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ÖgeA highorder finitevolume solver for supersonic flows(Lisansüstü Eğitim Enstitüsü, 2022) Spinelli, Gregoria Gerardo ; Çelik, Bayram ; 721738 ; Uçak ve Uzay MühendisliğiNowadays, Computational Fluid Dynamics (CFD) is a powerful tool in engineering used in various industries such as automotive, aerospace and nuclear power. More than ever the growing computational power of modern computer systems allows for realistic modelization of physics. Most of the opensource codes, however, offer a secondorder approximation of the physical model in both space and time. The goal of this thesis is to extend this order of approximation to what is defined as highorder discretization in both space and time by developing a twodimensional finitevolume solver. This is especially challenging when modeling supersonic flows, which shall be addressed in this study. To tackle this task, we employed the numerical methods described in the following. Curvilinear meshes are utilized since an accurate representation of the domain and its boundaries, i.e. the object under investigation, are required. Highorder approximation in space is guaranteed by a Central Essentially NonOscillatory (CENO) scheme, which combines a piecewise linear reconstruction and a kexact reconstruction in region with and without discontinuities, respectively. The usage of multistep methods such as RungeKutta methods allow for a highorder approximation in time. The algorithm to evaluate convective fluxes is based on the family of Advection Upstream Splitting (AUSM) schemes, which use an upwind reconstruction. A central stencil is used to evaluate viscous fluxes instead. When using highorder schemes, discontinuities induce numerical problems, such as oscillations in the solution. To avoid the oscillations, the CENO scheme reverts to a piecewise linear reconstruction in regions with discontinuities. However, this introduces a loss of accuracy. The CENO algorithm is capable of confining this loss of accuracy to the cells closest to the discontinuity. In order to reduce this accuracy loss Adaptive Mesh Refinement (AMR) is used. This algorithm refines the mesh near the discontinuity, confining the loss of accuracy to a smaller portion of the domain. In this study, a combination of the CENO scheme and the AUSM schemes is used to model several problems in different compressibility regimes, with a focus on supersonic flows. The scope of this thesis is to analyze the capabilities and the limitations of the proposed combination. In comparison to traditional implementations, which can be found in literature, our implementation does not impose a limit on the refinement ratio of neighboring cells while utilizing AMR. Due to the high computational expenses of a highorder scheme in conjunction with AMR, our solver benefits from a shared memory parallelization. Another advantage over traditional implementations is that our solver requires one layer of ghost cells less for the transfer of information between adjacent blocks. The validation of the solver is performed in different steps. We assess the order of accuracy of the CENO scheme by interpolating a smooth function, in this case the spherical cosine function. Then we validate the algorithm to compute the inviscid fluxes by modeling a Sod shock tube. Finally, the Boundary Conditions (BCs) for the inviscid solver and its order of accuracy are validated by modeling a convected vortex in a supersonic uniform flow. The curvilinear mesh is validated by modeling the flow around a NACA0012 airfoil. The computation of the viscous fluxes is validated by modeling a viscous boundary layer developing on a flat plate. The BCs for viscous flows and the curvilinear implementation are validated by modeling the flow around a cylinder and a NACA0012 airfoil. The AUSM schemes are tested for shock robustness by modeling an inviscid hypersonic cylinder at a Mach number of 20 and a viscous hypersonic cylinder at a Mach number of 8.03. Then, we validate our AMR implementation by modeling a twodimensional Riemann problem. All the validation results agree well with either numerical or experimental results available in literature. The performance of the code, in terms of computational time required by the different orders of approximation and the parallel efficiency, is assessed. For the former a supersonic vortex convection served as an example, while the latter used a twodimensional Riemann problem. We obtained a linear speedup until 12 cores. The highest speedup value obtained is 20 with 32 cores. Furthermore, the solver is used to model three different supersonic applications: the interaction between a vortex and a normal shock, the double Mach reflection and the diffraction of a shock on a wedge. The first application resembles a strong interaction between a vortex and a steady shock wave for two different vortex strengths. In both cases our results perfectly match the ones obtained by a Weighted Essentially NonOscillatory (WENO) scheme documented in literature. Both schemes are approximating the solution with the same order of accuracy in both, time and space. The second application, the double Mach reflection, is a challenging problem for highorder solvers because the shock and its reflections interact strongly. For this application, all AUSMschemes under investigation fail to obtain a stable result. The main form of instability encountered is the Carbuncle phenomenon. Our implementation overcomes this problem by combining the AUSM+M scheme with the formulation of the speed of sound of the AUSM+up scheme. This combination is capable of modeling this problem without instabilities. Our results are in agreement with those obtained with a WENO scheme. Both, the reference solutions and our results, use the same order of accuracy in both, time and space. Finally, the third example is the diffraction of a shock past a delta wedge. In this configuration the shock is diffracted and forms three different main structures: two triple points, a vortex at the trailing edge of the wedge and a reflected shock traveling upwards. Our results agree well with both, numerical and experimental results available in literature. Here, a formation of a vortexlet is observed along the vortex slipline. This vorticity generation under inviscid flow condition is studied and we conclude that the stretching of vorticity due to compressibility is the reason. The same formation is observed when the angle of attack of the wedge is increased in the range of 030. In general, the AUSM+up2 scheme performed best in terms of accuracy for all problems tested here. However, for configurations, in which the Carbuncle phenomenon may appear, the combination of the AUSM+M scheme and the computation of the speed of sound formula of the AUSM+up scheme is preferable for stability reasons. During our computations, we observe a small undershooting right behind shocks on curved boundaries. This is imputable to the curvilinear approximation of the boundaries, which is only secondorder accurate. Our experience shows that the smoothness indicator formula in its original version, fails to label uniform flow regions as smooth. We solve the issue by introducing a threshold for the numerator of the formula. When the numerator is lower than the threshold, the cell is labeled as smooth. A value higher than 10^7 for the threshold might force the solver to apply highorder reconstruction across shocks, and therefore will not apply the piecewise linear reconstruction which prevents oscillations. We observe that the CENO scheme might cause unphysical states in both inviscid and viscous regime. By reconstructing the conservative variables instead of the primitive ones, we are able to prevent unphysical states for inviscid flows. For the viscous flows, temporarily reverting to firstorder reconstruction in the cells where the temperature is computed as negative, prevents unphysical states. This technique is solely required during the first iterations of the solver, when the flow is started impulsively. In this study the CENO, the AUSM and the AMR methods are combined and applied successfully to supersonic problems. When modeling supersonic flow with highorder accuracy in space, one should prefer the combination of the AUSM schemes and the CENO scheme. While the CENO scheme is simpler than the WENO scheme used in comparison, we show that it yields results of comparable accuracy. Although it was beyond the scope of this study, the AUSM can be extended to real gas modeling which constitutes another advantage of this approach.

ÖgeA modified anfis system for aerial vehicles control(Lisansüstü Eğitim Enstitüsü, 2022) Öztürk, Muhammet ; Özkol, İbrahim ; 713564 ; Uçak ve Uzay MühendisliğiThis thesis presents fuzzy logic systems (FLS) and their control applications in aerial vehicles. In this context, firstly, type1 fuzzy logic systems and secondly type2 fuzzy logic systems are examined. Adaptive NeuroFuzzy Inference System (ANFIS) training models are examined and new type1 and type2 models are developed and tested. The new approaches are used for control problems as quadrotor control. Fuzzy logic system is a humanly structure that does not define any case precisely as 1 or 0. The Fuzzy logic systems define the case with membership functions. In literature, there are very much fuzzy logic applications as data processing, estimation, control, modeling, etc. Different Fuzzy Inference Systems (FIS) are proposed as Sugeno, Mamdani, Tsukamoto, and ¸Sen. The Sugeno and Mamdani FIS are the most widely used fuzzy logic systems. Mamdani antecedent and consequent parameters are composed of membership functions. Because of that, Mamdani FIS needs a defuzzification step to have a crisp output. Sugeno antecedent parameters are membership functions but consequent parameters are linear or constant and so, the Sugeno FIS does not need a defuzzification step. The Sugeno FIS needs less computational load and it is simpler than Mamdani FIS and so, it is more widely used than Mamdani FIS. Training of Mamdani parameters is more complicated and needs more calculation than Sugeno FIS. The Mamdani ANFIS approaches in the literature are examined and a new Mamdani ANFIS model (MANFIS) is proposed. Training performance of the proposed MANFIS model is tested for a nonlinear function and control performance is tested on a DC motor dynamic. Besides, ¸Sen FIS that was used for estimation of sunshine duration in 1998, is examined. This ¸SEN FIS antecedent and consequent parameters are membership functions as Mamdani FIS and needs to defuzzification step. However, because of the structure of the ¸Sen defuzzification structure, the ¸Sen FIS can be calculated with less computational load, and therefore ¸Sen ANFIS training model has been created. These three approaches are trained on a nonlinear function and used for online control. In this study, the neurofuzzy controller is used as online controller. Neurofuzzy controllers consist of simultaneous operation of two functions named fuzzy logic and ANFIS. The fuzzy logic function is the one that generates the control signal. It generates a control signal according to the controller inputs. The other function is the ANFIS function that trains the parameters of the fuzzy logic function. Neurofuzzy controllers are intelligent controllers, independent of the model, and constantly adapting their parameters. For this reason, these controllers' parameters values are constantly changing according to the changes in the system. There are studies on different neurofuzzy control systems in the literature. Each approach is tested on a DC motor model that is a singleinput and singleoutput system, and the neurofuzzy controllers' advantages and performances are examined. In this way, the approaches in the literature and the approaches added within the scope of the thesis are compared to each other. Selected neurofuzzy controllers are used in quadrotor control. Quadrotors have a twostage controller structure. In the first stage, position control is performed and the position control results are defined as angles. In the second stage, attitude control is performed over the calculated angle values. In this thesis, the neurofuzzy controller is shown to work perfectly well in single layer control structures, i.e., there was not any overshooting, and settling time was very short. But it is seen from quadrotor control results that the neurofuzzy controller can not give the desired performance in the twolayered control structure. Therefore, the feedback error learning control system, in which the fuzzy controller works together with conventional controllers, is examined. Fundamentally, there is an inverse dynamic model parallel to a classical controller in the feedback error learning structure. The inverse dynamic model aims to increase the performance by influencing the classical controller signal. In the literature, there are a lot of papers about the structure of feedback error learning control and there are different proposed approaches. In the structure used in this work, fuzzy logic parameters are trained using ANFIS with error input.The fuzzy logic control signal is obtained as a result of training. The fuzzy logic control signal is added to the conventional controller signal. This study has been tested on models such as DC motor and quadrotor. It is seen that the feedback error learning control with the ANFIS increases the control performances. Antecedent and consequent parameters of type1 fuzzy logic systems consist of certain membership functions. A type2 FLS is proposed to better define the uncertainties, because of that, type2 fuzzy inference membership functions are proposed to include uncertainties. The type2 FLS is operationally difficult because of uncertainties. In order to simplify type2 FLS operations, interval type2 FLS is proposed as a special case of generalized type2 FLS in the literature. Interval type2 membership functions are designed as a twodimensional projection of general type2 membership functions and represent the area between two type1 membership functions. The area between these two type1 membership functions is called Footprint of Uncertainty (FOU). This uncertainty also occurs in the weight values obtained from the antecedent membership functions. Consequent membership functions are also type2 and it is not possible to perform the defuzzification step directly because of uncertainty. Therefore, type reduction methods have been developed to reduce the type2 FLS to the type1 FLS. Type reduction methods try to find the highest and lowest values of the fuzzy logic model. Therefore, a switch point should be determined between the weights obtained from the antecedent membership functions. Type reduction methods find these switch points by iterations and this process causes too much computation, so many different methods have been proposed to minimize this computational load. In 2018, an iterativefree method called Direct Approach (DA) was proposed. This method performs the type reduction process faster than other iterative methods. In the literature, studies such as neural networks and genetic algorithms on the training for parameters of the type2 FLS still continue. These studies are also used in the interval type2 fuzzy logic control systems. There are proposed interval type2 ANFIS structures in literature, but they are not effective because of uncertainties of interval type2 membership functions. FLS parameters for ANFIS training should not contain uncertainties. However, the type2 FLS should inherently contain uncertainty. For this reason, KarnikMendel algorithm is modified, which is one of the typereduction methods, to apply the ANFIS on interval type2 FLS. The modified KarnikMendel algorithm gives the same results as the KarnikMendel algorithm. The modified KarnikMendel algorithm also gives exact parameter values for use in ANFIS. One can notice that the ANFIS training of the interval type2 FLS has been developed successfully and has been used for system control.

