Abstract
This paper demonstrates time-resolved stagnation temperature measurements in a shock tunnel at a frequency of 25 kHz using emission spectroscopy in air and nitrogen test conditions. The two most important parameters for determining the flow conditions generated in a shock tunnel experiment are the stagnation pressure and temperature of the flow just upstream of the supersonic nozzle. While the pressure can be measured using a wall-mounted transducer with relative ease, the measurement of the temperature requires a optical technique such as time resolved emission spectroscopy. Knowledge of the transient stagnation temperature behavior is critical to all subsequent expansion tube flow processes. The driver gas emission spectrum data at the post-shock condition shows continuum and atomic line radiation. The continuum radiation can be described by a black body radiator with the individual spectra showing sufficient continuum information for accurately fitting Planck functions. Atomic line radiation was excluded by skipping those data from the measured spectra. The fitting routine shows clear differences in determined temperatures including and neglecting atomic line radiation. These measurements allow for the exact determination of the shock tunnel flow conditions in combination with pressure transducer data. The flow condition used in the experiment corresponds to a nominal Mach-10 condition at an altitude of 65 km, however, the technique is not limited to this condition and can be used for a large range of flow conditions.
Highlights
The difficulties associated with directly measuring the stagnation temperature using these techniques have led scientists to look for easier options to determine the flow’s total temperature
The stagnation temperature is typically inferred from a measurement of the shock speed in the shock tube
The shock speed, inferred from the temporal pressure traces using these two transducers, can be used with the normal shock equations to calculate the total temperature of the flow if the initial pressure and gas composition is known
Summary
The difficulties associated with directly measuring the stagnation temperature using these techniques have led scientists to look for easier options to determine the flow’s total temperature. The shock speed, inferred from the temporal pressure traces using these two transducers, can be used with the normal shock equations to calculate the total temperature of the flow if the initial pressure and gas composition is known. Numerical codes such as NASA CEA or STUBE can be used to account for the appropriate thermochemical gas properties. The use of shock tunnels for aerodynamic and aerothermodynamic investigations of high-speed vehicles is interesting as the real flight conditions can be correctly reproduced [1]. Studies of side-jet missile control, of free-flight trajectory measurements for aerodynamic coefficient estimation, of projectile control by plasma discharge, of heatflux measurements at missile noses to predict heat loads on missile surfaces and structures and of the atmospheric dispersion of droplets are summarized
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