LEE- Uçak ve Uzay Mühendisliği-Yüksek Lisans

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http://hdl.handle.net/11527/19474

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Failure analysis of adhesively bonded cfrp joints

(Graduate School, 2021-01-04) Daylan, Seda ; Mecitoğlu, Zahit ; 511171169 ; Aeronautical and Astronautical Engineering

Joints are critical areas where load transfer occurs and should be designed to provide maximum strength to the structure. The adhesive bonding process is widely used as a structural joining method in aerospace applications. There are many advantages of using adhesively bonding joints instead of classical mechanical fastening. Some of these can be listed as joining of similar and dissimilar materials (metal-to-composite, metal-to-metal, metal-to-glass), providing a more uniform stress distribution with a significant decrease in the stress concentration in the structure since there will be no fastening holes, a considerable weight gain compared to mechanical fasteners, strong in terms of fatigue strength due to the absence of fastener holes in the structure. In addition to the above-mentioned positive aspects of using adhesives as a structural joining method, strength prediction is vital for an optimum design process in the initial sizing and critical design phases. The fact that adhesively bonded joints have various failure modes makes failure predictions complex. According to ASTM D5573, adhesively bonded composite joints have seven typical failure modes, but they can be listed under three main headings: adhesive failure, cohesive failure, and adherend failure. Adhesive failure occurs at the adherend and adhesive interface, and usually, the adhesive remains on an adherend. These failures are generally attributed to the poor-quality bonding process, environmental factors, and insufficient surface preparation. The other kind of failure, adherend failure, occurs when the structural integrity of the adherend breaks down before the joint structure and means that the strength of the joint area exceeds the strength of the adherend. On the other hand, cohesive failure is the type of failure expected after an ideal design and bonding process, where failure occurs within the adhesive structure. After cohesive failure, the adhesive material is seen on both adherends. Structural joining with adhesive has been used in the aerospace industry since the early 1970s and 1980s. Since these dates, many analytical and numeric methods have been used to study the failures of adhesively bonding joints. Analytical method studies to analyze the failures of adhesively bonded single lap joints, known in the literature, started with Volkersen in 1938. Volkersen did not include the eccentricity factor in the calculations due to the geometric nonlinearity of the single lap joint. This factor was first taken into account by Goland and Reissner in their calculations in 1944. Goland and Reissner made a remarkable study in analysing the adhesively bonded single lap joint, calculating the loads in the joint area and subsequently the stress on the adhesive. Afterwards, analytical method studies were continued by Hart Smith, Allman, Bigwood & Crocombe and more. In addition to analytical method studies, the continuum mechanic approach, fracture mechanic approach, and damage mechanic approach can be given examples to the numerical method studies. The fracture mechanics approach used in this thesis examines the initial crack propagation in the adhesive under three different loading modes. Crack propagation occurs when the adhesive's critical strain energy release rate equals the strain energy release rate under that load. After the three different modes' strain energy release rate values are calculated separately, an evaluation is made according to the power-law failure criterion. There are many types of joint configurations in the literature, and the common ones can be summarized as single lap joints, double lap joints, stepped joints etc. The single-lap joint type is the most widely used joint type in terms of ease of design and effectiveness. Within the scope of this thesis, it is aimed to obtain a general solution that can be applied to all joints after first making a study for the single lap joint geometry and validating the results of this study experimentally. Studies have been carried out to predict the failure load of adhesively bonding CFRP joints. They include two main steps, which are to find the loads at the edges of the joint area and to evaluate the failure criteria by calculating the strain energy release rate with these loads. As the first step, loads at the joint edges are found analytically and with the finite element method, respectively. While calculating the loads analytically, the Modified Goland and Reissner theory is used, which differs from the classical Goland and Reissner theorem by taking the adhesive thickness into account. While calculating the loads with the finite element method, the modelling technique first studied by Loss and Kedward and then described by Farhad Tahmasebi in his work published with NASA is used. The primary purpose of using this modelling technique is to simulate load transfers in overlap regions accurately for complex and analytically challenging to calculate geometries. Especially in aerospace, since modelling the large components with solid elements is not effective in terms of time and resources, a practical modelling technique that can produce results with high accuracy is needed. In the modelling technique used in the thesis, adherends are modelled with shell elements while the adhesive region is modelled between coincident nodes with three spring elements to provide stiffnesses in the shear and peel directions, and the nodes of the adhesive elements are connected to the adherends with rigid elements. The modulus values of the adhesive material are used in the stiffness calculation of the spring elements. After obtaining the loads with the analytical and finite element method, the second step, the calculation of the strain energy release rate values on the adhesive material, is carried out with reference to two different studies. Firstly, linear fracture mechanics formulations were studied by Williams, assuming that the energy required to advance an existing crack unit amount is equal to the difference of performed external work with internal strain energy, and the laminate containing crack performs linear elastic behaviour is used. Conventional beam theory is used for the 1D case, as the deformation will occur like beam deformation. Using beam theory, he formulated the external work and internal strain energy at the beginning and end of the crack. And using these two equations, he found energy release rate formulations in relation to bending moment and axial load. Then, mode separation is made to calculate the energy release rates in the mode I and II directions separately because the critical strain energy release rate value in these two directions is different and needs to be evaluated independently. This study's disadvantage is that the transverse shear load is ignored, and calculations are made only with bending moment and longitudinal force. Within the scope of the thesis, the strain energy release rate is calculated both with the loads found analytically and with the loads found by the finite element method. Shahin and Taheri did the other reference work, and with overlap edge loads, the stress on the adhesive first and then the strain energy release rate is calculated. In this study, two assumptions are made, and the first is that the shear and peel stress change is zero along with the thickness of the adhesive, and the other is that the stress on the adhesive is as much as the displacement difference of the adherends. As a result of the derivations, the stress distribution on the adhesive is found in the joint structure consisting of CFRP adherend and adhesive. Then, according to Irwin's VCCI approach, as if there is a virtual crack, the integration crack length is rewritten so that it converges to zero and the displacements are in stress. Thus, the stress and energy relationship equation is obtained, and strain energy release rates in mode I and II directions of the adhesive are calculated. As a result of all these studies, mode I and mode II strain energy release rate calculations are made according to two different methods with the loads found analytically and with the finite element method. The strain energy release rate values found and the critical strain energy release rate values, which are allowable, are evaluated according to the power-law failure criteria, and failure load predictions are made. For specimens with different overlap lengths, experimental failure load values and predicted failure load values are compared, and inferences are made about the accuracy of the FEM modelling technique and the methods used in SERR calculation. All these results are interpreted in detail, and it is obtained that the FEM modelling technique gives high accuracy results with Method 2 used in SERR calculation. Finally, a bonding analysis tool has been developed with the python programming language. This tool first detects the finite elements corresponding to upper and lower adherends in the model from NASTRAN .bdf file. Then reads the element loads from the .pch file, which is a NASTRAN output and contains the element loads, then calculates the SERR using Method 2 and calculates the reserve factor and failure load, respectively. This tool has been prepared so that these calculations can be made in a short time and accurately for tens of elements in the overlap zone in complex and large models.
Experimental investigation of leading edge suction parameter on massively separated flow

(Graduate School, 2021-05-10) Aydın, Egemen ; Yıldırım Çetiner, Nuriye Leman Okşan ; 511171150 ; Aerospace Engineering ; Uçak ve Uzay Mühendisliği

The study aims to investigate and understand the Leading Edge Suction Parameter (LESP) application on the massively separated flow. The experiment was done by gathering force data from the downstream flat plate and the visualization of the flow structures is done by Digital Particle Image Velocimetry. The experiments are conducted in free surface, closed-circuit, large scale water channel located in Trisonic Laboratory of Istanbul Technical University's Faculty of Aeronautics and Astronautics. The velocity of the tunnel is equal to 0.1 m/s which results in a 10.000 Reynolds Number. During the experiment, the flat plate at the downstream of the gust generator (plat plate) is kept constant angle of attack and the test cases are selecting to show that the LESP parameter that derived from only one force component works for different gust interaction with the flat plate. As already discussed in the literature, the critical LESP parameter depends on only airfoil shape and its ambient Reynolds Number. Also, the critical LESP number is calculated in literature as equal to 0,05 for plat plate at the 10,000 Reynolds Number. We did not perform an experiment to find critical LEPS numbers as our experiment was done with a flat plate on 10,000 Re. A different angle of attack and different gust impingement combination has been shown that the LESP parameter works even in a highly unstable gust environment. Flow structures around the airfoil leading edge are behaving as expected from the LESP theory (leading-edge vortex separation and unification).
Implementation of propulsion system integration losses to a supersonic military aircraft conceptual design

( 2021-10-07) Karaselvi, Emre ; Nikbay, Melike ; 511171151 ; Aeronautics and Astronautics Engineering ; Uçak ve Uzay Mühendisliği