ÖgeA multidisciplinary design approach for conceptual sizing of advanced rotor blades(Lisansüstü Eğitim Enstitüsü, 20220719) İbaçoğlu, Hasan ; Arıkoğlu, Aytaç ; 511072102 ; Aeronautics and Astronautics EngineeringRotorcrafts are versatile vehicles with their unique hovering flight capability. However, their forward flight speed limitations and high noise levels are shortened to their usage in much wider areas. Therefore, the rotorcraft industry working on advanced rotorcraft, which are called compound rotorcrafts, development projects increasingly to overcome these problems. The conceptual design phase is the beginning of a development project where the most critical decisions are taken in this stage. So, vehiclelevel optimization algorithms are needed for decisionmaking to lead the project correctly. On the other hand, simplified lowlevel approaches must be used during conceptual design optimization because of too many design parameters to avoid impractical solution times. Furthermore, rotorcrafts with advanced rotors require advanced design approaches to obtain superior performance, structural, and noiselevel characteristics. Therefore, advanced conceptual design approaches are needed to overcome this contradiction. The rotor is the most critical component, which is also the source of the most problems of a rotorcraft such as lack of performance and noise. Therefore, rotor blade optimization is the main issue in the conceptual design phase at the beginning of a project. A multidisciplinary rigid rotor blade design optimization approach that is suitable for the conceptual design, sizing, and evaluation stages of helicopter development processes is suggested. Performance, structural strength of the blade, and noiselevel predictions are considered for the objective function. Blade outer surface and structure are represented by a geometrical model in which the chord, thickness ratio, chamber ratio, and twist distributions along the blade radial stations can be defined as linear or nonlinear functions. The distribution of the number of layers for both skin and spar was also defined in the presented model parametrically. Lowlevel but sufficient fidelity analysis methods were chosen to be able to reduce the computing time. Performance analysis and sizing of the vehicle were obtained by Blade Element Momentum Theory (BEMT) based inhouse developed helicopter sizing code called ROTAP. A trim algorithm for compound helicopters that may have additional lifting surfaces and thrust components is suggested. Airfoil Characteristics are calculated by the wellknown panel method code Xfoil. Both these codes are modified and embedded in the code developed for this study. Structural analysis was obtained using the 1D FEM approach. Crosssectional properties of the composite beam are calculated by VABS and displacements under the loads are calculated by GEBT. Reduced FfowcsWilliamsHawkings equations are used to estimate loading, thickness, and highspeed impulsive noise levels. A hybrid optimization algorithm is suggested to get optimal results. Sequential Quadratic Programming (SQP) can be used to find local optimal points. And then the global optimal point is searched by RSM over local optimal points iteratively. RSMbased surrogate modeling, evaluation, and optimization tool was also developed for manual inspection of the design space. As a case study, multiobjective aerodynamic performance optimization of aircraft propeller is performed.

ÖgeA numerical approach for plasma based flow control(Graduate School, 20230405) Ata, Reşit Kayhan ; Şahin, Mehmet ; 511132114 ; Aeronautics and Astronautics EngineeringIn the present study, a novel numerical method has been developed to solve incompressible magnetohydrodynamics (MHD) and electrohydrodynamics (EHD) flow problems in a parallel monolithic (fullycoupled) approach. To solve the fluid flow, incompressible NavierStokes equations are discretized using face/edge centered unstructured Finite Volume Method (FVM). The same formulation is used for the magnetic transport equation to model the magnetic effects. The sidecentered approach, where the velocity and magnetic field components are placed at the center of each cell face while pressure and Lagrange variables are placed at the center of the control volume, provides a stable numerical algorithm without the need of modifications for pressurevelocity coupling. The discretization of both MHD and EHD equations described above results in saddle point problem in fully coupled (monolithic) form. In order to solve this problem an upper triangular right preconditioner is used and restricted additive Schwarz preconditioner with FGMRES algorithm is employed to solve the system. Domain decomposition is handled by METIS library. For these numerical algorithms PETSc software package is used. For the solution of incompressible MHD flow problems, the continuity, incompressible NavierStokes, magnetic induction equation are solved along with the divergence free condition of magnetic field. Due to the interaction between magnetic field and conducting fluids, Lorentz force term is added to the fluid momentum equation. For the numerical stability, a Lagrange multiplier term is used in the magnetic induction equation, which has no physical meaning nor effect on the solution. The original approach satisfies the mass conservation within each element but it is not necessarily satisfied in the momentum control volume. Two modifications are proposed as a remedy. First, the convective fluxes are computed over the twoneighbouring elements which then resulted in improved mass conservation over the momentum control volume and increased stability. The second modification applies to only twodimensional MHD flows. The Lorentz force term in the momentum equation is replaced with $\sigma [\textbf{E} + \textbf{u} \times \textbf{B}] \times \textbf{B}$. Neglecting $\textbf{E}$ makes this term similar to mass matrix if $\textbf{B}$ is taken from the previous time step. Therefore, this modification improves the preconditioning of the monolithic approach. The developed solver is first validated for twodimensional Hartmann flow of which the analytical solution is known. Then liddriven cavity and backward facing step problems are investigated under external magnetic field both in 2D and 3D with insulating walls. Threedimensional MHD flow in ducts is another case where analytic solutions exist. Both conducting and insulating wall boundary conditions are employed and validated. Finally twodimensional flow over circular cylinder and NACA 0012 profile are investigated for vertical/horizontal external magnetic field and insulating/conducting boundaries. The eletrohydrodynamics (EHD) flow problems involve the interaction between electric field and charged particles inside the fluid. In the present study, the effect of plasma on the flow over lifting bodies is investigated and the working fluid is air, which is neutral at standard conditions. Therefore, a device called Dielectric Barrier Discharge (DBD) is used to ionize the air in a small volume near the surface. DBD consists of two electrodes separated by a dielectric layer. When a voltage is applied to the electrodes, ionization takes place. In order to simulate this phenomenon, Suzen\&Huang model is employed in which Poisson equation is solved for electric potential and charge density, separately. Once potential and charge density are known Coulumb force can be calculated and added as a body force term in the incompressible NavierStokes equation. The sidecentered approach is used for the velocity components and pressure is placed at the element center for the momentum and continuity equations. For the solution of Poisson equation the charge density and electric potential are placed at the element center while gradients are defined at the edge centers. The solver is first applied to an EHD flow in quiescent air and compared with both experimental and numerical solutions. Then, two electrodes are placed at the bottom wall of 2D cavity with a moving lid to investigate the effect of electric field on classical cavity problem. Finally, EHD flow over NACA 0012 airfoil at angle of attacks up to $\alpha=7$ is investigated in terms of flow structure, lift and drag coefficients.

ÖgeA study on optimization of a wing with fuel sloshing effects(Graduate School, 20220124) Vergün, Tolga ; Doğan, Vedat Ziya ; 511181206 ; Aeronautics and Astronautics Engineering ; Uçak ve Uzay MühendisliğiIn general, sloshing is defined as a phenomenon that corresponds to the free surface elevation in multiphase flows. It is a movement of liquid inside another object. Sloshing has been studied for centuries. The earliest work [48] was carried out in the literature by Euler in 1761 [17]. Lamb [32] theoretically examined sloshing in 1879. Especially with the development of technology, it has become more important. It appears in many different fields such as aviation, automotive, naval, etc. In the aviation industry, it is considered in fuel tanks. Since outcomes of sloshing may cause instability or damage to the structure, it is one of the concerns about aircraft design. To prevent its adverse effect, one of the most popular solutions is adding baffles into the fuel tank. Still, this solution also comes with a disadvantage: an increase in weight. To minimize the effects of added weight, designers optimize the structure by changing its shape, thickness, material, etc. In this study, a NACA 4412 airfoilshaped composite wing is used and optimized in terms of safety factor and weight. To do so, an initial composite layup is determined from current designs and advice from literature. When the design of the initial system is completed, the system is imported into a transient solver in the Ansys Workbench environment to perform numerical analysis on the time domain. To achieve more realistic cases, the wing with different fuel tank fill levels (25%, 50%, and 75%) is exposed to aerodynamic loads while the aircraft is rolling, yawing, and dutch rolling. The aircraft is assumed to fly with a constant speed of 60 m/s (~120 knots) to apply aerodynamic loads. Resultant force for 60 m/s airspeed is applied onto the wing surface by 1Way FluidStructure Interaction (1Way FSI) as a distributed pressure. Using this method, only fluid loads are transferred to the structural system, and the effect of wing deformation on the fluid flow field is neglected. Once gravity effects and aerodynamic loads are applied to the wing structure, displacement is defined as the wing is moving 20 deg/s for 3 seconds for all types of movements. On the other hand, fluid properties are described in the Ansys Fluent environment. Fluent defines the fuel level, fluid properties, computational fluid dynamics (CFD) solver, etc. Once both structural and fluid systems are ready, system coupling can perform 2Way FluidStructure Interaction (2Way FSI). Using this method, fluid loads and structural deformations are transferred simultaneously at each step. In this method, the structural system transfers displacement to the fluid system while the fluid system transfers pressure to the structural system. After nine analyses, the critical case is determined regarding the safety factor. Critical case, in which system has the lowest minimum safety factor, is found as 75% filled fuel tank while aircraft dutch rolling. After the determination of the critical case, the optimization process is started. During the optimization process, 1Way FSI is used since the computational cost of the 2Way FSI method is approximately 35 times that of 1Way FSI. However, taking less time should not be enough to accept 1Way FSI as a solution method; the deviation of two methods with each other is also investigated. After this investigation, it was found that the variation between the two methods is about 1% in terms of safety factors for our problem. In the light of this information, 1Way FSI is preferred to apply both sloshing and aerodynamic loads onto the structure to reduce computational time. After method selection, thickness optimization is started. Ansys Workbench creates a design of experiments (DOE) to examine response surface points. Latin Hypercube Sampling Design (LHSD) is preferred as a DOE method since it generates noncollapsing and spacefilling points to create a better response surface. After creating the initial response surface using Genetic Aggregation, the optimization process is started using the MultiObjective Genetic Algorithm (MOGA). Then, optimum values are verified by analyzing the optimum results in Ansys Workbench. When the optimum results are verified, it is realized that there is a notable deviation in results between optimized and verified results. To minimize the variation, refinement points are added to the response surface. This process is kept going until variation comes under 1%. After finding the optimum results, it is noticed that its precision is too high to maintain manufacturability so that it is rounded into 1% of a millimeter. In the end, final thickness values are verified. As a result, optimum values are found. It is found that weight is decreased from 100.64 kg to 94.35 kg, which means a 6.3% gain in terms of weight, while the minimum safety factor of the system is only reduced from 1.56 to 1.54. At the end of the study, it is concluded that a 6.3% reduction in weight would reflect energy saving.

