Aerothermodynamic Simulation of Subscale models of the FIRE II and Titan Explorer Vehicles in Expansion Tubes

Capra, Bianca Rose (2007). Aerothermodynamic Simulation of Subscale models of the FIRE II and Titan Explorer Vehicles in Expansion Tubes PhD Thesis, School of Engineering, University of Queensland.

       
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Author Capra, Bianca Rose
Thesis Title Aerothermodynamic Simulation of Subscale models of the FIRE II and Titan Explorer Vehicles in Expansion Tubes
School, Centre or Institute School of Engineering
Institution University of Queensland
Publication date 2007
Thesis type PhD Thesis
Supervisor Professor Richard Morgan
Abstract/Summary Scale models of the terrestrial re-entry probe FIRE 11, and the proposed Titan aerocapture vehicle Titan Explorer, were tested with regards to their aerothermodynamic environment in the expansion tubes at The University of Queensland. Test models were sized from their respective flight vehicles using binary scaling to ensure similitude was maintained between the flight and experimental flowfields. The FIRE I1 model had a length scale of 1:27.7, maximum diameter of 24.28mm and nose radius 33.78mm. This model was orientated at a 0" angle of attack and was designed to simulate the flight heating environment 21.75s after re-entry, where radiation contributed an estimated 17% - 36% to the total heat transfer. The Titan Explorer test model was a 70" sphere-cone and was designed to simulate a peak heating condition occurring 253s into an aerocapture manoeuvre at Titan, where radiative heat transfer was estimated to contribute more than 80% to the aerothermodynamic heating. The model was sized to a length scale of 1:40.8 giving a maximum diameter of 94.46mm with nose radius of 22.37mm. The Titan Explorer model was tested at both a 0" and 16" angle of attack with lift-up in a simulated Titan atmosphere consisting of 5% CH4 and 95% N2. Both models were manufactured from steel and contained forebody heat flux instrumentation for measurement of the aerothermodynamic heating. A new gauge was developed for the separation and detection of the radiative component of heat transfer to the Titan Explorer model where non-negligible amounts of radiation were present on the test model. This gauge contained two thin-film nickel temperature sensing elements housed in a brass holder. Sensing elements were separated from direct contact with the flow via appropriate optical windows. All reflecting surfaces of the radiation gauge, with the exception of the nickel element, were painted in flat black paint to minimise spurious reflections. Total heat transfer on both models was measured with fast response type-E surface mounted thermocouples. From an analysis of both a transparent and absorptive gas it was proposed that absolute aerothermodynamic radiative heat transfer will remain invariant between an experimental model and the flight vehicle, provided binary scaling was maintained, and radiation was uncoupled or weakly coupled to the flow. This necessitated the treating of each mode of heat transfer independently when scaling flight data to the experimental models and vice versa. A parameter referred to as the reduced flight value was developed that allowed experimentally measured heat transfer to be directly compared with the appropriately scaled flight heating rates. The reduced flight value also allowed for estimation of flight vehicle heating rates from experimentally measured total and radiative heat transfer. Stagnation point results from the FIRE I1 testing successfully demonstrated the ability of expansion tubes to recreate the aerothermodynamic environment of flight vehicles. At this location an average of 14.28&7%kW/cm2 of total heat transfer was experimentally measured. This was shown to have a good agreement of between 5% - 15% with both empirical convective heating correlations and the reduced flight values scaled from the flight data. Total heat transfer values measured at two radial locations 11.4" and 18.7" from the nose were 17.11&7%kW/m2 and l7.14f 7%kW/cm2 respectively and were substantially higher than both the stagnation point and reduced flight values. The high heating rates experimentally measured at these locations were attributed to the boundary layer becoming turbulent on the test model as a result of surface roughness and was estimated likely to occur at approximately 8.4' from the stagnation point. Aerothermodynamic radiative heat transfer was successfully separated and measured on the Titan Explorer test model. It was shown that this measured radiative heat transfer was a direct result of the addition of CH4 into the test gas which promoted the formation of the strong radiating particle cyanogen, CN, in the shock layer. At the design orientation of a 16" angle of attack with lift-up in a simulated Titan atmosphere, 20.8% 14%W/m2 of surface radiative heat transfer was measured at the flow stagnation point. This represented 4% of the total heat transfer of 510&7%W/cm2 recorded by the stagnation point thermocouples. A similar contribution of radiative heat transfer was measured with the model orientated at a 0" AOA when the radiation gauge was positioned 19.6mm vertically below the sphere-cone apex. In this orientation 21.6% 14%W/m2 of radiative heat transfer was measured compared to the local total heat transfer of 420W/cm2.

 
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Created: Fri, 21 Nov 2008, 15:53:03 EST