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At least 613 records · Page 34

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NASA NTRS1990

Numerical solutions of the linearized Euler equations for unsteady vortical flows around lifting airfoils

A linearized unsteady aerodynamic analysis is presented for unsteady, subsonic vortical flows around lifting airfoils. The analysis fully accounts for the distortion effects of the nonuniform mean flow on the imposed vortical disturbances. A frequency domain numerical scheme which implements this linearized approach is described, and numerical results are presented for a large variety of flow configurations. The results demonstrate the effects of airfoil thickness, angle of attack, camber, and Mach number on the unsteady lift and moment of airfoils subjected to periodic vortical gusts. The results show that mean flow distortion can have a very strong effect on the airfoil unsteady response, and that the effect depends strongly upon the reduced frequency, Mach number, and gust wave numbers.

Scott, James R.↗

NASA NTRS1990

Numerical simulations of supersonic flow through oscillating cascade sections

A finite difference code was developed for modeling inviscid, unsteady supersonic flow by solution of the compressible Euler equations. The code uses a deforming grid technique to capture the motion of the airfoils and can model oscillating cascades with any arbitrary interblade phase angle. A flat plate cascade is analyzed, and results are compared with results from a small perturbation theory. The results show very good agreement for both the unsteady pressure distributions and the integrated force predictions. The reason for using the numerical Euler code over a small perturbation theory is the ability to model real airfoils that have thickness and camber. Sample predictions are presented for a cascade of loaded airfoils and show appreciable differences in the unsteady surface pressure distributions when compared with the flat plate results.

Huff, Dennis L.↗

NASA NTRS1989

Three-dimensional aerodynamics of an annular airfoil cascade including loading effects

A series of experiments are described which investigate and quantify the effect of loading on the three-dimensional flow through a subsonic annular cascade of cambered airfoils. At two levels of loading, detailed data quantify the cascade inlet velocity, the intrapassage flow field, the airfoil surface pressure distributions, the exit flow field, and the total pressure loss distributions. Aerodynamic loading is shown to strengthen the radial pressure gradient, the passage vortex structure, the vortex-endwall boundary layer interactions, and the losses.

Fleeter, S.↗

NASA NTRS1990

Numerical solutions of the linearized Euler equations for unsteady vortical flows around lifting airfoils

Scott, James R.↗

NASA NTRS1990

Numerical solutions of the linearized Euler equations for unsteady vortical flows around lifting airfoils

Scott, James R.↗

NASA NTRS1989

Wind-tunnel investigations of wings with serrated sharp trailing edges

Exploratory wind-tunnel force measurements are presented for two wing geometries with small-scale planar and nonplanar serrated trailing-edge devices (chord-Reynolds numbers ranged from 1.0 - 3.7 x 10 to the 6th). The planar serrated trailing-edge extensions reduced the drag at conditions when trailing-edge separation occurred at low angles of attack. The introduction of serrations reduced or eliminated the drag penalty due to the small (1-2 percent of the chord length) nonplanar trailing-edge flaps, while maintaining the effects of increase in camber. The presence of streamwise vortices immediately downstream of the serrated trailing edges is believed to have favorably affected the boundary-layer flow approaching the trailing edge and the near-wake development, resulting in reduced pressure drag.

Vijgen, P. M. H. W.↗

NASA NTRS1990

Cascade aerodynamic gust response including steady loading effects

To predict the unsteady convected gust aerodynamic response of a cascade comprised of arbitrary thick and cambered aerofoils in an incompressible, inviscid, flow field, a complete first-order model is formulated. The flow is analyzed by considering a periodic flow channel. The velocity potential is separated into steady and unsteady harmonic components, each described by a Laplace equation. The strong dependence of the unsteady aerodynamics on the steady effects of aerofoil and cascade geometry and incidence angle is manifested in the coupling of the unsteady and steady flow fields through the unsteady boundary conditions. Analytical solutions in individual grid elements of a body-fitted computational grid are then determined, with the complete solution obtained by assembly of these local solutions. The validity and capabilities of this model and solution technique are then demonstrated by analyzing the steady and unsteady aerodynamics of both theoretical and experimental cascade configurations.

Chiang, Hsiao-Wei D.↗

NASA NTRS1989

The use of a Navier-Stokes code in the wing design process

The feasibility was determined of incorporating the Navier-Stokes computational code, CFL3D, into the supersonic wing design process. The approach taken is of two steps. The first step was to calibrate CFL3D against existing experimental data sets obtained on thin sharp edged delta wings. The experimental data identified six flow types which are dependent on the similarity parameters of Mach number and angle of attack normal to the leading edge. The calibration showed CFL3D capable of simulating these various separated and attached flow conditions. The second step was to use CFL3D to study the initial formation of leading edge separation over delta wings at supersonic speeds. This consisted of examining solutions obtained on a 65 deg delta wing at Mach number of 1.6 with varying cross sectional shapes. Reynolds number was held constant at 1000000 and the Baldwin-Lomax turbulence model was used. The study showed that through the use of leading edge radius and/or camber, the onset of leading edge separation can be delayed to a higher angle of attack than observed on a flat sharp edged wing. Based on the geometries studied, three wind tunnel models are being designed to verify these results.

