Published e ight data for boundary-layer transition at hypersonic speeds are surveyed. The survey is limited to measurements reported in the open literature and carried out at hypersonic and high-supersonic speeds, on vehicles for which ablation is believed to be negligible or small. The emphasis is on work that may be suitable for validation of advanced transition-estimation methods that are based on simulation of the physical mechanisms, such as e N , the parabolized stability equations, and direct numerical simulations. Brief discussions are presented for each report. Known comparisons to the advanced simulation methods are also presented. Nomenclature Me = Mach number at the boundary-layer edge Res = Reynolds number at transition onset, based on arc length from the leading edge and local conditions at the boundary-layer edge ReT = Reynolds number at transition onset, based on arc length and conditions at the boundary-layer edge ReT/ft = unit Reynolds number per foot, at the boundary-layer edge at transition onset Reµ = Reynolds number at transition, usually onset, based on momentum thickness and conditions at the boundary-layer edge Te = temperature at the boundary-layer edge Tr = recovery temperature at the wall Tw = wall temperature xT = arc length to transition onset, from the nose µc = cone half-angle, deg

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Flight Data for Boundary-Layer Transition at Hypersonic and Supersonic Speeds

Nomenclature a = constant; Fig. 1 C , C 0 = constants k = roughness element height, ft N k = average roughness element height, ft L = vehicle length, ft M = Mach number n = exponent; Fig. 1 N R = Poll’s transition parameter; Eq. (1) Reke = roughnessReynolds number based on height k and edge conditions Rekk = roughnessReynolds number based on height k and conditions at k Reμ = Reynolds number based on height μ and edge conditions U = velocity component parallel to test surface or velocity component perpendicularto attachment line, ft /s V = velocity component parallel to attachment line, ft/s X = generalized disturbanceparameter or axial coordinate along windward centerline x = coordinate perpendicular to attachment line, ft Y = generalized transition parameter ® = angle of attack, deg ± = smooth-wall laminar boundary-layer thickness, ft = Poll’s length scale [Eq. (2)], ft μ = smooth-wall laminar boundary-layermomentum thickness, ft 1 = viscosity, lbm/ft ¢ s o = kinematic viscosity, ft2/s 1⁄2 = density, lbm/ft

Review and synthesis of roughness-dominated transition correlations for reentry applications

Boundary-layer trip devices for the Hyper-X forebody have been experimentally examined in several wind tunnels.Fivedifferenttripconegurationswerecomparedinthreehypersonicfacilities:theNASALangleyResearch Center 20-Inch Mach 6 Air and 31-Inch Mach 10 Air tunnels and in the HYPULSE Reeected Shock Tunnel at the General Applied Sciences Laboratory. Heat-transfer distributions, utilizing the phosphor thermography and thin-elm techniques, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. Parametric variations include angles of attack of 0, 2, and 4 deg; Reynolds numbers based on model length of 1.2 ££ 10 6‐15.4 £ 10 6 ; and inlet cowl door simulated in both open and closed positions. Comparisons of boundary-layer transition as a result of discrete roughness elements have led to the selection of a trip coneguration for the Hyper-X Mach 7 eight vehicle.

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Hypersonic Boundary-Layer Trip Development for Hyper-X

An experimental investigation of the effects of two-dimensional roughness on hypersonic boundary layers was carried out at the JAXA 0.5 m hypersonic wind tunnel using a 5 deg half-angle sharp cone at a freestream Mach number of 7.1 and a wide range of stagnation conditions. Aerodynamic heating distributions and surface pressure fluctuations were measured with and without roughness elements applied to the cone. A wavy wall roughness with a wavelength of 2� located well upstream of the breakdown region had the effect of delaying transition. Surface pressure spectrum densities indicate that disturbances of the roughness wavelength are amplified compared to the smooth wall case. At a lower stagnation temperature, however, wavy wall roughness configurations had the same effect as regular roughness, and no discernible differences between wavy wall roughness and spherical roughness were observed, either in the transition point or in the pressure fluctuation spectrum.

Experiment of the Two-Dimensional Roughness Effect on Hypersonic Boundary-Layer Transition

