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At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Numerical simulation of the edge tone phenomenon

Time accurate Navier-Stokes computations were performed to study a class 2 (acoustic) whistle, the edge tone, and to gain knowledge of the vortex-acoustic coupling mechanisms driving production of these tones. Results were obtained by solving the full Navier-Stokes equations for laminar compressible air flow of a two dimensional jet issuing from a slit interacting with a wedge. Cases considered were determined by varying the distance from the slit to the wedge. Flow speed was kept constant at 1,750 cm/s as was the slit thickness of 0.1 cm, corresponding to conditions in the experiments of Brown. The analytical computations revealed edge tones to be present in four harmonic stages of jet flow instability over the wedge as the jet length was varied from 0.3 to 1.6 cm. Excellent agreement was obtained in all four edge tone stage cases between the present computational results and the experimentally obtained frequencies and flow visualization results of Brown. Specific edge tone generation phenomena and further confirmation of certain theories and empirical formulas concerning these phenomena were brought to light in this analytical simulation of edge tones.

Dougherty, N. S.↗
NASA NTRS1991

Nonlinear Instability of Hypersonic Flow past a Wedge

The nonlinear stability of a compressible flow past a wedge is investigated in the hypersonic limit. The analysis follows the ideas of a weakly nonlinear approach. Interest is focussed on Tollmien-Schlichting waves governed by a triple deck structure and it is found that the attached shock can profoundly affect the stability characteristics of the flow. In particular, it is shown that nonlinearity tends to have a stabilizing influence. The nonlinear evolution of the Tollmien-Schlichting mode is described in a number of asymptotic limits.

Seddougui, Sharon O.↗
NASA NTRS2010

An Analytical Explanation for the X-43A Flush Air Data Sensing System Pressure Mismatch Between Flight and Theory

Following the successful Mach 7 flight test of the X-43A, unexpectedly low pressures were measured by the aft set of the onboard Flush Air Data Sensing System s pressure ports. These in-flight aft port readings were significantly lower below Mach 3.5 than was predicted by theory. The same lower readings were also seen in the Mach 10 flight of the X-43A and in wind-tunnel data. The pre-flight predictions were developed based on 2-dimensional wedge flow, which fails to predict some of the significant 3-dimensional flow features in this geometry at lower Mach numbers. Using Volterra s solution to the wave equation as a starting point, a three-dimensional finite wedge approximation to flow over the X-43A forebody is presented. The surface pressures from this approximation compare favorably with the measured wind tunnel and flight data at speeds of Mach 2.5 and 3.

Ellsworth, Joel C.↗
NASA NTRS1991

Development of a CFD Code for Analysis of Fluid Dynamic Forces in Seals

The aim is to develop a 3-D computational fluid dynamics (CFD) code for the analysis of fluid flow in cylindrical seals and evaluation of the dynamic forces on the seals. This code is expected to serve as a scientific tool for detailed flow analysis as well as a check for the accuracy of the 2D industrial codes. The features necessary in the CFD code are outlined. The initial focus was to develop or modify and implement new techniques and physical models. These include collocated grid formulation, rotating coordinate frames and moving grid formulation. Other advanced numerical techniques include higher order spatial and temporal differencing and an efficient linear equation solver. These techniques were implemented in a 2D flow solver for initial testing. Several benchmark test cases were computed using the 2D code, and the results of these were compared to analytical solutions or experimental data to check the accuracy. Tests presented here include planar wedge flow, flow due to an enclosed rotor, and flow in a 2D seal with a whirling rotor. Comparisons between numerical and experimental results for an annular seal and a 7-cavity labyrinth seal are also included.

Athavale, Mahesh M.↗
NASA NTRS1972

Computer program for calculating laminar and turbulent boundary layer development in compressible flow

A computer program is described which performs a numerical integration of the equations of motion for a compressible two-dimensional boundary layer. Boundary layer calculations may be carried out for both laminar and turbulent flow for arbitrary Reynolds number and free stream Mach number distribution on planar or axisymmetric bodies with wall heating or cooling, longitudinal wall curvature, wall suction or blowing, and a rough or a smooth wall. A variety of options are available as initial conditions. The program can generate laminar initial conditions such as Falkner-Skan similarity solutions (so that initial wedge flows can be simulated including Blasius or stagnation point flow) or approximate equilibrium turbulent profiles. Alternatively, initial profile input data can be utilized.

