LEE- Uçak ve Uzay Mühendisliği Lisansüstü Programı

Bu topluluk için Kalıcı Uri

http://hdl.handle.net/11527/19473

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A feedback star identification algorithm via regularized pattern recognition using a unique feature extraction

(Graduate School, 2024-07-11) Özyurt, Erdem Onur ; Aslan, Alim Rüstem ; 511172118 ; Aeronautics and Astronautics Engineering

This thesis presents a star identification algorithm integrated with preprocessing. Star sensors, which are highly reliable for attitude determination use of spacecrafts and satellites, relies on star identification algorithms. The star identification algorithm proposed in this study is capable of functioning either in lost-in-space method or recursive method. Both methods utilize a unique feature extraction scheme. This novel approach of feature extraction method extracts a single vector from each captured image instead of treating each star as a separate object. This cumulative approach aims to save a significant amount of memory space while taking the entire catalog into account for elevated accuracy. A database containing stars from the catalog is constructed using the unconventional features extracted from each corresponding field-of-view. The databases may differ in size and detail dependent on the parameters of overlapping ratio and brightness threshold. These parameters have a significant effect on accuracy and complexity of the method. The method aims to estimate the inertial boresight vector and the rotation angle about it. This is a novel approach that is carried out by matching frames but not matching individual stars, star pairs, star triangles or star polygons. Both star identification methods rely on pattern recognition and regularization successively. First, a 1NN classifier is used to perform a coarse estimation with limited accuracy specified by the characteristics of the database with predetermined parameters. The coarse estimation is exactly the database vector that is most similar to the observation vector. Subsequently, a dictionary is generated using the neighbor database vectors of the most similar database vector. The final estimation is obtained by conducting a regularization method for fine estimation. A solution coefficient vector is yielded through regularization. The estimates of boresight vector and rotation angle are retrieved using the solution coefficient vector. This is the output of the lost-in-space star identification method. Since the lost-in-space algorithm is very sensitive to false stars, an additional false star filtering algorithm is developed. This algorithm is based on density-based clustering. A disparity list is created using two successive image frames. After estimating true stars by implementing density-based clustering on the disparity list, false stars are removed. Using the successive frames containing only estimated true stars, an affine transformation matrix is obtained by a regression analysis procedure. In order to overcome the issues tackling the lost-in-space star method, the recursive star identification method is developed. Apart from the algorithmic structure taken from the lost-in-space method, it possesses an update mechanism that ensures usage a much smaller portion of the database to reduce computational complexity and average run time. Also, the false star filtering avoids sensitivity to false stars.
A high-order finite-volume solver for supersonic flows

(Lisansüstü Eğitim Enstitüsü, 2022) Spinelli, Gregoria Gerardo ; Çelik, Bayram ; 721738 ; Uçak ve Uzay Mühendisliği

Nowadays, 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 open-source codes, however, offer a second-order 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 high-order discretization in both space and time by developing a two-dimensional finite-volume 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. High-order approximation in space is guaranteed by a Central Essentially Non-Oscillatory (CENO) scheme, which combines a piece-wise linear reconstruction and a k-exact reconstruction in region with and without discontinuities, respectively. The usage of multi-step methods such as Runge-Kutta methods allow for a high-order 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 high-order schemes, discontinuities induce numerical problems, such as oscillations in the solution. To avoid the oscillations, the CENO scheme reverts to a piece-wise 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 high-order 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 two-dimensional 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 two-dimensional Riemann problem. We obtained a linear speed-up 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 Non-Oscillatory (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 high-order solvers because the shock and its reflections interact strongly. For this application, all AUSM-schemes 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 vortex-let is observed along the vortex slip-line. 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 0-30. 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 second-order 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 high-order reconstruction across shocks, and therefore will not apply the piece-wise 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 first-order 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 high-order 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.
A modified anfis system for aerial vehicles control

(Lisansüstü Eğitim Enstitüsü, 2022) Öztürk, Muhammet ; Özkol, İbrahim ; 713564 ; Uçak ve Uzay Mühendisliği

