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At least 145 records · Page 8

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DOE OSTI2021

Influence of turbulence intensity on vortex pattern for a rigid cylinder with turbulence stimulation in flow induced oscillations

The effect of ambient turbulence intensity on the flow induced oscillations of rigid circular cylinders with symmetric and asymmetric local turbulence stimulation is studied. Cylinders oscillate in one degree of freedom transversely to a steady uniform flow. Two-dimensional unsteady Reynolds-Average Navier-Stokes equations with the Spalart-Allmaras turbulence model are used to solve the problem numerically. Three freestream turbulence intensity values (0.2%, 1%, 5%) are used to study the influence of turbulence intensity on vortex pattern. Simulation results are compared with experimental data measured in the Marine Renewable Energy Laboratory of the University of Michigan in the range of 30,000 ≤ Re ≤ 110,000. The amplitude ratio, lift coefficient, vortex modes, and the interactions between vortices and cylinders are observed and discussed. In conclusion, major conclusions are: (a) Vortex patterns strongly depend on the freestream ambient turbulence intensity. (b) Low turbulence intensity can generate multi-vortex patterns at high Re. (c) High turbulence intensity results in fewer separations of a shed vortex into multiple vortices inducing emergence of elongated vortex patterns.

42 ENGINEERING↗

NASA NTRS2012

Study of Convective Flow Effects in Endwall Casing Treatments in Transonic Compressor Rotors

The unsteady convective flow effects in a transonic compressor rotor with a circumferential-groove casing treatment are investigated in this paper. Experimental results show that the circumferential-groove casing treatment increases the compressor stall margin by almost 50% for the current transonic compressor rotor. Steady flow simulation of the current casing treatment, however, yields only a 15% gain in stall margin. The flow field at near-stall operation is highly unsteady due to several self-induced flow phenomena. These include shock oscillation, vortex shedding at the trailing edge, and interaction between the passage shock and the tip clearance vortex. The primary focus of the current investigation is to assess the effects of flow unsteadiness and unsteady flow convection on the circumferential-groove casing treatment. Unsteady Reynolds-averaged Navier-Stokes (URANS) and Large Eddy Simulation (LES) techniques were applied in addition to steady Reynolds-averaged Navier-Stokes (RANS) to simulate the flow field at near-stall operation and to determine changes in stall margin. The current investigation reveals that unsteady flow effects are as important as steady flow effects on the performance of the circumferential grooves casing treatment in extending the stall margin of the current transonic compressor rotor. The primary unsteady flow mechanism is unsteady flow injection from the grooves into the main flow near the casing. Flows moving into and out of the grooves are caused due to local pressure difference near the grooves. As the pressure field becomes transient due to self-induced flow oscillation, flow injection from the grooves also becomes unsteady. The unsteady flow simulation shows that this unsteady flow injection from the grooves is substantial and contributes significantly to extending the compressor stall margin. Unsteady flows into and out of the grooves have as large a role as steady flows in the circumferential grooves. While the circumferential-groove casing treatment seems to be a steady flow device, unsteady flow effects should be included to accurately assess its performance as the flow is transient at near-stall operation.

Hah, Chunill↗

NASA NTRS1991

Development of a perturbation generator for vortex stability studies

Theory predicts vortex instability when subjected to certain types of disturbances. It was desired to build a device which could introduce controlled velocity perturbations into a trailing line vortex in order to study the effects on stability. A perturbation generator was designed and manufactured which can be attached to the centerbody of an airfoil type vortex generator. Details of design tests and manufacturing of the perturbation generator are presented. The device produced controlled perturbation with frequencies in excess of 250 Hz. Preliminary testing and evaluation of the perturbation generator performance was conducted in a 4 inch cylindrical pipe. Observations of vortex shedding frequencies from a centerbody were measured. Further evaluation with the perturbation generator attached to the vortex generator in a 2 x 3 foot wind tunnel were also conducted. Hot-wire anemometry was used to confirm the perturbation generator's ability to introduce controlled frequency fluctuations. Comparison of the energy levels of the disturbances in the vortex core was made between locations 42 chord lengths and 15 chord lengths downstream.

Riester, J. E.↗

NASA NTRS1954

On the Development of Turbulent Wakes from Vortex Streets

Wake development behind circular cylinders at Reynolds numbers from 40 to 10,000 was investigated in a low-speed wind tunnel. Standard hot-wire techniques were used to study the velocity fluctuations. The Reynolds number range of periodic vortex shedding is divided into two distinct subranges. At r=40 to 150, called the stable range, regular vortex streets are formed and no turbulent velocity fluctuations accompany the periodic formation of vortices. The range r=150 to 300 is a transition range to a regime called the irregular range, in which turbulent velocity fluctuations accompany the periodic formation of vortices. The turbulence is initiated by laminar-turbulent transition in the free layers which spring from the separation points on the cylinder. The transition first occurs in the range r=150 to 300. Spectrum and statistical measurements were made to study the velocity fluctuations.

