Boundary Layer Edge Conditions Research Articles

Upstream Influence in Shock Wave/Transitional Boundary Layer Interactions at Mach 1.8

This work expands on previous efforts that observed high-intensity resonance in the dynamic shock motion generated by shock wave/boundary layer interactions with Mach 1.8 edge conditions when the inflow boundary layer is in transition. Specifically, this effort characterizes the behavior of a thickening of the boundary layer upstream of the separation shock location that appears to strongly correlate with the motion of the upstream influence shock; this work also verifies that the focused spectra observed in measurements with a cylinder on a flat plate are also observed in interactions generated by a blunt fin or on a cone with similar boundary layer edge conditions. The dynamic motion of the shock waves in the transitional interactions are characterized using high-speed schlieren and statistical information derived by tracking the location of various features of the interaction within the frame sequence.

Numerical Study of Supersonic Boundary-Layer Transition due to Sonic Wall Injection

The boundary-layer transition on a generic hypersonic forebody at Mach 6 downstream of a sonic wall injection is investigated by means of implicit large-eddy simulation, Fourier analysis, and dynamic mode decomposition of numerical data. An academic Mach 4.2 flat-plate configuration with boundary-layer edge conditions matching those of the forebody is considered. Two empirical correlations are proposed to predict the penetration of the underexpanded tripping jet into the boundary layer (that is, the Mach disk height) as a function of the pressure ratio. Different transition mechanisms are observed downstream of the injection port, depending on the jet penetration. The dynamic mode decomposition reveals the spatial structures, temporal frequencies, and growth rates of three-dimensional instability modes. These modes, arising from the jet counter-rotating vortices, are either varicose or sinusoidal, with the latter being more efficient than the former for tripping the boundary layer at low-injection pressure. Recent experiments in the Boeing/U.S. Air Force Office of Scientific Research Mach 6 quiet tunnel at Purdue University show similar transition patterns, but also some discrepancies. A more representative configuration including crossflow effects, and more resolved simulations, is needed for a quantitative comparison.

Global Mode Analysis in the L/T Transition of a Hypersonic Boundary Layer Forced by Wall Injection

The laminar-turbulent transition of a Mach 4.6 flat-plate boundary layer forced by wall injection is investigated using Implicit Large Eddy Simulation. The boundary layer edge conditions match those of a 1:3 scale hypersonic vehicle forebody at Mach number 6 in the T-313 blow-down wind tunnel of ITAM - Russian Academy of Science. Two different injection pressures are considered. The breakdown to turbulence is observed through instantaneous Q-criterion visualizations and wall pressure fluctuations. Global instability analysis is performed by the computation of Koopman modes from 2D pressure and density snapshots of the ILES flow field. Some coherence is found in the spatial structures and temporal frequencies of global modes computed on two different grids, indicating a weak dependency on grid resolution. Physical insight is gained, confirming different conjectured transition mechanisms for the different injection pressures.

Hypersonic Boundary-Layer Transition Forced by Wall Injection: A Numerical Study

The laminar-turbulent transition of a Mach 4.6 flat-plate boundary layer forced by a wall underexpanded jet is investigated using implicit large-eddy simulation based on a fifth-order weighted essentially nonoscillatory scheme. The boundary-layer edge conditions match those of a -scale hypersonic vehicle forebody at Mach 6 in the T-313 conventional blowdown wind tunnel of the Institute of Theoretical and Applied Mechanics, Russian Academy of Science at Novosibirsk. The effects of injection pressure and jet pulsation on the structure and the stability of the near-injection flow are investigated. Downstream breakdown to turbulence is analyzed through -criterion visualization, mean velocity profiles, and skin friction coefficients. Results show that a low-pressure injection is enough for effective tripping, and that a nonpulsed jet is paradoxically more efficient than a pulsed one.

PSE를 이용한 익형 위 경계층 안정성 해석

경계층 안정성 방정식인 PSE를 이용하여 익형 표면에 형성되는 경계층의 안정성 해석을 수행하였다. 압축성 비점성 유동 해석으로 경계층 가장자리 조건을 얻고, 일반좌표계에서의 압축성 경계층 방정식을 4차 정확도로 계산하여 층류 경계층 유동장을 얻었다. 층류 경계층 데이터를 PSE의 입력으로 하여 안정성 해석을 수행하고 교란의 증폭률을 얻어 안정성 특성을 고찰 하였다. 마하수 0.5의 NACA0012 및 HSNLF(1)-0213 익형에 대한 해석을 수행하여 교란 주파수에 따른 증폭률 및 위치에 따른 교란의 진폭 분포 특성을 파악하였다. 익형의 윗면과 아랫면에서 받음각에 따른 안정성 특성을 각각 증폭률의 크기와 주파수 범위에 대해 분석하였다. 또한 중립안정성 곡선, 마하수에 따른 안정성 특성을 살펴보았으며 익형의 종류에 따른 안정성 특성 차이를 분석하였다. In this study, stability analysis of boundary layers on airfoils is performed by using parabolized stability equations(PSE). Boundary layer edge conditions are obtained by compressible inviscid flow calculations. Mean velocity and temperature profiles of the laminar boundary layer are obtained by solving compressible boundary layer equations in generalized curvilinear coordinates with fourth order accuracy in the wall normal direction. Laminar mean flow profiles are used as input data for PSE to investigate growth rates of disturbances and stability characteristics. For the cases of boundary layer on NACA0012 and HSNLF(1)-0213 airfoils at Mach number 0.5, growth rates with respect to disturbance frequencies and profiles of disturbance amplitude are investigated. The effect of angle of attack on stability characteristics are examined at both upper and lower surfaces. The neutral stability curves, effect of Mach number and effect of airfoil section shapes are also analyzed.

