In this thesis, optimal placement of adhesively bonded piezoelectric patches on laminated plates and the determination of geometry of the bonding area to maximize actuation effect are studied. A new finite element model, in which each layer is considered to be a separate plate, is developed. The adhesive layer is modeled as a distributed spring system. In this way, relative transverse normal and shear motion of the layers are allowed. Effect of delamination on the adhesive layer stresses is also studied and investigated through several case studies. Optimization problems, having single and multiple objectives, are investigated for both actuator placement and selective bonding examples. In these case studies, 2D and 3D Pareto fronts are also obtained. &amp / #8216 / Hide and Seek Simulated Annealing&amp / #8217 / method is adapted for discrete problems and used as the optimization technique for single-objective problems. Finally, Multiple Cooling Multi Objective Simulated Annealing optimization algorithm is adapted and used in multi-objective optimization case studies.

The purpose of this thesis is to implement Finite Volume Weighted Essentially Non-Oscillatory (FV-WENO) scheme to solution of one and two-dimensional discretised Euler equations with different flux algorithms. The effects of the different fluxes on the solution have been tested and discussed. Beside, the effect of the grid on these fluxes has been investigated. Weighted Essentially Non-Oscillatory (WENO) schemes are high order accurate schemes designed for problems with piecewise smooth solutions that involve discontinuities. WENO schemes have been successfully used in applications, especially for problems containing both shocks and complicated smooth solution structures. Fluxes are used as building blocks in FV-WENO scheme. The efficiency of the scheme is dependent on the fluxes used in scheme The applications tested in this thesis are the 1-D Shock Tube Problem, Double Mach Reflection, Supersonic Channel Flow, and supersonic Staggered Wedge Cascade. The numerical solutions for 1-D Shock Tube Problem and the supersonic channel flow are compared with the analytical solutions. The results for the Double Mach Reflection and the supersonic staggered cascade are compared with results from literature.

UAVs employed for low altitude jobs are more liable to collide with the urban structures on their way to the goal point. In this thesis, the problem of reactive obstacle avoidance is addressed by an innovative partial integrated guidance and control (PIGC) approach using the Six-DOF model of real UAV unlike the kinematic models used in the existing literatures. The guidance algorithm is designed which uses the collision cone approach to predict any possible collision with the obstacle and computes an alternate aiming direction for the vehicle. The aiming direction of the vehicle is the line of sight line tangent to the safety ball surrounding the obstacle. The point where the tangent touches the safety ball is the aiming point. Once the aiming point is known, the obstacle is avoided by directing the vehicle (on the principles of pursuit guidance) along the tangent to the safety ball. First, the guidance algorithm is applied successfully to the point mass model of UAV to verify the proposed collision avoidance concept. Next, PIGC approach is proposed for reactive obstacle avoidance of UAVs. The reactive nature of the avoidance problem within the available time window demands simultaneous reaction from the guidance and control loop structures of the system i.e, in the IGC framework (executes in single loop). However, such quick maneuvers cause the faster dynamics of the system to go unstable due to inherent separation between the faster and slower dynamics. On the contrary, in the conventional design (executes in three loops), the settling time of the response of different loops will not be able to match with the stringent time-to-go window for obstacle avoidance. This causes delay in tracking in all the loops which will affect the system performance adversely and hence UAV will fail to avoid the obstacle. However, in the PIGC framework, it overcomes the disadvantage of both the IGC design and the conventional design, by introducing one more loop compared to the IGC approach and reducing a loop compared to the conventional approach, hence named as Partial IGC. Nonlinear dynamic inversion technique based PIGC approach utilizes the faster and slower dynamics of the full nonlinear Six-DOF model of UAV and executes the avoidance maneuver in two loops. In the outer loop, the vehicle guidance strategy attempts to reorient the velocity vector of the vehicle along the aiming point within a fraction of the available time-to-go. The orientation of the velocity vector is achieved by enforcing the angular correction in the horizontal and vertical flight path angles and enforcing turn coordination. The outer loop generates the body angular rates which are tracked as the commanded signal in the inner loop. The enforcement of the desired body rates generates the necessary control surface deflections required to stir the UAV. Control surface deflections are realized by the vehicle through the first order actuator dynamics. A controller for the first order actuator model is also proposed in order to reduce the actuator delay. Every loop of the PIGC technique uses nonlinear dynamic inversion technique which has critical issues like sensitiveness to the modeling inaccuracies of the plant model. To make it robust against the parameter inaccuracies of the system, it is reinforced with the neuro-adaptive design in the inner loop of the PIGC design. In the NA design, weight update rule based on Lyapunov’s theory provides online training of the weights. To enhance fast and stable training of the weights, preflight maneuvers are proposed. Preflight maneuvers provide stabilized pre-trained weights which prevents any misapprehensions in the obstacle avoidance scenario. Simulation studies have been carried out with the point mass model and with the Six-DOF model of the real fixed wing UAV in the PIGC framework to test the performance of the nonlinear reactive guidance scheme. Various simulations have been executed with different number and size of the obstacles. NA augmented PIGC design is validated with different levels of uncertainties in the plant model. A comparative study in NA augmented PIGC design was performed between the pre-trained weights and zero weights as used for weight initialization in online training. Various comparative study shows that the NA augmented PIGC design is quite effective in avoiding collisions in different scenarios. Since the NDI technique involved in the PIGC design gives a closed loop solution and does not operate with iterative steps, therefore the reactive obstacle avoidance is achieved in a computationally efficient manner.

