41
Archived Theses
0
DOIs Assigned
0%
DOI Rate
Discipline
Investigation of thermal effect on aeroelastic flutter speed
Aeroelasticity is a discipline that examines the behavior of elastic structures in fluid flow. One of the critical phenomena within this field is flutter, a potentially catastrophic instability where aerodynamic forces couple with structural vibrations, leading to uncontrolled oscillations. Determining flutter speed is vital for aircraft safety, as exceeding this limit can result in structural failure. Therefore, flutter speed must remain outside the flight envelope. Traditionally, thermal effects are often neglected in flutter calculations, despite their potential influence on structural characteristics. This study investigates the effects of thermal loads on the flutter characteristics of aircraft structures. Initially, a modal analysis is conducted by using MSC Nastran Software to determine the natural frequencies under varying thermal loads. The numerical method is validated through experimental vibration tests performed at different temperatures, comparing the experimentally obtained natural frequencies with those derived numerically. Subsequently, a generic fighter aircraft mission profile is developed to create a thermal map for an AGARD wing. Thermal computational fluid dynamics (CFD) analyses are performed by using STAR-CCM+ software to simulate temperature distribution across the wing structure. The generated thermal loads are then applied as pre-loads onto the AGARD wing, and modal analyses are conducted by using the validated numerical methodology in MSC Nastran. Finally, flutter speeds for different thermal load conditions are calculated using the ZAERO software. Comparations reveal that increased thermal loads result in a reduction of both flutter speed and flutter frequency. The findings emphasize the necessity of incorporating thermal effects into flutter analysis for a more accurate assessment of aircraft structural stability.
Evaluation of the V-tail configuration on a machine learning and CFD based design framework
Recent works show that unconventional tail configurations like Vee-Tails might be good alternatives to conventional tails in terms of providing less weight and reduced drag in commercial airplanes. In this study, a conventional tail configuration consisted of an horizontal and a vertical stabilizer has been replaced by a Vee-Tail. The baseline airplane is the CSR-01, which is an hypothetical aircraft from the CeRAS repository. Various sizing cases for the Vee-Tail configuration have been considered. The sizing variables have been selected as the following 4 geometric parameters that define the Vee-Tail: The span length, the root chord, the taper ratio and the dihedral angle. The flows over the whole airplane have been computed using a Reynolds-Averaged Navier-Stokes solver. The Spalart-Allmaras turbulence model has been employed to compute the eddy viscosity. All the flow computations have been done in a parallel computation environment. Once the flows have been computed, the aerodynamic performance of the Vee-Tail has been established using the Support Vector Regression, which is a machine learning algorithm. A Vee-Tail geometry has been determined matching the lift, drag and pitching characteristics of the original configuration as close as possible.
Aerodynamic investigation of very closely flying fighters to develop new formation types
New generation fighters act as key-player in the operational environment within the scope of today's technological capabilities. At present, the old generation fighters have to be more challenging against the 5th generation fighters. Considering the improvements for the new generation fighters, to be more successful in an operational area, except those the technological developments, the fighters need to differentiate their level of performance, the performance of subsystems, and the most crucial thing, the concept of operations. As part of system and performance capabilities for fighters, many effective tactical formation sequences used to increase the performance of the fighters in operational environment. This thesis covers the studies to increase the "Survivability" and "Lethality" factors for Friendly (Blue) Force systems by using a new close flight formation sequence against the problems encountered in the targeting and detection algorithms of the Opponent (Red) Force fighters' airborne radar systems. For this scope, single F-16 aerodynamic analyses were performed using CFD methods on ANSYS Fluent to decide on triple and quadruple formation sequences depend on aerodynamic and formation requirement limits. Aerodynamic analyses of close-flight formation of older generation fighters were performed using CFD methods for triple and quadruple F-16 arrays. Finally, RCS calculations were made for the new formation sequences and the predicted results were obtained. The results have been interpreted with the Operational Analysis perspective to analyse the effectiveness in the operational environment, and new tactical developments and operational perspectives were gained on the effectiveness of the 4.5th Generation Fighters and older generations in the operational environment.
Numerical investigation of inclined fluidic oscillator in microchannel heat exchanger
The primary aim of this research was to investigate the enhancement of momentum and thermal energy transfer in a microchannel heat exchanger through the use of fluidic oscillators. To this end, a sweeping jet impingement heat transfer performance was evaluated with varying impact angles, distances between fluidic oscillators, and Re numbers. This study employed two fluidic oscillators to investigate heat transfer in a microchannel heat exchanger. To enhance the effectiveness of heat transmission in jet cooling, the fluid was oscillated. This allowed it to be calculated how much the flow rate and, consequently, the local Nu value increased, increasing the heat transfer performance. The study was based on the work of Wu Y. et al. validation studies were carried out on the study from the base study and model is validated for 3 Re numbers. Following the validation research, angles of 30, 45, and 60 degrees were evaluated between the fluid oscillator and the channel. Additionally, the distance between the two oscillators was investigated at distances of 7D, 7.5D, and 8D from the center of the outlet hole. A steady heat flux was applied to the channel surface, which served as the impact surface in the model. The fluid's Reynolds range was identified as 3000 to 10000. The study's findings revealed that increasing the oscillation of the jet flow process improves heat transfer efficiency. Additionally, it was found that when the impact angle grew, the pressure drop and Nu number likewise increased, with a 12 percent improvement in heat transfer being the largest. The heat transmission is found to be positively impacted by shortening the space between the strings, with a maximum improvement of 5.7 percent.
Weight reduction in aircraft with topology optimization
One of the most important parameters in airplanes is weight. Aircraft must be light enough to perform the tasks and capabilities expected of them, while at the same time being structurally resistant to the loads they encounter. Historically, different methodologies have been tried to make lighter aircraft designs. Topology optimization has recently become a preferred weight reduction method in the aviation industry. Lighter designs are obtained by removing the material that does not provide sufficient strength to the structure from the structure. In this study, weight reduction was made in the wing area of Cessna 172 aircraft by topology optimization. Before the topology optimization phase of the wing structure, the initial weight and the forces acting on it were calculated analytically. Based on the real-life weight of the wing structure, a finite element model of its pre-optimization state was created, and its structural strength was checked. A valid model was obtained before the topology optimization by making validation analysis of the created finite element model. In topology optimization, two different topology optimizations were applied to 3-dimensional and 2-dimensional structures. In these structures, weight reduction was achieved by determining the design areas and constraints for optimization. The structure obtained from topology optimization was statically evaluated again in order not to fail.
