Weng, Shangkun. / Thesis (M.Phil.)--Chinese University of Hong Kong, 2009. / Includes bibliographical references (p. 73-78). / Abstracts in English and Chinese. / Abstract --- p.i / Acknowledgement --- p.iii / Chapter 1 --- Introduction --- p.1 / Chapter 2 --- Existence and Uniqueness results of transonic shock solution to full Euler system in a large variable nozzle --- p.11 / Chapter 2.1 --- The mathematical description of the transonic shock problem and main results --- p.11 / Chapter 2.2 --- The reformulation on problem (2.1.1) with (2.1.5)-(2.1.9) --- p.18 / Chapter 2.3 --- An Iteration Scheme --- p.30 / Chapter 2.4 --- A priori estimates and proofs of Theorem 2.2.1 and Theorem 2.1.1 --- p.39 / Chapter 3 --- A monotonic theorem on the shock position with respect to the exit pressure --- p.50 / Chapter 4 --- Discussions and Future work --- p.64 / Chapter 5 --- Appendix --- p.66 / Chapter 5.1 --- Appendix A: Background solution --- p.66 / Chapter 5.2 --- Appendix B: An outline of the proof of Theorem 2.1.2 --- p.67

Two extensions to the proper orthogonal decomposition (POD) technique are considered for steady transonic aerodynamic applications. The first is to couple the POD approach with a cubic spline interpolation procedure in order to develop fast, low-order models that accurately capture the variation in parameters, such as the angle of attack or inflow Mach number. The second extension is a POD technique for the reconstruction of incomplete or inaccurate aerodynamic data. First, missing flow field data is constructed with an existing POD basis constructed from complete aerodynamic data. Second, a technique is used to develop a complete snapshots from an incomplete set of aerodynamic snapshots. / Singapore-MIT Alliance (SMA)

Simultaneous Perturbation Stochastic Approximation method has attracted considerable application in many different areas such as statistical parameter estimation, feedback control, simulation-based optimization, signal & image processing, and experimental design. In this paper, its performance as a viable optimization tool is demonstrated by applying it first to a simple wing geometry design problem for which the objective function is described by an empirical formula from aircraft design practice and then it is used in a transonic fan blade design problem in which the objective function is not represented by any explicit function but is estimated at each design iteration by a computational fluid dynamics algorithm for solving the Navier-Stokes equations / Singapore-MIT Alliance (SMA)

This dissertation provides numerical studies to improve bladed disk assembly design for preventing blade high cycle fatigue failures. The analyses are divided into two major subjects. For the first subject presented in Chapter 2, the mechanisms of transonic fan flutter for tuned systems are studied to improve the shortcoming of traditional method for modern fans using a 3D time-linearized Navier-Stokes solver. Steady and unsteady flow parameters including local work on the blade surfaces are investigated. It was found that global local work monotonically became more unstable on the pressure side due to the flow rollback effect. The local work on the suction side significantly varied due to nodal diameter and flow rollback effect. Thus, the total local work for the least stable mode is dominant by the suction side. Local work on the pressure side appears to be affected by the shock on the suction side. For the second subject presented in Chapter 3, sensitivity studies are conducted on flutter and forced response due to frequency mistuning and aerodynamic asymmetry using the single family of modes approach by assuming manufacturing tolerance. The unsteady aerodynamic forces are computed using CFD methods assuming aerodynamic symmetry. The aerodynamic asymmetry is applied by perturbing the influence coefficient matrix. These aerodynamic perturbations influence both stiffness and damping while traditional frequency mistuning analysis only perturbs the stiffness. Flutter results from random aerodynamic perturbations of all blades showed that manufacturing variations that effect blade unsteady aerodynamics may cause a stable, perfectly symmetric engine to flutter. For forced response, maximum blade amplitudes are significantly influenced by the aerodynamic perturbation of the imaginary part (damping) of unsteady aerodynamic modal forces. This is contrary to blade frequency mistuning where the stiffness perturbation dominates. / Dissertation