In aircraft with jet propulsion engine intakes at supersonic speed, strong pressure waves referred to as shockwaves occur, which may interact with any present boundary layers along the intake surface. The adverse pressure gradients associated with Shock Wave-Boundary Layer Interaction (SWBLI) may cause boundary layer flow separation, which can result in disturbances of the flow that can be harmful to the device or decrease engine performance. A common way in dealing with the adverse effects of SWBLI is through removal of low-momentum flow in the boundary layer, a process referred to as boundary layer bleed. In the process of bleed, the boundary layer is subjected to a pressure difference promoting flow out of the system, through a porous surface, and into a plenum. The porous surfaces used in the mass flow removal process contain orifices in small scales. Thus, in Computational Fluid Dynamics (CFD), creating a mesh resolving both the orifice scales and the bulk flow is a cumbersome task, and the computational cost becomes substantially increased. To this end, several boundary conditions which effectively model the large-scale effects of bleed have been developed. The aim of this study is to implement the Boundary Condition (BC) developed by John W. Slater into M-EDGE, the in-house compressible CFD-solver of SAAB Aeronautics. The bleed boundary condition model is based on a dimensionless surface sonic flow coefficient, which is derived from empirical wind-tunnel measurements of the bleed mass flow. In previous work, the Slater bleed BC has been shown to correlate well with wind-tunnel data. Furthermore, a simple transpiration law formulated by Reynald Bur was implemented in order get familiarized with the M-EDGE Fortran source code. However, this model is expected to yield unsatisfactory results, as reported in previous work in the field. The implemented Slater BC is tested on two different two-dimensional flow cases; flow over a flat plate without SWBLI, and flow including a shock wave generator creating SWBLI. In the flat plate case, simulations were run at Mach numbers 1.27, 1.58, 1.98 and 2.46 over a 6.85cm plate of 19% porosity. In the SWBLI-case, only flow at Mach 2.46 was considered, with a 9.53cm plate of 21% porosity. The Reynolds number range used throughout was 1.39−1.76·10^7/m. Simulations were run at different bleed rates over a structured grid using steady state RANS with the Spalart-Allmaras one-equation turbulence model. The boundary condition performance was assessed by its ability to recreate the sonic flow coefficients on which it is based. Further, the shape of downstream pitot pressure profiles are compared with experimental data. Results from the studies indicate that the implementation manages to recreate the data for the sonic flow coefficient with small error margins. The implementation can be used to simulate porous plates of different dimensions and porosities, even though the bleed model is based on empirical mass flow measurements of a 6.85cmplate of 19% porosity. The implementation is able to predict global bleed effects in the flow field, as indicated by comparisons of pitot pressure profiles at various downstream reference planes, despite differences in reference boundary layer intake profiles. Further, the overall flow field was compared visually with other simulation-studies, indicating that the global Mach distributions of the geometries were in accordance with the reference data. However, pitot profiles should be further studied with better matched intake boundary layer profiles. The main limitation of the boundary condition is that it relies on the wind-tunnel data of the surface sonic flow coefficients for specific bleed plate configurations. Furthermore, the implementation has only been verified to work within specific Mach number range of the underlying empirical measurements. In future work, the generality of the model could be increased by extending the data to other configurations and Mach numbers by conducting new experiments or using other published empirical data.

The present study considers the numerical simulation of the different flow characteristics involved in the conjugate heat transfer analysis of an internally cooled gas turbine blade. Conjugate simulations require full coupling of convective heat transfer in fluid regions to the heat diffusion in solid regions. Therefore, accurate prediction of heat transfer quantities on both external and internal surfaces has the uppermost importance and highly connected with the performance of the employed turbulence models. The complex flow on both surfaces of the internally cooled turbine blades is caused from the boundary layer laminar-to-turbulence transition, shock wave interaction with boundary layer, high streamline curvature and sequential flow separation. In order to discover the performances of different turbulence models on these flow types, analyses have been conducted on five different experimental studies each concerned with different flow and heat transfer characteristics. Each experimental study has been examined with four different turbulence models available in the commercial software (ANSYS FLUENT13.0) to decide most suitable RANS-based turbulence model. The Realizable k-&epsilon / model, Shear Stress Transport k-&omega / model, Reynolds Stress Model and V2-f model, which became increasingly popular during the last few years, have been used at the numerical simulations. According to conducted analyses, despite a few unreasonable predictions, in the majority of the numerical simulations, V2-f model outperforms other first-order turbulence models (Realizable k-&epsilon / and Shear Stress Transport k-&omega / ) in terms of accuracy and Reynolds Stress Model in terms of convergence.

