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Unsteady turbulent flow simulations were performed based on the Reynolds-averaged Navier–Stokes (RANS) equations to investigate flow oscillation due to three-dimensional (3D) configuration of a Mach 1.5 supersonic open cavity flow with a length-to-depth ratio of 3. Two-dimensional (2D) and 3D unsteady simulation results were analyzed and compared with experimental data and Rossiter’s empirical prediction data. The three-dimensional cavity width-to-depth ratio (W/D) was 1, 3.8 and 7.6. Computational results indicated that pressure oscillation in the 2D flow was generated by a single-flow structure, whereas a multiple-flow structure generated multiple oscillation peaks in the 3D flow. The flow structure in the 3D cavity was investigated. For the 2D flow case, the cavity internal pressure wave was directly synchronized with the free shear layer. In the 3D flow case, an unstable spanwise flow due to the sidewall was observed. This spanwise fluctuation produced additional pressure oscillations coupled with the streamwise internal pressure wave. The numerical results indicate that the spanwise flow reduces the propagation speed of the internal pressure waves and the intensity of the corresponding pressure fluctuation.

In this paper, we present a computational aeroelastic analysis of flexible flapping wings in the vicinity of solid walls. The wall effects change the aerodynamic forces and moments of the wings, and thus the aeroelastic behavior. The numerical simulation is carried out using a fluid–structure interaction framework by coupling the computational fluid dynamics and computational structural dynamics. A preconditioned Navier–Stokes solver based on a finite volume method is used for the aerodynamic analysis. The structural analysis is performed using a nonlinear structural model based on a geometrically exact beam formulation. The method is validated using previous numerical and experimental results. The aeroelastic characteristics of the flexible wings with and without the walls are computed and compared.

The RANS equations are widely used to analyze complex flows over aircraft. The equations require a turbulence model for turbulent flow analyses. A suitable turbulence must be selected for accurate predictions of aircraft aerodynamic characteristics. In this study, numerical analyses of three-dimensional aircraft are performed to compare the results of various turbulence models for the prediction of aircraft aerodynamic characteristics. A 3-D RANS solver, MSAPv, is used for the aerodynamic analysis. The four turbulence models compared are the Sparlart–Allmaras (SA) model, Coakley’s $$q-\omega $$ model, Huang and Coakley’s $$k-\varepsilon $$ model, and Menter’s $$k-\omega $$ SST model. Four aircrafts are considered: an ARA-M100, DLR-F6 wing–body, DLR-F6 wing–body–nacelle–pylon from the second drag prediction workshop, and a high wing aircraft with nacelles. The CFD results are compared with experimental data and other published computational results. The details of separation patterns, shock positions, and $$C_{p}$$ distributions are discussed to find the characteristics of the turbulence models.

In this paper, a triangular planar element is developed for a geometrically nonlinear structural analysis, which includes the drilling degrees of freedom using a corotational framework. Based on the assumptions of a small degree of strain and large displacement, the corotational framework allows an accurate geometrically nonlinear structural analysis. The presently improved corotational framework accommodates in-plane rotational behavior (that is, the drilling degrees of freedom) by using the corotational framework corresponding to a solidlike planar element. It focuses on triangular planar elements that will be useful for three-dimensional analysis using a reduced number of degrees while targeting a structure with a complex geometry, such as a flapping wing. Regarding the present analysis, validation by solving both static and time-transient problems is conducted. The fluid–structure interaction framework is then developed by using the present structural analysis. During this validation procedure, the pr...

Realistic computations of flows induced by synthetic jets in multiple flow conditions are studied with an unsteady preconditioned Reynolds-averaged Navier–Stokes solver with deforming meshes. Deforming meshes suitable for structured grid systems are used to simulate the motion of diaphragms driven at a given frequency.With deforming meshes, a boundary-conditionmodel is not necessary for synthetic jet calculations. To obtain time-accurate solutions with a grid system under deformation, the geometric conservation law is employed. Numerical calculations of three cases of the Langley Research Center Workshop Computational Fluid Dynamics Validation of Synthetic Jets and Turbulent Separation Control workshop are carried out, and the computational results are compared with the experimental data. The results indicate the effectiveness of the current method in synthetic jet flow computations.

