Can fundamental shock-wave/boundary-layer interaction research be relevant to inlet aerodynamics?

The important influence of shock waves on supersonic inlet performance has led to much time and effort being expended in the area of shock wave/boundary layer interaction research (SWBLI) and SWBLI control. In this short review, the impact of SWBLIs on supersonic inlet aerodynamic research is discussed and is contrasted with fundamental SWBLI research. Inlet research focussed on internal flow performance is reviewed, based on the salient results, conclusions, and the limitations of such work. The role of fundamental SWBLI research in relation to supersonic inlet research is considered, and the possible positive impact of improving the link between fundamental SWBLI research and inlet design is considered. A simple flow-field is discussed which is thought to be able to simulate at least some more of the flow physics found in a typical inlet. A brief review of real inlet parameters is then given to help determine appropriate fundamental experimental parameters such as incoming Mach number, incoming boundary-layer thickness and subsonic difiuser angle. Copyright © 2012 by N. Titchener, H. Babinsky, and E. Loth.

The influence of geometry on jet plume development

Our recent efforts of using large-eddy simulation (LES) type methods to study complex and realistic geometry single stream and co-flow nozzle jets and acoustics are summarized in this paper. For the LES, since the solver being used tends towards having dissipative qualities, the subgrid scale (SGS) model is omitted, giving a numerical type LES (NLES). To overcome near wall streak resolution problems a near wall RANS (Reynolds averaged Navier-Stokes) model is smoothly blended in the LES making a hybrid RANS-NLES approach. Several complex nozzle geometries including the serrated (chevron) nozzle, realistic co-axial nozzles with eccentricity, pylon and wing-flap are discussed. The hybrid RANS-NLES simulations show encouraging predictions for the chevron jets. The chevrons are known to increase the high frequency noise at high polar angles, but decrease the low frequency noise at lower angles. The deflection effect of the potential core has an important mechanism of noise reduction. As for co-axial nozzles, the eccentricity, the pylon and the deployed wing-flap are shown to influence the flow development, especially the former to the length of potential core and the latter two having a significant impact on peak turbulence levels and spreading rates. The studies suggest that complex and real geometry effects are influential and should be taken into count when moving towards real engine simulations. © 2012 Elsevier Ltd. All rights reserved.

Large-eddy simulation of complex geometry jets

Computations are made for chevron and coflowing jet nozzles. The latter has a bypass ratio of 6:1. Also, unlike the chevron nozzle, the core flow is heated, making the inlet conditions reminiscent of those for a real engine. A large-eddy resolving approach is used with circa 12 × 10 6 cell meshes. Because the codes being used tend toward being dissipative the subgrid scale model is abandoned, giving what can be termed numerical large-eddy simulation. To overcome near-wall modeling problems a hybrid numerical large-eddy simulation-Reynolds-averaged Navier-Stokes related method is used. For y + ≤ 60 a Reynolds-averaged Navier-Stokes model is used. Blending between the two regions makes use of the differential Hamilton-Jabobi equation, an extension of the eikonal equation. For both nozzles, results show encouraging agreement with measurements of other workers. The eikonal equation is also used for ray tracing to explore the effect of the mean flow on acoustic ray trajectories, thus yielding a coherent solution strategy. © 2011 by Cambridge University.

THE SENSITIVITY OF 2D COMPRESSOR INCIDENCE RANGE TO IN-SERVICE GEOMETRIC VARIATION

The loss mechanisms which control 2D incidence range are discussed with an emphasis on determining which real in-service geometric variations will have the largest impact. For the majority of engine compressor blades (Minlet>0.55) both the negative and positive incidence limits are controlled by supersonic patches. It is shown that these patches are highly sensitive to the geometric variations close to, and around the leading edge. The variations used in this study were measured from newly manufactured as well as ex-service blades. Over most the high pressure compressor considered, it was shown that manufacture variations dominated. The first part of the paper shows that, despite large geometric variations (~10% of leading edge thickness), the incidence range responded in a linear way. The result of this is that the geometric variations have little effect on the mean incidence range of a row of blades. In the second part of the paper a region of the design space is identified where non-linear behavior can result in a 10% reduction in positive incidence range. The mechanism for this is reported and design guidelines for its avoidance offered. In the final part of the paper, the linear behavior at negative incidence and the transonic nature of the flow is exploited to design a robust asymmetric leading edge with a 5% increase in incidence range.

Normal shock interactions in rectangular channels

Experiments have been conducted to examine the mechanisms behind the coupling between corner separation and centreline separation when holding a normal shock in a rectangular channel. The study has focused on a M ∞ = 1.5 normal shock held in a wind tunnel with a parallel rectangular cross-section. The primary mechanism explaining the link between the corner separation size and the centreline separation appears to be the generation of compression waves which act to smear the adverse pressure gradient imposed upon other parts of the flow. In addition, the origin of the λ-foot leading leg appears to be depended upon the size of the corner separations. Experimental results indicate that the alteration of the λ-region, which occurs in the supersonic portion of the SBLI, is more important than the generation of any blockage in the subsonic region downstream of the shock wave. Copyright © 2012 by H. Babinsky, D.M.F. Burton.

