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Abstract
This thesis introduces a new experimental technique developed for prescribing flight representative, non-uniform temperatures on surfaces of hypersonic impulsive wind-tunnel models. The thesis is presented in three sections: (i) analyses of temperature profiles likely to occur on hypersonic vehicle surfaces, (ii) development of an experimental technique to generate desired, non-uniform temperatures, and (iii) experimental validation of the technique.
The range of temperature profiles exhibited on hypersonic vehicles was determined from thermal finite-element analyses of representative hypersonic vehicle structures using analytically derived hypersonic heat-flux models. Structural, material, and flight condition information from eight hypersonic aircraft programs were used to ensure the simulations were representative of actual aircraft. The most severe temperature profile calculated was used to define non-uniform temperature requirements for the experimental technique.
The developed experimental technique consisted of contoured electro-resistive heating elements embedded into the surface of hypersonic models. Equations were derived for calculating heating element thickness contours to achieve desired surface temperature profiles, with two heating modes designed: steady-state operation and transient operation. Preferable material property combinations for heating element materials were determined, with graphite, C/C, and C/SiC found to be suitable. Electrical power supply requirements were also derived.
For validation of the experimental technique, a heated aerodynamic model was designed and tested. The model consisted of an 80 mm wide flat plate, with an insertable graphite heating element positioned along the model's top surface. The model was tested at the University of Southern Queensland's TUSQ facility using their Mach 6 flow condition. For validating the contouring method, one steady-state operation heating element was tested but produced temperature errors up to 18%. Two transient operation heating elements were tested, with each achieving temperature errors up to 6.3% and 11.1% respectively. Aerothermodynamic tests were conducted with the model using three heating element configurations: (i) unheated, (ii) with a uniform temperature of 1023 K, and (iii) with non-uniform temperatures from 675 K to 1129 K. Schlieren results showed changes in boundary-layer thicknesses for the different wall temperature conditions were in approximate agreement with theory.