Military aircraft technologies play an essential role in ensuring combat superiority from the past to the present. That is why the air forces of many countries constantly require the development and procurement of advanced aircraft technologies. A fifth-generation fighter aircraft is expected to have significant technologies such as stealth, low-probability of radar interception, agility with supercruise performance, advanced avionics, and computer systems for command, control, and communications. As the propulsion system is a significant component of an aircraft platform, we focus on propulsion system and airframe integration concepts, especially in addressing integration losses during the early conceptual design phase. The approach is aimed to be appropriate for multidisciplinary design optimization practices. Aircraft with jet engines were first employed during the Second World War, and the technology made a significant change in aviation history. Jet engine aircraft, which replaced propeller aircraft, had better maneuverability and flight performance. However, substituting a propeller engine with a jet engine required a new design approach. At first, engineers suggested that removing the propellers could simplify the integration of the propulsion system. However, with jet engines for fighter aircraft, new problems arose due to the full integration of the propulsion system and the aircraft's fuselage. These problems can be divided into two parts: designing air inlet, air intake integration, nozzle/afterbody design, and jet interaction with the tail. The primary function of the air intake is to supply the necessary air to the engine with the least amount of loss. However, the vast flight envelope of the fighter jets complicates the air intake design. Spillage drag, boundary layer formation, bypass air drag, and air intake internal performance are primary considerations for intake system integration. The design and integration of the nozzle is a challenging engineering problem with the complex structure of the afterbody and the presence of jet and free-flow mix over control surfaces. The primary considerations for the nozzle system are afterbody integration, boat-tail drag, jet flow interaction, engine spacing for twin-engine configuration, and nozzle base drag. Each new generation of aircraft design has become a more challenging engineering problem to meet increasing military performances and operational capabilities. This increase is due to higher Mach speeds without afterburner, increased acceleration capability, high maneuverability, and low visibility. Tradeoff analysis of numerous intake nozzle designs should be carried out to meet all these needs. It is essential to calculate the losses caused by different intakes and nozzles at the conceptual design of aircraft. Since the changes made after the design maturation delay the design calendar or changes needed in a matured design cause high costs, it is crucial to accurately present intake and nozzle losses while constructing the conceptual design of a fighter aircraft. This design exploration process needs to be automated using numerical tools to investigate all possible alternative design solutions simultaneously and efficiently. Therefore, spillage drag, bypass drag, boundary layer losses due to intake design, boat-tail drag, nozzle base drag, and engine spacing losses due to nozzle integration are examined within the scope of this thesis. This study is divided into four main titles. The first section, "Introduction", summarizes previous studies on this topic and presents the classification of aircraft engines. Then the problems encountered while integrating the selected aircraft engine into the fighter aircraft are described under the "Problem Statement". In addition, the difficulties encountered in engine integration are divided into two zones. Problem areas are examined as inlet system and afterbody system. The second main topic, "Background on Propulsion," provides basic information about the propulsion system. Hence, the Brayton cycle is used in aviation engines. The working principle of aircraft engines is described under the Brayton Cycle subtitle. For the design of engines, numbers are used to standardize engine zone naming to present a common understanding. That is why the engine station numbers and the regions are shown before developing the methodology. The critical parameters used in engine performance comparisons are thrust, specific thrust and specific fuel consumption, and they are mathematically described. The Aerodynamics subtitle outlines the essential mathematical formulas to understand the additional drag forces caused by propulsion system integration. During the thesis, ideal gas and isentropic flow assumptions are made for the calculations. Definition of drag encountered in aircraft and engine integration are given because accurate definitions prevent double accounting in the calculation. Calculation results with developed algorithms and assumptions are compared with the previous studies of Boeing company in the validation subtitle. For comparison, a model is created to represent the J79 engine with NPSS. The engine's performance on the aircraft is calculated, and given definitions and algorithms add drag forces to the model. The results are converged to Boeing's data with a 5% error margin. After validation, developed algorithms are tested with 5th generation fighter aircraft F-22 Raptor to see how the validated approach would yield results in the design of next-generation fighter aircraft. Engine design parameters are selected, and the model is developed according to the intake, nozzle, and afterbody design of the F-22 aircraft. A model equivalent to the F-119-PW-100 turbofan engine is modeled with NPSS by using the design parameters of the engine. Additional drag forces calculated with the help of algorithms are included in the engine performance results because the model is produced uninstalled engine performance data. Thus, the net propulsive force is compared with the F-22 Raptor drag force Brandtl for 40000 ft. The results show that the F-22 can fly at an altitude of 40000 ft, with 1.6M, meeting the aircraft requirements. In the thesis, a 2D intake assumption is modeled for losses due to inlet geometry. The effects of the intake capture area, throat area, wedge angle, and duct losses on motor performance are included. However, the modeling does not include a bump intake structure similar to the intake of the F-35 aircraft losses due to 3D effects. CFD can model losses related to the 3D intake structure, and test results and thesis studies can be developed. The circular nozzle, nozzle outlet area, nozzle throat area, and nozzle maximum area are used for modeling. The movement of the nozzle blades is included in the model depending on the boattail angle and base area. The works of McDonald & P. Hughest are used as a reference to represent the 2D-sized nozzle. The method described in this thesis is one way of accounting for installation effects in supersonic aircraft. Additionally, the concept works for aircraft with conventional shock inlets or oblique shock inlets flying at speeds up to 2.5 Mach. The equation implementation in NPSS enables aircraft manufacturers to calculate the influence of installation effects on engine performance. The study reveals the methodology for calculating additional drag caused by an engine-aircraft integration in the conceptual design phase of next-generation fighter aircraft. In this way, the losses caused by the propulsion system can be calculated accurately by the developed approach in projects where aircraft and engine design have not yet matured. If presented, drag definitions are not included during conceptual design causing significant change needs at the design stage where aircraft design evolves. Making changes in the evolved design can bring enormous costs or extend the design calendar.
Helikopter yer rezonansı kararsızlığının çözümü

(Lisansüstü Eğitim Enstitüsü, 2022) Keser, Oğuzhan ; Türkmen, Halit Süleyman ; 736937 ; Uçak ve Uzay Mühendisliği Ana Bilim Dalı

Bu çalışmada seçilen örnek bir helikopterin "Aérospatiale Gazelle" yer rezonansı analizi yapılmıştır. Rotorun ağırlık merkezinin pallerdeki ilerleyen pal hızı ile gerileyen pal hızının farkından dolayı dönme ekseninden kaçması sonucu bir merkezkaç kuvveti oluşur. Oluşan bu kuvvetin frekansı gövde frekansı ile çakıştığı durumda yer rezonansı durumu meydana gelir. Bu rezonans durumu bir kararsızlık/stabilite problemidir ve büyük genliklere sebebiyet verebilir. Böyle bir durumda helikopter ana yapısında büyük kırımlara yol açabilir. Bu tezde önce literatür araştırmasıyla yapılan çalışmalar incelenmiştir. Literatürdeki ağırlıklı çalışmalar ise lineer olmayan sönüm önerileri olan çalışmalardır. Eş merkezli çift rotor sistemin rezonansı, iniş takımlarının rezonansa etkisi veya lineer olmayan diferansiyel denklemler aracılığıyla yer rezonansındaki yapılan çalışmalardan bazılarını oluşturmaktadır. Yer rezonansını daha iyi kavrayabilmek için önce titreşim teorisi basit olarak ayrık sistemlerde ve serbest titreşim altında incelenmiştir. Serbest titreşim altında sönümsüz ve sönümlü sistemin doğal frekansları özdeğer çözümü ile bulunmuştur. Zorlanmış titreşim altında sistemin cevabı matematiksel olarak elde edilmiştir. Sonrasında ise ayrık sistemden farklı olarak sürekli bir sistemde kiriş örneği ele alınmış, sürekli sistemin genel kısmi diferansiyel denklemi zamana ve konuma göre çözülmüştür. Sürekli sistemin zorlanmış titreşim altındaki yer değiştirmesi de genel diferansiyel denklemden türetilmiş ve örnek bir kiriş problemi ele alınarak sistemin frekans ve deplasman cevabı analitik olarak hesaplanmış, ABAQUS (sonlu elemanlar modeli) ile de doğrulanmıştır. Sonrasında ise Nahas'ın Coleman'ın denklemlerinden yararlanılarak oluşturduğu sekiz serbestlik dereceli (altı gövde frekansı, iki rotor frekansı) ile dinamik denklemin özdeğer çözümü gösterilmiştir. Kurulan bu dinamik denklemde kütle matrisi helikopter kütlesinden ve üç eksendeki (x-y-z) atalet terimlerinden oluşmaktadır. Berklik matrisi ise gövdeye kıyasla iniş takımlara daha esnek olduğu için iniş takımının terimlerinden oluşmaktadır. Sönüm matrisi ise daha sonra ortaya çıkacak olan kararsızlıktan sonra gövde frekansı, rotor hızı ve palin kütlesinin helikopter kütlesine oranından hesaplanacak bir matristir. Seçilen Gazelle helikopterinin dış yapısı için bir yüzey çizimi bulunmuş ve bu yüzey kullanılarak sonlu elemanlarda helikopterin dış geometrisinin modeli oluşturulmuştur. Daha sonra bu modele floor, basınç duvarı, longeron ve frame ler eklenerek gerçek bir helikopter modeline benzetilmeye çalışılmıştır. Helikopterin toplam kalkış ağırlığı bilindiği için motor ağırlığı, pilot ağırlıkları ve kokpitteki diğer ağırlıklar konsantre yük olarak helikopterin içine yerleştirilmiş ve böylece gerçek helikopterdekine yakın bir ağırlık dağılımı ve gerçekçi bir ağırlık merkezi hedeflenmiştir. Buradaki amaç helikopterin atalet değerlerini ve ağırlık merkezinin konumunu gerçeğe en yakın temsil etmektir. Pale ait değerlerin ise literatürdeki çalışmalardan alınmıştır.
Değişken malzemeli ve noktasal kütle taşıyan kirişlerin termal etki altındaki titreşim davranışının incelenmesi

