Development of the US3D Code for Advanced Compressible and Reacting Flow Simulations

Candler, Graham V.; Johnson, Heath B.; Nompelis, Ioannis; Subbareddy, Pramod K.; Drayna, Travis; Gidzak, Vladimyr M.; Barnhardt, Michael

doi:10.2514/6.2015-1893

Cited by 147 publications

(44 citation statements)

References 47 publications

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“…The cone wall is modeled as a smooth body with a flare radius of 3 m. The initial half-angle is 1.5 deg, the total length is 516.89 mm, the half-base is 57.15 mm, and the nose-tip radius is 0.15 mm. Basic state simulations are conducted (the same as those considered in [28]) using US3D [29] to represent the Mach 6 Quiet Tunnel conditions at Purdue University [30,31]. The grid consists of roughly 1.25 million total grid points, with 1000 streamwise points, 626 wall-normal points, and a 1 deg of rotation in the azimuthal direction with symmetry boundary conditions to simulate 0 angles of attack.…”

Section: Base State Flowmentioning

confidence: 99%

Thermoacoustic Interpretation of Second-Mode Instability

Kuehl

2018

AIAA Journal

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A thermoacoustic interpretation of Mack's second-mode instability is proposed. It is demonstrated that the fundamental mechanism of second-mode wave growth in hypersonic boundary layers is consistent with standingwave thermoacoustically driven instability. The resonant nature of such modes is sustained by an acoustic impedance well between the wall (infinite impedance) and near the sonic line (secondary peak in impedance). A Lagrangian approach is adopted to show that such resonant standing waves derive energy from the base flow through thermoacoustic Reynolds stress, which results from the divergence of acoustic power inside the impedance well, and thermodynamic work. This treatment does not represent a complete energy closure (due to the neglect of viscosity), but it does provide insight toward a fundamental energy source and the physical mechanisms governing Mack's acoustic second mode.

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Section: Base State Flowmentioning

confidence: 99%

Thermoacoustic Interpretation of Second-Mode Instability

Kuehl

2018

AIAA Journal

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“…For more complex kinetics models, the EVM approach will be correspondingly more efficient. Furthermore, the present numerical method has the same numerical stability properties as the well-established detailed simulation approach [27].…”

Section: Computational Costmentioning

confidence: 88%

“…Therefore, implicit time-integration methods are used in the present work. For RANS simulations, the data-parallel line-relaxation method [45] is used, and for the wallmodeled LES, a second-order-accurate version of an implicit pointrelaxation method [46] is employed; see also the work of Candler et al [27,39]. As stated previously, the derivatives of the evolutionvariable source term ζ are tabulated and used to form the Jacobian of the source term for use in the implicit operator.…”

Section: Time Integrationmentioning

confidence: 99%

“…In the present work, we use standard approaches for the first three elements, and we focus on the accurate and efficient modeling of turbulent combustion at conditions relevant to high-speed flight. To this end, we modify an existing computational fluid dynamics code [27] for the wall-modeled LES of compressible reacting flows to represent the detailed thermochemical state of the reacting fuel-air mixture. The focus is on the development of a numerical method that is rigorously consistent with the thermochemical model so that the numerical method correctly reflects the variations in the thermodynamic state.…”

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confidence: 99%

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Wall-Modeled Large-Eddy Simulation of Autoignition-Dominated Supersonic Combustion

