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NASA NTRS1985

Supercritical airfoil drag reduction by passive shock wave/boundary layer control in the Mach number range .75 to .90

Airfoils operating in the transonic region are subject to large increases in drag due to shock wave/boundary layer interactions. The concept of passive shock wave/boundary layer control seeks drag reduction by placing a thin cavity with a porous top surface at the airfoil chordwise position where a shock wave would normally occur. The higher pressure behind the shock wave circulates flow through the cavity to the lower pressure ahead of the shock wave. The effects from this circulation prevent boundary layer separation and reduce entropy increases through the shock wave. In this investigation this concept is studied at a freestream Mach number range of .75 and .90. The Mach number distributions over the model, the wake impact pressure surveys used to determine profile drag and schlieren photographs for 2.8 percent porosity and solid airfoil cases are presented and compared. Results indicate that the profile drag coefficient can be reduced by as much as 40 percent through the use of this passive drag control system.

Nagamatsu, H. T.↗

NASA NTRS2010

Analysis of Nozzle Jet Plume Effects on Sonic Boom Signature

An axisymmetric full Navier-Stokes computational fluid dynamics (CFD) study was conducted to examine nozzle exhaust jet plume effects on the sonic boom signature of a supersonic aircraft. A simplified axisymmetric nozzle geometry, representative of the nozzle on the NASA Dryden NF-15B Lift and Nozzle Change Effects on Tail Shock (LaNCETS) research airplane, was considered. The highly underexpanded nozzle flow is found to provide significantly more reduction in the tail shock strength in the sonic boom N-wave pressure signature than perfectly expanded and overexpanded nozzle flows. A tail shock train in the sonic boom signature, similar to what was observed in the LaNCETS flight data, is observed for the highly underexpanded nozzle flow. The CFD results provide a detailed description of the nozzle flow physics involved in the LaNCETS nozzle at different nozzle expansion conditions and help in interpreting LaNCETS flight data as well as in the eventual CFD analysis of a full LaNCETS aircraft. The current study also provided important information on proper modeling of the LaNCETS aircraft nozzle. The primary objective of the current CFD research effort was to support the LaNCETS flight research data analysis effort by studying the detailed nozzle exhaust jet plume s imperfect expansion effects on the sonic boom signature of a supersonic aircraft. Figure 1 illustrates the primary flow physics present in the interaction between the exhaust jet plume shock and the sonic boom coming off of an axisymmetric body in supersonic flight. The steeper tail shock from highly expanded jet plume reduces the dip of the sonic boom N-wave signature. A structured finite-volume compressible full Navier-Stokes CFD code was used in the current study. This approach is not limited by the simplifying assumptions inherent in previous sonic boom analysis efforts. Also, this study was the first known jet plume sonic boom CFD study in which the full viscous nozzle flow field was modeled, without coupling to a sonic boom propagation analysis code, from the stagnation chamber of the nozzle to the far field external flow, taking into account all nonisentropic effects in the shocks, boundary layers, and free shear layers, and their interactions at distances up to 30 times the nozzle exit diameter from the jet centerline. A CFD solution is shown in Figure 2. The flow field is very complicated and multi-dimensional, with shock-shock and shockplume interactions. At the time of this reporting, a full three-dimensional CFD study was being conducted to evaluate the effects of nozzle vectoring on the aircraft tail shock strength.

Bui, Trong↗

NASA NTRS1987

Structure of supersonic turbulent flow past a sharp fin

The three-dimensional shock-wave/turbulent-boundary-layer interaction generated by a sharp fin is examined both experimentally and theoretically at Mach 3 for a fin angle 20 deg and Reynolds number 900,000. This study represents an extension of previous research for the sharp fin configuration to stronger interactions. The experimental data include surface pressure profiles, surface streamline patterns, and boundary-layer profiles of pitot pressure and yaw angle. Two separate theoretical approaches or 'models' were employed. Both models utilize the three-dimensional compressible Navier-Stokes equations in mass-averaged variables. The theoretical approach of Knight (1984 and 1985) employs the algebraic turbulent eddy-viscosity model of Baldwin and Lomax (1978), and the theoretical model of Horstman (1984) employs the two-equation turbulence model of Jones and Launder (1972) coupled with the wall function model of Viegas et al. (1985).

