Browsing by Author "Qin, N."

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    The Aerodynamics of High Speed Aerial Weapons
    (Cranfield University, 1999-09) Prince, Simon A.; Qin, N.; Birch, T.
    The focus of this work is the investigation of the complex compressible flow phenomena associated with high speed aerial weapons. A three dimen- sional multiblock finite volume flow solver was developed with the aim of studying the aerodynamics of missile configurations and their component structures. The first component of the study involved the aerodynamic investigation of the isolated components used in the design of conventional missile config- urations. The computational study of nine ogive-cylinder body experimental test cases is presented together with a new interpretation of the complex vortical flow including the windward appearance of a "vortex shock wave". In addition, a simple modification to improve the accuracy of the Baldwin- Lomax/Degani-Schi fl`' turbulence model is put forward, and the phenomenon of "phantom vorticity" in Euler solutions and its alleviation are described. Inclined Delta Wings in supersonic flow were computed in order to study the aerodynamics of wings alone, and in particular the vortex-shock interactions which occur on their leeward surfaces. The second component of the study was the computational and experimen- tal investigation of a generic cruciform missile configuration. The compli- cated interactions between shock waves and boundary/shear layers that are seen to occur around and in the wake of the cruciform wing arrangement were studied and described. The third component of the research involved an assessment of the pre- diction technologies used in the design of modern weapons. In particular the role of Computational Fluid Dynamics in the process of design.
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    A discrete Navier-Stokes adjoint method for aerodynamic optimisation of BlendedWing-Body configurations
    (Cranfield University, 2002-12) Le Moigne, Alan; Qin, N.
    An aerodynamic shape optimisation capability based on a discrete adjoint solver for Navier- Stokes flows is developed and applied to a Blended Wing-Body future transport aircraft. The optimisation is gradient-based and employs either directly a Sequential Quadratic Programming optimiser or a variable-fidelity optimisation method that combines low- and high-fidelity models. The shape deformations are parameterised using a B´ezier-Bernstein formulation and the structured grid is automatically deformed to represent the design changes. The flow solver at the heart of this optimisation chain is a Reynolds averaged Navier- Stokes code for multiblock structured grids. It uses Osher’s approximate Riemann solver for accurate shock and boundary layer capturing, an implicit temporal discretisation and the algebraic turbulence model of Baldwin-Lomax. The discrete Navier-Stokes adjoint solver based on this CFD code shares the same implicit formulation but has to calculate accurately the flow Jacobian. This implies a linearisation of the Baldwin-Lomax model. The accuracy of the resulting adjoint solver is verified through comparison with finitedifference. The aerodynamic shape optimisation chain is applied to an aerofoil drag minimisation problem. This serves as a test case to try and reduce computing time by simplifying the fidelity of the model. The simplifications investigated include changing the convergence level of the adjoint solver, reducing the grid size and modifying the physical model of the adjoint solver independently or in the entire optimisation process. A feasible optimiser and the use of a penalty function are also tested. The variable-fidelity method proves to be the most ef- ficient formulation so it is employed for the three-dimensional optimisations in addition to parallelisation of the flow and adjoint solvers with OpenMP. A three-dimensional Navier- Stokes optimisation of the ONERA M6 wing is presented. After describing the concept of Blended Wing-Body and the studies carried out on this aircraft, several aerodynamic optimisations are performed on this geometry with the capability developed in this thesis.
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    Efficient upwind algorithms for solution of the Euler and Navier-stokes equations
    (Cranfield University, 1995-10) McNeil, C. Y.; Roe, P. L.; Qin, N.
