Showing papers in "AIAA Journal in 2010"
TL;DR: The use of large-eddy simulation (LES) methods for calculation of turbulent flows has increased substantially in recent years as discussed by the authors, and the availability of LES and hybrid Reynolds-averaged Navier-Stokes (RANS)/LES in general-purpose codes is discussed.
Abstract: Usage of large-eddy simulation (LES) methods for calculation of turbulent flows has increased substantially in recent years. This paper attempts to 1) provide an assessment of the current capabilities of LES, 2) outline some recommended practices for using LES, and 3) identify future research needs. The assessment considers flow problems for which LES can be successfully applied today and flow problems for which LES still has limitations. The availability of LES and hybrid Reynolds-averaged Navier-Stokes (RANS)/LES in general-purpose codes is discussed. Several important issues for which the LES community has not yet reached a consensus are discussed. These include grid sensitivity studies, application of unstructured grid methods, upwind-biased solvers, and turbulence (subgrid) modeling including continuous hybrid RANS/LES approaches. A section on recommended practices and key considerations tries to provide guidance on some of the important items that need to be addressed in using LES. The paper concludes with a discussion of future research directions, with a focus on work needed to advance the capabilities and reliability of LES for analysis of turbulent flows.
TL;DR: In this paper, the tradeoff between computational cost and accuracy is evaluated for aerothermoelastic analysis based on either quasi-static or time-averaged dynamic fluid-thermal-structural coupling, as well as computational fluid dynamics based reduced-order modeling of the aerodynamic heat flux.
Abstract: The field of aerothermoelasticity plays an important role in the analysis and optimization of airbreathing hypersonic vehicles, impacting the design of the aerodynamic, structural, control, and propulsion systems at both the component and multi-disciplinary levels. This study aims to expand the fundamental understanding of hypersonic aerothermoelasticity by performing systematic investigations into fluid-thermal-structural coupling, and also to develop frameworks, using innovative modeling strategies, for reducing the computational effort associated with aerothermoelastic analysis. Due to the fundamental nature of this work, the analysis is limited to cylindrical bending of a simply-supported, von K arm an panel. Multiple important effects are included in the analysis, namely: 1) arbitrary, nonuniform, in-plane and through-thickness temperature distributions, 2) material property degradation at elevated temperature, and 3) the effect of elastic deformation on aerodynamic heating. It is found that including elastic deformations in the aerodynamic heating computations results in non-uniform heat flux, which produces non-uniform temperature distributions and non-uniform material property degradations. This results in reduced flight time to the onset of flutter and localized regions in which the material temperature limits may be exceeded. Additionally, the trade-off between computational cost and accuracy is evaluated for aerothermoelastic analysis based on either quasi-static or time-averaged dynamic fluid-thermal-structural coupling, as well as computational fluid dynamics based reduced-order modeling of the aerodynamic heat flux. It is determined that these approaches offer the potential for significant improvements in aerothermoelastic modeling in terms of efficiency and/or accuracy.
TL;DR: In this article, a large-eddy simulation of an underexpanded sonic jet injection into supersonic crossflows is performed to obtain insights into key physics of the jet mixing.
Abstract: Large-eddy simulation of an underexpanded sonic jet injection into supersonic crossflows is performed to obtain insights into key physics of the jet mixing. A high-order compact differencing scheme with a recently developed localized artificial diffusivity scheme for discontinuity-capturing is used. Progressive mesh refinement study is conducted to quantify the broadband range of scales of turbulence that are resolved in the simulations. The simulations aim to reproduce the flow conditions reported in the experiments of Santiago and Dutton [Santiago, J. G., and Dutton, J. C., "Velocity Measurements of a Jet Injected into a Supersonic Crossflow," Journal of Propulsion and Power, Vol.132,1997, pp. 264―273] and elucidate the physics of the jet mixing. A detailed comparison with these data is shown. Statistics obtained by the large-eddy simulation with turbulent crossflow show good agreement with the experiment, and a series of mesh refinement studies shows reasonable grid convergence in the predicted mean and turbulent flow quantities. The present large-eddy simulation reproduces the large-scale dynamics of the flow and jet fluid entrainment into the boundary-layer separation regions upstream and downstream of the jet injection reported in previous experiments, but the richness of data provided by the large-eddy simulation allows a much deeper exploration of the flow physics. Key physics of the jet mixing in supersonic crossflows are highlighted by exploring the underlying unsteady phenomena. The effect of the approaching turbulent boundary layer on the jet mixing is investigated by comparing the results of jet injection into supersonic crossflows with turbulent and laminar crossflows.
TL;DR: In this article, the flowfield downstream of a strut-based injection system in a supersonic combustion ramjet is investigated using large-eddy simulation with a new localized dynamic subgrid closure for compressible turbulent mixing.
