A one-equation turbulence model for aerodynamic flows

Boundary-fitted coordinate systems for numerical solution of partial differential equations—A review

The intent of this study on micro-array flow control is to demonstrate the viability and economy of Response Surface Methodology (RSM) to determine optimal designs of micro-array actuation for controlling the shock wave turbulent boundary layer interactions within supersonic inlets and compare these concepts to conventional bleed performance. The term micro-array refers to micro-actuator arrays which have heights of 25 to 40 percent of the undisturbed supersonic boundary layer thickness. This study covers optimal control of shock wave turbulent boundary layer interactions using standard micro-vane, tapered micro-vane, and standard micro-ramp arrays at a free stream Mach number of 2.0. The effectiveness of the three micro-array devices was tested using a shock pressure rise induced by the 10 shock generator, which was sufficiently strong as to separate the turbulent supersonic boundary layer. The overall design purpose of the micro-arrays was to alter the properties of the supersonic boundary layer by introducing a cascade of counter-rotating micro-vortices in the near wall region. In this manner, the impact of the shock wave boundary layer (SWBL) interaction on the main flow field was minimized without boundary bleed.

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Optimal Control of Shock Wave Turbulent Boundary Layer Interactions Using Micro-Array Actuation

Recent research progress on unstart mechanism, detection and control of hypersonic inlet

The objective of this article is to report progress toward the development of an Euler analysis procedure for predicting the unstart tolerance of supersonic inlets. As an aid to understanding boundary condition issues, a one-dimensional, linear-analysis procedure was developed and used to analyze inlet unstart behavior. Using these results as a guide, an Euler analysis procedure was extended through the addition of a new bleed boundary condition, a new compressor face boundary condition, and an engine demand model for the simulation of unsteady inlet flows caused by freestream flow disturbances. Five unstart conditions were identified with the Euler analysis of the axisymmetric inlet for both 20- and 90-deg throat bleed configurations. Results show that both increases and decreases in temperature or velocity will unstart the inlet, whereas only pressure decreases will unstart the inlet. It was also found that 90-deg throat bleed improves the unstart tolerance relative to 20-deg throat bleed for freestream pressure decreases, temperature increases, and changes in velocity.

Prediction of supersonic inlet unstart caused by freestream disturbances

Computations were performed to investigate the physics of three-dimensi onal, shock-wave/boundary-layer interactions on a flat plate in which fluid in the boundary layer was bled through a circular hole into a plenum to control shock-wave induced flow separation. This study revealed two underlying mechanisms by which bleed holes can control shock-wave/boundary-layer interactions. It also showed how bleed-hole placement relative to where the incident shock wave impinges affects upstream, spanwise, and downstream influence lengths. This study is based on the ensemble-averaged, full compressible Navier-Stokes equations closed by the BaldwinLomax turbulence model. Solutions to these equations were obtained by an implicit, partially split, two-factored method with flux-vector splitting on a chimera overlapping grid.

Three-dimensional shock-wave/boundary-layer interactions with bleed

This numerical study investigates the effectiveness of bleed in controlling shock-wave/boundary-layer interactions on a flat plate with a focus on understanding how bleed-hole angle, presence of upstream and downstream bleed holes, and pressure ratio across bleed holes affect structure of barrier shock, surface pressure distribution, and bleed rate (in terms of flow coefficient). The bleed-hole angles investigated are 30 deg slanted and 90 deg normal, which give rise to two different types of barrier shocks. The influence of upstream and downstream bleed holes were investigated by studying the bleed process through an isolated hole and through three holes arranged in tandem along the streamwise direction. The plenum/freestream pressure ratios investigated range from 0.3 to 1.7, which produced choked and unchoked flows in the bleed holes. This study is based on the ensemble-averaged, full compressible Navier-Stokes equations closed by the Baldwin-Lomax model with solutions obtained by an implicit finite volume method on an overlapping Chimera grid.

Control of shock-wave/boundary-layer interactions by bleed

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.

Calculation of Unsteady Transonic Flow over an Airfoil

Computations were performed to investigate three-dimensional, shock-wave/bo undary-layer interactions on a flat plate with bleed through four staggered rows of circular holes that discharge into the same plenum The focus of the computations was to examine how bleed through rows of holes affect bleed rate and the pressure and Mach number distributions The effects of the following parameters on the flow were investigated: 1) with and without shock-wave impingement on the boundary layer and 2) spacings between bleed holes in the streamwise and spanwise directions Results show that just two rows of bleed holes arranged in a staggered fashion placed upstream of the incident shock are adequate in blocking the shock-induced adverse pressure gradient from propagating further upstream Results also show that the spacings between the centers of holes can exceed the hole diameter not just in the streamwise direction, but also in the spanwise direction, and still be able to control shock-induced flow separation This study is based on the ensemble-averaged, "full compressible" Navier-Stokes (N-S) equations closed by the Baldwin-Lomax algebraic turbulence model Solutions to the ensemble-averaged N-S equations were obtained by an implicit, partially split, two-factored algorithm with flux-vector splitting in the streamwise direction on a chimera overlapping grid

Shock-wave/boundary-layer interactions with bleed through rows of holes

A numerical study was conducted to investigate how bleed through a two-dimensional slot affects shock-wave induced, boundary-layer separation on a flat plate. This study is based on the ensemble-averaged, compressible, Navier-Stokes equations closed by the Baldwin-Lomax, algebraic turbulence model. The algorithm used to obtain solutions was the implicit, partially split, two-factored scheme of Steger. This study examined the effects of the following parameters in controlling shock-wave induced flow separation: location of slot in relation to where the incident shock wave impinged on the boundary layer, size of slot in relation to the boundary-layer thickness, number of slots, spacings between slots, and strength of the incident shock wave. This study also showed the nature of the very complex flowfield about the slot or slots and how the plenum affects the bleed process. The results of this study are relevant to problems where bleed is used to control shock-wave induced, boundary-layer separation (e.g., inside jet engine inlets and wind tunnels).

W. J. Chyu

Papers

Three-dimensional shock-wave/boundary-layer interactions with bleed

Control of shock-wave/boundary-layer interactions by bleed

Calculation of Unsteady Transonic Flow over an Airfoil

Shock-wave/boundary-layer interactions with bleed through rows of holes

Numerical Study of Shock-Wave/Boundary-Layer Interactions with Bleed