This thesis presents fuzzy logic systems (FLS) and their control applications in aerial vehicles. In this context, firstly, type-1 fuzzy logic systems and secondly type-2 fuzzy logic systems are examined. Adaptive Neuro-Fuzzy Inference System (ANFIS) training models are examined and new type-1 and type-2 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 neuro-fuzzy controller is used as online controller. Neuro-fuzzy 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. Neuro-fuzzy 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 neuro-fuzzy control systems in the literature. Each approach is tested on a DC motor model that is a single-input and single-output system, and the neuro-fuzzy 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 neuro-fuzzy controllers are used in quadrotor control. Quadrotors have a two-stage 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 neuro-fuzzy 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 neuro-fuzzy controller can not give the desired performance in the two-layered 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 type-1 fuzzy logic systems consist of certain membership functions. A type-2 FLS is proposed to better define the uncertainties, because of that, type-2 fuzzy inference membership functions are proposed to include uncertainties. The type-2 FLS is operationally difficult because of uncertainties. In order to simplify type-2 FLS operations, interval type-2 FLS is proposed as a special case of generalized type-2 FLS in the literature. Interval type-2 membership functions are designed as a two-dimensional projection of general type-2 membership functions and represent the area between two type-1 membership functions. The area between these two type-1 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 type-2 and it is not possible to perform the defuzzification step directly because of uncertainty. Therefore, type reduction methods have been developed to reduce the type-2 FLS to the type-1 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 iterative-free 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 type-2 FLS still continue. These studies are also used in the interval type-2 fuzzy logic control systems. There are proposed interval type-2 ANFIS structures in literature, but they are not effective because of uncertainties of interval type-2 membership functions. FLS parameters for ANFIS training should not contain uncertainties. However, the type-2 FLS should inherently contain uncertainty. For this reason, Karnik-Mendel algorithm is modified, which is one of the type-reduction methods, to apply the ANFIS on interval type-2 FLS. The modified Karnik-Mendel algorithm gives the same results as the Karnik-Mendel algorithm. The modified Karnik-Mendel algorithm also gives exact parameter values for use in ANFIS. One can notice that the ANFIS training of the interval type-2 FLS has been developed successfully and has been used for system control.
A multi-disciplinary design approach for conceptual sizing of advanced rotor blades

(Lisansüstü Eğitim Enstitüsü, 2022-07-19) İbaçoğlu, Hasan ; Arıkoğlu, Aytaç ; 511072102 ; Aeronautics and Astronautics Engineering

Rotorcrafts 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, vehicle-level optimization algorithms are needed for decision-making to lead the project correctly. On the other hand, simplified low-level 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 noise-level 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 noise-level 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. Low-level 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 in-house 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 well-known 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. Cross-sectional properties of the composite beam are calculated by VABS and displacements under the loads are calculated by GEBT. Reduced FfowcsWilliams-Hawkings equations are used to estimate loading, thickness, and high-speed 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. RSM-based surrogate modeling, evaluation, and optimization tool was also developed for manual inspection of the design space. As a case study, multi-objective aerodynamic performance optimization of aircraft propeller is performed.
A multiscale approach to understand the effects of design parameters on the elastic behavior of 3D orthogonally woven composites

(Graduate School, 2024-11-04) Erkoç, Hilal ; Cebeci, Hülya ; 511201167 ; Aeronautical and Astronautical Engineering

This study aims to investigate the effect of various parameters on the elastic constants of three-dimensional (3D) orthogonally woven composites. Two-dimensional (2D) laminated composites exhibit high in-plane stiffness and strength; however, they are inadequate in applications subjected to out-of-plane loads, particularly in engine fan blades, aircraft fuselage structures, and wind turbine blades. With an innovative approach, 3D orthogonally woven composites effectively overcome the limitations of traditional 2D laminates. The usage of 3D orthogonally woven composites in these structures can be beneficial because 3D orthogonally woven composites are more resistant to out-of-plane loading than 2D laminates, due to their improved mechanical properties through the thickness. In addition to this, improved impact damage tolerance, higher delamination resistance, and reduced assembly and production costs through single-piece fabric production are advantages of 3D orthogonally woven composites. 3D orthogonally woven composites, in spite of their advantages, present certain challenges in application. One of the significant challenges is the complex nature of their manufacturing process, which demands specialized equipment and skilled personnel, leading to high production costs. Their complex structure can also complicate design, analysis, and simulation, requiring advanced computational models. Additionally, the complex architecture of these composites can present challenges in repair and maintenance procedures. 3D orthogonally woven structures consist of three interwoven sets of yarns arranged in orthogonal directions, where the warp and weft yarns remain straight while the binder yarns interlace them to create a multidimensional architecture. This complex architecture of 3D orthogonally woven composites plays an important role in determining the mechanical properties of the structure. Since differences in cross-section configurations, yarn arrangements, and fiber interactions significantly influence the load-carrying capacity, stiffness, and overall performance of the composite, an in-depth examination of the structural architecture is critical to optimizing the mechanical properties of the material. Several analytical studies have examined the effects of binder-to-weft and binder-to-warp ratios on the elastic properties of 3D orthogonally woven composites. These analyses employ representative volume elements (RVEs) to model the material behaviors. The binder-to-weft ratio characterizes the number of wefts of yarn a binder yarn encircles before reversing direction within the weft layer. Similarly, the binder-to-warp ratio represents the proportion of warp yarns per layer relative to the total number of warps encompassed by the RVE. However, a key limitation of these existing studies is based on the absence of a comparative analysis between analytical solutions and numerical simulations. Furthermore, the impact of RVE thickness on its elastic coefficients has not been thoroughly investigated. Here, the effects of changing thickness on the tensile response of the structure, as obtained through analytical solutions and numerical simulations, are presented. Elastic constants of 3D fiber-reinforced composites were estimated using a multi-scale homogenization technique based on meso-macro homogenization with good correlation. Numerical simulations were performed using ABAQUS software to analyze the behavior of the models. Through the optimization of the geometrical parameters of RVE, 3D orthogonally woven composites can be effectively implemented across a diverse range of engineering applications, especially in the aviation field.
A numerical approach for plasma based flow control

