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A parametric numerical study was performed to study effects of the nonuniform premixing of a hydrogen-air mixture on the structures of oblique detonation and shock-induced combustion formed on the surface of a hypersonic spherical projectile. Axisymmetric two-dimensional Navier-Stokes equations were solved with a detailed chemical kinetic mechanism involving nine species. A smooth-front oblique detonation formed on the hypersonic projectile at a Mach number of 6.46 under a uniformly stoichiometric condition was taken as the completely premixed case. For nonuniform cases, hydrogen mole fraction was distributed at the inlet boundary based on the Gaussian function. As the nonuniformity was increased, the following occurred: 1) deformed oblique detonation, 2) oscillating shock-induced combustions with nonuniform corrugated structures, 3) steady shock-induced combustions with (and without) the minimum induction length located outside the centerline, and 4) nose-confined combustion. Analyses on distributions of induction lengths combined with zero-dimensional additional simulations demonstrated that a reactivity gradient became strongly influential on the nonuniformly premixed structures as the nonuniformity was increased, and finally the postshock flowfield became the most determinant factor. [ABSTRACT FROM AUTHOR]

The numerical simulation of meso-, convective-, and microscale atmospheric flows requires the solution of the Euler or the Navier-Stokes equations. Nonhydrostatic weather prediction algorithms often solve the equations in terms of derived quantities such as Exner pressure and potential temperature (and are thus not conservative) and/or as perturbations to the hydrostatically balanced equilibrium state. This paper presents a well-balanced, conservative finite difference formulation for the Euler equations with a gravitational source term, where the governing equations are solved as conservation laws for mass, momentum, and energy. Preservation of the hydrostatic balance to machine precision by the discretized equations is essential because atmospheric phenomena are often small perturbations to this balance. The proposed algorithm uses the weighted essentially nonoscillatory and compact-reconstruction weighted essentially nonoscillatory schemes for spatial discretization that yields high-order accurate solutions for smooth flows and is essentially nonoscillatory across strong gradients; however, the well-balanced formulation may be used with other conservative finite difference methods. The performance of the algorithm is demonstrated on test problems as well as benchmark atmospheric flow problems, and the results are verified with those in the literature. [ABSTRACT FROM AUTHOR]

A possible way to decrease interaction tone noise of a contra-rotating open rotor is through the application of trailing-edge blowing by reducing the momentum deficit of the front rotor wake, and therefore its interaction with the aft rotor. In this contribution, an assessment of front-rotor trailing-edge blowing is presented for the reduction of contra-rotating open-rotor interaction noise. For this purpose, the German Aerospace Center designed a generic contra-rotating open rotor that has been modified to include trailing-edge blowing at the front rotor blades. With the German Aerospace Center computational fluid dynamics code TAU, unsteady Reynolds-averaged Navier-Stokes simulations have been made of the baseline and the trailing-edge blowing configuration. Subsequently, an aeroacoustic analysis has been performed with the Ffowcs-Williams/Hawkings tool APSIM+ for both configurations. The results show negligible differences of the aerodynamic performance, with significantly lower unsteady loading of the aft rotor when the trailing-edge blowing technique is active. The aeroacoustic results illustrate that the interaction tones are notably ameliorated; that is, a reduction over almost the complete polar angle is seen when trailing-edge blowing is active. With respect to the A-weighted overall sound pressure level, a reduction of approximately 2.5 dBA in the maximum level for the trailing-edge blowing configuration is achieved for the considered geometry. Noise reductions up to 5 dBA are displayed in the up- and downstream directions, which mainly arise from reduced interaction tones caused by the trailing-edge blowing technique. The decreased noise radiation with trailing-edge blowing mainly stems from a reduction in the first two interaction tones. [ABSTRACT FROM AUTHOR]

