127
Archived Theses
0
DOIs Assigned
0%
DOI Rate
Discipline
Development of methods for reducing scheduled maintenance downtime of aircraft
Maximization of aircraft utilization is one of the most important issues for airline companies. By accumulating more flight hours in a certain period of time, direct operating costs per flight hour can be reduced. Achieving more flight hours depends on aircraft availability for flight. One of the methods for increasing aircraft availability for flight is to reduce scheduled maintenance downtime of the aircraft. Methods of reducing maintenance downtime of an aircraft were examined in the scope of this thesis. By means of these methods, scheduled maintenance downtime of the aircraft is reduced, aircraft availability for flight is increased and a cost based structure has been established for the airline companies. Today many airline companies perform maintenance checks in predetermined intervals. Maintenance tasks given in maintenance programs are performed during these maintenance checks. However, this method used by many airline companies remains currently insufficient to prevent earlier accomplishment of maintenance tasks. Due to the rigid structure of this maintenance concept and fixed intervals, such airlines are subjected to enormous loss of material, man hour etc… This cause the aircraft to be kept much more time on gorund for maintenance purposes, consequently the aircraft availability for flight decreases. In the method proposed in this thesis, a more flexible structure has been developed to perform the maintenance not only during periodic checks but also whenever the aircraft is on the ground for any reason. In this method, a single task-oriented maintenance concept has been proposed. The method proposed in this thesis has been supported with some case studies performed in an airline company. In the scope of this thesis, a software which will help to increase the aircraft availability for flight has been developed to calculate more accurate and appropriate letter check alternatives according to real-time utilization values accumulated by each aircraft. This software allows us to use real-time utilization values for all aircraft and therefore it may suggest different appropriate letter check alternatives for each aircraft. Maintenance tasks which are planned in maintenance checks without taking into account the real-time utilization of the aircraft cause losses because of potential earlier accomplishment of these tasks and therefore decrease the aircraft availibility eventually. For these purposes, the accomplishment of maintenance tasks in the most appropriate letter check alternatives according to real-time utilization values and the suggestion of different appropriate letter check alternatives for each aircraft are studied in the software developed in this thesis. Single task-oriented maintenance concept, the accomplishment of maintenance tasks in the most appropriate letter check alternatives according to the real-time utilization values and the calculation of different appropriate letter check alternatives for each aircraft are subject of this thesis. With this method and maintenance concept, it is intended to decrease maintenance scheduled downtime of the aircraft, to accomplish maintenance tasks on time and to prevent losses and earlier accomplishment, to reduce maintenance cost of airline companies, to increase aircraft availability for flight and to maximize aircraft utilization, therefore to increase the flight hour in a certain period of time.
Structural health monitoring of frp composites using fbg sensors
Use of composite materials have become prevalent in aerospace, marine, energy and automotive industries. The well known advantages of composites make them very attractive for complex load bearing structures. However, there is still a long way to fully exchange metallic materials with composites; because unknown factors like long term fatigue behavior necessitates relatively higher safety factors that should be applied for composite structures during the design phase. This study was performed to develop structural health monitoring (SHM) methods using Fiber Bragg Grating (FBG) sensors, which may reduce safety-oriented design perspective by providing continuous information about the integrity of structure during the service. An FBG sensor behaves like an optical filter that reflects a specific wavelength of light back to the light source. The reflected light's peak wavelength is dependent on the distance between gratings. Once the distance between gratings change due to a mechanical load, peak wavelength of reflected light shifts. This property of FBG sensors make them perfect for strain measurement. Advantages such as small diameter, immunity to electromagnetic fields, multiplexed usage opportunity and embeddable nature of FBG sensors make them unique for the use of SHM applications. Embedding FBG sensors into composite materials is not a straight-forward task, since it is a delicate process especially for manufacturing technique like Resin Transfer Molding (RTM) method which utilizes a closed mold. Specimens manufactured by RTM method should be designed in such a way that the fiber optical cable must be taken out from the end of the specimen in one piece and should be protected during the entire manufacturing process. In the beginning of this study, testing the specimens that contain embedded FBGs were performed by special apparatus because of the method that was used to take the fiber optics out of the specimen. Studies showed that additional apparatus was complicating the experimental procedure and most of the time caused immature failure of the specimens during fatigue tests. To overcome this, initial research was focused on developing an effective method for taking the optical sensor out of the specimen that makes both manufacturing and testing more feasible. In this context, L shaped geometry was developed for specimens that contain FBG sensors which resolved all complications related to testing of FBG embedded specimens by eliminating apparatus use. Embedded FBG sensors are ideally suited for continuous monitoring of composites under cyclic loading conditions, which can be very useful for SHM applications. Remaining useful life (RUL) of composite structures under fatigue loading, which have been investigated by many researchers in literature, was utilized in this thesis study for FBG embedded composites by calculating released strain energy at each load cycle by using FBG sensor data. Initially fatigue constants of GFRP material derived experimentally by testing sensorless specimens and calculating strain energy release rate (SERR) in a wide variety of strain ratios. Following to that FBG embedded GFRP specimens were manufactured and tested under high cycle fatigue conditions. Research results showed that FBG sensors are very effective to obtain strain from the interior of structure during fatigue experiments. It has been seen that shorter sensor lengths are more reliable in terms of signal quality as an embedded sensor since long FBGs are prone to peak splitting due to the variable strain fields within the composites. Sinusoidal strain behavior for each load cycles were clearly obtained using embedded FBGs at 4Hz of cyclic loading. Predicting the remaining life under some assumptions was successfully performed by analyzing the FBG data. As an interesting outcome, it was revealed that the local behavior of composites monitored by embedded FBG sensors differs from the global observations made by external sensors. It should be noted that FBG sensors stayed intact until the fracture of specimens, which proves the method's usability for actual applications. Two FBG sensors on the same optical fiber were used as a multiplexed sensor and embedded into specimens to utilize a biaxial sensor configuration. This method was enabled to collect strain data both in axial (i.e. loading) and transverse directions using embedded FBG sensors under cyclic loading conditions. By analyzing the fatigue strain data of both FBGs, Poisson's ratio of GFRP specimens was utilized at each load cycle. Poisson's ratio degradation trend of tested specimens was revealed using embedded FBG data during fatigue loading. It has been seen that Poisson's ratio of composite materials can be simultaneously monitored using embedded FBGs during the entire life of the structure, which contains critical data about the current state of the structure in terms of integrity. Both extensometer and displacement controlled test types were used during the experiments. It has been revealed that the local strain trends of composites observed by FBG sensors was different than the global behavior throughout the gage length obtained by extensometer and LVDT in a similar manner with previous study. Strain-gage sensors were also used during the experiments and it has been seen that these sensors are not as durable as FBG sensors especially for cyclic loading conditions. More than 3.5 millions of load cycles were performed in a low strain fatigue experiment and both FBG sensors were working flawlessly when the test was ended intentionally without any visible damage in the specimen. Temperature effect on FRP composites and embedded FBG sensors was also studied within the scope of this thesis study. FRP composites manufactured by RTM and hand layup with Vacuum Curing&Debulking Table (VCDT) methods were mechanically compared under cold, room and elevated temperature conditions. Products of RTM method were performed better for all temperature conditions. It has been seen that cold temperature (i.e. -40C) was increased while elevated temperature (i.e. 70C) was deceresed the ultimate strength of the composite materials compared to room temperature conditions. Overall, the mechanical behavior of both specimen types did not heavily differentiate especially considering the fiber volume fraction difference. Surface quality of the specimens was totally different for both techniques. Hand layup specimens manufactured using VCDT were containing many bubbles and the surfaces of them were heavily porous while RTM products' surfaces were perfectly smooth without any porosity and visible bubbles. Results revealed that wet hand layup technique was not suitable for embedding FBGs since the sensors are very sensitive to discontinuities within the hosting material. Moreover, porous nature of wet hand lay up products were not appropriate for fatigue testing since it would reduce the fatigue life substantially. Therefore, it was decided to investigate elevated temperature behavior of FBG embedded specimens manufactured by RTM method under fatigue loading. Experiment was performed at 50C with a 0.27 peak strain ratio. Specimen was tested more than 2 million cycles under these conditions and both specimen and FBG sensor were stayed intact without any failure. The data of FBG sensor was analyzed for SERR calculations and no any adverse effect of elevated temperature was noticed on FBG data. FRP composites containing open holes were also investigated from the SHM point of view. GFRP specimens were manufactured that contains central hole notch with a 4mm of diameter. Static experiments were performed on specimens without sensors in the beginning to understand the mechanical behavior. Following to that, three specimens manufactured with biaxially embedded FBG sensors. Axial and lateral sensors were located at 4mm, 7mm and 10mm away from the center of the holes of each specimen. Opening the holes using CNC for all specimens were monitored by the embedded FBGs. Release of the residual strains resulted by impregnation was observed within the analyzed data during the drilling process. Specimen that contains FBGs at 7mm from the hole center was tested statically. Both sensors were collected strain data around the hole notch as expected till failure of the composite. Other two specimens that contains biaxially located FBG sensors were fatigue tested. Experiments revealed that the sensors at 2mm distance from the hole edge (i.e. 4mm from the center) were able to survive only about 28-40 \% of total fatigue life due to the progressive damage caused by cyclic loading. Other specimen's FBG sensors, which were located at 8mm distance from the hole edge (i.e. 10mm from center),were survived during the entire fatigue life of the specimen (i.e. around 50000 cycles). Analyzed FBG data was indicated that a very unstable strain behavior exists around the hole during the experiment. It can be stated that monitoring damage dependent strain field around an open hole was the most challenging environment for embedded FBG sensors since they were directly exposed to the growing cracks in the vicinity of the hole edges. Nevertheless, embedded sensors were able to perform their monitoring function under cyclic loading condition at a designated distance from the hole edges. FBG sensor behavior in open hole composites was also studied using FEM analysis. Unidirectional carbon/epoxy composite specimens were considered as host material about 1mm thick, which were used for fatigue monitoring with FBG sensors and performed very well. After performing experiments for obtaining material constants, three FEM models were run to simulate behavior of FBG sensors embedded into open hole specimens under tensile loading. Results showed that a significant amount of axial strain (i.e approx. 900 microstrain) variation may develop along an FBG sensor which is located near the hole and aligned with center point of specimen in longitudinal direction. The revealed uneven strain field in the vicinity of an open hole is already a challenging environment for an FBG sensor even ignoring the growing damage in its surroundings under cyclic loads. These results clarify complicated behavior of FBG sensors that were embedded into open-hole specimens and tested under fatigue loading in previous studies. Overall, different techniques were investigated that effectively utilize FBG sensors in this study for static and fatigue monitoring of FRP composites. It can be stated that the used methods of monitoring FRP composites with embedded FBG sensors have created a strong basis for future SHM systems.