(Lisansüstü Eğitim Enstitüsü, 2022) Kıroğlu, İbrahim ; Kaya, Metin Orhan ; 769232 ; Uçak ve Uzay Mühendisliği Bilim Dalı

Kiriş yapılarının başta havacılık olmak üzere otomotiv ve inşaat gibi birçok sektörde oldukça yaygın bir kullanım alanı vardır. Genel olarak, boyuna ve enine dik yükleri destekleyen kiriş yapılarının uzunluğu kesit ölçülerine göre oldukça büyüktür. Kullanım alanı oldukça geniş olan kiriş yapılarının analizinde çeşitli metotlar ve teoriler bulunmaktadır. Bu çalışma kapsamında Euler-Bernoulli ve Timoshenko kiriş teorileri incelenmiş, kirişlerin titreşim davranışının incelenmesi amacıyla titreşim denklemleri elde edilmiştir. Bir dizi diferansiyel denklemden oluşan titreşim denklemlerinin çözümü için oldukça yaygın ve pratik bir çözüm yöntemi olan Diferansiyel Dönüşüm Metodu (DDM) kullanılmış, sonuçlar analitik çözüm ile kıyaslanmıştır. Diferansiyel Dönüşüm Metodu için MATLAB kodu oluşturulmuş, çözümler geliştirilen kod yardımıyla elde edilmiştir. Bu teorilerin haricinde günümüzde kullanımı yaygınlaşan Sonlu Elemanlar Yöntemi (SEY) ile de sonuçlar elde edilmiş ve analitik sonuçlar ile karşılaştırma yapılmıştır. Sonlu Elemanlar Yöntemi'nde ABAQUS paket programı kullanılmış, kiriş modellemeleri bu yazılım ile yapılmıştır. Uçak yapıları, yüksek mukavemete ve yorulma dayanımına sahip oldukça hafif yapılardan oluşmaktadır. Bu bağlamda tez çalışmasında incelenmek üzere kiriş yapı malzemeleri olarak çeşitli metalik ve kompozit malzemeler belirlenmiştir. Kompozit malzemelerin elastik karakteristiklerinin belirlenmesi için kompozit teorisi kullanılarak laminaların mikromekanik ve makromekanik analizi yapılmıştır. Sıcaklıkla ilgili fenomenlerin yapılar üzerindeki etkisi oldukça geniş bir çalışma alanına sahiptir. Sıcaklıktaki değişim, kirişin titreşim davranışında büyük bir farklılığa sebep olabilir. Bu tür yapıların dinamik davranışları yapının termal genleşmesine ve malzeme özelliklerine bağlı olarak sıcaklık etkisiyle değişmektedir. Bu tez çalışmasında sıcaklık değişiminin kirişlerin titreşim davranışına etkisini incelemek amacıyla, elde edilen titreşim denklemine sıcaklık terimi eklenmiş ve çözümler yinelenmiştir. Belirlenen beş farklı sıcaklık değişimi için sonuçlar karşılaştırmalı olarak verilmiştir.
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.
Experimental investigation of underexpanded transverse jet interaction with supersonic crossflow

(Graduate School, 2022) Malkoçoğlu, Utkun Erinç ; Yıldırım Çetiner, Okşan ; 732995 ; Aeronautical and Astronautical Engineering Programme

Transverse jet interaction with crossflow is one of the most canonical and studied flow phenomena in fluid dynamics. It is possible come across to this kind of flow even in nature and daily life; smoke blowing out of the chimney in windy weather could be given as a typical example. Despite different flow structures observed, this interaction is matter of interest in both subsonic and supersonic crossflows. As regards to aerospace discipline, this particular flow event becomes prominent in high-speed applications. Two main research area comprises most of the studies in this context; lateral thrust vectoring without any control surface of supersonic missiles and effective injection for fuel mixing in combustion chamber of supersonic combustion ramjets which have key role in modern aerospace systems. Even minor changes in incidence angles of supersonic missile control surfaces could lead to complex flow structures and prevent effective maneuverability. Lateral control with jets injected to supersonic crossflow becomes advantageous on account of much less response time in comparison with conventional control surfaces. On the other hand, maxiumum mixing of fuel with flow in combustion chamber is a must for scramjet engines. Flow structures, their stability and penetration to flow domain is very crucial. Furthermore, pressure gradients; thus, momentum losses and their minimization in the downstram region of the jet should be evaluated. Realization of jet interaction with supersonic crossflow will contribute to competitive new-generation aerospace solutions. As a result of interaction, various zones occur in both upstream and downstream regions of the jet. The most distinct one is the bow shock which occurs at a certain upstream distance from the jet exit. High-pressure jet behaves just like an obstacle against the crossflow and forces it to go around. Since jet mixing evolves with increasing distance from the surface, its dominance and resistance against crossflow diminishes. As a result, bow shock is bent towards the surface. Bow shock induces flow separation with adverse pressure gradient. Therefore; a recirculation zone, and more importantly, a horseshoe vortex is observed. On the other hand, jet accelerates by expansion in the vicinity of injection surface. Then, it is surrounded by a barrel shock, which defines compression. As a result, supersonic nature of the jet concludes with a Mach disk which is normal to trajectory of the jet. In downstream of the Mach disk, vortical events take place. The most characteristic one is a counter-rotating vortex pair which enlarges as distance to jet exit increases. When surface activities are inspected, a V-shaped separation zone draws attention. When moved to streamwise symmetry axis of the jet, reattachment occurs. In streamwise direction, this zone terminates and then, reflection shocks appear. Briefly, flow structures as a result of jet interaction with supersonic crossflow are summarized as such.
Experimental and numerical studies on low velocity impact behavior of Glare panels

(Graduate School, 2022) Mazı, Oğuzhan ; Doğan, Vedat Ziya ; 775917 ; Aeronautical and Astronautical Engineering Programme

The aerospace industry is always striving to improve aircraft efficiency and strength capacities. New materials are constantly being researched in order to produce more durable and lighter structural parts. Glare materials, which are obtained by laying glass fiber resins in certain directions between thin aluminum sheet metal parts, are very promising, especially in terms of fatigue and impact damage resistance. Glare, which was used in the fuselage panels of the Airbus A380 aircraft, has provided many advantages to the aircraft in many respects. It is known that aircraft are subjected to many low-speed impact damages during their manufacturing and service life such as tool drop, impact of foreign objects on the runway. These damages are also a design criteria to be considered during aircraft design. In this thesis, low velocity impact tests were performed on test specimens made of Glare 4A-2/1-0.3 material in accordance with the standards. Calibration tests were performed to determine the critical damage level and then verification tests were performed to examine the critical energy level. At the same time, a numerical model was prepared with the finite element method to verify the tests in Abaqus/Explicit program. In the model, three-dimensional solid elements were used and the interlaminar behavior was created using cohesive surfaces. The solution algorithm of the Abaqus/explicit program was created with the help of a VUMAT code and the damage criterion for composite plates was embedded in this code. The results of the simulation studies were compared with the results of the experimental studies and consistent results were observed within certain error rates. After the numerical results were verified, the finite element models were updated and the effects of various parameters such as plate thickness, energy level, metal thickness and impact angle on low velocity impact damage were investigated. As a result of the studies carried out, the parameters examined were evaluated and preliminary evaluations were made regarding the use of Glare 4A-2/1-0.3 material for aircraft structure in terms of low-velocity impact resistance. Considering the low velocity impact damage, it was concluded that Glare 4A materials can be evaluated in addition to traditional metallic structures and composite structures at the material selection stage for aircraft structural design. The studies concluded that increasing laminate thickness results in more lightweight structures than increasing outer Al thickness. Moreover, considering oblique impact conditions, it was seen that dent depth and panel failure is proportional to the impact perpendicularity. Finally, it was stated that there are many research areas that need to be examined regarding Glare materials. Some suggestions for the future research and studies were mentioned.
Mechanical response of the carbon fiber reinforced polymer composite sandwich structures with pyramidal lattice core

(Graduate School, 2022) Önal, Gürkan ; Mecitoğlu, Zahit ; 732994 ; Aeronautics and Astronautics Engineering Programme