(Graduate School, 2023-04-05) Ata, Reşit Kayhan ; Şahin, Mehmet ; 511132114 ; Aeronautics and Astronautics Engineering

In the present study, a novel numerical method has been developed to solve incompressible magnetohydrodynamics (MHD) and electrohydrodynamics (EHD) flow problems in a parallel monolithic (fully-coupled) approach. To solve the fluid flow, incompressible Navier-Stokes 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 side-centered 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 pressure-velocity 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 Navier-Stokes, 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 two-neighbouring elements which then resulted in improved mass conservation over the momentum control volume and increased stability. The second modification applies to only two-dimensional 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 two-dimensional Hartmann flow of which the analytical solution is known. Then lid-driven cavity and backward facing step problems are investigated under external magnetic field both in 2D and 3D with insulating walls. Three-dimensional MHD flow in ducts is another case where analytic solutions exist. Both conducting and insulating wall boundary conditions are employed and validated. Finally two-dimensional 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 Navier-Stokes equation. The side-centered 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.
A numerical investigation of total temperature probes measurement performance

(Graduate School, 2023) Meriç, Erdem ; Edis, Fırat Oğuz ; 810043 ; Aeronautics and Astronautics Engineering Program

In almost every industrial application, the temperature is measured for development and condition monitoring purposes. The accuracy of these measurements is crucial to avoid misunderstandings about the current condition and misguidance in the development phase. The most practical mean of temperature measurement in industrial applications is using a thermocouple. Thermocouples are very flexible structures so they can be applied in many different regions for solid and fluid temperature measurements. It is also possible to design measurement probe geometries using thermocouples as sensing elements. In machines involving high-speed gas flow, the kinetic energy of fluid can't be neglected in energy interaction calculations so flow must be adiabatically stagnated before temperature measurement. The temperature a flowing fluid gains because of adiabatic stagnation is called stagnation or total temperature. A stationary probe geometry measures the total temperature of flow but there may be deviations in the temperature of the sensing point due to the flow physics. These deviations lead to errors in measurement. These errors are classified as recovery error, conduction error and radiation error. Recovery error originated from the non-adiabatic stagnation of flow on the surface of the thermocouple (TC) junction. Recovery error is characterized by a parameter called recovery factor which shows the degree of dynamic temperature recovery on the measurement. Conduction and radiation errors arise due to solid boundary conditions which are different from the flow total temperature around the probe. These different temperature zones cause heat interaction via conduction and radiation heat transfer modes between the TC junction and surroundings giving rise to deviations in measurement. Special probe designs are used to prevent these errors. In this study, an experimental case was selected from the literature to create a conjugate heat transfer (CHT) methodology. This CHT methodology served to investigate flow physics around and inside total temperature probes and the nature of heat interaction between flow and probe geometry. This experimental case contains a total temperature probe calibration setup which investigates the measurement performance of probe geometry under different Mach number flows. In the simulations, the measurement probe geometry was modelled and exposed to the flow at the same speed as the test conditions. The main observed parameter during simulations was TC junction temperature which determines the performance of the total temperature probe. The results of simulations were observed to be in harmony with experimental data. Then, flow structures around and inside the total temperature probe were investigated in detail using the outputs of simulations. The main aim of total temperature probe geometry is to decrease flow velocity inside the shield to decrease thermal conduction in the boundary layer. In simulations, this aim was observed to be accomplished. The flow velocity vectors were investigated to understand the nature of flow around and inside the total temperature probe. No flow separation was observed on the shield inlet.
A study on optimization of a wing with fuel sloshing effects

(Graduate School, 2022-01-24) Vergün, Tolga ; Doğan, Vedat Ziya ; 511181206 ; Aeronautics and Astronautics Engineering ; Uçak ve Uzay Mühendisliği

In 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 airfoil-shaped 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 1-Way Fluid-Structure Interaction (1-Way 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 2-Way Fluid-Structure Interaction (2-Way 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, 1-Way FSI is used since the computational cost of the 2-Way FSI method is approximately 35 times that of 1-Way FSI. However, taking less time should not be enough to accept 1-Way 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, 1-Way 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 non-collapsing and space-filling points to create a better response surface. After creating the initial response surface using Genetic Aggregation, the optimization process is started using the Multi-Objective 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.
A study on static and dynamic buckling analysis of thin walled composite cylindrical shells

(Graduate School, 2022-01-24) Özgen, Cansu ; Doğan, Vedat Ziya ; 511171148 ; Aeronautics and Astronautics Engineering ; Uçak ve Uzay Mühendisliği

