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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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Wind-tunnel experiments on divergence of forward-swept wings

An experimental study to investigate the aeroelastic behavior of forward-swept wings was conducted in the Langley Transonic Dynamics Tunnel. Seven flat-plate models with varying aspect ratios and wing sweep angles were tested at low speeds in air. Three models having the same planform but different airfoil sections (i.e., flat-plate, conventional, and supercritical) were tested at transonic speeds in Freon 12. Linear analyses were performed to provide predictions to compare with the measured aeroelastic instabilities which include both static divergence and flutter. Six subcritical response testing techniques were formulated and evaluated at transonic speeds for accuracy in predicting static divergence. Two "divergence stoppers" were developed and evaluated for use in protecting the model from structural damage during tests.

Ricketts, R. H.↗
NASA NTRS1988

Prediction of oscillating thick cambered aerofoil aerodynamics by a locally analytic method

The effects of mean-flow incidence, airfoil camber, and airfoil thickness on the incompressible aerodynamics of an oscillating airfoil are investigated theoretically, developing and applying a first-order FEM based on locally analytical solutions (LASs). Laplace equations are used to describe the steady and unsteady harmonic velocity potentials; a body-fitted computational grid is employed; grid-element solutions for both potentials are determined using a numerical LAS method; and the LASs are then assembled to obtain a complete solution. Results for a series of flat-plate and Joukowski airfoils are presented in extensive graphs and discussed in detail.

Chiang, Hsiao-Wei D.↗
NASA NTRS2000

Validation of a High-Order Prefactored Compact Scheme on Nonlinear Flows with Complex Geometries

Three benchmark problems are solved using a sixth-order prefactored compact scheme employing an explicit 10th-order filter with optimized fourth-order Runge-Kutta time stepping. The problems solved are the following: (1) propagation of sound waves through a transonic nozzle; (2) shock-sound interaction; and (3) single airfoil gust response. In the first two problems, the spatial accuracy of the scheme is tested on a stretched grid, and the effectiveness of boundary conditions is shown. The solution stability and accuracy near a shock discontinuity is shown as well. Also, 1-D nonlinear characteristic boundary conditions will be evaluated. In the third problem, a nonlinear Euler solver will be used that solves the equations in generalized curvilinear coordinates using the chain rule transformation. This work, continuing earlier work on flat-plate cascades and Joukowski airfoils, will focus mainly on the effect of the grid and boundary conditions on the accuracy of the solution. The grids were generated using a commercially available grid generator, GridPro/az3000.

Hixon, Ray↗
NASA NTRS1993

Subsonic aerodynamic characteristics of the HL-20 lifting-body configuration

The HL-20 is proposed as a possible future manned spacecraft. The configuration consists of a low-aspect-ratio body with a flat undersurface. Three fins (a small centerline fin and two outboard (tip) fins set at a dihedral angle of 50 deg) are mounted on the aft body. The control system consists of elevon surfaces on the outboard fins, a set of four body flaps on the upper and lower aft body, and an all-movable center fin. Both the elevons and body flaps were capable of trimming the model to angles of attack from -2 deg to above 20 deg. The maximum trimmed lift-drag ratio was 3.6. Replacing the flat-plate tip fins with airfoil tip fins increased the maximum trimmed lift-drag ratio to 4.2. The elevons were effective as a roll control, but they produced about as much yawing moment as rolling moment because of the tip-fin dihedral angle. The body flaps produced less rolling moment than the elevons and only small values of yawing moment. A limited investigation of the effect of varying tip-fin dihedral angle indicated that a dihedral angle of 50 deg was a reasonable compromise for longitudinal and lateral stability, longitudinal trim, and performance at subsonic speeds.

Ware, George M.↗
NASA NTRS1985

A review and analysis of boundary layer transition data for turbine application

A number of data sets from the open literature that include heat transfer data in apparently transitional boundary layers, with particular application to the turbine environment, were reviewed and analyzed to extract transition information. The data were analyzed by using a version of the STAN5 two-dimensional boundary layer code. The transition starting and ending points were determined by adjusting parameters in STAN5 until the calculations matched the data. The results are presented as a table of the deduced transition location and length as functions of the test parameters. The data sets reviewed cover a wide range of flow conditions, from low-speed, flat-plate tests to full-scale turbine airfoils operating at simulated turbine engine conditions. The results indicate that free-stream turbulence and pressure gradient have strong, and opposite, effects on the location of the start of transition and on the length of the transition zone.

Gaugler, R. E.↗
NASA NTRS1985

Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfroils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Shaw, L. M.↗
NASA NTRS1986

Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfols were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a clasical usnteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Shaw, L. M.↗
NASA NTRS1945

Wind-Tunnel Investigation of a Rectangular NACA 2212 Airfoil with Semispan Ailerons and with Nonperforated, Balanced Double Split Flaps for Use as Aerodynamic Brakes

Flat-plate flaps with no wing cutouts and flaps having Clark Y sections with corresponding cutouts made in wing were tested for various flap deflections, chord-wise locations, and gaps between flaps and airfoil contour. The drag was slightly lower for wing with airfoil section flaps. Satisfactory aileron effectiveness was obtained with flap gap of 20% wing chord and flap-nose location of 80 percent wing chord behind leading edge. Airflow was smooth and buffeting negligible.

