Optical and Pressure Measurements in Shock Tunnel Testing of a Model Scramjet Combustor

McIntyre, Timothy J.; Houwing, A. F. P.; Palma, P. C.; Rabbath, P. A. B.; Fox, J.S.

doi:10.2514/2.5196

Cited by 26 publications

(12 citation statements)

References 12 publications

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“…In previous experimental efforts, planar laser-induced fluorescence using the hydroxyl molecule as the fluorescing species (OHPLIF) has been applied to combusting flows in shock tunnels 12,13 and to a model scramjet combustor 14 . However, in those experiments the OHPLIF technique was used to visualize regions of combustion in the flow, or flow features and flow mixing, and not for velocity measurements.…”

Section: Choice Of Measurement Techniquementioning

confidence: 99%

Stereoscopic Particle Image Velocimetry Measurements in Scramjet Combustors

Smith¹

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The prospect of atmospheric flight at five times the speed of sound, or greater, has captivated researchers for many years. One type of hypersonic propulsion engine capable of providing such astounding flight-speeds is the Dual-mode scramjet (DMSJ). This engine is especially attractive because a single flowpath is used over a wide range of flight Mach numbers. However, the complex combustion process in a DMSJ is not well understood. The flowfield of a DMSJ combustor is highly turbulent and highly threedimensional due to fuel injection schemes and massive flow separation. These characteristics, along with temperatures in excess of 1200K, make experimental measurements extremely challenging.Therefore, three-component velocity

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Section: Choice Of Measurement Techniquementioning

confidence: 99%

Stereoscopic Particle Image Velocimetry Measurements in Scramjet Combustors

Smith¹

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“…[6][7][8][9][10][11][12][13][14][15][16][17][18][19][20][21][22]. Because of the large stagnation enthalpies (greater than 3 MJ/kg) associated with high ight Mach numbers beyond Mach 8, currently only impulse facilities are capable of providing the required stagnation temperature and Mach number to replicate a combustor environment.…”

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confidence: 99%

Characterization of Expansion Tube Flows for Hypervelocity Combustion Studies

Ben‐Yakar

Hanson

2002

Journal of Propulsion and Power

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An expansion tube with an 89 mm inner diameter and a 12 m length was characterized for basic studies of the near-eld mixing and ame-holding characteristics of gaseous fuels injected in high-stagnation-enthalpy supersonic ows. The expansion tube provides a wide range of freestream conditions with little perturbation of test gas chemical composition. The latter is critical for supersonic-combustion studies in the high-stagnationenthalpy ows associated with hypersonic airbreathing propulsion systems. Efforts focused on achieving three operating points, simulating ight Mach 8, 10, and 12 stagnation-enthalpyconditions at the entrance of a supersonic combustor. The ability of the expansion tube to provide a steady-ow test time of adequate duration and a core ow of suf cient size for 2-mm jet-in-cross ow studies was veri ed. A core-ow size of 25 mm could be achieved in 6.2 MJ/kg stagnation-enthalpy ows with 400 ¹s of steady test time ( ight Mach 12 simulation). The primary effects of the boundary layer on the expansion tube ow were to extend the useful test time (in contrast with conventional shock-tube performance) and to decrease the core-ow size. NomenclatureDa = Damköhler number J = jet to freestream momentum ux ratio, (°pM 2 / j =.°pM 2 / 1 L = characteristic length L 1 = characteristic distance from edge of the injection plate L 2 = characteristic recirculation region length M = Mach number M s = shock Mach number M w = molecular weight p = static pressure q = dynamic pressure Re = Reynolds number T = static temperature t = time U = velocity x = distance°= speci c heat ratio°c = average speci c heat ratio, 1 2 .°0 C°3/ c = compression ef ciency through a hypersonic vehicle diffuser µ = characteristic temperature for the ignition delay time º i = gas composition ½ = density ¿ c = chemical time ¿ i = ignition delay time Subscripts j , jet = jet properties at the injection exit 0 = atmosphere conditions 1 = driven section conditions

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“…Although laser-based techniques like coherent anti-Stokes Raman scattering (CARS) (Eckbreth et al 1988, Vereschagin et al 2001, Cutler et al 2003, O'Byrne et al 2007) and planar laser-induced fluorescence (PLIF) (Allen et al 1993, McMillin 1993, McIntyre et al 1997, Gruber et al 2004) have proved to be very useful in laboratory scramjet tests because of their nonintrusive, species-specific measurement capabilities, these techniques require large, expensive pulsed laser systems that cannot realistically be put onto a flight experiment.…”

Section: Introductionmentioning

confidence: 99%

Diode Laser Sensor for Scramjet Inlets

O’Byrne¹,

Wittig²,

Kurtz³

et al. 2011

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Report Documentation PageForm Public reporting burden for the collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson Davis Highway, Suite 1204, Arlington VA 22202-4302. Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to a penalty for failing to comply with a collection of information if it does not display a currently valid OMB control number. REPORT DATE JUN 20112. REPORT TYPE This work presents test results of a diode-laser sensor for an axially symmetric inlet in a Mach 6 hypersonic flow facility. The conical inlet design presented here is more realistic than the flat plate design presented in the previous report, and is similar to the design that will be chosen as part of the Australian Space Research Program SCRAMSPACE flight test scheduled for 2012. The optical system was changed for this study to make the pitch and catch optics less sensitive to vibration, and to fit into a smaller volume than for the previous test. In addition, wavelength modulation detection was attempted with the system. Initial tests of the conical inlet design showed that it was not starting, as indicated using flow visualization images. The second version of the sensor has similar velocity sensitivity to the system used for the flat plate inlet test. The conical system was, however, still too susceptible to amplitude noise, which hinders the ability of the system to measure the free stream temperature or the angle of attack. Electrical problems with the log-ratio amplifier prevented temperature measurements for the vehicle. Preliminary vibration tests have shown that the problem is caused by independent motion between the laser and collimator when the tunnel fires. A preliminary design of a low-power lock-in amplifier for use in field experiments was completed, and tests with current inputs in place of photodiodes proved successful. SUBJECT TERMS

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Optical and Pressure Measurements in Shock Tunnel Testing of a Model Scramjet Combustor

Cited by 26 publications

References 12 publications

Stereoscopic Particle Image Velocimetry Measurements in Scramjet Combustors

Stereoscopic Particle Image Velocimetry Measurements in Scramjet Combustors

Characterization of Expansion Tube Flows for Hypervelocity Combustion Studies

Diode Laser Sensor for Scramjet Inlets

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