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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1977

An experimental and computational investigation of the steady and unsteady transonic flow fields about an airfoil in a solid-wall test channel

An experimental and computational investigation of the steady and unsteady transonic flow field about a thick airfoil is described. An operational computer code for solving the two-dimensional, compressible Navier-Stokes equations for flow over airfoils was modified to include solid-wall, slipflow boundary conditions to properly assess the code and help guide the development of improved turbulence models. Steady and unsteady flow fields about an 18% thick circular arc airfoil at Mach numbers of 0.720, 0.754, and 0.783 and a chord Reynolds number of 11 x 10 to the 6th are predicted and compared with experiment. For the first time, computed results for unsteady turbulent flows with separation caused by a shock wave were obtained which qualitatively reproduce the time-dependent aspects of experiments. Features such as the intensity and reduced frequency of airfoil surface-pressure fluctuations, oscillatory regions of trailing-edge and shock-induced separation, and the Mach number range for unsteady flows were all qualitatively reproduced.

Levy, L. L., Jr.↗
NASA NTRS1993

A hybrid structured-unstructured grid method for unsteady turbomachinery flow computations

A hybrid grid technique for the solution of 2D, unsteady flows is developed. This technique is capable of handling complex, multiple component geometries in relative motion, such as those encountered in turbomachinery. The numerical approach utilizes a mixed structured-unstructured zonal grid topology along with modeling equations and solution methods that are most appropriate in the individual domains, therefore combining the advantages of both structured and unstructured grid techniques.

Mathur, Sanjay R.↗
NASA NTRS1994

Progress In Incompressible Pump Flow Calculations

Steady and unsteady flows for propulsion systems are efficiently simulated by solving the incompressible Navier-Stokes equations. The solution method is based on the pseudo compressibility approach and uses an implicit-upwind differencing scheme together with the Gauss-Seidel line relaxation method. Current computations use one equation Baldwin-Barth turbulence model which is derived from a simplified form of the standard kappa - epsilon model equations. The resulting computer code is applied to the flow analysis inside an advanced rocket pump impeller in steadily rotating reference frames. Numerical results are compared with experimental measurements. The effects of exit and shroud cavities with the leak-age flow are investigated. Time-accurate incompressible Navier-Stokes formulation with the overlapped grid scheme capability was evaluated by using MIT flapping foil experiment. The grid dependency, turbulence model effects, and the effect of order of differencing were investigated. Numerical results were compared against experimental data. The resulting procedure were applied to unsteady flapping foil calculations. Two upstream NACA 0025 foils perform high frequency synchronized motion and generate unsteady flow conditions to the downstream larger stationary foil. Comparison between unsteady experimental data and numerical results from two different moving boundary procedures will be presented.

Kiris, Cetin↗
NASA NTRS1973

A note on the unsteady cavity flow in a tunnel

The unsteady internal cavitating flow such as the one observed in a pump or a turbine is studied for a simple two-dimensional model of a base-cavitating wedge in an infinite tunnel and it is shown how the cavitation compliance can be calculated using the linearized free streamline theory. Numerical values are obtained for the limiting case of a free jet. Two important features are: First, the cavitation compliance is found to be of complex form, having additional resistive and reactive terms beyond the purely inertial oscillation of the whole channel in 'slug flow.' Second, the compliance has a strong dependence on frequency.

Kim, J. H.↗
NASA NTRS1981

Calculation of Unsteady Transonic Flow over an Airfoil

An implicit finite-difference solver for either the Euler equations or the "thin-layer" Navier-Stokes equations was used to calculate a transonic flow over the NACA 64A010 airfoil pitching about its one-quarter chord. An unsteady automatic grid-generation procedure that will improve significantly the computational efficiency of various unsteady flow problems is described. The calculated results for both inviscid and viscous flows at Much number 0.8 over the airfoil oscillating with reduced frequency referenced to one-half chord, 0.2, are compared with experimental data measured in the Ames 11 x 11-ft Transonic Wind Tunnel. Nonlinear, unsteady effects of the flow on the surface pressure variations, shock-wave excursions, and overall airloads are examined. Good agreements between the results of computations and experiments were obtained. In the shock-wave region, however, the results of the viscous-flow computations showed closer agreement with the experimental data.