Knight, Doyle D.↗

NASA NTRS1988

An entropy and viscosity corrected potential method for rotor performance prediction

An unsteady Full-Potential Rotor code (FPR) has been enhanced with modifications directed at improving its drag prediction capability. The shock generated entropy has been included to provide solutions comparable to the Euler equations. A weakly interacted integral boundary layer has also been coupled to FPR in order to estimate skin-friction drag. Pressure distributions, shock positions, and drag comparisons are made with various data sets derived from two-dimensional airfoil, hovering, and advancing high speed rotor tests. In all these comparisons, the effect of the nonisentropic modification improves (i.e., weakens) the shock strength and wave drag. In addition, the boundary layer method yields reasonable estimates of skin-friction drag. Airfoil drag and hover torque data comparisons are excellent, as are predicted shock strength and positions for a high speed advancing rotor.

Bridgeman, John O.↗

NASA NTRS1998

Numerical Simulations of Unsteady Low-Reynolds-Number Flows Over the APEX Airfoil

Laminar and transitional separation bubbles are an important feature of low-Reynolds-number flows over airfoils. The separation bubbles are unsteady and have a significant impact on the aerodynamic properties of the airfoils. In this paper unsteady low-Reynolds-number separated flows over the APEX airfoil are calculated using a Navier-Stokes solver. The numerical results show the presence of unsteady separation bubbles in the flowfield. An analysis of the numerical results shows that flowfield disturbances are amplified significantly in the separation bubble, leading to periodic vortex shedding. A linear stability analysis of the separated boundary layer is performed and the results show that the dominant wavenumber and frequency in the numerical simulations agree with the most unstable wavenumber and frequency from the linear stability analysis. The numerical results also show the growth and interaction of disturbance waves in the separation bubble. For transonic flows over the APEX airfoil, the calculations show that the presence of shocks causes significant changes in the separation location and consequently, the overall flowfield.

Tatineni, Mahidhar↗

NASA NTRS2001

The CE/SE Method: a CFD Framework for the Challenges of the New Millennium

The space-time conservation element and solution element (CE/SE) method, which was originated and is continuously being developed at NASA Glenn Research Center, is a high-resolution, genuinely multidimensional and unstructured-mesh compatible numerical method for solving conservation laws. Since its inception in 1991, the CE/SE method has been used to obtain highly accurate numerical solutions for 1D, 2D and 3D flow problems involving shocks, contact discontinuities, acoustic waves, vortices, shock/acoustic waves/vortices interactions, shock/boundary layers interactions and chemical reactions. Without the aid of preconditioning or other special techniques, it has been applied to both steady and unsteady flows with speeds ranging from Mach number = 0.00288 to 10. In addition, the method has unique features that allow for (i) the use of very simple non-reflecting boundary conditions, and (ii) a unified wall boundary treatment for viscous and inviscid flows. The CE/SE method was developed with the conviction that, with a solid foundation in physics, a robust, coherent and accurate numerical framework can be built without involving overly complex mathematics. As a result, the method was constructed using a set of design principles that facilitate simplicity, robustness and accuracy. The most important among them are: (i) enforcing both local and global flux conservation in space and time, with flux evaluation at an interface being an integral part of the solution procedure and requiring no interpolation or extrapolation; (ii) unifying space and time and treating them as a single entity; and (iii) requiring that a numerical scheme be built from a nondissipative core scheme such that the numerical dissipation can be effectively controlled and, as a result, will not overwhelm the physical dissipation. Part I of the workshop will be devoted to a discussion of these principles along with a description of how the ID, 2D and 3D CE/SE schemes are constructed. In Part II, various applications of the CE/SE method, particularly those involving chemical reactions and acoustics, will be presented. The workshop will be concluded with a sketch of the future research directions.

Chang, Sin-Chung↗

NASA NTRS1987

Optical studies of shock generated transient supersonic base flows

A shock tube employing interferometric and schlieren techniques is used to study transient base flow phenomena following shock wave passage over two plane bluff bodies: a hemicircular cylinder and a cylinder with the Galileo Jovian probe profile. An attempt is made to understand the physics of transition from transient to steady state flow, and to provide code verification for a study employing the Illiac IV computer. Transient base flow interactions include a series of shock diffraction, regular, and Mach reflections, coupled with boundary layer development, separation, and recompression. Vorticity generation and transport underlie these features. The quantitative verification of the computer code includes comparisons of transient pressure and density fields, near wake geometries, and bow shock standoff distances.