Herring, H. J.↗
NASA NTRS2019

CFD Simulations of the IHF Arc-Jet Flow: 9-Inch Nozzle, Flow Surveys, LEAF Wedge Calibration Data

This paper reports computational analyses and flow characterization studies in a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using a wedge model placed in a free jet downstream of new 9-inch diameter conical nozzle in the Ames 60-MW Interaction Heating Facility. Both the nozzle and wedge model were specifically designed for testing in the new Laser-Enhanced Arc-jet Facility. Data were obtained using stagnation calorimeters and wedge models placed downstream of the nozzle exit. Two instrumented wedge calibration plates were used: one water-cooled and the other RCG-coated tile plate. Experimental surveys of arc-jet test flow with pitot and heat flux probes were also performed at three arc-heater conditions, providing assessment of the flow uniformity and valuable data for the flow characterization. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, including the models tested, and comparisons with the experimental measurements. By taking into account nonuniform total enthalpy and mass flux profiles at the nozzle inlet as well as the expansion waves emanating from the nozzle exit and their effects on the model flowfields, these simulations approximately reproduce the probe survey data and predict the wedge model surface pressure and heat flux measurements.

Arc-jets↗
NASA NTRS1989

Calculation of winged-body-like flow fields using an implicit upwind space-marching code

The Mach-8 flow past a 60-deg swept fin mounted on a 12-degree ramp has been simulated using a parabolized Navier-Stokes solver employing an upwind algorithm in order to investigate the interference patterns that develop when a bow shock impinges on a wing shock. Good agreement with experimental data is found downstream of the interaction for each of three meshes of varying grid point density and computed surface pressure and heat transfer. In the wedge flow region, some disagreement with experimental data is noted for both pressure and heat transfer.

Lawrence, Scott L.↗
NASA NTRS1988

Hypersonic turbulent wall boundary layer computations

The Baldwin-Lomax (1978) algebraic turbulence model was modified for hypersonic flow conditions. Two coefficients in the outer-layer eddy-viscosity model were determined as functions of Mach number and temperature ratio. By matching the solutions from the Baldwin-Lomax model to those from the Cebeci-Smith (1974) model for a flat plate at hypersonic speed, the new values of the coefficients were obtained. The results show that the values of C(cp) and C(kleb) are functions of both Mach number and wall temperature ratio. The C(cp) and C(kleb) variations with Mach number and wall temperature were used for the calculations of both a 4-deg wedge flow at Mach 18 and an axisymmetric Mach 20 nozzle flow. The Navier-Stokes equations with thin-layer approximation were solved for the above hypersonic flow conditions and the results were compared with existing experimental data. The agreement between the numerical solutions and the existing experimental data were good. The modified Baldwin-Lomax model thus is useful in the computations of hypersonic flows.

Kim, S. C.↗
NASA NTRS1988

Hypersonic turbulent wall boundary layer computations

The Baldwin-Lomax algebraic turbulence model was modified for hypersonic flow conditions. Two coefficients in the outer layer eddy viscosity model were determined as functions of Mach number and temperature ratio. By matching the solutions from the Baldwin-Lomax model to those from the Cebeci-Smith model for a flat plate at hypersonic speed, the new values of the coefficient were obtained. The results show that the values of C sub cp and C sub kleb are functions of both Mach number and wall temperature ratio. The C sub cp and C sub kleb variations with Mach number and wall temperature were used for the calculations of both a 4 deg wedge flow at Mach 18 and an axisymmetric Mach 20 nozzle flow. The Navier-Stokes equations with thin layer approximation were solved for the above hypersonic flow conditions and the results were compared with existing experimental data. The agreement between the numerical solutions and the existing experimental data were good. The modified Baldwin-Lomax model thus is useful in the computations of hypersonic flows.

Kim, S. C.↗
NASA NTRS1974

A comparison of numerical solutions to the inviscid equations of fluid motion

The second-order MacCormack method and the third-order Rusanov and Kutler-Warming-Lomax methods are applied to the inviscid 1-d Burgers' equation, wedge flow, and the problem of shock reflection from a rigid boundary. The numerical solution in each case is compared to the exact solution and the quantitative estimates of accuracy are obtained. Results of this study show that the third-order Kutler-Warming-Lomax or the Rusanov methods tuned for minimum dissipation or minimum dispersion provide the most accurate solution in each of the examples considered.

Anderson, D. A.↗
NASA NTRS2009

Turbulent Wing-Leading-Edge Correlation Assessment for the Shuttle Orbiter

This study was conducted in support of the Orbiter damage assessment activity that takes place for each Shuttle mission since STS-107 (STS - Space Transportation System). As part of the damage assessment activity, the state of boundary layer (laminar or turbulent) during reentry needs to be estimated in order to define the aerothermal environment on the Orbiter. Premature turbulence on the wing leading edge (WLE) is possible if a surface irregularity promotes early transition and the resulting turbulent wedge flow contaminates the WLE flow. The objective of this analysis is to develop a criterion to determine if and when the flow along the WLE experiences turbulent heating given an incoming turbulent boundary layer that contaminates the attachment line. The data to be analyzed were all obtained as part of the MH-13 Space Shuttle Orbiter Aerothermodynamic Test conducted on a 1.8%-scale Orbiter model at Calspan/University of Buffalo Research Center in the Large Energy National Shock Tunnels facility. A rational framework was used to develop a means to assess the state of the WLE flow on the Orbiter during reentry given a contaminated attachment-line flow. Evidence of turbulent flow on the WLE has been recently documented for a few STS missions during the Orbiter s flight history, albeit late in the reentry trajectory. The criterion developed herein will be compared to these flight results.