Buckling of metallic skin-stringer and sandwich panel structures: A comparative numerical study
In this study, global buckling behaviour of the metallic skin-stringer and honeycomb sandwich panels, which are used on the helicopter fuselage sections, is analysed. The metallic skin-stringer having different stringer cross sectional areas (z-, c-, and omega-section) and the metallic honeycomb sandwich panels in three variants (panels with longitudinal doublers, with transverse doublers, and without doublers) are examined with finite element simulations. The linear eigenvalue buckling analysis is used to calculate the critical global buckling load for simply supported boundary conditions. The most durable skin-stringer panel against the global buckling failure is determined to be the panel reinforced with the omega stringers. The panel with z stringers is found to be the least durable configuration against global buckling. The buckling strength of the panel with omega stringer is 13.1 and 15.6 percent higher than the panels with c and z stringers, respectively. The most durable honeycomb sandwich panel configuration against global buckling is determined to be the panel with transverse doublers. The panel without any doublers is found to be the least durable configuration against global buckling. The buckling strength of the panel with transverse doublers is 11.8 and 12.6 percent higher than the panels with longitudinal doublers and without any doublers, respectively. Honeycomb sandwich panels exhibit higher buckling strength compared to the skin-stringer panels. Critical buckling loads of the sandwich panels per kilogram are greater than those of the skin-stringer panels. Considering the structural, manufacturing, and environmental requirements, metallic honeycomb sandwich panels can be utilized as side panels of a helicopter fuselage and they can give better buckling performance per unit weight. Two panel concepts also can be used on different sections of a helicopter fuselage as long as they meet the structural requirements.
Passive flow control around the airfoil with slot
In this thesis, numerical analysis was performed to control of flow over NACA 0018 airfoil with slot at variant angles of attack. The Reynolds number was set to Re=3x10^5. Inlet of slot was posiniated on the pressure side of airfoil at chordwise location of x/c=5% while outlet of slot was located on the suction side of airfoil at three different chordwise location. Numerical analyzes were carried out using two different turbulence models, K-ω SST and K-ε RNG. Obtained results from the numerical analyzes were evaluated in terms of effect of the slot on flow structure and aerodynamic forces. It was observed that the location of slot remarkably affect flow over airfoil and delays stall angle. The model 40 suppress the formation of recirculation region formed over the airfoil. Furthermore, it was observed that while the stall angle of base airfoil is α=16°, it was delayed up to α=20° for Model 40. Also, it was observed that the lift coefficient was increased by 154% for Model 40 at α=19°. On the other hand, it was observed that Model 20 significantly effective on the flow control in comparison with the base airfoil at angles of attack of α ≥13°, while it has insignificant effect at low angles of attack. Similar observation was also obtained for Model 60 at low angles of attack. According to obtained results from both turbulence models, the most efficient model was determined as Model 40 in order to control flow around the airfoil especially for post stall angles.
Vision-based landing site detection for a UAV: From theory to application
This study addresses the critical need to enhance the safety of multirotor UAVs during emergency flights by addressing the challenges of identifying suitable landing spots and avoiding collisions with obstacles. Drones encounter difficulties due to their elevated height, leading to potential collisions, inaccurate distance estimation, and weak radio signals. Researchers are diligently developing a cost-effective and efficient landing system to mitigate these issues and enhance overall safety. Key gaps in vision-based landing site detection for UAVs include adapting to varying environmental conditions, achieving real-time processing, precise obstacle recognition, ensuring robustness, and scalability, handling data anomalies, integrating with other sensors, obtaining diverse training data, complying with regulations, and maintaining cost-effectiveness. To address these challenges, the proposed model combines a U-Net, a deep Convolutional Neural Network (CNN) architecture, with a ResNet 34 backbone. The model builds a labelled aerial photo database using image segmentation techniques for CNN training, determines landing areas through 2D dataset image analysis, and plans paths from the UAV's position to the designated area centre using the A* algorithm. The entire research is implemented in Python, harnessing TensorFlow's capabilities. Impressively, the model achieves a training accuracy of 96.6% and a validation accuracy of 96.34% while utilizing approximately 80% of the dataset, comprising 1480 images, during the training phase, with a recorded training loss of 0.1034. These outcomes underscore the model's exceptional accuracy, establishing it as a cornerstone for subsequent landing area selection and path planning by the algorithm.
Investigation on the effects of dielectric material on the thrust performance of annular DBD plasma ion thrusters
In this thesis, the effects of dielectric material on the propulsion performance of annular DBD plasma ion thrusters are investigated. Briefly, the history of electric propulsion system, types of electric thrusters and DBD ion thrusters are given. In order to maximize the propulsion performance, geometrical parameters (annular and sinusoidal), electrical parameters and material types were considered and experiments were carried out. In all experiments, voltages of 4 kVpp, 6 kVpp and 8 kVpp were applied respectively. An electronic thrust measurement device was used to measure the thrust generated by the DBD plasma actuator during the experiments. As a result of the experiments, it is stated that the thrust value increases while the voltage is increased. Moreover, it is observed that for sine geometries with the same outer diameter, the thrust efficiency decreases as the diameter increases.
Investigation of ice accretion effects on aerodynamic performance of NACA 2415 airfoil at low reynolds number
This thesis examines the effects of three distinct ice accumulations on the aerodynamic performance of the NACA 2415 airfoil. A brief discussion is provided of icing environments, icing types and the icing effects on the airfoil and also on the aircraft components. Besides, some necessary information is provided for icing prevention (anti-icing) and de-icing techniques. The experiments are carried out for the base airfoil as well as glaze, rime, and horn ice accumulated airfoils at Re=1.2×105. For this study, force and velocity measurements are taken. Additionally, the oil and smoke-wire flow visualization techniques are also performed. The selected range of angle of attack for force measurements is defined as -7° to 23°. Additionally, considering the results of force measurements, seven distinct angles of attack values including 0°, 4°, 8°, 13°, 14°, 17° and 20° are chosen for velocity measurements and flow visualization experiments. Results showed that, CLmax of ice accreted airfoils are decreased as expected. 20%, 23% and 88% of decrease in maximum lift coefficient is observed in the GIA, RIA and HIA models, respectively. An increase in drag coefficients at α = 0° for the GIA, RIA and HIA airfoils is observed to be around 14%, 4% and 307%, respectively. A sudden increase in lift coefficient of the GIA airfoil is observed at post-stall. Moreover, a double-peak behavior is noted for the GIA, RIA airfoil and also for base airfoil in velocity measurements.
Hibrit kompozit malzemelerin balistic modellemesi
Hybrid composites have become one of the essential and highly important engineering materials for advanced and complex applications in today's industry. In this study, the behaviours of composite materials under the same ballistic effect of layer type and layer arrangement were investigated experimentally, provided that the number of layers remained the same. . In the laboratory, 10 resin-supported glass fibre and carbon doped hybrid composite samples were obtained by vacuum infusion production method. In order to examine the ballistic effects of the produced composites, shooting was carried out in the polygon in accordance with STANAG 2920 standards. As a result of these specified firing tests, the damages caused by the projectile on composite materials were analysed. In addition, it was analysed how the layer order and layer type in the material affect the projectile velocity.
