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At least 217 records · Page 12

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NASA NTRS1984

Measurements of large scale disturbances in rectangular wall jets

Detailed measurements of the velocity field and wall surface pressure distribution of a tone-generating wall jet were obtained in the Ames 7- by 10-foot Wind Tunnel and anechoeic chamber. The jet issued from a 0.508 cm by 10.16 cm nozzle adjacent to a 3.81 cm wall, with an initially parabolic velocity profile at the exit. The jet was operated at Reynolds numbers, R, of 4910, 7740, and 9410. At R = 4910 and 7740, the flow appeared as a double-rowed vortex array which formed upstream of the edge and convected downstream without change in structure. At R = 9410, apparent vortex pairing occurred at the edge location. These flow characteristics were observed in both velocity field measurements and flow visualization records. The unsteady vorticity field of the jet was computed from phase-averaged hot-wire anemometer measurements of the velocity field. The associated acoustic field was estimated by the method of Green's functions, and was found to be similar to measurements of the sound field. These observations tend to support theoretical predictions that vortex shedding at the edge tends to cancel the effect of acoustic sources in other parts of the flow field.

Horne, W. C.↗

NASA NTRS1985

Scaling of airfoil self-noise using measured flow parameters

Data from an airfoil broadband self-noise study are reported. Attention here is restricted to two-dimensional models at zero angle of attack to the flow. The models include seven NACA 0012 airfoil sections and five flat plate sections with chordlengths ranging from 2.54 to 60.96 cm. Testing parameters include flow velocity to 71.3 m/s and boundary-layer turbulence through natural transition and by tripping. Detailed aerodynamic measurements are conducted in the near-wake of the sharp trailing edges. The noise spectra of the self-noise sources are determined by the use of a cross-spectral technique. The acoustic data are normalized using the measured aerodynamic parameters in order to evaluate a commonly used scaling law. An examination of the Reynolds number dependence of the normalized overall levels has revealed a useful scaling result. This result appears to quantify the transition between turbulent boundary-layer trailing-edge noise and laminar boundary-layer vortex shedding noise.

Brooks, T. F.↗

NASA NTRS1986

Aerodynamic Prediction for Supersonic Canard-Tail Missiles

LRCDM2 computer program developed to calculate pressure distribution at points on surfaces of complete supersonic missile. Missile comprises up to two finned sections attached to axisymmetric body of circular cross section. Includes effects of vortex shedding due to forebody and forward fins, providing more accurate rolling moments. LRCDM2 written in FORTRAN IV.

Dillenius, M. F.↗

NASA NTRS1985

Nonlinear problems in flight dynamics involving aerodynamic bifurcations

Aerodynamic bifurcation is defined as the replacement of an unstable equilibrium flow by a new stable equilibrium flow at a critical value of a parameter. A mathematical model of the aerodynamic contribution to the aircraft's equations of motion is amended to accommodate aerodynamic bifurcations. Important bifurcations such as, the onset of large-scale vortex shedding are defined. The amended mathematical model is capable of incorporating various forms of aerodynamic responses, including those associated with dynamic stall of airfoils.

Murray Tobak↗

NASA NTRS1987

Measurements of the unsteady flow field within the stator row of a transonic axial-flow fan. 1: Measurement and analysis technique

This two-part paper presents laser anemometer measurements of the unsteady velocity field within the stator row of a transonic axial-flow fan. The objective is to provide additional insight into unsteady blade-row interactions within high speed compressors which affect stage efficiency, energy transfer, and other design considerations. Part 1 describes the measurement and analysis techniques used for resolving the unsteady flow field features. The ensemble-average and variance of the measured velocities are used to identify the rotor wake generated and unresolved unsteadiness, respectively. (Rotor wake generated unsteadiness refers to the unsteadiness generated by the rotor wake velocity deficit and the term unresolved unsteadiness refers to all remaining contributions to unsteadiness such as vortex shedding, turbulence, mass flow fluctuations, etc.). A procedure for calculating auto and cross correlations of the rotor wake generated and unresolved unsteady velocity fluctuations is described. These unsteady-velocity correlations have significance since they also result from a decomposition of the Navier-Stokes equations. This decomposition of the Navier-Stokes equations resulting in the velocity correlations used to describe the unsteady velocity field will also be outlined in this paper.