Composite materials are widely used for many years in various industries. Depending upon technological developments, fiber reinforced composite materials such as carbon, glass and aramid fiber, have been introduced. Also, they can be classified based on the matrix type or the reinforcement type. Matrix-based composite materials can be divided by ceramic, organic, metal matrix; while, reinforcement based classification covers fiber-reinforced, particulate and structural composite materials. It is also explicitly known that sandwich structures are established by a core, upper and lower skin (a.k.a facing or face sheet). Core of a sandwich structures conventionally is chosen as honeycomb or foam. The sandwich structures which includes foam or honeycomb, are remarkable common as structural composite. Surely, this yields having plenty of research in the literature. In the context of thesis herein, a special group of structural composites rather than foam or honeycomb, have been investigated. It is the carbon fiber reinforced polymer composite sandwich structures with pyramidal lattice (CPL) core. Along with demanding to high stiffness or strength to weight ratio, it is needed to arise to replace conventional cores by their counterparts. Promising ones might be given as Kagom'e core, X-Type core, Y-Type core, V Type core, tetrahedral, diamond textile, diamond collinear, square collinear, pyramidal. As to CPL core, it has been examined by a number of researcher from different perspectives such as compressive behaviour under quasi static loading, shear behaviour, bending behaviour, enhancing its analytical models, improving its manufacturing method, hierarchical CPL core, its node designing, searching appropriate failure criteria or etc. It is also indicated that works related to the CPL core in literature are not extensive. As the other core types, the CPL core could be characterized by its relative density. This parameter is basically defined as the ratio of volume occupied by the material within a cell to volume of the cell. Also, it determines failure mode of the CPL core. For instance, the CPL core with lower relative density, tends to fail by means of Euler buckling in the case of compression loading. However, it is possible to have delamination failure if the CPL core has higher relative density. Within the scope of the current thesis a relative density formulation has been derived. This is exact definition despite of that approximate formulations have been introduced. Moreover, it is note that the CPL cores which have been studied in present thesis, have square cross section. This fact has been holds by the exact definition of relative density. Also, it can be indicated that relative density is a parameter which can be utilized by comparing any type of core between each other. Subsequently, two different mechanical behaviour of sandwich structures with the CPL core, have been studied in this work such as out of plane compression and flexural based shear. For each behavior, the specimens have been designed so as to have two different relative densities like 2.863% and 0.725%. While the former relative density stands for Design 1, the latter relative density represents Design 2.
Design and optimization of two stage launch vehicles with the same liquid propellant rocket engines in both stages

(Graduate School, 2022) Özçelik, Kubilay ; Aslan, Ali Rüstem ; 714559 ; Aeronautical and Astronautical Engineering Programme

Space exploration is an important technological catalyst for humanity. While researching space and its practical uses, it accelerates the development of new technologies. Reaching orbit is a difficult and complex problem. To get the speed required to stay in orbit, launch vehicles need to have very high propellant mass fraction ratios and high performing propulsion systems. Reaching the performance limits to reach orbit needs high technology and expensive materials to be used. Because of this it is very expensive to put payload into orbit. In the recent years private space companies are entering to the launch vehicle market. These privately funded companies try to drop the prices to be able to compete with existing launch service companies to insert payloads into orbit. To do so they try to reuse the same liquid rocket engines in all stages to drop the development and manufacturing costs. Most of the private launch vehicle companies are designing only one rocket engine and are using them in their 1st and 2nd stages. While the 1st stage engines are bundled together using engines that have sea level optimized nozzle. The same engine is used in the 2nd stage with a vacuum optimized nozzle. Doing so, they reduce the development costs, complexity and manufacturing costs of their launch vehicle. Also the new trend is to design the launch vehicle as reusable as possible. This allows for cost reductions that make the launch vehicle more competitive in the market. Some companies that use this approach are SpaceX, RocketLab USA and Relativity Space. In this thesis, a launch vehicle optimization tool is developed specifically for two stage to orbit vehicles that use the same liquid propellant rocket engines for all stages with only minor modifications. In the 1st stage many sea level optimized engines are bundled together and in the 2nd stage a single vacuum optimized engine is used. It can design launch vehicles for different propellant combinations and liquid rocket engine cycles. Most launch vehicle design methods estimate the stage properties and try to distribute the mass of the stages based on estimations. After finding a viable solution it is designed in detail and the assumed performances of the stages cannot be achieved. This causes an iterative design loop that is resource draining. To solve this problem in this thesis the liquid propellant engines and stages are designed in detail. Firstly, the liquid propellant rocket engine is designed in detail and after that the stage is created by adding tanks and pressurization system. The stage design tool is connected and implemented such that it can design stages with bundled engines for the 1st stage and modifies the same engine as vacuum optimized for the 2nd stage to create the desired launch vehicle. The stage design tool is connected to an optimization algorithm and launch vehicle design tool to create the specified launch vehicle design tool necessary for this thesis. One of the most important design parameters for a launch vehicle is the required delta V for the selected mission. But without simulating the launch trajectory making a good estimate for required delta V is difficult. Therefore, to validate the designed launch vehicles, an orbital trajectory simulation code is developed based on MATLAB. Using this simulator, the designed launch vehicles are trajectory simulated and if successful they are validated or if they are unsuccessful the design parameters are updated accordingly in launch vehicle design code and the process is repeated to find good performing launch vehicles. Designing a launch vehicle is a complex multi-disciplinary and multi objective problem. To rapidly design the launch vehicle the most important parameters are selected as payload capacity, vehicle delta V capacity and T/W ratio at liftoff. The payload capacity and delta V capacity mostly influence the mass of the launch vehicle. Whereas the T/W ratio at liftoff determines the engine thrust and orbital launch performance of the launch vehicle. The optimization algorithm is developed such that it searches for the launch vehicle with minimum liftoff mass while ensuring the design input parameters are met with minimal error.
Fused filament fabrication of PETG :Investigation of the mechanical properties through the parameter optimization

(Graduate School, 2022) Parlak, Buket ; Cebeci, Hülya ; 736752 ; Uçak ve Uzay Mühendisliği Bilim Dalı