Düşük reynolds sayisinda naca 0012 kanat profi̇li̇ni̇n aerodi̇nami̇k özelli̇kleri̇ üzeri̇ndeki̇ hücum kenar deği̇şi̇mi̇ni̇n etki̇si̇
In this study, it is aimed to enhance the aerodynamic performance of the NACA 0012 airfoil with the morphing leading edge method. The investigation was conducted using an airfoil with a chord length of 120 mm under Reynolds number conditions of Re = 105. In addition to the base airfoil, a MATLAB code was developed to generate modified airfoil geometries by specifying the deflection angle and the deflection initiation point (expressed as a percentage of the chord length) as input parameters. The computational analysis considered deflection angles of 2°, 4°, and 6°, initiated at 10%, 20%, and 30% of the chord length. All configurations were analyzed using ANSYS FLUENT and compared aerodynamically with the base case (0° deflection). The performance evaluation focused on maximizing lift coefficient (CL), minimizing drag coefficient (CD), and optimizing the lift-to-drag ratio (CL/CD) as a measure of aerodynamic efficiency. The results demonstrate that airfoils with 6° deflection achieved the highest aerodynamic efficiency relative to the base model in the post-stall regime (e.g., α = 18°), with observed reductions in laminar separation bubbles (LSB) and trailing-edge vortex intensity. At moderate angles of attack (α = 10°), airfoils with 4° deflection exhibited optimal performance, yielding a 73.61% improvement in CL/CD compared to the base model. While stall characteristics revealed limited improvement, all modified geometries demonstrated reduced drag compared to the base model.
The effect of partial flexibility on aerodynamic performance of finite wing
In this thesis, the aerodynamic effects of partially flexible surface coverage applied to a NACA 0018 airfoil operating in low Re regimes were experimentally investigated. The flexible surface, made of latex material, was placed along a specific chordwise region of the wing, and four different wings with varying spanwise flexibility ratios were tested. Wind tunnel experiments were conducted in the Re range of 3 x 10⁴ to 10 x 10⁴, measuring lift (CL) and drag (CD) coefficients. Additionally, surface oil and tuft flow visualization techniques were used to analyze flow structures. The findings revealed that the partially flexible surface significantly suppressed the laminar separation bubble (LSB), delayed stall, and enhanced lift coefficient, thereby improving aerodynamic efficiency, particularly at lower Re. However, the effectiveness of the flexible surface diminished as the Re increased. These results contribute to the literature by demonstrating the potential of partial flexibility in optimizing passive flow control methods for UAV applications.
Yüksek hızlı ı̇ç ve dış akışlar ı̇çı̇n ısıl aşınmanın modellenmesı̇
In this thesis study, the ablation of graphitic materials is modeled and two different analysis tools are developed. In the first model, the computations are conducted with a decoupled approach. Secondly, a more generalized calculation method is introduced via coupling the flow field and solid conduction analysis codes. In the modeling studies, both the gas and solid domains are assumed to be axisymmetric and the physical domains are discretized with structured finite volume cells. For the flow field analyses, Godunov type approximate Riemann solvers are used for the convective flux calculations and the gradient terms of the viscous fluxes are calculated with the help of grid transformation metrics. Mass diffusion terms are included in the Navier-Stokes equations. Baldwin-Lomax, Baldwin-Barth and Spalart-Allmaras turbulence models are implemented and the results are compared for the coupled approach. A linear solver is developed for the computation of gas phase chemical reaction source terms. The effects of the solid thermal properties with temperature are not neglected in the computations. The models are tested for hypersonic air and Solid Rocket Motor (SRM) nozzle flow conditions. Wall equilibrium is assumed for graphite in air problem and an implicit solver is developed for the calculation of surface species mass fractions rather than using previously prepared lookup tables. Both equilibrium and finite rate surface thermochemistry approaches are tested for the internal flows of the SRM nozzles.
Isıl bariyer kaplamalarda oksidasyon ve plastısıte etkılı hasarın modellemesi
Computational analysis and simulation of multi-physics phenomena taking place in coating systems is still a challenging task. Specifically, for ceramic coatings used as protective systems for base materials against elevated temperatures, known as thermal barrier coating (TBC) systems, construction of continuum level models which can express coupled nonlinear phenomena has attracted great attention. Thermal stresses, oxidation, creep and numerous other mechanisms and phenomena make it even harder to model and simulate the behavior of TBCs and bring a need for the development of premier models. In this effort, a new numerical model that allows simulation of oxidation and thermally grown oxide (TGO) in bond-coat is presented. Phase-field theory is used with finite strain formulation and implemented using user element subroutine (UEL) in ABAQUS software for the finite element method. Results are compared with experimental data available for TGO in the literature. Initiation and development of damage were also modeled considering plasticity induced failure. Results were also compared with experimental data available for damage and stress development in isothermal conditions.
Helikopterler için gürültü minimal & yeşil yörünge ve uçuş profili eniyilemesi
The main aim of this study is to provide a multi-disciplinary optimization and track environment to generate acoustic optimal trajectories through waypoints that ensures the rotorcraft of interest can follow at practical effort, safety, fuel consumption and speed. Rotorcraft noise annoyance remains as a challenge to solve complex, three dimensional and coupled rotary wing aerodynamics, aeroacoustics and flight dynamics interactively. Two essential paths can be acknowledged in order to reduce annoyance. One is the more sophisticated option, optimized new rotorcraft design, whereas the other option is to benefit from the directivity characteristic of sound and perform trajectory optimization to minimize noise impact at noise sensitive premises. This study focuses on the second yet with consideration of the potential trade-offs between low noise signature and other performance parameters. Eventually, the main aim of this study is to develop a trajectory optimization and track framework for rotorcrafts providing minimal noise, low emission i.e. lower fuel consumption, safe and trackable, in other words "green" flight profiles. In this scope, a Lagrangian CFD solver specialized for rotor/propellers is developed, coupled with rotorcraft mathematical model and an aeroacoustics solver to build a high fidelity, accuracy and resolution rotorcraft comprehensive modeling environment. The developed methodology is validated with wind tunnel, whirl tower test data, PIV results and benchmark commercial tools. The developed comprehensive tool provides free flight trim, high fidelity modeling and analysis capability for conventional and unconventional rotorcraft configurations with unsteady wake dynamics covering blade-vortex, rotor-wake and rotor-rotor interactions. Further in the study, the comprehensive model is extended into a real-time computable simulation model. Then a model predictive control -an optimal control- approach is developed to simultaneously optimize the trajectory and control input to track the generated trajectory. The multi-disciplinary objective function including acoustics, performance, fuel, safety, comfort and mission concerns provides the so called "green" trajectory with reduced noise impact at desired locations. Various simulations were performed to further test the aerodynamic modeling, aeroacoustics analysis and trajectory optimization capabilities of the developed framework. It is concluded that the proof of concept, i.e. the potential of reduced noise impact and fuel consumption over the same mission through trajectory optimization, is achieved. Developed methodology can be utilized to generate optimal flight routes and procedures specific to rotorcraft configuration, which are currently rather generic for all types of rotorcrafts, especially for booming e-VTOL platforms that will mostly operate over urban areas or for re-planning of legacy flight routes
Bir jet eğitim uçağı kanadının dayanım ve direngenlik kısıtları altında yapısal eniyilemesi
This thesis presents structural optimization studies for a jet trainer aircraft wing structure. The main purpose is to select the most convenient rib/spar layout for T-38 wing like geometry which consists of metallic and composite components by using a finite element modeling and analysis tool of MSC NASTRAN optimization capabilities. In this study, in order to decrease the number of design variables and as well as to be able to obtain smooth thickness transitions between adjacent zones, design variable linking method is used by applying different shape functions. First of all, after creating the outer geometry in CATIA, the finite element model (FEM) is prepared using MSC PATRAN. Similarly, all design zones where design variables are free to change are designated. To be able to link the thicknesses of the components to each other with regards to their locations on the wing, all coordinates of design zones are transformed to natural coordinate system as their center points are positioned as located between 0 and 1 in span-wise direction and -0.5 and 0.5 in chord-wise direction using scripts coded in Python. Aerodynamic loads are then calculated by using MSC FLDS (FlightLoads) tool and summed in predefined monitor stations to distribute onto the structure using rigid body elements. Obtained shear force and bending moment distributions are also compared with another method called as Schrenk's Approximation. In the optimization studies, thicknesses of the metallic structures and principal composite ply thicknesses are considered as design variables. There are five design constraints which are von Mises stress for the metallic structures, failure index for strength check of composite structures, global buckling, damping and natural frequencies which are used to control the flutter speed. The objective in this research is to obtain a minimum weight in design while providing predefined constraints via investigating various design candidate geometries having different layouts but same outer geometry. Finally, by using the scripts developed within this thesis study, the design variable linking method used according to the locations of the design zones become suitable also for the Nastran Sol 200 users.
