This study investigates the effect of chordwisely modulated tip injection on the flow and turbulence characteristics of the tip vortex through experimental measurements downstream of a rectangular half-wing that has an aspect ratio of three. This injection technique involves spanwise jets at the tip that are issued from a series of holes along the chord line normal to the freestream flow direction. The injection mass flow rate from each hole is individually controlled using computer driven solenoid valves and therefore the flow injection geometrical pattern at the tip can be adjusted to any desired waveform shape, with any proper injection velocity. The measurements are performed in a blow-down wind tunnel using Constant Temperature Anemometry and Kiel probe traverses as well as Stereoscopic Particle Image Velocimetry. Current data show consistent trends with v previously observed effects of steady uniform tip injection such as the upward and outward motion of the vortex as well as increased levels of turbulence within the vortex core. The vortex size gets bigger with injection and the total pressure levels get reduced significantly near the vortex core. The injection pattern also seems to affect the size of the wing wake as well as the wake entrainment characteristics of the tip vortex. Depending on the injection waveform pattern and injection momentum coefficient the helicoidal shape of the tip vortex also seems to get affected.

This thesis presents a finite element method based comparative study of linear and geometrically non-linear progressive failure analysis of thin walled composite aerospace structures, which are typically subjected to combined in-plane and out-of-plane loadings. Different ply and constituent based failure criteria and material property degradation schemes have been included in a PCL code to be executed in MSC Nastran. As case studies, progressive failure analyses of sample composite laminates with cut-outs under combined loading are executed to study the effect of geometric non-linearity on the first ply failure and progression of failure. Ply and constituent based failure criteria and different material property degradation schemes are also compared in terms of predicting the first ply failure and failure progression. For mode independent failure criteria, a method is proposed for the determination of separate material property degradation factors for fiber and matrix failures which are assumed to occur simultaneously. The results of the present study show that under combined out-of-plane and in-plane loading, linear analysis can significantly underestimate or overestimate the failure progression compared to geometrically non-linear analysis even at low levels of out-of-plane loading.

This thesis presents structural analysis of a cockpit of a jet trainer type aircraft and the correlation studies performed by using ground pressurisation test results. For this purpose, first the response of the complex integrated fuselage structure is investigated under the complex type cabin pressure load. Then, cockpit part of the fuselage structure is modelled using commercial finite element software MSC/PATRAN&reg / and MSC/NASTRAN&reg / . The finite element model (FEM) of the cockpit structure is improved by the examination of the ground pressurisation test data and is finalised after achieving a good correlation between the finite element analysis (FEA) and the test results. This final form of the FEM of the cockpit structure serving as a benchmark is proved to be reliable for any future modifications.

This study presents the results of an experimental study that investigates the effects of uniform/waveform tip injection along the camberline on the total pressure loss characteristics downstream of a row of Low Pressure Turbine (LPT) blades. The experiments are performed in a low speed cascade facility. This injection technique involves spanwise jets at the tip that are issued from a series of holes along the camber line normal to the freestream flow direction. The injection mass flow rate from each hole is individually controlled using computer driven solenoid valves and therefore the flow injection geometrical pattern at the tip can be adjusted to any desired waveform shape, and can be uniform as well as waveform along the camber. Measurements involve Kiel probe traverses for different injection scenarios 0.5 axial chords downstream of the blades. Results show that, instead of performing uniform mass injection along the camberline, by selecting an appropriate waveform injection pattern one can reduce the total loss levels of the blade, including the tip leakage loss as well as the wake losses.

Presented in this thesis is the development of a high-fidelity aerothermal model for GE T700 turboshaft engine. The model was constructed using thermodynamic relations governing change of flow properties across engine components, and by applying real component maps for the compressor and turbines as well as empirical relations for specific heats. Included in the model were bleed flows, turbine cooling and heat sink effects. Transient dynamics were modeled using inter-component volumes method in which mass imbalance between two engine components was used to calculate the inter-component pressure. This method allowed fast, high-accuracy and iteration-free calculation of engine states. Developed simulation model was successfully validated against previously published simulation results, and was applied in the simulation of inlet distortion and engine deterioration. Former included simulation of steady state and transient hot gas ingestion as well as transient decrease in the inlet total pressure. Engine deterioration simulations were performed for four different cases of component deterioration with parameters defining engine degradation taken from the literature. Real time capability of the model was achieved by applying time scaling of plenum volumes which allowed for larger simulation time steps at very little cost of numerical accuracy. Finally, T700 model was used to develop a generic model by replacing empirical relations for specific heats with temperature and FAR dependent curve fits, and scaling T700 turbine maps. Developed generic aerothermal model was applied to simulate steady state performance of the Lycoming T53 turboshaft engine.