ÖgeA study on static and dynamic buckling analysis of thin walled composite cylindrical shells(Graduate School, 20220124) Özgen, Cansu ; Doğan, Vedat Ziya ; 511171148 ; Aeronautics and Astronautics Engineering ; Uçak ve Uzay MühendisliğiThinwalled structures have many useage in many industries. Examples of these fields include: aircraft, spacecraft and rockets can be given. The reason for the use of thinwalled structures is that they have a high strength weight ratio. In order to define a cylinder as thinwalled, the ratio of radius to thickness must be more than 20, and one of the problems encountered in the use of such structures is the problem of buckling. It is possible to define the buckling as a state of instability in the structure under compressive loads. This state of instability can be seen in the load displacement graph as the curve follows two different paths. The possible behaviors; snap through or bifurcation behavior. Compressive loading that cause buckling; there may be an axial load, torsional load, bending load, external pressure. In addition to these loads, buckling may occur due to temperature change. Within the scope of this thesis, the buckling behavior of thinwalled cylinders under axial compression was examined. The cylinder under the axial load indicates some displacement. When the amount of load applied reaches critical level, the structure moves from one state of equilibrium to another. After some point, the structure shows high displacement behavior and loses stiffness. The amount of load that the structure will carry decreases considerably, but the structure continues to carry loads. The behavior of the structure after this point is called postbuckling behavior. The critical load level for the structure can be determined by using finite elements method. Linear eigenvalue analysis can be performed to determine the static buckling load. However, it should be noted here that eigenvalueeigenvector analysis can only be used to make an approximate estimate of the buckling load and input the resulting buckling shape into nonlinear analyses as a form of imperfection. In addition, it can be preferred to change parameters and compare them, since they are cheaper than other types of analysis. Since the buckling load is highly affected by the imperfection, nonlinear methods with geometric imperfection should be used to estimate a more precise buckling load. It is not possible to identify geometric imperfection in linear eigenvalue analysis. Therefore, a different type of analysis should be selected in order to add imperfection. For example, an analysis model which includes imperfection can be established with the Riks method as a nonlinear static analysis type. Unlike the NewtonRapson method, the Riks method is capable of backtracking in curves. Thus, it is suitable for use in buckling analysis. In Riks analysis, it is recommended to add imperfection in contrast to linear eigenvalue analysis. Because if the imperfection is added, the problem will be bifurcation problem instead of limit load problem and sharp turns in the graph can cause divergence in analysis. Another nonlinear method of static phenomena is called quasistatic analysis which is used dynamic solver. The important thing to note here is that the inertial effects should be too small to be neglected in the analysis. For this purpose, kinetic energy and internal energy should be compared at the end of the analysis and kinetic energy should be ensured to be negligible levels besides internal energy. Also, if the event is solved in the actual time length, this analysis will be quite expensive. Therefore, the time must be scaled. In order to scale the time correctly, frequency analysis can be performed first and the analysis time can be determined longer than the period corresponding to the first natural frequency. For three analysis methods mentioned within this study, validation studies were carried out with the examples in the literature. As a result of each type of analysis giving consistent results, the effect of parameters on static buckling load was examined, while linear eigenvalue analysis method was used because it was also sufficient for cheaper analysis method and comparison studies. While displacementcontrolled analyses were carried out in the static buckling analyses mentioned, loadcontrolled analyses were performed in the analyses for the determination of dynamic buckling force. As a result of these analyses, they were evaluated according to different dynamic buckling criteria. There are some of the dynamic buckling criteria; Volmir criterion, BudianskyRoth criterion, HoffBruce criterion, etc. When BudianskyRoth criterion is used, the first estimated buckling load is applied to the structure and displacement  time graph is drawn. If a major change in displacement is observed, it can be assumed that the structure is dynamically buckled. For HoffBruce criterion, the speed  displacement graph should be drawn. If this graph is not focused in a single area and is drawn in a scattered way, it is considered that the structure has moved to the unstable area. As in static buckling analyses, dynamic buckling analyses were primarily validated with a sample study in the literature. After the analysis methods, the numerical studies were carried out on the effect of some parameters on the buckling load. First, the effect of the stacking sequence of composite layers on the buckling load was examined. In this context, a comprehensive study was carried out, both from which layer has the greatest effect of changing the angle and which angle has the highest buckling load. In addition, the some angle combinations are obtained in accordance with the angle stacking rules found in the literature. For those stacking sequences, buckling forces are calculated with both finite element analyses and analytically. In addition, comparisons were made with different materials. Here, the buckling load is calculated both for cylinders with different masses of the same thickness and for cylinders with different thicknesses with the same mass. Here, the highest force value for cylinders with the same mass is obtained for a uniform composite. In addition, although the highest buckling force was obtained for steel material in the analysis of cylinders of the same thickness, when we look at the ratio of buckling load to mass, the highest value was obtained for composite material. In addition, the ratio of length to diameter and the effect of thickness were also examined. Here, as the length to diameter ratio increases, the buckling load decreases. As the thickness increases, the buckling load increases with the square of the thickness. In addition to the effect of the length to diameter ratio and the effect of thickness, the loading time and the shape of the loading profile are also known in dynamic buckling analysis. In addition, the critical buckling force is affected by imperfections in the structure, which usually occur during the production of the structure. How sensitive the structures are to the imperfection may vary depending on the different parameters. The imperfection can be divided into three different groups as geometric, material and loading. Cylinders under axial load are particularly affected by geometric imperfection. The geometric imperfection can be defined as how far the structure is from a perfect cylindrical structure. It is possible to determine the specified amount of deviation by different measurement methods. Although it is not possible to measure the amount of imperfection for all structures, an idea can be gained about how much imperfection is expected from the studies found in the literature. Both the change in the buckling load on the measured cylinders and the imperfection effect of the buckling load can be measured by adding the measured amount of imperfection to the buckling load calculations. In cases where the amount of imperfection cannot be measured, the finite element can be included in the analysis model as an eigenvector imperfection obtained from linear buckling analysis and the critical buckling load can be calculated for the imperfect structure using nonlinear analysis methods. In this study, studies were carried out on how imperfection sensitivity changes under both static and dynamic loading with different parameters. These parameters are the the lengthtodiameter ratio, the effect of the stacking sequence of the composite layers and the added imperfection shape. The most important result obtained in the study on imperfection sensitivity is that the effect of the imperfection on the buckling load is quite high. Even geometric imperfection equal to thickness can cause the buckling load to drop by up to half.

ÖgeAdvanced energy and exergy analysis on aircraft jet engines(Graduate School, 20231208) Fawal, Sara ; Kodal, Ali ; 511212113 ; Aeronautics and Astronautics EngineeringA comparative performance analysis for various optimization criterion functions is to be carried out for an irreversible Brayton cycle applicable to aircraft jet engines: Ramjet, Turbojet (No Afterburner), Turbojet (With Afterburner), TurboRamjet. Newly defined parameters are introduced as power loss parameter (PLOS), effective power loss parameter (EPLOS) and CarnotBrayton shape factor (CBSF) for a better assessment of the performance and power losses throughout the operation of the engine cycle. In addition, optimization functions, such as maximum power (MP), maximum power density (MPD), ecological coefficient of performance (ECOP) and ecological function (ECOL) are considered and their optimal operation conditions are compared with respect to each other. This research studied the effects on the prescribed optimization criterions targeted towards the aviation industry under variations of compressor pressure ratio θ_c, compressor and turbine efficiencies (η_c and η_t respectively), cycle temperature ratio / maximum cycle temperature, altitude and flight Mach number M_∞ where applicable with respect to the jet engine being considered. Therefore, the classical irreversible Brayton cycle is extended and applied to airbreathing engines; which included effects of all the engine components (from free stream to inlet to outlet) as part of the thermodynamic cycle model. While many researchers have carried out performance analysis for internal combustion engines including gas turbine engine, this study is an extension of the available optimization functions such as MP, MPD, ECOP and ECOL for aircraft jet engines. As mentioned, power density is defined as the ratio of power to the maximum specific volume in the cycle. Whereas ECOP is defined as the ratio of power output to the loss rate of availability and ECOL as the power output minus the loss rate of availability. In order to extend the classical irreversible Brayton cycle to airbreathing engines applicable for aircrafts, further development studies must be carried out to obtain: higher propulsion efficiency and higher ratios of power output with respect to engine weight, volume, and frontal area. The objective is to obtain a larger power output to engine size (weight) in a more thermodynamically efficient manner for a real turbojet cycle where maximum ECOP, ECOL, power density and power conditions can be used as a basis for the determination of optimal operating conditions and preliminary design constraints for real turbojet engines at flight conditions. The comparative performance analysis for various optimization criterion functions used for the aircraft engine cycle will be applied to ramjet, turbojet without afterburner and tubojet with afterburner to reach the final intended application of turboramjet engine. The turboramjet engine cycle is identified as Turbine Based Combined Cycle Engines (TBCC). Such hybrid cycle engines can be applied to UAV's, UCAV's and powering future hypersonic flight vehichles. The software to be used for the comparative performance analysis for the irreversible Brayton cycle applicable to aircraft jet engine cycles is the academic version of MATLAB 2018b provided by the MathWorks group. The emissions and radiative forcing (RF) from the aviation industry and its effects on air pollution and the ecology are an important concern, where aviation ranks as one of the top ten emitters. The major greenhouse gas emitters that contribute to RF are: carbon dioxide CO2, carbon monoxide CO, water H2O, nitrous oxide NOX, sulphur oxides SOX and volatile organic compounds VOCs. Thus, performance evaluation of aircraft propulsion systems must be assessed with respect to environmental and ecological conditions as well as power and fuel consumption considerations. Therefore, various optimization criterion functions which can be used as tools by the aviation industry to design 'new generation engines' which are economically and ecologically favourable. It is anticipated that this research would provide valuable insight in the preliminary design of airbreathing engines (Ramjet, Turbojet: No Afterburner, Turbojet: With Afterburner and TurboRamjet) and set a stage for exploration towards adaptive engine components and cycles for the conception of truly intelligent engines; an engine that can assess its current operating state and work under the most efficient power regime (ECOL or ECOP or MP or MPD) to achieve the designers and engine's intended performance potential.

ÖgeAeroacoustic investigations for a refrigerator air duct and flow systems(Graduate School, 20220216) Demir, Hazal Berfin ; Çelik, Bayram ; 511181186 ; Aeronautics and Astronautics EngineeringNoise has become an important public health problem with industrialization, and has become a crucial design problem for engineering. For this reason, noise reduction studies have became the focus, especially in the white goods, automotive and aviation sectors, which requires interaction with human. Among the vehicles and products in the aforementioned sectors, the refrigerators, unlike the others, are located in the center of the living area and work throughout the day. Therefore, possible sound problems are observed more quickly by the users and are found to be disturbing. At this point, the investigation and reduction of the acoustic propagation of existing products by various numerical and experimental methods is a valuable contribution to both industry and literature. Within the scope of this thesis, the freezer compartment of a refrigerator with a No frost cooling system was investigated from an aeroacoustic perspective. The freezer compartment consists of three drawers where food will be placed, an axial fan that provides air flow, an evaporator cover that separates the evaporator pipes and the interior volume, and plastic walls surrounding them. The main source of air flow noise in the system is the axial fan. For this reason, in the first step of the study, solo aeroacoustic examination of the axial fan was made. Afterwards, the entire freezer volume was examined and the study was completed with three different model proposals in which acoustic emission was reduced. The flow field analysis of the axial fan with an operational speed of 1200 rpm was carried out with commercial software ANSYS Fluent. In this numerical model, Shear Stress Transport 𝑘 – 𝜔 turbulence model was used. Governing equations was solved under threedimensional, transient, viscous, incompressible flow assumptions. The rotation of the fan was defined by the sliding mesh method. The numerical flow solution was validated with experimental volumetric flow rate data. According to the numerical and experimental results, the flow rate of the axial fan under the specified conditions was determined as 19 L/s. A hybrid aeroacoustic model is created by giving the pressure outputs of the flow solution as input to the acoustic model. For the acoustic solution, Ffowcs Williams & Hawkings (FWH) model defined in ANSYS Fluent was used and the result of the solution was compared with the sound pressure data collected in the full anechoic acoustic room. Although there is some difference between the numerical and experimental sound pressure curves, it was observed that the hybrid model established to understand the general trend and to catch the blade passing frequency was successful. It was predicted that the difference between experimental and numerical measurements occurred for two reasons. The first is absence of the fan motor in the numerical analysis. Another reason is that the acoustic propagation resulting from the excitation of the air flow to the system structures cannot be predicted with this model. In the second step of the study, the model validated with axial fan solutions was applied to the freezer compartment. The aim here is to reveal the air flow distribution in the freezer volume and to identify the regions where turbulence effects increase. In the numerical model, the axial fan was rotated at an operational speed of 1200 rpm and this rotation was achieved by the sliding mesh method. As a result of the analysis, it was seen that the turbulence formation started at the wing tips as observed in the solo fan analyses, and the vortices coming out of the trailing edge tips were especially concentrated in the region between the upper wall of the freezer volume and the upper two drawers. In addition, a turbulent area was detected at the bottom of the evaporator cover (which is the fan suction area). As a result of the hybrid aeroacoustic model solution, the sound pressure data collected from 1 meter away from the front, rear and side surfaces of the freezer and the sound pressure data collected from the same locations in the full anechoic acoustic room were compared. When the total sound pressure in the range of 1010000 Hz is compared, it is seen that there is a difference of 37 dBA between the numerical model and the experimental results. As a result of the investigations of the axial fan in the solo and freezer volume, three different freezer models have been proposed to improve air flow, reduce turbulence and reduce the resulting noise caused by air flow. In the fist suggested model, the bottom part of the evaporator cover has changed and the acostic propagation has decreased 0.24 dBA at 1200 rpm rotational speed. The position of the axial fan and its distance from the structures in the suction and discharge directions are the parameters affecting the acoustic propagation. In the second model, it is aimed to provide acoustic gain by changing the fan position. In this context, the fan was moved on the shaft by 5 mm and brought closer to the blowing region. With this modification, total sound power level was decreased 2.18 dBA. The final model is the superposition of the first two models. Here, it was aimed to see the combined effect of two mentioned model. At 1200 rpm rotational speed, 3.27 dBA gain was achived by the third model.

