X-33 experimental aeroheating at Mach 6 using phosphor thermography

Abstract

The goal of the NASA Reusable Launch Vehicle (RLV) technology program is to mature and demonstrate essential, cost effective technologies for next generation launch systems. The X-33 flight vehicle presently being developed by Lockheed Martin is an experimental Single Stage to Orbit (SSTO) demonstrator that seeks to validate critical technologies and insure applicability to a full scale RLV. As with the design of any hypersonic vehicle, the aeroheating environment is an important issue and one of the key technologies being demonstrated on X-33 is an advanced metallic Thermal Protection System (TPS). As part of the development of this TPS system, the X-33 aeroheating environment is being defined through conceptual analysis, ground based testing, and computational fluid dynamics. This report provides an overview of the hypersonic aeroheating wind tunnel program conducted at the NASA Langley Research Center in support of the ground based testing activities. Global surface heat transfer images, surface streamline patterns, and shock shapes were measured on 0.013 scale (10-in.) ceramic models of the proposed X-33 configuration in Mach 6 air. The test parametrics include angles of attack from -5 to 40 degs, unit Reynolds numbers from 1x10 6 to 8x10 6 /ft, and body flap deflections of 0, 10, and 20 deg. Experimental and computational results indicate the presence of shock/shock interactions that produced localized heating on the deflected flaps and boundary layer transition on the canted fins. Comparisons of the experimental data to laminar and turbulent predictions were performed. Laminar windward heating data from the wind tunnel was extrapolated to flight surface temperatures and generally compared to within 50 deg F of flight prediction along the centerline. When coupled with the phosphor technique, this rapid extrapolation method would serve as an invaluable TPS design tool. Nomenclature h heat transfer coeff. (lbm/ft 2 -sec), Ç q /(H aw -H w )

Aerothermodynamic Characteristics of Boundary Layer Transition and Trip Effectiveness of the HIFiRE Flight 5 Vehicle

“…This pattern is indicative of the switch from in-flow towards out-flow as the angle of attack is increased for the model and is similar to the patterns exhibited on vehicles such as the X-33 in Ref. 26.…”

“…On the basis of these resemblances, this transition pattern is presumed to be the result of crossflow. 26 As shown in Fig. 5, the centerline heating departed from the laminar heating conditions at all unit Reynolds number tested, though at the lowest unit Reynolds number, 1.1x10 6 /ft, heating rates did not reach fully turbulent heating conditions before the end of the model.…”

X-33 hypersonic boundary layer transition

“…9c) 7 figuration and appears to be related to changes in the direction of the forebody streamlines, as reported in Ref. 7. At the lower angles of attack the forebody generates surface streamlines which curve predominately in towards the centerline (inflow).…”

Critical Hypersonic Aerothermodynamic Phenomena

“…Two articles, by Horvath et al (2001) and Murphy et al (2001), presented information relating to objective 1: "Obtain data to define the aerodynamic forces and moments and/or the heat-transfer distributions for complete configurations whose complex flow fields resist computational modeling." Two arti cles by Hollis et al (2001a,b) presented information relating to objective 4: "Obtain measurements of parameters, such as the heat transfer and the drag, to be used in comparison with computed flow-field solutions over a range of configuration ge ometries and of flow conditions (code calibration)."…”