Additive manufacturing methods are in increasing demand every year due to their low cost, production of complex parts, rapid prototyping possibilities, and accessibility, and they can be preferred over other traditional methods (casting, forging). Additive manufacturing; is used effectively in many fields, especially in aviation. In addition, it is available in the literature that wax patterns (wax-patterns) used in precision casting, with its rapid prototyping feature, are obtained by the Fused Filament Fabrication (FFF) method. However, it is seen that the choice of polymer used here is very important. The polymer having high strength, low thermal expansion coefficient, and no causing shrinkage and warping during production are the desired properties. There are wax models produced with PLA and ABS in the literature. It is seen that parts produced with FFF are used not only in prototyping, but also in unmanned aerial vehicles. Additive manufacturing methods are classified according to the type of material as metal, ceramic and polymer-based. According to the ISO/ASTM 52900:2015 standard, material types are also divided into their sub-headings. The basic working principle of additive manufacturing is based on the principle that the feeder (it can be a powder or polymer filament) is melted with the help of a melting source and reassembled on a table at the desired dimensions. First of all, the CAD (Computer Aided Design) models of the part are created as .STL (Standard Triangle Language) file format and it is combined with the parameter information to be used in the printer with the help of a slicer and the g-code is created. This generated g-code is uploaded to the printer and the processes are started. Parameter selections play an important role in determining the mechanical properties of the polymer parts. The most important parameters used in the FFF method are as follows; the infill ratio, the layer height, the layer thickness, the width of the raster, the infill pattern, the air gap ratio, the raster orientation, the build direction, the printer speed, the printer temperature and the nozzle diameter. The choice of polymer type is another important parameter. In this study, PETG polymer was used because of its high resistance to chemicals, fatigue resistance, high toughness, and low shrinkage during production compared to other polymers and its easy production. This study aimed to examine the effects of the negative air gap, selected infill pattern and tensile sample standard, annealing heat treatment temperature and time on tensile properties (Ultimate Tensile Strength (UTS) and Elastic Modulus (E)). For the first parameter set, 60 samples were produced. 20 of these samples were concentric and produced by ASTM D638 Type IV standard. The remaining 20 samples were also concentric ASTM D3039. To examine the infill pattern difference in the last 20 samples, they were produced in rectilinear infill in accordance with the ASTM D3039 standard. All 5 samples were produced to have a 0%, 10%, 15%, and 20% negative air gap. As a result of the comparison of the infill patterns, it was seen that the concentric filling resulted in 29,65-50,54% higher results in E and 33,06%-47,88% higher results in UTS than the rectilinear infill. Another comparison was made between samples produced by ASTM D638 Type IV with concentric infill and 0%, 10%, 15% and 20% negative air gap, and samples produced according to ASTM D3039. According to the comparison of the different infill patterns, the concentric infill samples produced according to ASTM D638 Type IV showed the highest properties of 16,33% in E and 20,69% - 48,16% in UTS compared to those produced according to ASTM D3039. The effect of increased negative air gap was also investigated in both concentric (ASTM D638 Type IV and ASTM D3039) and rectilinear (ASTMD3039) samples. In all comparisons, the samples with a 0% negative air gap were compared with the samples produced with 10%, 15%, and 20% air gaps. As the negative air gap ratio increased, ASTM D638 Type IV concentric samples showed an increase of 11,38% - 31,54% in E and 37,18% - 63,89% in UTS. The effect of the air gap was found to be negative in the concentric filling produced according to the ASTM D3039 standard (as the gap increased, there was a decrease between 2,84% and 10,20% in E, while a decrease between 4,9% and 8,14% in UTS was observed. The effect of the air gap was found to be positive in the rectilinear filling produced according to the ASTM D3039 standard (when the negative air gap increased, there was a decrease between 2,51% and 32,44% in E, and an increase between 6,24% and 17,45% in UTS). According to all these results, the parameter set that gave the best results was obtained in the sample with ASTM D638 Type IV, concentric infill, and 15% negative air gap (E: 1.87 Gpa and UTS: 41,84 Mpa). Another aim of this study is to examine the post-process effects. To examine the effects of annealing heat treatment, 20 samples of ASTM D638 Type IV, concentric and 15% negative air gap were produced. This study was planned for two annealing temperatures and two selected times. The selection of the tensile test specimen is still a controversial issue, and the effect of the two standards was examined and discussed in this study. In this study, the importance of the effect of heat treatment temperature and time on mechanical properties was emphasized. The effects of two temperatures, 80°C, and 55°C, were investigated. At these temperatures, each sample was held in the furnace for 1 hour and 4 hours. Samples that were heat treated at 80°C were first compared with those that were heat treated at 55 °C. The tensile test results of the samples annealed at 55°C for 1 hour are higher 17,94% in E and 13,73% in UTS than the samples kept at 80°C for 1 hour. In the same way, the tensile test results of samples that were heat treated at 55 °C kept for 4 hours, are higher 17,10% in E and 13,67% in UTS than compared to 80°C. In order to see the effect of the time, the temperature was kept constant and the samples were held for 1 hour and 4 hours. According to the results obtained, there was no high increase in E and UTS as the holding time increased. All results were compared with non-heat-treated concentric specimens produced with 15% air gap and treated according to ASTM D638 Type IV. As a result of this comparison, while a 14,32% decrease was observed in E in the samples kept at 80°C 1 hour, this decrease was recorded as 2,16% in UTS. In the samples kept at 55°C 1 hour, it increased up to 4,42% in E and up to 13,41% in UTS. These results were also compared with the data in the literature, and the results were also compatible with the literature. In the samples processed at 80°C for 4 hours, a decrease of 13,77% was observed in E, while this decrease was recorded as 0,39% in UTS. In samples processed at 55°C for 4 hours, it increased by 4,01% in E and 15,38% in UTS. In the literature, 7% increase in E and 6% increase in UTS were obtained in the heat treatment of line infill samples produced with ASTM D638 Type I, 100% infill, and held at 55°C for 1 hour. The reason for the higher increase in literature compared to the samples produced with 100% infill is the effects of the negative air gap. The mechanical properties of samples produced with FFF are always lower than those obtained by injection molding, due to molding defects (like voids) and anisotropy. It is known that due to the nature of the FFF method, there are many voids inside the structure in the parts printed with 100% infill ratio. All the results obtained in this thesis were also compared with the mechanical properties obtained by injection molding. As a result of this comparison, it was observed that the highest difference was in the rectilinear produced samples (57,76%-44,06% in E, 68,96%-61,21% in UTS). In concentric samples produced according to ASTM D3039, this difference was between 23,31% and 14,60% in E and 36,91%-42,05% in UTS. Samples produced according to ASTM D638 Type IV it was found to be lower in 10,78-32,18% in E; 14,14% and 47,6% in UTS compared to the samples produced by injection. It was determined that the results of the samples produced by injection were approached with the annealing heat treatment at 55°C at most. The difference was recorded as 6,84% in E; 5,09% in UTS for 1 hour and 7,21% in E; 3,44% in UTS for 4 hours. The novel approach of this study is that reach the injected molded part results with appropriate parameter optimization studies. After the tensile tests, the fracture surfaces of the samples were also examined, and it was observed that 2 of the 60 samples were fractured in the GAT (G: Failure Type A: Failure Area T: Failure Location) rupture mode. It was observed that 20 samples produced according to ASTM D638 Type IV were broken from the inner narrow length, except for 2 of them. In addition, PETG is an advantageous polymer; no delamination and shrinkage problems were encountered compared to other polymers. In this study, it has been seen that the effects of infill, tensile specimen standard selection and negative air GAP, heat treatment, time, and selected temperature have significant effects on mechanical properties.
Autonomous heading control of a fixed-wing aircraft with deep reinforcement learning

(Graduate School, 2022) Sarıgül, Fatih Ahmet ; Beyazit, İsmail ; 771377 ; Aeronautics and Astronautics Engineering Programme

Autonomous control has become more reachable with recent advancements and new techniques such as deep reinforcement learning (DRL) and is getting more and more popular in every field including flight control. Autonomous flight is an important trait to attain for an aircraft because it provides to get rid of external involvement in control and it has the potential to excel in human skills. In fully autonomous flight, the aircraft is needed to complete its all tasks without any human involvement. However, instead of fully autonomous flight, this work focuses only on heading control because making the heading control autonomous by a learning algorithm means that it is possible to make other flight tasks autonomous by using a similar learning algorithm. Therefore, in this work, devising a learning algorithm for autonomous heading control is the main goal of the work. In this work, it is desired to attain autonomous control for a case that demands a complex environment and dynamics like the heading maneuvers of a fixed-wing aircraft. However, because the focus of the work is mainly at the algorithmic level, it is not a good idea to dive into the complex environment and dynamics directly. Therefore, firstly the Dubins model is used in this work to represent the fixed-wing aircraft while testing the learning algorithm. Dubins model can be simply considered as just a point that has a heading and a velocity. It can be implemented in both 2D and 3D environments easily and it can represent a car driving or an aircraft flight in a very basic manner. However, with some constraints and some addition to its dynamics, it can be converted easily into a good representation of a fixed-wing aircraft. This new representation is much simpler than a fully described dynamics of a fixed-wing aircraft but it is still a good representation to see how the learning algorithm works. Therefore, in this work, a simplified fixed-wing aircraft model is obtained from a 3D Dubins Airplane model and the learning algorithm is tested in this simplified model. Implementation of the learning algorithm to a fully described 6-degree of freedom dynamics is left for future works. It seems that the state-of-art DRL algorithms can provide a solution for the autonomous heading control of a fixed-wing aircraft. So, finding the most appropriate state-of-art algorithm and implementing it to the problem can offer a solution. But, in this work, it is not the way that is followed. Instead of jumping directly to the state-of-art methods, this work starts with a basic learning algorithm, and it is tested in a simple environment, after getting satisfactory results at this level, the environment is rendered more complex status and the algorithm is made more advanced to deal with new conditions. In this way, it can be seen why the previous methods lack and what is needed to make learning possible in this new condition. By repeating this process, it is aimed at obtaining a DRL algorithm to solve the problem while having the opportunity to make improvements in the algorithm at every level. In this work, the DRL algorithm is obtained by using the aforementioned technique for heading control. This DRL algorithm consists of a combination of DQN and Actor-Critic methods. It meets the requirement of dealing with continuous state and action spaces and it has some unique approaches which do not exist in other algorithms. The new algorithm that has been obtained in this work is tested on the Dubins model and simplified model to see its validity and whether it can be used for more complex tasks and dynamics. The promising results show that the algorithm can be enhanced to deal with other flight tasks also, and it may offer solutions to complex real-world problems.
Robust trajectory optimization of constrained re-entry flight via stochastic collocation based ensemble pseudospectral optimal control

(Graduate School, 2022) Selim, Akan ; Özkol, İbrahim ; 768672 ; Aeronautics and Astronautics Graduate Programme

A new computational framework for constrained robust trajectory optimization problems in hypersonic flight and re-entry has been developed, called Stochastic Collocation based Ensemble Pseudospectral Optimal Control (SC-EPOC). Uncertainty space has been simplified by utilizing Sparse Grid methods, one of which is the Conjugate Unscented Transformation. Then, the uncertainty-aware optimal control problem (OCP) is rewritten as a computationally tractable deterministic OCP by utilizing Ensemble Optimal Control. The resulting problem is solved via Pseudospectral Optimal Control Methodology and Nonlinear Programming. A tailored Pseudospectral Optimal Control Software has been developed, and validated against benchmark cases taken from literature with greater results, thanks to the in-house developed hyper-dual differentiation library and sparse calculation of Jacobian and Hessian matrices of the resulting optimization problem for Nonlinear Programming. A mesh-refinement algorithm has been developed for singular OCPs where wild oscillations occur over the boundary arc, which is unacceptable for safety-critical mission design. The algorithm is tested and validated against theoretically derived optimality conditions on a simple thrust programming landing rocket problem. To make SC-EPOC computationally tractable, the software has been upgraded by leveraging vectorization and parallelization technologies for sparse calculation of 3D Jacobian and Hessian matrices. As a result, the software has been shown convergent under high uncertainties with more than a million of optimization variables and constraints. Furthermore, SC-EPOC is applied to a strictly constrained re-entry problem where, a re-entry vehicle is being commanded to steer from an Entry Interface with uncertainties towards a TAEM area under angle-of-attack limits, dynamic load and heat flux constraints while maximizing the cross-range. Problem has been solved for different combinations of uncertainties including model uncertainties and state uncertainties and the results, including the variations of optimized states, control, costates derived by utilizing Covector Mapping Theorem, their boundary values and Hamiltonian have been given. Their optimality conditions have been derived by utilizing ensemble optimal control theory. Computational results showed excellent agreement between theory. In the end, future research direction has been depicted, and a theoretical investigation has been conducted for what is called as the Integrated Ensemble Pseudospectral Guidance and Recovery Control (IEPG&RC) to optimize the robust trajectories by incorporating the control term, that makes it possible to incorporate higher uncertainties and steer the initial distribution towards the same endpoint.
Yüksek devirli rulmanlarda açık yağlama sistemi için sprey enjektör tasarımı