Compressible Equilibrium Turbulent Boundary Layers at Nonadiabatic Wall Conditions

Experimental data have been obtained to define test conditions necessary for a compressible, naturally developing, equilibrium turbulent boundary layer. The model consisted of a flat plate that was tested in the NASA Langley Research Center 8-ft high temperature tunnel. For the present study, the nominal boundarylayer edge Mach numbers were 5.0 and 6.2. The local Reynolds number based on boundary-layer edge conditions and plate length ranged from 8 x 106 to 39 x 10 6. The nominal ratio of adiabatic wall to actual wall temperature was 5.4, due to both the high total temperature of 3300°R and the heat-sink characteristics of the model. This temperature ratio is considerably higher than that obtained in previous compressible boundary-layer studies and simulates hypersonic flight under highly cooled wall conditions, which are anticipated for hypersonic cruise vehicles. The data indicate that a momentum thickness Reynolds number of at least 4000 is required for an equilibrium turbulent boundary layer, which is in approximate agreement with incompressible studies. This criterion was determined by comparing the transformed wake strength of the boundary layer with the incompressible data of Coles. Also, several shape factors were examined and found to support the trend shown by the wake strength.

Reply by Author to K. F. Stetson

Stetson, K. F., Nosetip Bluntness Effects on Cone Frustum Boundary-Layer Transition in Hypersonic Flow, AIAA Paper 831763, July 1983. Reshotko, E. and Khan, M. M. S., ''Stability of the Boundary Layer on a Blunted Plate in Supersonic Flow, presented at IUTAM Symposium on Laminar-Turbulent Transition, Stuttgart, FRG, Sept. 1979. Wright, R. L. and Zoby, E. V., Flight Boundary-Layer Transition Measurements on a Slender Cone at Mach 20, AIAA Paper 77719, June 1977. Stetson, K. F., Thompson, E. R., Donaldson, J. C., and Siler, L. G., Laminar Boundary-Layer Stability Experiments on a Cone at Mach 8, Part 2: Blunt Cone, AIAA Paper 84-0006, Jan. 1984. Stetson, K. F., Thompson, E. R., Donaldson, J. C., and Siler, L. G., Laminar Boundary-Layer Stability Experiments on a Cone at Mach 8, Part 1: Sharp Cone, AIAA Paper 83-1761, July 1983. via the change in the boundary-layer edge conditions through the ''entropy wake generated by the curved bow shock. This change of inviscid flow characteristics is well predicted by the embedded Newtonian theory,' as has been demonstrated recently. Changes in wall temperature and unit Reynolds number do, of course, affect the viscous flow characteristics for both the sharp and the blunted cone. However, the effect of a geometric change from a sharp to a blunted cone is still inviscid in nature. Keeping the above in mind, a careful reading of Ref. 1 will provide all of the answers to the questions raised by Mr. Stetson, keeping in mind that the ratio dN/dB tells whether the conic frustum is long enough to experience the inviscid flow effects imprinted on the boundary layer by the entropy wake.

Calculation of Three-Dimensional Boundary Layers on Rotating Turbine Blades

An assessment has been made of the applicability of a three-dimensional boundary-layer analysis to the calculation of heat transfer and streamline flow patterns on the surfaces of both stationary and rotating turbine passages. In support of this effort, an analysis has been developed to calculate a general nonorthogonal surface coordinate system for arbitrary three-dimensional surfaces and also to calculate the boundary-layer edge conditions for compressible flow using the surface Euler equations and experimental pressure distributions. Using available experimental data to calibrate the method, calculations are presented for the endwall, and suction surfaces of a stationary cascade and for the pressure surface of a rotating turbine blade. The results strongly indicate that the three-dimensional boundary-layer analysis can give good predictions of the flow field and heat transfer on the pressure, suction, and endwall surfaces in a gas turbine passage.

Numerical Computation of Space Shuttle Laminar Heating and Surface Streamlines

Exact inviscid flowfield codes are used, together with a quasi-three-dimensional boundary-layer analysis, to provide estimates of the windward surface heating and streamline patterns of the shuttle orbiter vehicle under laminar flow conditions. The accuracy and limitations of the methods are established by comparison with available wind-tunnel experiments and with more exact numerical solutions for simple flows. Flight predictions are presented showing the effects of finite-rate (nonequilibrium) chemical reactions, and the effects of varying boundary-layer edge conditions due to the growth of the boundary layer into the inviscid flow (entropy-layer swallowing). Differences between flowfield predictions at wind-tunnel and nominal flight conditions are discussed.