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At least 451 records · Page 25

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NASA NTRS1972

Satellite Aerodynamics and Density Determination from Satellite Dynamic Response

The aerodynamic drag and lift properties of a satellite are first expressed as a function of two parameters associated with gas-surface interaction at the satellite surface. The dynamic response of the satellite as it passes through the atmosphere is then expressed as a function of the two gas-surface interaction parameters, the atmospheric density, the satellite velocity, and the satellite orientation to the high speed flow. By proper correlation of the observed dynamic response with the changing angle of attack of the satellite, it is found that the two unknown gas-surface interaction parameters can be determined. Once the gas-surface interaction parameters are known, the aerodynamic properties of the satellite at all angles of attack are also determined.

Karr, G. R.↗

NASA NTRS1974

Results of investigations (OA20A) on a 0.015-scale 140A/B configuration space shuttle vehicle orbiter model in the NASA/Langley Research Center unitary plan wind tunnel

Data obtained from a wind tunnel test of a 0.015-scale 140A/B configuration SSV orbiter model are presented. This test was conducted with runs at Mach numbers of 2.5, 3.9, and 4.6 for a constant Reynolds number of 2.5 x 1,000,000/foot. Only one model configuration, the complete 140A/B orbiter vehicle, was investigated; various control-surface settings were run through angles-of-attack from -4 to +42 degrees at 0 and +3 degrees of yaw and through angles-of-sideslip from -4 to +6 degrees at 0, +10, +20, and +30 degrees pitch. The purpose of this test was to establish and verify longitudinal and lateral-directional stability and control characteristics for the updated SSV configuration.

Nichols, M. E.↗

NASA NTRS1976

Design of an all-attitude flight control system to execute commanded bank angles and angles of attack

A flight control system for use in air-to-air combat simulation was designed. The input to the flight control system are commanded bank angle and angle of attack, the output are commands to the control surface actuators such that the commanded values will be achieved in near minimum time and sideslip is controlled to remain small. For the longitudinal direction, a conventional linear control system with gains scheduled as a function of dynamic pressure is employed. For the lateral direction, a novel control system, consisting of a linear portion for small bank angle errors and a bang-bang control system for large errors and error rates is employed.

Burgin, G. H.↗

NASA NTRS1979

Leading-edge slat optimization for maximum airfoil lift

A numerical procedure for determining the position (horizontal location, vertical location, and deflection) of a leading edge slat that maximizes the lift of multielement airfoils is presented. The structure of the flow field is calculated by iteratively coupling potential flow and boundary layer analysis. This aerodynamic calculation is combined with a constrained function minimization analysis to determine the position of a leading edge slat so that the suction peak on the nose of the main airfoil is minized. The slat position is constrained by the numerical procedure to ensure an attached boundary layer on the upper surface of the slat and to ensure negligible interaction between the slat wake and the boundary layer on the upper surface of the main airfoil. The highest angle attack at which this optimized slat position can maintain attached flow on the main airfoil defines the optimum slat position for maximum lift. The design method is demonstrated for an airfoil equipped with a leading-edge slat and a trailing edge, single-slotted flap. The theoretical results are compared with experimental data, obtained in the Ames 40 by 80 Foot Wind Tunnel, to verify experimentally the predicted slat position for maximum lift. The experimentally optimized slat position is in good agreement with the theoretical prediction, indicating that the theoretical procedure is a feasible design method.

Olson, L. E.↗

NASA NTRS1980

Experimental and computational study of transonic flow about swept wings

An experimental investigation of NACA 0010 and 10% circular arc wing models, swept at 45 deg, spanning a channel, and at zero angle of attack is described. Measurements include chordwise and spanwise surface pressure distributions and oil-flow patterns for a range of transonic Mach numbers and Reynolds numbers. Calculations using a new three-dimensional Navier-Stokes code and a two-equation turbulence model are included for the circular-arc wing flow. Reasonable agreement between measurements and computations is obtained.

Bertelrud, A.↗

NASA NTRS1981

An experimental study of the flow in a wing with a partial span drooped leading edge

The flow field produced by a low aspect ratio wing (AR = 3.0) with a partial span leading edge droop was investigated in a series of low speed wind tunnel tests (Reynolds number based on 17.8 cm chord = 560,000). Photographs were obtained of surface oil flow patterns over an angle of attack range of alpha = 0 to 29 deg. Flow field surveys of the partially stalled wing at alpha = 25 deg were completed using a hot-wire probe, a split-film probe and a Conrad probe. The flow field survey data was presented using a color video display. The data indicated regions of apparent reversed flow in the separation region behind the wing and indicated the general cross-sectional shape of the separated wake flow.