Roshko, Anatol↗

NASA NTRS1994

Three Dimensional Compressible Turbulent Flow Computations for a Diffusing S-Duct With/Without Vortex Generators

Numerical investigations on a diffusing S-duct with/without vortex generators and a straight duct with vortex generators are presented. The investigation consists of solving the full three-dimensional unsteady compressible mass averaged Navier-Stokes equations. An implicit finite volume lower-upper time marching code (RPLUS3D) has been employed and modified. A three-dimensional Baldwin-Lomax turbulence model has been modified in conjunction with the flow physics. A model for the analysis of vortex generators in a fully viscous subsonic internal flow is evaluated. A vortical structure for modeling the shed vortex is used as a source term in the computation domain. The injected vortex paths in the straight duct are compared with the analysis by two kinds of prediction models. The flow structure by the vortex generators are investigated along the duct. Computed results of the flow in a circular diffusing S-duct provide an understanding of the flow structure within a typical engine inlet system. These are compared with the experimental wall static-pressure, static- and total-pressure field, and secondary velocity profiles. Additionally, boundary layer thickness, skin friction values, and velocity profiles in wall coordinates are presented. In order to investigate the effect of vortex generators, various vortex strengths are examined in this study. The total-pressure recovery and distortion coefficients are obtained at the exit of the S-duct. The numerical results clearly depict the interaction between the low velocity flow by the flow separation and the injected vortices.

Cho, Soo-Yong↗

NASA NTRS1974

Beading and spiking phenomena in the M551 metals melting experiment

A study was made regarding the beading and spiking phenomena observed in the M551 metals melting experiment conducted during the Skylab I mission in June 1973. An analysis was made of the beading phenomenon based on the Karman vortex shedding theory. The results tend to support the hypothesis that beading which occurred in the stainless and tantalum samples was a Karman vortex street formation. A dynamic model of cavity oscillation is discussed to explain the spiking phenomenon which was observed in the stainless steel and tantalum samples. Calculations of spiking frequency indicate that the intensity of spiking depends primarily on the vapor pressure and surface tension properties of the material, and is only slightly affected by the level of gravitation acceleration.

Fan, C.↗

NASA NTRS1991

Control of unsteady separated flow associated with the dynamic pitching of airfoils

Although studies have been done to understand the dependence of parameters for the occurrence of deep stall, studies to control the flow for sustaining lift for a longer time has been little. To sustain the lift for a longer time, an understanding of the development of the flow over the airfoil is essential. Studies at high speed are required to study how the flow behavior is dictated by the effects of compressibility. When the airfoil is pitched up in ramp motion or during the upstroke of an oscillatory cycle, the flow development on the upper surface of the airfoil and the formation of the vortex dictates the increase in lift behavior. Vortex shedding past the training edge decreases the lift. It is not clear what is the mechanism associated with the unsteady separation and vortex formation in present unsteady environment. To develop any flow control device, to suppress the vortex formation or delay separation, it is important that this mechanism be properly understood. The research activities directed toward understanding these questions are presented and the results are summarized.

Ahmed, Sajeer↗

NASA NTRS1977

Vortex interactions and decay in aircraft wakes

The dynamic interactions of aircraft wake vortices are investigated using both inviscid and viscous models. The phenomenon of vortex merging resulting in the rapid aging of a vortex wake is examined in detail. It is shown that the redistribution of vorticity from convection and diffusion during merging is a mechanism effective in reducing the hazard of a wake. Inviscid computations show that the merging phenomenon may be sensitive to small changes in spanwise load distribution and that the fuselage vortex shed from the wing-fuselage junction can play a significant role in promoting merging of wing-tip and flap vortices. Vortex-wake-merging computations using a second-order closure model of turbulent transport indicate that a low-hazard wake occurs when the generating aircraft flap and wing-tip vortices of the same strength and sign. This optimum is achieved when the flap vortex is located outboard approximately 40% of the distance to the tip vortex.