This paper presents computations of the dynamic derivatives of three dimensional flight vehicle configurations using CFD. The pitch dynamic derivatives are computed from the pitch sinusoidal motion, while the roll damping is computed based on steady state calculation using a non-inertial frame method. The Basic Finner and the SDM(Standard Dynamic Model) are chosen for the benchmark tests against other numerical and experimental results. For the flow calculations, a 3-D Euler solver that can be run both on the non-inertial frame and on the inertial frame is developed. A dual-time stepping method is applied for the unsteady time accurate simulations. A good agreement of pitch-roll dynamic derivatives with previously published numerical results and the experimental results is observed.

This paper investigates the convergence characteristics of the modified artificial compressibility method proposed by Turkel. In particular, a focus is mode on the convergence characteristics due to variation of the preconditioning factor (αu) and the artificial compressibility ( β) in conjunction with an upwind method. For the investigations, a code using the modified artificial compressibility is developed. The code solves the axisymmetric incompressible Reynolds averaged Navier-Stokes equations. The cell-centered finite volume method is used in conjunction with Roe’s approximate Riemann solver for the inviscid flux, and the central difference discretization is used for the viscous flux. Time marching is accomplished by the approximated factorizationalternate direction implicit method. In addition, Menter’s k-ω shear stress transport turbulence model is adopted for analysis of turbulent flows. Inviscid, laminar, and turbulent flows are solved to investigate the accuracy of solutions and convergence behavior in the modified artificial compressibility method. The possible reason for loss of robustness of the modified artificial compressibility method with αu >1.0 is given.

Off-line 6-DOF simulation program for store separation analysis has been developed. The developed program enables to predict a trajectory of a store from the database which was constructed by wind tunnel testing or CFD analysis. The flow angle method was applied to the program for predicting aerodynamic coefficients from the database and the ejector forces and constraints were enabled to incorporate the equations of motion for computing the trajectory. Using the program, the trajectories were calculated and the results are compared with the CTS results.

A numerical procedure for aerodynamic load analysis of long span bridges is presented. The preconditioned Reynolds averaged Navier-Stokes equations are adopted to compute flows over the bridges. To capture the turbulent characteristics of the flows, two equation turbulence models, Coakley’s q − ω model and Menter’s k − ω SST model, are used to compute the turbulent viscosity. A dual time stepping method in conjunction with the AF-ADI method is used to advance the solution in time. A loosely coupled method of the preconditioned RANS equations with the turbulence model equations is employed for fast computation without losing numerical stability. The numerical method for the aerodynamic load analysis is verified against well-known benchmark problems. Aerodynamic loads of two real bridges are computed with the method to demonstrate the usefulness of the method.

Euler codes or Navier-Stokes codes for compressible flows suffer severe degradation in convergence as Mach number approaches zero. The convergence problem arose from the wide disparity in characteristic speeds can be solved using preconditioning methods without large modifications. In this paper, a preconditioned RANS(Reynolds Averaged Navier-Stokes) solver is developed for analysis of low Mach number flows. In order to validate the method, computational examples are chosen and the results are compared with the experimental data and the existing computed results showing a good accuracy and convergence characteristics for steady inviscid, laminar and turbulent flows at low Mach number.

6-DOF simulation program is developed in order to increase the efficiency of the store separation analysis. This S/W is much faster than a method based on CFD(Computational Fluid Dynamics) technology, and allows the simulation of stores with fixed shape as well as with extensible wings, because it uses aerodynamic databases which are prepared beforehand. In this paper, aerodynamic databases of stores are obtained with MSAP(Multi-body Separation Analysis Program), and unsteady damping coefficients are modeled with Missile Datcom. These databases and the 6-DOF simulation program are used to predict the trajectory of an external store, while its wings are being deployed. The analysis results indicate that the safe separations of the store can be achieved not only with the wing fixed but with the wings being deployed.

A study on the test, design and fabrication of wind tunnel model for measurement of air load distribution on wing surfaces is presented. 447 pressure taps are installed normal to the wing surfaces, and measured by PSI-8400 system using total 8 ESPs modules installed in the model. The test was performed at 50 m/sec constant speed in the low speed wind tunnel of Agency for Defense Development. Tests were carried out to determine effects of angle of attack, angle of sideslip and flap and stores for the load distribution of wing. The test results in this paper can be applied to the design optimization of structure and validation of computational fluid dynamics.