    An efficient three-dimensionasl tructured solver for the Euler and Navier-Stokese quations is developed based on a finite volume upwind algorithm using Roe fluxes. Multigrid and optimal smoothing multi-stage time stepping accelerate convergence. The accuracy of the new solver is demonstrated for inviscid flows in the range 0.675 :5M :5 25. A comparative grid convergence study for transonic turbulent flow about a wing is conducted with the present solver and a scalar dissipation central difference industrial design solver. The upwind solver demonstrates faster grid convergence than the central scheme, producing more consistent estimates of lift, drag and boundary layer parameters. In transonic viscous computations, the upwind scheme with convergence acceleration is over 20 times more efficient than without it. The ability of the upwind solver to compute viscous flows of comparable accuracy to scalar dissipation central schemes on grids of one-quarter the density make it a more accurate, cost effective alternative. In addition, an original convergencea cceleration method termed shock acceleration is proposed. The method is designed to reduce the errors caused by the shock wave singularity M -+ 1, based on a localized treatment of discontinuities. Acceleration models are formulated for an inhomogeneous PDE in one variable. Results for the Roe and Engquist-Osher schemes demonstrate an order of magnitude improvement in the rate of convergence. One of the acceleration models is extended to the quasi one-dimensiona Euler equations for duct flow. Results for this case d monstrate a marked increase in convergence with negligible loss in accuracy when the acceleration procedure is applied after the shock has settled in its final cell. Typically, the method saves up to 60% in computational expense. Significantly, the performance gain is entirely at the expense of the error modes associated with discrete shock structure. In view of the success achieved, further development of the method is proposed.
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    Implicit multi-block Euler/Navier-Stokes simulations for hovering helicopter rotor
    (Cranfield University, 2003-02) Zhong, B.; Qin, N.
    A three dimensional implicit multiblock Navier-Stokes solver for hovering rotor vortical flow simulations has been developed. The governing equations used are cast in an attached blade rotating frame. Two formulations of the governing equations using the relative or absolute velocity as variables respectively are employed and investigated. The Osher's approximate Riemann solver is used for the convective fluxes evaluation. A modified MUSCL scheme is employed for improving the accuracy of the discretisation for the in viscid fluxes. A Block Incomplete Lower and Upper Decomposition (BILU) is adopted for solving the linear system resulted from the use of an implicit scheme. Special treatment for the terms, including extra flux terms and source terms, arising from the non- inertial reference system are implemented. A multiblock technique is used to obtain the exibility for quality grid generation. The suitability of different grid topologies for vortex wake capturing is demonstrated. Numerical tests show that significant improvement in computational efficiency is achieved by utilising the BILU implicit scheme in both fixed wing and hovering rotor calculations. Numerical simulations also demonstrate Navier-Stokes solutions give more accurate results than that from Euler solutions, especially in transonic tip speed cases. Computed results including surface pressure distributions and tip vortex trajectories are compared with the experimental data, which shows that the developed solver and the numerical scheme can simulate hovering rotor flows with good accuracy.
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    Monotone integrated large eddy simulation of supersonic boundary layer flows
    (2001-02) Chong, Yon Han; Qin, N.
    For simulations of supersonic flows shock-capturing schemes have to be used. A shock-capturing scheme produces more dissipation than a central difference scheme. In fact, the numerical dissipation produced by shock-capturing schemes is problematic when performing Large Eddy Simulation of supersonic flows with shock-waves. Another train of thought is to turn the numerical dissipation to our advantage. If the numerical dissipation of a numerical method can mimic the dissipation of the subgrid-scale(SGS) eddies, not only is SGS modelling unnecessary, but the numerical dissipation will be a positive contribution to the calculation. This approach is called MILES. As a reference case, a zero-pressure-gradient, flat-plate boundary-layer flow was chosen as there are analytical, experimental, DNS and LES results available. The freestream conditions are a Mach number of 2.25 and a Reynolds number of 1.613 x 104/in or 6.007 x 103 based on the displacement thickness. The central difference scheme, Osher’s scheme and Roe’s scheme are tested for suitability in MILES. The central difference scheme is found to be numerically^ too non-dissipative without SGS modelling. Osher’s scheme is too dissipative so that it hinders the development of turbulence. Roe’s scheme without use of a limiter seems to have the right amount of numerical dissipation to mimic a SGS model. Two popular slope limiters were also tested, but both affected turbulence development when no shockwave was present.
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    A numerical investigation of the flows in and around clustered module plug nozzles
    (Cranfield University, 2001-08) Perigo, D. A.; Qin, N.
