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At least 775 records · Page 43
All sourcesNASA NTRSDOE OSTI
NASA NTRS2006

Estimation of Longitudinal Unsteady Aerodynamics of a Wing-Tail Combination From Wind Tunnel Data

This paper presents an initial step toward model identification from wind tunnel data for an airliner configuration. Two approaches to modeling a transport configuration are considered and applied to both steady and large-amplitude forced-oscillation wind tunnel data taken over a wide range of angles of attack. Only limited conclusions could be drawn from this initial data set. Although model estimated time histories of normal force and pitching moment agree reasonably well with the corresponding measured values, model damping parameters did not, for some cases, have values consistent with small amplitude oscillatory data. In addition, large parameter standard errors implied poor information content for model structure determination and parameter estimation. Further investigation of the modeling problem for more general aerodynamic models is recommended with close attention to experiment design for obtaining parameters with high accuracy.

Murphy, Patrick C.↗
NASA NTRS1971

A numerical method for computing three dimensional viscous supersonic flow fields about slender bodies

A numerical method is presented for the calculation of steady, three-dimensional, viscous, compressible flow fields about slender bodies at angle of attack and at supersonic speeds. Approximations are introduced in modeling the flow in the longitudinal direction. Accordingly, the flow fields calculated with the program were computed with a model that permits viscous crossflow together with inviscid axial flow. An analysis of the errors introduced by such a treatment is presented. Numerical calculations were made and compared with experimental results for an ogive-cylinder and an airplane fuselage configuration. Generally, good agreement with experiment was obtained. However, boundary layer separation and body vortex positions differed from experimental locations on the ogive-cylinder, and the shock induced by the fuselage canopy was predicted at a slightly different location.

Walitt, L.↗
NASA NTRS1974

Effect of external tank nose shape on the Rockwell International space shuttle vehicle 3, integrated configuration (IA37B)

Tests of several tank nose shapes were conducted as a part of the investigation of configuration changes to reduce drag for the integrated vehicle. The effect on the integrated vehicle aerodynamic characteristics of several tank nose shapes was investigated. The tank nose shapes investigated were the 600 inch (baseline) and 1204 inch radius ogives, and the 600 inch ogive plus a spike 360 inches long and 12.0 inches in diameter. Data were obtained over a Mach number range of 0.6 through 4.96 and for angles-of-attack and sideslip from -10 through +10 degrees. The model used was the 0.004-scale integrated vehicle model number 34-OTS.

Allen, E. C.↗
NASA NTRS1977

Force testing manual for the Langley 20-inch Mach 6 tunnel

Data reduction and procedures for conducting force tests in a 20 inch Mach 6 tunnel are described. A discussion of pretest and testing phases are included. Items that are to be checked during model design and construction are outlined as well as safety requirements, starting loads tests, instructions for data acquisition and model installation. Measurement of balance and model misalignment and instructions for calibrating the angle of attack screen are covered. Procedures for making reference pressure, attitude tare, and data runs are included. The 20 inch tunnel force program is examined, and a description of data recording system input and load contrast sheets is given. An appendix presents a description, operating characteristics, and Mach number calibration of the tunnel, as well as tunnel characteristics.

Keyes, J. W.↗
NASA NTRS1977

Aerodynamics of slender finned bodies at large angles of attack

In certain missions finned missiles perform slewing maneuvers. Here, large angles of attack are attained. Experimental data needed to understand the aerodynamics of such vehicles are presented. The purpose of this investigation was to study the interaction of the body flow field with that produced by the fins and the resulting effects on the aerodynamic forces and moments. The experiments were conducted at a nominal Mach number of 2.7 and angles of attack from 0 to 50 deg, with two different models. The tests were performed in a range of Reynolds number from 1.5 x 10 to the 6th to 4 x 10 to the 7th per foot (to cover both the laminar and fully turbulent regimes.) Several fin roll angles were investigated. Static pressures on both body and fin surfaces are reported.