Optimizasyon yöntemleri ile helikopter manevralarının gerçekleştirilmesi
In order to certificate a helicopter, aviation safety agencies must know that all designed helicopter configurations can withstand all loads resulting from maneuvers defined in certification standards. In other words, the maneuvers defined in regulations must be performed for each appropriate combination of weight and center of gravity. Then, designers have to prove that the helicopter can fly safely across the entire flight spectrum. Therefore, load engineers perform all maneuvers defined in the helicopter usage spectrum in order to analyze all possible load values. In traditional methods, a trial and error approach is used to reach the maneuver. However, performing these maneuvers with a trial and error approach not only causes expensive computing but also requires more engineering effort. Additionally, it may cause some defects in the maneuvers. Therefore, this thesis study aims to achieve the desired helicopter maneuvers using optimization methods in order to reduce the calculation cost and engineering efforts and perform the maneuvers accurately. For this purpose, only selected maneuvers are performed in this thesis. In addition, various optimization methods in different configurations have been applied to solve these maneuvers. Thus, the most useful one among them has been decided by making the necessary comparisons. Finally, the most useful optimization method has been applied to maneuvers frequently performed by helicopters throughout its lifetime.
Jenerik bir çırpan kanat mekanizmasının aerodinamik ve atalet karakterizasyonu
The primary focus of this thesis is to conduct an aerodynamical and inertial characterization of a generic flapping wing mechanism. Firstly, a CAD model of the utilized flapping mechanism is created, and a custom test platform is developed for the performed flapping-wing test sequence. Experimental validation of the modal analysis performed for the developed test platform is presented. In the inertial force characterization study, the standalone effect of the inertial forces that occur during flapping motion on the produced total aerodynamic lift force is aimed to investigate. In this framework, the wing membrane is eliminated, and a flapping wing test is performed to measure the inertial force generated due to the motion of the prescribed kinematics in the lift direction. The dynamic calibration of the utilized sensor is carried out to ensure the fidelity of the acquired data during the flapping test. Verification studies regarding to the experimental inertial force investigation are conducted by performing theoretical and dynamic finite element analysis approaches. For the aerodynamic characterization of the proposed system, a nylon wing membrane is attached to the utilized flapping wing mechanism, and a flapping wing experiment is performed in the air medium. To verify the experimental results, a numerical validation study is conducted and, a fluid-structure interaction analysis model is presented for the corresponding model. Besides, the effect of wing-flexion on aerodynamic performance is investigated for the corresponding model. In this regard, a distinct fluid-structure interaction analysis model, including a flapping wing with high-rigidity material properties, is created to minimize the deflection that occurs on the wing structure. Experimental, theoretical, and numerical approaches conducted for the inertial and aerodynamic characterization of the utilized flapping mechanism showed great resemblance to each other. A comparative study performed for the effect of wing-flexion shows that wing elasticity enhances the thrust force for the proposed model.
Doğal dinamik kullanılarak hale yörüngede ve dünya etrafında seyır eden uzay araçları için kol uçuşu tasarımı
This thesis presents the studies performed for spacecraft formation flight design and analyses. Two main design problems are addressed in this thesis: First the formation flight design of satellites flying near Sun-Earth collinear libration points; the second the formation flight design of satellites flying at Low Earth Orbit. Thus, formation flight design near Sun-Earth L1 and L2 libration points is investigated first, where solar radiation pressure as well as the gravitational disturbances of the planets are taken into account, for different Julian dates. The periodicity of the relative motion in formation flight is taken as a design criterion and convenient initial conditions are computed for each deputy satellite for desired formation configuration. It is desired that the required formation is maintained without the need of any correction maneuvers for formation keeping. In the second part, the method presented in this thesis for formation flight design is applied to Low Earth Orbit satellites. The results are also compared to the results obtained using current methods available in the literature. Results show that proposed method gives more consistent results and provides flexibility on the orbit design for formation in terms of formation keeping and fuel consumption needs as compared with the currently available methods. Finally, it can be stated that the trajectory and orbit computations done using the method presented in this thesis provide long term formation flight for space missions at L1, L2 libration points and for Low Earth Orbit missions. The main contribution of this method is the inclusion of all disturbancing forces acting on the satellite as a time variant discrete model. The initial conditions are found iteratively that ensures the periodic trajectory. Here, the usage of time variant discrete model to obtain periodic relative motion is a feature that distinguishes the current study from the existing methods in the literature.
Entegre güçlendirilmiş yapısal paneller için yapay sinir ağı bazlı burkulma yükleri belirleme aracı
The sudden change in the load carrying capacity under compresive loading, called buckling, may cause catastrophic failures. Therefore, determination of the first buckling and collapse loads of structural elements is essential in preliminary design stages. Finite element (FE) analyses and structural testing are used to determine buckling characteristics of a structural element. However, in early design stages, FE analyses are time consuming and structural testing is costly. In this study, an artificial neural network tool (ANN) is used to reduce computational effort to determine buckling loads of integrally stiffened structural panels in early design stages. Reuslts of FE analyses are employed to train the ANN. Moreover, Latin Hypercube Sampling (LHS) methodology is used to reduce the number of required FE analyses to generate database that artificial neural network is based on. Finally, a Multi-fidelity sampling algorithm that uses FE models with different mesh resolutions is implemented for generation of the ANN database in order to reduce computational time spent for finite element analyses. Mean errors and fit performance model results are compared to determine accuracy of the neural network results.