Heat transfer modes enhancement by using magnetic nanofluids
Ferrofluids, a mixture of nonmagnetic fluids with magnetic nano-sized particles, offer advantages such as adjustable thermal conductivity and ability to be controlled via an external magnetic field. Although referred to as not a matter of life and death, ferrofluids have various medical applications which might serve as one soon. An example for promising medical application of nanofluids is cancer treatments, where external magnetic field is utilized to guide special magnetic nanoparticles directly to the malign tumors giving less harm to the surrounding tissues. Their not totally understood superiority for heat conductance also offers effective solutions for a variety of different cooling applications. This work focuses on investigation of flow characteristics of this new engineered fluids in channel flows. Study here involves numerical calculations rather then experimentation. Thus, efficient tools to solve the system equations plays a vital role. Nanoparticle specific properties of the flow has been simulated by the available emprical models in the literature.
Uçak kanat aerodinamiğinin had yöntemi ile analizi ve doğrulama
HAD yöntemi 1970' lerin sonunda 2 boyutlu bilgisayar kod ve çeşitli algoritmalarla başlayarak önemli bir atılım gerçekleştirmiştir. HAD yöntemi günümüzde araştırma, geliştirme, optimizasyon, doğrulama vb. gibi işlemler için bir hayli kullanılmaktadır. HAD yönteminin kullanılma amacı varolan veya olmayan bir nesnenin akış karakteristiği hakkında bilgi sahibi olunması isteğidir. Model kompleksliliğiyle doğru orantılı olarak kaynak, yani gerekli olan bilgisayar yatırımı da artmaktadır. HAD yöntemi teoride hayatı oldukça kolaylaştıran bir alet olsada tecrübeyle sonuç yorumlama olmadan sadece renkli görüntülerden oluşur. Hatanın nerede olduğunun farkedilmesi ve çözülmesi bu yöntemin en önemli parçasıdır. HAD yöntemi teorik yöntemleri kullanan akademik veya ticari yazılımların numerik hesaplamalar yapması sonucu elde edilmektedir. Teorik hesaplamalarda probleme bağlı olarak Navier-stokes ve süreklilik denklemlerinde bazı parametreler ihmal edildiği gibi aynı şekilde HAD yönteminde de aynı kabulleri girdilere göre yazılım yaparak hesabı numerik olarak çözmektedir. Bu tez çalışmasında HAD yönteminin günümüzde kullanılma sebeplerinden biri olan sonuç doğrulama hakkında araştırma yapılmıştır. İki adet ayrı geometri bu işleme tabii tutulmuştur. Bunlar NATO' nun birçok deneyinde kullanılan ONERA M6 kanadı ve rüzgâr tünel kalibrasyon modeli AGARD C' dir. ONERA M6 kanadı 1972 yılında ONERA S2MA rüzgâr tünelinde transonik hızlarda 30 adet farklı teste maruz bırakılmıştır. Kanat üzerinde 7 adet kesit düzlem oluşturulmuş ve bu düzlemlere ölçüm delikleri yerleştirilmiştir. Ölçüm deliklerinden basınç dağılımı okumaları yapılmıştır. Son olarak basınç katsayıları deliğin konumu bazında hesaplanarak her test için grafikleri çizdirilmiştir. AGARD C modeli ise, rüzgâr tünel kalibrasyonlarında kullanılan bir modeldir. Transonik hızlarda çeşitli testlere tabi tutulan bu modelin varyasyonları birçok enstitü tarafından ele alınmıştır. Test sonuçları 3 adet hucüm açısı ve enstitülerce oluşturulan farklı modellerin Mach sayıları bazında moment, sürükleme, kaldırma kuvveti katsayıları hesaplanarak her test için grafikleri çizdirilmiştir. Bu tezde bilgisayar ortamındaki katı modeli çizilen cisimlerin HAD yöntemiyle akış analizi yapılarak elde edilen sonuçları ile deney sonuçlarının karşılaştırılma yapılması amaçlanmıştır. Geliştirilmiş çeşitli viskoz modellerinden ikisi ile seçilen testler çözülmüştür. Basınç, sürükleme, kaldırma kuvveti katsayıları karşılaştırılmıştır. Karşılaştırmalar sonucu HAD yöntemi sonuçlarında çeşitli hatalar elde edilmiştir. Bunun nedenleri çözüm ağındaki bozukluk olmasına kanaat getirilmiştir. Spalart-Almaras viskoz modeli çözümlerinde tek denklem kullanmaktadır. Bu modelin belirli bir transonik hız limitine kadar oldukça doğru çözüm verdiği görülmüştür. Fakat 1.1 Mach sayısından yüksek çözümlerde çözüm ağının hatalarıyla da birlikte çözümdeki hata oranı ciddi şekilde artmaktadır. SST (Shear Stress Transport) k-omega modelinde ise girdap ve viskoz etkilerinin daha iyi hesaplandığı bilinmektedir. Yalnız bu yöntem çözüm ağındaki hatalardan oldukça etkilendiği için hesaplamalar yarım olarak yapılabilmiştir.
Termoelastik bir kirişin method of lines ile incelenmesi
Bu çalışmada ki asıl amaç termoelastik bir kirişin method of lines ile incelemesini yapmaktır. Bu amaçla ilk önce termoelastisitenin ne olduğu anlatılmıştır. Bir malzeminin elastik davranışını incelerken çoğu mühendislik problemi için ısıl etkiler ihmal edilir. Bunun nedeni dünyada kullanılan malzemelerin dünyada ki yerçekimine uygun olarak daha dayanıklı ve doğal frekanslarının daha büyük olmasıdır. Bu durum enerji denkleminde gerinmenin yaptığı işi ihmal etmemize olanak verir. Çünkü bir malzeme de ısıl etki ile bir deformasyonun olabilmesi için malzemenin ısıl yüklemeye karşı ısıl tepki süresinin göreli olarak hızlı olması gerekir. Esas olarak bir malzeme için ısıl etkilerin hesaba katılıp katılmaması yani elastisite denklemlerinin termoelastisite denklemlerine dönüştürülmesi malzemenin ısıl tepki süresi ile alakalıdır. Eğer malzemenin ısıl tepki süresi yani sıcaklık değişimi süresi elastik deformasyon süresine yakın ya da ondan daha küçük ise ısıl etkiler hesaba katılmalıdır. Boley 1957 yılında ki makalesinde bu durum için bir formül ortaya koymuştur. Uçak ve uzay mühensiliği alanı için malzemelerde ısıl yükleme sonucu ortaya çıkan ısıl gerilmenin hesaplanması önemlidir, çünkü bu alanda kullanılan malzemelerin çoğu hafif ve esnek malzemelerdir. Malzemenin hafif ve esnek olması ısıl geçirgenliğini arttırır buda dikkate aldığımız karakterisitik uzunluk boyunca sıcaklık farklarına neden olur. Oluşan bu sıcaklık farkları ısıl moment formülünde yer aldığı için diğer bir değişle ısıl moment oluşması sıcaklık farkına bağlı olduğu için malzemede ısıl yükleme kaynaklı gerilme ve bunun sonucunda yerdeğiştirmeler oluşur. Bu nedenle yapısal dinamik denklemler ile ısı denklemi birbirine bağlanır (coupled) ve eş zamanlı olarak çözülür. Bu denklem sistemine ise malzemenin termoelastik denklemleri denir. Havacılık alanındaki en temel yapı kiriş yapısı olduğu için bizvde termoelastisite denklemlerini uygulamak için bu yapıyı seçtik. Bu amaç ile literatür taraması yapıldı ve giriş bölümünde literatürde bu konu ile ilgili yapılan çalışmalardan bahsedildi. Daha sonra bir ankastre kiriş için termoelastisite denklemlerinin çıkarılışı ayrıntılı olarak anlatıldı. Denklemler çıkartılırken sıcaklık dağılımı 2 boyutlu zamana bağlı, enine yer değiştirme ise 1 boyutlu zamana bağlı olarak düşünüldü. Denklemlerin zamana bağlı olması sistemi lineer olmayan bir hale getirdi. Lineer olmayan bu denklem sistemi analitik olarak çözülemediği için sayısal olarak çözüldü. Sayısal çözüm yapılırken method of lines kullanıldı ve gerekli kod mathematica programında yazıldı. Method of lines kısmi diferansiyel denklem sistemini adi diferansiyel denklem sistemine dönüştürmemize imkan verir. Böylece elimizde bir adi diferansiyel denklem sistemi olur. Bu sistem özdeğer problemi gibi düşünülerek gerekli çözüm yapılır. Tezimizde method of lines yönteminden ayrıntılı olarak bahsedilmiştir. Kullandığımız kiriş teorisi Euler-Bernoulli kiriş teorisidir. Bu teorideki kabuller ve denklemlerin oluşturulma mantığından tezimizde bahsedilmiştir. Bütün denklemler method of lines'a göre ayrıklaştırıldıktan sonra çözümleri ayrıntılı olarak anlatılmıştır. Ayrıca kullanılan kodda teze eklenmiştir. İlerideki çalışmalarda bu kod daha da geliştirilip daha kompleks yapılar için kullanılabilir. Son bölümde ise termoelastisite denklemleri çözüldükten sonra elde ettiğimiz sonuçlar grafikler yardımı ile gösterilmiştir. Daha sonra grafikler ile görselleştirilen sonuçlar tartışılmıştır. Çıkan sonuçların fiziksel olarak doğru olup olmadığı üzerinde özellikle durulmuş ve belli varsayımlar yapılıp bu varsayımlara göre çıkan sonuçlar yorumlanmıştır. Son olarak varsaydığımız durumların sayısal çözüm ile uyumlu olduğu gösterilmiştir.
Dynamic and aeroelastic analysis of a wind turbine blade modeled as a thin-walled composite beam
In this thesis, dynamic and aeroelastic analyses of a wind turbine blade designed as a rectangular thin-walled composite beam are carried out. The response of the blade is investigated by setting proper geometrical linear formulation. The blade is modeled as a uniform (untapered) thin-walled composite beam with rectangular cross-section. This composite beam is consisted of anisotropic material and contains warping and transverse shear effects. Also, there exists some elastic couplings due to the directionality of composite material. In order to gather this coupled behavior, circumferentially asymmetric stiffness (CAS) is used for lay-up configuration. Firstly, free vibration analyses are performed for both of the nonshearable and shearable models. Then, static analysis is realized to find out divergence while the dynamic analysis is done for the flutter behavior. The unsteady aerodynamic loads are represented by Theodorsen function in frequency domain. Moreover, stiffness coefficients are found for different ply-angles and rotational speeds. The results are compared with the literature and it is seen that a good compatibility is provided.