Thin-walled 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 thin-walled structures is that they have a high strength weight ratio. In order to define a cylinder as thin-walled, 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 thin-walled 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 post-buckling 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 eigenvalue-eigenvector 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 Newton-Rapson 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 quasi-static 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 displacement-controlled analyses were carried out in the static buckling analyses mentioned, load-controlled 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, Budiansky-Roth criterion, Hoff-Bruce criterion, etc. When Budiansky-Roth 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 Hoff-Bruce 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 length-to-diameter 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.
Advanced energy and exergy analysis on aircraft jet engines

(Graduate School, 2023-12-08) Fawal, Sara ; Kodal, Ali ; 511212113 ; Aeronautics and Astronautics Engineering

A 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), Turbo-Ramjet. Newly defined parameters are introduced as power loss parameter (PLOS), effective power loss parameter (EPLOS) and Carnot-Brayton 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 Turbo-Ramjet) 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.
Advanced visual odometry and depth estimation techniques for unmanned aerial systems (UAS) in U-Space environments

(Graduate School, 2024-12-11) Roghani Seyed, Seyed Erfan ; Koyuncu, Emre ; 511182113 ; Aeronautical and Astronautical Engineering

This thesis explores advanced techniques in visual odometry (VO) and depth estimation for Unmanned Aerial Systems (UAS), specifically within the context of U-Space environments. U-Space, as a European initiative, aims to ensure the safe, efficient, and secure integration of UAS into airspace. This work contributes to this goal by addressing two critical aspects of UAS navigation: precise visual odometry and reliable depth estimation. Chapter 1 - Introduction: The introduction presents the context of U-Space, outlining its evolution and the services it offers, with a focus on emergency management. The challenges of autonomous contingency planning in UAS operations are highlighted, particularly in relation to visual odometry and depth estimation. Chapter 2 - Canonical Trinocular Feature-Based Visual Odometry: This chapter proposes a novel trinocular camera configuration to enhance VO for UAS. The research compares two trinocular setups—inline and 45-degree—with traditional binocular setups, testing them in various scenarios (horizontal, vertical, hybrid, and long). The results demonstrate that the 45-degree trinocular configuration with a standard lens offers significant improvements in both accuracy and computational efficiency, reducing the computational effort to 40\% of that required by binocular systems while delivering more accurate results. However, when a fisheye lens is used, the benefits are less pronounced, particularly in vertical and long scenarios. Chapter 3 - Fine-Tuning Monocular Depth-Estimator Artificial Neural Networks Trained on Synthetic RGB-D Datasets for Real Scenes: This chapter addresses the challenge of depth estimation for UAS using monocular cameras, which are cost-effective but typically less reliable than stereo cameras. The research investigates the effectiveness of fine-tuning deep-learning models trained on synthetic data with small real-world datasets. The results show that complete fine-tuning of all model parameters, as opposed to just the decoder, yields the best performance, especially when the available real data is limited to less than 12.5\% of the data required for optimal model performance. This finding is crucial for applications where only limited real-world data is available. Conclusion: The thesis concludes that the proposed trinocular VO configuration significantly enhances the accuracy and efficiency of UAS navigation, particularly in complex U-Space environments. Additionally, it establishes the importance of fine-tuning depth estimation models with real-world data, even when such data is scarce, to improve the reliability of UAS in operational scenarios. These advancements contribute to the broader goal of integrating UAS into airspace, ensuring they can operate safely and effectively under various conditions.
Aeroacoustic investigations for a refrigerator air duct and flow systems

(Graduate School, 2022-02-16) Demir, Hazal Berfin ; Çelik, Bayram ; 511181186 ; Aeronautics and Astronautics Engineering

Noise 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 three-dimensional, 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 (FW-H) 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 10-10000 Hz is compared, it is seen that there is a difference of 3-7 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.
Aerodynamic and structural optimization a male class unmanned aerial vehicle wing with genetic algorithm

(Graduate School, 2023) Ün, Kağan ; Yıldız, Kaan ; 511191209 ; Aeronautical and Astronautical Engineering Programme

