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At least 433 records · Page 24
All sourcesNASA NTRSDOE OSTI
NASA NTRS1976

Correction factory techniques for improving aerodynamic prediction methods

A method for correcting discrete element lifting surface theory to reflect given experimental data is presented. Theoretical pressures are modified such that imposed constraints are satisfied while minimizing the changes to the pressures. Several types of correction procedures are presented and correlated; (1) scaling of pressures; (2) scaling of downwash values; and (3) addition of an increment to the downwash that is proportioned to pressure. Some special features are included in these methods and they include: (1) consideration of experimental data from multiple deflection modes, (2) limitation of the amplitudes of the correction factors, and (3) the use of correction factor mode shapes. These methods are correlated for cases involving all three Mach Number ranges using a FORTRAN IV computer program. Subsonically, a wing with an oscillating partial span control surface and a wing with a leading edge droop are presented. Transonically a two-dimensional airfoil with an oscillating flap is considered. Supersonically an arrow wing with and without camber is analyzed. In addition to correction factor methods an investigation is presented dealing with a new simplified transonic modification of the two-dimensional subsonic lifting surface theory. Correlations are presented for an airfoil with an oscillating flap.

Giesing, J. P.↗
NASA NTRS1975

Transonic aerodynamic characteristics of a wing/body combination incorporating jet flaps

A 0.25-scale semispan wing/body model with two types of jet flaps was tested in the Ames 11- by 11-Foot Transonic Wind Tunnel. The objective of that testing was to measure the static aerodynamic forces and moments and wing pressure distributions on six configurations differentiated by wing camber, jet flap type, and jet flap angle. Maximum thrust coefficients were limited to 0.12. Angle of attack was varied from -4 deg to 15 deg for Mach numbers between 0.6 and 0.95 at a constant unit Reynolds number of 18.0 million/m (5.5 million/ft). More refined designs and considerably more testing will be required to establish the practicability of the total-exhausting jet flap concept.

Holmberg, J. L.↗
NASA NTRS1976

Applications of vortex lattice theory to preliminary aerodynamic design

Some applications of the vortex-lattice theory to the preliminary aerodynamic design and analysis of subsonic aircraft were presented. These methods include the Rockwell-Tulinius vortex-lattice theory for estimating aerodynamic characteristics, a Trefftz plane optimization procedure for determining the span loads for minimum induced drag, and a modification of the Trefftz plane procedure to estimate the induced drag for specified span loads. The first two methods are used to aerodynamically design aircraft planforms, twists, and cambers, and the latter method is used to estimate the drag for components such as flaps and control surfaces. Results from the theories for predicting lift and pitching moment, drag due to lift, and the drag of control surfaces are compared with experimental data. This data was obtained on a general aviation model with flaps and a close-coupled canard-wing model.

Paulson, J. W., Jr.↗
NASA NTRS1926

Lattice arrangements for rapid convergence

A simple, systematic, optimized vortex-lattice approach is developed for application to lifting-surface problems. It affords a significant reduction in computational costs when compared to current methods. Extensive numerical experiments have been carried out on a wide variety of configurations, including wings with camber and single or multiple flaps, as well as high-lift jetflap systems. Rapid convergence as the number of spanwise or chordwise lattices are increased is assured, along with accurate answers. The results from this model should be useful not only in preliminary aircraft design but also, for example, as input for wake vortex roll-up studies and transonic flow calculations.

Hough, G. R.↗
NASA NTRS1976

A subvortex technique for the close approach to a discretized vortex sheet

The close-approach problem associated with vortex-lattice methods was examined numerically with the objective of calculating velocities at arbitrary points, not just at midpoints, between the vortices. The objective was achieved using a subvortex technique in which a vortex splits into an increasing number of subvortices as it is approached. The technique, incorporated in a two-dimensional potential flow method using "submerged" vortices and sources, was evaluated for a cambered Joukowski airfoil. The method could be extended to three dimensions, and should improve non-linear methods, which calculate interference effects between multiple wings and vortex wakes, and which include procedures for force-free wakes.

Maskew, B.↗
NASA NTRS1976

A computer program to calculate the longitudinal aerodynamic characteristics of wing-flap configurations with externally blown flaps

A vortex lattice lifting-surface method is used to model the wing and multiple flaps. Each lifting surface may be of arbitrary planform having camber and twist, and the multiple-slotted trailing-edge flap system may consist of up to ten flaps with different spans and deflection angles. The engine wakes model consists of a series of closely spaced vortex rings with circular or elliptic cross sections. The rings are normal to a wake centerline which is free to move vertically and laterally to accommodate the local flow field beneath the wing and flaps. The two potential flow models are used in an iterative fashion to calculate the wing-flap loading distribution including the influence of the waves from up to two turbofan engines on the semispan. The method is limited to the condition where the flow and geometry of the configurations are symmetric about the vertical plane containing the wing root chord. The calculation procedure starts with arbitrarily positioned wake centerlines and the iterative calculation continues until the total configuration loading converges within a prescribed tolerance. Program results include total configuration forces and moments, individual lifting-surface load distributions, including pressure distributions, individual flap hinge moments, and flow field calculation at arbitrary field points.