The field of Computational Fluid Dynamics (CFD) is in a continual state of advancement due to new numerical techniques, optimization of existing codes, and the increase in memory and processing speeds of computers. In this thesis, the solution technique for a pre-existing Navier-Stokes flow solver is adapted from an explicit Runge Kutta method to a Lower-Upper Symmetric Gauss-Seidel (LU-SGS) implicit time integration method. Explicit time integration methods were originally used in CFD codes because these methods require less memory. Information needed to advance the flow in time is localized to each grid point. These explicit methods are, however, restricted by time step sizes due to stability criteria. In contrast, implicit methods are unaffected by a large time step sizes but are restricted by memory requirements due to the complexities of unstructured grids. The implementation of LU-SGS performs grid re-ordering for unstructured meshes because of the coupling of grid points in the integration method's solution. The explicit and implicit flow solvers were tested for inviscid flows in incompressible, compressible, and transoinc flow regimes. The results found by comparing the implicit and explicit algorithms revealed a significant speed up in convergence to steady state by the LU-SGS method in terms of iteration number and CPU time per iteration.

<p>Vehicles with internal combustion (IC) engines fueled by hydrocarbon compounds have been used for more than 100 years for ground transportation. During the years and in particular in the last decade, the environmental aspects of IC engines have become a major political and research topic. Following this interest, the emissions of pollutants such as NO_x, CO₂ and unburned hydrocarbons (UHC) from IC engines have been reduced considerably. Yet, there is still a clear need and possibility to improve engine efficiency while further reducing emissions of pollutants. The maximum efficiency of IC engines used in passenger cars is no more than $40\%$ and considerably less than that under part load conditions. One way to improve engine efficiency is to utilize the energy of the exhaust gases to turbocharge the engine. While turbocharging is by no means a new concept, its design and integration into the gas exchange system has been of low priority in the power train design process. One expects that the rapidly increasing interest in efficient passenger car engines would mean that the use of turbo technology will become more widespread. The flow in the IC-engine intake manifold determines the flow in the cylinder prior and during the combustion. Similarly, the flow in the exhaust manifold determines the flow into the turbine, and thereby the efficiency of the turbocharging system. In order to reduce NO_x emissions, exhaust gas recirculation (EGR) is used. As this process transport exhaust gases into the cylinder, its efficiency is dependent on the gas exchange system in general. The losses in the gas exchange system are also an issue related to engine efficiency. These aspects have been addressed up to now rather superficially. One has been interested in global aspects (e.g. pressure drop, turbine efficiency) under steady state conditions.In this thesis, we focus on the flow in the exhaust port and close to the valve. Since the flow in the port can be transonic, we study first the numerical modeling of such a flow in a more simple geometry, namely a bump placed in a wind tunnel. Large-Eddy Simulations of internal transonic flow have been carried out. The results show that transonic flow in general is very sensitive to small disturbances in the boundary conditions. Flow in the wind tunnel case is always highly unsteady in the transonic flow regime with self excited shock oscillations and associated with that also unsteady boundary-layer separation. To investigate sensitivity to periodic disturbances the outlet pressure in the wind tunnel case  was varied periodically at rather low amplitude. These low amplitude oscillations caused hysteretic behavior in the mean shock position and appearance of shocks of widely different patterns. The study of a model exhaust port shows that at realistic pressure ratios, the flow is transonic in the exhaust port. Furthermore, two pairs of vortex structures are created downstream of the valve plate by the wake behind the valve stem and by inertial forces and the pressure gradient in the port. These structures dissipate rather quickly. The impact of these structures and the choking effect caused by the shock on realistic IC engine performance remains to be studied in the future.</p> / CICERO

Thesis (Ph. D.)--Virginia Polytechnic Institute and State University, 1997. / Vita. Includes bibliographical references (p. 115-121).

