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At least 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRS2004

Navier-Stokes Computations of Longitudinal Forces and Moments for a Blended Wing Body

The object of this paper is to investigate the feasibility of applying CFD methods to aerodynamic analyses for aircraft stability and control. The integrated aerodynamic parameters used in stability and control, however, are not necessarily those extensively validated in the state of the art CFD technology. Hence, an exploratory study of such applications and the comparison of the solutions to available experimental data will help to assess the validity of the current computation methods. In addition, this study will also examine issues related to wind tunnel measurements such as measurement uncertainty and support interference effects. Several sets of experimental data from the NASA Langley 14x22-Foot Subsonic Tunnel and the National Transonic Facility are presented. Two Navier-Stokes flow solvers, one using structured meshes and the other unstructured meshes, were used to compute longitudinal static stability derivatives for an advanced Blended Wing Body configuration over a wide range of angles of attack. The computations were performed for two different Reynolds numbers and the resulting forces and moments are compared with the above mentioned wind tunnel data.

Pao, S. Paul↗
NASA NTRS2005

Navier-Stokes Computations of Longitudinal Forces and Moments for a Blended Wing Body

The object of this paper is to investigate the feasibility of applying CFD methods to aerodynamic analyses for aircraft stability and control. The integrated aerodynamic parameters used in stability and control, however, are not necessarily those extensively validated in the state of the art CFD technology. Hence, an exploratory study of such applications and the comparison of the solutions to available experimental data will help to assess the validity of the current computation methods. In addition, this study will also examine issues related to wind tunnel measurements such as measurement uncertainty and support interference effects. Several sets of experimental data from the NASA Langley 14x22-Foot Subsonic Tunnel and the National Transonic Facility are presented. Two Navier-Stokes flow solvers, one using structured meshes and the other unstructured meshes, were used to compute longitudinal static stability derivatives for an advanced Blended Wing Body configuration over a wide range of angles of attack. The computations were performed for two different Reynolds numbers and the resulting forces and moments are compared with the above mentioned wind tunnel data.

Pao, S. Paul↗
NASA NTRS1985

Wind-tunnel evaluation of a 21-percent-scale powered model of a prototype advanced scout helicopter

An exploratory wind tunnel investigation of a 21 percent scale powered model of a prototype advanced scout helicopter was conducted in the Langley 4 by 7 Meter Tunnel. The investigation was conducted to define the overall aerodynamic characteristics of the Army Helicopter Improvement Program (AHIP), to determine the effects of the rotor on the aerodynamic characteristics and to evaluate the effect of a mast mounted sight on the aircraft stability characteristics. Tests covered a range of thrust coefficients, advance ratios, angles of attack and angles of sideslip and were run for both rotor on and rotor off configurations. Results of the investigation showed that the prototype configuration was longitudinally unstable with angle of attack for all configurations tested. The instability was due to unfavorable interference effects between the horizontal tail and the wake shed from the engine pylon and rotor hub, which caused a loss of horizontal tail effectiveness. The addition of the mast mounted sight had little effect on the stability of the model, but it caused an alteration in the rotor lift distribution that resulted in substantial interference drag for the sight.

Phelps, A. E., III↗
NASA NTRS1990

Evaluation of an adaptive unstructured remeshing technique for integrated fluid-thermal-structural analysis

An adaptive unstructured remeshing technique is evaluated for integrated fluid-thermal-structural analysis. The technique is combined with the finite element method to solve: (1) the Navier-Stokes equations for high-speed compressible flow; (2) the energy equation for the structural-thermal response; and (3) the quasi-static equilibrium equations for the structural response. The remeshing technique and the analysis solution procedure are described. The effectiveness of the approach is evaluated with two application studies. The flow analysis of Mach 8 shock-shock interference on a three-inch-diameter cylinder is used as the first application study to demonstrate the capability of the remeshing technique and to examine proper remeshing indicators for the inviscid and boundary layer regions. The applicability of the approach for the thermal and structural analyses of the structure is evaluated in the second application study of a 0.25-inch-diameter convectively cooled leading edge subjected to intense aerodynamic heating. Issues associated with remeshing indicators for thermal stress problems are identified.