Taşıyıcı yüzey sistemlerinin girdap kafes yöntemi ile analizi ve genetik algoritma ile eniyilemesi
Hava araçlarının aerodinamik tasarımı, analizi ve eniyilemesi, akışkanlar dinamiği bilimiyle uğraşan araştırmacıların üzerinde durduğu karmaşık ve esasında çok-disiplinli mühendislik problemleridir. Bu kapsamda, günümüzde uygulanan Hesaplamalı Akışkanlar Dinamiği (CFD) çözümleri tasarım sürecinin ilk aşamalarında birçok sebeple uygun görülmemektedir. Onun yerine daha basit çözümler sunan, akış alanıyla ilgili temel yaklaşımlarda bulunanarak çözümü hızlandıran yöntemlerin kullanımı daha yaygındır. Hesaplamalı aerodinamik uygulamalarında sıkça kullanılan Girdap Kafes Yöntemi (VLM) taşıyıcı yüzeylerin üzerindeki potansiyel akışı çözmektedir. Ludwig Prandtl'ın 1918 yılında tanıttığı taşıyıcı çizgi modeli (PLL) sonlu düz kanatlar için analitik bir çözüm sağlarken, yine Prandtl'ın ortaya koyduğu teoriyi esas alan VLM ile herhangi bir konfigürasyondaki çok elemanlı taşıyıcı yüzey sisteminin incelemesi girdap-kafes modeli ile mümkün olmaktadır. Bu modelde yüzeyler çözümlemenin yapılacağı panellere bölünür ve her panele taşıyıcı özelliğini veren atnalı girdapları yerleştirilir. Potansiyel akım sınır şartı kullanılarak çözülen girdap şiddetleri ile taşıyıcı yüzey üzerine etki eden aerodinamik kuvvetler Kutta-Joukowski Kanunu ile hesaplanır. Konsept ve ön tasarım süreçlerinde ilgili hava aracının taşıyıcı yüzeyleriyle ilgili kanat açıklığı, açıklık oranı, kök ve uç veter uzunlukları gibi birçok parametrenin belirlenmesi gerekmektedir. Ancak, her bir parametrenin belirlenmesi başlı başına bir eniyileme çalışması gerektirmektedir. Bunun sebebi ise taşıyıcı yüzeyler birbirleri arasında aerodinamik etkileşim içindedir. Bir yüzeye ait parametre diğer bir yüzey üzerinde etki yaratmaktadır ve hava aracının genel performansı açısından bu etkinin incelenmesi önem arz etmektedir. Doğanın genetik mekanizmalarını açıklayan Darvinci Evrim Teorisi (DET), yapay sistemlerin eniyilemesinde de etkin bir şekilde kullanılabilmektedir. Evrimsel Algoritmaların (EA) bir alt sınıfı olan Genetik Algoritmalar (GA), doğanın kanunlarını yapay sistemlere uygulayarak, zorluklar karşısında kendi kendini geliştiren sistemler üretmektedir. Bu uygulamanın temellerini John H. Holland 1975 yılında atmıştır. Holland'ın yürüttüğü çalışmalar hem yapay sistemlerin eniyilemesi hem de doğal sistemlerin bilimine büyük katkı sağlamıştır. Bu tez kapsamında çok elemanlı taşıyıcı yüzeylerin nümerik incelemesi için VLM yönteminin bir bilgisayar programını geliştirilmiştir. MATLAB programlama dili kullanılarak geliştirilen yazılımın öncelikle PLL yöntemi ile düz bir kanat analizi için karşılaştırılmış ve elde edilen sonuçlar doğrulanmıştır. Çok elemanlı sistemler içinse ek doğrulama çalışmaları yürütülmüştür. Ters açılı bir çift kanat kullanılarak yer etkisi analizi yapılmış ve kanatlar arası uzaklık artırıldıkça yer etkisinin azaldığı görülmüştür. Bir diğer doğrulama çalışmasında ise bir kanat arkasında yerleştirilen yatay kuyruğun analizi yapılmıştır. Bu analizde aşağı sapmanın etkisiyle kuyruk üzerindeki etkin hücum açısının serbest akımın geliş açısından daha düşük hesaplanması incelenmiştir. Doğrulama çalışmaları neticesinde, taşıyıcı yüzeylerin incelenmesi amacıyla geliştirilen VLM yazılımı evrimsel döngüler sırasında çağırılabilecek bir fonksiyona dönüştürülmüştür. Evrimsel döngülerin değerlendirme aşamasında, her bir birey için bu fonksiyon çağrılmış ve aerodinamik katsayılar hesaplanmıştır. Hesaplanan katsayılar güçlü bireylerin belirlenmesine yarayan uygunluk fonksiyonlarında kullanılmıştır. Çaprazlama ve mutasyon evrelerinden sonra elde edilen çocuk bireylerin sonraki nesle nasıl aktarılacağı konusunda farklı yaklaşımlar denenmiş olup, elde edilen sonuçlar karşılaştırılmıştır. Sonuç olarak geliştirilen taşıyıcı yüzey sistemlerini eniyileme yazılımı uygulamalarda kullanılmış ve elde edilen bulgular paylaşılmıştır.
Recoverability envelope analysis of nonlinear control laws for agile maneuvering aircraft
This theses proposes a novel framework to examine the effectiveness of controllers, which are designed for agile maneuvering aircraft upset recovery. The introduced framework incorporates two main assessment criteria. First, the angular rate envelope extent of the controller is determined, which provides a measure of functionality region of the controller. Second, time histograms are obtained in order to understand how fast does the controller recover the aircraft if the controller can recover the aircraft. Such recovery time analysis also lead this theses to cover a novel methodology which yields the recovery of an agile maneuvering aircraft from upset conditions as swiftly as possible. In this work, a linear quadratic regulator, a nonlinear dynamic inversion (NDI) based angular rate controller, double loop NDI controller, and another NDI based switching controller are considered for comparison. The latter controller presented in this theses is a new strategy for upset recovery of an agile maneuvering aircraft. As it is mentioned earlier, this theses introduces a novel methodology in order to recover an agile maneuvering aircraft from upset conditions as swiftly as possible. The current methodology based on an angular rate grid, which consists of angular velocity nodes. The aircraft is said to start its motion from an initial condition while the initial angular velocity corresponds to a certain node within the aforementioned angular rate grid. A nonlinear dynamic inversion (NDI) based controller is employed to reach transition angular rates, which are the desired rates before the recovery initiates. Once the desired transition node is reached, the controller produces necessary input signals to recover the aircraft. The time elapsed throughout the whole recovery process is recorded. The same sequence of control actions are repeated for each initial and transition nodes of the grid. In this way, a recovery time graph is obtained. The time graph is then used to determine the reference angular rate trajectory, which provides the minimum recovery time for the given rate envelope. Simulations are carried out by employing a nonlinear high fidelity F-16 model. Simulation results show that the switching NDI controller has the superior recovery performance in terms of angular rate envelope extent. Moreover it is revealed that there exist such angular rate trajectories results with faster recovery. Therefore, the aforementioned results point out non-intuitive solutions. That is, for certain initial states it is not suitable to directly regulate the angular rates when recovery time is considered. Rather, for some cases, it is more convenient to increase the roll rate first, and then regulate the angular rates. That being said, the recovery methodology proposed in this theses yields faster recovery.
Uçak motoru komponent tasarımı ve optimizasyonu
Bu çalışmanın temel amacı uçak motorlarında dönen parçalar olarak nitelendirilen disklerin optimizasyonunun gerçekleştirilerek, infilak hızının hesaplanması ve bu anlamda bir metodolojinin ortaya konmasıdır. Gerçek bir disk geometrisinde bulunan 3 boyutlu yapılar tasarıma dahil edilmemiştir. 3 boyutlu yapılara örnek olarak civata delikleri, kanadın diske bağlandığı boşluklu yapılar örnek gösterilebilir. Bir disk tasarımı için LCF, HCF, çatlak mekaniği ve sürünme gibi konuların detaylarına bu çalışmada değinilmemiştir. Tasarım, ön boyutlandırma açısından yapı üzerinde görünen von Mises eşdeğer gerilme değerinin malzemenin akma dayanımı ile karşılaştırılması ve infilak hız tayini hesaplarını içermektedir. Bu amaçla literatür taramasından elde edilen veriler ışığında ilk olarak analitik hesaplamar gerçekleştirilmiş ve bu hesaplamalar Excel Solver aracı vasıtasıla optimize edilmiştir. Analitik hesaplamalar süresince ayrıklaştırma metodu kullanılarak, gerçek bir disk geometrisine yakınsanmaya çalışılmıştır. Gerçek disk geometrisi sabit kalınlıkta farklı disklerden oluşacak şekilde parçalara bölünmüştür. Denge ve uygunluk denklemleri ve disk sınır şartları yardımıyla hesaplamalara gidilmiştir. Ayrıklaştırma metodu ile elde edilen geometrik veriler sonlu elemanlar analizi yardımıyla tekrar analiz edilerek analitik hesaplamalarla kıyaslamaya gidilmiştir. Sonuçlara göre analitik hesaplama yöntemi olan ayrıklaştırma metodu yaklaşık %10 farkla uygun sonuçlar vermiştir. Analitik ve sonlu elemanlar analizi yardımıyla disk üzerinde oluşan gerilme dağılımları incelenmiş ve disk tasarımı için dikkat edilmesi gereken kritik bölgeler belirlenmiştir. Bu duruma göre bu tez çalışmasında oluşturulan geometrinin gerilmeler açısından kritik olan bölgeleri en iç çap olan „‟bore‟‟ bölgesi ve en dış çap olan „‟rim‟‟ bölgesine geçiş yarı çapı olarak ifade edilebilir. Disk optimizasyonunun yanı sıra havacılık otoriteleri tarafından ortaya konulan regülasyonlar sebebiyle disk infilak hızı hesaplanmıştır. Hesaplamalarda Hallinan denklemi olarak ifade edilen denklem kullanılmıştır. Disk infilak hız hesaplamalarında seçilen malzemenin elastik ve plastik davranışları incelenmiş ve buna bağlı olarak infilak hızının değişimi ortaya konmuştur. Malzemenin infilak hesaplamalarına yönelik elastik ve plastik davranışı S faktör olarak nitelendirilen bir katsayının değişimi vasıtasıyla elde edilmiştir. S faktörü deneysel bir veridir ve çentikli bir numunenin çekme testinden elde edilir. Yüksek plastik davranışı gösteren bir malzeme için S faktörü 1 iken elastik davranış gösteren bir malzeme için 0‟dır. S faktörü 1, 0.5 ve 0 değerleri alınarak kritik olarak ön görülen bölgelerde infilak hızı hesaplanmıştır. Buna göre yüksek yüklemeler altında kesit alanı boyunca yük paylaşımı yapan durum için infilak hız değeri yük paylaşımı yapmadan kopmaya giden duruma göre daha yüksek çıkmıştır.