Agnone, A. M.↗
NASA NTRS1978

Visualization of the separation and subsequent transition near the leading edge of airfoils

A visual study was performed using the low speed smoke wind tunnels with the objective of obtaining a better understanding of the structure of leading edge separation bubbles on airfoils. The location of separation, transition and reattachment for a cylindrical nose constant-thickness airfoil model were obtained from smoke photographs and surface oil flow techniques. These data, together with static pressure distributions along the leading edge and upper surface of the model, produced the influence of Reynolds number, angle of attack, and trailing edge flap angle on the size and characteristics of the bubble. Additional visual insight into the unsteady nature of the separation bubble was provided by high speed 16 mm movies. The 8 mm color movies taken of the surface oil flow supported the findings of the high speed movies and clearly showed the formation of a scalloped spanwise separation line at the higher Reynolds number.

Arena, A. V.↗
NASA NTRS1979

Pressure distributions on three different cruciform aft-control surfaces of a wingless missile at Mach 1.60, 2.36, and 3.70. Volume 2: Clipped delta tail

Pressure coefficients were obtained in the Langley Unitary Plan wind tunnel for a wingless missile with a clipped delta tail. The angle of attack was varied from -4 deg to 20 deg, model roll angle was varied from 0 deg to 90 deg in 22.5 deg increments, and tail deflections were 0 deg to - 15 deg. The pressures were measured on two adjacent tail surfaces using 91 pressure orifices per tail surface. Results are presented in plotted and tabular form.

Lamb, M.↗
NASA NTRS1979

The effects of configuration changes on spin and recovery characteristics of a low-wing general aviation research airplane

A fully instrumented, low-wing, single-engine general aviation airplane has been spin tested. Several tail configurations, wing leading-edge modifications, fuselage modifications, moment-of-inertia variations, center-of-gravity positions, and control inputs have been tested to determine their effect on spinning and spin recovery. Results indicate that wing airfoil design can significantly influence airplane spin and recovery characteristics and can overpower the effects of tail design. Results also point out a need to determine limitations of such factors as Reynolds number in model spin test techniques and high angle-of-attack aerodynamics.

Stough, H. P., III↗
NASA NTRS1984

Wind tunnel wall interference correction for aircraft models

Wall interference correction procedures seek to determine the required changes in certain flow or geometric parameters so that the difference between the flow properties at the model's surface in the tunnel and free air are minimized. A transonic and a linear correction procedure were developed for aircraft models. In addition to Mach number and angle of attack corrections, an estimate of the accuracy of the corrections is provided by the transonic correction procedure. Lift, pitching moment and pressure measurements near the tunnel walls are required. The efficiency and accuracy of the correction procedure are improved. Moreover, correction of both the wing and tail angles of attack is allowed. The procedure is valid for transonic as well as subcritical flows. However, for subcritical flows further approximations and simplifying assumptions are made, leading to a very simple and efficient correction procedure.

Rizk, M. H.↗
NASA NTRS1987

Airfoil flutter model suspension system

A wind tunnel suspension system for testing flutter models under various loads and at various angles of attack is described. The invention comprises a mounting bracket assembly affixing the suspension system to the wind tunnel, a drag-link assembly and a compound spring arrangement comprises a plunge spring working in opposition to a compressive spring so as to provide a high stiffness to trim out steady state loads and simultaneously a low stiffness to dynamic loads. By this arrangement an airfoil may be tested for oscillatory response in both plunge and pitch modes while being held under high lifting loads in a wind tunnel.

Reed, Wilmer H.↗
NASA NTRS1987

Experimental flowfield visualization of a high alpha wing at Mach 1.62

Experimental oil-flow and tuft patterns and vapor-screen flow-visualization data were obtained on a cambered wing model at Mach 1.62 for an angle-of-attack range of 0 to 14 deg. These data were used as flow diagnostic tools along with surface-pressure and force data and full-potential theory calculations. A large separation bubble was found on the lower wing surface at low angle of attack. The high-angle-of-attack flowfield was characterized by a large attached-flow leading-edge expansion followed by a cross-flow shock. At 14 deg, the cross-flow shock apparently induced discrete regions of streamwise separated flow which were clearly indicated in the vapor-screen and oil-flow photographs.

Pittman, James L.↗
NASA NTRS1987

Experimental flowfield visualization of a high alpha wing at Mach 1.62

Experimental oil-flow and tuft patterns and vapor-screen flow-visualization data were obtained on a cambered wing model at Mach = 1.62 for an angle-of attack range of 0-14 deg. These data were used as flow diagnostic tools along with surface-pressure and force data and full-potential theory calculations. A large separation bubble was found on the lower wing surface at low angle of attack. The high-angle-of-attack flowfield was characterized by a large attached-flow leading-edge expansion followed by a crossflow shock. At alpha = 14 deg the crossflow shock apparently induced discrete regions of streamwise separated flow, which were clearly indicated in the vapor-screen and oil-flow photographs.