ÖgeAn ALE framework for multiphase flows(Graduate School, 20220824) Güventürk, Çağatay ; Şahin, Mehmet ; 511162103 ; Aeronautical and Astronautical EngineeringAn Arbitrary Lagrangian Eulerian (ALE) framework which combines the advantages of both Lagrangian and Eulerian methods is developed to solve incompressible multiphase flow problems. The divstable side centered unstructured finite volume formulation is used for the discretization of the incompressible isothermal NavierStokes equations along with the isothermal constitutive equations for OldroydB and FENECR fluids. In this approach, the velocity vector components are defined at the midpoint of each cell face, while the pressure term and extra stress tensor are defined at element centroids. The present arrangement of the primitive variables leads to exact total mass conservation at machine precision due to the present stable numerical discretization with no adhoc modifications. In addition, a special attention is given to satisfy global discrete geometric conservation law (DGCL) at discrete level for the application of the interface kinematic boundary condition in order to conserve the total mass for each species for multiphase flow problems. Furthermore, the pressure field and extra stress field are treated to be discontinuous across the interface with the discontinuous treatment of density and viscosity and jump conditions are satisfied. Surface tension force is treated as a tangent force and discretized in a semiimplicit form. Two different approaches for the computation of unit normal vector have been implemented: the least squares biquadratic surface fitting (LSBSF) and the mean weighted by sine and edge length reciprocals (MWSELR). The combination of MWSELR method and discontinuous treatment of density and viscosity reduced the parasitic currents to the machine precision. The resulting large system of algebraic equations is solved in a fully coupled manner in order to improve the time step restrictions. As a preconditioner, an approximate matrix factorization similar to that of the projection method is employed and the parallel algebraic multigrid solver BoomerAMG provided by the HYPRE library, which is accessed through the PETSc library, has been utilized for the scaled discrete Laplacian of pressure and the diagonal blocks of mesh deformation equations. The present calculations verify that the mass of the bubble can be conserved at machine precision independent of spatial and temporal resolutions. The accuracy of the proposed method is initially validated on the static bubble problem, since the surface tension force is highly sensitive to the accurate evaluation of the unit normal vector and the inaccuracies significantly contribute to unphysical velocities, called parasitic currents. The calculations indicate that the parasitic currents can be reduced to machine precision for the MWSELR method. The MWSELR approach, as far as our knowledge goes, has not been used for the evaluation of normal vectors in multiphase flows. In the second benchmark case, the proposed approach is applied to the single bubble rising in a viscous quiescent liquid for both low and high density ratios. The calculations produce accurate predictions of the bubble shape, center of mass, rise velocity, etc. Furthermore, the mass of each species is conserved at machine precision and discontinuous pressure field is obtained in order to avoid errors due to the incompressibility restriction in the vicinity of liquidliquid interfaces at large density and viscosity ratios. The third benchmark case is rising of a Taylor bubble in 2D and in 3D. Taylor bubbles are large bullet shaped bubbles whose crosssection almost fill the crosssectional area of the channel. Therefore, this benchmark case is numerically harder than the previous cases. It is seen, 3D bubble rises faster due to the smaller blockage effect (i.e. cross section of the bubble/cross section of the tube) of the bubble in three dimension with respect to the 2D bubble. In addition, drag force of the bubble decreases due to the threedimensional relieving effect. The results are compared with the results available in the literature and it is shown that the obtained bubble shape and velocity field in the vicinity of the Taylor bubble are similar to that of the literature. In the fourth test case, rise of a single bubble in a quiescent, viscoelastic fluid due to buoyancy is simulated in 2D and the viscoelastic fluid is modeled as OldroydB. By changing the size of the bubble, domain, placing the bubble to the different locations and changing the fluid properties, many simulations are performed and the change in bubble shape, rise velocity, circularity and sphericity are inspected. It is seen that the existence of the wall highly effects the outcome. In addition, the cusp at the trailing edge of the bubble and negative wake behind the bubble are observed in some cases. Therefore, it is shown that a viscoelastic fluid model that exhibits shear thinning is not essential for negative wake to arise. This result contradicts with the some published papers in the literature but is also supported by the others. The final benchmark case is similar to the previous one but this time viscoelastic fluid is modeled as FENECR and the problem is in 3D. Besides, subsequent simulations are performed for Newtonian bubble and Newtonian continuum, Newtonian bubble and viscoelastic continuum, and viscoelastic bubble and Newtonian continuum. It is observed that the bubble has a slight cusp at the trailing edge for Newtonian bubble and viscoelastic continuum. On the other hand, the bubble has a dimple at the trailing edge for the viscoelastic bubble and Newtonian continuum. In addition, it is shown that the results are in a good agreement with the result available in the literature. Finally, the methods used to develop and test the present multiphase solver for both Newtonian and viscoelastic fluids are summarized. Advantages and the drawbacks of the present solver are addressed with possible future applications.

ÖgeAnalysis of aircraft landing gear brake induced vibrations(Graduate School, 20230123) Altınbağ, Öner ; Balkan, Demet ; 511191131 ; Aeronautics and Astronautics EngineeringToday, aviation systems are the product of more than 100 years of work. The most groundbreaking process in these studies was experienced during the cold war years. The achievements of many engineering activities today are based on the knowledge gained in these years. Some major problems have been completely resolved in this progress, and some of them still continue to be active problems. The landing gear system is always critical to aircraft and is the engineering solution for almost all functions on the ground. In recent years engineers have been trying to optimize previous achievements within the framework of weight reduction, reliability, integration, energy consumption, noise reduction, cost reduction, and maintenance activities. One of the most important problems related to landing gear systems from the past to the present is the vibration problem, which we can examine under noise reduction. In this study, the causes of vibrations originating from the landing gear braking system were examined together with previous studies in the literature. A comparative approach to brakeinduced vibrations, which is still seen as a problem today, has been sought as a solution using today's tools. In this context, the parameters required for an aircraft landing gear model were calculated with the preliminary design activities used in the literature and industry. With these calculations, a model was created using MSC ADAMS software. Tire models in multibody dynamics simulations for vehicle dynamics were examined. As a result, the most suitable tire model was selected for the scope of the study. The parameters of the relevant tire model have been modified from the result of the tire sizing calculations. Two different vibration frequencies were investigated under four different longitudinal velocity conditions in order to make a valuable comparison. The results obtained from the model were compared and interpreted by using the previous studies from the literature.

ÖgeAnalysis of bird strike on metallic panels(Graduate School, 20230615) Çayhan, Kenan ; Balkan, Demet ; 511201133 ; Aeronautics and Astronautics EngineeringThis thesis investigates the phenomenon of bird strikes, using a combination of literature analysis, statistical analysis, and theoretical models. The study focuses on the potential damage that bird strikes can cause to various parts of an aircraft, which are windfacing components such as wings, stabilizers, engines, and windshields. The variety of possible outcomes from a bird strike poses a significant threat to aviation safety, as bird strikes account for 90% of Foreign Object Damage (FOD) incidents. As a result, aviation regulations require aircraft to meet specific levels of bird strike tolerance for critical components, and there are a number of certification requirements that airplanes must meet to be regarded safe to fly. To investigate the bird strikes on aircraft, the study uses numerical models, including the Smooth Particle Hydrodynamics (SPH) model, which was used to simulate sandwich plate bird impact experiments. The study concludes that the SPH model may be useful for finite element bird strike case analyses, which can help to improve aviation safety by identifying potential vulnerabilities and developing effective prevention measures. When using a new numerical approach, it is important to compare the results to experimental data to ensure that the simulation accurately reflects reality. Many research studies have included both numerical simulations and experimental data to understand how well the simulation corresponds to realworld scenarios. Experimental studies have traditionally guided aircraft designers in creating structures that are tough enough to withstand bird strikes. However, as aircraft components have become more complex, it has become necessary to develop bird strike simulation programs to design aircraft parts that are both airworthy and can be produced quickly and economically. Furthermore, the optimization process typically involves many iterative steps, which makes computerbased analyses more efficient and cheaper than experiments. However, conducting experiments with real birds, which are often dead or drugged chickens, presents a number of issues. The reproducibility of experiments, the health of researchers, and the availability of suitable bird models are all concerns. Real bird torsos vary greatly, making it difficult to obtain consistent results. While certification regulations only define the mass properties of the bird, different bird species have different densities, leading to variations in pressure loads between tests. As a result of these difficulties, researchers have begun using substitute bird materials instead of real birds. Advancements in computer technology have led to the development of cheaper and more advanced finite element software since the 1980s. This has allowed scientists to analyze bird strikes numerically due to the low cost, speed, and repeatability of the analyses. Various substitute bird models have been investigated in studies, and results have been compared with experimental data. The simple cylinder geometry is still a valuable approach to compare simulation results with experimental data. Different geometries such as spheres, cylinders with flat or hemispherical ends, and ellipsoids may also be used in simulations. When birds are struck at high speeds, their behavior is different from that of a simple elastic solid, and it is the responsibility of scientists and engineers to study the behavior of bird materials both theoretically and experimentally. Statistical data related to bird strikes is provided in the thesis, and it is emphasized that frontfacing components of aircraft are the most critical as they are most likely to encounter a direct bird strike. The most frequently struck parts of an aircraft are the fuselage, nose, radome, windshield, wing, rotor, and jet engine. Approximately 70% of bird strikes occur at altitudes between zero and 152 meters, which is primarily during takeoff and landing. This information is useful in avoiding bird strike accidents. As the altitude of an aircraft increases, the natural habitats of birds become further from the plane. The velocity of the projectile has a significant impact on how it responds upon impact. The behaviour of the projectile can be divided into five categories based on the internal stresses it experiences: elastic impact, plastic impact, hydrodynamic impact, sonic impact, and explosive impact. Elastic impact occurs when the projectile material strength is well above the internal stresses caused by the low speeds and accelerations, resulting in the projectile bouncing back from the surface. As the impactor velocity increases, the projectile enters the plastic behavior region, yet the velocity is still low enough to maintain fluidlike flow behavior, causing the bird to spread in every direction parallel to the plate, and the load to expand to a larger area. The theory behind bird strike at velocities that cause the bird to act in the hydrodynamic region is investigated. When the impactor with the initial velocity hits a surface, materials in contact with the rigid plate would immediately come to rest, generating a shock wave with velocity normal to the plate and towards the impactor body. There would be a significant pressure gradient at the outer surface because there is shock load pressure on the inner side and free surface pressure on the outer side. Soft objects impacted at high velocities behave differently than at low velocities, such that even elastic solids behave like liquids. However, testing with real birds can yield scattered data and it is not ethical to kill animals for scientific purposes. Gelatine has been found to be a suitable artificial substitute material with uniform characteristics and can be shaped into simple geometries such as cylinders and spheres for easy handling. Finite element programs offer various solution methods for bird strike simulations. Lagrangian method involves nodes attached to the material while Eulerian method uses fixed nodes in a defined space where material flows through it. Arbitrary Lagrangian Eulerian method is another option that allows for the defined space to change with the material flow, leading to faster computation time. Additionally, the meshless method called smooth particle hydrodynamics allows for particles to move freely without mass distortion. Various basic shapes of birds can be examined for bird strike impacts, including a cylinder, a cylinder with hemispherical ends, an ellipsoid, or a sphere. For a bird with a mass of 1.8 kg and specific geometric parameters, the density of the bird can be determined to be 900 kilograms per cubic meter. Conversely, by using a standard density of 950 kilograms per cubic meter and entering the mass of the bird, a specific volume value can be determined and used to specify the bird's geometry. Honeycomb materials provide stiffness to the structure while not adding too much mass. Hence, honeycombs are a kind of deformable shock absorbers that is widely used in the aircraft industry. In the reference tests, they used single and double core honeycomb sandwich metal plates as specimens under bird strike. They made a correlation between test results and simulation results which can be beneficial. Modelling the material of honeycomb in LSDYNA has a number of challenges. Firstly, honeycomb has a complex geometry which is expensive to model and simulate with shell elements. Therefore, its effective behavior can be modelled under homogenized solid elements. Out of plane stress strain curve up to crushing was given at reference. Which can be inserted as a stress strain curve to the solid elements. Particle node quantity for the bird impactor and element number for the aluminum sheets and honeycomb is limited with the computer power. Therefore, node numbers are generally about 20519 for the bird material. The simulations provide spatial displacement values and nominal strain curve values that are generally similar to the experimental results. However, there are slight differences, which may be due to errors in both the simulations and the tests. Overall, the strain values align well with the experimental data for both simulations. Therefore, the SPH method can be effectively used to simulate bird strikes on honeycomb sandwich plates, which is advantageous since experimental studies can be timeconsuming and costly, especially in the initial design phase of aerospace vehicles.