Chyu, W. J.↗
NASA NTRS1981

Calculation of Unsteady Transonic Flow Over an Airfoil

An implicit finite-difference solver for either the Euler equations or the "thin-layer" Navier-Stokes equations was used to calculate a transonic flow over the NACA 64A010 airfoil pitching about its one-quarter chord. An unsteady automatic grid-generation procedure that will improve significantly the computational efficiency of various unsteady flow problems is described. The calculated results for both inviscid and viscous flows at Mach number 0.8 over the airfoil oscillating with reduced frequency referenced to one-half chord, 0.2, are compared with experimental data measured in the Ames 11 x 11 ft Transonic Wind Tunnel. Nonlinear, unsteady effects of the flow on the surface pressure variations, shock-wave excursions, and overall airloads are examined. Good agreements between the results of computations and experiments were obtained. In the shock-wave region, however, the results of the viscous-flow computations showed closer agreement with the experimental data.

Chyu, W. J.↗
NASA NTRS1982

Algorithm for Unsteady Potential Flow About Airfoils

Implicit finite-difference scheme efficiently computes unsteady potential flow about airfoils. Formulation uses density and velocity potential as dependent variables. Conservation form is retained to assure that shock wave location and speed are computed correctly. Scheme fills need for method to calculate efficiently unsteady potential flow about airfoils and to predict flutter and other unsteady aeroelastic phenomena in transonic flow regimes.

Chipman, R.↗
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 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 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 NTRS1986

Unsteady transonic flow calculations for wing-fuselage configurations

Unsteady transonic flow calculations are presented for wing-fuselage configurations. Calculations are performed by extending the XTRAN3S unsteady transonic small-disturbance code to allow the treatment of a fuselage. Details of the XTRAN3S fuselage modeling are discussed in the context of the small-disturbance equation. Transonic calculations are presented for three wing-fuselage configurations with leading edge sweep angles ranging from 0 deg to 46.76 deg. Simple bending and torsion modal oscillations of the wing are calculated. Sectional lift and moment coefficients for the wing-alone and wing-fuselage cases are compared and the effects of fuselage aerodynamic interference on the unsteady wing loading are revealed. Tabulated generalized aerodynamic forces used in flutter analyses, indicate small changes in the real in-phase component and as much as a 30% change in the imaginary component when the fuselage is included in the calculation. These changes result in a 2 to 5% increase in total magnitude and a several degree increase in phase.

Batina, J. T.↗
NASA NTRS1988

Calculation of unsteady Euler flows in turbomachinery using the linearized Euler equations

The model used by the present unsteady cascade flow calculation method is based on the linearized unsteady Euler equations and accounts for blade loading and geometry, shock motion, and wake motion. The steady flow that must be determined prior to the unsteady flow is ascertained by means of a Newton iteration procedure. A noteworthy feature of the procedure is the use of shock-fitting to determine steady and unsteady shock positions; the use of the Euler equations in conjunction with the Rankine-Hugoniot shock-jump conditions is found to directly model the generation of entropy and vorticity at shocks. Results are presented for both channel and cascade flows.

Hall, Kenneth C.↗
NASA NTRS1985

Unsteady transonic flow calculations for interfering lifting surface configurations

Unsteady transonic flow calculations are presented for aerodynamically interfering lifting surface configurations. Calculations are performed by extending the XTRAN3S (Version 1.5) unsteady transonic small-disturbance code to allow the treatment of an additional lifting surface. The research was conducted as a first-step toward developing the capability to treat a complete flight vehicle. Grid generation procedures for swept tapered interfering lifting surface applications of XTRAN3S are described. Transonic calculations are presented for wing-tail and canard-wing configurations for several values of mean angle of attack. The effects of aerodynamic interference on transonic steady pressure distributions and steady and oscillatory spanwise lift distributions are demonstrated. Results due to wing, tail, or canard pitching motions are presented and discussed in detail.