Mühimmat sürüsü için güdüm yöntemlerinin değerlendirilmesi
In this study, collective flight, mission execution and decision-making algorithm for airdrop micro munitions with two different swarm algorithm alternatives are examined. Through a concept munition design is created and an aerodynamic database is prepared with the Missile-Datcom software. Six degrees of freedom flight mechanics model, subsystems of munition and environment are modeled. A linear model is obtained and controller is designed. Pole placement method is used in the controller design. Full state feedback assumption may not suitable for the actual problem. For this reason, the autopilot gains of the designed controller are converted into gain values suitable for measurements by using the projection control method. The designed autopilot tracks the commands of the swarm and proportional navigation guidance. The swarm algorithms coefficients are optimized for the nominal state by particle swarm optimization. Alternative algorithms obtained as a result of optimization are tested under various distortions by Monte Carlo analysis. Sensitivity analysis and performance evaluation are examined as well.
Sünek hasar tahmini için gözenekli bir plastisite modelinin sayısal uygulaması ve analizi
Ductile damage and fracture are known to be driven by the microvoid nucleation, growth, and coalescence. Porous micromechanical description of the ductile metals led to many phenomenological material models, which are used to predict the damage and fracture in engineering structures. In this thesis, the assessment of a rate-independent porous plasticity model is done through the representative volume element (RVE) calculations. The model is based on the formalism presented in [1] which is implemented as a user material subroutine through a prediction-correction scheme similar to a classical J2 plasticity framework. In this context, RVE's are taken from a periodic array of spherical voids surrounded by an elastoplastic matrix material with isotropic exponential hardening, and they are deformed under a constant triaxial stress state with a displacement controlled method. The implementation of the model and the method of the RVE calculations are explained in detail. Limitations of the original model are discussed, and a heuristics extensions to the constitutive framework is proposed to obtain a better fit between the porous model and the unit cell results in terms of volumetric void growth and equivalent stress-strain relation. Numerical analyses show the possibility of achieving a compact framework with a straightforward implementation that agrees well with the RVE simulations for a wide range of stress triaxiality values. The present framework is compared with the widely used Gurson-Tvergaard-Needleman (GTN) model and the differences are discussed. A simple void coalescence relation is added to this framework to simulate the final failure phase of ductile deformation. Additionally, tension simulations with smooth and blunt notched specimens are performed with the GTN model, the present porous plasticity model, and the Johnson-Cook uncoupled damage model to address the model's performance in a ductile fracture simulation. Results show that the present framework and the GTN model can yield almost identical results in notched simulations in terms of engineering stress-strain response and the porosity evolution. The thesis is concluded with an outlook and possible future improvements.
Genel maksat helikopteri için makas tipli kuyruk rotorunun aerodinamik ve aeroakustik incelemesi
The noise of the rotorcraft has become nonignorable with their increasing use in daily life. The aerodynamically active components such as the main and tail rotors are the primary sources of noise, and therefore, their evaluation during the design phase must also be carefully included. Pressure variations caused by several aerodynamic mechanisms on the rotor blades are responsible for the generation of the noise, which may be obtained by computational fluid dynamics. Along with rotor locked self pressure fields, the tail rotor being in the main rotor wake and exposed to strongly nonstationary airflow also need attention. Hence, full helicopter modeling is required to have an accurate evaluation on the tail rotor. However, this modeling and solution are costly for computational fluid dynamics. Therefore, an alternative approach for acoustic analysis is developed in this study. Accordingly, the goal of the thesis is to propose a coupling methodology between comprehensive rotorcraft analysis and an aeroacoustic solver, and to implement of the methodology for investigation of scissors tail rotor. Firstly, the comprehensive analysis approach is validated using available required power data and sectional blade airloads. Then, the proposed pressure distribution methodology is implemented to couple comprehensive analysis approach and acoustic solver. A CFD generated airfoil pressure database is used to find pressure distribution over each blade section. Additionally, this coupling is validated using sectional loads and acoustic pressure. Furthermore, aerodynamic and aeroacoustic behavior of the tail rotor are evaluated for isolated and main rotor interacted conventional tail rotor configuration. Finally, effects of vertical distance and scissors angle on aerodynamic and aeroacoustic behavior of the main rotor interacted scissors type tail rotor configuration are investigated. Results indicate that the main rotor wake has a significant impact on the tail rotor at certain flight conditions conditions, and vertical distance and scissors angle changes the oscillatory loads, and there are scissors tail rotor configurations which offers lower noise level than conventional tail rotor.
Küçük uydularda görüş tabanlı üç eksenli yönelim tahmin algoritmasının tasarımı
Equipment and budget limitations of satellite missions encourages researchers to use on-board equipment for multiple tasks. Camera stands out as an ideal equipment that can be used for multiple purposes in satellites making Earth observation. In this thesis, an attitude estimation algorithm is designed to aid the coarse attitude estimates with the attitude information obtained from Earth images. Attitude information is extracted from Earth images according to perspective geometry laws. This estimated attitude is fed into a multiplicative extended Kalman filter (MEKF). Coarse attitude information is obtained by pre-processing magnetometer and sun sensor measurements within the QUEST algorithm and also used in the measurement update process of the filter to ensure the sustainability of the attitude estimates. The algorithm is tested for a hypothetical nano satellite in different scenarios depending on the availability of Earth images at different times.
Kaplamalı takviyeye sahip polimer nanokompozitler için iki seviyeli bir homojenleştirme yöntemi
Polymer nanocomposites are materials which have polymer matrices and nano-scale reinforcement elements. Other than the matrix and the reinforcement element, another phase in nanocomposite systems has been observed and some of the property enhancement or deficiency has been attributed to this phase. The third phase in the composite system is known as the interphase. The local properties of polymer nanocomposites are not easy to distinguish since the length scale is very small and characterization at the nano-scale is not a simple procedure. Macroscopic behavior of nanocomposites can be deducted by homogenization of a representative volume element in many different ways. The aim of the thesis is to propose a new homogenization approach to model the polymer nanocomposites with coated inclusions. Understanding the structure of nanocomposites, formation of the interphase and different characterization techniques for the interphase are important for this aim. Interphase properties affect the overall macroscopic mechanical behavior. So it is important to model the interphase and obtain the effective composite properties accordingly. Different homogenization techniques and applications related to this problem are studied and a new approach is proposed. This method aims to model load transfer between the matrix and the reinforcement element through the interphase in a correct way. Furthermore, a computational gain is aimed with the proposed model in comparison to a full three-phase finite element simulation. Another goal is to extend the proposed method to viscoelasticity. For this purpose two-dimensional and three-dimensional finite element codes are developed for a viscoelastic homogenization framework. The formulations regarding the viscoelastic homogenization are derived and presented.