As demonstrated by Parent, B., et al., ("Electron and Ion Transport Equations in Computational Weakly-Ionized Plasmadynamics," Journal of Computational Physics, Vol. 54, 2014, pp. 51-69), the computational efficiency of the drift-diffusion plasma model can be increased significantly by recasting the equations such that the potential is obtained from Ohm's law rather than Gauss's law and by adding source terms to the ion transport equations to ensure that Gauss's law is satisfied. Not only did doing so reduce the stiffness of the system, leading to faster convergence, but it also resulted in a higher resolution of the converged solution. The combined gains in convergence acceleration and resolution amounted to a hundredfold increase in computational efficiency when simulating nonneutral plasmas with significant quasi-neutral regions. In this paper, it is shown that such a recast of the drift-diffusion model has yet another advantage: its lack of stiffness permits the electron and ion transport equations to be integrated in coupled form along with the Favre-averaged Navier-Stokes equations. Test cases relevant to plasma aerodynamics (including nonneutral sheaths, magnetic field effects, and negative ions) demonstrate that the proposed coupled system of equations can be converged in essentially the same number of iterations as that describing nonionized flows while not sacrificing the generality of the drift-diffusion model. [ABSTRACT FROM AUTHOR]

This work presents multiple high-fidelity large-eddy simulations of flow over a cylindrical turret with a flat window oriented at two angles, 90 and 100 deg. For the 100 deg case, additional computations were performed to investigate the effectiveness of multiple types of flow control including rows of short pins, tall pins, and steady blowing wall jets inserted upstream of the turret as well as a steady suction slot at the leading edge of the flat window. The large-eddy simulation computations were obtained using a well-validated high-order Navier-Stokes flow solver employing a sixth-order compact spatial discretization in conjunction with an eighth-order low-pass spatial filter. Overall, large-eddy simulation solutions compared favorably to experimental time mean and fluctuating velocity profiles as well as the general flow structure at both angles. Additionally, valuable insight was obtained on how rows of pins, blowing jets, and steady slot suction control flow separation. Slot suction at the upwind aperture lip was determined to be more effective than rows of pins and blowing jets in controlling flow separation, which is critical in reducing aero-optical aberrations. Steady slot suction was capable of eliminating massive separation at a larger look angle of 120 deg. [ABSTRACT FROM AUTHOR]

The paper presents the developments of novel mesh generation algorithms over complex glaze-ice shapes containing multicurvature ice-accretion geometries, such as single/double ice horns. The twofold approaches tackle surface geometry discretization as well as field mesh generation. First, an adaptive curvilinear curvature control algorithm is constructed, solving a one-dimensional elliptic partial differential equation with periodic source terms. This method controls the arc length grid spacing, so that high convex and concave curvature regions around ice horns are appropriately captured, and is shown to effectively treat the grid shock problem. Second, a novel blended method is developed by defining combinations of source terms with two-dimensional elliptic equations. The source terms include two common control functions, Sorenson and Spekreijse, and an additional third source term to improve orthogonality. This blended method is shown to be very effective for improving grid quality metrics for complex glaze-ice meshes with Reynolds-averaged Navier-Stokes resolution. The performance in terms of residual reduction per nonlinear iteration of several solution algorithms (point-Jacobi, Gauss-Seidel, alternating direction implicit, point, and line Successive Over-Relaxation) are discussed within the context of a full multigrid operator. Details are given on the various formulations used in the linearization process. It is shown that this performance of the solution algorithm depends on the type of control function used. Finally, the algorithms are validated on standard complex experimental ice shapes, demonstrating the applicability of the methods. [ABSTRACT FROM AUTHOR]

The effects of plate angle on acoustic waves from a supersonic jet impinging on an inclined flat plate at angles of 30, 45, and 60 deg are numerically investigated. Three-dimensional compressible Navier-Stokes equations are solved using the modified weighted compact nonlinear scheme. Similar to previous studies, the acoustic fields indicate that there are at least three types of acoustic waves in all of the cases considered herein: 1) Mach waves generated from the shear layer of the main jet, 2) acoustic waves generated from the impingement region, and 3) Mach waves generated from the shear layer of the supersonic flow downstream of the jet impingement. Acoustic waves (2) are generated from two different acoustic sources: 1) the interaction between the plate shock wave and the shear layer, and 2) the interaction between the bubble-induced shock waves and the shear layer. The frequency characteristics of acoustic waves are related to the thickness of the shear layer in the impingement region. The results of the present study indicate the source location and the characteristics of acoustic waves (2) for various flat-plate angles. [ABSTRACT FROM AUTHOR]