Batina, J. T.↗
NASA NTRS1985

Unsteady transonic flow calculations for interfering lifting surface configurations

Unsteady transonic flow calculations are presented for aerodynamically interfering lifting surface configurations. Calculations are performed by extending the XTRAN3S (Version 1.5) unsteady transonic small-disturbance code to allow the treatment of an additional lifting surface. The research was conducted as a first-step toward developing the capability to treat a complete flight vehicle. Grid generation procedures for swept tapered interfering lifting surface applications of XTRAN3S are described. Transonic calculations are presented for wing-tail and canard-wing configurations for several values of mean angle of attack. The effects of aerodynamic interference on transonic steady pressure distributions and steady and oscillatory spanwise lift distributions are demonstrated. Results due to wing, tail, or canard pitching motions are presented and discussed in detail.

Batina, J. T.↗
NASA NTRS1978

Three dimensional steady and unsteady asymmetric flow past wings of arbitrary planforms

The nonlinear discrete vortex method is extended to treat the problem of asymmetric flows past a wing with leading edge separation, including steady and unsteady flows. The problem is formulated in terms of a body fixed frame of reference and the nonlinear-discrete vortex method is modified accordingly. Although the method is general, only examples of flows past delta wings are presented due to the availability of experimental data as well as approximate theories. Comparison of results with experimental results for a delta wing undergoing a steady rolling motion at zero angle of attack demonstrate the superiority of the present method over existing approximate theories in obtaining highly accurate loads. Numerical results for yawed wings at large angles of attack are also presented. In all cases, total load coefficients, pressure distributions, and shapes of the free vortex sheets are shown.

Kandil, O. A.↗
NASA NTRS1983

Application of unsteady laminar triple-deck theory to viscous-inviscid interactions from an oscillating flap in supersonic and subsonic flow

Unsteady triple-deck theory is applied to analyze the local viscous-inviscid interaction of an idealized oscillating flap with a laminar boundary layer in either supersonic or subsonic external flow. For small flap amplitudes and small-to-moderate nondimensional frequencies, linearized analytical solutions by means of Fourier transformation are given for the pressure and shear distributions ahead of and behind the flap hinge. In the supersonic case, the predicted unsteady viscous effects reduce the pressure amplitude and spread it out upstream while causing it to lag the quasi-steady inviscid prediction (flap motion); this also results in an unsteady delay of incipient separation at the hinge to a higher flap amplitude. In contrast, the lowered unsteady viscous pressure distribution in purely subsonic flow slightly leads the flap motion, resulting in incipient separation at a smaller flap amplitude than the quasi-steady value. These trends are shown to be in qualitative agreement with available experimental data.

Inger, G. R.↗
NASA NTRS1994

Unsteady jet flow computation towards noise prediction

An attempt has been made to combine a wave solution method and an unsteady flow computation to produce an integrated aeroacoustic code to predict far-field jet noise. An axisymmetric subsonic jet is considered for this purpose. A fourth order space accurate Pade compact scheme is used for the unsteady Navier-Stokes solution. A Kirchhoff surface integral for the wave equation is employed through the use of an imaginary surface which is a circular cylinder enclosing the jet at a distance. Information such as pressure and its time and normal derivatives is provided on the surface. The sound prediction is performed side by side with the jet flow computation. Retarded time is also taken into consideration since the cylinder body is not acoustically compact. The far-field sound pressure has the directivity and spectra show that low frequency peaks shift toward higher frequency region as the observation angle increases from the jet flow axis.

Soh, Woo-Yung↗