Abstract: The flowfield downstream of a strut-based injection system in a supersonic combustion ramjet is investigated using large-eddy simulation with a new localized dynamic subgrid closure for compressible turbulent mixing. Recirculations are formed at the base of the strut in the nonreacting flow and trap some of the injected fluid. The high levels of turbulence along the underexpanded hydrogen jets and in the shear layer lead to a high level of mixing of fuel and freestream fluids. Furthermore, the shear layer unsteadiness permits efficient large-scale mixing of freestream and injected fluids. In the reacting flowfield, the flame anchoring mechanism is, however, found to depend more on a recirculation region located downstream of the injectors than on their sides. A region of reverse flow is formed that traps hot products and radicals. Intermittent convection of hot fluid toward the injector occurs and preheats the reactants.
TL;DR: In this article, a point-collocation non-intrusive polynomial chaos technique is used for uncertainty propagation in computational fluid dynamics simulations, where the input uncertainties are propagated with both the non-inrusive Polynomial Chaos method and Monte Carlo techniques to obtain the statistics of various output quantities.
Abstract: This paper describes a point-collocation nonintrusive polynomial chaos technique used for uncertainty propagation in computational fluid dynamics simulations. The application of point-collocation nonintrusive polynomial chaos to stochastic computational fluid dynamics is demonstrated with two examples: 1) a stochastic expansion-wave problem with an uncertain deflection angle (geometric uncertainty) and 2) a stochastic transonic-wing case with uncertain freestream Mach number and angle of attack. For each problem, input uncertainties are propagated with both the nonintrusive polynomial chaos method and Monte Carlo techniques to obtain the statistics of various output quantities. Confidence intervals for Monte Carlo statistics are calculated using the bootstrap method. For the expansion-wave problem, a fourth-degree polynomial chaos expansion, which requires five deterministic computational fluid dynamics evaluations, has been sufficient to predict the statistics within the confidence interval of 10,000 crude Monte Carlo simulations. In the transonic-wing case, for various output quantities of interest, it has been shown that a fifth-degree point-collocation nonintrusive polynomial chaos expansion obtained with Hammersley sampling was capable of estimating the statistics at an accuracy level of 1000 Latin hypercube Monte Carlo simulations with a significantly lower computational cost. Overall, the examples demonstrate that the point-collocation nonintrusive polynomial chaos has a promising potential as an effective and computationally efficient uncertainty propagation technique for stochastic computational fluid dynamics simulations.
TL;DR: In this article, the characteristics of the pulsed-plasma jet issuing into stagnant air at a pressure of 35 torr were determined using optical emission spectroscopy (OES).
Abstract: thepulsed-plasmajet,thejetisinjectednormallyintoaMach3crossflowandthepenetrationdistanceismeasuredby using schlieren imaging. These measurements show that the jet penetrates 1.5 boundary-layer thicknesses into the crossflow andthe jet-to-crossflowmomentum fluxratioisestimated to be0.6.Aseries of experiments wasconducted to determine the characteristics of the pulsed-plasma jet issuing into stagnant air at a pressure of 35 torr. These resultsshowthattypicaljetvelocitiesofabout250 m=scanbeinducedwithdischargeenergiesofabout30mJperjet. Furthermore, the maximum pulsing frequency was found to be about 5 kHz, because above this frequency the jet beginstomisfire.Themisfiringappearstobeduetothe finitetimeittakesforthecavitytoberechargedwithambient air between discharge pulses. The velocity at the exit of the jet is found to be primarily dependent on the discharge current and independent of other discharge parameters such as cavity volume and orifice diameter. Temperature measurementsaremadeusingopticalemissionspectroscopyandrevealthepresenceofconsiderablenonequilibrium between rotational and vibrational modes. The gas heating efficiency was found to be 10% and this parameter is shown to have a direct effect on the plasma jet velocity. These results indicate that the pulsed-plasma jet creates a sufficiently strong flow perturbation that holds great promise as a supersonic flow actuator.
TL;DR: The major conclusion is that accuracies of the node centered and the best cell-centered schemes are comparable at equivalent number of degrees of freedom.