ÖgeBir savaş uçağının burun iniş takımı yapısal analizi(Lisansüstü Eğitim Enstitüsü, 20230123) Aydın, Gözde ; Özkol, İbrahim ; 511191200 ; Uçak ve Uzay MühendisliğiHiç şüphesiz uçak tasarımında uçağın her bir komponent ayrı bir mühendislik süreci ve ciddi bir zaman gerektirmektedir. Uçaklarda iniş takımı önemli bir ana mekanik sistemdir. Bu tez çalışmasında da bir savaş uçağının burun iniş takımı tasarımı gerçekleştirilerek yapısal analizi yapılmıştır. Tez çalışması süresince, birçok kaynak incelenmiştir ve uzun bir literatür araştırma süreci gerçekleştirilmiştir. Havacılık endüstrisinde yüksek dayanımlı ve hafif bir yapı tasarlamak en kritik parametrelerdendir. İniş takımları uçakların toplam ağırlığının yaklaşık %6' sını oluşturur. Yani uçak ağırlığının büyük bir kısmını oluşturur. Dayanım/ağırlık oranı yüksek iniş takımı tasarlamak en önemli tasarım gerekliliğidir. İniş takımları, iniş ve kalkış sırasında uçağa gelen dinamik ve statik yüklere maruz kalır. Bu nedenle iniş takımı sisteminin bu yüklemelere karşı dayanımlı bir yapıya sahip olması gerekir. Yüklere dayanamadığı takdirde iniş takımı ve uçakta ciddi yapısal hasarlar meydana gelebilir. Havacılık tarihinden bu yana birçok farklı çeşitte iniş takımları tasarlanmıştır. İlk başta tasarımlarda sabit iniş takımları kullanılırken zaman içerisinde bu tip iniş takımlarının aerodinamik açıdan dezavantajlı olduğu görülmüştür. Uçaklarda daha yüksek hız ve daha uzun havada kalma süresi gibi isterleri karşılayabilmek için katlanabilir iniş takımları tasarlanmıştır. Daha kompleks bir yapı olmasına karşın uçaklarda performans isterleri de göz önüne alındığında katlanabilir iniş takımlarının kullanımı zamanla yaygınlaşmıştır. İniş takımı tipine karar verildikten sonra ana ve burun iniş takımının konumuna karar verilirken, ağırlık merkezinin konumu göz önüne alınarak uçağın yerde hareketi, devrilmemesi, yan rüzgar etkisini azaltması, iniş ve kalkış sırasında manevra kabiliyetine izin vermesi sağlanmalıdır. Öncelikle, iniş takımı analizi için bir tasarım hazırlanmıştır. Bu tasarım için bilinmesi gereken belli parametreler vardır. Bu parametreler uçağın ağırlık merkezi, yerden yüksekliği, ortalama veter uzunluğu, iniş takımları arasındaki mesafe gibi sıralanabilir. Literatürdeki savaş uçaklarının bir çoğu incelenerek bu parametreler ile ilgili veriler toplanmıştır. Ortalama bir değer seçilerek kavramsal tasarım için gerekli bilgiler elde edilmiştir. Böylelikle iniş takımının uçağın ağırlık merkezine göre yerleşimi yapılmıştır. Daha sonra, uçak yerleşimine göre iniş takımlarına gelen yüklemeler hesaplanmıştır. Yük hesaplamalarında literatürdeki kitaplardan faydalanılmıştır. Yüklere göre lastik boyutuu, amortisör stroğu ve dikme çapı belirlenmiştir. Parça çizimleri ve montajda Siemens NX programı kullanılmıştır. Ayrıntılı boyutlandırma ve çizim yapıldıktan sonra kritik yük koşullarını belirleyebilmek için farklı iniş koşullarında iniş takımına gelen üç eksendeki kuvvetler hesaplanmıştır. Farklı kuvvet ve doğrultularda en kiritik üç koşul seçilerek analizler bu iniş koşullarında gerçekleştirilmiştir. Sonlu elemalar yöntemi ile burun iniş takımı ANSYS Workbench programı kullanılarak analiz edilmiştir. Yapısal analizi gerçekleştirilen iniş takımında malzeme değişikliği yapılarak yapıların VonMises gerilme ve deformasyon değerleri elde edilmiştir. Yapıların maruz kaldığı yüklemeler yüksek olduğu için komponentlerde yüksek gerilmeler görülmüştür. Bu nedenle, iniş takımlarında malzeme seçilirken yüksek dayanım ve uzun ömre sahip olması önemlidir. Son olarak, yapılan analiz sonuçlarına göre parçaların ağırlıkları, deformasyon miktarları ve dayanımları karşılaştırılmıştır. Bu sonuçlar doğrultusunda tasarım kriterleri de göz önüne alınarak malzeme seçimi yapabilir veya tasarımda değişiklik kararı alınabilir. Bu şekilde yapılan analizler serisi ile optimum bir burun iniş takımı tasarımına ulaşmak mümkündür.

ÖgeCoherent structures and energy transfer in decelerated turbulent boundary layers(Graduate School, 20230210) Güngür, Taygun Recep ; Güngör, Ayşe Gül ; Maciel, Yvan ; 511162103 ; Aeronautical and Astronautical EngineeringThis thesis aims to expand our knowledge about turbulent boundary layers (TBLs) developing under adverse pressure gradients (APG). The main focus of this thesis is coherent structures and energy transfer mechanisms in APG TBLs with small and large velocity defects. For this, two novel nonequilibrium APG TBL direct numerical simulation databases are generated. The first database is a nonequilibrium APG TBL with $Re_\theta$ reaching 8000 and a shape factor spanning between approximately $1.4$ and $3.2$. It is the main database utilized throughout the thesis. The second database has identical domain and boundary conditions to the first one. The difference between them is that turbulence in the inner layer of the second database is artificially eliminated. This second database is generated to examine the effect of the inner layer on the outer layer turbulence. For comparison purposes, a channel flow case, two zero pressure gradient (ZPG) TBLs and two homogeneous shear turbulence (HST) databases from the literature are employed. The energycarrying and –transferring structures are examined using the spectral distributions and twopoint correlations. The analysis reveals that energycarrying structures in small defect APG TBLs and canonical flows have similar spatial and spectral features. In the large defect case, turbulence in the inner layer, which is the dominant region in canonical flows and small defect APG TBLs, loses its importance and outerlayer turbulence becomes dominant. The inner peak in the $\langle u^2\rangle$ spectra does not exist in the largedefect case. Moreover, twopoint correlations show that the spatial organization becomes different in the largedefect case as well. Regarding the energytransferring structures, production, pressurestrain and dissipation structures behave in a similar fashion to the energycarrying structures. The spectral distributions show that the canonical flows and small defect APG TBLs behave very similarly. The shape of the spectra is qualitatively similar in both cases. In the large defect case, the wallnormal distributions of production and pressurestrain become different since the outer layer becomes dominant. However, the shape of 2D spectra and the aspect ratio of structures are alike in all cases. The production and pressurestrain structures are analyzed in more detail using the relative size and wallnormal positions with respect to each other and energetic structures using spectral distributions. The results show that production and pressurestrain spectra have similar features in both the inner and outer layers regardless of the velocity defect, despite the differences in energetic structures. In the inner layer, the results suggest that the nearwall cycle or another mechanism with similar spectral features exists in large defect APG. As for the outer layer, an interesting result is that in largedefect APG TBLs it acts more like a free shear layer than in smalldefect APG TBLs or canonical flows. Besides that, production and intercomponent energy transfer mechanisms are similar in all cases regardless of velocity defect. No inflection point instability in the outer layer of the largedefect APG TBLs was detected. The effect of the nearwall region on the outerlayer layer structures is examined through Reynoldsshearstress carrying structures' spatial features by detecting individual structures using spatiotemporal volumetric data. The results show that the outer layer is not significantly affected by the innerlayer turbulent activity. The structures' spatial features mostly depend on the mean shear. The aspect ratio of Reynoldsshearstress carrying structures remains almost identical in the outer layer when the innerlayer turbulence is eliminated. Moreover, the aspect ratio follows a similar trend in both outer layers of APG TBLs and HSTs when the structures' size is normalized with the Corrsin length scale. The overall conclusion is that energy transfer mechanisms remain the same within one layer regardless of the velocity defect. The reason why the wallnormal distribution of energy and energy transfer dramatically changes in the large defect case is probably the change in the mean shear profile due to the increasing velocity defect.