AIRFOILS - NACA 2212 - CHAR & MOMENTS - WING FLAPS↗
NASA NTRS1989

Multigrid solution of the Navier-Stokes equations on triangular meshes

A Navier-Stokes algorithm for use on unstructured triangular meshes is presented. Spatial discretization of the governing equations is achieved using a finite element Galerkin approximation, which can be shown to be equivalent to a finite volume approximation for regular equilateral triangular meshes. Integration steady-state is performed using a multistage time-stepping scheme, and convergence is accelerated by means of implicit residual smoothing and an unstructured multigrid algorithm. Directional scaling of the artificial dissipation and the implicit residual smoothing operator is achieved for unstructured meshes by considering local mesh stretching vectors at each point. The accuracy of the scheme for highly stretched triangular meshes is validated by comparing computed flat-plate laminar boundary layer results with the well known similarity solution, and by comparing laminar airfoil results with those obtained from various well-established structured quadrilateral-mesh codes. The convergence efficiency of the present method is also shown to be competitive with those demonstrated by structured quadrilateral-mesh algorithms.

Mavriplis, Dimitri J.↗
NASA NTRS1990

Multigrid solution of the Navier-Stokes equations on triangular meshes

A Navier-Stokes algorithm for use on unstructured triangular meshes is presented. Spatial discretization of the governing equations is achieved using a finite element Galerkin approximation, which can be shown to be equivalent to a finite volume approximation for regular equilateral triangular meshes. Integration steady-state is performed using a multistage time-stepping scheme, and convergence is accelerated by means of implicit residual smoothing and an unstructured multigrid algorithm. Directional scaling of the artificial dissipation and the implicit residual smoothing operator is achieved for unstructured meshes by considering local mesh stretching vectors at each point. The accuracy of the scheme for highly stretched triangular meshes is validated by comparing computed flat-plate laminar boundary layer results with the well known similarity solution, and by comparing laminar airfoil results with those obtained from various well-established structured quadrilateral-mesh codes. The convergence efficiency of the present method is also shown to be competitive with those demonstrated by structured quadrilateral-mesh algorithms.

Mavriplis, D. J.↗
NASA NTRS1994

Localized and distributed boundary-layer receptivity to convected unsteady wake in free stream

Receptivity to a model convected disturbance in the presence of localized and distributed variations in wall geometry and wall-suction velocity is examined. The model free-stream disturbance corresponds to the time-harmonic wake of a vibrating ribbon that is placed at a suitable distance above the surface of a thin airfoil. The advantages of using this disturbance for experiments on receptivity to convected disturbances are outlined. A brief parametric study is presented for a flat-plate boundary layer. The study quantifies the effect of wake position as well as wake width; in addition, it should be helpful in the choice of an optimal setting for a controlled experiment of the above type, which the above parametric study shows as feasible.

Choudhari, Meelan↗
NASA NTRS1986

On the application of a linearized unsteady potential-flow analysis to fan-tip cascades

A linearized potential flow analysis, which accounts for the effects of nonuniform steady flow phenomena on the unsteady response to prescribed blade motions, has been applied to five two-dimensional cascade configurations. These include a flat-plate cascade and three cascades which are representative of the tip sections of current fan designs. Here the blades are closely spaced, highly staggered, and operate at low mean incidence. The fifth configuration is a NASA Lewis cascade of symmetric biconvex airfoils for which experimental measurements are available. Numerical solutions are presented that clearly illustrate the effects and importance of blade geometry and mean blade loading on the linearized unsteady response at high subsonic inlet Mach number and high blade-vibrational frequency. In addition, a good qualitative agreement is shown between the analytical predictions and experimental measurements for the cascade of symmetric biconvex airfoils. Finally, recommendations on the research needed to extend the range of application of linearized unsteady aerodynamic analyses are provided.

Usab, W. J., Jr.↗
NASA NTRS1988

Aerodynamic characteristics of wings designed with a combined-theory method to cruise at a Mach number of 4.5

A wind-tunnel study was conducted to determine the capability of a method combining linear theory and shock-expansion theory to design optimum camber surfaces for wings that will fly at high-supersonic/low-hypersonic speeds. Three force models (a flat-plate reference wing and two cambered and twisted wings) were used to obtain aerodynamic lift, drag, and pitching-moment data. A fourth pressure-orifice model was used to obtain surface-pressure data. All four wing models had the same planform, airfoil section, and centerbody area distribution. The design Mach number was 4.5, but data were also obtained at Mach numbers of 3.5 and 4.0. Results of these tests indicated that the use of airfoil thickness as a theoretical optimum, camber-surface design constraint did not improve the aerodynamic efficiency or performance of a wing as compared with a wing that was designed with a zero-thickness airfoil (linear-theory) constraint.