This thesis focuses on the development of guidance algorithms in order to avoid a prescribed obstacle primarily using the Collision Cone Method (CCM). The Collision Cone Method is a geometric approach to obstacle avoidance, which forms an avoidance zone around the obstacles for the vehicle to pass the obstacle around this zone. The method is reactive as it helps to avoid the pop-up obstacles as well as the known obstacles and local as it passes the obstacles and continue to the prescribed trajectory. The algorithm is first developed for a 2D (planar) avoidance in 3D environment and then extended for 3D scenarios. The algorithm is formed for the optimized CCM as well. The avoidance zone radius and velocity are optimized using constraint optimization, Lagrange multipliers with Karush-Kuhn-Tucker conditions and direct experimentation.

This thesis presents the results of an experimental study that investigates the effects of steady vortex generator jets (VGJs) integrated to a low pressure turbine blade to control the laminar separation bubble occurring on the suction surface of the blade at low Reynolds numbers. The injection technique involves jets issued from the holes located near the suction peak of the test blade which is in the middle of a five-blade low speed linear cascade facility. Three injection cases are tested with different blowing ratio values ranging from low to high. Surface pressure and particle image velocimetry (PIV) measurements are performed. The results show that steady VGJ is effective in eliminating the laminar separation bubble. Also it is observed that to have fully developed attached boundary layer, blowing ratio should be chosen accordingly since a very thin separation zone still exists at low blowing ratios.

In this thesis, store separation from two different configurations are solved using computational methods. Two different commercially available CFD codes / CFD-FASTRAN, an implicit Euler solver, and an unsteady panel method solver USAERO, coupled with integral boundary layer solution procedure are used for the present computations. The computational trajectory results are validated against the available experimental data of a generic wing-pylon-store configuration at Mach 0.95. Major trends of the separation are captured. Same configuration is used for the comparison of unsteady panel method with Euler solution at Mach 0.3 and 0.6. Major trends are similar to each other while some differences in lateral and longitudinal displacements are observed. Trajectories of a fueltank separated from an F-16 fighter aircraft wing and full aircraft configurations are found at Mach 0.3 using only the unsteady panel code. The results indicate that the effect of fuselage is to decrease the drag and to increase the side forces acting on the separating fueltank from the aircraft. It is also observed that the yawing and rolling directions of the separating fueltank are reversed when it is separated from the full aircraft configuration when compared to the separation from the wing alone configuration.

The objective of this thesis is to accomplish the three dimensional parallel thin-layer Navier-Stokes solutions for low aspect ratio rectangular flat wings in compressible flow. Two block parallel Navier Stokes solutions of an aspect ratio 1.0 flat plate with sharp edges are obtained at different Mach numbers and angles of attack. Reynolds numbers are of the order of 1.0E5-3.0E5. Two different grid configurations, the coarse and the fine grids, are applied in order to speed up convergence. In coarse grid configuration, 92820 total grid points are used in two blocks, whereas it is 700,000 in fine grid. The flow field is dominated by the vortices and the separated flows. Baldwin Lomax turbulence model is used over the flat plate surface. For the regions dominated by the strong side edge vortices, turbulence model is modified using a polar coordinate system whose origin is at the minimum pressure point of the vortex. In addition, an algebraic wake-type turbulence model is used for the wake region behind the wing. The initial flow variables at the fine grid points are obtained by the interpolation based on the coarse grid results previously obtained for 40000 iterations. Iterations are continued with the fine grid about 20000-40000 more steps. Pressures of the top surface are predicted well with the exception of leading edge region, which may be due to unsuitable turbulence model and/or grid quality. The predictions of the side edge vortices and the size of the leading edge bubble are in good agreement with the experiment.

In this thesis a 3-D Euler flow solver is coupled with a finite element program in order to solve static aeroelastic problems involving aircraft wings. A loosely coupled solution approach based on an iterative solution procedure is used to solve the coupled field problem. Because of the deformation of the underlying surface over which the flow is solved, Computational Fluid Dynamics mesh has to move at each computational aeroelastic iteration in order to comform to the new shape of the aerodynamic surface. As a part of this work, a procedure is developed in order to move fluid grid points, which views the whole computational domain as an isotropic elastic medium and solves it using finite element method. A matching discrete interface is defined / displacement and pressure data exchange is accomplished at this interface. AGARD Wing 445.6 and an elastic supercritical wing is modelled and solved with the developed computational aeroelastic procedure and the obtained results are compared with numerical and wind tunnel data.