Deduction of temperature fluctuations in transient compression wind tunnels using incompressible turbulent flow data

Wind tunnels and other aero-thermal experimental facilities are likely to make a contribution to the optimisation of energy and propulsion systems for the foreseeable future. Short duration wind tunnels such as shock tunnels and gun tunnels rely on a transient compression rocess and are likely to generate significant turbulent fluctuations in the nozzle reservoir region. In the present study, the magnitude of likely stagnation temperature fluctuations in two such facilities is inferred from incompressible temperature fluctuations data obtained by other workers. The friction velocity Reynolds numbers for the gun tunnel and shock tunnel cases considered presently were ReT= 31,579 and 24,975 respectively. The RMS stagnation temperature fluctuations, when averaged over the pipe flow diameter, are estimated to be 15.3 and 291 K for the gun tunnel and shock tunnel cases respectively. The estimated RMS value in the case of the gun tunnel is significantly larger than the experimental value previously measured on the centre line of the gun tunnel nozzle of 2.3 K. The difference observed between the inferred and measured temperature fluctuations in the gun tunnel case may be related to spatial variations in the temperature fluctuations. In the case of the shock tunnel, the magnitude of the fluctuations is demonstrated to be significant for supersonic combustion experiments. The present approach for estimating the magnitude temperature fluctuations should be refined, but more detailed measurements of temperature fluctuations in such facilities are also required

Visible and near infrared spectroscopy of Hayabusa re-entry using semi-autonomous tracking

A ground-based tracking camera and co-aligned slit-less spectrograph were used to measure the spectral signature of visible radiation emitted from the Hayabusa capsule as it entered into the Earth's atmosphere in June 2010. Good quality spectra were obtained that showed the presence of radiation from the heat shield of the vehicle and the shock-heated air in front of the vehicle. An analysis of the black body nature of the radiation concluded that the peak average temperature of the surface was about (3100±100) K

Absolute concentration measurements of OH* in an axisymmetric hydrogen-air premixed flame adjacent to a hot graphite model

Absolute concentration of the chemiluminescent radical OH* was determined in an axisymmetric hydrogen-air premixed flame adjacent to a resistively-heated graphite surface. Two-dimensional images of the axisymmetric chemiluminescence from the excited-state of OH were recorded by an ICCD camera with a narrow-band filter at approximately 310 nm. A temperature of around 1800 K was achieved on the graphite surface using an electrical heating power of 5.5 kW. Surface temperatures were measured using a two-color ratio pyrometry (TCRP) technique. The line-of-sight-integrated chemiluminescent emissions that were imaged using the ICCD device were transformed to radial distributions through an Abel inversion method. A new method for calibration of the absolute number density of the radiating radical OH* is proposed based on the intensity ratio of the measured OH* chemiluminescence and the radiation emitted from the hot graphite surface. This is a convenient approach in the present work because adequate signal magnitudes from both these phenomena are acquired by the ICCD device simultaneously during testing

Radiometric temperature analysis of the Hayabusa spacecraft re-entry

Hayabusa, an unmanned Japanese spacecraft, was launched to study and collect samples from the surface of the asteroid 25143 Itokawa. In June 2010, the Hayabusa spacecraft completed it’s seven year voyage. The spacecraft and the sample return capsule (SRC) re-entered the Earth’s atmosphere over the central Australian desert at speeds on the order of 12 km/s. This provided a rare opportunity to experimentally investigate the radiative heat transfer from the shock-compressed gases in front of the sample return capsule at true-flight conditions. This paper reports on the results of observations from a tracking camera situated on the ground about 100 km from where the capsule experienced peak heating during re-entry

Shock tunnel studies of scramjet phenomena, supplement 8

Reports by the staff of the University of Oueensland on various research studies related to the advancement of scramjet technology are presented. These reports document the tests conducted in the reflected shock tunnel T4 and supporting research facilities that have been used to study the injection, mixing, and combustion of hydrogen fuel in generic scramjets at flow conditions typical of hypersonic flight. In addition, topics include the development of instrumentation and measurement technology, such as combustor wall shear and stream composition in pulse facilities, and numerical studies and analyses of the scramjet combustor process and the test facility operation. This research activity is Supplement 8 under NASA Grant NAGW-674

Shock tunnel studies of scramjet phenomena, supplement 7

Reports by the staff of the University of Queensland on various research studies related to the advancement of scramjet technology are presented. These reports document the tests conducted in the reflected shock tunnel T4 and supporting research facilities that have been used to study the injection, mixing, and combustion of hydrogen fuel in generic scramjets at flow conditions typical of hypersonic flight. In addition, topics include the development of instrumentation and measurement technology, such as combustor wall shear and stream composition in pulse facilities, and numerical studies and analyses of the scramjet combustor process and the test facility operation. This research activity is Supplement 7 under NASA Grant NAGW-674

Nitrous oxide decomposition for supersonic combustion experiments in the USQ Ludwieg tube facility

Wind tunnel facilities are required to support the development of the scramjet engine. The performance of a short duration wind tunnel which uses the thermal decomposition of nitrous oxide to augment the flow stagnation enthalpy is simulated in the present work. In its normal model of operation, the wind tunnel being considered uses a free piston compression process to compress the test gas to a moderate stagnation temperature and pressure, suitable for cold flow hypersonic aerodynamics experiments. However, by introducing a mixture of nitrous oxide and nitrogen ahead of the piston, it should be possible to provide a test gas which closely simulates the properties of air. A simulation which models the thermo-chemistry of the free piston compression process with nitrous oxide is described. Results indicate that relatively large magnitude fluctuations in the test gas are likely to arise due to the rapid decomposition process, rendering the compressed gas unsuitable for supersonic combustion testing. Piston braking or some other change to the configuration would be necessary in order to successfully use nitrous oxide decomposition to enhance the enthalpy of the test gas

Transient temperature probe measurements in a Mach 4 nitrogen jet

Stagnation temperature measurements have been obtained in a Mach 4 free jet of nitrogen using a technique based on transient thin film heat flux probe measurements. The uncertainty in the stagnation temperature measurements depends on the probe location within the jet but is typically around ±5 K at the centre of the jet. The thin film heat flux probe technique also provides a measurement of the heat transfer coefficient of the thin film probes with an uncertainty of around ±4% at the centre of the jet. Pitot pressure measurements were also obtained within the jet. Analysis of the heat transfer coefficient results yields the Mach number and velocity profiles which are compared with results from the pitot probe measurements. Jet velocities identified using the thin film probe and the pitot probe techniques produce results with uncertainties of less than ±2% at the centre of the jet. Measurements of RMS stagnation temperature fluctuations indicate values of around 3 K at the centre of the jet to more than 10 K in the shear layer