Mack, Robert J.↗
NASA NTRS1961

Considerations of Methods of Improving Helicopter Efficiency

Recent NASA helicopter research indicates that significant improvements in hovering efficiency, up to 7 percent, are available from the use of a special airfoil section formed by combining an NACA 632A015 thickness distribution with an NACA 230 mean line. This airfoil should be considered for flying-crane-type helicopters. Application of standard leading-edge roughness causes a large drop in efficiency; however, the cambered rotor is shown to retain its superiority over a rotor having a symmetrical airfoil when both rotors have leading-edge roughness. A simple analysis of available rotor static-thrust data indicates a greatly reduced effect of compressibility effects on the rotor profile-drag power than predicted from calculations. Preliminary results of an experimental study of helicopter parasite drag indicate the practicability of achieving an equivalent flat-plate parasite-drag area of less than 4 square feet for a rotor-head-pylon-fuselage configuration (landing gear retracted) in the 2,000-pound minimum-flying-weight class. The large drag penalty of a conventional skid-type landing (3.6 square feet) can be reduced by two-thirds by careful design. Clean, fair, and smooth fuselages that tend to have narrow, deep cross sections are shown to have advantages from the standpoint of drag and download. A ferry range of the order of 1,500 miles is indicated to be practicable for the small helicopter considered.

Dingeldein, Richard C.↗
NASA NTRS1959

Analysis and Evaluation of Supersonic Underwing Heat Addition

The linearized theory for heat addition under a wing has been developed to optimize wing geometry, heat addition, and angle of attack. The optimum wing has all of the thickness on the underside of the airfoil, with maximum-thickness point well downstream, has a moderate thickness ratio, and operates at an optimum angle of attack. The heat addition is confined between the fore Mach waves from under the trailing surface of the wing. By linearized theory, a wing at optimum angle of attack may have a range efficiency about twice that of a wing at zero angle of attack. More rigorous calculations using the method of characteristics for particular flow models were made for heating under a flat-plate wing and for several wings with thickness, both with heat additions concentrated near the wing. The more rigorous calculations yield in practical cases efficiencies about half those estimated by linear theory. An analysis indicates that distributing the heat addition between the fore waves from the undertrailing portion of the wing is a way of improving the performance, and further calculations appear desirable. A comparison of the conventional ramjet-plus wing with underwing heat addition when the heat addition is concentrated near the wing shows the ramjet to be superior on a range basis up to Mach number of about B. The heat distribution under the wing and the assumed ramjet and airframe performance may have a marked effect on this conclusion. Underwing heat addition can be useful in providing high-altitude maneuver capability at high flight Mach numbers for an airplane powered by conventional ramjets during cruise.

wings↗
NASA NTRS1959

Flutter Tests of Some Simple Models at a Mach Number of 7.2 in Helium Flow

Results of hypersonic flutter tests on some simple models are presented. The models had rectangular plan forms of panel aspect ratio 1.0, no sweepback, and bending-to-torsion frequency ratios of about 1/3. Two airfoil sections were included in the tests; double wedges of 5-, 10-, and 15-percent thickness and flat plates with straight, parallel sides and beveled leading and trailing edges. The models were supported by a cantilevered shaft. The double-wedge wings were tested in helium at a Mach number of 7.2. An effect of airfoil thickness on flutter speed was found, thicker wings requiring more stiffness to avoid flutter. A few tests in air at a Mach number of 6.9 showed the same thickness effect and also indicated that tests in helium would predict conservative flutter boundaries in air. The data in air and helium seemed to be correlated by piston-theory calculations. Piston-theory calculations agreed well with experiment for the thinner models but began to deviate as the thickness parameter MT approached and exceeded 1.0. A few tests on flat-plate models with various elastic-axis locations were made. Piston-theory calculations would not satisfactorily predict the flutter of these models, probably because of their blunt leading edges.

Morgan, Homer G.↗
NASA NTRS1995

An experimental study of a turbulent boundary layer in the trailing edge region of a circulation-control airfoil

This report describes the development of a new near-wall LDV technique and its use to measure the full turbulent stress tensor in a highly three-dimensional separated boundary layer flow. Laser Doppler velocimetry, with small diameter laser beams allowing small beam-crossing angles (i.e., larger fringe spacings), was used, resulting in measurements close to solid surfaces. Dennis Johnson of NASA/Ames combined the advantages offered by FFT signal processors with a beam-steering probe inserted in the flow to devise a new self-aligning, near-wall LDV approach capable of making direct measurements of the transverse velocity component. This new technique was applied to a 2D, zero-pressure-gradient, flat-plate boundary layer in the High Reynolds Number Pilot Channel at Ames. Accurate and repeatable measurements of U and uu to within 30 mm of the wall were obtained. Next, the near-wall LDV technique was employed in a 3D separated boundary layer flow, consisting of a 2-dimensional wall boundary layer approaching a right-angle wedge. The resulting lateral and adverse streamwise pressure gradients generated significant spanwise velocities and turbulent stresses, and a large reverse-flow region. The near-wall LDV was modified further to facilitate measurements in the 3D Wedge Flow.

Brown, Jeffrey↗