Boundary-layer transition plays an important role in aerodynamic heating of blunt capsule shapes. Transition may occur from disturbances created by surface roughness and fluctuations of the freestr...

To address the hitherto unknown mechanism of boundary-layer transition on blunt reentry capsules, the role of roughness-induced disturbance growth on a spherical-section forebody is assessed via optimal transient growth theory and direct numerical simulations (DNS). Optimal transient-growth studies have been performed for the blunt capsule experiments at Mach 5.9 in the Hypersonic Ludwieg tube Braunschweig (HLB) of the Technische Universität Braunschweig, which included measurements behind a patch of controlled, distributed micron-sized surface roughness. Transient-growth results for the HLB capsule indicate similar trends as the corresponding numerical data for a Mach 6 experiment in the Actively Controlled Expansion (ACE) facility of the Texas A&M University (TAMU) at a lower Reynolds number. Both configurations indicate a similar dependence on surface temperature ratio, and more important, rather low values of maximum energy gain. DNS are performed for the conditions of the HLB experiment to understand the generation of stationary disturbances by the roughness patch and the accompanying evolution of unsteady perturbations. However, no evidence of either modal or nonmodal disturbance growth in the wake of the roughness patch is found in the DNS data; thus, the physical mechanism underlying the observed onset of transition still remains unknown.

The effect of uniformly distributed surface roughness on the onset of boundary-layer transition for a capsule geometry was experimentally investigated in a conventional hypersonic blowdown wind tunnel. The thin-walled scale model of the Orion Crew Exploration Vehicle used in this study was mounted at a 28 deg angle of attack in the Adjustable Contour Expansion wind tunnel. A test matrix of 24 runs was developed to experimentally estimate the location of transition and to relate the flow properties at this location and the roughness height to established transient growth scaling. Tests were conducted for four different Reynolds numbers at six quasi-uniform surface roughness height distributions. The tunnel was nominally set to Mach 6.0 and a total temperature of 430 K for each run. Infrared thermometry was used to measure the surface temperature distribution. Heat transfer rates generated through a series of surface temperature maps were processed into normalized Stanton number centerline plots. These plots directly indicated the boundary-layer transition location. Analysis of the results indicated that the experimental data matched the established transient growth scaling.

A scalar effective viscosity method is used to calculate the incompressible flow in the gap between two infinite, parallel smooth disks, one of them rotating and the other stationary. The flow is calculated for two values of the gap Reynolds number omega(s squared)/nu, namely 296 and 2852, where omega is the disk angular velocity, s the gap width, and nu the kinematic viscosity of the fluid; and over a range of radial Reynolds number omega(r squared)/nu from 0 to 10 million, where r is the radius from the axis of rotation. A numerical shooting technique is used to match the two boundary-layer flows arising on the disk and wall. Gradual transition is assumed for radial Reynolds numbers from 160,000 to 250,000. The results are in good agreement with Daily and Nece's measurements of velocity profiles and disk friction drag for enclosed rotating disks, and at sufficiently large radius the local skin friction approaches that of turbulent Couette flow.

Much has been learned about the physics underlying the transition process at supersonic and hypersonic speeds through years of analysis, experiment and computation. Generally, the application of this knowledge has been restricted to simple shapes like plates, cones and spherical bodies. However, flight reentry vehicles are in reality never simple. They typically are highly complex geometries flown at angle of attack so three-dimensional effects are very important, as are roughness effects due to surface features and/or ablation. This paper will review our present understanding of the physics of the transition process and look back at some of the recent flight test programs for their successes and failures. The goal of this paper is to develop rationale for new hypersonic boundary layer transition flight experiments. Motivations will be derived from both an inward look at what we believe constitutes a good flight test program as well as an outward review of the goals and objectives of some recent US based unclassified proposals and programs. As part of our recommendations, this paper will address the need for careful experimental work as per the guidelines enunciated years ago by the U.S. Transition Study Group. Following these guidelines is essential to obtaining reliable, usable data for allowing refinement of transition estimation techniques.

Surface roughness can have a profound effect on boundary layer transition and there is a vast empirical literature on this topic. However, the mechanisms responsible for transition with three-dimensional distributed roughness have been elusive. Various T-S based mechanisms have been investigated in the past but have been shown not to be applicable. More recently, the applicability of transient growth theory to roughness induced transition has been studied and a model developed that makes use of computational results based on the spatial transient growth theory. The resulting transition relations account for the separate roles of roughness height, Mach number and surface temperature level on the transition behavior. These results and also roughness receptivity issues are here reviewed.