Bir iletişim uydusunda titreşim test sonuçları kullanılarak sonlu elemanlar modelinin iyileştirilmesi
In this thesis, the vibration analysis of a communication satellite is performed, and the improvement of the finite element model by using vibration test results is presented. First, the satellite finite element model is generated using MSC/PATRAN and MSC/NASTRAN commercial software. With the natural frequency and frequency response analysis, the expected frequency values and response amplitudes in accelerometers are calculated in vibration tests. The results obtained in the vibration test are compared with the analysis results using the HYPERGRAPH commercial software and their similarity with the modal assurance criteria calculation, which is developed using the FORTRAN programming language. The frequency values converged to the measured values as a result of the finite element model improvement, which is performed using error sensitivity assessment at the relevant frequencies.
Yaklaşma tapalı füzeler için öldürme analizi aracı
In this thesis study, a developed kill analysis tool for the missiles having proximity fuzes is presented. Some of today's missiles have proximity fuze, and this type of missiles' warhead can be detonated not when the missile hits the target but at a certain distance. The warhead consists of fragments, and the fragments scattered by the explosion's effect are intended to penetrate the target. A tool is needed for doing kill analysis in order to find the blasting time that the maximum number of fragments hit the target. It is aimed to detonate the warhead at the right time during the missile-target engagement by doing kill analyses for different engagement conditions with the help of the developed tool. Then, the kill analysis results are embedded into the missile computer before the missile is launched.
Kanat profillerinde buffet araştirmasi
The aim of the present thesis is to investigate buffet onset characteristics in two-dimensional flow. NACA0012 airfoil was chosen as the geometry in this study since the profile is frequently used for the lifting surfaces of aircraft, which are exposed to high flight loads during transonic flight and this causes buffet and flutter. In addition, it is one of the very few models which was studied for the buffet onset investigation in the literature and detailed wind tunnel test data is available. In the present study, after determining the optimum mesh resolution and the appropriate turbulence model, the results of the steady flow numerical analysis for low and high Reynolds numbers were validated with wind tunnel data. The transient flow is investigated for the buffet onset region of the NACA0012 airfoil and the results are verified with wind tunnel data. The oscillating behavior of the shock wave on the airfoil is investigated through the frequency and the Strouhal numbers of lift, drag, moment and static pressure distributions through time and frequency. The study investigates the primary effectors that stimulate buffet onset and the behavior of the flow during buffet through numerical analyses using Unsteady Reynolds-Averaged Navier-Stokes (U-RANS) Equations. The effects of Mach number, angle of attack and Reynolds number on buffet onset is investigated for the examination of buffet characteristics in detail.
Sesaltı gömülü hava alıkları için adjoint tabanlı aerodinamik şekil optimizasyonu
In this study the aerodynamic shape optimization of subsonic submerged intake is performed. The total pressure at the aerodynamic interface plane and lift are employed as objective functions. Drag is constrained and employed as penalty objective. Open-ware platforms are used including SALOME for solid modeling, GMSH for the hybrid mesh generation and SU2 for flow solutions and adjoint based shape optimization. Free form deformation box is employed for shape parameterization and surface deformation. Initially NACA intake placed on a flat plate is solved and result is compared with reference study at subsonic flow conditions. Thereafter single objective shape optimization is employed which enhances the $C_L/C_D$ from 0.08 to 0.53 at the end of design cycles. The second case is of trapezoidal-entrance intake placed in slender aerodynamic body. The results are compared with reference study and baseline case for optimization is established. The shape is optimized using multi objective function and for multiple free form deformation boxes. The intake remains flush with the body during the shape deformation process. The intake internal surface is then allowed to deform along the optimization steps. The pressure recovery increases by 3\%. The third case is similar but with circular entrance in slender body. A single free form deformation box encloses the intake region and the deformation is allowed such that optimization creates a semi-submerged shape. The total pressure recovery increases by 2\% and $C_L/C_D$ also increases by 10\%. In the last case again the trapezoidal entrance intake is employed. The optimization process is similar to the third case. The optimum shape produces a substantial deformation at the intake surface and a semi-submerged intake is finally formed. The total pressure recovery improves by 6\% and an additional $C_L/C_D$ ratio of 0.04 is achieved.
Dokuma ve çapraz-katlı bükümlü kompozit yapılardakı hasarın deneysel yöntemlerle incelenmesi
Laminated curved-shape composite parts which are used in the spar and ribs in aircraft and wind turbine blades are subjected to high interlaminar tensile and shear stresses. These stresses cause delamination and subsequent reduction in load-carrying capacity. In this study, failure mechanism of cross-ply and fabric curved composite laminates under pure transverse loading are examined experimentally using an in-house designed test fixture. Stress field over the curved beam is obtained with finite element analyses (FEA) and analytic solution where multilayered theory is used. Dynamic delamination of the cross-ply and fabric curved specimens is recorded with a high-speed camera where the failure sequence is captured and the crack tip speeds are calculated. For the cross-ply laminates, von Mises strain field is obtained with digital image correlation (DIC) method where the strain fields are found to agree well with elastic FEA up to a specific loading at which a population of matrix cracks nucleate. Detailed fractography of the tested specimens is carried out with digital microscope. Finite element analysis with 2D Hashin Failure criteria is also successfully predict the radial cracks observed in the micrographs of cross-ply curved composite laminates. However, the meandering crack path could not be properly predicted with 2D Hashin Failure Criteria. In fabric curved composite laminates, the failure is observed to occur due to inter-ply and intra-ply crack growth. In fabric laminates, the crack tip travels in a fluctuating manner at speeds reaching intersonic speeds whereas in cross-ply laminates the crack tip speed reaches about Rayleigh wave speed.
İyileştirilmiş aerodinamik performansa ve azaltılmış radar izine sahip ses altı hava alıkları için tasarım yöntemi geliştirilmesi
In this thesis, a design method for a subsonic intake with high aerodynamic performance and reduced radar signature is developed. Intakes are key components of aircrafts' propulsion systems that create open ended cavities resulting in increased radar signature. To achieve reduced radar signature characteristics, a parameterized double curved intake is proposed. Due to the nature of this multi-objective and multi-disciplinary engineering problem, surrogate based analysis and optimization approach is taken to reduce the computationally expensive and time consuming analyses. The thesis consists of three main sections. In the first section, validation studies for the intake performance evaluation methods are conducted. Mesh independence and turbulence model selection studies are carried out for the computational fluid dynamics analyses and radar cross section evaluation studies are carried out for the electromagnetic analyses. In the second section, a design problem is constructed and the verified analyses methods are used for the performance evaluation of the design points required by the surrogate model. Then the best performing design in terms of aerodynamics and radar cross section is obtained with an optimization study. In the last section, analyses with high fidelity prediction tools are conducted for the validation of the obtained design and the design method. Results showed that the surrogate based model is successful at the prediction of performance indicators and can be used for intake optimization purposes. Lastly the performance of the best design is investigated with post-process outputs revealing competitive results regarding aerodynamics and radar signature.