Abstract: Cell-centered and node-centered approaches have been compared for unstructured finite-volume discretization of inviscid fluxes. The grids range from regular grids to irregular grids, including mixed-element grids and grids with random perturbations of nodes. Accuracy, complexity, and convergence rates of defect-correction iterations are studied for eight nominally second-order accurate schemes: two node-centered schemes with weighted and unweighted least-squares (LSQ) methods for gradient reconstruction and six cell-centered schemes two node-averaging with and without clipping and four schemes that employ different stencils for LSQ gradient reconstruction. The cell-centered nearest-neighbor (CC-NN) scheme has the lowest complexity; a version of the scheme that involves smart augmentation of the LSQ stencil (CC-SA) has only marginal complexity increase. All other schemes have larger complexity; complexity of node-centered (NC) schemes are somewhat lower than complexity of cell-centered node-averaging (CC-NA) and full-augmentation (CC-FA) schemes. On highly anisotropic grids typical of those encountered in grid adaptation, discretization errors of five of the six cell-centered schemes converge with second order on all tested grids; the CC-NA scheme with clipping degrades solution accuracy to first order. The NC schemes converge with second order on regular and/or triangular grids and with first order on perturbed quadrilaterals and mixed-element grids. All schemes may produce large relative errors in gradient reconstruction on grids with perturbed nodes. Defect-correction iterations for schemes employing weighted least-square gradient reconstruction diverge on perturbed stretched grids. Overall, the CC-NN and CC-SA schemes offer the best options of the lowest complexity and secondorder discretization errors. On anisotropic grids over a curved body typical of turbulent flow simulations, the discretization errors converge with second order and are small for the CC-NN, CC-SA, and CC-FA schemes on all grids and for NC schemes on triangular grids; the discretization errors of the CC-NA scheme without clipping do not converge on irregular grids. Accurate gradient reconstruction can be achieved by introducing a local approximate mapping; without approximate mapping, only the NC scheme with weighted LSQ method provides accurate gradients. Defect correction iterations for the CC-NA scheme without clipping diverge; for the NC scheme with weighted LSQ method, the iterations either diverge or converge very slowly. The best option in curved geometries is the CC-SA scheme that offers low complexity, second-order discretization errors, and fast convergence.
TL;DR: In this paper, an evaluation of computational models for flight dynamics simulations on low-speed aircraft with very-flexible high-aspect ratio wings is carried out for flight simulation.
Abstract: An evaluation of computational models is carried out for flight dynamics simulations on low-speed aircraft with very-flexible high-aspect ratio wings. Structural dynamic models include displacement-based, strain-based, and intrinsic (first-order) geometrically-nonlinear composite beams, while thin-strip and vortex lattice methods are considered for the unsteady aerodynamics. It is first shown that all different beam finite element models (previously derived in the literature from different assumptions) can be consistently obtained from a single set of equations. This approach has been used to expand existing strain-based models to include shear effects. Comparisons are made in terms of numerical efficiency and simplicity of integration in flexible aircraft flight dynamics studies. On the structural modeling, it was found that intrinsic solutions can be several times faster than conventional ones for aircraft-type geometries. For the aerodynamic modeling, thin-strip models based on indicial airfoil response are found to perform well in situations dominated by small amplitude dynamics around large quasi-static wing deflections, while large-amplitude wing dynamics require three-dimensional descriptions (e.g. vortex lattice).
TL;DR: In this article, an efficient gradient-based aerodynamic shape optimization is presented, which consists of several components, including a novel integrated geometry parameterization and mesh movement, a parallel Newton-Krylov flow solver, and an adjoint-based gradient evaluation.
Abstract: An efficient gradient-based algorithm for aerodynamic shape optimization is presented. The algorithm consists of several components, including a novel integrated geometry parameterization and mesh movement, a parallel Newton―Krylov flow solver, and an adjoint-based gradient evaluation. To integrate geometry parameterization and mesh movement, generalized B-spline volumes are used to parameterize both the surface and volume mesh. The volume mesh of B-spline control points mimics a coarse mesh; a linear elasticity mesh-movement algorithm is applied directly to this coarse mesh and the fine mesh is regenerated algebraically. Using this approach, mesh-movement time is reduced by two to three orders of magnitude relative to a node-based movement. The mesh-adjoint system also becomes smaller and is thus amenable to complex-step derivative approximations. When solving the flow-adjoint equations using restarted Krylov-subspace methods, a nested-subspace strategy is shown to be more robust than truncating the entire subspace. Optimization is accomplished using a sequential-quadratic-programming algorithm. The effectiveness of the complete algorithm is demonstrated using a lift-constrained induced-drag minimization that involves large changes in geometry.
TL;DR: In this article, a framework is developed to treat the most general approach that considers the largest possible design space, where the use of lamination parameters efficiently defines stiffness variation over a structural domain with the minimum number of variables.
Abstract: With the large-scale adoption of advanced fiber placement technology in industry, it has become possible to fully exploit the anisotropy of composite materials through the use of fiber steering. By steering the composite fibers in curvilinear paths, spatial variation of stiffness can be induced resulting in beneficial load and stiffness distribution patterns. One especially relevant area in which fiber steering has proved its effectiveness is in improving buckling loads of composite panels. Previous research used predefined forms of fiber angle variations and the coefficients of these analytic expressions were used as design variables. Alternatively, the local ply angles were used as design variables directly. In this paper, a framework is developed to treat the most general approach that considers the largest possible design space. The use of lamination parameters efficiently defines stiffness variation over a structural domain with the minimum number of variables. A conservative reciprocal approximation scheme is introduced. The inverse buckling factor is expanded linearly in terms of the in-plane stiffness and in terms of the inverse bending stiffness. The new approximation scheme is convex in lamination parameter space. Numerical results demonstrate improvements in excess of 100% in buckling loads of variable-stiffness panels compared to the optimum constant stiffness designs. Buckling load improvements are attributed primarily to in-plane load redistribution, which is confirmed both by the prebuckling stress distribution as well as by comparing the performance of designs optimized with variation of both in-plane and bending stiffness to those optimized with only bending stiffness variation. A tradeoff study between in-plane stiffness and buckling performance is also presented and shows the benefits of variable-stiffness design in enlarging the design possibilities of composite panels.