Thermo-elastic analysis and multi objective optimal design of functionally graded flywheel for energy storage systems
In this study, thermo-elastic analysis and a multi objective optimal design of a functionally graded Flywheel rotor that can be used in a Flywheel Energy Storage System (FESS) has been investigated. Due to the rapid mechanization and technological innovations that have taken place since the industrial revolution, our energy need has been increasing at an unprecedented rate. On the other hand, limited availability of known energy resources makes it necessary to find solutions that enables more efficient use of energy. One of the solutions to enhance energy efficiency is to store excess production for future usage. In this context, Flywheel energy storage applications for storing electrical energy have gained a great interest in recent years. Flywheel energy storage systems are in favorable position among others due to their strengths in high energy capacity, fast response, long life, low maintenance and environmentally friendly characteristics. Recent advancements in lightweight high strength materials, power electronics and magnetic bearings have made high speed FESS applications possible today. Obviously, those high speed applications such as turbo machinery components, high speed compressors, brake disks, gears, pumps, generators are of practical interest in engineering. Therefore, development of FESS systems has become a popular research topic for scientists and engineers nowadays. Flywheel energy systems are based on a rotating mass principle. In a FESS, electrical energy is converted into the mechanical energy and stored in the Flywheel rotor as rotational kinetic energy. Energy conversion is achieved by means of an electric motor and a bi-directional power converter. Flywheel rotor is the key element to determine the overall performance and capacity of the system. Energy capacity of a Flywheel rotor can be increased by using materials that is capable of resisting loads at high rotational speeds. Functionally graded materials which is a special kind of composites can be used for this purpose. For this reason, the subjected Flywheel rotor in this study has been assumed to be made of a functional graded material. For the thermo elastic analysis, a complex nonlinear differential equilibrium equation has been established. The complexity came from several factors such as non uniform thickness Flywheel geometry, non homogenous material content and by taking into account of the thermal effects on material properties. The governing equation of motion solved numerically by using generalized differential quadrature method (GDQM). Results of numerical solution has been validate by comparison with an axisymmetric two dimensional finite element model. A perfect correlation of both approaches has been demonstrated. Flywheel shape optimization task has been carried by using non dominated sorting genetic algorithm (NSGA II). Optimal designs have been searched for 18 design spaces having different speed, thickness and radius limits. Depending on the bearing type, maximum speed of the flywheel can be two types such as low speed and high speed. In 9 of the design spaces it has been assumed that conventional mechanical bearings were used. Hence, these designs had a maximum rotational speed limit of 10000 rpm. The other 9 design spaces assumed to have magnetic bearings. Therefore, those designs had a maximum rotational speed limit of 100000 rpm. The optimization problem is defined to maximize the average kinetic energy that the flywheel can store between an ambient temperature range without failure and without exceeding rotational speeds allowed by the bearings while minimizing the mass. The genotype of individuals have been constructed by a mixed type encoding such that thickness (h) and volume fraction coeffficient (m) has been encoded by 15 bit Gray coding and materials are encoded by value encoding. For generating child populations a multi point crossover has been used. Additionally, bit invertion and value changing mutations has been applied to the chromosomes. Genetic algorithm code have generated 1000 populations each of them having 1000 individuals. Therefore 1 million potential candidates have been evaluated for each design space. For the material content, random combinations of two materials out of three possibles (Ti-6Al-4V, Inconel 718, AISI 4340) have been selected for each designs. Optimization results have been achieved by using the Numerical solver and the Genetic algorithm code simultaneously and iteratively for each solution. Optimization study has presented following results, - At constant speed, size of the outer radius is the most effective design parameter in terms of the energy capacity. For reaching a higher energy capacity a bigger Flywheel size is reqiured. However, as the size gets bigger the stresses induced due to the centrifugal force increases proportional to the square of the radius of the Flywheel. - At high speeds the stress concerns becomes more evident such that, maximum design speed can not be achieved due to the violation of the stress limit. In such cases, algorithm starts another loop to determine the maximum allowable speed before continuing the other steps. In other words at high speed designs maximum rotational speed becomes another design variable. Therefore, computational efforts highly increases for high speed design searches. Consequently the convergence speed substantially decreases. - For low speed designs, maximum design speed (10000rpm) is achievable without violating any optimization criteria. Whereas, for the high speed applications, maximum achievable speed can not be attained due to stress constrains. For high speed designs maximum achievable speed depends on the Flywheel material and geometry considered. - For the optimal shape, the algorithm tries to increase the mass near the outer edge to increase the energy capacity. Whereas, near the inner edge mass has been increased in order to reduce stresses. In middle sections, algorithm has kept as minimum thickness as possible to reduce the weight of the design. - For material composition, algorithm allocates the lighter materials near outer radius. Whereas, the heavier materials are placed as possible as at the inner radius region in order to reduce stresses caused by the centrifugal loads. Additionally, algorithm tends to allocate AISI4340 to increase the stored energy due to its high density and strength. Whereas, for low energy capacity designs Ti-6Al-4V has been used to minimize the mass. Inco 718 has been used together with these materials to fine tune the designs with its medium properties. Some of the designs seem to have only one material content instead of a mixed FGM structure. Therefore, it can be said for some optimal designs homogenous material structure is the best choice to provide the optimal energy storage. - Some of the designs seem unrealistic due to very thin web sections. For such solutions the study may be developed further to take into account of resonance vibration effects.
Uzun menzilli uçuşların performans ve planlama yöntemlerine dayalı maliyet iyileştirmesi
Havayolu taşımacılığı, 20.yüzyıl içerisinde önemini her geçen yıl arttırmış ve herkes tarafından tercih edilen bir noktaya gelmiştir. Artan talep ve rekabet ortamı, havayolu şirketlerini maliyet hesaplamalarında ciddi çalışmalara iterek özellikle yakıt tasarrufuyla maliyet iyileştirmesi hedeflerin başını çekmektedir. Bu zamana kadar maliyet iyileştirmesi için birçok çalışma yapılmış ve olumlu etkileri görülerek havayolu şirketlerine kar sağlanmıştır. Literatürde yer alan çalışmalar, yakıt tasarrufunda birçok etken parametre olduğunu göstermektedir. Toplam kazanç eldesinde, uçağın yapısal ve operasyonel ağırlığı düşürülerek maksimum ticari yük miktarı kapasitesiyle uçuşun gerçekleştirilmesi en etkin yollardan biri olmuştur. Seyir uçuş sürati ve irtifası optimizasyonu, üstgeçiş maliyetleri, meteorolojik şartlar ve hava sahalarının kapalılıkları ele alınarak gerçekleştirilen rota optimizasyonları, periyodik motor bakımları, tek motor ile yapılan taksi planlamaları, uçağın bekleme esnasında yedek motor yerine harici elektrik ve hava ünitelerin kullanımı, uçak aerodinamik yapısındaki geliştirmelerin uygulanması, uçak ağırlığına etki eden uçak içi eğlence sistemleri, kitaplar, dergiler ve kullanılabilir su kaynaklarının azaltılması gibi çalışmalar toplam uçuş maliyetini olumlu yönde etkilemektedir. Yapılan çalışmalar ve araştırmalar uzun menzilli uçuşlarda oluşturulan uçuş planlarının meteorolojik şartlara, uçuş sürati ve irtifasına göre optimize edilmesi ve bu planlamaların gerçeğe en yakın tahminler ile yapılması toplam uçuş maliyetinde önemli ölçüde azalma sağladığını göstermektedir. Bu nedenle havayolu şirketleri günümüzde piyasada bulunan ve maliyet açısından fayda sağlayan ve optimizasyonların en etkin biçimde yapılabildiği ticari uçuş planlama paket programlarını kullanmaktadır. Bu çalışmada, Boeing 777-300 ER uçağının maksimum kalkış ağırlığı ile gerçekleştirilen uzun menzilli İstanbul-Los Angeles uçuşu performans ve planlama yöntemlerine dayalı maliyet iyileştirmesi amaçlanmıştır. Örnek uçuş planlamaları, günümüzde dispeçer eğitimlerinde ve sanal havacılık platformlarında kullanılan, ticari uçuş planlama yazılımlarına yakın sonuçlar veren "Professional Flight Planner X - PFPX" paket programı kullanılarak oluşturulmuştur. Literatürde yer alan maliyet iyileştirmelerinde en etkin performans yöntemlerinden biri olan Cost Indeks metodu ve planlama yöntemi olarak da Redispatch metodu İstanbul – Los Angeles uzun menzil uçuşunda incelenmiştir. Öncelikle Büyük Daire Mesafesi planlama yöntemi ile İstanbul - Los Angeles arası uçuş menzili belirlenerek en kısa mesafe tespit edilmiştir. Bu bilgiden yola çıkarak navigasyonel veriler doğrultusunda 3 noktadan oluşan büyük daire mesafesine en yakın menzile sahip GCD rotası ve bu zamana kadar yapılan gerçek uçuş planlamalarının menzil ortalaması olarak ele alınan OPT rotası oluşturulmuştur. Çalışmada, tırmanma ve süzülüş performans verileri kullanılan paket programının veri tabanı dahilinde ele alınacaktır. Her iki rotaya etki eden ortalama rüzgar şiddeti ve ISA sapması, "Climate Reanalizer" yaz ve kış dönemi veri tabanı incelenerek belirlenmiş ve planlamalarda kullanılmıştır. Elde edilen planlama parametreleri kullanılarak oluşturulan İstanbul - Los Angeles uçuş planlarında farklı Cost Indeks değerleri uygulanarak en düşük maliyetli uçuş rotası ve Cost Indeks değeri saptanmıştır. Bir havayolu şirketinin filosundaki uçak sayısı ve yıllık uçuş sayısı toplam maliyeti etkilediğinden dolayı bu çalışmadaki yaklaşımlar tek bir uçak ve rota üzerinden değerlendirilmiştir. Cost Indeks metoduyla en düşük uçuş maliyetine ulaşılan rota ve dönem ele alınarak Redispatch metodu aynı planlamaya uygulanmıştır. Hedef, standart planlamada elde edilen ticari yük miktarı sabit tutularak Redispatch metoduyla aynı uçuş planındaki yedek (Contingency) yakıt miktarını azaltmak olmuştur. Bu sayede uçuş için gerekli yakıt miktarı azaltılacak ve tasarruf sağlanacaktır. Tüm bu veriler bir araya getirilerek sonuçlar karşılaştırıldığında rota optimizasyonunda ve planlama yöntemlerinde kullanılan parametrelerin uzun menzil uçuşlarında maliyete ne kadar büyük etkisi olduğu açıkça görülmektedir. Ayrıca halihazırda uygulanan uçuş planlarına doğru hesaplamalar sonucu Cost Indeks ve Redispatch metodu uygulandıktan sonra tek bir uzun menzil uçuşunun yıllık operasyonu ele alındığında, en kötü şartlarda bile 120000$ maliyet iyileştirmesi sağlandığı sonuçlarda görülmektedir. Bu yüzden elde edilen sonuçlar, uzun menzil uçuşlarında performans ve planlama yöntemlerinin havayolu şirketleri tarafından imkanlar dahilinde uygulanması ve derinlemesine çalışmalar yapılması gerektirdiğini kanıtlar haldedir.