A new fully automated store-trajectory simulation code has been developed. The method couples a highly robust computational fluid dynamics method and a cut-paste algorithm for Chimera domain decomposition to enhance the efficiency of store-trajectory simulation. The time-step size limitation caused by grid movement is relaxed with the interpolation of solutions in the previous time step. A treatment for orphan cells is also devised to improve the robustness of the method

A new systemized procedure for chimera-domain decomposition is presented. This procedure consists of a new cut-paste algorithm for optimal mesh interface and a two-step search method for donor cell identification. It is fully automated and requires minimal user input. The cut-paste algorithm is based on the advancing front technique in which the fronts are iteratively determined from the initial fronts, which are a collection of fringe points obtained from conventional Chimera hole cutting. The final fronts are determined iteratively in such a way that the overlapping region is minimized. With this method, interpolation points are located away from solid walls where flow gradient is high. We also can reduce the error that may arise from interpolation in stiff gradient regions. Two- and three-dimensional examples are chosen to demonstrate the effectiveness of this new procedure

In a flapping wing micro air vehicle (MAV), inspired by an organism of either insects or birds, flexibility of the wing structure induces a crucial effect upon the vehicle performance. Thus, in an analysis upon the flapping wing MAV, coupling between aerodynamics and structural dynamics considering the wing flexibility will be a critical component. This paper presents an accurate computational approach to simulate a flapping wing by coupling between CFD and CSD. Non-linear structural analysis based on the geometrically exact beam formulation was used. Such non-linear beam analysis was coupled with preconditioned Navier-Stokes solutions. For a grid deformation in the aerodynamic analysis, the mesh shearing methodology was used. A coupling between the structural and aerodynamic analyses was conducted by adopting the implicit coupling approach. After that, an aeroelastic analysis was performed and the results are compared with the experimental results. However, the flapping wing configuration is not slender in reality and their vein section geometry is complex generally. Thus, to consider those features, the finite element analysis, beam and shell, based on a co-rotational (CR) theory was developed in parallel. Currently, the CR beam analysis with a warping DOF was developed and validated by comparing it with NASTRAN in static condition.

This paper describes a computational study of the prediction of stability derivatives for the standard dynamic model(SDM) using three-dimensional Euler and Reynolds-averaged Navier-Stokes(RANS) computation fluid dynamics(CFD) solvers. Directional and longitudinal stability derivatives are computed in pitch and yaw harmonic oscillations, and a dual time stepping method is applied to evaluate time-accurate solutions. Lateral derivatives are evaluated by a steady state method in a non-inertial reference frame. Source terms due to the transformation of the governing equations to the non-inertial coordinate are added for the steady state method. Simulations are performed at various flight conditions in terms of Mach numbers, angles of attack, and oscillation amplitudes. The simulations show that the nonlinear characteristics of the stability derivatives appear at high angles of attack as well as in the transonic region.

Numerical simulations of 3D aircraft configurations are performed in order to understand the effects that turbulence models have on the aerodynamic characteristics of an aircraft. An in-house CFD code that solves 3D RANS equations and 2-equation turbulence model equations is used for the study. The code applies Roe’s approximated Riemann solver and an AF-ADI scheme. Furthermore van Leer’s MUSCL extrapolation with van Albada’s limiter is adopted. Various versions of Menter’s k-omega SST turbulence models as well as Coakley’s q-omega model are incorporated into the CFD code. Menter’s k-omega SST models include the standard model, the 2003 model, the model incorporating the vorticity source term, and the model containing controlled decay. Turbulent flows over a wing are simulated in order to validate the turbulence models contained in the CFD code. The results from these simulations are then compared to computational results of the 3rd AIAA CFD Drag Prediction Workshop. Moreover, numerical simulations of the DLR-F6 wing-body and wing-body-nacelle-pylon configurations are conducted and compared to computational results of the 2nd AIAA CFD Drag Prediction Workshop. Especially, the aerodynamic characteristics as well as flow features with respect to the turbulence models are scrutinized. The results obtained from each simulation incorporating Menter’s k-omega SST turbulence model variations are compared with one another.Copyright © 2011 by KSME

A complete three-dimensional mathematical model has been developed governing the steady, laminar flow of an incompressible fluid subjected to a magnetic field and including internal heating due to the Joule effect, heat transfer due to conduction, and thermally induced buoyancy forces. The thermally induced buoyancy was accounted for via the Boussinesq approximation. The entire system of eight partial differential equations was solved by integrating intermittently a system of five fluid flow equations and a system of three magnetic field equations and transferring the information through source-like terms. An explicit Runge-Kutta time-stepping algorithm and a finite difference scheme with artificial compressibility were used in the general non-orthogonal curvilinear boundary-conforming co-ordinate system. Comparison of computational results and known analytical solutions in two and three dimensions demonstrates high accuracy and smooth monotone convergence of the iterative algorithm. Results of test cases with thermally induced buoyancy demonstrate the stabilizing effect of the magnetic field on the recirculating flows.