Pittman, James L.↗
NASA NTRS1988

A full-potential theory analysis of the supersonic aerodynamics of a 60-deg delta wing-body configuration

Experimental spanwise pressure distributions for a 60-deg delta wing/body of approximate fineness ratio 7.6 have been obtained and compared to predictions using full-potential theory. Analysis was performed at Mach 1.6 for angles of attack in the range 0.8 to 10 deg, and for Mach numbers ranging from 1.4 to 1.8 at lift coefficients 0.3 and 0.4. The intent of the study was to examine an attached flow approach for maneuver wing design in the presence of a fuselage. For the Mach number, angle-of-attack conditions considered, the full-potential theory accurately modeled the pressure distributions provided the flow remained attached. By combining the full-potential theory results with an empirical shock-induced separation criterion, it was found that the onset of shock-induced separation can be predicted. The investigation showed that, if an attached-flow approach is used with an empirical method of indicating shock-induced separation, the full-potential method is capable of being used as an effective tool for designing maneuver wings.

Rose, O. C.↗
NASA NTRS1989

Optimal guidance and propulsion control for transatmospheric vehicles

Problems associated with on-board trajectory optimization and with the synthesis of guidance laws are addressed for ascent to LEO of an air-breathing, single-stage-to-orbit vehicle. A multimode propulsion system is assumed which incorporates turbojet, ramjet, scramjet, and rocket engines. An energy-state approximation is applied to a four-state dynamic model for flight of a point mass over a spherical nonrotating earth. An algorithm for generating fuel-optimal climb profiles is derived via singular perturbation theory. This algorithm results from application of the minimum principle to a low-order dynamic model that includes general functional dependence on angle of attack and a normal component of thrust. Switching conditions are derived which, under appropriate assumptions, govern optimal transition from one propulsion mode to another. The use of bank angle to modulate the magnitude of the vertical component of lift is shown to improve the index performance. Numerical results illustrate the nature of the resulting fuel-optimal climb paths.

Corban, J. E.↗
NASA NTRS1990

Design of a three-component wall-mounted balance

The design and evaluation of a three-component, wall-mounted pyramidal balance for a small wind tunnel is discussed. The balance was designed to measure lift, drag, pitching moment, and angle of attack. The specific design of each component and mathematical models used to design the balance are covered. Balance evaluation consisted of calibration, tare, and interaction analysis.

Winkelmann, Allen E.↗
NASA NTRS1992

Efficient simulation of incompressible viscous flow over multi-element airfoils

The incompressible, viscous, turbulent flow over single and multi-element airfoils is numerically simulated in an efficient manner by solving the incompressible Navier-Stokes equations. The computer code uses the method of pseudo-compressibility with an upwind-differencing scheme for the convective fluxes and an implicit line-relaxation solution algorithm. The motivation for this work includes interest in studying the high-lift take-off and landing configurations of various aircraft. In particular, accurate computation of lift and drag at various angles of attack, up to stall, is desired. Two different turbulence models are tested in computing the flow over an NACA 4412 airfoil; an accurate prediction of stall is obtained. The approach used for multi-element airfoils involves the use of multiple zones of structured grids fitted to each element. Two different approaches are compared: a patched system of grids, and an overlaid Chimera system of grids. Computational results are presented for two-element, three-element, and four-element airfoil configurations. Excellent agreement with experimental surface pressure coefficients is seen. The code converges in less than 200 iterations, requiring on the order of one minute of CPU time (on a CRAY YMP) per element in the airfoil configuration.

Rogers, Stuart E.↗
NASA NTRS1993

The use of subscale models to predict self-induced oscillations of flight vehicles

The suitability of utilizing subscale dynamic wind tunnel experiments to predict wing rock for slender aircraft is presented. As observed in experiments, both flight and wind tunnel, there is significant interaction of the forebody and leading-edge extension vortices. It is concluded that subscale models can be employed to predict high angle of attack dynamic behavior such as wing rock provided proper precautions are taken.

Nelson, R. C.↗