Large-eddy simulations of a wing-body junction flow experimentally studied are presented. Wall junction flows are common in engineering applications, but the flow physics at the corners of the junction is not well understood. Moreover, in these types of flows, the performance of the subgrid-scale models frequently used for large-eddy simulations is not well characterized. To address these issues, large-eddy simulations of the wing-body junction are performed, with multiple levels of grid resolution. Two-dimensional turbulent profiles are generated every time step and introduced to the inlet plane of the computation domain to mimic the unsteady turbulent boundary layer. A Reynolds-averaged Navier-Stokes calculation is used as a precursor to initialize the flowfield of the large-eddy simulations. Then, large-eddy simulations with the Vreman model is performed and the simulation results are analyzed with respect to three specific goals: compare the computed and measured turbulence statistics in the junction region, investigate the flow physics in the corner region, and discuss the potential reasons for the inaccuracies of the subgrid-scale model. The sensitivity of the predicted results to inflow conditions and to subgrid-scale modeling is also investigated. [ABSTRACT FROM AUTHOR]

The paper presents a computational and an experimental investigation of base flow of a circular cylinder at hypersonic speeds. Effects of chemistry and wall temperature on the flow in the base region, at low to high enthalpies, are discussed. The experiments were conducted in a shock tunnel at a nominal Mach number of 10. Freestream Reynolds numbers based on cylinder diameter were 0.97×104 and 3.74×104, respectively, and the total specific enthalpies were 13.35 and 3.94 MJ/kg, respectively. The test gas was air. The surface pressure and heat flux were measured using a cold wall model. Equilibrium and thermal as well as chemical nonequilibrium numerical simulations were performed using a Navier-Stokes equations-based computational fluid dynamics code. Both a cold wall and adiabatic wall were considered. Particular emphasis was placed on the wake structure, vorticity distribution, wake centerline aerothermodynamic properties, and surface data. The existing low-enthalpy cold hypersonic wind-tunnel experimental data are included for comparison. The simulations predicted the effect of chemistry on the near wake to be negligible for the low-enthalpy, high Reynolds number flow but more significant for the high-enthalpy, low Reynolds number flow. [ABSTRACT FROM AUTHOR]

his study presents hybrid Reynolds-averaged Navier-Stokes/large-eddy simulations of the unsteady flow and noise-generation phenomena in the slat cove of a high-lift wing profile. These computations are part of a joint numerical/experimental aeroacoustics collaborative program dedicated to slat-flow analysis. A dedicated two-element wing profile (slat plus main body) has been designed to isolate slat noise from other possible sources (e.g., the flap), while minimizing mean flow deflection effects, to improve the fidelity of open-jet wind-tunnel measurements. The design of this two-element airfoil has been performed numerically, using an optimization process based on steady Reynolds-averaged Navier-Stokes calculations. This airfoil has been investigated experimentally at the École Centrale de Lyon open jet facility. Unsteady zonal hybrid Reynolds-averaged Navier-Stokes/large-eddy simulations have been performed to provide a comprehensive description of the unsteady flow inside the slat cove, focusing on the noise-generation processes. A detailed analysis of the physics of the unsteady flow inside the slat cove is presented as well as a comparison of numerical results with available experimental data. The pressure spectra associated with the slat-cove flow are characterized by several tonal peaks emerging from the underlying broadband content. The existence of such peaks is attributed to a feedback loop involving the main shear layer inside the slat cove. A theoretical law is proposed to predict the associated tonal frequencies and assessed at the end of the paper. [ABSTRACT FROM AUTHOR]

A numerical study of the transient deformation and drag properties of impulsively started decelerating drops in axisymmetric flows is conducted. The incompressible Navier-Stokes equations coupled with the volume of fraction method are solved. The differences in the deformation and drag characteristics between the accelerating and decelerating drops are elucidated. The Weber number We has a significant influence on the drag properties of the round drops; however, the effect of the Ohnesorge number Oh is small. A dynamic droplet drag model is provided based on the computational fluid dynamics results to predict the drag coefficient of a decelerating drop. In addition, the flow of liquid drops past an RAE2822 airfoil is considered, and its drag is evaluated using the drag model. [ABSTRACT FROM AUTHOR]