A single board computer design for flight control systems
Single board computers are compact and widespread solutions for commercial, industrial and military projects. Although the technical specs, capabilities, peripherals and interfaces differ, a standalone hardware including at least a processor which an operating system run on it with desired peripherals and interfaces such as RAM's, memories and various ports can be called as a single board computer. In this thesis, a single board computer hardware developed in TUBITAK-BILGEM-ILTAREN (The Scientific and Research Council of Turkey - Informatics and Information Security Research Center - Advanced Technologies Research Institute) as a part of a project of the Digital Hardware Design Group and which can be used in flight control systems is presented. Beginning from the literature survey and market research for component selection; architectural design and detailed schematic designs are carried out as well as PCB design and implementation. This thesis includes critical design decisions, comparison of alternative designs, methods and topologies for architectural, schematic and PCB design phases along with the signal integrity considerations and necessary calculations and simulations in order to accomplish the design. Though the designed hardware is manufactured, assembled and the board bring-up process is completed successfully, all these stages are excluded from the scope as the thesis focuses on the design. This design is based on Xilinx Zynq-7000 XC7Z045, an all programmable SoC with on-chip dual-core ARM® Cortex-A9 processor and 350K programmable logic cells. Therefore the ultimate design actually includes intellectual property implementations for the FPGA side of the SoC and firmware/software development including the algorithm implementations as well. These are also out of the scope of the thesis and only the board design stages of the computer are demonstrated. Furthermore, while the hardware design including architectural, schematics and PCB design stages is completely accomplished by the author of the thesis, digital designs for the FPGA and firmware/software development processes are carried out by other team members. This is another reason why they are excluded in addition to the fact that the thesis focuses only on the hardware development. Zynq-7000 SoC, DDR-3 RAM's, NAND flash memory, external SD memories and VGA, RS-232, RS-422 and Gigabit Ethernet interfaces as well as the power distribution network are the main blocks of the computer. Hardware design details about each block beside deep background information are given in the thesis.
Fault detection, isolation and reconfiguration in navigation system of Boeing-747 aircraft
In this thesis; Boeing 747 aircraft mathematical modelling, INS, GPS, INS error modelling, Kalman Filtering, INS/GPS integration, fault detection, isolation and reconfiguration are widely explained. In the study, Kalman filter is used to estimation of Boeing 747 aircraft states. The INS error modelling and integration of INS/GPS are examined and Kalman Filter is used to estimate states for INS error and integrated INS/GPS models. Fault detection, isolation and reconfiguration on Boeing 747 aircraft model and integrated INS/GPS is done by Kalman filter. All these simulations are generated on MATLAB simulation environment. There are two common methods for navigation, one of them is Dead Reckoning and other one is Position Fixing. INS is a type of Dead Reckoning navigation and most preferable, while GPS is more popular in Position Fixing method. Global Positioning System (GPS) and Inertial Navigation System (INS) provide navigation solutions to the aircrafts, ships, spacecraft, missile and other vehicles. INS is comprised by navigation computers and inertial sensors such as accelerometers and gyroscopes. INS procures information about aircraft's positions, velocities and attitudes. Nowadays, different type of aircraft systems such as Flight Management System, Electronic Flight Instrument System, Flight Guidance and Control System use INS output data. The inertial navigation system have different type of errors such as initiliazation errors and inertial sensors errors. The main problem of the system is that these errors expand with time, so that INS is time dependent. INS gives high accuracy for short time because INS errors increase with time and system accuracy decreases. GPS is well-known system for navigation and it is used in different areas such as mobile phones, aircrafts, ships, automobiles and so forth. GPS is consisted of three different segments which are control, user and space. The user segment is GPS receiver and it calculates own position by using GPS signals which are send from GPS satellites on space. The control segment monitor and make correction to GPS data. Mostly used GPS has some type of errors due to different causes. For instance signal quality of GPS may decrease due to signal interference and navigation information may not be satisfied. As a consequence, stand-alone INS or GPS does not sufficient solution for navigation applications which need high precision such as aircraft navigation. These systems have pros and cons, so that these systems are integrated and their advantages are combined for better navigation solution. The integration of INS and GPS is carried out by Kalman Filter. Kalman filter has recursive data processing algorithm and the main features of the filter are integrates more than one systems, diagnose system noises and computes unknown variables and minimizes error of measurements. In this study there are different type of problems are solved by using Kalman Filter. These solutions are that the filter is used to estimate state variables of Boeing 747 aircraft, Kalman filter minimizes the measurement error such inertial errors and find best value of measured parameters, the filter make prediction for INS errors and it is used for integration of INS/GPS, and Kalman filter is used to fault detection, isolation and reconfiguration. Faults may occur any time in the systems, this affects system reliability and sustainability. Therefore, systems should be designed to minimizing fault effects. In this study, it is examined and observed that the filter is reliable and providing correct data, when faults happen. The aim of fault detection and isolation is that when fault is occured, fault should be successfully detected and isolated. In MATLAB simulation environment, faults are given to the systems which are integrated INS/GPS and Boeing 747 aircraft model. Kalman filter is used to fault detection, isolation and reconfiguration. Simulation results show that Kalman filter is very effective and reliable in faulty condition. When fault has occured in a system, fault is detected and isolated by using Kalman filter and the filter maintains correct estimation for states.
Numerical investigation and thermal analysis of solar cavity receiver
The purpose of this thesis is to investigate steady, laminar, combined pure natural convection and surface radiation in a tilted 2D solar square cavity receiver, perform thermal analysis by using ANSYS Fluent, and compare the results with published ones and to be able to offer a new solution to the investigated reasons of thermal losses. This study is important to propose an optimum thermal design for solar concentrators. 2D solar square cavity receiver has a constantly heated left wall, and insulated top and bottom wall and an opening right wall. The temperature of the heated left wall is 350 K and temperature of the air coming from opening wall is 300 K, so the temperature difference investigated for the solar cavity receiver is 50 K. Different Rayleigh numbers from 102 to 106 were studied. Parameter changed for Rayleigh number depends on the factor investigated in the different cases for this study, such as it could be gravity or dimension. 2D, laminar, steady state flow, with SIMPLE scheme, presssure-based solver was studied and for air density, Boussinesq approach was applied. Operating temperature, density and pressure were defined as ambient airflow and for Radiation mode, S2S method was used. The default values for the relaxation factors were applied. The residual for continuity, x-momentum, y- momentum convergence criteria was taken as 10-3, on the other hand, for energy convergence it was taken as 10-6. In this thesis, in order to study the solar cavity receiver, six types of validation studies were performed for validating the mathematical model and boundary conditions used at both numerical calculations and ANSYS Fluent. The numerical and thermal analysis for both pure natural convection and combined natural convection and surface radiation were studied for finding the main reasons of thermal losses. In all the validation studies, the computed Nusselt numbers were validated with the published ones. The first validation study was for a pure natural convection analysis in a closed square cavity. The results indicated that, as Rayleigh number increases, the fluid velocity increases and fluid gets more close to the walls of the cavity and thermal boundary layer and velocity boundary layer thicknesses decrease. The heated fluid circulates in clockwise direction in the cavity because of the buoyancy effect. The second validation study was to investigate the pure natural convection heat transfer in a fully open square cavity. A fully open, 2D, square cavity thermal analysis was performed by just changing the opening wall boundary condition as pressure inlet. The velocity vector results indicated that fluid enters from bottom part of the cavity, reaches to the hot wall, gets heated, and exits from right side of the top wall of the cavity by buoyancy effect. There forms both thermal boundary layer and velocity boundary layer near the hot wall but as the Rayleigh number increases the thicknesses of these boundary layers' decrease. The third validation study was to investigate the tilt angle of open cavity effect on the pure natural convection heat transfer. The result showed that, the minimum natural convection heat loss occurs when φ= 180o, whereas, the maximum heat loss occurs when φ = 60o for the Rayleigh numbers of 103 and 104, and when φ = 75o for the Rayleigh numbers of 105 and 106. As the tilt angle increases up to φ = 60o for Ra= 103 and 104 and up to φ = 75o for Ra= 105 and 106, the heat loss increases slightly and then after that angles the heat loss decreases as dramatically. The fourth validation was to investigate the effect of opening ratio of cavity aperture on the natural convection heat transfer. The Nusselt numbers calculated for a partially open square cavity were compared with the ones for a fully open square cavity and then the optimum one was discussed for the heat loss reduction. In addition, isotherms and streamlines of a partially open square cavity were validated by comparing with the published results of study of Bilgen and Oztop. The result indicated that increasing the aperture size of the cavity increases the heat loss, so that using a partially open square cavity can reduce the heat loss. Finally, the last two validation studies were to validate the radiative model (S2S) used and to validate the Nusselt numbers found from combined heat transfer analysis by comparing with the published ones. The validation succeeded and the results indicated that, as the Rayleigh number increases, heat loss increases and radiation heat loss becomes dominant at high Rayleigh numbers. After validation studies, thermal losses in a solar cavity receiver were investigated by analyzing different parameters such as temperature difference, cavity dimension and cavity tilt angle. The temperature difference and dimension changing of cavity were studied just for Ra=106 and the results were discussed for investigating the heat loss. The results indicated that, as temperature difference increases, the heat loss increases mostly due to radiation heat loss. The Nusselt number for radiation increases dramatically as increasing the temperature of the hot wall, on the other hand, the Nusselt number for natural convection remains almost same. The increase in the length of the solar square cavity makes radiation heat loss dominant because the heat loss of convection decreases from % 97 to % 42. The results of the study performed for investigation of the tilt angle effect on the heat loss in solar cavity indicated that there is a nonlinear relationship with heat loss and tilt angle. To come up, background of a cavity such as a closed cavity and then an open cavity were studied. Different cases of open cavity were studied such as a fully open cavity, a partially open cavity, an inclined open cavity and all the results were compared with the published similar studies, and the best optimum cavity model at the highest thermal performance was proposed. Both pure natural and combined heat losses were studied and the dominant one was determined.