Aerodynamic parameter estimation from closed loop data has been developed as another research area since control and stability augmentation systems have been mandatory for aircrafts. This thesis focuses on aerodynamic parameter estimation of an air to ground missile from closed loop data using separate surface excitations. A design procedure is proposed for designing separate surface excitations. The effect of excitations signals to the system is also analyzed by examining autopilot disturbance rejection performance. Aerodynamic parameters are estimated using two different estimation techniques which are ordinary least squares and complex linear regression. The results are compared with each other and with the aerodynamic database. An application of the studied techniques to a real system is also given to validate that they are directly applicable to real life.

Aerodynamic parameter estimation from closed loop data has been developed as another research area since control and stability augmentation systems have been mandatory for aircrafts. This thesis focuses on aerodynamic parameter estimation of an air to ground missile from closed loop data using separate surface excitations. A design procedure is proposed for designing separate surface excitations. The effect of excitations signals to the system is also analyzed by examining autopilot disturbance rejection performance. Aerodynamic parameters are estimated using two different estimation techniques which are ordinary least squares and complex linear regression. The results are compared with each other and with the aerodynamic database. An application of the studied techniques to a real system is also given to validate that they are directly applicable to real life.

This thesis presents the calibration processes of the pitot-static system, which is a part of the air data system of a military transport aircraft through flight tests. Tower fly-by method is used for air data system calibration. Altitude error caused by the position of the static port on the aircraft is determined by analyzing the data collected during four sorties with different weight, flap and landing gear configurations. The same data has been used to determine the airspeed measurement error. It has been shown that both the altitude and airspeed errors are within the allowable limits specified by FAR 25. Same method is also used for trailing cone calibration that is used for high altitude test flights for RVSM certification.

In this thesis, development of a generic transient aero-thermal gas turbine model is presented. A simulation code, gtSIM is developed based on an algorithm which is composed of a set of differential equations and a set of non-linear algebraic equations representing each gas turbine engine component. These equations are the governing equations which represents the aero-thermodynamic process of the each engine component and they are solved according to a specific solving sequence which is defined in the simulation code algorithm. At each time step, ordinary differential equations are integrated by a first-order Euler scheme and a set of algebraic equations are solved by forward substitution. The numerical solution process lasts until the end of pre-defined simulation time. The objective of the work is to simulate the critical transient scenarios for different types of gas turbine engines at off-design conditions. Different critical transient scenarios are simulated for two di&reg / erent types of gas turbine engine. As a first simulation, a sample critical transient scenario is simulated for a small turbojet engine. As a second simulation, a hot gas ingestion scenario is simulated for a turbo shaft engine. A simple proportional control algorithm is also incorporated into the simulation code, which acts as a simple speed governor in turboshaft simulations. For both cases, the responses of relevant engine parameters are plotted and results are presented. Simulation results show that the code has the potential to correctly capture the transient response of a gas turbine engine under different operating conditions. The code can also be used for developing engine control algorithms as well as health monitoring systems and it can be integrated to various flight vehicle dynamic simulation codes.

In this study, anisotropic shells of revolution subject to symmetric and unsymmetrical static loads are analysed. In derivation of governing equations to be used in the solution, first order transverse shear effects are included in the formulation. The governing equations can be listed as kinematic equations, constitutive equations, and equations of motion. The equations of motion are derived from Hamilton&rsquo / s principle, the constitutive equations are developed under the assumptions of the classical lamination theory and the kinematic equations are based on the Reissner-Naghdi linear shell theory. In the solution method, these governing equations are manipulated and written as a set called fundamental set of equations. In order to handle anisotropy and first order transverse shear deformations, the fundamental set of equations is transformed into 20 first order ordinary differential equations using finite exponential Fourier decomposition and then solved with multisegment method of integration, after reduction of the two-point boundary value problem to a series of initial value problems. The results are compared with finite element analysis results for a number of sample cases and good agreement is found. Case studies are performed for circular cylindrical shell and truncated spherical shell geometries. While reviewing the results, effects of temperature and pressure loads, both constant and variable throughout the shell, are discussed. Some drawbacks of the first order transverse shear deformation theory are exhibited.