Experimental Characterization for Hypersonic Testing (original) (raw)
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An Overview on Hypersonic Flow Research with Infrared Thermography
QIRT 06, 2006
The technological development achieved in instruments and methodology concerning both flights and ground hypersonic experiment (employed in space plane planning) goes towards an updating and a standardization of the heat flux technical measurements. In fact, the possibility to simulate high enthalpy flow relative to reentry condition by hypersonic arc-jet facility needs devoted methods to measure heat fluxes. Aim of this work is to demonstrate that InfraRed (IR) thermographic measurements with new heat flux sensor (IR-HFS) can be used as powerful tool in hypersonic high enthalpy flow research.
CARS temperature measurements in a hypersonic propulsion test facility
1990
Static-temperature measurements performed in a reacting vitiated air-hydrogen Mach-2 flow in a duct in Test Cell 2 at NASA LaRC by using a coherent anti-Stokes Raman spectroscopy (CARS) system are discussed. The hypersonic propulsion Test Cell 2 hardware is outlined with emphasis on optical access ports and safety features in the design of the Test Cell. Such design considerations as
Hypersonic Test Analysis by Means of Aerothermal Coupling Methodology and Infrared Thermography
AIAA Journal, 2013
In the current analysis, surface temperature measurements, carried out with infrared thermography in a plasma wind tunnel experiment on a wing-leading-edge thermal-protection-system demonstrator, are performed and compared with the results of the test numerical rebuilding carried out through an integrated procedure coupling the external aerodynamic field to the internal thermal state of the structure. For the measurement campaign, an infrared mirror system is developed to allow the viewing of the leading-edge region, which was not visible from the standard optical access of the facility. The optical calibration of the camera with a special target is performed. The twodimensional infrared images are rebuilt on the three-dimensional surface grid of the test article used for computational fluid dynamics computations, thus providing a useful tool for the direct comparison of numerical and experimental data. The comparison between numerical and experimental data is quite satisfactory even though some discrepancies, critically discussed, are present.
47th AIAA Aerospace Sciences Meeting including The New Horizons Forum and Aerospace Exposition, 2009
This paper describes the analytical methods used for temperature sensitive paint (TSP) measurements to obtain quantitative global heat-flux diagnostics in the hypersonic environment on important physical phenomena like laminar-turbulent transition, near-surface stationary vortices and separation. This problem is generally considered as an inverse heat transfer problem. The exact analytical solution of the 1D unsteady heat conduction for a polymer layer (TSP or TSP/insulating-layer) on a base of any material is obtained and used to calculate heat flux into the polymer surface as an explicit function of a time-dependent surface temperature change measured by TSP in hypersonic tunnels. In addition, an analytical solution of the 3D unsteady heat conduction equation is given to include the effect of the lateral heat conduction.
International Journal of Aeronautical and Space Sciences, 2010
Through experimental investigations utilizing hypersonic shock tunnel-coaxial thermocouples as well as blow down hypersonic wind tunnel-temperature sensitive paints, the heat flux and the temperature over a protuberance were measured and analyzed. The experimental data were subsequently compared to heat flux data that was obtained by using blow down hypersonic wind tunnel and heat flux gauges. According to the comparison, both sets of data illustrated correlation with one another. The measured heat flux was large when the height of the protuberance was large. Experimental results show that heat flux measurements taken at higher locations were greater than those taken at lower locations. For high protuberances, a severe jump in the heat flux was observed, ranging in values within 0.6-0.7 of the height of the protuberances. However, when the protuberance was sufficiently short, a rise in the heat flux was rarely observed as the protuberance was totally submerged under the separation region.
Journal of Thermophysics and Heat Transfer, 2010
This paper describes the analytical method used for temperature-sensitive-paint measurements to obtain quantitative global heat flux diagnostics in hypersonic tunnels. An exact inverse solution of the one-dimensional unsteady heat conduction equation for a polymer layer (temperature-sensitive paint or temperature-sensitive paint and an insulator layer) on a base of any material is obtained to calculate heat flux into the polymer surface as an explicit function of a time-dependent surface temperature change measured by temperature-sensitive paint. In addition, an analytical solution of the three-dimensional unsteady heat conduction equation is given to consider the effect of the lateral heat conduction. Simulations and experiments are conducted to examine the analytical method and assess the relevant factors regarding the measurement uncertainty.
Sadhana, 2006
Aerodynamic forces and fore-body convective surface heat transfer rates over a 60 • apex-angle blunt cone have been simultaneously measured at a nominal Mach number of 5·75 in the hypersonic shock tunnel HST2. An aluminum model incorporating a three-component accelerometer-based balance system for measuring the aerodynamic forces and an array of platinum thin-film gauges deposited on thermally insulating backing material flush mounted on the model surface is used for convective surface heat transfer measurement in the investigations. The measured value of the drag coefficient varies by about ±6% from the theoretically estimated value based on the modified Newtonian theory, while the axi-symmetric Navier-Stokes computations overpredict the drag coefficient by about 9%. The normalized values of measured heat transfer rates at 0 • angle of attack are about 11% higher than the theoretically estimated values. The aerodynamic and the heat transfer data presented here are very valuable for the validation of CFD codes used for the numerical computation of flow fields around hypersonic vehicles.
Unsteady simulation of hypersonic flow around a heat flux probe in ground testing conditions
International Journal of Heat and Mass Transfer, 2017
In order to investigate the accuracy of the rebuilding code for the free stream conditions and the total enthalpy in the Longshot Hypersonic facility at the von Karman Institute (VKI), a series of unsteady CFD simulations of axisymmetric hypersonic flow over a heat flux probe have been performed. The free stream and the total conditions provided by experiments as a function of time were used as unsteady inlet boundary conditions for those simulations. The full duration (15 ms) of the test window is simulated by means of an implicit time accurate finite volume solver. The numerical method and the time integration procedure used during this investigation are described. The numerical results are within 6% of the experimental data for both the stagnation pressure and the heat flux at the reference test time conditions.
Stagnation temperature measurements in the USQ hypersonic wind tunnel
2010
A thermocouple probe with a heated shield has been used to measure stagnation temperature at the nozzle exit of the University of Southern Queensland hypersonic wind tunnel. The thermocouple probe consisted of a welded junction T-type thermocouple mounted within a heated tube with a vent hole downstream of the junction. Pressure transducers within the barrel of the wind tunnel have also been used to obtain the pressure history during the free piston compression process. The pressure measurements have been used to provide a theoretical value for the flow stagnation temperature for direct comparison with the thermocouple measurements. Assuming isentropic compression of the test gas, the flow stagnation temperature would be about 571 K for the current operating condition. After applying a response-time correction for the thermocouple signals, a stagnation temperature value of about 495 K was obtained from the measurements. The measured stagnation temperature of the test gas is somewhat lower than the isentropic value because of heat loss from the test gas to the barrel during the test gas compression and discharge process.