Liang, P.-Y.↗

NASA NTRS1985

Conical similarity of shock/boundary-layer interactions generated by swept and unswept fins

A parametric experimental investigation has been made of the class of three-dimensional shock wave/turbulent boundary layer interactions generated by swept and unswept leading-edge fins. The fin sweepback angles were 0-65 deg at 5, 9, and 15 deg angles of attack. Two equilibrium two-dimensional turbulent boundary layers with a freestream Mach number of 2.95 and a Reynolds number of 6.3 x 10 to the 7th/m were used as incoming flow conditions. All of the resulting interactions were found to possess conical symmetry of the surface flow patterns and pressures outside of an initial inception zone. Further, these interactions were found to obey a simple conical similarity rule based on inviscid shock wave strength, irrespective of fin sweepback or angle of attack. This is one of the first demonstrations of similarity among three-dimensional interactions produced by geometrically dissimilar shock generators.

Settles, G. S.↗

NASA NTRS1983

Using second-sound shock waves to probe the intrinsic critical velocity of liquid helium II

A critical velocity truly intrinsic to liquid helium II is experimentally sought in the bulk fluid far from the apparatus walls. Termed the 'fundamental critical velocity,' it necessarily is caused by mutual interactions which operate between the two fluid components and which are activated at large relative velocities. It is argued that flow induced by second-sound shock waves provides the ideal means by which to activate and isolate the fundamental critical velocity from other extraneous fluid-wall interactions. Experimentally it is found that large-amplitude second-sound shock waves initiate a breakdown in the superfluidity of helium II, which is dramatically manifested as a limit to the maximum attainable shock strength. This breakdown is shown to be caused by a fundamental critical velocity. Secondary effects include boiling for ambient pressures near the saturated vapor pressure or the formation of helium I boundary layers at higher ambient pressures. When compared to the intrinsic critical velocity discovered in highly restricted geometries, the shock-induced critical velocity displays a similar temperature dependence and is the same order of magnitude.

Turner, T. N.↗

NASA NTRS1993

Supersonic flow and shock formation in turbine tip gaps

Shock formation due to overexpansion of supersonic flow at the inlet to the tip clearance gap of a turbomachine has been studied. As the flow enters the tip gap, it accelerates around the blade pressure-side corner creating a region of minimum static pressure. The 'free streamline' separates from the wall at the corner; and, for Mach numbers greater than about 1.3, it curves back to intersect the blade tip. At this point, the freestream flow is abruptly turned parallel to the surface, giving rise to an oblique shock. The results are consistent with compressible sharp-edged orifice flow calculations found in the literature and with the theory of oblique shock wave formation in supersonic flow over a wedge. For freestream Mach numbers of 1.4 to 1.8, wave angles are 43 to 54 deg, and turning angles are 9 to 20 deg; as the Mach number increases, the angle of turn also increases. It appears that in a turbine, after separating from the inlet corner, the flow reattaches on the blade tip and an oblique shock is formed at 0.4-1.4 tip gap heights into the clearance gap. The resulting shock-boundary layer interaction may contribute to further enhancement of already high heat transfer to the blade tip in this region. This in turn could lead to higher blade temperatures and adversely affect blade life and turbine efficiency.

Moore, John↗

NASA NTRS2015

Revisiting Turbulence Model Validation for High-Mach Number Axisymmetric Compression Corner Flows

Two axisymmetric shock-wave/boundary-layer interaction (SWBLI) cases are used to benchmark one- and two-equation Reynolds-averaged Navier-Stokes (RANS) turbulence models. This validation exercise was executed in the philosophy of the NASA Turbulence Modeling Resource and the AIAA Turbulence Model Benchmarking Working Group. Both SWBLI cases are from the experiments of Kussoy and Horstman for axisymmetric compression corner geometries with SWBLI inducing flares of 20 and 30 degrees, respectively. The freestream Mach number was approximately 7. The RANS closures examined are the Spalart-Allmaras one-equation model and the Menter family of kappa − omega two equation models including the Baseline and Shear Stress Transport formulations. The Wind-US and CFL3D RANS solvers are employed to simulate the SWBLI cases. Comparisons of RANS solutions to experimental data are made for a boundary layer survey plane just upstream of the SWBLI region. In the SWBLI region, comparisons of surface pressure and heat transfer are made. The effects of inflow modeling strategy, grid resolution, grid orthogonality, turbulent Prandtl number, and code-to-code variations are also addressed.

