Transverse jet mixing and combustion experiments in hypervelocity flows

Bélanger, Jacques; Hornung, H. G.

doi:10.2514/3.24009

Cited by 27 publications

(10 citation statements)

References 6 publications

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“…6 Various kinds of injection concepts have been investigated in the literature to reduce separation zones and accelerate penetration into the freestream. 7 ¡ 10 Transverse fuel injections in most cases cause a strong blockage of the main ow, 11 which, associatedwith irreversibilitiesof the resultingshock waves, lead to strong thrust losses. 5 On the other hand, good near-eld mixing is possible.…”

Section: Introductionmentioning

confidence: 99%

Numerical Investigation of Hydrogen Strut Injections into Supersonic Airflows

Gerlinger

Brüggemann

2000

Journal of Propulsion and Power

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The in uence of different injector geometries on the mixing behavior of planar supersonic jets is investigated. Hydrogen is injected in the ow direction through the blunt end of a strut. Different lip thicknesses at the injector end induce different extents of recirculation zones. Additionally, changes are caused within the shock-wave/expansion fan pattern at the injector exit that have an important in uence on loss in total pressure. The accuracy of the used numerical scheme is demonstrated for one injector geometry by comparison with experimental data. Based on this con guration, modi cations in lip thickness and injection Mach number are investigated numerically and are assessed using several calculated performance parameters. Results show that the chosen lip thickness has a much stronger in uence on loss in total pressure than moderate modi cations in injector length and height. For the very thin hydrogen jets that were investigated, the mixing ef ciency is nearly independent on injector lip thickness.

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Section: Introductionmentioning

confidence: 99%

Numerical Investigation of Hydrogen Strut Injections into Supersonic Airflows

Gerlinger

Brüggemann

2000

Journal of Propulsion and Power

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“…For this reason, aerodynamic effects related to this kind of actuator have been accurately investigated in the last decades. Because of the considerable experimental [3,5,6] and numerical work [7][8][9][10][11][12] accomplished on this subject, the nature of these effects and the structure of the interference between the lateral jet and crossflow are currently well known [13][14][15][16][17][18][19][20][21].…”

Section: Introductionmentioning

confidence: 99%

Substitution of Hot-Gas Lateral Jets by Cold-Gas Jets in Supersonic Flows

Sourgen

Gauthier

Leopold

et al. 2011

Journal of Spacecraft and Rockets

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The supersonic flow around a rocket piloted by thrusters has been investigated. Steady Reynolds-averaged Navier-Stokes computations and pressure measurements in a wind tunnel, using pressure taps and pressuresensitive paint, have been performed. For experimental investigations, real hot-gas thrusters have been replaced by cold-gas jets using a lower total pressure and a different gas. The uniqueness of this work involved calibrating those cold-gas thrusters in order to reproduce an aerodynamic interference very similar to the hot-gas jets case. This has been achieved by modifying the section ratio of the nozzles in order to control the wave celerity ratio between freestream and jet flows. A good agreement has been obtained between numerical simulations of hot-gas and cold-gas jets flows, as well as between computations and measurements of the pressure coefficient around the body in the wind tunnel. Thus, this method has provided a good approximation for the interaction between hot-gas jets and crossflow by using lower-pressure and lower-temperature jets.the body, m J = momentum flux similarity number JR = momentum flux ratio M = mass flux similarity number M th = thruster ejection Mach number M 1 = freestream Mach number NCR = nozzle celerity ratio NJR = nozzle momentum flux ratio NMR = nozzle mass flux ratio NPR = nozzle pressure ratio NrTR = nozzle temperature ratio P = pressure similarity number PR = total pressure ratio P j = nozzle exit static pressure, Pa P o , P 1 = freestream static pressure, Pa P t = total pressure, Pa P ti = thruster stagnation pressure, Pa Re D = diameter-based Reynolds number r th = specific thruster gas constant, J kg 1 K 1 r 1 = specific freestream gas constant, J kg 1 K 1 T = temperature similarity number T j = nozzle exit static temperature, K T t = total temperature, K T ti = thruster stagnation temperature, K T 1 = freestream static temperature, K U = velocity component along the x axis, m s 1 U o , U 1 = freestream velocity component along the x axis, m s 1 u f = friction velocity component along the x axis, m s 1 W = velocity component along the mean ejection direction, m s 1 W j = nozzle exit velocity component along the mean ejection direction, m s 1 x = horizontal coordinate along the body axis, m y = horizontal coordinate perpendicular to the body axis, m y = dimensionless wall distance z = vertical coordinate, m th = thruster gas isentropic coefficient 1= freestream gas isentropic coefficient = kinematic viscosity, m 2 s 1 = density, kg m 3 j = nozzle exit density, kg m 3 1 = freestream density, kg m 3 w = wall shear, Pa

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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%

“…Examples of existing impulse-type facilities include 1) the free-piston re ected shock tunnel, T5, located at the Guggenheim AeronauticalLaboratory,California Institute of Technology 11 ; 2) Calspan re ected shock tunnel 9 ; and 3) the expansion tube HYPULSE located at the General Applied Science Laboratories. 6¡8 In these facilities, generic combustor models with hydrogen injection have been tested using conventional measurement techniques such as pressure measurements along the combustor and ow visualization with differential interferometry.…”

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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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Transverse jet mixing and combustion experiments in hypervelocity flows

Cited by 27 publications

References 6 publications

Numerical Investigation of Hydrogen Strut Injections into Supersonic Airflows

Numerical Investigation of Hydrogen Strut Injections into Supersonic Airflows

Substitution of Hot-Gas Lateral Jets by Cold-Gas Jets in Supersonic Flows

Characterization of Expansion Tube Flows for Hypervelocity Combustion Studies

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