A new method for enhancing the convergence rates of iterative schemes for the numerical integration of systems of partial differential equations has been developed. It is termed the distributed minimal residual (DMR) method. The DMR method has been applied to incompressible Navier-Stokes equations with heat transfer. Alt numerical lest cases were obtained using explicit four-stage Runge-Kutla or Euler implicit lime integration. The DMR method was found to reduce compulation lime by 207ndash;60%, depending on the test case. The formulation for the DMR method is general in nature and can be applied to explicit and implicit iterative algorithms for arbitrary systems of partial differential equations.

A new automated Chimera method has been developed for the prediction of store trajectory. The method uses a highly robust CFD method and our cut-paste algorithm for the Chimera domain decomposition to enhance the efficiency of store trajectory simulation. The time-step size limitation due to the grid movement is relaxed with the interpolation of solutions in the previous time step. A treatment for orphan cells is also devised to improve the robustness of method. Computational results show that the current method is capable of simulating the store separation without user interruption and that a timestep size of 0.005 set, which is more than three times larger than ever reported, can be used for the Eglin Wing/Pylon/Store separation problem. Introduction Since the Chimera grid technique’ was introduced to CFD(Computational Fluid Dynamics) community, it is gaining its popularity not only because the grid generation over complex geometries is relatively simple compared to the multi-block grid generation technique but because the Chimera method is the most successful method in handling relative movement of multiple bodies. The Chimera grid technique is naturally suitable for moving body problems. One of moving body problems frequently encountered in aerodynamic applications is the trajectory prediction of a store separating from mother aircraft. There have been several reports that successfully predict the trajectory of the well-known Eglin Wing/Pylon/Store problem234 .5 Recently, the simulations of ripple release of multiple stores have been reported6.’ However, the trajectory prediction with CFD is not a routine engineering application. The ‘Senior Researcher tPrincipa1 Researcher, Member AIAA z Doctoral candidate 5 Associate Professor, Senior Member AIAA. copyright @ lgqg by the American Institute of Aerondcs and Astronautics, IX. AU tights reserd. major stumbling block against the goal comes from the fact that most Chimera methods rely on interactive approaches for the Chimera hole cutting8.g The computations, furthermore, are often hindered by the lack of the efficiency and the robustness of the CFD tools. In this paper, we present our efforts toward the goal. Our efforts include implementing an automated Chimera hole cutting method, cut-paste algorithm,‘O combined with two step donor cell search. Cut-paste algorithm is the advancing front technique for the Chimera hole cutting. The solid walls or non-penetrable boundaries are used as the initial hole-cutting boundary so that the user input is minimized. With this method, the Chimera hole cutting can be performed automatically for unsteady applications. Also, two step donor cell search ensures fast and reliable search. As stated by Lijewski,4 the computation often fails due to the presence of orphan cells. A proper handling of orphan cells without loss of accuracy is a “must” for full unsteady simulations. Furthermore, the time-step size is limited by the presence of the points which are initially hole points and become normal points as the grid moves without being updated. We devise methods to handle such problems to enhance the robustness of our Chimera method. In addition, a highly robust CFD method” is used for aerodynamic forces and moment calculations. The object of this paper is to present our attempt to enhance the efficiency and the robustness of the computational method and the Chimera method so that the store prediction method can be used as an engineering tool in near future. The trajectory computations of Eglin Wing/Pylon/Store and ripple release from TER (Tripple Ejector Rack) are presented showing the current method satisfies the objective of our study. . Chimera Hole Cutting The existing Chimera methods need user specified hole cutting surfaces for the Chimera hole cutting. In PEGASUS,8 for example, the user must specify the hole cutting-surfaces as well as the hole grid, while in