Bilanin, A. J.↗

NASA NTRS1977

Simplified unsteady aerodynamic concepts, with application to parameter estimation

A simplified aerodynamic force model based on the physical principle of Prandtl's lifting line theory and trailing vortex concept has been developed to account for unsteadiness in the aircraft dynamics. The wake is assumed to be compressed to a single shed vortex element of appropriate strength moving downstream at a speed sufficient to approximate the Wagner function. Results are presented illustrating the ability of the simplified theory to duplicate exact solutions in unsteady aerodynamics. Further, consideration is given to the utility of the model in a parameter identification application.

Wells, W. R.↗

NASA NTRS1976

Summary of some recent studies of subsonic vortex lift and parameters affecting the leading-edge vortex stability

Various subsonic configurations which develop vortex lift are examined herein. Comparisons are made with data and the combination of direct and indirect edge forces through use of the suction analogy. For most configurations, the use of the indirect or augmented vortex lift leads to improved agreement with data. The studies conducted showed that it is possible for the leading-edge vortex to exhibit bursting on a cropped wing and the wing aerodynamic characteristics not show the effect. It was further found that the distributions of leading-edge suction correlate well with the maintenance of vortex flow aerodynamic characteristics. Lastly, a method is presented for relating the initial value of circulation and axial velocity of the leading-edge shed vortex to the wing geometry.

Lamar, J. E.↗

NASA NTRS1995

The effects of vortex modeling on blade-vortex interaction noise prediction

The use of a blade vortex interaction noise prediction scheme, based on CAMRAD/JA, FPR and RAPP, quantifies the effects of errors and assumptions in the modeling of the helicopter's shed vortex on the acoustic predictions. CAMRAD/JA computes the wake geometry and inflow angles that are used in FPR to solve for the aerodynamic surface pressures. RAPP uses these surface pressures to predict the acoustic pressure. Both CAMRAD/JA and FPR utilize the Biot-Savart Law to determine the influence of the vortical velocities on the blade loading and both codes use an algebraic vortex model for the solid body rotation of the vortex core. Large changes in the specification of the vortex core size do not change the inplane wake geometry calculated by CAMRAD/JA and only slightly affect the out-of-plane wake geometry. However, the aerodynamic surface pressure calculated by FPR changes in both magnitude and character with small changes to the core size used by the FPR calculations. This in turn affects the acoustic predictions. Shifting the CAMRAD/JA wake geometry away from the rotor plane by 1/4 chord produces drastic changes in the acoustic predictions indicating that the prediction of acoustic pressure is extremely sensitive to the miss distance between the vortex and the blade and that this distance must be calculated as accurately as possible for acceptable noise predictions. The inclusion or exclusion of a vortex in the FPR-RAPP calculation allows for the determination of the relative importance of that vortex as a BVI noise source.

Gallman, Judith M.↗

NASA NTRS1998

Numerical Experiment with Time and Spatial Accuracy of Navier-Stokes Computation For Helicopter Problems

Helicopter flowfields are highly unsteady, nonlinear and three-dimensional. In forward flight and in hover, the rotor blades interact with the tip vortex and wake sheet developed by either itself or the other blades. This interaction, known as blade-vortex interactions (BVI), results in unsteady loading of the blades and can cause a distinctive acoustic signature. Accurate and cost-effective computational fluid dynamic solutions that capture blade-vortex interactions can help rotor designers and engineers to predict rotor performance and to develop designs for low acoustic signature. Such a predictive method must preserve a blade's shed vortex for several blade revolutions before being dissipated. A number of researchers have explored the requirements for this task. This paper will outline some new capabilities that have been added to the NASA Ames' OVERFLOW code to improve its overall accuracy for both vortex capturing and unsteady flows. To highlight these improvements, a number of case studies will be presented. These case studies consist of free convection of a 2-dimensional vortex, dynamically pitching 2-D airfoil including light-stall, and a full 3-D unsteady viscous solution of a helicopter rotor in forward flight In this study both central and upwind difference schemes are modified to be more accurate. Central difference scheme is chosen for this simulation because the flowfield is not dominated by strong shocks. The feature of shock-vortex interaction in such a flow is less important than the dominant blade-vortex interaction. The scheme is second-order accurate in time and solves the thin-layer Navier-Stokes equations in fully-implicit manner at each time-step. The spatial accuracy is either second and fourth-order central difference or third-order upwind difference using Roe-flux and MUSCLE scheme. This paper will highlight and demonstrate the methods for several sample cases and for a helicopter rotor. Preliminary computations on a rotor were performed by using this method and are in the process of documentation.