TL;DR: In this article, a hybrid approach to the development of a hybrid prediction methodology for jet noise is described, where a Gaussian function model for the two-point cross correlation of the fourth-order velocity fluctuations in the acoustic source is presented.
Abstract: A novel approach to the development of a hybrid prediction methodology for jet noise is described Modeling details and numerical techniques are optimized for each of the three components of the model Far-field propagation is modeled by solution of a system of adjoint linear Euler equations, capturing convective and refraction effects using a spatially developing jet mean flow provided by a Reynolds-averaged Navier―Stokes computational fluid dynamics solution Sound generation is modeled following Goldstein's acoustic analogy, including a Gaussian function model for the two-point cross correlation of the fourth-order velocity fluctuations in the acoustic source Parameters in this model describing turbulent length and time scales are assumed to be proportional to turbulence information also taken from the Reynolds-averaged Navier―Stokes computational fluid dynamics prediction The constants of proportionality are, however, not determined empirically, but extracted by comparison with turbulence length and time scales obtained from a large eddy simulation prediction The large eddy simulation results are shown to be in good agreement with experimental data for the fourth-order two-point cross-correlation functions The large eddy simulation solution is then used to determine the amplitude parameter and also to examine which components of the cross correlation are largest, enabling inclusion of all identified dominant terms in the Gaussian source model The acoustic source description in the present approach is therefore determined with no direct input from experimental data This model is applied to the prediction of sound to the experimental configuration of the European Union JEAN project, and gives encouraging agreement with experimental data across a wide spectral range and for both sideline and peak noise angles This paper also examines the accuracy of various commonly made simplifications, for example: a locally parallel mean flow approximation rather than consideration of the spatially evolving mean jet flow and scattering from the nozzle; the assumption of small radial variation in Green function over the turbulence correlation length; the application of the far-field approximation in the Green function; and the impact of isotropic assumptions made in previous acoustic source models
TL;DR: In this article, the effect of the interaction strength on the unsteady behavior of a planar shock wave impinging on a low Reynolds turbulent boundary layer is investigated by means of a variation in incident shock angle under otherwise constant flow conditions.
Abstract: The effect of the interaction strength on the unsteady behavior of a planar shock wave impinging on a low Reynolds turbulent boundary layer is investigated. This is achieved by means of a variation in incident shock angle under otherwise constant flow conditions. In addition, the effect of an order-of-magnitude variation in the Reynolds number is considered. This has been done for equivalent interaction strength, based on a similar probability of occurrence of instantaneous flow separations. The measurement technique employed is two-component planar particle image velocimetry. Common mechanisms for the large-scale reflected-shock unsteadiness are deduced by means of conditional statistics based on the separation bubble height. The results indicate that both upstream and downstream mechanisms are at work, the dominant mechanism depending on the interaction strength. No significant dependence on the Reynolds number was observed for interactions with a similar probability of instantaneous flow separations.
TL;DR: In this article, a reduced-order nonlinear unsteady aerodynamic modeling approach suitable for analyzing pitching/plunging airfoils subject to fixed or time-varying freestream Mach numbers is described.
Abstract: A reduced-order nonlinear unsteady aerodynamic modeling approach suitable for analyzing pitching/plunging airfoils subject to fixed or time-varying freestream Mach numbers is described. The reduced-order model uses kriging surrogates to account for flow nonlinearities and recurrence solutions to account for time-history effects associated with unsteadiness. The resulting surrogate-based recurrence framework generates time-domain predictionsofunsteadylift,moment,anddragthataccuratelyapproximate computational fluiddynamicssolutions, but at a fraction of the computational cost. Results corresponding to transonic conditions demonstrate that the surrogate-based recurrence framework can mimic computational fluid dynamics predictions of unsteady aerodynamic responses when flow nonlinearities are present. For an unsteady aerodynamic modeling problem considered in this study, an accurate reduced-order model was generated by the surrogate-based recurrence framework approach with significantly fewer computational fluid dynamics evaluations compared to results reported in the literature for a similar problem in which a proper-orthogonal-decomposition-based approach was applied. Furthermore, the results show that the surrogate-based approach can accurately model time-varying freestream Mach number effects and is therefore applicable to rotary-wing applications in addition to fixed-wing applications.
TL;DR: In this paper, large-eddy simulation is used to analyze supersonic flow, mixing, and combustion in a two-stage fuel injector strut, which has been carefully validated in a large number of other studies.