Mendenhall, M. R.↗
NASA NTRS1975

Cascade with subsonic leading-edge locus

The paper investigates a two-dimensional oscillating cascade with a subsonic leading edge locus in a supersonic flow. The blades are assumed to be of small thickness and camber, and are undergoing small amplitude-harmonic oscillations. The problem is reduced to the solution of a functional integral equation, and an expression is given for the kernel function.

Goldstein, M. E.↗
NASA NTRS1976

A higher order panel method for linearized supersonic flow

A higher order panel method using linearly varying source and quadratically varying doublet distributions for computing linearized supersonic flow over arbitrary wings and bodies is described. With tangential mass flux boundary conditions on circular cones at angles of attack, the source panel method gave pressure distributions which are in excellent agreement with exact nonlinear theory. The numerical source panel method was demonstrated to be very stable by randomly panelling an axially symmetric body. For wings with thickness and camber, both the source and doublet panel methods yield excellent agreement with exact linearized solutions. Computational experiments indicate that the use of exact surface paneling in lieu of linearized boundary conditions for wings provides better agreement with higher order theory and with experiment.

Ehlers, F. E.↗
NASA NTRS1976

An improved method for the prediction of completely three-dimensional aerodynamic load distributions of configurations with leading edge vortex separation

The application of a higher-order subsonic potential flow panel method to the solution of three-dimensional flow about wing and wing-body combinations with leading-edge vortex separation is presented. The governing equations are the linear flow differential equation and nonlinear boundary conditions which require that the flow be parallel to the wing and body surfaces and that the free vortex sheet, springing from the leading and trailing edges, be aligned with the local flow and support no pressure jump. The vortex core is modeled as a simple line vortex which receives vorticity from the free sheet through a connecting sheet. The Kutta condition is imposed on all appropriate edges of the wing. This set of nonlinear equations is solved by an iterative procedure. The Goethert rule accounts for compressibility. The method has been programmed for the CDC 6600. Delta wings, gothic wings, arrow wings, cambered wings, and wing with body have been analyzed. Initial studies involving variations of panel density, vortex sheet sizing, Jacobian update, and initial geometry demonstrate that the present method generally exhibits good convergence characteristics.

Rubbert, p. E.↗
NASA NTRS1976

Inverse transonic airfoil design including viscous interaction

A numerical technique was developed for the analysis of specified transonic airfoils or for the design of airfoils having a prescribed pressure distribution, including the effect of weak viscous interaction. The method uses the full potential equation, a stretched Cartesian coordinate system, and the Nash-MacDonald turbulent boundary layer method. Comparisons with experimental data for typical transonic airfoils show excellent agreement. An example shows the application of the method to design a thick aft-cambered airfoil, and the effects of viscous interaction on its performance are discussed.

Carlson, L. A.↗
NASA NTRS1977

Two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90

An investigation was conducted in the Langley 6- by 28-inch transonic tunnel and the 6- by 19-inch transonic tunnel to determine the two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90. The airfoils differed in thickness, thickness distribution, and camber. The FX69-H-098, the BHC-540, and the NACA 0012 airfoils were investigated in the 6- by 28-inch tunnel at Reynolds numbers (based on chord) from about 4.7 to 9.3 million at the lowest and highest test Mach numbers respectively. The FX69-H-098, the NLR-1, the BHC-540, and the NACA 23012 airfoils were investigated in the 6- by 19-inch tunnel at Reynolds numbers from about 0.9 to 2.2 million at the lowest and highest test Mach numbers respectively.

Noonan, K. W.↗
NASA NTRS1977

Assessment of existing analytic methods for prediction of high angle-of-attack loads on delta wings at supersonic speeds

An assessment of the applicability of four loading prediction methods to high angle-of-attack conditions for simplified wing-body configurations is provided. The methods are: The tangent wedge approximation, the linear theory methods of Middleton and Woodward, and a shock-fitting finite-difference technique. Estimates obtained by these methods were compared with experimental pressure data on delta wings to examine the effects of Mach number, camber, sweep angle, and angle of attack. Results indicate that all of the methods provided reasonable estimates at moderate angles of attack. At these moderate angles of attack, the methods of Middleton and Woodward provided good estimates at Mach numbers higher than those usually associated with linear theory. Only the finite-difference method provided reasonable load estimates at high angles of attack.