Dechaumphai, Pramote↗
NASA NTRS1990

Evaluation of an adaptive unstructured remeshing technique for integrated fluid-thermal-structural analysis

An adaptive unstructured remeshing technique is evaluated for integrated fluid-thermal-structural analysis. The technique is combined with the finite element method to solve: (1) the Navier-Stokes equations for high-speed compressible flow, (2) the energy equation for the structural-thermal response, and (3) the quasi-static equilibrium equations for the structural response. The remeshing technique and the analysis solution procedure are described. The effectiveness of the approach is evaluated with two applications studies. The flow analysis of Mach 8 shock-shock interference on a three-inch-diameter cylinder is used as the first application study to demonstrate the capability of the remeshing technique and to examine proper remeshing indicators for the inviscid and boundary layer regions. The applicability of the approach for the thermal and structural analyses of the structure is evaluated in the second application study of a 0.25-inch-diameter convectively cooled leading edge subjected to intense aerodynamic heating. Issues associated with remeshing indicators for thermal stress problems are identified.

Dechaumphai, Pramote↗
NASA NTRS1981

Inviscid transonic flow over axisymmetric bodies

Axisymmetric transonic flow is of interest not only because of its practical application to missile and launch vehicle aerodynamics but also because of its relation, in terms of area rule, to fully three dimensional flow. RAXBOD computer program analyzes steady, inviscid, irrotational, transonic flow over axisymmetric bodies in free air. RAXBOD uses finite-difference relaxation method to solve numerically exact formulation of disturbance velocity potential with exact surface boundary conditions. Agreement with available experimental results has been good in cases where viscous effects and wind-tunnel wall interference are not important.

South, J. C., Jr.↗
NASA NTRS1992

Aircraft Drag Reduction

This paper reviews aerodynamic drag reduction for friction, wave and vortex drag associated with supersonic cruise aircraft and suggests approaches and research directions which cover the spectrum from possibly significant to revolutionary. Various synergisms are also included, especially in regard to suction laminar flow control and flow separation control at cruise. The former may also enable improved low speed, high lift systems, lift to drag ratio for subsonic cruise, reduced parasitic viscous drag for favorable interference wave drag reduction approaches, and turbulent skin friction reduction via slot injection. Flow separation control at cruise proffers opportunities for increased leading edge thrust, increased lift increment from upper surface, increased fuselage lift/camber for wave drag due to lift reduction, improved performance of various favorable interference wave drag reduction schemes, as well as possibly better low speed, high-schemes and wing cruise performance. More speculative approaches include multi-stage aircraft, thrust vectoring for lift augmentation, as well as trim, and a number of conceptual wave and vortex drag reduction schemes.

D M Bushnell↗
NASA NTRS1994

Plasma heating for containerless and microgravity materials processing

A method for plasma heating of levitated samples to be used in containerless microgravity processing is disclosed. A sample is levitated by electrostatic, electromagnetic, aerodynamic, or acoustic systems, as is appropriate for the physical properties of the particular sample. The sample is heated by a plasma torch at atmospheric pressure. A ground plate is provided to help direct the plasma towards the sample. In addition, Helmholtz coils are provided to produce a magnetic field that can be used to spiral the plasma around the sample. The plasma heating system is oriented such that it does not interfere with the levitation system.

Leung, Emily W.↗
NASA NTRS1973

An analytical procedure for evaluating shuttle abort staging aerodynamic characteristics

An engineering analysis and computer code (AERSEP) for predicting Space Shuttle Orbiter - HO Tank longitudinal aerodynamic characteristics during abort separation has been developed. Computed results are applicable at Mach numbers above 2 for angle-of-attack between plus or minus 10 degrees. No practical restrictions on orbiter-tank relative positioning are indicated for tank-under-orbiter configurations. Input data requirements and computer running times are minimal facilitating program use for parametric studies, test planning, and trajectory analysis. In a majority of cases AERSEP Orbiter-Tank interference predictions are as accurate as state-of-the-art estimates for interference-free or isolated-vehicle configurations. AERSEP isolated-orbiter predictions also show excellent correlation with data.