In this thesis, a genetic algorithm based airfoil, planform and wing structure design is utilized for a 1500-2000 kg class fixed wing reconnaissance MALE (Medium Altitude Long Endurance) type UAV (Unmanned Aerial Vehicle). Due to their mission descriptions, these UAV are generally designed to have long range and high loiter time. For these reasons, an aircraft to be designed for observation purposes must have high aerodynamic efficiency and high fuel load ratios (low empty mass) to maximize the range and loiter time. To achieve high aerodynamic efficiency, these aircraft have high wing span ratios, and their wings contain specially designed airfoils. These airfoils are generally designed to have high lift-to-drag ratios. On the other hand, due to long wing structures of high-span aircraft, the wings of such aircraft have high bending forces especially during maneuvers. Therefore, design of airfoils of such aircraft have a compromise between fitting the ideal lightweight inner skeleton for the wing and providing ideal aerodynamics. For these reasons, the wing profile located at the root of the wing is generally chosen to be thicker than the profile used at the tips of the same wing. To simplify the design process, each airfoil are constructed from two Bezier curves that create centreline and thickness distribution of airfoils. Control points of the Bezier curves are the most of the input parameters of the genetic algorithm program. From sets of control points, the airfoils are created. Then, the airfoils are analysed in XFoil by the interface of the function made in MATLAB. After analysis of airfoils, a planform that uses these root-tip airfoils is tested for having sufficiently high lift and low drag for the cruise altitude, cruise speed and cruise power. Then the airfoilplanform combination that pass the basic requirements are sorted by their maximum lift-to-drag ratios. The airfoil-planform combinations with higher maximum lift-todrag ratios are selected for creating the next generation, and the cycle continues. When the maximum number of generations are achieved, the best airfoil-planform combination of the last generation is selected as the best candidate. The fitness criterion of this first phase is the lift-to-drag ratio of the airfoil-planform combination. After the winner airfoil-planform combination is created, inner structure optimization process for the wing begins. Inner structure of the planform consists of four ribs and a twin-box spar structure made of 7068 aluminium alloy, the strongest commercial aluminium alloy available. The lift force and torsion moment of each wing segment is transferred to the spar by the ribs of the wing. The cross section of the spar consists of two closed cells with the support of four stiffeners and eight flanges. Vertical walls have thickness of 2.5 mm, while upper and lower walls have thickness of 1.5 mm. The flanges have cross section value of 400 mm2 , and are set to the upper and lower ends of vertical walls, filling the corners of each cells. The stiffeners have cross section value of 200 mm2 , and are set to the middle of the upper and lower walls, in between the vertical walls. While the stiffeners and flanges carry the tensile and compressive loads, the walls primarily carry the shear loads. To simplify the structural analysis, several assumptions are made. The main spar is assumed as a serial combination of smaller spar sections with constant cross sections. The stresses on cross sections are analysed with structural idealization method. The cross section is assumed as collections of idealized shear force carrying panels and normal force carrying area members called booms. At first, the effective areas at the positions of booms are found by adding effective boom area of wall sections to the actual stiffener or flange areas. From the effective areas and positions of each boom, area moments of inertia and bending moment centre are found. From area moments of inertia of the section and applied bending moment, the compressive and tensile forces of each boom are calculated. After the calculation of tensile and compressive forces, shear forces on the walls are calculated from the area and wall thicknesses of each cell, torsion moment inflicted on section, and compressive and tensile forces of each boom. After the stress calculations are made for each section, a selection process is carried out to control the stresses on the cross sections on each wing rib. If the stresses on the wing at any point is larger than safe limits, the spar is considered as infeasible specimen. Otherwise the fitness value for the spar is compared to the other successful specimens and the fittest specimen is chosen for each generation to create new specimens. The weight of the spar is the fitness value for the second phase. At the end of the process, the ideal cross section is obtained and the program finishes working.
Aktif kontrol uygulamalarında firar kenarı emiş yüzeyi manipülasyonunun akış gürültüsü üzerindeki etkilerinin incelenmesi

(Lisansüstü Eğitim Enstitüsü, 2024-02-05) Toksavul, Atila ; Zafer, Baha ; 511191185 ; Uçak ve Uzay Mühendisliği