Abstract: In this study, large-eddy simulation is used to analyze supersonic flow, mixing, and combustion in a supersonic combustor equipped with a two-stage fuel injector strut. The present study focuses on mixing, ignition, and flame stabilization and the degree of detail required by the reaction mechanism in the large-eddy simulation model framework. An explicit large-eddy simulation model, using a mixed subgrid model and a partially stirred reactor turbulence-chemistry interaction model, is used in an unstructured finite volume setting. The model, and its components, has been carefully validated in a large number of other studies. To bestow further validation and to provide supplementary information about the physics of mixing and supersonic combustion, experimental data from the National Aerospace Laboratory of Japan's supersonic combustor, equipped with the two-stage strut injector and connected to ONERA's vitiation air heater, are employed. The large-eddy simulation predictions are compared with the experimental centerline wall pressure distribution and the planar laser-induced fluorescence imaging of hydroxide-ion radicals distributions in several cross sections of the combustor, showing excellent qualitative and quantitative agreements. The large-eddy simulation results are furthermore used to elucidate the complicated flow, mixing, and combustion physics imposed by the multi-injector two-stage injector strut. The importance of the combustion chemistry appears weaker than expected but with the one-step mechanism resulting in a too early ignition (caused by local shock wave heating) and a more stable flame, as compared with the more detailed two- and seven-step mechanisms.
TL;DR: In this article, a large-eddy simulation framework for turbulence modeling is used and real-gas effects are accounted for through a cubic equation of state and appropriate viscosity and conductivity coefficients.
Abstract: This paper presents the numerical computation of a turbulent jet of nitrogen into nitrogen under supercritical pressure. The large-eddy simulation framework for turbulence modeling is used and real-gas effects are accounted for through a cubic equation of state and appropriate viscosity and conductivity coefficients. The purpose of this paper is to evaluate how low-pressure large-eddy simulation equations coupled with real-gas thermodynamics and transport compare with experiments. Although this approach does not take into account the impact of high density gradients and nonlinear thermodynamics on turbulence modeling, the results show reasonable agreement with available experimental data and reveal the importance of numerics for such computations. The simulations indicate a limited influence of the density ratio and the thermodynamic conditions on the jets spreading rate and pseudosimilarity behavior.
TL;DR: In this article, a computational-fluid-dynamics-based computational methodology for fast on-demand aeroelastic predictions of the behavior of a full aircraft configuration at variable flight conditions is presented.
Abstract: This paper describes a computational-fluid-dynamics-based computational methodology for fast on-demand aeroelastic predictions of the behavior of a full aircraft configuration at variable flight conditions and demonstrates its feasibility. The methodology relies on the offline precomputation of a database of reduced-order bases and models associated with a discrete set of flight parameters, and its training for an interpolation method suitable for reduced-order information. The potential of this near-real-time computational methodology for assisting flutter flight testing is highlighted with the aeroelastic identification of an F-16 configuration in the subsonic, transonic, and supersonic regimes.
Abstract: The dynamics of unstart in a floor-mounted inlet-isolator model in a Mach 5 flow are investigated experimentally using particle image velocimetry and fast-response wall pressure measurements The inlet compression is obtained with a 6-deg ramp and the isolator is a rectangular straight duct that is 254 mm high by 508 mm wide by 2423 mm long Unstart is initiated from the scramjet mode (fully supersonic in the isolator) by deflecting a motorized flap at the downstream end of the isolator With the flap fully down, the particle image velocimetry data of the started flow capture the characteristics of the isolator boundary layers and the initial inlet reflected shock system During unstart, the unstart shock system propagates upstream through the inlet-isolator The particle image velocimetry data reveal a complex, three-dimensional flow structure that is strongly dependent on viscous mechanisms Particularly, the unstart shock system propagates upstream and induces significant boundary-layer separation Side-view particle image velocimetry data show that the locations of strongest separation during unstart correlate with the impingement locations of the initial inlet shock as it reflects down the isolator For example, in the middle of unstart, the unstart shock system is associated with massive separation of the ceiling boundary layer that begins where the first inlet shock reflection impinges on the ceiling The observation that separation increases at the inlet shock reflection impingement locations is likely due to the fact that the boundary layers in these locations are subject to larger adverse pressure gradients, thus making them more susceptible to separation During the unstart process, large regions of separated flow form near the floor and ceiling with reverse flow velocities up to about 04U ∞ These regions of separated, subsonic flow appear to extend to the isolator exit, creating a path by which the isolator exit boundary condition can be communicated upstream Plan-view particle image velocimetry data show the unstart process begins with separation of the isolator sidewall boundary layers Overall, the unstart flow structure is highly three-dimensional
TL;DR: In this paper, the authors evaluated the refinement of some classical theories, such as the Kirchhoff and Reissner-Mindlin theories, adding generalized displacement variables (up to fourth-order) to the Taylor-type expansion in the thickness plate direction.