Ahmad, Jasim↗

NASA NTRS2021

The Near and Intermediate Wake of a Thick Flat Plate with a Circular Trailing Edge

The near and intermediate wake regions of a thick flat plate with a circular trailing edge (TE) are investigated with a direct numerical simulation (DNS). The focus is on the region x/D < 43.0(D is the diameter of the TE). The upper and lower separating boundary layers are both turbulent and are statistically identical; the resulting wake is symmetric in the mean. The present study is a continuation of an earlier investigation (Case AX). The earlier research dealt with shed vortex structure, in particular the persistence of the coherence of these vortices with increasing streamwise distance (except for infrequent, random disruption/rupture). In addition, across-centerline movement (ACM)of shed vortices, that is prevalent past x/D = 20.0 in a cylinder case at a comparable Reynolds number, was found to be essentially absent in Case AX. In the present investigation, an underlying cause for the occasional disruption/rupture of shed vortices is explored. Strong interactions between subsequent shed vortices in the trailing edge region is found to be an important contributor to this phenomenon. The distributions of phase-averaged streamwise and cross-stream components of intensity and the corresponding budget terms in the near wake are compared with those obtained in the intermediate wake here. In addition, the main features of the distribution of the phase-averaged velocity pressure-gradient term are compared with those of the production term, for the two intensity components. Except for a change in sign, the important features in the two budget terms are qualitatively similar for both components.

Direct Numerical Simulations of Flat Plate Wake↗

NASA NTRS1991

Time accurate Navier-Stokes computations of solid rocket motor internal field joint cavity flows

The low-speed flow of combustion gases in the Space Shuttle Solid Rocket Motor (SRM) following ignition are investigated with a time-accurate CFD technique. The flow of the combustion gases is examined in the forward joint section of the SRM at 80 s after ignition with attention given to the amplification of existing 'organ-pipe' modes of 15, 30, and 45 Hz resulting from frequency lock-in with vortex-shedding modes. The CFD code employed is a multizone compressible Navier-Stokes flow solver using a finite-volume, upwind-biased, total-variation-diminishing scheme. The vortex-generated oscillations generated by the inhibitor and the cavity at the forward field joint are described in terms of amplitude and frequency and related to the flow dynamics in the SRM. Oscillations are shown to occur spontaneously when the trailing edge of the cavity at the joint is sharp and can be amplified if the unburned portion of the inhibitor protrudes into the flow.

Farr, Rebecca A.↗

NASA NTRS1986

Application of a computational model for vortex generators in subsonic internal flows

A model for the analysis of vortex generators in a fully viscous subsonic internal flow is evaluated. A vorticity source term is used in a modified form of the Parabolized Navier-Stokes equations to model the shed vortex. Computed results are compared with idealized flow vortex paths, and with experimental data for vortex generators embedded in a thick turbulent boundary layer. The analysis is also compared with experimental data for a separated diffusing S-duct and for a diffusing S-duct with vortex generators. Quantitative comparisons are shown for the latter three cases. Emphasis is placed on verifying the ability of the model to predict global distortions in the flow field.

Kunik, W. G.↗

NASA NTRS1986

Application of a computational model for vortex generators in subsonic internal flows

Kunik, W. G.↗

NASA NTRS1985

Dynamic ground effects on a two-dimensional flat plate

The effect of time-variant vortex shedding is simulated by a sequence of discrete vortices convecting downstream in the wake of a two-dimensional flat plate whose lifting condition is modeled by means of the quasi-vortex lattice method. The boundary condition of this problem is specified in such a way that the tangency condition on the surface of the flat plate is satisfied; the boundary condition also takes into account the effect of airfoil motion relative to the ground. Significant lift changes are shown to occur, due to the dynamic ground effect, that are crucial in aircraft takeoff and landing transitions.

Chen, Y.-S.↗

NASA NTRS1979

Attenuation of sound in a low Mach number nozzle flow

The energy conversion mechanisms which govern the emission of low frequency sound from an axisymmetric jet pipe of arbitrary nozzle contraction ratio in the case of low Mach number nozzle flow are discussed. The energy of the incident sound which flows through the nozzle is used to maintain two distinct characteristic disturbances in the exterior fluid. First, there is the emitted radiation which has the directivity equivalent to that of a monopole-dipole combination. Second, essentially incompressible vortex waves are induced on the jet by vortex shedding from the lip of the nozzle and may involve the excitation of instability modes. Two linearized analytical models are examined to determine the partition of the emitted energy between the radiation field and the vortex waves. One of these is an exact linear theory in which the jet boundary is treated as a vortex sheet. The second model assumes that the width of the mean shear layer of the jet cannot be neglected. The results are discussed with reference to recent nozzle attenuation measurements.

Howe, M. S.↗