Hava araçları ve rüzgâr türbinlerinin aerodinamik ve aeroakustik özelliklerinin iyileştirilmesi, performansları ve işleyişleri açısından önemli bir yer tutar. Havacılık alanında gerçekleşen gelişmeler ile birlikte, uçan yolcu sayısı her geçen gün artmaktadır. Benzer bir şekilde; hızla artan şehirleşme, beraberinde yüksek enerji ihtiyacını getirmektedir. Dolayısıyla, rüzgâr çiftlikleri gibi alternatif enerji kaynaklarının kullanımı her geçen gün artmaktadır. Bu sebeple; gürültü kontrolü ve gürültünün önlenmesi, günümüzde giderek önem kazanmaktadır. Bu çalışma kapsamında; aktif akış kontrol tekniklerinin kullanımının, bir kanadın aerodinamik ve aeroakustik performansı üzerindeki etkileri incelenmektedir. Aktif kontrol metodu olarak, firar kenarına yakın bölgelerde emme akışı uygulamasına gidilmiştir. Firar kenarı gürültüsünün azaltılmasında, bahsedilen aktif akış kontrol metodunun etkinliği araştırılmıştır. Bu doğrultuda, oluşturulan farklı senaryolar için Hesaplamalı Akışkanlar Dinamiği (HAD) analizleri gerçekleştirilmiştir. Sayısal verilerin alınması sürecinde, literatürde yaygın olarak kullanılması ve doğrulama verilerine ulaşılabilmesi nedeniyle, bir NACA (National Advisory Committee for Aeronautics) kanat profili olan ve körlenmiş firar kenarına sahip NACA 0012 profili kullanılmıştır. Kullanılan NACA 0012 profili için veter uzunluğu 0.2m, firar kenarı körlüğü (h/δ*) ise 0.196 olarak alınmıştır. Simülasyonlar, ANSYS Fluent yazılımı üzerinde Reynolds-Averaged Navier–Stokes (RANS) ile Large Eddy Simulation (LES) karma modeli olan Stress-Blended Eddy Simulation (SBES) türbülans modeli kullanılarak gerçekleştirilmiştir. Türbülans alt modeli (subgrid-scale) olarak, Smagorinsky-Lilly modeli kullanılmıştır. Basınç- hız bağıntısının sağlanması için Pressure-Implicit with Splitting of Operators (PISO) algoritması kullanılmıştır. Öncelikle, çalışmada kullanılacak HAD analizleri için oluşturulan sayısal modelin geçerliliği teyit edilmiştir. Bu süreçte, doğrulama analizleri gerçekleştirilmiştir. Gerçekleştirilen doğrulama analizleri sırasında herhangi bir aktif kontrol metodu kullanılmamıştır. Fakat, türbülansa bağlı etkilerin daha iyi gözlemlenebilmesi için, NACA 0012 profili üzerinde hücum kenarına yakın bölgeler üzerinde tetikleme akışı uygulanmıştır. Tetikleme akışı için kullanılan değerler, literatürde yer alan akış verilerini baz almaktadır ve deneysel sonuçlara yakınsamak için sayısal analizlerde kullanılmak üzere oluşturulmuştur. Sayısal modelin doğrulanması, NACA 0012 profili üzerinden elde edilen basınç değerleri ve akustik verilerin literatür ile karşılaştırılması ile sağlanmıştır. Bu süreçte, veter uzunluğuna göre Reynolds sayısı 400000 olarak alınmıştır. Profil üzerinden elde edilen negatif basınç katsayıları, literatür ile karşılaştırılmıştır. Akustik veriler ise, firar kenarına yakın konumlar üzerinden hesaplanarak referans alınan değerler ile karşılaştırılmıştır. Elde edilen sonuçlar neticesinde, kullanılan sınır şartlarının ve oluşturulan sayısal modelin geçerliliği sağlanmıştır. Doğrulama analizleri sonrasında, firar kenarını referans alan doğrultusal gürültü dağılımları, hesaplanmıştır. Elde edilen sonuçlar üzerinde gürültünün kuadrupol etkileri yeterince gözlemlenemediğinden dolayı serbest akış hızı arttırılmıştır. Veter uzunluğuna göre 530000 Reynolds sayısına sahip akış üzerinde aktif gürültü kontrol metotları uygulanmıştır. Farklı aktif kontrol parametrelerinin profil üzerinde uygulandığı senaryolar oluşturulmuş, bulunan sonuçlar birbiri ile karşılaştırılmıştır. Senaryolar, literatürde yer alan aktif gürültü kontrol parametresi baz alınarak oluşturulmuştur. Parametre, aktif kontrol yüzeyinin uzunluğu ve kontrol akışının hızı ile doğru, serbest akış hızı ile ters orantılıdır. Dolayısıyla; serbest akış koşulları değişmediği sürece, birbirine denk kontrol akışı debilerine sahip senaryoların sahip olduğu kontrol parametreleri de birbirine eşit olacaktır. Çalışmada, kontrol yüzeyinin manipüle edilmesi ile oluşturulan farklı senaryolar için gürültü değerleri incelenmiş, birbirine denk kontrol parametrelerine sahip senaryolar üzerinden elde edilen sonuçlar karşılaştırılmıştır. HAD analizleri için oluşturulan modellerde kontrol parametreleri; 0.268, 0.537, 1.611 ve 3.222 olarak belirlenmiştir. Her bir kontrol parametresi için, NACA 0012 profili yüzeyine, birbirine denk merkezli ancak farklı uzunluklarda kontrol yüzeylerine sahip üç farklı durumda aktif kontrol senaryosu tanımlanmıştır. Kontrol parametresinin 0.268 olarak tanımlandığı durumda oluşturulan tüm senaryolar için gerçekleştirilen analizlerde, firar kenarına göre 90° ve 270° doğrultularında gürültü iyileştirme gözlemlenmiştir. Kontrol yüzeyinin daraltıldığı senaryoda bu etkiler artmış, tüm açısal doğrultularda efektif bir gürültü iyileştirme elde edilmiştir. Kontrol parametresinin 0.537 olarak tanımlandığı durumda ise farklı olarak, kontrol yüzeyinin genişletildiği senaryoda etkin bir gürültü iyileştirme gözlemlenmiştir. Tüm doğrultularda efektif olarak gürültü iyileştirmenin gözlendiği senaryolarda; 0° doğrultusunda yani iz bölgesinde gözlenen gürültü miktarındaki azalmanın, diğer doğrultularda gözlenen değerlere göre daha düşük olduğu saptanmıştır. Kontrol parametrelerinin 1.611 ve 3.222 olarak belirlendiği senaryolar ise, elde edilen verilerin doğrulaması, dipol ve kuadrupol etkilerin daha iyi gözlenebilmesi için analiz edilmiştir. Kontrol parametresinin 1.611 olarak belirlendiği senaryolarda iz bölgesi gürültüsünün önlenmesinde etkili sonuçlar elde edilmiştir, ancak bu durum 90° ve 270° doğrultularında gürültü önleme performansında düşüş yaratmıştır. Kontrol parametresinin 3.222 olarak belirlendiği senaryolarda ise ölçüm alınan tüm doğrultular üzerinde gürültü önleme performansında düşüş gözlemlenmiştir.
An ALE framework for multiphase flows