    This thesis aims to make advances in the accurate simulation of the ows in and around clustered module plug nozzles. The resulting simulations presented in this thesis are, as far as can be ascertained from available data, the most detailed to date in Europe. A comparison is made with results from other sources for clarication of this point. In the process of producing these solutions, two ow solvers have been developed. NSAXIMB is a general 2D multi-block ow solver,developed by the author, for the axisymmetric, Reynolds averaged Navier-Stokes equations. It was developed to allow simulation of axisymmetic plug nozzle congurations and the investigation of the effects of turbulence modelling on such ows. MERLIN is a general 3D, implicit, multi-block ow solver again for the RANS equations. MERLIN was developed by the Centre for Computational uid Dynamic at Craneld. Signicant input from this work has included a large portion of the structure of the mean ow solver and the extension of the advanced two equation turbulence modelling, incorporated in NSAX- IMB, to three dimensions. Of the turbulence models investigated the zonal models of Menter prove to be most effective in reproducing experimental results. These models out perform a more advanced non-linear eddy viscosity formulation, based on the work of Abid. In an effort to improve solution accuracy, grid adaptation software, based on node redistribution techniques has been developed for use in conjunction with the 3D ow solver. This work is demonstrated in conjunction with a basic test case before application to the clustered module plug nozzle conguration. Results for the complex block topology adopted in the 3D test case are shown to cause the adaptation process to fail. Further, it is shown that such a process may not be generalised for arbitrary topologies.
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    Numerical study of the unsteady aerodynamics of helicopter rotor aerofoils
    (Cranfield University, 1999-03) Shaw, Scott; Qin, N.
    A two-dimensional model of the aerodynamics of rotor blades in forward flight is proposed in which the motion of the blade is represented by periodical variations of the freestrearn velocity and incidence. A novel implicit methodology for the solution of the compressible Reynolds averaged Navier-Stokes equations and a twoequation model of turbulence is developed. The spatial discretisation is based upon Osher's approximate Riernann solver, while time integration is performed using a Newton-Krylov method. The method is employed to calculate the steady transonic aerodynamics of two supercritical aerofoils and the unsteady aerodynamics of pitching aerofoils. Comparison with experiment and independent calculations for these test cases is satisfactory. Further calculations are performed for the self-excited periodic flow around a biconvex aerofoil. Comparison of quasi-steady and unsteady calculations suggests that the flow instability responsible for the self-excited flow is due to the presence of a shock induced separation bubble in the corresponding steady flow. Finally the method is used to predict the aerodynamics of aerofoils performing inplane and combined inplane-pitching motions. Results show that quasi-steady aerodynamic models are unsuitable at conditions representative of high-speed forward flight. For shock free flows, the unsteady effects of freestrearn oscillations can be represented by a simple phase lag. For transonic flows the influence of unsteadiness on shock wave dynamics is shown to be complex. Calculations for indicial motion show that the unsteady behaviour of the flow is related to the finite time taken by disturbance waves to travel to the shock wave from the leading and trailing edges of the aerofoil.
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    Response surface aerodynamic optimisation for blended wing body aircraft
    (Cranfield University, 2005-02-08) Vavalle, Armando; Qin, N.
    This study is concerned with a methodology for the aerodynamic analysis and preliminary design of a novel configuration for high subsonic civil transport, based on the flying wing concept, known a Blended Wing Body (BWB). A response surface based optimisation method is developed, enabling the designer to monitor the effect of shape modification on the controllability of the aircraft in both longitudinal and lateral/directional motion and on the Wing structural weight, while maximising the aerodynamic efficiency. The design aspects considered included high- speed aerodynamics, flight static-stability and trim characteristics. The response surface Scheme employs a space filling design of experiment technique to build least square fitting quadratic polynomials, used in place of the original computational modules in a gradient based search. A optimisation test indicated that the present method is more effective in leading the design near to the global optimum as opposed to a conventional gradient method with direct search, despite that the constructed approximation may not represent accurately the actual surface. With this system, multiple constrained optimisation problems are successfully solved in the favourable case of smooth objective/constraint function. Where these functions may exhibit high non-linear trends, an iterative response surface method refining both approximation and bounds of the design space is proposed. The capabilities of such a technique are shown for transonic aerofoil optimisation problems, demonstrating that the proposed method is more efficient and more effective than some other state-of- the-art methods. As a result of these studies, the aerodynamic efficiency of a large capacity BWB configuration has been considerably improved by re-designing the external shape to generate a spanwise loading intermediate between triangular and elliptic. The longitudinal static stability analysis revealed that the aircraft is stable except at low- weights with zero-payload. The lateral/directional analyses showed that the aircraft is stable in roll, but unstable in yaw. Despite that the winglets are found to stabilise the aircraft, it is directionally unstable without additional vertical stabilisers. I