Winkelmann, A. E.↗

NASA NTRS1982

Analytical study of vortex flaps on highly swept delta wings

This paper highlights some current results from ongoing analytical studies of vortex flaps on highly swept delta wings. A brief discussion of the vortex flow analysis tools is given along with comparisons of the theories to vortex flap force and pressure data. Theoretical trends in surface pressure distribution for both angle-of-attack variation and flap deflection are correctly predicted by Free Vortex Sheet theory. Also shown are some interesting calculations for attached-flow and vortex-flow flap hinge moments that indicate flaps utilizing vortex flow may generate less hinge moment than attached flow flaps. Finally, trailing-edge flap effects on leading-edge flap thrust potential are investigated and theory-experiment comparisons made.

Frink, N. T.↗

NASA NTRS1983

Aerodynamic characteristics of a series of twin-inlet air-breathing missile configurations. 2: Two-dimensional inlets at supersonic speeds

A series of air-breathing missile configurations was investigated to provide a data base for the design of such missiles. The model could be configured with either twin axisymmetric or two dimensional inlets. Three circumferential inlet locations were investigated: 90 deg, 115 deg, and 135 deg from the top center. Two vertical wing locations, as well as wingless configurations, were used. Three tail configurations were formed by locating the tail surfaces either on the inlet fairings or on fairings on the body. The surfaces were used to provide pitch control. Two dimensional inlets with extended compression surfaces, used to improve the angle-of-attack performance of the inlets for wingless configurations, were also investigated. The two dimensional inlet configurations are covered.

Hayes, C.↗

NASA NTRS1983

Aerodynamic characteristics of a series of twin-inlet air-breathing missile configurations. 3: Axisymmetric and two-dimensional inlets at subsonic-transonic speeds

A series of air-breathing missile configurations was investigated to provide a data base for the design of such missiles. The model could be configurated with either twin axisymmetric or two dimensional inlets. Three circumferential inlet locations were investigated: 90 deg, 115 deg, and 135 deg from the top center. Two vertical wing locations, as well as wingless configurations, were used. Three tail configurations were formed by locating the tail surfaces either on the inlet fairing or on the inlet fairing or on fairings on the body. The surfaces were used to provide pitch control. Two dimensional inlets with extended compression surfaces, used to improve the angle-of-attack performance of the inlets for wingless configurations, were also investigated. The twin axisymmetric two dimensional inlet types without internal flow are covered, and the boost configuration of an air-breathing missile is simulated.

Hayes, C.↗

NASA NTRS1985

Numerical simulation of viscous supersonic flow over a generic fighter configuration

A procedure is presented, as well as some results, to calculate the flow over a generic fighter configuration. A parabolized marching Navier-Stokes code is used to obtain the solution over a wing-canopy body. The flow conditions simulate supersonic cruise with a freestream Mach number of 2.169 and angles of attack of 4 and 10 deg. The body surface is considered to be an adiabatic wall and the flow is assumed to be turbulent for the given Reynolds number.

Chaussee, D. S.↗

NASA NTRS1985

An experimental investigation of apex fence flaps on delta wings

The effects on leading edge vortex formation produced by apex fences mounted on two delta wings were examined in a subsonic wind tunnel. Pressure data were collected on the upper surface of the test wings and oil flow and bubble flow visualizations were performed. The deflections studied were 74, 90 and 65 deg in one-sided and two-sided configurations. Full span and semi-span delta wing planforms were used. Relatively small, symmetrically deployed apex fences provided enhanced upper surface suction which, in some angle-of-attack situations, could lead to a pitch-up force. Asymmetrically-deployed fences showed promise for yaw control at high angles of attack, althouh the force effectiveness has yet to be quantified.

Vess, R. J.↗

NASA NTRS1986

Numerical simulation of vortical flow over an elliptical-body missile at high angles of attack

Numerical solutions to the Reynolds-averaged Navier-Stokes equations are given for the flow about an elliptical body missile (3:1 ellipse) at a Mach number of 2.5 and a unit Reynolds number of 6.56 x 10 to the 6th/m. At high angles of attack, the flow is dominated by large-scale free vortices which occur in the lee-side flow field due to crossflow boundary-layer separation. Emphasis is focused on the accurate prediction of the lee-side vortical flow. Solutions are presented for both symmetric and asymmetric (body rolled 45 deg) configurations at 10 deg and 20 deg angle of attack. The computed results are compared with experimental surface pressure coefficients and vapor-screen photographs. Excellent agreement is obtained in all cases.

Newsome, R. N.↗

NASA NTRS1988

Measurement of local convective heat transfer coefficients from a smooth and roughened NACA-0012 airfoil: Flight test data

Wind tunnels typically have higher free stream turbulence levels than are found in flight. Turbulence intensity was measured to be 0.5 percent in the NASA Lewis Icing Research Tunnel (IRT) with the cloud making sprays off and around 2 percent with cloud making equipment on. Turbulence intensity for flight conditions was found to be too low to make meaningful measurements for smooth air. This difference between free stream and wing tunnel conditions has raised questions as to the validity of results obtained in the IRT. One objective of these tests was to determine the effect of free stream turbulence on convective heat transfer for the NASA Lewis LEWICE ice growth prediction code. These tests provide in-flight heat transfer data for a NASA-0012 airfoil with a 533 cm chord. Future tests will measure heat transfer data from the same airfoil in the Lewis Icing Research Tunnel. Roughness was obtained by the attachment of small, 2 mm diameter hemispheres of uniform size to the airfoil in three different patterns. Heat transfer measurements were recorded in flight on the NASA Lewis Twin Otter Icing Research Aircraft. Measurements were taken for the smooth and roughened surfaces at various aircraft speeds and angles of attack up to four degrees. Results are presented as Frossling number versus position on the airfoil for various roughnesses and angles of attack.