(Lisansüstü Eğitim Enstitüsü, 2022) Yalçın, Erdem Çağatay ; Edis, Fırat Oğuz ; 811151118 ; Uçak ve Uzay Mühendisliği Yüksek Lisans Programı

Hava araçlarının veya roketlerin itki sistemlerinde kullanılabilen turbo jet motorların kullanıldığı aracın planlanan kullanım ömrüne uygun tasarlanabilmesi mühendislik açısından zorlayıcı bir optimizasyon problemidir. Motorlarda döner elemanların yataklandığı rulmanlar sistemin en kritik noktasını oluşturmakta dolayısıyla tasarlanan motorların kullanım ömrünü de bu kritik noktaları olan rulmanların kullanım ömrü belirlemektedir. Rulmanların kullanım ömrünün uzatılması ve daha pürüzsüz bir şekilde görevini yerine getirebilmesi için tarihi taş devrine dayanan, günümüze kadar ilk kullanılmaya başlandığı andan itibaren önemini yitirmemiş, uygarlığımızın temeli olan aletlerin ve makinelerin sorunsuz çalışmasında ve uzun ömürlü olmasında en büyük katkıya sahip; yağlama sistemi kullanılmaktadır. Yağlama sisteminin kalitesini belirleyen ve sistem gereksinimlerine uygun yağlama yapılabilmesini sağlayan en önemli tasarım noktası doğru enjektörün seçimidir. Bu tez çalışmasında yüksek devirli bir rulmanda açık yağlama sistemi gerekliliklerinin belirlenmesi ve bu gerekliliklere uygun yağlama enjektörü tasarımı çalışması anlatılacaktır. Bu tez kapsamında, yağlama enjektörü tasarımına varan ön tasarım çalışmaları 3 aşamada incelenmiş, bu ön tasarım çalışmaları sonucu elde edilen veriler ışığında enjektör tasarımı yapılmış, tasarlanan enjektörün testleri yapılmış ve deneysel sonuçlar elde edilmiş, bu sonuçlar teorik sonuçlar ile kıyaslanmış ve bu kıyaslamaya dayanarak çıkarımlar oluşturulmuştur. Ön tasarım çalışmasının ilk adımı, yağlama gereksinimlerini karşılamak için rulmanların ürettiği ısının belirlenmesidir. Rulmanlarda üretilen ısı, rulman geometrisine, rulman üzerindeki yüke ve rulmanın dönme hızına bağlıdır. Rulmanlarda ısı oluşumu dönen bilyelerden kaynaklanmaktadır. Bu nedenle hesaplamalarda kullanılacak kritik rulman geometrisi bilgileri bilyelerin bulunduğu bölgede toplanır. Rulmana ait gerekli hız ve boyut bilgileri bu geometri bilgileri kullanılarak hesaplanabilir. Yükler ve bilye hızları hesaplandıktan sonra sürtünmeden kaynaklı kaybedilen enerjinin tamamının ısıya dönüştüğü kabulü ile üretilen ısı hesaplanır. Ön tasarım çalışmasının ikinci adımı olarak rulmana gönderilecek olan yağlayıcının debi hesabı yapılmalıdır çünkü rulmanda üretilen ısının sistemden dışarıya atılması gerekmektedir ki rulman sıcaklığının kritik seviyelere ulaşması önlenebilsin. Rulmanın bu kritik seviyelerdeki sıcaklığa ulaşmasını önleyebilmek için rulmana yağlayıcı sıvı gönderilmektedir. Gönderilecek yağlayıcının debisini belirlemek için ısı transferi hesabı yapılabilir. Bu çalışmanın başlangıcında, gönderilen yağlayıcı ve havanın rulman çıkışında aynı sıcaklığa ulaşacağı kabulüyle, rulmanda üretilen ısının hava-yağ karışımına aktarıldığı ısıl denge kurulmuştur. Böylece karışımın rulman çıkış sıcaklığı tespit edilmiştir. Çıkış sıcaklığı belirlendikten sonra rulman sıcaklığını bulmak için iteratif bir çalışma yapılacaktır. Rulmanlara gönderilecek debi belirlendikten sonra, ön tasarım çalışmasının üçüncü ve son adımı doğru tip enjektör seçimidir. Yağlama için kullanılacak enjektörler sprey ve jet enjektörler olmak üzere iki seçenekte incelenmiştir. Enjektör tipine karar verebilmek için literatür taraması yapılarak sprey ve jet enjektör karşılaştırılmıştır. Literatür taramasından elde edilen bulgular göstermektedir ki, hava akışı hızlandığında jet akış büyük tanecikler halinde parçalandığı için hava akışı ile tam anlamı ile taşınamayıp kendi ağırlığı ile akış içerisinde dağılmaktadır. Bu durum jet akışın istenilen hedefe gitmemesine sebep olmaktadır. Rulmana doğru şekilde ulaşmayan yağlayıcının, rulmandan geçmeden ortamdan atılma olasılığı oldukça yüksektir. Ayrıca kapalı sistemlerde sızdırmazlık elemanları sayesinde rulman dışında gidecek bir yere sahip olmayan yağlayıcı, açık sistemde bu sızdırmazlık elemanlarının bulunmaması sebebi ile doğru taşınma kapasitesine sahip olmaz ise, yağlayıcı rulmana ulaşamayacak ve yağlama görevi başarı ile gerçekleştirilemeyecektir. Dolayısıyla bu tezde söz konusu yüksek devirli rulmanlarda açık yağlama sistemi kullanıldığında yağlayıcının tanecik yapısını küçültmek, dağılımını arttırmak ve ikincil hava akışı ile taşınabilir hale gelmesi için sprey enjektör tercih edilmiştir. Literatür çalışmalarında görülen, enjektör tasarımında kullanılan denklemlerin çoğunun ampirik yaklaşımlar varsayılarak yapılan testler sonucunda oluşturulan denklemler olduğudur. Bu tezde, tasarlanacak sistemin çalışma aralığındaki ve düşük hata oranlarına sahip denklemler referans alınmıştır. Bu tezde bahsedilen enjektör için kullanılan tasarım parametreleri; tahliye katsayısı, dağılma uzunluğu, film kalınlığı, yağ çıkış hızı, dağılma rejimleri, sprey dağılma açısı, Sauter ortalama çapı ve yay hesabıdır. Enjektör tasarımı enjektör gövdesi, döner parça, yay, yay tutucu, filtre ve yüksük olmak üzere 6 parçadan oluşmaktadır. Enjektör boyutları, küçük motorlara uyacak şekilde kasıtlı olarak küçüktür. Her ne kadar bu durum üretimsel bazı sıkıntılara sebep verse de kimi motorlar için küçük boyutlara sahip bir enjektör tasarımı zorunluluk olarak ortaya çıkmaktadır. Bu sebeple küçük boyuttaki enjektörlerin incelenmesi ve bu konuda çalışmalar yapılması bir gereklilik olarak ortadadır. Test sisteminde enjektörlerin akış ve basınç ilişkisinin belirlenmesi amaçlanmaktadır. Ayrıca sprey dağılım açısı farklı debilerde gözlemlenmiş ve sprey kalitesinin bir göstergesi olarak kullanılmıştır. Testler kademeli olarak artırılan farklı debilerde gerçekleştirilmiş ve her adımda kamera kayıtları alınarak sprey formları incelenmiştir. Yapılan testler sonucunda düşük basınçlarda püskürtme kalitesinin kötü olduğu ve istenilen debilerin fazla gönderildiği görülmüştür. Düşük basınçta akış enjektörde girdap şeklinde hızlanamadığı için enjektörden çıkan akışkan saçılarak dışarı çıkamaz. Bu nedenle akış birincil rüzgâr rejiminde gerçekleştiği için yeterli sprey formu elde edilememektedir. Sprey kalitesini artırmak için basıncın artırılması gerektiği açıktır. Ancak bu debiyi artıracağından, debi değeri ile sprey kalitesi arasındaki denge gözetilerek tasarım yapılmalıdır.
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.
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.
Fonksiyonel derecelendirilmiş malzemeden üretilen plakların mekanik ve ısıl yükler altındaki burkulma analizi

(Lisansüstü Eğitim Enstitüsü, 2022-01-27) Aktaş, İbrahim Utku ; Doğan, Vedat Ziya ; 511171115 ; Uçak ve Uzay Mühendisliği