Mcmillin, S. Naomi↗

NASA NTRS1990

Navier-Stokes analysis of transonic cascade flow

A new kind of C-type grid is proposed, this grid is non-periodic on the wake and allows minimum skewness for cascades with high turning and large camber. Reynolds-averaged Navier-Stokes equations are solved on this type of grid using a finite volume discretization and a full multigrid method which uses Runge-Kutta stepping as the driving scheme. The Baldwin-Lomax eddy-viscosity model is used for turbulence closure. A detailed numerical study is proposed for a highly loaded transonic blade. A grid independence analysis is presented in terms of pressure distribution, exit flow angles, and loss coefficient. Comparison with experiments clearly demonstrates the capability of the proposed procedure.

Arnone, A.↗

NASA NTRS1991

Compressible flows with periodic vortical disturbances around lifting airfoils

A numerical method is developed for solving periodic, three-dimensional, vortical flows around lifting airfoils in subsonic flow. The first-order method that is presented fully accounts for the distortion effects of the nonuniform mean flow on the convected upstream vortical disturbances. The unsteady velocity is split into a vortical component which is a known function of the upstream flow conditions and the Lagrangian coordinates of the mean flow, and an irrotational field whose potential satisfies a nonconstant-coefficient, inhomogeneous, convective wave equation. Using an elliptic coordinate transformation, the unsteady boundary value problem is solved in the frequency domain on grids which are determined as a function of the Mach number and reduced frequency. The numerical scheme is validated through extensive comparisons with known solutions to unsteady vortical flow problems. In general, it is seen that the agreement between the numerical and analytical results is very good for reduced frequencies ranging from 0 to 4, and for Mach numbers ranging from .1 to .8. Numerical results are also presented for a wide variety of flow configurations for the purpose of determining the effects of airfoil thickness, angle of attack, camber, and Mach number on the unsteady lift and moment of airfoils subjected to periodic vortical gusts. It is seen that each of these parameters can have a significant effect on the unsteady airfoil response to the incident disturbances, and that the effect depends strongly upon the reduced frequency and the dimensionality of the gust. For a one-dimensional (transverse) or two-dimensional (transverse and longitudinal) gust, the results indicate that airfoil thickness increases the unsteady lift and moment at the low reduced frequencies but decreases it at the high reduced frequencies. The results show that an increase in airfoil Mach number leads to a significant increase in the unsteady lift and moment for the low reduced frequencies, but a significant decrease for the high reduced frequencies.

Scott, James R.↗

NASA NTRS1991

Transonic cascade flow calculations using non-periodic C-type grids

A new kind of C-type grid is proposed for turbomachinery flow calculations. This grid is nonperiodic on the wake and results in minimum skewness for cascades with high turning and large camber. Euler and Reynolds averaged Navier-Stokes equations are discretized on this type of grid using a finite volume approach. The Baldwin-Lomax eddy-viscosity model is used for turbulence closure. Jameson's explicit Runge-Kutta scheme is adopted for the integration in time, and computational efficiency is achieved through accelerating strategies such as multigriding and residual smoothing. A detailed numerical study was performed for a turbine rotor and for a vane. A grid dependence analysis is presented and the effect of artificial dissipation is also investigated. Comparison of calculations with experiments clearly demonstrates the advantage of the proposed grid.

Arnone, Andrea↗

NASA NTRS1991

Unified aeroacoustics analysis for high speed turboprop aerodynamics and noise. Volume 4: Computer user's manual for UAAP turboprop aeroacoustic code

The Unified AeroAcoustic Program (UAAP) code calculates the airloads on a single rotation prop-fan, or propeller, and couples these airloads with an acoustic radiation theory, to provide estimates of near-field or far-field noise levels. The steady airloads can also be used to calculate the nonuniform velocity components in the propeller wake. The airloads are calculated using a three dimensional compressible panel method which considers the effects of thin, cambered, multiple blades which may be highly swept. These airloads may be either steady or unsteady. The acoustic model uses the blade thickness distribution and the steady or unsteady aerodynamic loads to calculate the acoustic radiation. The users manual for the UAAP code is divided into five sections: general code description; input description; output description; system description; and error codes. The user must have access to IMSL10 libraries (MATH and SFUN) for numerous calls made for Bessel functions and matrix inversion. For plotted output users must modify the dummy calls to plotting routines included in the code to system-specific calls appropriate to the user's installation.

Menthe, R. W.↗

NASA NTRS1991

RWF rotor-wake-fuselage code software reference guide

The RWF (Rotor-Wake-Fuselage) code was developed from first principles to compute the aerodynamics associated with the complex flow field of helicopter configurations. The code is sized for a single, multi-bladed main rotor and any configuration of non-lifting fuselage. The mathematical model for the RWF code is based on the integration of the momentum equations and Green's theorem. The unknowns in the problem are the strengths of prescribed singularity distributions on the boundaries of the flow. For the body (fuselage) a surface of constant strength source panels is used. For the rotor blades and rotor wake a surface of constant strength doublet panels is used. The mean camber line of the rotor airfoil is partitioned into surface panels. The no-flow boundary condition at the panel centroids is modified at each azimuthal step to account for rotor blade cyclic pitch variation. The geometry of the rotor wake is computers at each time step of the solution. The code produces rotor and fuselage surface pressures, as well as the complex geometry of the evolving rotor wake.