ÖgeDatadriven delay estimation and anomaly detection: A study on European and Turkish air traffic(Graduate School, 20230518) Aksoy, Muhammet ; Koyuncu, Emre ; 511201136 ; Aeronautical and Astronautical EngineeringAir traffic networks represent highly complex and interconnected physical systems. Unlike other transportation networks, air traffic is very heavily regulated and physically constrained. Although the airways and airspaces are somehow more flexible compared to land based transportation systems, the fact that aircrafts can only positioned on and operated by airports make them quite dependent on the operations of the airports. Air traffic is regulated to ensure safety, while also maintaining the throughput of travel from one location to another. While these regulations does a decent job on keeping the air travel safe and systematical, they fall short when there are disruptions among the network that hinders the air traffic. There are numerous reasons for disruptions in air transportation; weather conditions, accidents, capacity constraints, personnel strikes etc. Yet their negative effect to the air traffic is mostly the same: introducing delays. Due to the connected nature of the air traffic and airports, when a delay generating event occurs at one place, the other members of the network could experience the similar effects, if not at a larger degree. This delay propagation means there is a ripple effect through the network which can snowball the delay generations and cause very large congestions. To relieve the effects of delay generating events, air traffic federators regulate the air traffic in a reactionary way. This may include reducing the capacity on certain airports or airways, giving NOTAMs, holding aircrafts on the ground or in the air (with hold patterns). Since all these actions are \emph{reactionary}, they are set in place after the delays already propagates through the network since it is trivial to asses and quantify the propagations in a large and complex network system. This study hypothesis that if the air traffic network can be modeled so that the propagations can be accurately calculated, it becomes possible to take proactive actions instead of reactive ones. Proactive actions are significantly more important when there is a risk of snowballing and propagation. It allows to take action when the ill effects are still contained on fewer members with smaller intensities. This paves the way for a more effective and less costly approach. Hence, the study proposes a method with 3 main parts; first one is to model the air traffic network so that propagations can be quantified, second one is to estimate the parameters of this model to keep a shortsighted vision into the upcoming network state and third one is to come up with a comprehensive action generating model to find optimal proactive actions that can keep the delay spreading at minimum and improve system resiliency. The air traffic modeling part is done via adopting compartmental model from epidemiology. This model explains the tranmission of disease within a population. When it is applied to the physical network system, instead of disease and humans, the delay amount and aircrafts is used. Additionally with the meta population model, instead of considering aircrafts one by one, airports can be used as they are focused points of aircraft populations. By linking transmit rate to the flight frequency between airports and the recovery rate to the delay handling characteristics of the airport, The parameter estimation part is done by calculating the historic recovery rates of the airports and then using deep learning inference to predict the next time step's recovery rates. The other parameters of the air traffic model, such as the traffic flow, is already known before hand (flight plans). Therefore through the estimation of recovery rate the network state of the upcoming states can be accurately predicted. This prediction can then be fed to the action generating algorithm to make the most informed decision. The action generating algorithm therefore must fundamentally be a deterministic state to action mapper. Reinforcement learning approach is utilized to train this state to action mapper to make it capable of generating optimal decisions under a sufficiently large spectrum of conditions. The final part of this study concerns with anomalous flight detection in air traffic as these types of flights are one of the sources of disruptions in an air traffic network. Although flight paths naturally diverge from one another, they still adhere to a set of patterns that have been tested in various environments and are optimized for them. These patterns may or may not be simple, depending on a number of factors, such as airspace use, the cognitive complexity of controllers, the weather, and NOTAMs. It is a challenging task to accurately classify flights just by their trajectories into a desired set of categories based solely on its statistical properties because of the high variance. For this purpose, the study incorporates a statistical approach that takes into account the timebased characteristics of the flight trajectories to determine whether they are abnormal or not. This statistical method with LSTM autoencoders makes it possible to train the model with historical data and quickly predict the flight class, taking into account the timebased characteristics of a flight trajectory. LSTM autoencoders can capture the class of a flight with relatively shorter time windows (16 second intervals). Therefore the air space can be periodically sweeped for anomalies while the network model and action algorithm runs in parallel. The obtained results demonstrate that the suggested architecture is quite capable of classifying abnormal flight trajectories as it successfully detects simulated fighter aircraft trajectories in airspaces with high commercial flight density. With the applications of deep learning and reinforcement learning, this whole methodology ensembles is largely datadriven, however the introduction of the compartmental model from epidemiology lays out a strong and accurate mathematical formula to support these datacentric approach. As the results suggests, The whole network's resiliency, i.e. its ability to keep delays from spreading and absorbing them, significantly increases when the optimal actions are reflected on the parameters. Additionally with the help of unsupervised learning, anomalous flights are also detected and represented as a disruption source to the network. Possible biases and shortcomings due to the datadriven approach is recognized throughout the study yet the overall method is deemed to be of significant importance in terms of managing resiliency through air traffic networks.

ÖgeDeğişken açılı elyaf kompozitlerin uygulanabilirlik açısından yeni bir tasarım yaklaşımı ve diferansiyel evrim ile burkulma yükü optimizasyonu(Lisansüstü Eğitim Enstitüsü, 20220622) Beyazgül, Umut ; Balkan, Demet ; Mecitoğlu, Zahit ; 511181207 ; Uçak ve Uzay MühendisligiHavauzay araçları başta olmak üzere bir çok sektörde kullanılan elyaf kompozitlerin başlıca tercih edilme nedenleri önceden kullanılan muadillerine göre daha yüksek spesifik dayanım ve daha düşük ağırlık ile birlikte büyük bir maliyet tasarrufu sağlamasıdır. Bir yapının maruz kaldığı yük isterleri göz önüne alınarak yöne bağlı mekanik özellikleri sayesinde her katmanda uygun bir oryantasyon açısı ile istifleme dizisi tasarlanarak geleneksel elyaf kompozitlerin performansı artırılmaktadır. Bu tezdeki değişken oryantasyon açısıyla ifade edilen ise elyaf kompozitlerin değişken direngenliğini, elyaf açısını aynı katman içerisinde değiştirerek serim güzergahını değiştirmektir. Böylece, elyaf kompozit malzemenin yöne bağlı mekanik özelliklerini kullanarak optimizasyon kapsamını genişletmekte ve potansiyeline ulaşmasını sağlamaktadır. Düzlem içinde daha avantajlı bir yük dağılımı oluşturmaya imkan vermektedir. Bu tasarıma sahip kompozitlerin üretimi, uzun yıllardır geleneksel kompozit üretiminde de kullanılan, yüksek hassasiyet ve hızlı serim sağlayan otomatik elyaf serim cihazı (AFP) ve otomatik bant serim cihazı (ATL) ile planlanmaktadır. Bu çalışmada, mevcut laminasyon teorilerinde sabit sayı olarak geçen her bir katman açısı yerine konuma bağlı, lineer değişen bir oryantasyon açısı tanımı yapılmıştır ve tek eksende yükleme altında burkulma incelenmiştir. Yüksek dereceden doğrusal olmayan, türevlenebilirlik açısından klasikgradyan bazlı yöntemlerde uygulaması zor olan amaç fonksiyonlarının optimizasyonunda buluşsal aramaya bağlı stokastik özelliği olan yöntemler arasından göreceli olarak daha hızlı ve yüksek doğruluğu olan evrimsel algoritma kullanılmıştır. Tasarım modellemesi ve optimizasyon kodu python programlama dilinde yazılarak Abaqus doğrusal burkulma analizi amaç fonksiyonu olarak entegre edilmiştir. Üretim sonrası kusurlardan kaçınmak için simetrik elyaf kompozit ve dengelenmiş katman açıları dizilimi tasarımın sınırlarını oluşturmuştur. Ayrıca, açı parametrelerinin alt ve üst limitlerinin yanı sıra, üretim cihazları ve filament demetlerinden kaynaklı eğrilik yarıçapı kısıtlaması da kullanılmıştır ve eğrilik yarıçap kısıtlama denklemi türetilmiştir. Optimizasyon ve sayısal analizler sonucunda, farklı katman sayılarında ve farklı boyen oranlarında değişken açılı elyaf kompozitlerin kritik burkulma yükü bakımından avantajlı olduğu gösterilmiştir ve boyen oranına ve katman sayısına göre yük kazanım oranları arasındaki korelasyon analiz edilmiştir.

ÖgeDesign and optimization of variable stiffness composite structures modeled using Bézier curve(Graduate School, 20220609) Coşkun, Onur ; Türkmen, Halit S ; 511162115 ; Aeronautics and Astronautics EngineeringThe usage of advanced fiberreinforced polymer (FRP) matrix composites has been dramatically increased since the first carbon fiber patented in the 1960's. Particularly, the aerospace companies' interest has been gradually grown in carbon fiberreinforced polymer (CFRP) aircraft structures due to major performance improvements such as high strength and stiffness to weight ratios and reduced weight. The traditional design approaches and manufacturing methodologies of CFRP structures in various industries have been well established and applied for more than 50 years. They are mainly developed for straight fibers and the optimum design solutions have been achieved by the choice of constituent materials, different fiber orientation angles that are often limited to 0, ±45, and 90 degrees, laminate stacking sequence and total number of plies. However, increasing complexity of structure geometries have resulted in complex layups & contours; therefore, advanced manufacturing methodologies such as Automated Fiber Placement (AFP) and Tailored Fiber Placement (TFP) are developed to improve productivity and process reliability. Following the introduction of advanced manufacturing methods CFRP structures with complex geometry, complex layups & contours have been manufactured with improved productivity and process reliability. In addition to that, composite materials can be tailored more effectively to meet design requirements by changing the design approach from straight to curvilinear fibers. The composite structures designed with curvilinear fibers have spatially varying stiffness due to local fiber orientations in the ply, and accordingly they are named as variable stiffness (VS) structures. In this dissertation, the variable stiffness composite plates and circular cylindrical shells modeled using parametric Bézier curves as curvilinear fiber paths are designed and optimized. The design method with parametric Bézier curves covers a wide and complex design space from simple linear angle variation to constant curvature path to highly nonlinear angle variations. The designed VS composite structures are expressed with new layup definition conventions that use simple and intuitive variables such as segment/station angles and multipliers/curvatures. The optimum structural designs in the complex design space of plates and circular cylindrical shells are searched using a multistep optimization with multiobjective such as buckling and stiffness, and a novel pretrained multistep/cycle surrogatebased optimization (PMSO) framework with single objective, i.e. buckling, respectively. First, VS composite plates and circular cylinders are designed with 'Direct Fiber Path Parameterization' (DFPP) that uses continuous curve functions for fiber orientation angles at each point or grid in the laminate. The cubic and quadratic Bézier curves are used as curvilinear fiber path. The fiber paths as Bézier curves are constructed with approximation and interpolation formulations. The approximation curve captures the defined angles at the start point and the end point, and the shape of the curve changes with the position of the control points intuitively. On the other end, interpolation curve follows the exact positions of control points at the expense of control of the fiber angle. Therefore, fiber angles are different from the defined sector angles. Three types of parametric curves are formulated, i.e., cubic Bézier interpolation curve and quadratic and cubic Bézier approximation curves. Cubic Bézier approximation curves are specially formulated to define constant curvature fiber paths. Considering the characteristics of Bézier curves, intuitive conventions to define layups of laminated VS plates and shells are proposed. The position of course boundaries within each ply are calculated using the reference fiber path, and resulting courses are shifted along one direction to cover VS plate and cylindrical shell surfaces. The reference fiber paths are defined with design variables such as sector/station angles and multipliers/curvatures, which are used to calculate control points. Current proposal for layup definition allows one to move stations using multipliers within an interval, hence it is possible to find lower curvature fiber paths with the same sector angles. The minimum curvature value is a major characteristic of curvilinear fiber paths due to manufacturing constraints. Golden Section Search and Downhill Simplex methodologies are used depending on the design approach together with Bézier curve formulations. The Golden Section Search method, which is a technique for finding an extremum, (minimum or maximum) of a unimodal function, is applied to approximation curves, and Downhill Simplex method is applied to interpolation curves due to a multidimensional space with n multipliers. The curvature values are significantly minimized without changing the layup definitions; especially for quadratic Bézier approximation curves, the curvature distribution along characteristic length gets close to the constant curvature results. Three different geometries for VS plates (b/a ≈ 1.8) and two different geometries for VS circular cylinder Cylinder 1 (L/R ≈ 2.67) and Cylinder 2 (L/R = 2) are modeled. Considering the cylindrical coordinates, the courses laid on the cylinder are axially shifted to have circumferentially varying stiffness and strength; however, the effective width of the ply is modified to have continuous fiber paths around the circumference. To have averaged boundaries, which is called no gap condition, minimum effective course width is used as the reference shifting value. The layup process is completed on developed plane of the cylinder, and then translated into cylindrical coordinates. Second, finite element models of laminated VS plates and cylindrical shells are generated using Ansys Mechanical APDL codes. Four node Shell 181 quadrilateral elements with full integration are used to mesh the VS plate and the VS composite shells with Cylinder 1 geometry, and FE models of layered VS composite shells with Cylinder 2 geometry are generated using eight node Shell 281 elements with reduced integration. Both shell elements are based on the firstorder sheardeformation theory (referred to as MindlinReissner shell theory). These elements with six degrees of freedom at each node (translations in the nodal x, y, and z directions and rotations about the nodal x, y, and z axis) are usually used to analyze thin to moderatelythick shell structures. The mesh convergence studies of reference QI plate and VS circular shells and plates are performed, and reference element edge lengths are chosen considering accurate mapping of curvilinear fiber paths on finite element mesh, buckling results, and computational efficiency. The curvilinear fiber paths for each ply are then mapped to related element centroids by APDL functions. Next, the VS laminates and circular cylinders are optimized for maximum stiffness and/or buckling load using surrogatebased NSGAII algorithm. The NSGAII is an evolutionary algorithm and supports multiobjective optimizations. The design space development strategy is an important part of surrogate modeling to get optimal distribution of fewest number of points with maximum insight into the design. Thus, experimental designs are generated with Optimal Space Filling (OSF) algorithm according to specified intervals. Then, surrogate models are generated with Genetic Aggregation. The Genetic Aggregation selects the best solution from Full 2ndOrder Polynomials, NonParametric Regression, Kriging, and Moving Least Squares. The algorithm generates the population of all methods and then it applies single response surface or combination of response surfaces according to fitness functions. The assemblage of Genetic Aggregation surrogate model is constructed with weighted average of selected metamodels. The weight and the combination of metamodels depend on design of experiment method and the behavior of VS structures designed with the approximation and interpolation curves. Twocycle approach is used to increase the accuracy of the surrogate models. The first cycle consists of the design space between 80° and 80°, and the second cycle searches for ±20 degree of the optimum angle calculated at the first cycle. A better layup for Size 1 – Case 3 compared to results in literature is found by using reduced the domain in the second cycle. The best buckling performance is found for Size 3 plate with Case 3 boundary conditions that has 103% increase in buckling load against 44% reduction in equivalent stiffness compared to reference quasiisotropic laminate. It is clear that increase in plate size increases the buckling performance of VS plates. This is due to wider design space with relaxed curvature constraint that allows higher angle differences between edge and the middle of the plate, accordingly fiber angle at the plate edges can align closer to loading direction while the fiber angles far from edge converges to smaller angles. The quadratic Bézier approximation curve is found to be a good alternative of cubic Bézier approximation curve with constant curvature, as it has similar edge load distribution and buckling mode shapes. Additionally, the stations, which are fixed for cubic Bézier approximation curve with constant curvature, can be shifted for definition with quadratic approximation without changing the layup definition according to designer's need. Finally, a novel pretrained design optimization framework is proposed to optimize buckling load of VS composite circular cylinders under pure bending with curvature and strength constraints. By using Bézier curves, designers have more effective control on the design domain to improve the buckling performance in accordance with requirements such as curvature and strength. The strength constrain is calculated by using TsaiWu failure criterion. The optimizations are conducted using PMSO framework that utilizes NSGAII. The main benefit of this framework is to gather prior knowledge about the design space at the first step by conducting pretraining optimizations using laminated VS composite shells with single ply definition. This narrows down the design space significantly before conducting a full layup design optimization with large number of parameters at the second step. Moreover, multiple cycle approach at each step helps to reduce the complexity of the optimization together with increased surrogate model accuracy. The optimization is completed for four different laminate stackups that are made up of all VS plies and partial VS plies in combination with unidirectional fibers (±45°, 0° and 90°). The maximum increase in buckling load is found to be 31% for Laminate 1 and 41% for Laminate 4 compared to reference QI shells. This gives 14% and 16% higher buckling load than the literature studies, and the Laminate 4 results are achieved for two times more design variables using approximately same number of sampling points. The gain in buckling load is due to the redistribution of stresses on compression and tension side as a consequence of variable angle distribution within each ply. The fiber angles close to axial direction on the tension side increase the strength and stiffness of the structure, and angles close to circumferential axis on the compression side reduce the stiffness of buckling critical region to distribute the compressive loads onto wider region.