Comparative studies of different inflow turbulent forcing effects on detonation front dynamics and the flow statistics are conducted through direct numerical simulations of turbulence-detonation/shock interactions. High-resolution bandwidth-optimized weighted essentially nonoscillatory scheme of spatial discretization and total variation diminishing temporal integration are used to solve the three-dimensional chemically reactive Navier-Stokes equations. The turbulent inflow vertical and entropic forcing effects on the three-dimensional detonation front and cellular structures are first analyzed. The influence of the inflow forcing on the generated regular cell patterns with diamond shape is compared. The vortex behind the detonation is quite different from that behind shock waves with the same inflow fluctuations. This also results in the larger increase of velocity variances downstream of detonation than the turbulence-shock case. The similar trend can also be found for the Taylor microscales. Effects of different inflow turbulence intensities are investigated. The irregular cellular detonations under vertical inflow forcing are also analyzed. [ABSTRACT FROM AUTHOR]

This study gives an insight into the corner flow around a wing mounted on a flat plate. Reynolds-averaged Navier-Stokes simulations and experimental data are compared for a Reynolds number based on the wing chord (Rec equal to 3×105 and an angle of attack of 12 deg. The flat plate and wing boundary layers are fully turbulent, and the Reynolds number based on the momentum thickness of the incoming boundary layer is equal to Re=2200 It is shown that the Reynolds-averaged Navier-Stokes simulation using the Spalart-Allmaras model with the Boussinesq closure predicts a massive corner flow separation, whereas on the experimental side, the separation is confined to a small area. Better predictions are obtained when resorting to the Spalart-Allmaras model with the quadratic constitutive relation closure. A comprehensive investigation of the numerical results using the experimental data is performed to get a better understanding of the behavior of the Spalart-Allmaras quadratic constitutive relation model for the present corner flow. [ABSTRACT FROM AUTHOR]

A set of compact disturbance equations is developed for high-accuracy and efficient aeroacoustic simulations. The compact disturbance equations in their complete form are an exact rearrangement of the Navier-Stokes equations, but incorporate various linear and nonlinear disturbance equations such as the linearized Euler equations and the linearized Navier-Stokes equations. Their attractive mathematical properties facilitate the implementations of the full compact disturbance equations and the reduced equations in essentially the same form as in existing computational fluid dynamics solvers with minor modifications. A high-resolution unsteady simulation in a reduced domain can be embedded inside a less-expensive Reynolds-averaged Navier-Stokes solution of flow in a larger, possibly very complex, configuration. This results in a hybrid Reynolds-averaged Navier-Stokes/large-eddy simulation method with reduced meshing difficulties and computational costs, but improved grid qualities and more accurate boundary treatments for complex configurations. Furthermore, a seamless switch can be made between the embedded governing equations. This enables a novel closely coupled computational fluid dynamics/computational aeroacoustics approach for complex aeroacoustic applications, where the full Navier-Stokes equations can be recovered in the source region to capture the turbulent noise sources and the linearized Euler equations can be applied to simulate the noise propagation and possible reflection more accurately and efficiently. These benefits are demonstrated with three benchmark tests and excellent results are obtained. [ABSTRACT FROM AUTHOR]

The aerodynamic shape optimization of wings in transonic flow is an inherently challenging problem. In addition to the high computational cost of solving the Reynolds-averaged Navier-Stokes equations, there is a complex interdependence between the cross-sectional shape, wave drag, and viscous effects. Furthermore, it is necessary to perform multipoint optimizations to ensure good performance for a range of flight conditions. The choice of which flight conditions should be considered in a multipoint optimization and how many of these should be considered is still not well understood. This paper addresses this issue by solving a series of seven benchmark optimizations developed by the AIAA Aerodynamic Optimization Discussion Group. These optimization cases include a single-point optimization, four three-point optimizations, a nine-point optimization, and a five-point optimization. The optimizations consist in minimizing the weighted drag coefficient subject to lift, moment, thickness, and volume constraints. The optimizations were performed with respect to 768 shape design variables and an angle of attack for each flight condition. The single-point optimization was able to achieve an 8.1% drag reduction relative to the initial design, but it exhibited poor off-design performance. All the optimized designs were compared using a contour plot of ML/cD to evaluate the wing performance over the complete transonic flight operating envelope. Each of the four three-point optimizations successfully mitigated the poor off-design performance of the single-point design. However, the three-point optimization with widely spaced Mach numbers yielded a much more complex ML/cD contour with two distinct local maxima. Finally, the five- and nine-point optimizations yielded similar performance and the most robust off-design performance. [ABSTRACT FROM AUTHOR]