Turbine blade tip heat transfer in low speed and high speed flows

In this paper, high and low speed tip flows are investigated for a high-pressure turbine blade. Previous experimental data are used to validate a CFD code, which is then used to study the tip heat transfer in high and low speed cascades. The results show that at engine representative Mach numbers the tip flow is predominantly transonic. Thus, compared to the low speed tip flow, the heat transfer is affected by reductions in both the heat transfer coefficient and the recovery temperature. The high Mach numbers in the tip region (M>1.5) lead to large local variations in recovery temperature. Significant changes in the heat transfer coefficient are also observed. These are due to changes in the structure of the tip flow at high speed. At high speeds, the pressure side corner separation bubble reattachment occurs through supersonic acceleration which halves the length of the bubble when the tip gap exit Mach number is increased from 0.1 to 1.0. In addition, shock/boundary-layer interactions within the tip gap lead to large changes in the tip boundary-layer thickness. These effects give rise to significant differences in the heat-transfer coefficient within the tip region compared to the low-speed tip flow. Compared to the low speed tip flow, the high speed tip flow is much less dominated by turbulent dissipation and is thus less sensitive to the choice of turbulence model. These results clearly demonstrate that blade tip heat transfer is a strong function of Mach number, an important implication when considering the use of low speed experimental testing and associated CFD validation in engine blade tip design. Copyright © 2009 by ASME.

Shock wave/boundary-layer interaction control using a combination of vortex generators and bleed

To investigate whether vortex generators can be an effective form of passive flow control an experimental investigation has been conducted in a small-scale wind tunnel. With specific emphasis on supersonic inlet applications flow separation was initiated using a combined terminal shock wave and subsonic diffuser: a configuration that has been developed as a part of a program to produce a more inlet-relevant flowfield in a small-scale wind tunnel than previous studies. When flow control was initially introduced little overall flow improvement was obtained as the losses tended to be redistributed instead of removed. It became apparent that there existed a strong coupling between the center-span flow and the corner flows. As a consequence, only when flow control was applied to both the corner flows and center-span flow was a significant flow improvement obtained. When corner suction and center-span vortex generators were employed in tandem separation was much reduced and wall-pressure and stagnation pressure were notably improved. As a result, when applied appropriately, it is thought that vortex generators do have the potential to reduce the dependence on boundary-layer bleed for the purpose of separation suppression. Copyright © 2012 by Neil Titchener and Holger Babinsky. Published by the American Institute of Aeronautics and Astronautics, Inc.

Fluid-structure interaction simulations of a tension-cone inflatable aerodynamic decelerator

A series of fluid-structure interaction simulations of an aerodynamic tension-cone supersonic decelerator prototype intended for large mass payload deployment in planetary explorations are discussed. The fluid-structure interaction computations combine large deformation analysis of thin shells with large-eddy simulation of compressible turbulent flows using a loosely coupled approach to enable quantification of the dynamics of the vehicle. The simulation results are compared with experiments carried out at the NASA Glenn Research Center. Reasonably good agreement between the simulations and the experiment is observed throughout a deflation cycle. The simulations help to illuminate the details of the dynamic progressive buckling of the tension-cone decelerator that ultimately results in the collapse of the structure as the inflation pressure is decreased. Furthermore, the tension-cone decelerator exhibits a transient oscillatory behavior under impulsive loading that ultimately dies out. The frequency of these oscillations was determined to be related to the acoustic time scale in the compressed subsonic region between the bow shock and the structure. As shown, when the natural frequency of the structure and the frequency of the compressed subsonic region approximately match, the decelerator exhibits relatively large nonaxisymetric oscillations. The observed response appears to be a fluid-structure interaction resonance resulting from an acoustic chamber (pistonlike) mode exciting the structure. Copyright © 2013 by Christopher Porter, R. Mark Rennie, Eric J. Jumper.

A canonical normal SBLI flow relevant to external compression inlets

The normal shock wave / boundary layer interaction (normal SBLI) is important to the operation and performance of a supersonic inlet, and the normal SBLI is particularly prominent in external compression inlets. To improve our understanding of such interactions, it is helpful to make use of fundamental flows which capture the main elements of inlets, without resorting to the level of complexity and system integration associated with full-geometry inlets. In this paper, several fundamental fiow-fleld configurations have been considered as possible test cases to represent the normal SBLI aspects found in typical external compression inlets, and it was found that the spillage-diffuser more closely retains the basic flow features of an external compression inlet than the other configurations. In particular, this flow-fleld allows the normal shock Mach number as well as the amount and rate of subsonic diffusion to be all held approximately constant mid independent of the application of flow control. In addition, a survey of several external compression inlets was conducted to quantify the flow and geometric parameters of the spillage-diffuser relevant to actual inlets. The results indicated that such a flow may be especially relevant if the terminal Mach number is about 1.3 to 1.4, the confinement parameter is around 10%, the width around twice or three times the height, and with the area expansion just downstream of the shock on the conservative side of the stall limit for incompressible diffusers. © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Canonical normal shock wave/boundary-layer interaction flows relevant to external compression inlets