Abstract: This work has evaluated the refinement of some classical theories, such as the Kirchhoff and Reissner-Mindlin theories, adding generalized displacement variables (up to fourth-order) to the Taylor-type expansion in the thickness plate direction. Isotropic, orthotropic, and laminated plates have been analyzed, varying the thickness ratio, orthotropic ratio, and stacking sequence of the layout. Higher-order theories have been implemented according to the compact scheme known as the Carrera unified formulation. The results have been restricted to simply-supported orthotropic plates subjected to harmonic distributions of transverse pressure for which closed-form solutions are available. For a given plate problem (isotropic, orthotropic, or laminated), the effectiveness of each employed generalized displacement variable has been established comparing the error obtained accounting for and removing the variable in the plate governing equations. A number of theories have therefore been constructed imposing a given error with respect to the available best results. Guidelines and recommendations that are focused on the proper selection of the displacement variables that have to be retained in refined plate theories are then furnished. It has been found that the terms that have to be used according to a given error vary from problem to problem, but they also vary when the variable that has to be evaluated (displacement, stress components) is changed. Diagrams (errors in terms of geometrical and orthotopic ratios) and graphical schemes have been built to establish the appropriate theories with respect to the data of the problem under consideration.
TL;DR: In this article, a theoretical analysis of the physical mechanisms driving mixing and combustion in supersonic airstreams is presented, where they are found to be different from those in the incompressible regime.
Abstract: Understanding the physics of supersonic combustion is the key to design a performing engine for scramjet-powered vehicles. Despite studies on supersonic combustion dating back to the 1950s, there are still numerous uncertainties and misunderstandings on this topic. The following questions need to be answered: How does compressibility affect mixing, flame anchoring, and combustion efficiency? How long must a combustor be to ensure complete mixing and combustion while avoiding prohibitive performance losses? How can reacting turbulent and compressible flows be modeled? Experimental results in the past have shown that supersonic combustion of hydrogen and air is feasible and takes place in a reasonable distance, which is a necessary requirement in actual hypersonic vehicles powered by supersonic combustion ramjets. These results are explained based on a theoretical analysis of the physical mechanisms driving mixing and combustion in supersonic airstreams, where they are found to be different from those in the incompressible regime. In particular, the classic Kolmogorov scaling is shown to be no longer strictly valid, and the flame regime is predicted to be significantly affected by compressibility and different from that of subsonic flames. This analysis is also supported by the results of the numerical simulations presented, showing that by generating sufficiently intense turbulence, a supersonic combustion flame is short and can indeed anchor within a small distance from fuel injectors, with the flame typically burning in the so-called flamelets-in-eddies regime.
TL;DR: In this paper, three properties for flux functions are proposed: shock stability/robustness, conservation of total enthalpy, and resolving boundary layer, and numerical experiments are performed for widely used or recently developed flux functions, and these fluxes are categorized into five major groups based on how they satisfy the three properties.
Abstract: In hypersonic flow computations, it is a key issue to predict surface heating accurately, though this is still challenging because there always are possibilities of resulting in anomalous solutions. In this paper, three properties for flux functions are proposed: 1) shock stability/robustness, 2) conservation of total enthalpy, and 3) resolving boundary layer. Then, numerical experiments are performed for widely used or recently developed flux functions, and these fluxes are categorized into five major groups based on how they satisfy the three properties. These tests reveal that no flux function investigated here possesses all the three properties. In particular, the first one is not satisfied by any flux functions, including flux-vector-splittings. Finally, contributions of those properties are compared inatwo-dimensional, viscous, hypersonicblunt-bodyproblem. Results showedthatthe firstandthe third properties are crucial, and the second one is preferred to predict hypersonic heating. A group of flux functions that best satisfies these properties is suggested, and they are recommended either to be used or designed for hypersonic heating computations.
TL;DR: In this paper, an immersed-boundary technique for compressible, turbulent flows was used to simulate the effects of micro vortex generators in controlling oblique-shock/turbulent boundary-layer interactions.
Abstract: This work presents an immersed-boundary technique for compressible, turbulent flows and applies the technique to simulate the effects of micro vortex generators in controlling oblique-shock/turbulent boundary-layer interactions. The Reynolds-averaged Navier-Stokes equations, closed using the Menter k-ω turbulence model, are solved in conjunction with the immersed-boundary technique. The approach is validated by comparing solutions obtained using the immersed-boundary technique with solutions obtained on a body-fitted mesh and with experimental laser Doppler anemometry data collected at Cambridge University for Mach 2.5 flow over single micro vortex generators. Simulations of an impinging oblique-shock boundary-layer interaction at Mach 2.5 with and without micro vortex-generator flow control are also performed, considering the development of the flow in the entire wind tunnel. Comparisons are made with experimental laser Doppler anemometry data and surface-pressure measurements from Cambridge University and an analysis of the flow structure is performed. The results show that three dimensional effects initiated by the interaction of the oblique shock with the sidewall boundary layers significantly influence the flow patterns in the actual experiment. The general features of the interactions with and without the micro vortex-generator array are predicted to good accord by the Reynolds-averaged Navier-Stokes/ immersed-boundary model.