Malzeme seçimi bütün mühendislik uygulamalarında çok önemli rol oynamaktadır. Neredeyse bütün mühendislik uygulamalarının gelişmesi ve ilerlemesi o alanda kullanılan malzemelerin gelişmişliği ile doğrudan ilişkilidir. Malzemelerin monolitik malzemeden alaşımlı malzemelere evrimi ve kompozit malzemelerin gelişimi, bir malzeme sınıfının çağın ihtiyaçlarına artık cevap veremiyor oluşundan doğmuştur. Çoğu mühendislik uygulamasında, monolitik bir malzemede bulunması imkânsız olan birbirleriyle çelişen özelliklere sahip malzemelerin kullanımına ihtiyaç duyulmaktadır. Ayrıca, farklı malzemelerin alaşımlanması, bileşen malzemelerin termodinamik davranışı ve bir malzemenin diğer malzemelerle karıştırılma derecesindeki kıstaslar ile sınırlıdır. Fonksiyonel derecelendirilmiş malzeme, iki malzemenin bir araya getirilmesi ve zorlu çalışma ortamlarına maruz kaldıktan sonra dahi işlevlerini yerine getirebilmesi ve özelliklerini koruyabilmesi gerekliliğinden doğmuştur. İşlevsel olarak derecelendirilmiş malzeme başlangıçta bir ısıl bariyer uygulaması ihtiyacı için geliştirilmiş olsa da, bu önemli gelişmiş malzemenin uygulaması artırılmış ve aşırı aşınma direnci ve korozyon direnci uygulamaları gibi mühendislik uygulamalarında bir dizi sorunu çözmek için kullanılmıştır. Bu yeni malzeme türünden havacılık, otomobil ve biyomedikal gibi uygulamalarda yararlanılmaktadır. Fonksiyonel derecelendirilmiş malzemeler, geleneksel kompozit malzemelerin zorlu çalışma ortamlarında kullanıldığında başarısız uygulamalara neden olmasının sonucunda ortaya çıkmıştır. Geleneksel kompozit malzemelerin mühendislik uygulamalarındaki başarısızlığı kompozit malzemeyi oluşturan katmanlar arasındaki belirgin bir şekilde tanımlanmış olan arayüzden kaynaklanmaktadır. Arayüz, bu bölgede yüksek bir gerilme yığılmasına sebebiyet vermekte ve kompozitin nihai başarısızlığına neden olan çatlak başlangıcını ve yayılmasını teşvik etmektedir. Bu çatlak oluşma ve ilerleme sürecine "delaminasyon" adı verilmektedir. Japonya' da bir uzay mekiği projesinde karşılaşılan ve fonksiyonel derecelendirilmiş malzemelerin ortaya çıkmasına ortam hazırlayan sorun, geleneksel kompozit malzemelerdeki bu belirgin arayüzün nasıl ortadan kaldırılabileceğini ve kompozitin istenen ısıl bariyer görevini nasıl yerine getirebileceği problemini ortaya koymuştur. Araştırmacılar, kademeli olarak değişen bir arayüz ile geleneksel kompozit malzemedeki keskin arayüzü sistematik olarak ortadan kaldırabildiler, böylece bu arayüzdeki gerilme yığılmasını azalttılar ve geliştirilen fonksiyonel derecelendirilmiş malzeme, zorlu çalışma koşullarında kırıma uğramadan ayakta kalabildi. Sonuç olarak malzemenin asıl geliştirilme amacı olan yapıya ısıl kalkan olması dışında çeşitli mühendislik uygulamaları için de fonksiyonel derecelendirilmiş malzemeler kullanılmıştır. Fonksiyonel derecelendirilmiş malzemeler, malzemenin hacmi boyunca değişen özelliklerle birlikte değişen bileşime sahip gelişmiş kompozit malzemelerdir. Havacılıkta kullanılan araçlar başta aerodinamik yükler olmak üzere birçok mekanik ve ısıl yüklere maruz kalmaktadır. Bu yükler hava aracının yapısallarının boyutlandırılmasında kullanılmaktadır. Güvenli bir hava aracı maruz kaldığı yükleri yapı içerisinde taşırken kırıma uğramayacak şekilde tasarlanmaktadır. Hava aracının yapısalları birçok farklı şekilde kırıma ya da hasara uğrayabilmektedir. Bunları öngörebilmek ve yapıyı ona göre tasarlamak hayati öneme sahiptir. Bununla beraber, yapıları kırıma uğratmayan fakat yapılarda yapısal kararsızlığa yol açan burkulma problemi havacılıkta çok önemli bir konudur. Örneğin bir uçağa gelen yükler kanat üzerindeki kabukların düzlem içi basma ya da çekme yüklerine maruz kalmasına sebep olabilmektedir. Kabuk elemanlarının basma yüküne maruz kaldığı durumlarda burkulma olayı gerçekleşebilir. Bu da hem kanat üzerindeki aerodinamik akışı bozabilmekte hem de yapının kararsız hale gelmesine sebep olabilmektedir. Bu gibi durumlarda yapının yük taşıma kapasitesi değişmekte ve burkulma sonrası hesaplamaların yapılması gerekmektedir. Bundan dolayı yapısal elemanların ne zaman burkulmaya uğrayabileceğini öngörebilmek büyük önem taşımaktadır. Bu tezde fonksiyonel derecelendirilmiş malzemeden üretilen plakların ısıl ve mekanik yüklemeler altındaki burkulma davranışları sistematik olarak ele alınacaktır. 1. Kısım' da yapılan çalışmadan genel olarak bahsedilip çalışmanın amacından ve isteğinden söz edilmiştir. 2. Kısım' da ise geçmişte fonksiyonel derecelendirilmiş plaklar üzerine yapılmış çalışmalar okuyucuya aktarılmıştır. Bu çalışmaları ifade etmeden önce temel burkulma probleminin tanımı yapılmıştır. Burkulma olayını tanımlamaya ilk olarak kolon ve kiriş elemanlarının burkulmasından başlanmış daha sonra plakların burkulması anlatılmıştır. Burkulma teorisinin alt yapısının okuyucuya bu şekilde verilmesi amaçlanmıştır. Ardından fonksiyonel derecelendirilmiş malzemelere kısa bir giriş yapılmış ve tarihçesinden bahsedilmiştir. Bu kısımda aynı zamanda fonksiyonel derecelendirilmiş malzemelerin burkulması üzerine yapılan akademik çalışmalardan da bahsedilmiştir. 3. Kısım' da fonksiyonel derecelendirilmiş malzemeden üretilen plakların mekaniğini anlamak adına geleneksel kompozit malzemeden üretilen plakların mekaniği okuyucuya aktarılmıştır. İlk olarak katmanlı kompozit plak teorilerinden kısaca bahsedilmiş ve sonra Klasik Kompozit Plaka Teorisi (KPT) ve Birinci Dereceden Kayma Şekil Değiştirme Teorisi (BKT) detaylı bir şekilde anlatılmıştır. Çünkü fonksiyonel derecelendirilmiş malzemeden üretilen plakların mekaniğini anlamak için geleneksel kompozit plakların mekaniğini iyice anlamak büyük önem taşımaktadır. 4. Kısım' da fonksiyonel derecelendirilmiş malzemelerin üretim yöntemlerinden kısaca bahsedilmiş ve etkin malzeme özelliklerinin nasıl modellendiği gösterilmiştir. 5. Kısım' a gelindiğinde daha önceden kısaca bahsedilen plakların burkulma problemi üzerinde durulmuş ve bu problemin belirli sınır koşulları altında analitik çözüm yöntemlerinden bahsedilmiştir. İlk olarak izotropik plakların burkulma probleminin çözümü, Navier ve Levy sınır koşullarını ayrı ayrı sağlayacak şekilde oluşturulan sınır koşulları altında çözülmüştür. Ardından Fonksiyonel derecelendirilmiş malzemeden üretilen plakların burkulma problemini çözebilmek için KPT kullanılarak analitik model oluşturulmuştur. Sonrasında oluşturulan analitik model her bir kenarından basit mesnetli kabul edilen fonksiyonel derecelendirilmiş plaklar için farklı yüklemeler altında MATLAB programında yazılan kod yardımı ile çözülmüştür. Bu yüklemeler mekanik ve ısıl yüklemeler olmak üzere ikiye ayrılmaktadır. Mekanik yüklemeler için üç farklı durum göz önüne alınmıştır. Bunlar: tek eksenli basma yükü, iki eksenli basma yükü ve iki eksenli basma – çekme yükü altındaki burkulma analizleridir. Isıl yükleme koşulları ise sıcaklığın kalınlık boyunca farklı şekillerde dağılımları göz önüne alınarak yine üç farklı şekilde yapıya uygulanacak ve burkulma analizi yapılmıştır. İlk olarak kalınlık boyunca sabit sıcaklık dağılımı için kritik burkulma sıcaklık farkı bulunmuştur. Ardından kalınlık boyunca doğrusal değişen sıcaklık dağılımı için burkulma analizi yapılıp kritik burkulma sıcaklık farkı elde edilmiş ve sonrasında ise kalınlık boyunca doğrusal olmayan sıcaklık dağılımı için bu analizler tekrarlanmıştır. Elde edilen tüm sonuçlar daha önceki çalışmalarla kıyaslanmış ve ince FD plaklar için KPT' nin oldukça başarılı sonuçlar verdiği görülmüştür. 6. Kısım' da ise sonlu elemanlar paket programı, PATRAN, NASTRAN yardımıyla burkulma analizleri gerçekleştirilmiş ve KPT ile elde edilen analitik sonuçlarla kıyaslanmıştır. Sonraki kısımlarda yapılan tüm çalışmalar kısaca değerlendirilmiş ve gelecekte bu konu üzerine yapılabilecek çalışmalardan bahsedilmiştir.
Flying and handling qualities oriented longitudinal robust control of a fighter aircraft in a large flight envelope

