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Harloff, G. J.

Publications and source records attributed to Harloff, G. J..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

Numerical simulation of supersonic flow using a new analytical bleed boundary condition

A new analytical bleed boundary condition is used to compute flowfields for a strong oblique shock wave/boundary layer interaction with a baseline and three bleed rates at a freestream Mach number of 2.47 with an 8 deg shock generator. The computational results are compared to experimental Pitot pressure profiles and wall static pressures through the interaction region. An algebraic turbulence model is employed for the bleed and baseline cases, and a one equation model is also used for the baseline case where the boundary layer is separated.

Harloff, G. J.↗
NASA NTRS1994

Design and operation of a supersonic annular flow facility

Supersonic annular flow passages exist in propulsion applications that include dual combustion ramjet engines where a supersonic annular flow (the outer flow) mixes with a sonic (or supersonic) gas generator flow (the inner flow) in the shock expansion zone downstream of the gas generator nozzle exit. Other engine designs include components in the form of annular ducts whose cross-sectional area varies in the streamwise flow direction. In some of these configurations, it is necessary to support the outer shroud (cowl) by means of struts positioned between the cowl and centerbody. To investigate the distorting influence of these struts on the local flow structure, it is first necessary to ensure that the intrinsic flow without struts is free of wave reflections and the effects of upstream disturbances. It is also necessary to demonstrate that the intrinsic flow exhibits the characteristics of a well-defined turbulent boundary layer flow, so that changes in the local flow structure induced by the presence of struts can be interpreted properly. The purpose of this Note is to demonstrate that a supersonic flow facility that meets these objectives has been developed.

Williams, K. E.↗
NASA NTRS1993

Navier-Stokes analysis of three-dimensional S-ducts

Full 3D Navier-Stokes computational results are presented for compressible flows within nondiffusing and diffusing S-ducts. The present study provides an understanding of the performance characteristics of typical S-ducts with attached and separated flows and provides a frame of reference for future computational fluid dynamic studies of internal flows with strong secondary flows. The predicted results, which were obtained using both H- and O-grids, are compared with the experimental wall pressures, static and total pressure fields, and velocity vectors. In addition, computed boundary-layer thickness, velocity profiles in wall coordinates and skin friction values are presented. The inviscid contributions to the secondary flows are quantified. The S-duct entrance Mach number was 0.6, and the Reynolds number was 1.76 x 10 exp 6 based on the upstream duct diameter.

Harloff, G. J.↗
NASA NTRS1993

A numerical investigation of supersonic strut/endwall interactions in annular flow with varying strut thickness

A full 3D Navier-Stokes numerical investigation has been conducted of the shock-wave/boundary-layer flow interactions caused by four diamond-shaped struts, of varying thickness, in an annular duct with Mach 3 core flow and turbulent boundary-layers on both walls. Secondary flows caused by weak-to-strong interactions are examined in the vicinity of a strut which is bounded by curved endwalls. The duct endwall boundary-layer separated for the strongest interaction. The struts studied had maximum thickness-to-chord ratios of 0.125, 0.188, 0.250, and 0.500. The duct gap height is 0.7 strut chords, the duct inner-to-outer wall radius ratio is 0.7, and the Reynolds number is 3 x 10 exp 5 based on the strut chord length which was held constant for all interactions considered. The effects of strut thickness on the secondary flows are discussed, including: trajectories for the leading and trailing edge horseshoe vortices, strut/endwall corner vortices, and boundary-layer separation. The line of coalescence discussed in the literature, previously ascribed to boundary-layer separation, is shown to be caused by the leading edge horseshoe vortex convecting along the shock front.

Williams, K. E.↗
NASA NTRS1993

Experimental and numerical investigation of supersonic turbulent flow in an annular duct

Experimental and numerical results are presented for developing supersonic turbulent flow in an annular duct formed by a circular centerbody and outer shroud. The experimental results are based on data taken in a new flow facility that was designed to generate a shock-free, supersonic annular flow. Numerical computations were performed using the Baldwin-Lomax turbulence model for comparison with experimentally measured profiles. The results demonstrate that computed and measured profiles are in excellent agreement, so that studies can now be conducted of shock wave/boundary layer interaction phenomena within the duct, such as those induced by changes in downstream duct geometry or by the placement of struts between the duct walls.