Shock control bumps (SCBs) are considered promising flow control devices for transonic commercial aircraft. By generating a λ-shock structure, 2D SCBs offer large drag savings, but perform poorly when that structure breaks down off-design. Milder-performing 3D devices produce weak vortices, that may offer some boundary layer control, and SCBs also affect buffet via direct impact on shock motions and separation. To date however, design studies have largely ignored complications from the swept, spanwise-varying flows, so this thesis tackles the question of whether SCB arrays can offer useful benefits to the performance of transonic commercial aircraft. Using a numerical infinite-wing model, a simple rotation adaptation is shown to redress deficient on-design drag performance of 3D SCBs under swept flows. With the correct rotation (dependent on height, planform and spacing) bumps follow performance-design trends similar to those in unswept flow. With this knowledge, an array design method is developed to tailor 2D and 3D devices to local flow conditions on an aircraft model, aiming to maximise on-design drag performance. Careful infinite-wing setup means the influence of rotation and array height on performance is replicated on the aircraft. Predicted array designs achieve 74-87% of their estimated local drag savings. However, with wave drag being a smaller percentage of the total, the influence of arrays on lift is more significant and makes the optimal designs shorter than predicted. Strategies for improving off-design drag performance are then evaluated. Stagger, an alternating chordwise translation applied to 3D arrays, broadens operating range and lowers drag penalties by better accommodating off-design shock movements, but offers a less favourable trade-off against on-design drag than simply reducing the array height. However, a 2D array can always outperform a 3D on drag objectives. Lastly, buffet performance is inferred using steady indicators based on trailing edge pressure and shock location. These disagree regarding the impact on buffet onset, unresolvably due to a lack of validation data, but agree that arrays could alleviate flow development post-onset. Optimal array designs depend on prioritised objectives: considering buffet severity and on-design drag, tall 2D (or 3D) arrays; for buffet and minimum off-design drag penalties (similar to the motivation behind vortex generator application), 3D arrays of varying height and stagger. A simple flight fuel consumption model utilising the computed drag data shows that many arrays are neutral or offer small savings (up to 0.3%) across a range of mission profiles. While likely too small to merit application for solely drag purposes, this implies buffet benefits without cost to efficiency. Unsteady tests and proper assessment of buffet onset are needed to confirm this.

abstract: Modern aircraft are expected to fly faster and more efficiently than their predecessors. To improve aerodynamic efficiency, designers must carefully consider and handle shock wave formation. Presently, many designers utilize computationally heavy optimization methods to design wings. While these methods may work, they do not provide insight. This thesis aims to better understand fundamental methods that govern wing design. In order to further understand the flow in the transonic regime, this work revisits the Transonic Similarity Rule. This rule postulates an equivalent incompressible geometry to any high speed geometry in flight and postulates a “stretching” analogy. This thesis utilizes panel methods and Computational Fluid Dynamics (CFD) to show that the “stretching” analogy is incorrect, but instead the flow is transformed by a nonlinear “scaling” of the flow velocity. This work also presents data to show the discrepancies between many famous authors in deriving the accurate Critical Pressure Coefficient (Cp*) equation for both swept and unswept wing sections. The final work of the thesis aims to identify the correct predictive methods for the Critical Pressure Coefficient. / Dissertation/Thesis / Masters Thesis Aerospace Engineering 2016

The purpose of this paper is to develop a general method for using Finite Elements in the Steepest Descent Method. The main application is to a partial differential equation for a Transonic Flow Problem. It is also applied to Burger's equation, Laplace's equation and the minimal surface equation. The entire method is tested by computer runs which give satisfactory results. The validity of certain of the procedures used are proved theoretically. The way that the writer handles finite elements is quite different from traditional finite element methods. The variational principle is not needed. The theory is based upon the calculation of a matrix representation of operators in the gradient of a certain functional. Systematic use is made of local interpolation functions.