(Graduate School, 2022-02-15) Kaçan, Zafer ; Koyuncu, Emre ; 511181144 ; Aeronautics and Astronautics Engineering

In the scope of this thesis study, robust control design apprach has been applied to F-16 aircraft which is aimed to satisfy Level 1 FHQ within the specified flight envelope. First, a brief information about the histoy of flight mentioned in the introduction chapter. This historical storyline starts from the early sketches of Leonardo da Vinci and extends along to Wright Brothers who had achieved the first sustainable, controlled heavier than air flight. Then the innovations in aerospace industry is mentioned along with the advances in technology at the same time. The milestone successes are explained which has brought us to realize the design of fly-by-wire flight control algoritms. Then a literature review of the documents about the F-16 aircraft, FHQ criteria, multivariable robust control applications and mathematical backgroud of this approach. The structure of the thesis has outlined. Then, F-16 aircraft has been presented along with the aerodynamic data and how the force and moment of the aircraft is related with the aerodynamic and thrust data. The presented data of the F-16 aircraft was obtained from the researches of NASA Langley Research Center which is based on wind tunnel test results of the F-16 aircraft. The mathematical model of the F-16 aircraft is introduced. This mathematical model includes the airframe spacifications, the mass data, the systems that represent actuator and sensor and the environmental data which gives the atmospheric properties with respect to the flight condition of the aircraft. Then, the trim and linearization algorithms are introduced for the steady-state wings level flight condition. The inputs, states and outputs related to the longitudinal motion of the aircraft has been identified and the resultant state space linear system which represents the characteristics of the aircraft is obtained. The longitudinal modes which are phugoid and the short-period mode are mentioned. Next, the flying and handling qualities to evaluate the performance of the aircraft are emphasized. The reason for the use of flying and handling qualities are determined and related with the pilot evaluations Cooper-Harper ratings. The suggested flying and handling qualities are explained for the use of both design guidance and evaluation criteria. It is mentioned that the CAP criterion is used as design guideline whereas the Bandwidth and Dropback criteria are used as evaluation criteria for the aircraft in the both frequency and time domain. The corresponding flying and handling qualitieslevels are detailed for the criteria and related intervals for the properties are supported with the graphical representations. The robust control approach is introduced while mentioning the background of the method. The norm definitions are done and the feedback properties are given in the related chapter of this thesis in order to associating the design purposes with the feedback properties. The relationships between the open-loop characteristics and closed loop results are identified and the loop-shaping aproach is emphasized. Yhen the uncertainty definitions are identified. The classes of uncertainty and where they are reasoned for is explained. An uncertainty definition which is suitable for the use in this thesis is mentioned. Then the H_∞ Loop Shaping approach is expressed. The normalized coprime factorization method is explained and the design of both one degree of freedom and two degrees of freedom H_∞ Loop Shaping approaches are detailed with the design steps. Then, the control structure used in this thesis is explained. It is aimed to design a pitch rate controller which will results in Level 1 flying and handling qualities within a specified flight envelope. The design has been made for one design point and then the resulted parameters are used for the whole flight envelope. This enables to overcome the complexity of gain scheduling manner and provides robustness against any probable loss of air data such as angle of attack. The controller architecture of NASA research was presented for the longitudinal axis. Then the optimization structure to find the design parameters which ensures that the pitch rate demand flight control law results in Level 1 flying and handling qualities within a specified flight envelope. The root mean square approach has been applied in optimization phase. It is purposed that the time responses of 5 different design points after step input should follow a desired transfer function response specified during the design of the two degrees of freedom H_∞ Loop Shaping algorithm as close as possible. Moreover, in order to satisfy the specified flying and handling qualities, time delay parameter is included in the optimization cost which makes the optimization multiobjective optimization with a weigthed sum cost function. The resultant optimized parameters for the design of two degrees of freedom H_∞ Loop Shaping architecture is given. The results of both nominal design point and the responses of 5 different design points along the flight envelope are presented. The flying and handling qualities evaluations are shown. Then the performance and stability robustness results are associated with the results. The comparison study between the two degrees of freedom H_∞ Loop Shaping algorithm and the NASA control structure which emphasizes a classical PI controller has been presented. The results are satisfactory as all the design points resulted in Level 1 flying and handling qualities responses in both frequency and time domain. It is seen that the control architecture is successful for performance and stability robustness as all uncertain plants are following the nominal response and no frequency response has crossed a nichols exclusion zone defined. The two degrees of freedom H_∞ Loop Shaping algorithm outperformed the NASA PI controller as Level 2 results are seen for NASA PI controller responses. The use of two degrees of freedom H_∞ Loop Shaping structure lowered the time delays as it was purposed in the optimization goals as the effective time delay results are less than the NASA PI controller.
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.
Kompozit panellerin genetik algoritma ile yapısal optimizasyonu

(Lisansüstü Eğitim Enstitüsü, 2022-05-13) Başsüllü, Tevfik Can ; Türkmen, Halit Süleyman ; 511171135 ; Uçak ve Uzay Mühendisliği

Havacılıkta kompozit malzemelerin kullanımı her geçen gün artmaktadır. Kompozit malzemeler sayesinde yüksek mukavemete ve düşük ağırlığa sahip yapılar elde edilebilmektedir. Bu tezdeki amaç kompozit panellerin optimizasyonu yapılarak belirli yüklemeler altında en yüksek dayanımı gösteren yapıların elde edilmesidir. Genetik algoritma yöntemi kullanılarak yüksek değişken sayısına sahip kompozit yapıların amaç fonksiyonu doğrultusunda kısıtlar altında global optimumlarını hızlı bir şekilde bulması hedeflenmiştir. Problemde yapının stabilitesi göz önüne alınarak katman dizilim optimizasyonu yapılmıştır. Stabilite için kullanılan gösterge burkulma dayanımı olacaktır. Verilen yükleme ve istenilen sınır koşulları altındaki yapının dayanabileceği en yüksek burkulma yüküne dayanabilmesi için optimize edilecektir. Plaka düz olarak düzlem içi yüklemelere maruz bırakılmıştır. Farklı sınır koşullarında sahip plakalar kullanılmıştır. Basit mesnetli, ankastre mesnetli ve üç kenarı basit mesnetli bir kenarı serbest plaka için çalışmalar yapılmıştır. Basit mesnetli plakada iki eksenli yükleme uygulanıp farklı yükleme oranları kullanılmıştır. Artan yükleme oranlarına göre plakanın burkulma yükünün dayanımının düştüğü gözlenmiştir. Ankastre mesnetli olduğunda basit mesnetliye göre yapının burkulma yükü dayanımında artış olmuştur. Bir kenarı serbest olan basit mesnetli yapı da ise burkulma yükü dayanımında büyük düşüş olmuştur. Plakalar gerinim ve Tsai-Wu göçme teorilerine göre değerlendirilmiştir. Burkulma yükünün büyüklüğü gerinim veya Tsai-Wu göçme teorilerine göre hesaplanmıştır. Bütün kenarların ankastre olduğu durumda burkulma yükü fazla geldiği için gerinim ve Tsai-Wu kriteri baskın olmuştur. Ancak bir kenarı serbest olan basit mesnetli plaka için de gerinim ve Tsai-Wu göçme teorilerini kısıtlarının hiçbir etkisi olmamıştır. Tüm kenarları basit mesnetli olan ise göçme teorileri burkulma yükünü sınırlandırıcı etki yapmıştır. Kısıtlara ardışık katman diziliminin de kontrol edilebilmesi için ardışık katman dizilimi eklenerek 0 ve 90 yönündeki katmanların bir araya toplanmasının önüne geçilmek istenmiştir. Ardışık katman dizilimi ise burkulma yükünü yükleme koşuluna ve sınır koşuluna göre etkilemektedir. Elde edilen tüm dizilimlerdeki burkulma dayanımları daha sonra FEM ile kıyaslanmıştır. Kısıtlar arttıkça elde edilen sonuçlardaki burkulma dayanım yükleri FEM sonuçları yakınsamaktadır. Böylelikle elde edilen sonuçların tutarlı gösterilmiştir. Genetik algoritma programlaması MATLAB kullanılarak hazırlanmıştır.

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Çıkarma tarihi ile LEE- Uçak ve Uzay Mühendisliği-Yüksek Lisans'a göz atma

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