An experimental investigation was conducted on a 5-degree half-angle cone and a 5-degree half-angle flared cone in a conventional Mach 6 wind tunnel to examine the effects of facility noise on boundary layer transition. The influence of tunnel noise was inferred by comparing transition onset locations determined from the present test to that previously obtained in a Mach 6 low disturbance quiet tunnel. Together, the two sets of experiments are believed to represent the first direct comparison of transition onset between a conventional and a low disturbance wind tunnel using a common test model and transition detection technique. In the present conventional hypersonic tunnel experiment, separate measurements of heat transfer and adiabatic wall temperatures were obtained on the conical models at small angles of attack over a range of Reynolds numbers, which resulted in laminar, transitional, and turbulent flow. Smooth model turbulent heating distributions are compared to that obtained with transition forced via discrete surface roughness. The model nosetip radius was varied to examine the effects of bluntness on transition onset. Despite wall to total temperature differences between the transient heating measurements and the adiabatic wall temperature measurement, the two methods for determining sharp cone transition onset generally yielded equivalent locations. In the 'noisy' mode of the hypersonic low disturbance tunnel, transition onset occurred earlier than that measured in the conventional hypersonic tunnel, suggesting higher levels of freestream acoustic radiation relative to the conventional tunnel. At comparable freestream conditions, the transition onset Reynolds number under low disturbance conditions was a factor of 1.3 greater than that measured on flared cone in the LaRC conventional hypersonic tunnel and a factor of 1.6 greater that the flared cone run in the low disturbance tunnel run 'noisy'. Navier-Stokes mean flow computations and linear stability analysis were conducted to assess the experimental results and have indicated N factors associated with sharp flared cone transition onset to be approximately a factor of 2 lower than that inferred from the corresponding low disturbance tunnel measurements.

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Boundary Layer Transition on Slender Cones in Conventional and Low Disturbance Mach 6 Wind Tunnels

The effect of isolated roughness on the windward surface boundary layer of the Shuttle Orbiter has been experimentally examined in the NASA Langley Research Center 20-InchMach 6 Tunnel. The size and location of isolated roughness elements (intended to simulate raised ormisalignedShuttleOrbiter Thermal Protection System tiles and protruding gap  ller material) were varied to systematically examine the response of the boundary layer. Global heat transfer images of the windward surface of a 0.75%-scaleOrbiter at an angle of attack of 40 deg were obtained over a range of Reynolds numbers using phosphor thermography and were used to infer the status of the boundary layer. Computationalpredictions were performed to provide both laminar and turbulent heating levels for comparison to the experimental data and to provide  ow eld parameters used for investigatingboundary-layer transition correlations. A variety of roughness heights and locations along the windward centerline were used. The roughness-transition correlation, using the predicted edge parameters Re /Me and k/ , was well behaved. The off-centerline results illustrate the potential for an asymmetric transition pattern to be isolated to one side of the vehicle, thereby causing the increased yawing moments experienced in  ight.

Shuttle Orbiter Experimental Boundary-Layer Transition Results with Isolated Roughness

Recent Orbiter wind tunnel data and flight data have been used to evaluate boundary-layer transition induced by discrete-roughness elements on the Orbiter windward surface. Orbiter flow field calculations have been used to compute transition parameters and disturbance parameters for correlating the results and comparing the trends. Existing transition correlations have been modified and applied to the Orbiter. These correlations provide a means to predict transition on the Orbiter given a known isolated roughness element. Furthermore, these data and data reduction methods provide information and guidance for the prediction of transition due to &screte-roughness elements on future winged reentry vehicles.

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Isolated Roughness Induced Boundary-Layer Transition: Shuttle Orbiter Ground Tests and Flight Experience

Conclusions The first series of metallic TPS tests in the NASA-LaRC Mach 7 High Temperature Tunnel has been completed. Additional testing is 'in progress and shall provide data for off-design configurations for the metallic TPS. The available data are being analyzed and being used to correlate analytical models to be used for X-33 flight design analysis. The final paper shan present additional data from these tests and comparisons between the data and analytical predictions.

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X-33 Metallic TPS Tests in NASA-LaRC High Temperature Tunnel

The spectral and total normal emittance of the Reinforced Carbon-Carbon (RCC) used on Space Shuttle Orbiter nose-cap and wing leading edges has been measured at room temperature and at surface temperatures of 1200 to 2100 K. These measurements were made on virgin and two flown RCC samples. Room temperature directional emittance data were also obtained and were used to determine the total hemispherical emittance of RCC as a function of temperature. Results of the total normal emittance for the virgin samples showed good agreement with the current RCC emittance design curve; however, the data from the flown samples showed an increase in the emittance at high temperature possibly due to exposure from flight environments.

Emittance measurements of Space Shuttle Orbiter Reinforced Carbon-Carbon

This article presents an introduction to ground testing of thermal protection materials (TPMs) used to protect spacecraft and hypersonic vehicles. These test facilities must simulate the aerothermodynamic environments expected during flight. Often these tests have been done in arc jet facilities. However, inductively coupled plasma (ICP) torches can generate the relevant test flow field and have advantages. Guidance for using these facilities for thermal protection material performance verification and flight qualification is provided.

Stanley A. Bouslog

Papers

Shuttle Orbiter Experimental Boundary-Layer Transition Results with Isolated Roughness

Isolated Roughness Induced Boundary-Layer Transition: Shuttle Orbiter Ground Tests and Flight Experience

X-33 Metallic TPS Tests in NASA-LaRC High Temperature Tunnel

Emittance measurements of Space Shuttle Orbiter Reinforced Carbon-Carbon

Plasma Torch Testing of Thermal Protection Materials (TPMs)