Shock Wave↗

NASA NTRS1987

Driving mechanism of unsteady separation shock motion in hypersonic interactive flow

Wall pressure fluctuations were measured under the steady separation shock waves in Mach 5 turbulent interactions induced by unswept circular cylinders on a flat plate. The wall temperature was adiabatic. A conditional sampling algorithm was developed to examine the statistics of the shock wave motion. The same algorithm was used to examine data taken in earlier studies in the Princeton University Mach 3 blowdown tunnel. In these earlier studies, hemicylindrically blunted fins of different leading-edge diameters were tested in boundary layers which developed on the tunnel floor and on a flat plate. A description of the algorithm, the reasons why it was developed and the sensitivity of the results to the threshold settings, are discussed. The results from the algorithm, together with cross correlations and power spectral density estimates suggests that the shock motion is driven by the low-frequency unsteadiness of the downstream separated, vortical flow.

Dolling, D. S.↗

NASA NTRS1994

Numerical Simulation of Shock-Induced Combustion Past Blunt Bodies Using Shock-Fitting Technique

Two-dimensional axisymmetric, reacting viscous flow over blunt projectiles is computed to study shock induced combustion at Mach 5.11 and Mach 6.46 in hydrogen-air mixture. A finite-difference, shock-fitting method is used to solve the complete set of Navier Stokes and species conservation equations. In this approach, the bow shock represents a boundary of the computational domain and is treated as a discontinuity across which Rankine-Hugoniot conditions are applied. All interior details of the flow such as compression waves, reaction front, and the wall boundary layer are captured automatically in the solution. Since shock-fitting approach reduces the amount of artificial dissipation, all the intricate details of the flow are captured much more clearly than has been possible with the shock-capturing approach. This has allowed an improved understanding of the physics of shock-induced combustion over blunt projectiles and the numerical results can now be explained more readily with one dimensional wave-interaction model than before.

Ahuja, J. K.↗

NASA NTRS1994

Numerical simulation of shock-induced combustion past blunt bodies using shock-fitting technique

Two-dimensional axisymmetric, reacting viscous flow over blunt projectiles is computed to study shock-induced combustion at Mach 5.11 and Mach 6.46 in hydrogen-air mixture. A finite-difference, shock-fitting method is used to solve the complete set of Navier-Stokes and species conservation equations. In this approach, the bow shock represents a boundary of the computational domain and is treated as a discontinuity across which Rankine-Hugoniot conditions are applied. All interior details of the flow such as compression waves, reaction front, and the wall boundary layer are captured automatically in the solution. Since shock-fitting approach reduces the amount of artificial dissipation, all the intricate details of the flow are captured much more clearly than has been possible with the shock-capturing approach. This has allowed an improved understanding of the physics of shock-induced combustion over blunt projectiles and the numerical results can now be explained more readily with one-dimensional wave-interaction model than before.

Ahuja, J. K.↗

NASA NTRSDate not supplied

Shock-Wave/Boundary-Layer Interaction (SWBLI) Experiments in the Presence of Transition-to-Turbulence on a Flat Plate Model in the NASA LaRC 31-Inch Mach 10 Air Tunnel