Suder, K. L.↗

NASA NTRS1987

Measurements of the unsteady flow field within the stator row of a transonic axial-flow fan. Part 2: Results and discussion

Unsteady velocity field measurements made within the stator row of a transonic axial-flow fan are presented. Measurements were obtained at midspan for two different stator blade rows using a laser anemometer. The first stator row consists of double circular-arc airfoils with a solidity of 1.68. The second features controlled-diffusion airfoils with a solidity of 0.85. Both were tested at design-speed peak efficiency conditions. In addition, the controlled-diffusion stator was also tested at near stall conditions. The procedures developed here are used to identify the rotor wake generated and unresolved unsteadiness from the velocity measurements (rotor wake generated unsteadiness refers to the unsteadiness generated by the rotor wake velocity deficit and unresolved unsteadiness refers to all remaining unsteadiness which contributes to the spread in the distribution of velocities such as vortex shedding, turbulence, etc.). Auto and cross correlations of these unsteady velocity fluctuations are presented to show their relative magnitude and spatial distributions. Amplification and attenuation of both rotor wake generated and unresolved unsteadiness are shown to occur within the stator blade passage.

Hathaway, M. D.↗

NASA NTRS1987

Numerical simulation of unsteady viscous flows

Most unsteady viscous flows may be grouped into two categories, i.e., forced and self-sustained oscillations. Examples of forced oscillations occur in turbomachinery and in internal combustion engines while self-sustained oscillations prevail in vortex shedding, inlet buzz, and wing flutter. Numerical simulation of these phenomena was achieved due to the advancement of vector processor computers. Recent progress in the simulation of unsteady viscous flows is addressed.

Hankey, Wilbur L.↗

NASA NTRS1986

Progress in tail rotor noise analysis

Helicopter tail rotor noise generated by interactions with the main rotor tip vortices and with the fuselage separation mean wake is investigated. The tail rotor blade-main rotor tip vortex interaction is modeled as an airfoil of infinite span cutting through a moving vortex. The present results are compared to the tail rotor loading and high speed thickness noise and are found to be significant. This noise mechanism is a function of the helicopter operating conditions and the location of the tail relative to the main rotor. Tail rotor self-generated noise due to turbulent vortex shedding from blunt trailing edges is also analyzed.

Chou, S.-T.↗

NASA NTRS1987

Numerical simulation of shock wave diffraction by TVD schemes

An upwind total variation diminishing (TVD) scheme and a predictor-corrector symmetric TVD scheme were used to numerically simulate the blast wave diffraction on a stationary object. The objective is to help design an optimum configuration so that lateral motion is minimized and at the same time vortex shedding and flow separation are reduced during a blast wave encounter. Results are presented for a generic configuration for both a coarse grid and a fine grid to illustrate the global and local diffraction flow fields. Numerical experiments for the shock wave reflection on a wedge are also included to validate the current approach. Numerical study indicated that these TVD schemes are more stable and produced higher shock resolution than classical shock capturing methods such as the explicit MacCormack scheme.

Young, Victor Y. C.↗

NASA NTRS1987

Energetics of oscillating lifting surfaces using integral conservation laws

The energetics of oscillating flexible lifting surfaces in two and three dimensions is calculated by the use of integral conservation laws in inviscid incompressible flow for general and harmonic transverse oscillations. Total thrust is calculated from the momentum theorem and energy loss rate due to vortex shedding in the wake from the principle of conservation of mechanical energy. Total power required to maintain the oscillations and hydrodynamic efficiency are also determined. In two dimensions, the results are obtained in closed form. In three dimensions, the distribution of vorticity on the lifting surface is also required as input to the calculations. Thus, unsteady lifting-surface theory must be used as well. The analysis is applicable to oscillating lifting surfaces of arbitrary planform, aspect ratio, and reduced frequency and does not require calculation of the leading-edge thrust.

Ahmadi, Ali R.↗

NASA NTRS1986

Tactical missile aerodynamics

The present conference on tactical missile aerodynamics discusses autopilot-related aerodynamic design considerations, flow visualization methods' role in the study of high angle-of-attack aerodynamics, low aspect ratio wing behavior at high angle-of-attack, supersonic airbreathing propulsion system inlet design, missile bodies with noncircular cross section and bank-to-turn maneuvering capabilities, 'waverider' supersonic cruise missile concepts and design methods, asymmetric vortex sheding phenomena from bodies-of-revolution, and swept shock wave/boundary layer interaction phenomena. Also discussed are the assessment of aerodynamic drag in tactical missiles, the analysis of supersonic missile aerodynamic heating, the 'equivalent angle-of-attack' concept for engineering analysis, the vortex cloud model for body vortex shedding and tracking, paneling methods with vorticity effects and corrections for nonlinear compressibility, the application of supersonic full potential method to missile bodies, Euler space marching methods for missiles, three-dimensional missile boundary layers, and an analysis of exhaust plumes and their interaction with missile airframes.

Hemsch, Michael J.↗

NASA NTRS1988

Vortex Suppressors Reduce Probe Vibrations

Lightweight addition to probe reduces cyclical stress aerodynamically. Life expectancy of instrumentation probe for high-speed flow increased by addition of helical strakes to cylindrical shield surrounding probe. Strakes disrupt vortex shedding normally occuring in fast flows around cylindrical body.

