Characterization of high enthalpy flows in the IISc free-piston driven shock tunnel using Two-Colour Ratio Pyrometry
High temperature and its associated effects set the hypersonic flow regime apart from other class of flows. Viscous dissipation raises the internal energy of the high-kinetic-energy gas as it slows down in the boundary layer. For slow heat conduction into the vehicle surface, the gas temperature increases drastically, leading to real gas effects, dissociation and ionization of molecules. Due to the thickened boundary layer, strong ‘viscous interaction’ exists between the outer inviscid shock layer and the boundary layer, leading to a pressure rise compared to an inviscid case. This pressure in turn tends to make the boundary layer thinner than expected and increases the temperature gradient (for a cold wall case) and therefore the aerodynamic heating of the vehicle. High temperature near the vehicle airframe may cause surface ablation of the heat shield, releasing energy into the shock layer. Short duration flights, with a small time of flight may not experience ablation because the heat load on such vehicles may be too little, despite the high temperature in their vicinity. Energy release affects its aerothermodynamics and its performance. Effect of energy deposition (by means like plasma heat addition, electric arc discharge, chromium oxidation, etc.) on wave drag has been well documented in open literature. However, no concrete measurements exist on the aerothermodynamic effects of heat addition. Apart from ablation, flow separation may occur due to shock-boundary layer interaction (SBLI) towards the aft of a blunt nosed vehicle. Temperature in the dead air region (separated region with circulatory flow) inside the shear layer influences the performance of the vehicle and must be characterized. Apart from SBLI, another cause of intense heating is shock-shock interaction as classified by Edney. Although studies on wave drag galore, no experiments exist to characterize the temperature distribution in the separated shear layer. Characterization of these phenomena in an in-house free piston driven shock tunnel (FPST; called HST3) is the subject of this work. These studies in shock tunnels require development of a temperature measurement technique and related equipment. Most equipment tested in shock tunnels are marred by disadvantages such as single point measurement, lack of knowledge of emissivity, etc. No attempt has been made towards spatially resolved temperature characterization of the high enthalpy gas layer in shock tunnels. A non-intrusive, optical temperature measurement technique called two-colour ratio pyrometry (TCRP) was developed in the laboratory. A commercial DSLR camera was used as the pyrometer. TCRP was demonstrated in a tube furnace, by comparing the results with those obtained from other methods. The camera was then used for time integrated (and line-of-sight (LOS) integrated) temperature characterization of the shock layer at a stagnation enthalpy of 5.2 MJ/kg and Mach 10.1. Peaks over a continuous broad band, obtained from emission spectroscopy, were accounted for during temperature characterization. The stagnation region temperature was about 14% lower than that predicted by a numerical code. Time resolved temperature measurements, during the test time of 260 μs, were done by acquiring images of the flow using a high-speed camera at 20,000 fps. A numerical model was also developed in Fluent to obtain the actual stagnation plane temperature from the path length integrated value. The time resolved temperature was higher by as much as 14% from the time integrated one, whereas the actual stagnation plane temperature was only about 2.5% higher than the LOS integrated value. Next, the effect of exothermic surface ablation and heat addition, by oxidation of a thin film of chromium coated on a large angle blunt cone in air flow, was studied on the aerothermodynamics of the shock layer-by characterizing its effects on the shock layer temperature, surface heat flux and shock stand-off distance. Experiments were done at a stagnation enthalpy of 6.31 MJ/kg and Mach number of 9.73. Heat release increased the shock layer temperature (obtained from TCRP) by 173 K, convective heat flux at the model surface by 25.5% and shock stand-off distance by 17%. The effect on shock layer aerothermodynamics was more with O2 as the test gas, as it contains a larger mole fraction of oxygen than air. Numerical simulations and calculations done to simulate real gas effects and wall surface reactions predicted a Cr regression rate of 0.074 mm/sec and a total heat deposition of about 215 W/cm2. Cr being carcinogenic and possessing a high activation energy, an organic compound, Nafion was tested as its replacement and found to have more prominent effects on the shock layer aerothermodynamics. A spiked body configuration was used to study the separated flow encountered at the aft of reentry vehicles. A time integrated image of the flow over the configuration was first acquired using a DSLR camera at 5.2 MJ/kg to locate the temperature hotspots. A maximum temperature of around 2920 K was encountered, in the region enveloped by the shear layer where the flow separates and recirculates. The temperature within the dead air region is high due to conversion of kinetic energy to internal energy. The flow was found to be unsteady in the separated zone and consists of collapse and inflation. Time resolved images of the unsteady flow were taken at an enthalpy of 6.31 MJ/kg. The unsteady results also showed low temperatures inside the conical shock and extreme temperatures, as high as 3300 K in the separated region. To conclude, two important phenomena pertaining to the design of hypersonic reentry vehicles were studied, and a temperature measurement technique to study them was developed and demonstrated in the laboratory.