been utilized in various turbomachinery applications in order to improve gas turbines performance. Jet pulsation is a promising technique because of the reduction in the amount of air removed from compressor. In this work two areas of pulsed jets applications were computationally investigated using the commercial code Fluent (ANSYS, Inc.); the first one is film cooling of High Pressure Turbine (HPT) blades and second one is flow separation control over Low Pressure Turbine (LPT) airfoil using Vortex Generator Jets (VGJ). Using pulsed jets for film cooling purposes can help to improve the effectiveness and thus allow higher turbine inlet temperature. Effects of the film hole geometry, blowing ratio and density ratio of the jet, pulsation frequency and duty cycle of blowing on the film cooling effectiveness were investigated. As for the low-pressure turbine (LPT) stages, the boundary layer separation on the suction side of airfoils can occur due to strong adverse pressure gradients. The problem is exacerbated as airfoil loading is increased. Active flow control could provide a means for minimizing separation under conditions where it is most severe (low Reynolds number), without causing additional losses under other conditions (high Reynolds number). The effects of the jet geometry, blowing ratio, density ratio, pulsation frequency and duty cycle on the size of the separated region were examined in this work. The results from Reynolds Averaged Navier-Stokes and Large Eddy Simulation computational approaches were compared with the experimental data.

Kartuzova, Olga V.↗
NASA NTRS1979

General Theory of Aerodynamic Instability and the Mechanism of Flutter

The aerodynamic forces on an oscillating airfoil or airfoil-aileron combination of three independent degrees of freedom were determined. The problem resolves itself into the solution of certain definite integrals, which were identified as Bessel functions of the first and second kind, and of zero and first order. The theory, based on potential flow and the Kutta condition, is fundamentally equivalent to the conventional wing section theory relating to the steady case. The air forces being known, the mechanism of aerodynamic instability was analyzed. An exact solution, involving potential flow and the adoption of the Kutta condition, was derived. The solution is of a simple form and is expressed by means of an auxiliary parameter k. The flutter velocity, treated as the unknown quantity, was determined as a function of a certain ratio of the frequencies in the separate degrees of freedom for any magnitudes and combinations of the airfoil-aileron parameters.

Theodorsen, Theodore↗