Meyer, R.↗
NASA NTRS2007

Development of a Rayleigh Scattering Diagnostic for Time-Resolved Gas Flow Velocity, Temperature, and Density Measurements in Aerodynamic Test Facilities

A molecular Rayleigh scattering technique is developed to measure time-resolved gas velocity, temperature, and density in unseeded turbulent flows at sampling rates up to 32 kHz. A high power continuous-wave laser beam is focused at a point in an air flow field and Rayleigh scattered light is collected and fiber-optically transmitted to the spectral analysis and detection equipment. The spectrum of the light, which contains information about the temperature and velocity of the flow, is analyzed using a Fabry-Perot interferometer. Photomultiplier tubes operated in the photon counting mode allow high frequency sampling of the circular interference pattern to provide time-resolved flow property measurements. An acoustically driven nozzle flow is studied to validate velocity fluctuation measurements, and an asymmetric oscillating counterflow with unequal enthalpies is studied to validate the measurement of temperature fluctuations. Velocity fluctuations are compared with constant temperature anemometry measurements and temperature fluctuations are compared with constant current anemometry measurements at the same locations. Time-series and power spectra of the temperature and velocity measurements are presented. A numerical simulation of the light scattering and detection process was developed and compared with experimental data for future use as an experiment design tool.

Mielke, Amy F.↗
NASA NTRS2007

Performance of Advanced Heavy-Lift, High-Speed Rotorcraft Configurations

The aerodynamic performance of rotorcraft designed for heavy-lift and high-speed cruise is examined. Configurations considered include the tiltrotor, the compound helicopter, and the lift-offset rotor. Design conditions are hover and 250-350 knot cruise, at 5k/ISA+20oC (civil) or 4k/95oF (military); with cruise conditions at 4000 or 30,000 ft. The performance was calculated using the comprehensive analysis CAMRAD II, emphasizing rotor optimization and performance, including wing-rotor interference. Aircraft performance was calculated using estimates of the aircraft drag and auxiliary propulsion efficiency. The performance metric is total power, in terms of equivalent aircraft lift-to-drag ratio L/D = WV/P for cruise, and figure of merit for hover.

Johnson, Wayne↗
NASA NTRS1991

Rotorcraft application of advanced computational aerodynamics

The objective was to develop the capability to compute the unsteady viscous flow around rotor-body combinations. In the interest of tractability, the problem was divided into subprograms for: (1) computing the flow around a rotor blade in isolation; (2) computing the flow around a fuselage in isolation, and (3) integrating the pieces. Considerable progress has already been made by others toward computing the rotor in isolation (Srinivasen) and this work focused on the remaining tasks. These tasks required formulating a multi-block strategy for combining rotating blades and nonrotating components (i.e., a fuselage). Then an appropriate configuration was chosen for which suitable rotor body interference test data exists. Next, surface and volume grids were generated and state-of-the-art CFD codes were modified and applied to the problem.

Stanaway, Sharon↗
NASA NTRS1974

Theoretical prediction of thick wing and pylon-fuselage-fanpod-nacelle aerodynamic characteristics at subcritical speeds. Part 1: Theory and results

The theoretical development and the comparison of results with data of a thick wing and pylon-fuselage-fanpod-nacelle analysis are presented. The analysis utilizes potential flow theory to compute the surface velocities and pressures, section lift and center of pressure, and the total configuration lift, moment, and vortex drag. The skin friction drag is also estimated in the analysis. The perturbation velocities induced by the wing and pylon, fuselage and fanpod, and nacelle are represented by source and vortex lattices, quadrilateral vortices, and source frustums, respectively. The strengths of these singularities are solved for simultaneously including all interference effects. The wing and pylon planforms, twists, cambers, and thickness distributions, and the fuselage and fanpod geometries can be arbitrary in shape, provided the surface gradients are smooth. The flow through nacelle is assumed to be axisymmetric. An axisymmetric center engine hub can also be included. The pylon and nacelle can be attached to the wing, fuselage, or fanpod.