Historically, there have been a limited number of studies involving Transitional Shock-Wave/Boundary-Layer Interactions (XSWBLI) in hypersonic flows, the majority of which have been performed at Mach Numbers of 7.5 and below. Additionally, there have been even fewer such experiments which incorporate non-intrusive and optical diagnostics. NASA Langley Research Center (NASA LaRC), in collaboration with The University of Texas at San Antonio (UTSA) and The University of Tennessee Space Institute (UTSI), is currently planning and preparing for a series of experimental aerodynamic tests using the NASA LaRC 31-inch Mach 10 Air Tunnel. The experiments will primarily focus on studying the dynamics of Shock-Wave/Boundary-Layer Interactions (SWBLI) in the presence of transition (XSWBLI) and turbulence (SWTBLI) on a large, flat plate model at a freestream Mach number of 10. Besides continued interest in understanding the flow at high Mach numbers, executing such tests at Mach 10 avoids some known concerns and will aid in solving new problems. XSWBLI are a highly unsteady phenomena and the generation of XSWBLI at lower Mach numbers can prove to be a significant challenge. This concern will be somewhat mitigated in the Mach 10 flow, as boundary layers become increasingly stabilized, but some difficulty in achieving transition is still expected. Modelling SWBLI in the presence of transitional and turbulent boundary layers has also proven to be difficult, so a high-Mach number experiment which applies non-intrusive and optical diagnostics will aid in solving a unique problem as well as advancing the understanding and characterization of an aerodynamic surface at Mach 10.The NASA LaRC 31-inch Wind Tunnel incorporates a test section having a cross-section of 31-inches × 31-inches and provides optical access to the test section via three (3) ultra-violet (UV) transmitting windows on the top, side, and bottom. The Mach 10 operating conditions will consist of pressures (P0) ranging from 2.4 to 10MPa (348 to 1,450 psi) at a temperature (T0) of 1,000 K (1,800 °R). The NASA LaRC 31-inch wind tunnel will deliver 1-minute blow-down runs, of which 30-45 seconds will be consumed to reach the aforementioned test conditions. The pump down time between runs will be approximately 45-60 minutes, which will provide for 6-10 test runs per day over the course of a weeklong entry. The proposed flat plate model will be constructed of stainless steel with geometry consisting of a10-inch × 30-inch upper surface and a 2-inch thickness. The model will be designed to support multiple leading-edge inserts in order to examine blunt, round, and sharp leading edges at Mach 10. Adjustments will be made to manage boundary layer thickness as well as the strength and size of the shock interaction region for each test. Measurements and diagnostics will be performed though high-speed Schlieren, on-body high-speed pressure transducers, IR thermography, and oil flow visualization. Hypersonic vehicle applications have and will continue to emerge at the forefront of aerospace. The NASA-UTSA-UTSI team realizes the value of characterizing SWBLI, XSWBLI, and SWTBLI behaviors well beyond single-digit hypersonic Mach numbers. As hypersonic vehicle speed capabilities continue to increase, measurement and diagnostic methods for Mach numbers of 10 and beyond will provide the foundation for a firm understanding of flow field behavior which will directly influence the advancement of technologies towards the design and manufacturing of high-speed aerodynamic surfaces, controls, thermal protection systems, acoustic treatments, and structural components. The use of non-intrusive and optical diagnostic methods in such experimentation is pivotal in developing the visualization and empirical data necessary to advance the aforementioned technology areas.

Shockwave/Boundary-Layer Interactions↗

NASA NTRSDate not supplied

PIV Measurements of Shock-Wave/Boundary-Layer Interactions at Mach 2.5 in a Circular Test Section

The particle image velocimetry (PIV) measurement technique has been applied to make detailed measurements in an impinging shock-wave/boundary-layer interaction (SWBLI) flowfield in a test section of circular cross-section. By virtue of this test section geometry, the flowfields studied are free from effects of the complex coupling of sidewall and corner interaction into the primary impinging shock interaction typical of a rectangular test section. The cone-cylinder shock generator hardware was arranged in both coaxial alignment with the test section to generate axisymmetric SWBLIs and in offset positions to generate swept SWBLIs. The angle of the conical section was also varied to alter the strength of the impinging shock wave. The application of the PIV technique in this test section required meticulous design to provide adequate optical access without disturbing the SWBLI flow. Details of the stereoscopic PIV apparatus and the velocity measurements obtained for a Mach 2.5 freestream are presented. Key features of the mean velocity and turbulent stress fields are highlighted. This data set should enable validation of computational fluid dynamics results for these flowfields.

Shockwave Boundary Layer Interaction↗

NASA NTRSDate not supplied

PIV Measurements of Shock-Wave/Boundary-Layer Interactions at Mach 2.5 in a Circular Test Section

Shockwave Boundary Layer Interaction↗