Williams, K. E.↗
NASA NTRS1993

Investigation of a strut/endwall interaction in supersonic annular flow

A combined experimental and numerical investigation of strut/endwall interactions within an annular duct having a supersonic core flow has been conducted. Four diamond-shaped struts with a 7 deg half angle were positioned circumferentially equidistant within an annular duct having a gap height of 0.7 strut chords, and an inner-to-outer wall radius ratio of 0.7. Turbulent boundary layers exist on both inner and outer walls of the duct, but have not merged. The core flow upstream of the struts is uniform at a nominal Mach number of 3.0 and a Reynolds number of 3 x 10 exp 5 based on the strut chord length. Experimental results, which include Pitot pressure distributions within the flow field, static pressure distributions on the inner and outer walls of the duct, and oil flow visualization on the centerbody and strut, are presented and compared with CFD predictions. Secondary flows associated with the interactions are examined including the trajectories of the horseshoe vortices formed at the leading and trailing edges of the strut and the trajectories of the vortices formed in the corner of the strut/endwall intersection.

Williams, K. E.↗
NASA NTRS1992

Navier-Stokes analysis and experimental data comparison of compressible flow within ducts

Many aircraft employ ducts with centerline curvature or changing cross-sectional shape to join the engine with inlet and exhaust components. S-ducts convey air to the engine compressor from the intake and often decelerate the flow to achieve an acceptable Mach number at the engine compressor by increasing the cross-sectional area downstream. Circular-to-rectangular transition ducts are used on aircraft with rectangular exhaust nozzles to connect the engine and nozzle. To achieve maximum engine performance, the ducts should minimize flow total pressure loss and total pressure distortion at the duct exit. Changes in the curvature of the duct centerline or the duct cross-sectional shape give rise to streamline curvature which causes cross stream pressure gradients. Secondary flows can be caused by deflection of the transverse vorticity component of the boundary layer. This vortex tilting results in counter-rotating vortices. Additionally, the adverse streamwise pressure gradient caused by increasing cross-sectional area can lead to flow separation. Vortex pairs have been observed in the exit planes of both duct types. These vortices are due to secondary flows induced by pressure gradients resulting from streamline curvature. Regions of low total pressure are produced when the vortices convect boundary layer fluid into the main flow. The purpose of the present study is to predict the measured flow field in a diffusing S-duct and a circular-to-rectangular transition duct with a full Navier-Stokes computer program, PARC3D, and to compare the numerical predictions with new detailed experimental measurements. The work was undertaken to extend previous studies and to provide additional CFD validation data needed to help model flows with strong secondary flow and boundary layer separation. The S-duct computation extends the study of Smith et al, and Harloff et al, which concluded that the computation might be improved by using a finer grid and more advanced turbulence models. The present study compares results for both the Baldwin-Lomas and k-epsilon turbulence models and is conducted with a refined grid. For the transition duct, two inlet conditions were considered, the first with straight flow and the second with swirling flow. The first case permits examination of the effects of the geometric transition on the flow field, while the second case includes the rotational flow effect characteristic of a gas turbine engine.

Harloff, G. J.↗
NASA NTRS1992

Three-dimensional compressible turbulent computations for a diffusing S-duct

The purpose of the present study was to evaluate the capability of the computational fluid dynamics computer program PARC3D to model flow in a typical diffusing subsonic S-duct, with strong secondary flows. This evaluation is needed to provide confidence in the analysis of aircraft inlets, which have similar geometries. The performance predictions include total pressure profiles, static pressures, velocity profiles, boundary layer data, and skin friction data. Flow in the S-duct is subsonic, and the boundary layers are assumed to be turbulent. The results for both H and O grid solutions, are compared with existing test data.

Smith, C. F.↗
NASA NTRS1992

Three-dimensional compressible turbulent computations for a nondiffusing S-duct

The PARC3D code was used to compute the compressible turbulent flow within a three dimensional, nondiffusing S-duct. A frame of reference is provided for future computational fluid dynamics studies of internal flows with strong secondary flows and provides an understanding of the performance characteristics of a typical S-duct with attached flow. The predicted results, obtained with both H- and O-grids, are compared with the experimental wall pressure, static and total pressure fields, and velocity vectors. Additionally, computed boundary layer thickness, velocity profiles in wall coordinates, and skin friction values are presented.