Tulinius, J. R.↗
NASA NTRS1965

Nasa research on the aerodynamics of jet vtol engine installations

This paper summarizes some of the more pertinent results of NASA investigations related to the aerodynamics of jet VTOL engine installations. It shows that there is a base loss in hovering due to suction forces creased on the underside of the fuselage by the entrainment of ambient air in the slipstream, and that the magnitude of this effect is related to the turbulence in the jet stream and its consequent rate of mixing with ambient air. It also shows that there are large lift losses and pitching moments due to jet-free-stream interference and that these characteristics can be significantly altered by proximity to the ground. And, finally, it shows that simple bellmouth inlets give good pressure recovery and low distortion for vertically mounted lift engines if the inlet lip radius is sufficiently large, but that such inlets are not suitable for windmill starting of the engines.

VTOL AIRCRAFT↗
NASA NTRS1936

Spinning Characteristics of Wings I : Rectangular Clark Y Monoplane Wing

A series of wind tunnel tests of a rectangular Clark Y wing was made with the NACA spinning balance as part of a general program of research on airplane spinning. All six components of the aerodynamic force and moment were measured throughout the range of angles of attack, angles of sideslip, and values omega b/2v likely to be attained by a spinning airplane; the results were reduced to coefficient form. It is concluded that a conventional monoplane with a rectangular Clark y wing can be made to attain spinning equilibrium throughout a wide range of angles of attack but that provision of a yawing moment coefficient of -0.02 (against the spin) by the tail, fuselage, and interferences will insure against attainment of equilibrium in a steady spin.

Bamber, M J↗
NASA NTRS1975

The self streamlining wind tunnel

A two dimensional test section in a low speed wind tunnel capable of producing flow conditions free from wall interference is presented. Flexible top and bottom walls, and rigid sidewalls from which models were mounted spanning the tunnel are shown. All walls were unperforated, and the flexible walls were positioned by screw jacks. To eliminate wall interference, the wind tunnel itself supplied the information required in the streamlining process, when run with the model present. Measurements taken at the flexible walls were used by the tunnels computer check wall contours. Suitable adjustments based on streamlining criteria were then suggested by the computer. The streamlining criterion adopted when generating infinite flowfield conditions was a matching of static pressures in the test section at a wall with pressures computed for an imaginary inviscid flowfield passing over the outside of the same wall. Aerodynamic data taken on a cylindrical model operating under high blockage conditions are presented to illustrate the operation of the tunnel in its various modes.

Goodyer, M. J.↗
NASA NTRS1992

Unsteady flows

The development status of tactical missile dynamic behavior prediction is evaluated, with a view to the formulation of guidelines for high angle-of-attack tests where highly nonlinear aerodynamics are encountered and strong cross-coupling arises between lateral and longitudinal degrees of freedom. Attention is given to the Ericsson (1979) inviscid computer code, which is entirely general within the geometric constraints of axisymmetry and low aspect ratio lifting surfaces. A serious problem in the prediction of full-scale high-alpha dynamics is the almost total dependence on subscale experimental data, whose use is severely restricted by Reynolds number scaling as well as by model support and wind tunnel wall interference.

Ericsson, L. E.↗
NASA NTRS1972

Space shuttle: Directional and lateral stability and interference effects of cruise engine location on a 0.015 scale space shuttle

The results are presented that were obtained from a wind tunnel tests to improve space shuttle booster baseline lateral-directional stability, control characteristics, and cruise engine location optimization. Tests were conducted in a 7 x 10-foot transonic wind tunnel. The model employed was a 0.015-scale replica of a space shuttle booster. The three major objectives of this test were to determine the following: (1) force, static stability, and control effectiveness characteristics for varying angles of positive and negative wing dihedral and various combinations of wing tip and centerline dorsal fins; (2) force and static stability characteristics of cruise engines location on the body below the high aerodynamic canard; and (3) control effectiveness for the low-mounted wing configuration. The wing dihedral study was conducted at a cruise Mach number of 0.40 and simulated altitude of 10,000 feet. Portions of the test were conducted to determine the control surfaces stability and control characteristics over the Mach number range of 0.4 to 1.2. The aerodynamic characteristics determined are based on a unit Reynolds number of approximately 2 million per foot. Boundary layer trip strips were employed to induce boundary layer transition.

Buchholz, R. E.↗