Procedures are discussed to correct for nonideality in a shock tube used in the reflected mode in conjunction with flash photolysis and atomic resonance absorption to measure chemical kinetics of atoms at high temperatures Experimentally, pressure time profiles for the incident and reflectedshock regions are made close to the location of the observation windows through which absorbance is measured The corresponding temperatures are calculated from the adiabatic equation of state Justification for this procedure is provided by extending Mirels' boundary layer theory to take intoaccount interaction of the reflected wave with the flowing gas in the free stream and in the boundary layer These theoretical methods are described for calculating the thermodynamic and hydrodynamic states behind the reflected wave from initial values of pressure and temperature and the measured velocity of the incident wave The implication of these results to kinetic measurements at high temperature is discussed

The thermodynamic state of the hot gas behind reflected shock waves: Implication to chemical kinetics†

Measurement of thermal rate constants by flash or laser photolysis in shock tubes: Oxidations of H2 and D2

The absorption spectrum of the negative ion of fluorine has been observed in shock-heated vapors of CsF. The photodetachment cross section of the fluorine ion has been measured in the wavelength region between 2100 and 3600 \AA{}. Results are found to be in reasonable agreement with the calculated values of Robinson and Geltman.

Electron Photodetachment Cross Section of the Negative Ion of Fluorine

Measurements and interpretation of shock tube ignition delay times in highly CO2 diluted mixtures using multiple diagnostics

Ignition, flame-holding, and mixing enhancement are fundamental aspects of supersonic combustion and are critical to the development of hypersonic airbreathing propulsion engines. High velocities associated with supersonic/hypersonic flight speeds constrain the performance of propulsion systems because of the limited flow residence time inside the combustor. A useful hypervelocity propulsion system therefore requires enhanced mixing of fuel and air, injection with very low drag penalty, and effective distribution of fuel over the burner cross-section. One of the simplest approaches is the transverse injection of fuel from wall orifices. The interesting but rather complicated flow-field dynamics of transverse jets injected into a supersonic crossflow has been studied by many supersonic combustion researchers since 1960’s, but with limited freestream flow conditions. Most of the previous research was performed in conventional wind tunnels by accelerating cold air into supersonic conditions, namely in low velocity and low total enthalpy flow conditions. However, a real supersonic combustor environment at flight speeds beyond Mach 8 can only be simulated using impulse facilities due to the required high total enthalpies. Among various impulse facilities, expansion tubes are especially useful in providing high total enthalpy flows with the proper chemical composition, namely the absence of dissociated species. This research is focused on studying the near-field mixing and ignition properties of transverse fuel jets injected into realistic supersonic combustor flows. We use advanced flow visualization techniques, namely planar laser-induced fluorescence (PLIF) imaging of the hydroxyl radical (OH) and ultra-fast-framing-rate schlieren imaging. While schlieren indicates the location of shock waves, jet penetration and large scale flow features, OH-PLIF is used to map the regions of ignition. The first objective of the present work is to characterize the expansion tube facility

Experimental investigation of mixing and ignition of transverse jets in supersonic crossflows

Shock tube flow nonuniformity is investigated in the limit where the shock and contact surface have reached their maximum separation. Ideal gases are considered. It is found that all fluid properties increase in value between the shock and contact surface. The nonuniformity is greatest when γ (ratio of specific heats) is large and Ms (shock Mach number) is low. For γ = 53 and Ms ≥ 3, the static temperature, density, and pressure increase by about 8, 12, and 20%, respectively; the stagnation temperature increases by about 35%, and the stagnation pressure, dynamic pressure, and stagnation point heat transfer increase by about a factor of 2. These results apply to turbulent as well as laminar boundary layers. The variation of flow properties with distance behind the shock, as well as particle time of flight, is given for both wholly laminar and wholly turbulent wall boundary layers. These results are particularly important for chemical rate and heat transfer studies.

H. Mirels

Papers

Flow Nonuniformity in Shock Tubes Operating at Maximum Test Times