(Graduate School, 2022-08-24) Güventürk, Çağatay ; Şahin, Mehmet ; 511162103 ; Aeronautical and Astronautical Engineering

An Arbitrary Lagrangian Eulerian (ALE) framework which combines the advantages of both Lagrangian and Eulerian methods is developed to solve incompressible multiphase flow problems. The div-stable side centered unstructured finite volume formulation is used for the discretization of the incompressible isothermal Navier-Stokes equations along with the isothermal constitutive equations for Oldroyd-B and FENE-CR fluids. In this approach, the velocity vector components are defined at the mid-point 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 ad-hoc 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 semi-implicit 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 liquid-liquid 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 cross-section almost fill the cross-sectional 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 three-dimensional 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 Oldroyd-B. 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 FENE-CR 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.
Analysis of aircraft landing gear brake induced vibrations

(Graduate School, 2023-01-23) Altınbağ, Öner ; Balkan, Demet ; 511191131 ; Aeronautics and Astronautics Engineering

Today, 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 brake-induced 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.
Analysis of bird strike on metallic panels

(Graduate School, 2023-06-15) Çayhan, Kenan ; Balkan, Demet ; 511201133 ; Aeronautics and Astronautics Engineering

This 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 wind-facing 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 real-world 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 computer-based 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 front-facing 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 fluid-like 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 LS-DYNA 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 time-consuming and costly, especially in the initial design phase of aerospace vehicles.
Analytical investigation of quasi-aeroservoelastic behaviour of an aircraft spoiler

(Graduate School, 2022) Kurtiş, Yiğit ; Mecitoğlu, Zahit ; Muğan, Ata ; 777765 ; Aeronautics and Astronautics Engineering Program

The application of science and mathematics to solve problems is called as engineering. In most of engineering process, accuracy is directly dependent on cost which can be defined as function of time and money. In the problem solving processes, there are a lot of assumptions in exchange for accuracy in order to reduce cost and find more solutions in a short time. Reducing solution time provides ability to enhance problem solving capabilities by increasing number of ways to solve problems, finding different sources of problems or optimizing solution methods. At the end, exact solution may not be reached, but more related problems can be solved with approximate solutions in limited time. With advanced technology in aviation industry, accurate designs are more important than before due to desire for better performance. In order to increase accuracy, research and development studies are performed, such as analytical formulizations and tests. Owing to high cost and long durations of test operations, analytical solutions are preferred to be supported if possible. Especially for aircraft design, due to safety consideration and aim for lightweight designs, designers have to balance time, weight and cost without any penalty for safety. In this condition, analytical solutions helps to reduce solution time for lightweight designs and create extra time for optimization studies. In this study, behavior of spoiler structures are investigated for desired deflection angle under external loads by means of analytical solutions. Spoiler is a control surface which can create drag and lift for aircraft. Spoiler structures have been implemented to aircrafts in order to improve control, especially while rolling, landing and braking. One of the main objectives of a spoiler structure is to increase drag for landing and braking applications. Additionally, spoilers can be used to increase roll rate for acrobatic or trainer aircrafts. Under aerodynamic load, as all structures deform, spoiler structures show a deformation. It affects spoiler deflection mechanism because points of mechanism changes when spoiler deformation occurs. In this case, spoiler rotates back towards to its original position where back rotation angle is usually not able to be considered in mechanism design. This condition creates dwindle for effectivity of spoiler surface which means reducing performance of aircraft. In this thesis, an analytical formulation study is performed in order to foresee back rotation angles of spoiler structures and gain ability to design mechanism for more convenient deflection angles for spoilers under aerodynamic loads. Result curves are created by curve fitting method in order to monitor and compare behavior of both analytical method analyses and finite element method analyses. Error functions are defined and calculated to find out tendency difference between analytical method and finite element method analyses under changing variables. For realistic deflection angles, the aim of this study is to accomplish accurate analytical results with error percentage below ±15% for back rotation angles and ±2% for final deflection angle compared to finite element method analyses. In the introduction section, engineering approaches for development studies are explained. Importance of accuracy for engineering application is tried to be stated by support of relations between accuracy and other engineering concerns. These concerns can be expressed with time, cost and other concerns, such as health issues, ethic concerns and safety. Also, scope and purpose of thesis are determined in this section. In the literature review section, spoiler structures and their duties on aircraft are stated. Dimensions are shown with examples and figures from aircraft industry. Grid stiffened spoiler concepts are explained in addition to commonly used structural architectures, such as composite and metal builtup structures.
Assessment and optimization of a hybrid rocket motor as the final stage of small satellite launch vehicle