Much has been learned about the physics underlying the transition process at supersonic and hypersonic speeds through years of analysis, experiment, and computation. The application has been principally to simple shapes like plates, cones, and spherical nose tips. But the shapes of the new entry vehicles are not simple. They will invariably be at angle of attack, and so three-dimensional effects will be very important, as will roughness effects due to ablation. This paper will review the physics basis of our present understanding of the transition process. Further, because of the complex geometries, it will address the need for careful experimental work as per the guidelines enunciated years ago by the U.S. Transition Study Group. Following these guidelines is essential to obtaining reliable, usable data for use in refining transition estimation techniques.

usually Reθ/Me = const. Also since Reθ varies as the square root of a length Reynolds number, any scatter in Reθ is greatly amplified when inverted to get a length transition Reynolds number or a length to transition. There is no apparent physics basis for this form of transition correlation. In this paper, the roles of modal (T-S, crossflow, Gortler, etc) and non-modal (transient growth) mechanisms on the transition process will be described. For modal disturbances, the physics based transition estimations are done by e methods. On the other hand, transient growth methods have been successfully applied in cases where the transition is due to surface roughness. These methods take into account many of the factors that can affect transition such as pressure gradient, surface temperature level, suction or blowing, surface roughness, etc. They also can suggest design changes that can lead to transition delay. Successful

Optimal disturbances for the supersonic flow past a sharp cone are computed to assess the effects due to flow divergence. This geometry is chosen because previously published studies on compressible optimal perturbations for flat plate and sphere could not isolate the influence of divergence alone, as many factors characterized the growth of disturbances on the sphere (How divergence, pressure gradient, centrifugal forces, and dependence of the edge parameters on the local Mach number). Flow-divergence effects result in the presence of an optimal distance from the cone tip for which the optimal gain is the largest possible, showing that divergence effects are stronger in the proximity of the cone tip. By properly rescaling the gain, wave number, and streamwise coordinate, due to the fact that the boundary-layer thickness on the sharp cone is root 3 thinner than the one over the flat plate, it is found that both the gain and the wave number compare fairly well. Moreover, results for the sharp cone collapse into those for the-flat plate when the initial location for the computation tends to the final one and when the azimuthal wave number is very large. Results show also that a cold wall enhances transient growth.

In the present work we revise results of transient growth in compressible boundary layers (∞at plate and sphere) to consider the complete Mack energy norm at the outlet, without the assumption that the out∞ow perturbation is comprised solely of streaky structures. Optimal perturbations are still in the form of counter-rotating streamwise vortices and this justifles the choice of the scaling in the governing equations. A strong efiect of the complete (full) energy norm at the outlet is found for the ∞at plate in supersonic regimes. No signiflcant efiects of the choice of the outlet norm can be appreciated for the sphere, in the range of parameters that are relevant to wind tunnel testing or ∞ight conditions.

Optimal steady perturbations in the boundary layer over a sphere are considered within the scope of the parabolized stability equations (PSE). The flow parameters at the edge of the boundary layer correspond to a high-speed free-stream flow of a calorically perfect gas, and the boundary layer velocity and temperature profiles are obtained using the local-similarity approximation. The governing PSE equations are derived from the linearized Navier-Stokes equations in spherical coordinates within the scope of the concept of optimal perturbations as streamwise vortices. Analysis of the transient growth phenomenon revealed that an increase of the sphere radius leads to an increase of the transient growth, and that the transient growth effect is stronger in the vicinity of the stagnation point. Similarly to previous results for compressible boundary layers, cooling of the wall destabilizes the boundary layer flow.

Significant research has been underway for several years in NASA Glenn Research Center's nozzle branch to develop advanced computational methods for simulating turbulent flows in exhaust nozzles. The primary efforts of this research have concentrated on improving our ability to calculate the turbulent mixing layers that dominate flows both in the exhaust systems of modern-day aircraft and in those of hypersonic vehicles under development. As part of these efforts, a hybrid numerical method was recently developed to simulate such turbulent mixing layers. The method developed here is intended for configurations in which a dominant structural feature provides an unsteady mechanism to drive the turbulent development in the mixing layer. Interest in Large Eddy Simulation (LES) methods have increased in recent years, but applying an LES method to calculate the wide range of turbulent scales from small eddies in the wall-bounded regions to large eddies in the mixing region is not yet possible with current computers. As a result, the hybrid method developed here uses a Reynolds-averaged Navier-Stokes (RANS) procedure to calculate wall-bounded regions entering a mixing section and uses a LES procedure to calculate the mixing-dominated regions. A numerical technique was developed to enable the use of the hybrid RANS-LES method on stretched, non-Cartesian grids. With this technique, closure for the RANS equations is obtained by using the Cebeci-Smith algebraic turbulence model in conjunction with the wall-function approach of Ota and Goldberg. The LES equations are closed using the Smagorinsky subgrid scale model. Although the function of the Cebeci-Smith model to replace all of the turbulent stresses is quite different from that of the Smagorinsky subgrid model, which only replaces the small subgrid turbulent stresses, both are eddy viscosity models and both are derived at least in part from mixing-length theory. The similar formulation of these two models enables the RANS and LES equations to be solved with a single solution scheme and computational grid. The hybrid RANS-LES method has been applied to a benchmark compressible mixing layer experiment in which two isolated supersonic streams, separated by a splitter plate, provide the flows to a constant-area mixing section. Although the configuration is largely two dimensional in nature, three-dimensional calculations were found to be necessary to enable disturbances to develop in three spatial directions and to transition to turbulence. The flow in the initial part of the mixing section consists of a periodic vortex shedding downstream of the splitter plate trailing edge. This organized vortex shedding then rapidly transitions to a turbulent structure, which is very similar to the flow development observed in the experiments. Although the qualitative nature of the large-scale turbulent development in the entire mixing section is captured well by the LES part of the current hybrid method, further efforts are planned to directly calculate a greater portion of the turbulence spectrum and to limit the subgrid scale modeling to only the very small scales. This will be accomplished by the use of higher accuracy solution schemes and more powerful computers, measured both in speed and memory capabilities.