TL;DR: In this article, a single dielectric barrier discharge plasma actuator for controlling turbulent boundary-layer separation from the deflected flap of a high-lift airfoil is investigated between Reynolds numbers of 240,000 (15 m/s) and 750,000(45 m/S).
Abstract: The efficacy of a single dielectric barrier discharge plasma actuator for controlling turbulent boundary-layer separation from the deflected flap of a high-lift airfoil is investigated between Reynolds numbers of 240,000 (15 m/s) and 750,000 (45 m/s). Momentum coefficients for the dielectric barrier discharge plasma actuator are approximately an order of magnitude lower than those usually employed for such studies, yet control authority is still realized through amplification of natural vortex shedding from the flap shoulder, which promotes momentum transfer between the freestream and separated region. This increases dynamic loading on the flap and further organizes turbulent fluctuations in the wake. The measured lift enhancement is primarily due to upstream effects from increased circulation around the entire model, rather than full reattachment to the deflected flap surface. Lift enhancement via instability amplification is found to be relatively insensitive to changes in angle of attack, provided that the separation location and underlying dynamics do not change. The modulation waveform used to excite low-frequency perturbations with a high-frequency plasma-carrier signal has a considerable effect on the actuator performance. Control authority decreases with increasing Reynolds number and flap deflection, highlighting the necessity for further improvement of plasma actuators for use in realistic takeoff and landing transport aircraft applications. These findings are compared to studies on a similar high-lift platform using piezoelectric-driven zero-net-mass flux actuation.
TL;DR: In this article, an aerodynamic shape optimization algorithm based on the Euler equations is proposed to minimize the induced drag of several nonplanar configurations using twist optimization to recover an elliptical lift distribution.
Abstract: The induced drag of several nonplanar configurations is minimized using an aerodynamic shape optimization algorithm based on the Euler equations. The algorithm is first validated using twist optimization to recover an elliptical lift distribution. Planform optimization reveals that an elliptical planform is not optimal when side-edge separation is present. Optimized winglet and box-wing geometries are found to have span efficiencies that agree well with lifting-line analysis, provided the bound constraints on the entire geometry are accounted for in the linear analyses. For the same spanwise and vertical bound constraints, a nonplanar split-tip geometry outperforms both the winglet and box-wing geometries, because it can more easily maximize the vertical extent at the tip. The performance of all the optimized geometries is verified using refined grids consisting of 88-152 million nodes.
TL;DR: In this paper, the relationship between flapping wing structure and the production of aerodynamic forces for micro air vehicle hovering flight by measuring full-field structural deformation and thrust generation was investigated.
Abstract: This experimental study investigates the relationship between flapping wing structure and the production of aerodynamic forces for micro air vehicle hovering flight by measuring full-field structural deformation and thrust generation. Results from four flexible micromembrane wings with different skeletal reinforcement demonstrate that wing compliance is crucial in thrust production: only certain modes of passive aeroelastic deformation allow the wing to effectively produce thrust. The experimental setup consists of a flapping mechanism with a single-degree-of- freedom rotary actuation up to 45 Hz at 70 deg stoke amplitude and with power measurement, a force and torque sensor that measures the lift and thrust, and a digital image correlation system that consists of four cameras capable of capturing the complete stroke kinematics and structural deformation. Several technical challenges related to the experimental testing of microflapping wings are resolved in this study: primarily, flapping wings less than 3 in. in length produce loads and deformations that are difficult to measure in an accurate and nonintrusive manner. Furthermore, the synchronization of the load measurement system, the vision-based wing deformation measurement system, and the flapping mechanism is demonstrated. Intensive data analyses are performed to extract useful information from the measurements in both air and vacuum.
TL;DR: In this paper, the accuracy of the correlation between a sphere in compressible multiphase simulations and a shock-tube was evaluated using the recent shock tube experiments of Jourdan et al.
Abstract: MPIRICALcorrelationsforthequasi-steadydragcoef!cientofa sphere in compressible "ow have been presented by severalauthors (e.g., Henderson  and Loth ). Such correlations areneeded in numerical simulations of compressible multiphase "owsinvolving spherical particles. In this Note, the accuracy of thecorrelationsofHendersonandLothareassessedusingthedatacollectedbyBaileyandStarr,andanimprovedcorrelationforthedrag coef!cient of a sphere in compressible "ow is developed. Theimproved correlation is validated for shock-particle interaction,using the recent shock-tube experiments of Jourdan et al. .
TL;DR: A novel surrogate modeling methodology designed specifically for propagating uncertainty from model inputs to model outputs and for performing a global sensitivity analysis is presented, which characterizes the contributions of uncertainties inmodel inputs to output variance while maintaining the quantitative rigor of the analysis.