Harloff, G. J.↗
NASA NTRS1991

Three-dimensional viscous flow computations of high area ratio nozzles for hypersonic propulsion

The PARC3D code was selected by the authors to analyze a variety of complex and high-speed flow configurations. Geometries considered for code validation include ramps and corner flows, which are characteristic of inlets and nozzles. Flows with Mach numbers of 3-14 were studied. Both two- and three-dimensional experimental data for shock-boundary-layer interaction were considered to validate the code. A detailed comparison of various flow parameters with available experimental data is presented; agreement between the solutions and the experimental data in terms of pitot pressure profiles, yaw-angle distributions, static pressures, and skin friction is found to be very good. In addition, two- and three-dimensional flow calculations were performed for a hypersonic nozzle. Comparison of the wall pressure results with the published solutions is made for the two-dimensional case.

Reddy, D. R.↗
NASA NTRS1990

Viscous three-dimensional analyses for nozzles for hypersonic propulsion

A Navier-Stokes computer code was validated using a number of two- and three-dimensional configurations for both laminar and turbulent flows. The validation data covers a range of freestream Mach numbers from 3 to 14, including wall pressures, velocity pressure, and skin friction. Nozzle flow fields computed for a generic scramjet nozzle from Mach 3 to 20, wall pressures, wall skin friction values, heat transfer values, and overall performance are presented. In addition, three-dimensional solutions obtained for two asymmetric, single expansion ramp nozzles at a pressure ratio of 10 consists of the internal expansion region in the converging/diverging sections and the external superonic exhaust in a quiescent ambient environment. The fundamental characteristics that were captured successfully include expansion fans; Mach wave reflections; mixing layers; and nonsymmetrical, multiple inviscid cell, supersonic exhausts. Comparison with experimental data for wall pressure distributions at the center planes shows good agreement.

Harloff, G. J.↗
NASA NTRS1990

Viscous three-dimensional analyses for nozzles for hypersonic propulsion

A Navier-Stokes computer code was validated using a number of two- and three-dimensional configurations for both laminar and turbulent flows. The validation data covers a range of freestream Mach numbers from 3 to 14, includes wall pressures, velocity profiles, and skin friction. Nozzle flow fields computed for a generic scramjet nozzle from Mach 3 to 20, wall pressures, wall skin friction values, heat transfer values, and overall performance are presented. In addition, three-dimensional solutions obtained for two asymmetric, single expansion ramp nozzles at a pressure ratio of 10 consists of the internal expansion region in the converging/diverging sections and the external supersonic exhaust in a quiescent ambient environment. The fundamental characteristics that were captured successfully include expansion fans; Mach wave reflections; mixing layers; and nonsymmetrical, multiple inviscid cell, supersonic exhausts. Comparison with experimental data for wall pressure distributions at the center planes shows good agreement.

Harloff, G. J.↗
NASA NTRS1988

Three-dimensional viscous flow computations of a circular jet in subsonic and supersonic cross flow

Three-dimensional viscous flow computations are presented for 90 deg injection angle jets in subsonic and supersonic cross flow. Comparisons with experimental data include jet centerline and vortex trajectories for the subsonic cross flow, and surface pressure measurement for the supersonic crossflow case. The vortices induced in the jet/freestream interaction are computed and illustrated. The vortices persist in subsonic flow and die out quickly in supersonic flow. The structure of the shocks in the unconfined supersonic flow is illustrated.

Harloff, G. J.↗
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 NTRS1988

Three-dimensional viscous flow computations of a circular jet in subsonic and supersonic cross flow

Three-dimensional viscous flow computations are presented for 90 deg. injection angle jets in subsonic and supersonic crossflow. Comparisons with experimental data include jet centerline and vortex trajectories for the subsonic crossflow, and surface pressure measurement for the supersonic crossflow case. The vortices induced in the jet/freestream interaction are computed and illustrated. The vortices persist in subsonic flow and die out quickly in supersonic flow. The structure of the shocks in the unconfined supersonic flow is illustrated.

Harloff, G. J.↗