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Automated thermal mapping techniques using chromatic image analysis

01 Apr 1989-Vol. 89, pp 25443
TL;DR: In this paper, a chromatic image analysis system and temperature sensitive coatings are used for thermal mapping and surface heat transfer measurements on aerothermodynamic test models in hypersonic wind tunnels.
Abstract: Thermal imaging techniques are introduced using a chromatic image analysis system and temperature sensitive coatings. These techniques are used for thermal mapping and surface heat transfer measurements on aerothermodynamic test models in hypersonic wind tunnels. Measurements are made on complex vehicle configurations in a timely manner and at minimal expense. The image analysis system uses separate wavelength filtered images to analyze surface spectral intensity data. The system was initially developed for quantitative surface temperature mapping using two-color thermographic phosphors but was found useful in interpreting phase change paint and liquid crystal data as well.

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Citations
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Journal ArticleDOI
TL;DR: In this article, boundary-layer trip devices for the Hyper-X forebody have been experimentally examined in several wind tunnels, including the NASALangleyResearch 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.
Abstract: 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.

186 citations

Journal ArticleDOI
TL;DR: In this article, the effects of discrete and distributed roughness elements on boundary layer transition, which included trip height, size, location, and distribution, both on and off the windward centerline, were investigated.
Abstract: Boundary layer and aeroheating characteristics of several X-33 configurations have been experimentally examined in the Langley 20-Inch Mach 6 Air Tunnel. Global surface heat transfer distributions, surface streamline patterns, and shock shapes were measured on 0.013-scale models at Mach 6 in air. Parametric variations include angles-of-attack of 20-deg, 30-deg, and 40-deg; Reynolds numbers based on model length of 0.9 to 6.6 million; and body-flap deflections of 0, 10 and 20-deg. The effects of discrete and distributed roughness elements on boundary layer transition, which included trip height, size, location, and distribution, both on and off the windward centerline, were investigated. The discrete roughness results on centerline were used to provide a transition correlation for the X-33 flight vehicle that was applicable across the range of reentry angles of attack. The attachment line discrete roughness results were shown to be consistent with the centerline results, as no increased sensitivity to roughness along the attachment line was identified. The effect of bowed panels was qualitatively shown to be less effective than the discrete trips; however, the distributed nature of the bowed panels affected a larger percent of the aft-body windward surface than a single discrete trip.

110 citations

Journal ArticleDOI
TL;DR: 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 as discussed by the authors.
Abstract: 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 Ž 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  owŽ 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.

76 citations

Journal ArticleDOI
TL;DR: In this paper, the authors provide an overview of the hypersonic aerothermodynamic wind tunnel program conducted at the NASA Langley Research Center in support of the X-38 development, which is intended to demonstrate the entire mission profile of returning Space Station crew members safely back to earth in the event of medical or mechanical emergency.
Abstract: The X-38 program seeks to demonstrate an autonomously returned orbital test flight vehicle to support the development of an operational Crew Return Vehicle for the International Space Station. The test flight, anticipated in 2002, is intended to demonstrate the entire mission profile of returning Space Station crew members safely back to earth in the event of medical or mechanical emergency. Integral to the formulation of the X-38 flight data book and the design of the thermal protection system, the aerothermodynamic environment is being defined through a synergistic combination of ground based testing and computational fluid dynamics. This report provides an overview of the hypersonic aerothermodynamic wind tunnel program conducted at the NASA Langley Research Center in support of the X-38 development. Global and discrete surface heat transfer force and moment, surface streamline patterns, and shock shapes were measured on scaled models of the proposed X-38 configuration in different test gases at Mach 6, 10 and 20. The test parametrics include angle of attack from 0 to 50 degs, unit Reynolds numbers from 0.3 x 10 (exp 6) to 16 x 10 (exp 6)/ ft, rudder deflections of 0, 2, and 5 deg. and body flap deflections from 0 to 30 deg. Results from hypersonic aerodynamic screening studies that were conducted as the configuration evolved to the present shape at, presented. Heavy gas simulation tests have indicated that the primary real gas effects on X-38 aerodynamics at trim conditions are expected to favorably influence flap effectiveness. Comparisons of the experimental heating and force and moment data to prediction and the current aerodynamic data book are highlighted. The effects of discrete roughness elements on boundary layer transition were investigated at Mach 6 and the development of a transition correlation for the X-38 vehicle is described. Extrapolation of ground based heating measurements to flight radiation equilibrium wall temperatures at Mach 6 and 10 were made and generally compared to within 50 deg F of flight prediction.

53 citations

Proceedings ArticleDOI
01 Jan 1998
TL;DR: A status review of the experimental and computational work performed to support the X-33 program in the area of hypersonic boundary-layer transition is presented in this article, where global transition fronts are visualized using thermographic phospor measurements.
Abstract: A status review of the experimental and computational work performed to support the X-33 program in the area of hypersonic boundary-layer transition is presented. Global transition fronts are visualized using thermographic phospor measurements. Results are used to derive transition correlations for "smooth body" and discrete roughness models and a computational tool is developed to predict transition onset for X-33 using these correlations. The X-33 thermal protection system appears to be conservatively designed based on these models. Additional study is needed to address concerns related to surface waviness before final conclusions can be made. A discussion of future test plans is included.

51 citations