Abstract: sampling) present an intractable computational burden. This paper presents a novel surrogate modeling methodology designed specifically for propagating uncertainty from model inputs to model outputs and for performing a global sensitivity analysis, which characterizes the contributions of uncertainties in model inputs to output variance, while maintaining the quantitative rigor of the analysis by providing confidence intervals on surrogate predictions. The approach is developed for a general class of models and is demonstrated on an aircraft emissions prediction model that is being developed and applied to support aviation environmental policy-making. The results demonstrate how the confidence intervals on surrogate predictions can be used to balance the tradeoff between computation time and uncertainty in the estimation of the statistical outputs of interest.
TL;DR: In this paper, an adjoint-based methodology for design optimization of unsteady turbulent flows on dynamic unstructured grids is described, and large-scale shape optimizations are demonstrated for turbulent flows over a tiltrotor geometry and a simulated aeroelastic motion of a fighter jet.
Abstract: An adjoint-based methodology for design optimization of unsteady turbulent flows on dynamic unstructured grids is described. The implementation relies on an existing unsteady three-dimensional unstructured grid solver capable of dynamic mesh simulations and discrete adjoint capabilities previously developed for steady flows. The discrete equations for the primal and adjoint systems are presented for the backward-difference family of time-integration schemes on both static and dynamic grids. The consistency of sensitivity derivatives is established via comparisons with complex-variable computations. The current work is believed to be the first verified implementation of an adjoint-based optimization methodology for the true time-dependent formulation of the Navier-Stokes equations in a practical computational code. Large-scale shape optimizations are demonstrated for turbulent flows over a tiltrotor geometry and a simulated aeroelastic motion of a fighter jet.
TL;DR: In this article, a large eddy simulation of the flow induced by a supersonic jet impinging on a flat plate in a stable regime was performed with an explicit third-order compressible solver using a unstructured mesh, a centered scheme, and the Smagorinsky model.
Abstract: This paper describes a numerical study based on large eddy simulation of the flow induced by a supersonic jet impinging on a flat plate in a stable regime. This flow involves very high velocities, shocks, and intense shear layers. Performing large eddy simulation on such flows remains a challenge because of the shock discontinuities. Here, large eddy simulation is performed with an explicit third-order compressible solver using a unstructured mesh, a centered scheme, and the Smagorinsky model. Three levels of mesh refinement (from 7 to 22 million cells) are compared in terms of instantaneous and averaged flowfields (shock and recirculation zone positions), averaged flow velocity and pressure fields, wall pressure, root mean square pressure fields, and spectral content using one and two-point analyses. The effects of numerical dissipation and turbulent viscosity are compared on the three grids and shown to be well controlled. The comparison of large eddy simulation with experimental data shows that the finest grid (a 22 million cell mesh) ensures grid-independent results not only for the mean and rms fields but also for higher statistics such as single and two-point correlation functions.
TL;DR: In this paper, the authors describe extensions and tests of characteristic methods for outlet boundary conditions in compressible solvers, based on the specification of incoming waves using one-and multidimensional approximations, extended to unstructured grids.
Abstract: This paper describes extensions and tests of characteristic methods for outlet boundary conditions in compressible solvers. Three methods based on the specification of incoming waves using one- and multidimensional approximations are extended to unstructured grids. They are first compared for weak to strong vortices propagating on low-to high-speed mean flows through outlet sections. A major issue is to determine the Mach number to be used in the specification of the transverse terms that must be taken into account in the incoming wave amplitude specifications. For the vortex computations, results show that the averaged Mach number leads to better results than its local value. The boundary conditions are then tested in a more complex case: the flow around a turbine blade. A reference solution using a long distance between the blade trailing edge and the outlet plane is first computed. For this solution, outlet boundary conditions have almost no effect on the flow around the blade. The distance between the trailing edge and the outlet plane is then shortened and the various characteristic treatments are compared, in which intense vortices cross the outlet plane. Results confirm the conclusions obtained on the simple vortex test case.
TL;DR: The trailing-edge noise model is more reliable for observer positions within ±30° from the fan-rotation plane, and should lead to a useful fast-running tool to be included in a blade-design process in an industrial context.
Abstract: This paper deals with the experimental validation of an analytical trailing-edge noise model dedicated to low-speed fans operating in free field. The model is intrinsically related to the aerodynamics of the blades and should lead to a useful fast-running tool to be included in a blade-design process in an industrial context. The investigations are made on a two-bladed low-speed axial fan without shroud, installed inside an anechoic room. The blades are instrumented with two sets of embedded small-size microphones (2.5 mm diam), and the wall-pressure signals are acquired via a slip ring mounted on the fan axis. The chord-based Reynolds number is about 200,000, and the tip Mach number about 0.07. The data base is completed by far-field measurements made with a single microphone on a moving support. The analytical model is based on a previously published extension of Amiet's trailing-edge noise theory. A blade is split into several strips in the spanwise direction, and the model is applied to each strip. For this the input data are interpolated from the measurements performed with the aforementioned sets of microphones. The trailing-edge noise model is more reliable for observer positions within ±30° from the fan-rotation plane.