Newton, James E.↗

NASA NTRS1988

Measurement of local convective heat transfer coefficients from a smooth and roughened NACA-0012 airfoil - Flight test data

Wind tunnels typically have higher free stream turbulence levels than are found in flight. Turbulence intensity was measured to be 0.5 percent in the NASA Lewis Icing Research Tunnel (IRT) with the cloud making sprays off and around 2 percent with cloud making equipment on. Turbulence intensity for flight conditions was found to be too low to make meaningful measurements for smooth air. This difference between free stream and wind tunnel conditions has raised questions as to the validity of results obtained in the IRT. One objective of these tests was to determine the effect of free stream turbulence on convective heat transfer for the NASA Lewis LEWICE ice growth prediction code. These tests provide in-flight heat transfer data for a NASA-0012 airfoil with a 533 cm chord. Future tests will measure heat transfer data from the same airfoil in the Lewis Icing Research Tunnel. Roughness was obtained by the attachment of small, 2 mm diameter hemispheres of uniform size to the airfoil in three different patterns. Heat transfer measurements were recorded in flight on the NASA Lewis Twin Otter Icing Research Aircraft. Measurements were taken for the smooth and roughened surfaces at various aircraft speeds and angles of attack up to four degrees. Results are presented as Frossling number versus position on the airfoil for various roughnesses and angles of attack.

Van Fossen, G. James↗

NASA NTRS1989

A computational challenge - Euler solution for ellipses

The Euler equations for flow past an ellipse are solved numerically. A computationally complex problem involving inviscid flow past an elliptical two-dimensional surface at subcritical Mach number and angle of attack is introduced. A lifting solution for any combination of grid and/or angle of attack which is nonsymmetric is obtained.

Pulliam, Thomas H.↗

NASA NTRS1990

Aerothermodynamic measurements on a proposed assured crew return vehicle (ACRV) lifting-body configuration at Mach 6 and 10 in air

A 0.02-scale model of a lifting-body concept for possible application to the Assured Crew Return Vehicle from Space Station Freedom was tested at Mach 6 and 10 in air. Thermal mappings and surface streamline patterns were obtained at angles of attack ranging from 0 to 30 deg and unit Reynolds numbers Re from 2 to 8 x 10 to the 6th/ft. Areas that experienced the highest heating were near the model nose and tip-fin leading edges. The effect of Re on windward centerline heating coefficients was negligible, whereas increases in angles of attack produced increases in heating. At Mach 6 and the highest unit Re, turbulent heat at the windward centerline was three to four times the laminar level. Leeward crossflow separation and vortex reattachment along the centerline are evident across the Re and angle-of-attack ranges tested, indicative of a complex flowfield.

Horvath, Thomas J.↗

NASA NTRS1991

An evaluation of three two-dimensional computational fluid dynamics codes including low Reynolds numbers and transonic Mach numbers

Full-potential, Euler, and Navier-Stokes computational fluid dynamics (CFD) codes were evaluated for use in analyzing the flow field about airfoils sections operating at Mach numbers from 0.20 to 0.60 and Reynolds numbers from 500,000 to 2,000,000. The potential code (LBAUER) includes weakly coupled integral boundary layer equations for laminar and turbulent flow with simple transition and separation models. The Navier-Stokes code (ARC2D) uses the thin-layer formulation of the Reynolds-averaged equations with an algebraic turbulence model. The Euler code (ISES) includes strongly coupled integral boundary layer equations and advanced transition and separation calculations with the capability to model laminar separation bubbles and limited zones of turbulent separation. The best experiment/CFD correlation was obtained with the Euler code because its boundary layer equations model the physics of the flow better than the other two codes. An unusual reversal of boundary layer separation with increasing angle of attack, following initial shock formation on the upper surface of the airfoil, was found in the experiment data. This phenomenon was not predicted by the CFD codes evaluated.

Hicks, Raymond M.↗

NASA NTRS1993

The X-15 airplane - Lessons learned

The X-15 rocket research airplane flew to an altitude of 354,000 ft and reached Mach 6.70. In almost 200 flights, this airplane was used to gather aerodynamic-heating, structural loads, stability and control, and atmospheric-reentry data. This paper describes the origins, design, and operation of the X-15 airplane. In addition, lessons learned from the X-15 airplane that are applicable to designing and testing the National Aero-Space Plane are discussed.

Aeroplane↗