Numerical model of flue gas desulfurization spray tower in power plant
Flue gas desulphurization system is designed by many companies for power plant system due to the environmental regulations. As coal is the main energy source, many power plants were built in coal-reserve areas of Turkey. Besides the plants were built, many FGD systems were also built by different global companies. Every company has its own FGD technology to clean the gas before releasing from the flue. These technologies are developed and protected by patent applications. The Babcock & Wilcox Company, Lenthes Bischoff GMBH, Mitsubishi Heavy industries and Noell-KRC Umwelttechnik GMBH are well-known FGD companies which they were built Kemerköy, Yatağan, Afşin and Orhaneli power plants in the 1990s. These plants powers are various from 210 MW to 360 MW. FGD systems have been installed in as many power plants as ever, but there is still no method to accurately predict their behavior. It is desired to study the flow in the column in detail to achieve optimum column design, which is its difficulty. Among of FGD systems design requirements, cleaning of flue gas at high (90% and above) rates, maximum desulphurization and minimum pressure drop, is a priority. In this thesis numerical model of the flue gas tower is applied to a real scale power plant and investigate the performance of the column. Some design factors are selected to compare the results. These are (1) adding spray level, (2) integrate perforated plate, like Babcock & Wilcox patented tray and (3) putting a ring to reduce leaks either flue gas or slurry liquid due to the column wall. Numerical analyses in FGD system designs have become one of the most widely used methods today due to the development of computational fluid dynamics (CFD) techniques and increase in computational power. As a result of the CFD analysis, detailed flow field estimates are obtained, and this provides a better understanding of dynamics. In the literature review given in the introduction part of the thesis, besides numerical modeling studies, laboratory scale testing and mathematical modeling studies were included. Studies on FGD systems in the literature have been investigated and referenced in the thesis content. In the section on physical systems and simplifications, the modeling steps of the FGD system are given in titles. To summarize, the modeling steps are respectively solved by creating the geometry, extracting the flow volume, dividing the flow volume into the finite elements, meshing, and the geometry that is ready for model definition. Modeling steps of the flue gas system have also been given in this order. The FGD system, the column geometry, and piping in the spray levels were primarily created with Siemens NX cad program. Spray levels are modeled as surface due to simplification of the geometry. The prepared geometry was transferred to Ansys-SpaceClaim program and necessary definitions were made and sent to Ansys-Meshing program. Appropriate size meshing was created and necessary surfaces in the boundary definitions were also named. In Fluent program, it was solved by making appropriate definitions and results were obtained. The information obtained from literature and with some simplifications, the flue gas desulphurization system was modeled on the power plant scale. These simplifications are that the density of the sprayed droplets is much greater than the flue gas, and as a result, the forces can be neglected other than drag force. In the calculation of the drag force acting on the droplets, it is assumed that the droplets are in spherical form. At the same time, it was considered two-way interaction between the gas and liquid phases in the flue-gas column. Eulerian and Lagrangian approaches generally used for mathematical modeling of two-phase flow. The advantage of Lagrangian approach is that droplet physical properties (diameter, density, etc.) can differ in. In this thesis, Euler-Lagrange approach is used for these reasons. When the gas phase is defined as Euler, the droplets are modeled as a discrete phase by Lagrange trajectory approach. The modeling of the discrete phase is defined detailly. Especially since working FGD model has too many nozzles, a total of 460 nozzles, must be defined together with the position informations and the properties of each injection. For this process, MATLAB program and Excel file that define the injector properties are used. How the droplets will behave after hitting the column wall is governed by the added UDF code. Reflection angles and velocities are calculated by using the velocity and incidence angle of each droplet that hit the column wall. Numerical and test results of Weiss and Wieltsch were taken as reference in these calculations. First, FGD column was analysed without the injectors to provide convergence for the gas phase, and the results are given as a comparison. After the gas phase converges, the discrete phase is activated with unsteady particle tracking. While the quencher level and the second spray level are operated in all analyses, the first spray level, which usually reserves in power plants, is operated in a single analysis condition. When the tray is modeled, the porous-jump boundary condition is applied on the surface where the tray is located. The coefficients necessary for porous-jumping were found out from the literature with perforated-plate values of 50% porosity and estimated the height of slurry flow values to be accumulated on the plate were entered. If FGD column has a ring geometry, a ring surface was created in the model geometry, and wall boundary condition was defined for this surface. The results and interpretations are given in the last two sections, with each analysis of the resulting velocity and pressure figures, and the effects of comparative graphics and design factors are summarized. From these sections, it is seen that the pressure drop in the flow of the flue gas is observed as a corrective effect of flue gas distributions. However, this pressure drop will increase operating costs. Efficiency reduction of the desulfurization systems is mainly caused by colliding gases in the column walls. The inability to place a nozzle near the wall areas and the flow of the sprayed slurry from the wall cause inefficiency. Therefore, the design proposed by the ring addition analysis clearly shows that the flue gas escaping without desulfurization accelerates to the center of the column and helps to meet with the slurry. In the appendix, simple implementations that can be done before the modeling of the flue gas systems are given. UDF code for controlling droplet-wall interaction and some numerical results which were used to create figures were given in the appendix, too. In addition, since the study requires for defining a hollow-cone nozzle, the parameters to be defined for the hollow-cone nozzle, have been summarized from the Ansys Fluent User's Guide document along with explanations.
Solar unmanned aerial vehicle design and production
This thesis specifies the requirements, design methods and material selection of a solar powered unmanned aircraft that designed for specific mission. The mission of this UAV defined as observation with payload of at least 500 g. The UAV flies 24 hours long to conduct observation mission continuously. Flight mission includes four legs: takeoff-climb leg, level flight leg, glide-climb leg and final glide. Solar powered UAV is propelled by the power generated by PV cells during daytime at the same time excess generated electric energy is delivered to Li-Ion batteries. After the day flight, solar powered UAV change its flight regime and it starts to use the power deposited by Li-Ion batteries. During night flight, UAV has also glide-climb pattern to reduce power that consumed by propulsion system. The design phase consists of three main parts as; conceptual, preliminary and detailed design. Firstly, mission requirements and design requirements described in conceptual design. Also, configuration selections were completed according to design criterias to provide most suitable design that provides maximum efficiency. In the preliminary design phase, aircraft starts to be formed basic sizing of aircraft and some weight estimations. Following stages, aerodynamic design is completed and analysis is conducted to meet mission requirements which needs to long hour flight strictly. After the power calculations, selection of components such as, motor and propeller, ESC, battery, MPPT and solar cells, were made. Aircraft dimensions were determined after calculations and performance analysis. Solar energy calculations were made for selected mission location and outputs of these calculations were compared to aircraft performance requirements. In the detailed design process, computer modelling of every components were made and proper layout was prepared. Main components detail design and material selection were maden to provide strength that is required to loads during flight. All processes were completed successfully to achieve requirements of solar powered long endurance mission.
The six degrees of freedom modeling and simulation of launch vehicles including sensitivity analysis
For the purpose of this thesis, a six degrees of freedom trajectory model was developed and verified for launch vehicle systems and an already operated launcher Minotaur-I was simulated with this tool. Moreover, the same tool was also used for sensitivity analysis for several important parameters of the vehicle. Launch vehicles or carrier rockets are used to transport one or more payloads from Earth's surface to space and launch vehicles are used for both commercial or military satellites, meteorological observation, and experimental researches. From the beginning of the space race, a variety of launch vehicle systems have been developed and operated to reach space, accomplish the assigned mission and gain global prestige. A six degrees of freedom equations of motion model using translational and angular motions was developed with its library functions such as coordinate systems transformation, mathematical operations, Euler and quaternions transformation, atmosphere and gravity models. US Standard Atmosphere Model 1976 and EGM2008 Gravity Model were implemented to use with equations of motion. Nine different check cases data for a simple sphere, brick and two-stage rocket verified by NASA were used for verification of the developed tool. The verification process included gravitational, translational, rotational equations of motion, inertial coupling, integration method of differential equations, rotation of the Earth, ellipsoidal Earth, Coriolis effects, staging and entire atmosphere verification steps, respectively. In addition to verified library functions and equations of motion with rotating oblate Earth, an example mission of Minotaur-I launch vehicle was simulated. For the modeling of Minotaur-I, CATIA for mass and inertia model and Missile DATCOM for the aerodynamic model were used and the desired parameters were achieved at the end of the simulation. The sensitivity analysis of some essential parameters for flight path and performance of launch vehicles were fulfilled after the successful simulation. The analysis parameters involved gross lift-off weight of the vehicle, specific impulse of each stage, pitch-over maneuver initial and final times, each stage ignition time and pitch-over maneuver angle. The analysis showed that the specific impulse of the first stage is the most dominant parameter among the whole variables. Overall results demonstrated that even small modifications in launch vehicle parameters play a substantial role in launcher trajectory, performance and reliability.