ÖgeDevelopment and testing novel guidance algorithms for visual drone interception(Graduate School, 20240613) Çetin, Ahmet Talha ; Koyuncu, Emre ; 511211106 ; Aeronautics and AstronauticsThis study tackles the challenge of guiding a quadrotor to intercept fastmoving targets visual and radar feedback by Visual Inertial Odometry or GPS respectively. Proposed system, designed as a counter UAV solution, utilizes onboard camera and radar information of the aerial threat. Target interception process has been divided into two parts. One is preterminal phase guidance where target information comes from radar feedback. Unless the target has not been seen at the camera, interceptor guided from radar feedback. Once the target is detected by the camera, the quadrotor switches to terminal phase guidance which is guiding counter drone to aerial target by visual feedback. For preterminal guidance, two different algorithms were developed. A Model Predictive Control based guidance algorithm has been designed for preterminal guidance. For preterminal guidance, parallel interceptions (toward the head or back) provide robustness to inevitable visual processing latency in terminal phase compared to lateral engagements. By addressing these issues, the proposed methodology mainly utilizes Model Predictive Control (MPC) method with added terminal constraints to satisfy engagement at the desired angle. While formulating the MPC, the objective function in the MPC is modified to reduce the interceptor's requirement for maneuvering at the end of the trajectory. MPC prediction horizon is calculated by considering vehicle limits to satisfy the feasibility of the problem. Another method is we use Bezier Splines to guide the quadrotor. Since quadrotors has limited onboard computational power, MPC might not be practical for some cases. By ensuring continuity with Bezier Splines, the system determines the optimal interception direction (towards the head or tail) and calculates the timetogo, considering in the target's position and velocity along with the interceptor's kinematic constraints. This method specifically addresses latency issues in target detection, crucial for intercepting highspeed targets effectively. Moreover, the delays introduced by target detection and localization pose significant challenges, particularly for small quadrotors with limited computational power. The proposed approach aims to achieve parallel engagement with the target's velocity vector, whether from the front or rear, thus minimizing delays and overcoming visual tracking difficulties before target is detected by onboard camera. This strategy reduces lateral acceleration within the image frame during the final stages of interception, resulting in smaller miss distances. This outcome is consistent with established guidance literature, which recognizes the advantages of reduced acceleration at the end of the interception path. When the target is detected by camera using object detection algorithms, terminal phase guidance is initiated. For detecting aerial threats, the object detection algorithm You Only Look Once (YOLO) is used. Maintaining detection and tracking by camera can be interrupted due to limitations such as motion blur, noise in the image and getting out of the camera field of view. When detection is interrupted, Kalman Filter is used for prediction of the target. For image based guidance we utilized proportional guidance with some modifications. For this work we assume that no stabilizing mechanism that preserve orientation of the camera is used. Since no stabilizing mechanism is used for the camera, we formulized propotional guidance rules in roll and pitch stabilized frame in order not to being affected from camera orientation. We employed two distinct navigation methods: GPSbased navigation and Visual Inertial Navigation for navigating towards to target at the preterminal phase. The wellestablished opensource ArduPilot platform was utilized for GPSbased navigation, while VINSMono was implemented for Visual Inertial Navigation. As for controllers, due to the differing frequencies of estimated odometry data from these systems, different position controllers were employed for each navigation solution. The ArduPilot builtin controller was utilized for GPSbased navigation, whereas a custom controller was designed and flighttested for handling VIO feedback. The aforementioned navigation and control methods allowed us to compare and evaluate their performance in different scenarios. The GPSbased navigation provided a reliable and accurate solution in environments with clear GPS signals, while the Visual Inertial Navigation offered a robust alternative in situations where GPS signals were weak or unavailable. The custom controller designed for VIO feedback was optimized to handle the unique characteristics of visual inertial data, ensuring smooth and precise control of the quadrotor. Through this approach, we were able to develop a comprehensive navigation system that can adapt to various operational conditions, enhancing the overall reliability and effectiveness of the quadrotor's guidance and control. Finally, real world flight tests were conducted for assessing overall performance of the system. To evaluate the performance of the GPSbased and VIObase navigation algorithms, interception flights tests were conducted separately and the performance of the guidance algorithm was assessed accordingly. In realworld flight tests, we tested the use of Bezier splines in the preterminal along and imagebased visual servoing for the terminal phase. In doing so, we examined the use of GPSbased and VIO based navigation algorithms. Results show performance of the proposed methodology.

ÖgeDevelopment of a fault tolerant flight control system for a UAV(Graduate School, 20220812) Vural, Sıtkı Yenal ; Hacızade, Cengiz ; 511082105 ; Aeronautics and Astronautics EngineeringIt is important for the unmanned aerial vehicles that are used for various purposes including military missions, surveillance, security, atmospheric data gathering etc. to be autonomous and control systems that can work flawless even when faults are present are needed for such systems. The methods for achieving fault tolerant control are under development and are still not used much in practical applications however developing a fault tolerant control system for all types of applications including aerial vehicle control systems seems to be an ultimate control aim. In this thesis, developing a fault tolerant control system for a UAV is aimed mainly for the given reasons. In the study, active and passive control methods and Kalman filter based fault detection and isolation techniques are used together to build a fault tolerant controller for an unmanned aerial vehicle. Also, a hybrid controller including both active and passive fault tolerant controllers is developed in order to benefit from their different characteristics in dealing with faults. Kalman filter based fault detection and isolation algorithm which can be used to detect and isolate the faults in sensors and actuators , to determine the source of the fault and to find the unbiased sensor measurements is developed in the study and its effectiveness is shown through simulations. To detect and isolate the faults occuring in sensors/actuators Kalman filter innovation sequence analysis is used. On the other hand, to determine the source of the fault, DoyleStein method based Kalman filter and to rectify the biased sensor measurements Kalman filter insensitive to measurement failures are employed in the study. Unmanned aerial vehicle model is used to simulate the fault cases and to show the successfulness of the built system. One of the methods used in the thesis to build a fault tolerant controller is the active fault tolerant control method. In this method, fault detection and isolation technique is used to determine the faults occuring in the system and to find the severity of the fault and this info is used to reconfigure the controller. The actuators would not work effectively if hydraulic pressure decrease, partial blockage of a control valve, voltage reduction in electrical servosystems etc. occur in the system. In these cases, the effectiveness of the actuators decrease. The change in the mentioned actuator effectiveness can be represented in the system as the control effectiveness factor related with the actuator. In the study, twostage Kalman filter is used to estimate the changes occurring in actuator control effectiveness factors which corresponds to the faults occurring in actuators. With the help of twostage Kalman filter in which a second biasestimation filter is used, the bias in the system can be estimated and the best state estimates can still be found. This type of filter,different than the augmented state filter in which all parameters are estimated in one stage, has the advantage of reducing calculation burden and thus giving results in small time period. In short, twostage Kalman filter consists of a biasfree state estimator that estimates the states, a bias estimator to estimate the bias, the residual vector and the covariance matrix calculation equations and coupling equations that are used to relate the filters and update the bias free state estimator. In the simulations, the faults in actuators are modelled as changes in control distribution matrix B and these changes are tried to be estimated using twostage Kalman filter and the reconfiguration of the controller is done using the determined new B matrix. In this method, one needs to determine when the fault occurs in the system and to decide when to reconfigure the controller. To that end, to determine the fault occuring in the system, weighted sumsquared bias estimate – WSSBE fault detection algortihm is used. This algorithm uses the statistical variables that are based on bias control effectiveness factor estimates. The ratio of the square of the bias estimate to its covariance matrix is summed in a predetermined window length which corresponds to an iteration period. The resultant value should be between determined theoretical values if there is no fault in the system. In the decisiongain update algorithm, convergence of the control effectiveness factors is important and mean value of the estimates can be used for this purpose. In the study, using the mentioned methods, control of heading and altitude in cases where actuator faults are present in aileron and elevators are realized. It is shown through simulations that the unmanned aerial vehicle can be effectively controlled using the active fault tolerant controller despite the decrease in actuator control effectiveness factors which corresponds to effectivity loss in actuator controls. Another method used in the thesis to design a fault tolerant controller is passive fault tolerant control method. In this method, the predesigned controller is relied upon in dealing with the faults occuring in the system. Thus, fault detection and isolation and controller reconfiguration are not needed in this scheme. Dynamic inversion technique is used together with robust integral of the signum of the error RISE method to design an asymptotic tracking passive fault tolerant controller that has the capability to cope with faults. The faults occuring in actuators are modelled as parametric uncertainties in control distribution matrix B. To build an asymptotic model following passive controller, control inputs that decrease the difference between model and system should be found. For this purpose, Lyapunov type functions are used and controller constants are determined as done in similar studies. In simulations, in longitudinal model, forward velocity, pitch rate and in lateral model, yaw rate and roll rate are the main states that are controlled. Using asymptotic tracking controller system that helps in maintaining control of mentioned states, an outer loop is also built which aranges heading and altitude changes by tuning reference model inputs using fedback state values from the main system. Simulations done using both longitudinal and lateral models show that the designed passive controller is effective in controlling the unmanned aerial vehicle at times when faults are present in actuators. Hybrid control method is also used in the study to build a fault tolerant controller. This method uses active and passive controllers at different times to achieve fault tolerant control. This way, at times when the fault detection and isolation algortihm based on twostage Kalman filter determines the fault but still needs time to find the severity of the fault, passive fault tolerant controller can be used and the system can be kept under control continously. Reconfiguration can later be done after the fault severity is determined. As passive and active controllers are shown to be effective in controlling the unmanned aerial vehicle, hybrid control can be used for controlling faulty plants continously. In simulations done using lateral model of the unmanned aerial vehicle,it is shown that the hybrid controller is successful in keeping the vehicle under control and tracking the heading inputs at times when actuator faults are present. To achieve this result, active and passive controllers are used at different times after fault occurrence. In conclusion, a fault tolerant controller is designed for the unmanned aerial vehicle and it is shown that it can be effectively used when actuator faults – actuator control effectiveness loss cases corresponding to the problems in actuators and/or sensor faults are present in the study.