Berry, John D.↗

NASA NTRS1989

Longitudinal aerodynamic characteristics of a subsonic, energy-efficient transport configuration in the National Transonic Facility

The Reynolds number, aeroelasticity, boundary layer transition, and nonadiabatic wall temperature effects, and data repeatability was determined in the National Transonic Facility (NTF) for a subsonic, energy efficient transport model. The model was tested over a Mach number range of 0.50 to 0.86 and a Reynolds number range of 1.9 million to approximately 23.0 million (based on mean geometric chord). The majority of the data was taken using cryogenic nitrogen (data at 1.9 million Reynolds number was taken in air). Force and moment, wing pressure, and wing thermocouple data are presented. The data indicate that increasing Reynolds number resulted in greater effective camber of the supercritical wing and horizontal tail, resulting in greater lift and pitching moment coefficients at nearly all angles of attack for M = 0.82. As Reynolds number was increased, untrimmed L/D increased, the angle of attack for maximum L/D decreased, drag creep was reduced significantly, and drag divergence Mach number increased slightly. Data repeatability for both modes of operation of the NTF (air and cryogenic nitrogen) was generally very good, and nonadiabatic wall effects were estimated to be small. Transition-free and transition-fixed configurations had significantly different force and moment data at M = 0.82 for low Reynolds number, and very small differences were noted at high Reynolds numbers.

Jacobs, Peter F.↗

NASA NTRS1989

Preliminary flight-determined subsonic lift and drag characteristics of the X-29A forward-swept-wing airplane

The X-29A subsonic lift and drag characteristics determined, met, or exceeded predictions, particularly with respect to the drag polar shapes. Induced drag levels were as great as 20 percent less than wind tunnel estimates, particularly at coefficients of lift above 0.8. Drag polar shape comparisons with other modern fighter aircraft showed the X-29A to have a better overall aircraft aerodynamic Oswald efficiency factor for the same aspect ratio. Two significant problems arose in the data reduction and analysis process. These included uncertainties in angle of attack upwash calibration and effects of maneuver dynamics on drag levels. The latter problem resulted from significantly improper control surface automatic camber control scheduling. Supersonic drag polar results were not obtained during this phase because of a lack of engine instrumentation to measure afterburner fuel flow.

Hicks, John W.↗

NASA NTRS1991

User's guide to a system of finite-element supersonic panel flutter programs

The utilization and operation of a set of six computer programs for the prediction of panel flutter at supersonic speeds by finite element methods are described. The programs run individually to determine the flutter behavior of a flat panel where the finite elements which model the panel each have four degrees of freedom (DOF), a curved panel where the finite elements each have four DOF, and a curved panel where the finite elements each have six DOF. The panels are assumed to be of infinite aspect ratio and are subjected to either simply-supported or clamped boundary conditions. The aerodynamics used by these programs are based on piston theory. Application of the program is illustrated by sample cases where the number of beam finite elements equals four, the in-plane tension parameter is 0.0, the maximum camber to panel length ratio for a curved panel case is 0.05, and the Mach number is 2.0. This memorandum provides a user's guide for these programs, describes the parameters that are used, and contains sample output from each of the programs.

Woolley, Christine L.↗

NASA NTRS1989

AFTI/F-111 airplane mission adaptive wing operational flight evaluation technique using uplinked pilot command cues

NASA and the USAF have conducted a program to investigate aircraft performance improvements utilizing a mission adaptive wing (MAW). The MAW was designed and developed for the AFTI/F-111 variable-sweep aircraft to provide a hydraulically driven, smooth, and continuous variable camber of the trailing and leading edges as a function of maneuvering requirements or of flight conditions. The remotely augmented vehicle facility (RAV) at the NASA DFRF, as utilized in the MAW investigations, is described. The RAV was a dedicated, ground based, general purpose facility capable of receiving a data stream downlinked from a test vehicle, processing this data stream in a digital computer, and transmitting processed data back to the test vehicle. It is shown that this method of flight testing provides a technique that can evaluate highly dynamic maneuvers.

Kempel, Robert W.↗

NASA NTRS1990

Numerical simulations of supersonic flow through oscillating cascade sections

A finite difference code has been developed for modeling inviscid, unsteady supersonic flow by solution of the compressible Euler equations. The code uses a deforming grid technique to capture the motion of the airfoils and can model oscillating cascades with any arbitrary interblade phase angle. A flat plate cascade is analyzed, and results are compared with results from a small-perturbation theory. The results show very good agreement for both the unsteady pressure distributions and the integrated force predictions. The reason for using the numerical Euler code over a small-perturbation theory is the ability to model 'real' airfoils that have thickness and camber. Sample predictions are presented for a cascade of loaded airfoils and show appreciable differences in the unsteady surface pressure distributions when compared with the flat plate results.

Huff, Dennis L.↗