Ignition and Flame Structure in a Compact Inlet/Scramjet Combustor Model

Gamba, Mirko; Miller, Victor A.; Mungal, M. G.; Hanson, Ronald K.

doi:10.2514/6.2011-2366

Cited by 14 publications

(17 citation statements)

References 9 publications

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“…The unstructured solver 4 has been used for preliminary non-reacting simulations of a scramjet inlet geometry (figure 10, based on the experiments by Gamba et al 19 ). High fidelity simulations of (detailed-chemistry) reacting turbulent flow in this geometry will provide very useful understanding of issues pertinent to scramjet engines.…”

Section: Summary and Future Workmentioning

confidence: 99%

A Numerical Method for DNS of Turbulent Reacting Flows Using Complex Chemistry

Asaithambi

Muppidi

Mahesh

2012

42nd AIAA Fluid Dynamics Conference and Exhibit

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A novel compressible Navier-Stokes solver for chemically reacting flows is proposed with an explicit Navier-Stokes predictor step and a semi-implicit corrector step for stiff chemical source terms. A modular code to read chemical mechanisms in the Chemkin format is coupled to the solver allowing the ability to simulate multiple fuels with minimal effort. This segregated approach allows the independent modification of the Navier-Stokes solver and the chemical source term integration algorithm. Validation of a well-stirred reactor, an unsteady unstrained diffusion flame, and results from a lifted non-premixed flame in vitiated coflow in two and three dimensional configurations are presented.

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Section: Summary and Future Workmentioning

confidence: 99%

A Numerical Method for DNS of Turbulent Reacting Flows Using Complex Chemistry

Asaithambi

Muppidi

Mahesh

2012

42nd AIAA Fluid Dynamics Conference and Exhibit

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“…With the enthalpy and entropy known (the entropy is found from the stagnation conditions since the flow is assumed isentropic) for the guessed value of throat velocity u*, the speed of sound can be obtained from the equilibrium air curve fit for a = a(h,s) 9 . Because sonic flow exists at the throat of an equilibrium chemically reacting nozzle, the velocity of the flow at the throat must be equal to the local speed of sound; therefore, if the calculated speed of sound at the throat, a, doesn't equal the guessed throat velocity, u*, an iterative root-finding algorithm must be applied until the difference between the calculated speed of sound and guessed throat velocity converges to an acceptable tolerance.…”

Section: ( )mentioning

confidence: 99%

“…To find the velocity at the exit plane of the nozzle A/A* must first be set to the value of A e /A t for the nozzle, meaning the product will correspond to properties at the nozzle exit, . A root finding algorithm and the equilibrium air curve fit for ρ = ρ(h,s) 9 , along with equation 1 rewritten for the enthalpy and velocity at the nozzle exit, are used iterating over the exit velocity, u e , until the ratio Minimum pressure on the right-hand side of equation 2 converges acceptably to the known ratio A e /A t . With the velocity, enthalpy, and entropy calculated at the nozzle exit the remaining properties are found via equilibrium air curve fits.…”

Section: ( )mentioning

confidence: 99%

Design of a Model Scramjet Combustor for Vortex-Enhanced Mixing and Combustion Studies

Ground

Zhu

Maddalena³

2014

19th AIAA International Space Planes and Hypersonic Systems and Technologies Conference

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The present work details the preliminary analysis and design of a model scramjet combustor intended to serve as a research platform for mixing and combustion studies in support of current non-reactive research. The shock tunnel located at the Aerodynamics Research Center of The University of Texas at Arlington, capable of producing test conditions simulating scramjet flight at Mach numbers ranging from ~5 to ~16 at altitudes of ~24 km to ~85 km, will serve as the test facility for the model combustor. Specifically, the modular design of the model scramjet combustor will allow for the study of supersonic combustion and flameholding characteristics of specific fuel injection strategies based on enhanced fuel-air mixing via pre-imposed selected modes of streamwise vortex interactions and consequent dynamics. In this paper, the fundamental requirements for the combustor design and the performance of the UTA hypersonic tunnel are discussed. A parametric sizing of the model combustor is utilized to define a suitable design that meets the specified requirements, and the inevitable constraints imposed by the limitation of ground testing facilities, discussed in the manuscript. Nomenclature= Area γ = Specific heat ratio h = Enthalpy M = Mach number P = Pressure ψ = Combustor to freestream flight static temperature ratio ρ = Density η = Adiabatic compression efficiency T = Temperature t = time θ = Inlet compression angle u = Velocity Subscripts 1 = Driven condition 4 = Driver condition 5 = Stagnation condition c = Combustor condition e = Nozzle exit condition i = ignition r = reaction * = Throat condition ∞ = Freestream condition

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“…Transverse jet in supersonic cross-flow (JISCF) is one of the most fundamental flow phenomena in supersonic combustion chambers [1][2][3][4][5], such as the actual flight vehicle Hyshot II [6,7] and the supersonic combustion experiments by Gamba et al [8,9]. In JISCF combustion chambers, there exist complex shock train, flame, turbulence and boundary layer interactions [6][7][8][9][10][11][12][13][14][15]. In addition to auto-ignition due to high-enthalpy and shock aerodynamic heating, local extinction caused by turbulent small-scale vortex and re-ignition, there is also strong coupling between the density, pressure, temperature and velocity.…”

Section: Introductionmentioning

confidence: 99%

Large-Eddy Simulations of Transverse Jet Mixing and Flame Stability in Supersonic Crossflow

Zhao

2021

AIAA Journal

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Large-eddy simulations (LES) combined with Partially Stirred Reactor (PaSR) combustion model are employed to investigate the turbulent mixing, combustion mode and flame stability of a sonic hydrogen jet injecting into high-enthalpy supersonic crossflow at three momentum flux ratios J, i.e. 0.71, 2.11 and 4.00, respectively. The LES accuracy in terms of the turbulent kinetic energy, power spectra density and sub-grid Damköhler number is carefully addressed against various LES resolution criteria and the experimental mean pressure distribution on the upper wall. The ignition processes with auto-ignition and shock compression effects are identified and analyzed. At J = 0.71, the shock-induced ignition occurs behind the reflected shock wave and the combustion heat release is dominated by the premixed combustion. While for the high jet to cross-flow momentum flux ratios, e.g., J = 2.11 and 4.00, ignition happens toward around the jet orifice due to the strong bow shock and reflected shock effects and the combustion heat releases are dominated by the non-premixed combustion. Furthermore, the mechanisms of flame stabilization, local extinction and re-ignition in the transverse jet combustion in supersonic cross-flow are further analyzed with the chemical explosive mode analysis (CEMA).

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Ignition and Flame Structure in a Compact Inlet/Scramjet Combustor Model

Cited by 14 publications

References 9 publications

A Numerical Method for DNS of Turbulent Reacting Flows Using Complex Chemistry

A Numerical Method for DNS of Turbulent Reacting Flows Using Complex Chemistry

Design of a Model Scramjet Combustor for Vortex-Enhanced Mixing and Combustion Studies

Large-Eddy Simulations of Transverse Jet Mixing and Flame Stability in Supersonic Crossflow

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