Jet eğitim uçağı ön gövdesinin cevap yüzey yöntemi ve genetik algoritma ile eniyilenmesi
The aerodynamic design of an aircraft's forebody geometry has a significant impact on its performance and stability. Although most studies on aircraft performance have focused on optimizing wing-like structures, the contribution of an effective forebody design to aircraft performance could be as significant as that of wing-like structures. A well-designed forebody can reduce the wave drag and improve the directional characteristics at high angles of attack. The forebody optimization of a jet trainer aircraft in terms of supersonic cruise performance and directional stability is investigated in this thesis. While doing so, two objectives are considered: the wave drag and directional stability. The response surface methodology is used to generate the aerodynamic database, and the non-dominated sorting genetic algorithm–II is used to search for Pareto-optimal solutions. The open-software flow solver SU2 is used to obtain turbulent flow solutions. It is shown that the optimization study enhances the aircraft's performance in terms of wave drag up to 2% and directional stability up to 30%.
Sınır tabakası geçişi ve boşluk akışı gürültüsünün geciktirilmiş ayrık-çevrinti benzetimine dayalı hesaplamaları
This dissertation investigates the boundary layer transition as well as the cavity flow noise prediction capabilities of a high-order in-house solver using various Delayed Detached-Eddy Simulation (DDES) frameworks. Before conducting the simulations, multiblock topology with a high-order overset grid technique is implemented into the solver, which makes mesh generation for complex geometries, such as the tunnel grids around blade sections, and cavity grids composed of two separate domains of the studied cases in this thesis. For the flow transition capability, the Baş-Çakmakçıoğlu (BCM) transition model is incorporated into DDES with a shear-layer-adapted (SLA) subgrid length scale, and applied to flowfields around a blade section and a cylinder. The results show that the BCM model captures the transition onset maintaining the laminar upstream flow while the SLA approach increases the turbulent content rapidly beyond transition. The collaboration between these two approaches enables capturing the aerodynamic coefficients of blade sections near the stall angles accurately. On the other hand, the SLA length scale is incorporated into the Improved DDES framework (IDDES-SLA) for computations of the M219 cavity flow, and its associated noise. The cavity problems are considered to have no physical lateral walls for reducing the demand for computational resources. The results show that for these specific cases the mean and turbulent flow fields could be captured reasonably without the lateral walls, when the cavity width is taken as at least one depth. In addition, unlike the standard one, the use of the SLA length scale helps capturing the Kelvin-Helmholtz instability dominated region. IDDES-SLA yields the best acoustic results among some other tested approaches, showing good agreement with reference studies. The absence of viscous lateral walls does not seem to have an impact on overall sound levels except near the front wall.
Delikli kompozit yapılarda boyutsal özelliklerin dayanım değerlerine etkisinin deneysel ve analitik olarak incelenmesi
In this thesis, we investigate the effect of size on the strength of open hole composite laminates using experimental and analytical approaches. The size effect consists of the effect of thickness and hole size while keeping width to diameter ratio constant. In the experimental part of this work, tensile tests are conducted on CFRP open hole specimens. In-situ images of the top and side surfaces of the test specimens are captured throughout the experiments and load displacement curves are acquired from the testing machine. The top surface images are captured using high resolution camera and analyzed with digital image correlation (DIC) technique. The side surface images are captured using high speed camera setup and these images and post-mortem microscope images are used to determine failure mechanisms. Finally, in agreement with the literature, size dependence is observed on a plot of failure stress vs. hole size. On the analytical part of this work, an analytical methodology is developed based on Tan's approach. The stress field is calculated using Lekhnitskii's Formulation and Classical Lamination Theory(CLT). The analytical strength prediction of the open hole composite laminate is done according to First-Ply-Fiber-Failure(FPFF) methodology. The analytical model performs well compared to other models in the literature in modeling the size dependence of the failure stress on the hole size.
Roket tipi geometrilerin aerodinamik kararlılık türevlerinin sayısal tahmini
Prediction of aircraft aerodynamic stability derivatives is crucial for evaluating its flight dynamic characteristics. By using the open source SU2 software, steady and unsteady Reynolds-Averaged Navier-Stokes analyzes are performed in order to determine the aerodynamic stability derivatives of a missile like projectile from literature. The employed prediction methods that are based on forced oscillatory motion and differential method are particularly emphasized for roll damping and pitch damping derivatives. Also, constant roll rate analyzes are performed with the options of rigid mesh motion and rotating reference frame. Differences between them are evaluated by comparing the results obtained from these different approaches with each other and experimental data. The effects of the Spalart Allmaras and Shear Stress Transport turbulence models on steady and unsteady simulations are examined in terms of accuracy and computational cost. The effects of oscillation method parameters such as reduced frequency, time step size, and oscillation amplitude are also investigated. The numerical prediction of the aerodynamic stability derivatives shows good agreement with the experimental data.
Strake-delta kanat konfigürasyonun adjoint tabanlı aerodinamik şekil optimizasyonu
Modern fighter aircraft compulsorily demand high maneuverability capability, which is mainly provided by an improved aerodynamic performance at high angles of attack. This is achieved by mostly employing strakes and canards. In this study, adjoint-based configuration and leading edge shape optimizations of a strake on a double-delta wing configuration are performed. SU2 is employed for flow and adjoint solutions. SU2 flow solutions are first verified on solutions for adaptive grids. In the configuration optimization, the sweep angle of the strake is considered as design variable. In the leading edge shape optimization, the free-form deformation box is employed. Remarkably, it is only allowed to modify the leading edge without changing the flat strake surface. Optimization studies are performed for both inviscid and turbulent flows at 10° and 22.5° angles of attack, respectively. It is shown that the sweep angle optimization based on turbulent flow solutions improves the L/D ratio by about 8.4% at 10° angle of attack and 2.7% at 22.5° angle of attack. In addition, the leading edge shape optimization based on turbulent flow solutions improves the L/D ratio by about 4.5% at 10° angle of attack and 3% at 22.5° angle of attack.