Landrum, E. J.↗
NASA NTRS1977

Aerodynamic characteristics at Mach 6 of a hypersonic research airplane concept having a 70 deg swept delta wing

The hypersonic aerodynamic characteristics of an air-launched, delta-wing research aircraft concept were investigated at Mach 6. The effect of various components such as nose shape, wing camber, wing location, center vertical tail, wing tip fins, forward delta wing, engine nacelle, and speed brakes was also studied. Tests were conducted with a 0.021 scale model at a Reynolds number, based on model length, of 10.5 million and over an angel of attack range from -4 deg to 20 deg. Results show that most configurations with a center vertical tail have static longitudinal stability at trim, static directional stability at angles of attack up to 12 deg, and static lateral stability throughout the angle of attack range. Configurations with wing tip fins generally have static longitudinal stability at trim, have lateral stability at angles of attack above 8 deg, and are directionally unstable over the angle of attack range.

Clark, L. E.↗
NASA NTRS1977

Force and pressure tests of the GA(W)-1 airfoil with a 20% aileron and pressure tests with a 30% Fowler flap

Wind tunnel force and pressure tests were conducted for the GA(W)-1 airfoil equipped with a 20% aileron, and pressure tests were conducted with a 30% Fowler flap. All tests were conducted at a Reynolds number of 2.2 and a Mach number of 0.13. The aileron provides control effectiveness similar to ailerons applied to more conventional airfoils. Effects of aileron gaps from 0% to 2% chord were evaluated, as well as hinge moment characteristics. The aft camber of the GA(W)-1 section results in a substantial up-aileron moment, but the hinge moments associated with aileron deflection are similar to other configurations. Fowler flap pressure distributions indicate that unseparated flow is achieved for flap settings up to 40 deg., over a limited angle of attack range. Theoretical pressure distributions compare favorably with experiments for low flap deflections, but show substantial errors at large deflections.

Wentz, W. H., Jr.↗
NASA NTRS1976

Characteristics of swept wings at high speeds, 30 January 1952

Some results of recent swept wing investigations are presented, that were undertaken to determine the effects of thickness and thickness distribution, camber and twist, nose-flap deflection, and devices or fixes for improving the wing pitching moment characteristics at high lift coefficients.

Donlan, C. J.↗
NASA NTRS1977

F-8 supercritical wing flight pressure, Boundary layer, and wake measurements and comparisons with wind tunnel data

Data for speeds from Mach 0.50 to Mach 0.99 are presented for configurations with and without fuselage area-rule additions, with and without leading-edge vortex generators, and with and without boundary-layer trips on the wing. The wing pressure coefficients are tabulated. Comparisons between the airplane and model data show that higher second velocity peaks occurred on the airplane wing than on the model wing. The differences were attributed to wind tunnel wall interference effects that caused too much rear camber to be designed into the wing. Optimum flow conditions on the outboard wing section occurred at Mach 0.98 at an angle of attack near 4 deg. The measured differences in section drag with and without boundary-layer trips on the wing suggested that a region of laminar flow existed on the outboard wing without trips.

Montoya, L. C.↗
NASA NTRS1977

Theoretical prediction of over-wing-blowing aerodynamics

A theoretical method is developed for predicting the aerodynamic characteristics of wings with over-wing-blowing jet. The method allows the jet to stay above the wing surface as well as to wash the surface. It accounts for the wing-jet interaction due to differences between the jet and freestream dynamic pressures and Mach numbers, in addition to the jet entrainment. For the former effect, the quasi-vortex-lattice method is used to satisfy the jet and wing boundary conditions. For the latter, a new theory was developed to calculate the jet entrained flow for given jet properties. Comparison of predicted results with available data of various configurations shows reasonably good agreement. Further theoretical analysis indicates that it is aerodynamically advantageous to locate the jet exit near and ahead of the wing leading edge, and that the camber shape has significant effect on the induced drag.

Lan, C. E.↗
NASA NTRS1977

Numerical solutions for viscous and potential flow about arbitrary two-dimensional bodies using body-fitted coordinate systems

A procedure for numerical solution of the time-dependent, incompressible Navier-Stokes equations for the flow about arbitrarily shaped two-dimensional bodies is given. This solution is based on a technique of automatic numerical generation of a curvilinear coordinate system having a coordinate line coincident with the body contour regardless of its shape. The implicit solution utilizes the vorticity-stream function formulation with a false-position iterative adjustment of the surface vorticity in satisfaction of the no-slip boundary condition. Excellent agreement with the Blasius boundary layer solution is obtained for a semi-infinite flat plate. Results are presented for Reynolds numbers up to 2000 for several airfoils and a cambered rock.

Thames, F. C.↗