Lightweight design for an engine structural part by using topology optimization for additive manufacturing
Lightweight design is becoming increasingly more important for vehicle design in aerospace and automotive industry. Although very important advances achieved in product design for both industry, some innovative designs could not be implemented due to the production restrictions. By means of emerging technologies in additive manufacturing, the ways of product design, development and logistics are transforming. Current advances in AM materials and creative design methods enable AM integration for lots of industries. One of the major tools of the creative design is topology optimization which was used for years in product optimization. However, lots of design restrictions observed due to the manufacturing constraints such as mold direction, draft angle, equal wall thickness, etc. By means of AM, topology optimization results have the possibility to be directly integrated as much as possible. Design freedom is the most important advantage of the AM. It is obvious that both designers and manufacturers have this freedom. Designers can create the designs in a wide range of alternatives, suppliers can produce the complex features easier than the conventional production methods. Although design implementation is easier than the conventional methods, design verification is a must and have to be applied as part and system level. The verificaiton methods are nearly same compared to the conventional designs. As digital verification, modal analysis and structural stress analysis have to be aplied on the design. In literature review, mainly focused on topology optimization and integration with AM. First, an automotive driveline bracket topology optimization example shared according to the conventional production methods. Design and non-design space definition, modal assesment and weight reduction studies showed in the example.It has been very common application in automotive and aeorospace application up to now. Then, a re-design study of an AM aerospace hinge bracket shared to show the transformation of the design while changing production method. Very significant weight reduction obtained by means of material change and design update. Lastly, different AM design approaches showed while integrating TO results to the conventional design. Direct integration of TO or partly integration of TO and standart CAD remodelling of the part is possible. Most important advantage of the partly integration is to having updatable 3D model to enable model updates in any phase of the product lifecycle. Topology optimization theory and methods explained shortly in second section. Although size, shape and topology optimization is used as the structural optimization methods, only topology opimization used as the optimization method in this thesis. Density and homogination methods explained also. Additive manufacturing history, definition, technologies, materials and dimensional accuracies explained in third section. Clear understanding of AM capabilities and currrent apllication is very crucial for a successful design integration in all industries. For a successful lighweight design integration, AM manufacturing methods have to be known very well to prevent design phases. Effective software tool usage are also very important for the design transformation and virtual verification of the updated design. Catia V5 used for design space creation and meshing works completed in Hypermesh. Then, FE model creation, modal analysis and stress analysis performed in Abaqus. As a post process tool, Hyperview used to view the analysis results. After completion of the design space analysis, TO initiated in Tosca which is compatible with Abaqus models. According to the TO results, first 3D model phase created and modal analysis performed to check whether the target achived or not. Due to the lower values of first mode, further phases were required. Afterwards, second 3D model phase performed to get higher frequency results. Finally, modal and stress analysis of second phase performed in Abaqus and targets achieved for both of them. Final model has very complex features for conventional production methods, also manufacturability of the final part is nearly impossible by using conventional manufacturing methods. On the other hand, AM feasibilty seems very promising with the help of advanced capability of the AM technologies.
Energy-based metrics for fighter aircraft assessment
Aerial combat is the cornerstone of military operations. One of the seminal elements that determine its outcome is aircraft maneuver. It is not easy to decide how fighter aircrafts will perform in combat, since reliable data regarding them are not usually available. Aircraft manufacturers and air forces frequently assert that their fighters are more superior than other equivalent fighters. However, they do not provide the real capabilities information of their fighters due to confidentiality reason. This study focuses on using existing flight envelopes and energy-maneuverability metrics for fighter aircraft assessment. The flight envelopes will be used to evaluate and compare fighter capabilities while the metrics will be applied to a decision process used to determine the best input strategy for various combat maneuvers. A fighter aircraft is hardly more superior than its adversary at every altitude and at all speeds. A well experienced pilot knows how to benefit from the weak points of the adversary. He/she may force the adversary at the points where his/her aircraft possesses obvious advantages and launches offensive actions. Pilot, who enjoys and can maintain this advantage, may not only control the combat but also determine its outcome. In this study, three diagrams that can be used to assess aircraft performance and maneuverability during combat will be discussed. They are V-n diagram, Specific Excess Power (P s ) diagram, and Energy-Maneuverability (E-M) diagram. V-n diagram shows the instantaneous maneuvering capability of an aircraft while P s diagram displays its sustained maneuvering capability. E-M diagram shows both the instantaneous and sustained maneuvering capabilities. Three comparison techniques using these envelopes, will be presented. They are V-n overlay, P s overlay, and E-M overlay. The fundamental rules for a successful aerial combat are quick kill, energy management, and assessment of transient characteristics. During combat, pilot must continuously evaluate his/her energy while at the same time assess the energy of the enemy. In an effort to gain energy advantage, both opponents will trade altitude for airspeed or airspeed for altitude many times. Energy should not be expended either too fast or too large. The assessment of aircraft transient characteristics allows us to give a final remark in determining the best input strategy for fighter maneuvers.
Optimal aircraft trajectory planning based on high-resolution actual weather data
The aim of the graduation thesis Optimal Aircraft Trajectory Planning based on High-resolution Actual Weather Data is to enhance the aircraft flight plans in terms of wind-optimal speed and altitude scheduling with the utilization of the spatial and temporal high-resolution wind and temperature data. It is tried to utilize the weather information and the optimization process from the viewpoint of ground-based aircraft trajectory prediction. To achieve the objectives of this thesis firstly, a bunch of the aircraft flight plans is precisely simulated using the BADA4-based high fidelity trajectory predictor and compared with the actual flown data retrieved from the respective QAR source to observe how closely the planned trajectories were followed in actual flights. Secondly, with use of the wind forecast in the flight plan, aircraft trajectories are optimized respectively and the potential fuel and/or time-saving are observed. Thirdly, With the investigation of the impact of the wind uncertainty, the aircraft trajectories are optimized utilizing the actual HRES wind data from QAR and the change in fuel consumption and elapsed time are observed. Next, utilizing the cost index adjustment algorithm as a Fine tuning process for the wind-optimal trajectories, a novel process is performed as a re-optimization process to define the wind-optimal CI optimized according to the wind data. The proposed dynamic wind-optimal CIs are cumulatively equal to the CI defined by the airlines. Finally, a detailed analysis over a set of 400 long-haul flights is done to show the impact of the wind uncertainty over the trip time and trip fuel. Also, the cumulative potential fuel and/or time savings for multiple case studies are illustrated in details along with the aggregate results. It is also emphasized that wind speed gradient with altitude should be considered as well. In result, it is seen that cruise Mach number and cruise altitude together are key parameters for fuel consumption under the wind effect. Hence, the uncertainty in the wind may cause flying at non-optimal conditions. Recovering this uncertainty and planning the best trajectory in strategical level could be advantageous for airline operations. Applying the novel idea of wind-optimal CI to the flown planned flights using high fidelity trajectory predictor, a fuel-saving purely related to the DCI method is being achieved. At this stage of the work, a short duration flight with a low number of waypoints is considered, which resulted in a fuel saving with maintaining the time of arrival. The analysis is done according to the operational flight plan and actual flight data. It is aimed to apply this method to different flights with various conditions to generalize and verify the proposed tool. The fuel savings achieved by the methodology described in this paper complies with the expected outlook by SESAR (300 - 500 kg on average). The wind uncertainty affects the strategic planning of flights. This study showed the differences between forecast and actual wind information leads to having different fuel consumptions. Considering how the actual wind deviates from the planned ones, fuel savings can be achieved by the methodology described in this paper. Therefore, having access to the accurate wind information, the approach in this work can be utilized to revise the strategic fight planning. The results show that this work is more suitable for applying to the long-haul flights. This graduation thesis is composed of six chapters, each of them dealing with different aspect of satisfying the objective of the thesis. Chapter one is introductory and describes the purpose of the thesis, the relevant works that have been done before and clearly defines the contribution of this work. The literature review is divided into three subsections; wind forecast and data assimilation, trajectory optimization and dynamic cost index to give detailed information about each part of this thesis. Chapter two is dedicated to the background information needed for this work and consists of three parts. part one analyzes the impact of the wind uncertainty over the aircraft trajectory and the along-track difference is defined as the metric for this analysis. part two focuses on the fundamental background information about the cruise phase of the flight and describes in details the cruise flight formulation, which is of the interest of this work. Part three is dedicated to the fuel-efficient trajectory optimization, divided into speed and altitude optimization. The speed optimization, which is the main part of this thesis is described in details with three different methods as Maximum Range Cruise (MRC), Long Range Cruise (LRC) and Economy/Cost Index (CI). In the end, the process of the optimal altitude generation is described in details. Chapter three deals with the data resources and data parsing process for this work. The acquired flight plans from Turkish Airlines are explained in details, then the actual flight data from the aircraft Quick Access Recorder (QAR) is described. In the end, the weather forecast, NOAA GFS data, are visualized and described. Chapter four focuses on the aircraft trajectory prediction tool and defines the process step by step. Firstly, it is described why the Eurocontrol BADA 4 is chosen as the core aircraft model of this work by comparing the Eurocontrol BADA 3 and BADA 4 functions such as aerodynamics, Engine thrust and fuel consumption model. Then, the aircraft motion modeling is defined as explaining the governing equations in details followed by modeling the aircraft dynamics as a point-mass model as in BADA 4. Chapter five is concerned with the methodology used in this work. Firstly, a detailed description of the proposed work and method is defined showing the workflow of this thesis, the inputs, processes, and the desired output. Then, the first core block, which is cost index adjustment is formulated and described in details. Speed optimization is also described in details defining the aircraft performance limits and how the solution of the optimization formula for the speed scheduling the optimal solution. Afterward, nine actual flight plans are simulated with proposed methods to show the performance of the developed trajectory predictor, which is desired and acceptable. Conclusions are drawn in chapter six along with the simulation results. Firstly, it is tried to show the exact impact of the wind uncertainty, the difference between the actual wind and the forecast wind, over the trip time and trip fuel over 400 long-haul flights. Then the simulation results of utilizing HRES wind data against low-resolution one is shown and discussed followed by the description of the generation of the artificial flight plans and analysis of speed schedule. Also, the simulation results of the CI adjustment is illustrated and discussed.
Stability analysis of nonlinear systems and upset recovery strategies for agile maneuvering aircraft
This thesis investigates the upset recovery problem of aircrafts as well as stability analysis of general nonlinear systems. In order to deal with upset recovery problem, the strategies which contains nonlinear control techniques are adopted. In order to validate through comparison, linear control technique which make use of linear quadratic regulator is also implemented. As a nonlinear control techniques nonlinear dynamic inversion (NDI) based angular rate regulator and double loop NDI controller for angle of attack, sideslip and bank angle are derived for High fidelity F-16 aircraft model. The choice of aircraft model enabled us to observe conditions with high angular rates which can be encountered in agile maneuver executions. Implemented controllers are evaluated through a novel framework where the recovery envelopes and recovery times are used as the performance criteria. Three dimensional recovery envelopes are constructed using three body angular rates which are used as initial points for simulations and included or excluded to recovery envelopes based on the success in recovery attempt. Successful recovery conditions are restricted so that the aircraft would not count as recovered if it exceed structural or altitude limits along with the constraints employed in actuator models. Based on the evaluations which are made by the proposed framework, another strategy is proposed which adopts the strong sides of single loop and double loop NDI controllers such that it utilizes the single loop NDI controller until the high angular rates are regulated below some predetermined threshold, then uses the double loop NDI controller to recover aerodynamic angles. This new strategy is called as switching NDI controller and its performance superiority is validated using the proposed framework. Thereafter conducting analysis based on simulation study, another approach is implemented to mitigate the computational burden which comes along with the simulations. Neural network model is constructed to reproduce the recoverability envelopes obtained via simulations. By this way, recoverability envelopes are created with high fidelity using only small portion of the whole simulations executed before.