The normal shock wave/boundary-layer interaction is important to the operation and performance of a supersonic inlet, and the normal shock wave/boundary-layer interaction is particularly prominent in external compression inlets. To improve understanding of such interactions, it is helpful to make use of fundamental flows that capture the main elements of inlets, without resorting to the level of complexity and system integration associated with full-geometry inlets. In this paper, several fundamental flowfield configurations have been considered as possible test cases to represent the normal shock wave/boundary-layer interaction aspects found in typical external compression inlets, and it was found that the spillage diffuser more closely retains the basic flow features of an external compression inlet than the other configurations. In particular, this flowfield allows the normal shock Mach number as well as the amount and rate of subsonic diffusion to all be held approximately constant and independent of the application of flow control. In addition, a survey of several external compression inlets was conducted to quantify the flow and geometric parameters of the spillage diffuser relevant to actual inlets. The results indicated that such a flow may be especially relevant if the terminal Mach number is about 1.3 to 1.4, the confinement parameter is around 10%, and the width is around twice or three times the height. In addition, the area expansion downstream of the shock should be limited to the conservative side of incipient stall based on incompressible diffusers. Copyright © 2013 by the authors.

A computational fluid-structure interaction study of inflation cycles of a tension cone decelerator

Inflatable aerodynamic decelerators present potential advantages for planetary entry in missions of robotic and human exploration. The design of these structures face many engineering challenges, including complex deformable geometries, anisotropic material response, and coupled shockturbulence interactions. In this paper, we describe a comprehensive computational fluid-structure interaction study of an inflation cycle of a tension cone decelerator in supersonic flow and compare the simulations with earlier published experimental results. The aeroshell design and flow conditions closely match recent experiments conducted at Mach 2.5. The structural model is a 16-sided polygonal tension cone with seams between each segment. The computational model utilizes adaptive mesh refinement, large-eddy simulation, and shell mechanics with self-contact modeling to represent the flow and structure interaction. This study focuses on the dynamics of the structure as the inflation pressure varies gradually, and the behavior of forces experienced by the flexible and rigid (the payload capsule) structures. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Cabaret finite-difference schemes for the one-dimensional Euler equations

In the present paper we consider second order compact upwind schemes with a space split time derivative (CABARET) applied to one-dimensional compressible gas flows. As opposed to the conventional approach associated with incorporating adjacent space cells we use information from adjacent time layer to improve the solution accuracy. Taking the first order Roe scheme as the basis we develop a few higher (i.e. second within regions of smooth solutions) order accurate difference schemes. One of them (CABARET3) is formulated in a two-time-layer form, which makes it most simple and robust. Supersonic and subsonic shock-tube tests are used to compare the new schemes with several well-known second-order TVD schemes. In particular, it is shown that CABARET3 is notably more accurate than the standard second-order Roe scheme with MUSCL flux splitting.

The aeroacoustics of transonic helicopter blades

One of the most important issues facing the helicopter industry today is helicopter noise, in particular transonic rotor noise. It is the main factor limiting cruise speeds, and there is real demand for efficient and reliable prediction methods which can be used in the rotor design process. This paper considers the Ffowcs Williams-Hawkings equation applied to a permeable control surface. The surface is chosen to be as small as possible, while enclosing both the blade and any transonic flow regions. This allows the problematic quadrupole term to always be neglected, and requires only near field CFD input data. It is therefore less computationally intensive than existing prediction methods, and moreover retains the physical interpretation of the sources in terms of thickness, loading and shock-associated noise. A computer program has been developed which implements the permeable surface form of retarded time formulation. The program has been validated and subsequently used to validate an acoustic 2-D CFD code. It is fast and reliable for subsonic motion, but it is demonstrated that it cannot be used at high subsonic or supersonic speeds. A second computer program implementing a more general formulation has also been developed and is presently being validated. This general formulation can be applied at high subsonic and supersonic speeds, except under one specific condition. © 2002 by the author(s). Published by the American Institute of Aeronautics and Astronautics, Inc.

Differential equation based length scales to improve DES and RANS simulations

Surprisingly expensive to compute wall distances are still used in a range of key turbulence and peripheral physics models. Potentially economical, accuracy improving differential equation based distance algorithms are considered. These involve elliptic Poisson and hyperbolic natured Eikonal equation approaches. Numerical issues relating to non-orthogonal curvilinear grid solution of the latter are addressed. Eikonal extension to a Hamilton-Jacobi (HJ) equation is discussed. Use of this extension to improve turbulence model accuracy and, along with the Eikonal, enhance Detached Eddy Simulation (DES) techniques is considered. Application of the distance approaches is studied for various geometries. These include a plane channel flow with a wire at the centre, a wing-flap system, a jet with co-flow and a supersonic double-delta configuration. Although less accurate than the Eikonal, Poisson method based flow solutions are extremely close to those using a search procedure. For a moving grid case the Poisson method is found especially efficient. Results show the Eikonal equation can be solved on highly stretched, non-orthogonal, curvilinear grids. A key accuracy aspect is that metrics must be upwinded in the propagating front direction. The HJ equation is found to have qualitative turbulence model improving properties. © 2003 by P. G. Tucker.