(Graduate School, 2024-06-24) Aksen, Ukde ; Aslan, Alim Rüstem ; 511182115 ; Aeronautics and Astronautics Engineering

In the realm of space exploration, launch vehicles or carrier rockets transport payloads from Earth to space for tasks such as commercial or military satellite deployment, meteorological observations, and experimental studies. These vehicles range from single-stage to reusable systems, classified by launch platforms on land, sea, or air, and categorized by payload size for targeted orbits. Hybrid rocket motors, combining solid propellants with liquid or gaseous oxidizers, provide flexibility and efficiency in propulsion. This thesis develops a six-degrees-of-freedom model for simulating small satellite launch vehicle orbits, including the replacement of the final stage of the Minotaur-I launch vehicle with a hybrid rocket motor. Multi-objective optimization addresses complex design parameters and mission objectives, utilizing hybrid propulsion for cost-effectiveness and environmental benefits. Additionally, response surface analysis evaluates key parameters influencing vehicle performance and trajectory optimization.
Attitude estimation and reaction wheels based control of an earth-pointing small satellite

(Graduate School, 2024-06-24) Gürsoy, Hakan ; Hacızade, Cengiz ; 511211123 ; Aeronautics and Astronautics Engineering

A satellite is an artificial object that is sent into orbit around a celestial body, usually the Earth. Depending on their mission design, they can perform various tasks from communication to military. Satellites consist of various subsystems to perform their assigned mission. Attitude Determination and Control Systems (ADCS) is responsible for orienting the satellite in the desired direction and maintaining its orientation in space. It uses various sensors and actuators to measure and control the satellite's attitude. The sensors, such as sun sensors, magnetometers and gyroscopes, provide data about the satellite's current orientation. The measurement data coming from these sensors are processed by various attitude estimation methods. The actuators, like reaction wheels, magnetorquers and thrusters, apply the necessary torques to adjust the satellite's attitude. The ADCS ensures that the satellite's payload is correctly positioned for its mission. If the satellite cannot be brought to the required orientation, it may not be able to fulfil its mission. In this thesis, some of the prominent vector measurement-based attitude estimation methods are compared. To make this comparison, the dynamic and kinematic model of an Earth-pointing satellite is derived and subsequently, this highly nonlinear model is linearized by using the Taylor series method. Following this, mathematical models of the sun sensor and magnetometer were also presented. Real sensor measurements are simulated by adding white noise to these mathematical models. The compared attitude estimation methods can be named as TRIAD, Q-Method, and SVD. The comparison is made through conducting simulations in the Matlab/Simulink environment. Root mean square errors of the estimation methods are computed by comparing their outputs with the output of the deterministic satellite system model. After comparing the attitude estimation methods, a comparison was made between the two types of the optimal controllers. The LQR controller is a feedback control method that aims to minimize a quadratic cost function, which is typically defined in terms of deviations from desired states and control inputs. It provides an optimal solution for controlling the satellite's attitude while considering both system dynamics and control effort. The LQG controller improves the capabilities of LQR by adding a Kalman filter, improving the estimation of system states from noisy sensor measurements. This integration improves the controller's ability to handle uncertainties and disturbances, making it more suitable for real-world satellite applications where sensor data may be prone to noise or inaccuracies. Reaction wheels are selected as the actuators. Reaction wheels are based on the principle of conservation of angular momentum. When it is necessary to change the orientation of the satellite, reaction wheels start to rotate according to the control signal; causing a change in angular momentum. The satellite body produces an angular velocity in the opposite direction to preserve the total momentum. In this way, the orientation of the satellite can be controlled. Although three reaction wheels are sufficient for attitude control in three axes, satellites generally use reaction wheel configurations consisting of four wheels. An extra fourth wheel creates actuator redundancy. If one of the reaction wheels fails, the other remaining three reaction wheels can still complete the mission. The internal dynamics of the reaction wheels are modelled as a DC motor on Simulink. There are various disturbance torques that affect the satellite along its orbit. For example, gravity-gradient torque comes from the Earth's unequal gravitational force acting on different parts of the satellite. The parts of the satellite that are closer to the Earth are exposed to more gravitational force. This imbalance creates a torque on the satellite. Disturbance torques like gravity-gradient torque however can also be used to keep the satellite stable. In this thesis, the gravity-gradient stability behaviour of the simulated satellite was also examined.

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