King, Rudolph A.↗
NASA NTRS1985

Refined numerical solution of the transonic flow past a wedge

A numerical procedure combining the ideas of solving a modified difference equation and of adaptive mesh refinement is introduced. The numerical solution on a fixed grid is improved by using better approximations of the truncation error computed from local subdomain grid refinements. This technique is used to obtain refined solutions of steady, inviscid, transonic flow past a wedge. The effects of truncation error on the pressure distribution, wave drag, sonic line, and shock position are investigated. By comparing the pressure drag on the wedge and wave drag due to the shocks, a supersonic-to-supersonic shock originating from the wedge shoulder is confirmed.

Liang, S.-M.↗
NASA NTRS1997

The Effect of Three-Dimensional Freestream Disturbances on the Supersonic Flow Past a Wedge

The interaction between a shock wave (attached to a wedge) and small amplitude, three-dimensional disturbances of a uniform, supersonic, freestream flow are investigated. The paper extends the two-dimensional study of Duck et al, through the use of vector potentials, which render the problem tractable by the same techniques as in the two-dimensional case, in particular by expansion of the solution by means of a Fourier-Bessel series, in appropriately chosen coordinates. Results are presented for specific classes of freestream disturbances, and the study shows conclusively that the shock is stable to all classes of disturbances (i.e. time periodic perturbations to the shock do not grow downstream), provided the flow downstream of the shock is supersonic (loosely corresponding to the weak shock solution). This is shown from our numerical results and also by asymptotic analysis of the Fourier-Bessel series, valid far downstream of the shock.

Duck, Peter W.↗
NASA NTRS1976

An experimental investigation of supersonic flow past a wedge-cylinder configuration

An experimental investigation of supersonic flow past double-wedge configurations was conducted. Over the range of geometries tested, it was found that, while theoretical solutions both for a Type V pattern and for a Type VI pattern could be generated for a particular flow condition (as defined by the geometry and the free-stream conditions), the weaker, Type VI pattern was observed experimentally. More rigorous flow-field solutions were developed for the flow along the wing leading-edge. Solutions were developed for the three-dimensional flow in the plane of symmetry of a swept cylinder (which represented the wing leading-edge) which was mounted on a wedge (which generated the "bow" shock wave). A numerical code was developed using integral techniques to calculate the flow in the shock layer upstream of the interaction region (i.e., near the wing root). Heat transfer rates were calculated for various free stream conditions. The present investigation was undertaken to examine the effects of crossflow on the resultant flow-field and to verify the flow model used in theoretical calculations.

Barnette, D. W.↗
NASA NTRS1988

Direct simulation of three-dimensional hypersonic flow about intersecting blunt wedges

The general three-dimensional direct simulation Monte Carlo method is used to study the hypersonic flow around two blunt wedges that intersect at a 90 deg angle. Results are obtained for the transitional flow regimes found at 85 and 100 km altitude with a reentry velocity of 7.5 km/s. The disturbance field in front of the double-wedge body is found to be larger than that produced by a single wedge. Surface pressures and flow densities are higher near the wedge intersection, whereas surface heating and shear streses are greater at locations removed from the corner. Results also show that three-dimensional flow structure occurs only near the wedge corner, and that the flow monotonically approaches the limiting two-dimensional wedge flow case in the spanwise direction.

Celenligil, M. Cevdet↗
NASA NTRS1983

The wedge hot-film anemometer in supersonic flow

A commercial wedge hot-film probe is studied to determine its heat transfer response in transonic to low supersonic flows of high unit Reynolds number. The results of this study show that its response in this flow regime differs from the response of cylindrical type sensors. Whereas the cylindrical sensor has the same sensitivity to velocity as to density for free-stream Mach numbers exceeding 1.3, the wedge probe sensitivity to velocity is always greater than its sensitivity to density over the entire flow regime. This property requires determination of three fluctuation components due to density, velocity, and temperature, in a transonic or supersonic turbulent flow. Sensitivity equations are derived based on the observed behavior of the wedge probe. Both the durability and the frequency response of the probe are excellent, the square wave insertion test indicating frequency response near 130 kHz. The directional response of the probe at sonic speed is poor and requires further examination before Reynolds stress measurements are attempted with dual sensor probes.

Seiner, J. M.↗