Hill, Arthur J.↗

NASA NTRS1988

Numerical simulation of unsteady incompressible viscous flows in generalized coordinate systems

Several numerical solutions of the three dimensional unsteady incompressible Navier-Stokes equations in generalized coordinate systems are presented. The governing equations are discretized by finite volumes with special care to the accurate approximation of the geometric quantities. The unknowns are the pressure and the volume fluxes over the computational cell faces. This formulation results in a robust fractional step solution method for solving discrete equations. Although this method is formulated for the three dimensional case, only two dimensional unsteady results are given. Results are presented for a lid driven two dimensional cavity flow at Reynolds number of 10,000, and for the flow over a circular cylinder with vortex shedding for several Reynolds numbers in the range 100 less than Re less than 1000.

Rosenfeld, Moshe↗

NASA NTRS1987

Computation and analysis of a cylinder wake flow

The Karman vortex wake of a circular cylinder at low Reynolds number was computed by a time-accurate, two-dimensional compressible Navier-Stokes equation solver which uses the MacCormack predictor-corrector finite-difference scheme and a nonreflecting boundary condition on the outer flow boundary. The results from a large number of time steps were analyzed using Fast Fourier Transform techniques to identify the important frequency components for comparison with published experimental data. A strong low-frequency component was found below the vortex shedding frequency and not harmonically related to it. The experimentally discovered low-frequency fluctuations in the cylinder wake are considered possibly to be precursors to transition from laminar to turbulent flow conditions. The present finding of similar frequencies in a computed wake tends to confirm their existence as a real wake phenomenon. This computational work provides a complementary means to experimental investigations of wake phenomena.

Townsend, J. C.↗

NASA NTRS1987

Details of the computed flowfield over a circular cylinder at Reynolds number 1200

The application of an upwind-biased implicit approximate factorization Navier-Stokes algorithm to the unsteady impulsive start-up flow over a circular cylinder at Reynolds number 1200 is described. The complete form of the compressible Navier-Stokes equations is used, and the algorithm is second-order accurate in both space and time. The development with time of the shape and size of the separated vortical flow region is computed, as well as the time-variation of several boundary layer parameters and profile shapes. Computations, in general, show excellent agreement with experiment, although the present method predicts a more rapid onset of reversed flow on the cylinder than evidenced in experiment. The changes that the vortical region behind the cylinder undergoes as the symmetric flow transitions to periodic vortex shedding are discussed. The flow becomes periodic with a Strouhal frequency of 0.222, which compares well with the experimental value of approximately 0.21. The effect of grid density on the development of the unsteady flow is also shown.

Rumsey, C. L.↗

NASA NTRS1987

Measurements of the unsteady flow field within the stator row of a transonic axial-flow fan. I - Measurement and analysis technique

This two-part paper presents laser anemometer measurements of the unsteady velocity field within the stator row of a transonic axial-flow fan. The objective is to provide additional insight into unsteady blade-row interactions within highspeed compressors which affect stage efficiency, energy transfer, and other design considerations. Part 1 describes the measurement and analysis techniques used for resolving the unsteady flow field features. The ensemble-average and variance of the measured velocities are used to identify the rotor wake generated and unresolved unsteadiness, respectively. (Rotor wake generated unsteadiness refers to the unsteadiness generated by the rotor wake velocity deficit and the term unresolved unsteadiness refers to all remaining contributions to unsteadiness such as vortex shedding, turbulence, mass flow fluctutions, etc.). A procedure for calculating auto and cross correlations of the rotor wake generated and unresolved unsteady velocity fluctuations is described. These unsteady-velocity correlations have significance since they also result from a decomposition of the Navier-Stokes equations. This decomposition of the Navier-Stokes equations resulting in the velocity correlations used to describe the unsteady velocity field will also be outlined in this paper.

Suder, K. L.↗

NASA NTRS1987

Measurements of the unsteady flow field within the stator row of a transonic axial-flow fan. II - Results and discussion

Hathaway, M. D.↗

NASA NTRS1988

Some numerical and physical aspects of unsteady Navier-Stokes computations over airfoils using dynamic meshes

An upwind-biased implicit approximate factorization algorithm is applied to several unsteady flows on dynamic meshes. The thin-layer form of the compressible Navier-Stokes equations is used to solve both laminar and turbulent flows over airfoils pitching about the quarter chord. Numerical aspects of the solutions are investigated, including grid and time step effects. Two methods for determining fluxes - flux-vector splitting and flux-difference splitting - are compared. Flux-difference splitting predicts results more accurately than flux-vector splitting on a coarse mesh, but both methods agree on a fine mesh. Physical aspects of the computations are also examined. An equilibrium turbulent boundary layer model computes generally better unsteady results in comparison with experiment than a nonequilibrium model for the transonic case analyzed. Also, the size and location of the primary shed vortex for an airfoil pitching up at a constant rate is calculated in good agreement with experiment for two pitch rates.

Rumsey, Christopher L.↗