2017

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Simulations of combustion in high-speed and supersonic flows need to account for autoignition phenomena, compressibility, and the effects of intense turbulence. In the present work, the evolution-variable manifold framework of Cymbalist and Dimotakis ("On Autoignition-Dominated Supersonic Combustion," AIAA Paper 2015-2315, June 2015) is implemented in a computational fluid dynamics method, and Reynolds-averaged Navier-Stokes and wall-modeled large-eddy simulations are performed for a hydrogen-air combustion test case. As implemented here, the evolution-variable manifold approach solves a scalar conservation equation for a reaction-evolution variable that represents both the induction and subsequent oxidation phases of combustion. The detailed thermochemical state of the reacting fluid is tabulated as a low-dimensional manifold as a function of density, energy, mixture fraction, and the evolution variable. A numerical flux function consistent with local thermodynamic processes is developed, and the approach for coupling the computational fluid dynamics to the evolution-variable manifold table is discussed. Wall-modeled large-eddy simulations incorporating the evolution-variable manifold framework are found to be in good agreement with full chemical kinetics model simulations and the jet in supersonic crossflow hydrogen-air experiments of Gamba and Mungal ("Ignition, Flame Structure and Near-Wall Burning in Transverse Hydrogen Jets in Supersonic Crossflow," Journal of Fluid Mechanics, Vol. 780, Oct. 2015, pp. 226-273). In particular, the evolutionvariable manifold approach captures both thin reaction fronts and distributed reaction-zone combustion that dominate high-speed turbulent combustion flows.Nomenclature a = speed of sound, m∕s C; X ; Z = progress variable, nonfuel mass fraction, and fuel mass fraction c v ; c v;s = mixture and s-species specific heats, J∕kg ⋅ K D = diffusion coefficient, m 2 ∕s E = total energy per unit volume, J∕m 3 e; e s = mixture and s-species specific energy, J∕kg F = convective flux vector h; h 0 = enthalpy and total enthalpy, J∕kg i = grid index J = jet momentum ratio j h = diffusive enthalpy flux k = kinetic energy, J∕kg L; R = values obtained from left and right data M s = s-species molar mass, kg∕kmol N s = number of species in detailed kinetics model n; n x ; n y ; n z = element face unit normal vector and components p = pressure, Pa q j = heat flux vector= vectors of conserved and primitive variables u; u; v; w = velocity vector and components, m∕s u 0 = face-normal velocity component, m∕s x; x j = position vector and its components, m Y s = s-species mass fraction α = dissipative flux factor δ R = reaction-zone thickness, m ε = dissipation rate ζ = evolution-variable source term, 1∕s η k = Kolmogorov length scale, m Λ; λ = diagonal matrix of eigenvalues and eigenvalue ν t ; ν = turbulence field variable and kinematic viscosity ρ; ρ s = density and s-species density, kg∕m 3 σ ij = viscous and Reynolds stress τ = evolution variable ϕ = stoichiometric fuel-air ratio χ = subgrid-s...

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“…The mean flow is computed using the US3D [94] CFD solver. The domain was a quarter-volume of the half-cylinder, with 200 grid points in the wall-normal direction and a total of 5.5 million cells.…”

Section: Computational Fluid Dynamics Solvermentioning

confidence: 99%

Hypervelocity shock layer emission spectroscopy with high temperature ablating models

Lewis¹

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During reentry into the atmosphere, a vehicle can encounter extreme convective and possibly radiative heating loads. Such spacecraft must rely on a carefully designed thermal protection system to maintain the integrity of the underlying structure, thus ensuring the survival of any cargo or crew it may contain. Despite many past developments and research efforts, to this day there remain large uncertainties associated with the prediction of heat shield performance. In particular, accurate modelling of ablation phenomena is critical to minimising uncertainties, reducing excess weight, and increasing safety. This includes thermochemical ablation by oxidation, nitridation, and sublimation, as well as the mechanical removal of material via spallation. Due to the high costs associated with in-flight testing, and the difficulty of mounting precision instruments and advanced diagnostics on flight vehicles, ground-based experiments stand as a cost-effective method for reducing modelling uncertainties.Ablation testing is generally conducted in long-duration facilities such as arc jets, which are well suited to investigating in-depth material response due to their ability to provide representative heating rates for extended periods. They are, however, unable to reproduce a realistic in-flight shock layer due to low Reynolds numbers and, depending on the facility type, a high degree of thermal and chemical nonequilibrium in the free-stream. In contrast, impulse facilities such as shock tubes and their variants are well suited to reproducing realistic flight conditions with free-streams that are a closer match for flows dominated by binary processes in terms of Reynolds number, temperature, and chemical composition. Due to their extremely short test times, however, ranging from the order of 10 µs to several milliseconds, they are unable to aerothermally heat test models up to representative in-flight temperatures, or recreate the quasi-steady heat and mass flux balance of flight. This key limitation of impulse facilities was recently overcome in experiments using the University of Queensland's X2 expansion tunnel by electrically pre-heating samples to temperatures approaching 2500 K immediately prior to a test.In this work, the pre-heating methodology was enhanced to generate maximum surface temperatures approaching 3300 K in order to study sublimation phenomena. It was found that although sublimation-regime temperatures could be achieved, the relevant surface processes themselves were suppressed due to the high pitot pressures of the conditions used. Instead, the nitridation process was studied by observing air shock layer CN emissions with pure graphite models heated to surface temperatures ranging from 1770-3300 K. These results were compared with numerical simulations which predicted a monotonic increase of emissions with surface temperature for all models considered, whereas the experiments showed an initial increase from 1770-2500 K and then remained relatively constant from 2500-3300 K. The simulations also su...

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Development of the US3D Code for Advanced Compressible and Reacting Flow Simulations

Cited by 147 publications

References 47 publications

Thermoacoustic Interpretation of Second-Mode Instability

Thermoacoustic Interpretation of Second-Mode Instability

Wall-Modeled Large-Eddy Simulation of Autoignition-Dominated Supersonic Combustion

Hypervelocity shock layer emission spectroscopy with high temperature ablating models

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