Kademeli ve eş merkezli çarpışma kutularının analizi: Tüp sayısının etkisi
In this study, stepped concentric crash tubes with circular AL6063 material are investigated using experimental and numerical methods. Crash tubes are used generally in automotive industry to as a passive safety component. In this study, feasibility of crash tubes in different areas are sought. So, it is designed almost ten times smaller than the conventional car crash tubes. The purpose of this study is to examine the energy absorption characteristics of stepped concentric crash tubes subjected to an axial impact load. The peak forces, internal energy and the total deflection are the selected parameters for the comparison of various crash tubes. The stepped concentric tubes have different lengths and different diameters, while the total mass and volume are kept constant. The area of the impacting object is taken as greater than the internal tube's cross sectional area to prevent passing the impacting object through the tube hole. First of all, a comprehensive study about articles on crash tubes presented. Almost 127 scientific articles are reviewed and tabulated information according to solution methods, geometry, material properties, impact velocity and impact mass properties. Table is given in Appendices section. In the next part, one selected model with AL6063 material is fabricated and tested at dynamic and quasi-static velocities. Dynamic tests are done by using INSTRON 8150 Drop Test Machine and quasi-static test is done by using MTS 647 Hydraulic Test Machine. Then, results of experiments compared and validated with results of analysis by using LS-DYNA, an engineering simulation software program In the second part, the effects of the number of the tubes and impact velocity are studied at high dynamic velocities by using LS-DYNA software program. To define optimal mesh quality, results of five different mesh sizes of a sample are investigated and chosen the most proper type of them to use in all analyses. After that, four different models are constituted with 3, 6, 9 and 12 stepped concentric tubes, respectively. All samples are analyzed at three different velocities: 50, 60 and 70 m/s and results are compared with one another. In the third part, stepped concentric crash tubes are modelled with aluminum foam and analyzed at 50 m/s impact velocity. Results of model with foam filled and empty tubes are compared and studied to demonstrate the advantages and disadvantages of foam material. All graphics are established using GRAPHER software program.
Teklif edilmiş sudan yer gözlem uydusu için haberleşme sistemi ön tasarımı
Sudan is one of the richest countries in the world in terms of natural resources, agricultural fields in particular. The cost of monitoring these resources from space is very high. Establishing a national satellite mission has large initial cost, but it has a large significance in the long run. SudaSat-1 is a proposed earth observation satellite mission funded by the government of Sudan to obtain high resolution images of various areas of the country and other countries as per request. Establishing reliable communication is an essential function for any spacecraft; the failure of communication means loss of mission, the goal of this thesis is to provide a preliminary design for the communication subsystem for SudaSat-1 satellite. The thesis provides a review of similar satellite missions to gain insight of the main considerations involved in the design. The review focused on the satellites with the same spatial resolution. The main characteristics of remote sensing satellites is reviewed, image acquisition modes are also explained. Satellite baseline mission design is introduced, a satellite platform is selected to satisfy overall mission requirements, imaging sensors specifications together with satellite orbit characteristics are also determined to provide the required ground sampling distance. Satellite orbit is designed and simulated using STK, orbit period and average communication time is determined. The thesis highlights mission design elements that influence the design and implementation of the communication subsystem. Communication subsystem-level requirements, functions and design components are overviewed. Components selection is based on requirements satisfaction and space heritage. Redundancy of critical components is put into consideration to increase the reliability of the subsystem. Separate downlink channels are set for imagery data (Xband) and engineering data (S-band). Satellite link power budget is calculated for both downlink channels and for uplink channel. An acceptable value of power margin is obtained. Communication subsystem architecture is presented showing the interconnection between main subsystem components. Satellite data transfer plan is determined illustrating the steps followed to transfer imagery and engineering data downlink and commands uplink
Teklif edilmiş sudan yüksek çözünürlüklü yer gözlem uydusu için uydu yönetim sistemi ön tasarımı
For Sudan as a country newly started its local researches in the field of space science, aerospace and satellite engineering, a preliminary design of a remote sensing satellite is to be made. This thesis is focusing in the satellite management subsystem and the mission design. In the beginning a review was done to the previous experiences in the field as a guide for the design, then the concept was surveyed, also the Remote Sensing basics and imaging process, the orbital dynamics and the management components and elements. A satellite platform was chosen to serve the mission, all supporting subsystems' requirements are agreed to be satisfied. A general revision for the baseline design was done. This platform is considered to be a Chinese one because the opportunity to start this project with china is relatively high, and if the project is implemented, it is most likely to be launched by a Chinese satellite launch provider. The platform is mainly made for Remote Sensing projects and it has a heritage in space. Then the mission design was started with naming the objectives and requirements for the satellite as all, from it the orbit parameters have been decided. Sensor specifications those can take the required high resolution images were selected. The operation modes have been determined to allow the calculations for the management subsystem. The satellite will be orbiting at an altitude of 650 kilometers in a polar sun-synchronous orbit. The satellite has to have the capability of taking two images per orbit. The images can be high resolution with narrow swath width or ten times low resolution with wider swath width. The architecture of the management subsystem was defined based on the mission and previous experiences. Mainly the centralized design is used. The OBC specification was set and the onboard computer software's requirements were cleared. For simplifying the design a separate computer is dedicated for the attitude determination and control operations. Telecommand types are differentiated with the priority. The data management processes were unclouded, a calculations for the data size and type was done. This allowed the minimum storage size to be known. The compression is done for all images while the encryption is an optional process. SpaceWire was picked out to be the internal bus because the high data rate needed when dealing with high resolution images taken by a high speed moving camera. The final design conclude a satellite with 1m GSD (ground sample distance) resolution for panoramic and a 4m GSD resolution for multispectral channels which is orbiting 650 km polar sun-synchronized orbit. The onboard computer has 2 GHz clock with storage can save up to 128 Gbytes. 60 Gbytes can be downloaded every pass over the main ground station. The rest of memory will be used as a redundancy storage and as a storage for images those will be downloaded to other ground stations. The main operation modes of the satellite are the acquisition mode, the battery charging mode, the recovery mode, the idle mode and the execution of a telecommand mode.
Zaman serisi analizi yöntemi kullanarak uçaklar için tahmine dayalı bakım: Trend eğilimleri ve gelecek tahminleri
Bu tez çalışması, havacılıkta uçuş emniyeti için kritik öneme sahip motor, flap ve yardımcı güç ünitesi (APU) sistemlerinde meydana gelen arızaların önceden tespitine yönelik tahmine dayalı bakım modelinin geliştirilmesini amaçlamaktadır. Araştırmada, Boeing 737 MAX8 uçaklarından elde edilen uçuş verileri kullanılmış; bu veriler ön işleme adımlarından geçirilerek zaman serisi analizi yöntemleriyle modellenmiş ve tahmin sonuçları MAE ve RMSE hata ölçütleriyle karşılaştırılmıştır. Ayrıca Power BI tabanlı geliştirilen etkileşimli rapor ile elde edilen sonuçlar görselleştirilmiş ve karar destek mekanizması için kullanılabilir hale getirilmiştir. Çalışma, havacılıkta yalnızca motor verilerine odaklanan geleneksel yaklaşımların ötesine geçerek flap ve APU sistemlerini de kapsayan bütünsel bir model sunmuş; uçuş emniyetini artırma, bakım maliyetlerini azaltma ve Aircraft on Ground (AOG) sürelerini kısaltma potansiyeli ortaya koymuştur. Elde edilen sonuçlar, tahmine dayalı bakımın havacılık sektöründe operasyonel verimlilik ve emniyet açısından önemli katkılar sağlayabileceğini göstermektedir.