ÖgeDevelopment of a nonlinear sonic boom propagation code(Graduate School, 20230124) Demiroğlu, Yusuf ; Nikbay, Melike ; 511191144 ; Astronautics and Aeronautics EngineeringCivil supersonic flight is still one of the most challenging research topics in the aerospace industry. Since Concorde's last flight in 2003, researchers tried to find efficient solutions to make supersonic flights more affordable and reliable. Meanwhile, with the advance of computational power, computational fluid dynamics (CFD) has been implemented in advanced optimization studies involved in elevating supersonic aircraft design processes with given operational criteria and requirements. However, reducing the cost of a supersonic flight by increasing aerodynamic efficiency is not the only concern in civil supersonic transport. The second most important factor for a supersonic aircraft is the noise produced on land due to the shock waves that propagate through the atmosphere to the ground. This phenomenon is called sonic boom which is addressed in this thesis study. A sonic boom generated by a supersonic aircraft can cause very loud noise on the ground that may exceed 100 decibels. This loudness value is not acceptable due to its effects on people's daily life. Therefore, to enable civil supersonic flight over land, sonic boom loudness must be eliminated or reduced below a certain level. This effort is called sonic boom minimization and there are several methodologies that are provided in this study. Lots of studies for sonic boom minimization utilize optimization algorithms that call sonic boom prediction tools along with the CFD solvers. Therefore, to reduce sonic boom loudness, a sonic boom propagation code that accurately predicts sonic boom loudness is essential for the multidisciplinary design optimization of civil supersonic aircraft. In this regard, a new nonlinear sonic boom prediction code, named ITUBOOM, is developed inhouse to be incorporated into our design optimization studies to achieve a lowboom aircraft geometry. ITUBOOM is developed in Python programming language for ease of implementation for design studies. A sonic boom calculation process can be broken down into three main steps; a nearfield solution with CFD to generate an initial acoustic signal, atmospheric propagation with acoustics methods, and loudness calculation. Unlike other sonic boom codes, ITUBOOM can also be used to generate a nearfield pressure directly from CFD outputs by surface slicing or inflow signature extraction. Then, it can be used to perform atmospheric propagation by taking into account nonlinear effects such as molecular relaxation and thermoviscous attenuation. Results of ITUBOOM are validated against NASA Langley Research Center's wellknown sBOOM code for different conditions on benchmark problems and presented in this thesis in detail.

ÖgeDevelopment of singleframe methods aided kalmantype filtering algorithms for attitude estimation of nanosatellites(Graduate School, 20210820) Çilden Güler, Demet ; Hacızade, Cengiz ; Kaymaz, Zerefşan ; 511162104 ; Aeronautics and Astronautics Engineering ; Uçak ve Uzay MühendisliğiThere is a growing demand for the development of highly accurate attitude estimation algorithms even for small satellite e.g. nanosatellites with attitude sensors that are typically cheap, simple, and light because, in order to control the orientation of a satellite or its instrument, it is important to estimate the attitude accurately. Here, the estimation is especially important in nanosatellites, whose sensors are usually lowcost and have higher noise levels than highend sensors. The algorithms should also be able to run on systems with very restricted computer power. One of the aims of the thesis is to develop attitude estimation filters that improve the estimation accuracy while not increasing the computational burden too much. For this purpose, Kalman filter extensions are examined for attitude estimation with a 3axis magnetometer and sun sensor measurements. In the first part of this research, the performance of the developed extensions for the state of art attitude estimation filters is evaluated by taking into consideration both accuracy and computational complexity. Here, singleframe methodaided attitude estimation algorithms are introduced. As the singleframe method, singular value decomposition (SVD) is used that aided extended Kalman filter (EKF) and unscented Kalman filter (UKF) for nanosatellite's attitude estimation. The development of the system model of the filter, and the measurement models of the sun sensors and the magnetometers, which are used to generate vector observations is presented. Vector observations are used in SVD for satellite attitude determination purposes. In the presented method, filtering stage inputs are coming from SVD as the linear measurements of attitude and their error covariance relations. In this step, UD is also introduced for EKF that factorizes the attitude angles error covariance with forming the measurements in order to obtain the appropriate inputs for the filtering stage. The necessity of the substep, called UD factorization on the measurement covariance is discussed. The accuracy of the estimation results of the SVDaided EKF with and without UD factorization is compared for the estimation performance. Then, a case including an eclipse period is considered and possible switching rules are discussed especially for the eclipse period, when the sun sensor measurements are not available. There are also other attitude estimation algorithms that have strengths in coping well with nonlinear problems or working well with heavytailed noise. Therefore, different types of filters are also tested to see what kind of filter provides the largest improvements in the estimation accuracy. Kalmantype filter extensions correspond to different ways of approximating the models. In that sense, a filter takes the nonGaussianity into account and updates the measurement noise covariance whereas another one minimizes the nonlinearity. Various other algorithms can be used for adapting the Kalman filter by scaling or updating the covariance of the filter. The filtering extensions are developed so that each of them is designed to mitigate different types of error sources for the Kalman filter that is used as the baseline. The distribution of the magnetometer noises for a better model is also investigated using sensor flight data. The filters are tested for the measurement noise with the best fitting distribution. The responses of the filters are performed under different operation modes such as nominal mode, recovery from incorrect initial state, short and longterm sensor faults. Another aspect of the thesis is to investigate two major environmental disturbances on the spacecraft close enough to a planet: the external magnetic field and the planet's albedo. As magnetometers and sun sensors are widely used attitude sensors, external magnetic field and albedo models have an important role in the accuracy of the attitude estimation. The magnetometers implemented on a spacecraft measure the internal geomagnetic field sources caused by the planet's dynamo and crust as well as the external sources such as solar wind and interplanetary magnetic field. However, the models that include only the internal field are frequently used, which might remain incapable when geomagnetic activities occur causing an error in the magnetic field model in comparison with the sensor measurements. Here, the external field variations caused by the solar wind, magnetic storms, and magnetospheric substorms are generally treated as bias on the measurements and removed from the measurements by estimating them in the augmented states. The measurement, in this case, diverges from the real case after the elimination. Another approach can be proposed to consider the external field in the model and not treat it as an error source. In this way, the model can represent the magnetic field closer to reality. If a magnetic field model used for the spacecraft attitude control does not consider the external fields, it can misevaluate that there is more noise on the sensor, while the variations are caused by a physical phenomenon (e.g. a magnetospheric substorm event), and not the sensor itself. Different geomagnetic field models are compared to study the errors resulting from the representation of magnetic fields that affect the satellite attitude determination system. For this purpose, we used magnetometer data from low Earthorbiting spacecraft and the geomagnetic models, IGRF and T89 to study the differences between the magnetic field components, strength, and the angle between the predicted and observed vector magnetic fields. The comparisons are made during geomagnetically active and quiet days to see the effects of the geomagnetic storms and substorms on the predicted and observed magnetic fields and angles. The angles, in turn, are used to estimate the spacecraft attitude, and hence, the differences between model and observations as well as between two models become important to determine and reduce the errors associated with the models under different space environment conditions. It is shown that the models differ from the observations even during the geomagnetically quiet times but the associated errors during the geomagnetically active times increase more. It is found that the T89 model gives closer predictions to the observations, especially during active times and the errors are smaller compared to the IGRF model. The magnitude of the error in the angle under both environmental conditions is found to be less than 1 degree. The effects of magnetic disturbances resulting from geospace storms on the satellite attitudes estimated by EKF are also examined. The increasing levels of geomagnetic activity affect geomagnetic field vectors predicted by IGRF and T89 models. Various sensor combinations including magnetometer, gyroscope, and sun sensor are evaluated for magnetically quiet and active times. Errors are calculated for estimated attitude angles and differences are discussed. This portion of the study emphasizes the importance of environmental factors on the satellite attitude determination systems. Since the sun sensors are frequently used in both planetorbiting satellites and interplanetary spacecraft missions in the solar system, a spacecraft close enough to the sun and a planet is also considered. The spacecraft receives electromagnetic radiation of direct solar flux, reflected radiation namely albedo, and emitted radiation of that planet. The albedo is the fraction of sunlight incident and reflected light from the planet. Spacecraft can be exposed to albedo when it sees the sunlit part of the planet. The albedo values vary depending on the seasonal, geographical, diurnal changes as well as the cloud coverage. The sun sensor not only measures the light from the sun but also the albedo of the planet. So, a planet's albedo interference can cause anomalous sun sensor readings. This can be eliminated by filtering the sun sensors to be insensitive to albedo. However, in most of the nanosatellites, coarse sun sensors are used and they are sensitive to albedo. Besides, some critical components and spacecraft systems e.g. optical sensors, thermal and power subsystems have to take the light reflectance into account. This makes the albedo estimations a significant factor in their analysis as well. Therefore, in this research, the purpose is to estimate the planet's albedo using a simple model with less parameter dependency than any albedo models and to estimate the attitude by comprising the corrected sun sensor measurements. A threeaxis attitude estimation scheme is presented using a set of Earth's albedo interfered coarse sun sensors (CSSs), which are inexpensive, small in size, and light in power consumption. For modeling the interference, a twostage albedo estimation algorithm based on an autoregressive (AR) model is proposed. The algorithm does not require any data such as albedo coefficients, spacecraft position, sky condition, or ground coverage, other than albedo measurements. The results are compared with different albedo models based on the reference conditions. The models are obtained using either a datadriven or estimated approach. The proposed estimated albedo is fed to the CSS measurements for correction. The corrected CSS measurements are processed under various estimation techniques with different sensor configurations. The relative performance of the attitude estimation schemes when using different albedo models is examined. In summary, the effects of two main space environment disturbances on the satellite's attitude estimation are studied with a comprehensive analysis with different types of spacecraft trajectories under various environmental conditions. The performance analyses are expected to be of interest to the aerospace community as they can be reproducible for the applications of spacecraft systems or aerial vehicles.