Numerical and experimental investigation of the impact performance of 3d lattices with negative poisson's ratio
Poisson's ratio is simply the ratio between strain values of any two directions. The Poisson's ratio tells many things about the deformation behaviour of the material and what should be expected from material under specific loading conditions. Generally materials have Poisson's ratio between 0 and 0.5, mostly around 0.3. The negative sign in front of the formula causes conventional materials to have positive value. Conventional materials tend to expand in the transverse direction when they are under compressive loading in longitudinal direction. Also, they shrink in the transverse direction when they are under tensile loading in longitudinal direction. However, these load applications are just the opposite for the auxetic materials. They shrink when they are compressed in any direction, and they expand in all directions when pulled in any direction. This causes auxetic materials too have negative Poisson's ratio. Normally there are not too many natural auxetic materials. There are only few examples such as, some ligaments of muscles and joints, tooth gum of some animals. Most of the auxetic materials are man made. Even if the raw material has positive Poisson's ratio, auxetic material can be produced from it. Generally, unique geometry is used to create auxetic material. Without the improving technology, the production of the complex shapes would be so hard to process. For this thesis study, 3D printer is used to produce complex and complicated geometries. Also, the technical devices are designed and produced according to the needs. In this study, two auxetic and one conventional geometries are used. First they are designed in CAD software programme and then analysed in finite element programme. After obtaining required knowledge about their strengths and weaknesses, these geometries are produced from 3D printer. Then impact tests and edge crush tests are performed on them to see their responses. Comparison between numerical analyses and experimental results are done. The importance and promising features of auxetic materials are detailly analysed.
Agile flight trajectory tracking through differential flatness
The trajectory planning and tracking problem in aerial vehicles with the constraints becomes highly-challenging due to the complexity of their dynamics. This paper focuses on a real-time agile flight trajectory generation methodology through differential flatness description. The differential flatness enables an effective dimension reduction, therefore, enables a real-time solution to the problem. The desired flat output is parametrized with b-spline and the control points of the splines are obtained through a sequential quadratic programming, which maximizes proximity of the b-spline curve to some predefined waypoints as well as smoothing at predefined level. Thus, trajectory tracking problem is translated into an optimization problem such that resulting optimal feasible trajectory meets all the given and dynamical constraints. For the simulation and validation purpose, we have shown application of this methodology to both fixed-wing and multi-rotor systems and compared their maneuvers for a given set of waypoints. Finally, through knot and control point insertion, real time trajectory re-planning is addressed.
Geomanyetik fırtına zamanlarında İstanbul üzerindeki iyonosferik elektron yoğunluğunun istatistiksel olarak incelenmesi ve GPS haberleşmesi üzerine etkisi
As being the conducting layer of the atmosphere, ionosphere plays a very important role in the propagation of the radio waves in the atmosphere and thus, in the long-distance communication over the globe and with the satellites in space. Especially, with the advances in GPS satellite technology, the predictions of the variations in the electron density of ionosphere become critically important in many of the areas that these technologies are used ranging from the military and navigation applications to aircraft traffic control systems and extending to the cellphones which are one of the extremely essential part of our daily lives, to even target finding in golf playing. Needless to say, finding the loss during the earthquakes and other hazardous natural phenomena is one other very important application of GPS information. The variations in ionospheric electron density occur in response to the variations in the solar radiation which changes over the solar cycle with the solar activity, and in the magnetosphere with the magnetic storm and magnetospheric substorm activity as well as due to the propagation of the waves into the ionospheric levels from the lower regions of the atmosphere. Variations in one of these factors cause both temporal and spatial variations in electron density. While the neutral winds in the thermosphere move electrons from one place to another, thus causing latitudinal and longitudinal variations in the electron density, the more periodic variations, such as daily, seasonal, and annual, occur in response to the variations in solar radiation input of the atmosphere. On the other hand, solar and magnetospheric activities, which are described under the term of Space Weather, cause irregular and non-periodic variations and thus result in unpredictable variations in the electron density. Electrons in the ionosphere are produced as result of ionization of the neutral atoms of the atmosphere at heights from approximately 60 km to 1000 km. The energetic particles from the Sun and from the magnetosphere in addition to the daily solar radiation are the major source of the ionization in the Earth's atmosphere. Solar photons create ionization that peak around the noon when the energy of the photons is the strongest, producing daily variation of the electron density. The seasonal and annual variations are the other two predictable variations caused by the solar radiation amount reaching the ionosphere. During the solar storms, the energetic particles enter the atmosphere, give rise to magnetic storm and magnetospheric substorms, and causing northern and southern lights (auroae) over the high latitudes and other phenomena over the equatorial regions. Depending on the energy of the particles, these generate ionization as well. The effect of the magnetic and magnetospheric storms on the electron density have been studied quiet widely. The effects and the physical mechanisms that cause these effects were to found to vary with the latitude, i.e. equatorial versus high latitudes, time of the day (universal time, UT), and magnetic local time (MLT). The background variations when there are no magnetic storms or magnetospheric storms are called quiet day variations while the variations in the presence of magnetic storm or magnetospheric substorm are called active day variations. Studies show both positive and/or negative effects when the total electron density or total electron content of the ionosphere (TEC) is higher or lower than that occurring during the quiet days. In general, positive effects are seen towards noon and the negative effects are observed in the afternoon after 16:00 UT. The effects of the magnetospheric storms and substorms are still an active area of research and mostly concentrated on the equatorial and high latitudes. Studies on the mid-latitude electron density variations, especially during the magnetospheric activity, are very limited. Also, most of the results obtained are based on the case studies while a complete statistical study is not available. Ionospheric electron density is measured using ionosonde radars from the ground and satellites in space. The ionosondes are high frequency (HF) radars that send radio signal to the ionosphere and measures the electron density using the reflected signal. They are indeed the oldest methods to measure electron density and gives amount of electrons at a fixed location. Despite oldest, they still operate and widely available over the globe and thus makes the global variation of ionospheric studies possible. They are the techniques that measure the electron density from ground. Recently, with the GPS technology, it is possible to obtain TEC using signal delays which is another active research area. In the absence of ionosonde observations, these provide good estimation of ionospheric electron density. In addition, there are several ionospheric models, such as International Reference Ionosphere (IRI), that give predictions of ionospheric parameters including electron density and TEC. Recently other models that include ionosphere-magnetosphere coupling possible become available such as TIEGCM, CTIPe, etc. While IRI does not include the effects of magnetospheric activity, TIEGCM and CTIPe type coupled models are expected to give better agreements with the observations. In this thesis study, it is our purpose to investigate the ionospheric electron density variation during the magnetospheric disturbances over Istanbul using both ionosonde and available ionospheric models, namely IRI and TIEGCM and determine the effects of variations in TEC on GPS range and time delays. In addition to case studies, this study presents the statistical investigation of the substorm time ionospheric electron density. The main motivation for our work comes from the fact that mid-latitude variations of electron density are not investigated in our region. Recently, we established an ionosonde radar within ITU campus in October 2012 and provides electron density and TEC data since then. One of our goals in this study is to present the observations of ITU ionosonde in order to understand the variability of ionosphere over Istanbul during the magnetic activities and then determine the differences from the variations seen during the quiet days. The strength and the time of the differences is important for the predictions of the electron density over a region. The comparisons were made with the results from the ionospheric models to determine how much and when the models differ the most from the observations and thus bring out the points that need to be improved in the models to decrease the discrepancies. As practical side of our study, we study the effects of the magnetospheric substorms on the GPS range delay and time delay predictions. Also, we compare ionosonde TEC observations during one of the geomagnetic activity event, March 17, 2015, with those obtained from GPS signal delay data. If we highlight our study; the first work is magnetospheric activity time analysis which was done by using 3 years of ionosonde data from 2013 to 2015, determine the variations in electron density during the magnetospheric storms.First recognize the differences in case studies, then expand data set to cover the magnetospheric storms that occur during these 3 years. The parameters we used are the F2 layer critical frequency (foF2), height of the maximum electron density in F2 layer (hmF2) and TEC. This study addresses the variations during the substorms statistically. Secondly, model comparisons are represented and lastly estimation of time and range delay and the estimation of TEC using GPS signals are given. Chapter 1 introduces the ionosphere, its structure, and its effects on the radio propagation and information on the GPS and how it is used to obtain TEC. Chapter 2 presents some literature work on the ionospheric electron density variations and GPS comparisons. In Chapter 3, we present the results from ionosonde data during the magnetospheric activity. A case study was presented and followed by statistical results. In Chapter 4, comparisons with the results from IRI and TIE-GCM models were presented. Chapter 5 will give summary of conclusions and suggestions for a future work.
İnsansız hava araçlarındaki servo eyleyiciler için güvenilirlik analizi
Bu tez çalışmasında güvenilirlik kavramı, temelleri, güvenilirlik analizinin neden ve nasıl yapılması gerektiği anlatılmıştır. Havacılık uygulamaları için genel bir çerçeve çizilmiş, daha sonra insansız hava aracına takılmış bir servo eyleyici için örnek analiz yapılmıştır. FAA ve EASA'nın yayımladığı havacılık regülasyonları, uçuşa elverişlilik sertifikası için güvenilirlik analizinin yapılmasını şart koşmaktadır. ARP 4754A, ARP 4761 ve askeri standart MIL-STD 1629A gibi tavsiye niteliğindeki diğer belgeler, güvenilirlik analizlerin kapsamını ve nasıl yapılacağını tarif etmektedir. Bağlayıcı regülasyonlar ise bu tavsiye niteliğindeki belgelere atıfta bulunmakta ve bu belgelerin takip edilmesini tavsiye etmektedir. Bahsi geçen belgelerde yapılması tavsiye edilen İşlev zarar analizi (FHA: functional hazard analysis), uçak güvenliği ön değerlendirmesi (PSSA: preliminary system safety assesment), hata ağacı analizi (FTA: Fault tree analysis), arıza biçimleri, etkileri ve kritiklik analizi (FMECA: failure mode, effects and criticality analysis) gibi yöntemler hakkında genel bilgi verilmiştir. Analizler, işlev zarar analizi (FHA) ile başlar. Uçağın ana işlevleri belirlenir ve bu işlevlerin kaybedilmesinin güvenlik üzerindeki etkileri incelenir. Güvenlik üzerinde küçük etkileri olan işlev kayıplarının görece yüksek olasılıklı olması kabul edilebilir, ancak birden çok kişinin öldüğü, çoğunda uçağın düşüşü ile sonuçlanan felaket durumuna (catastrophic) neden olacak işlev kayıplarına yol açacak arızaların gerçekleşme olasılığı, 10-9/saat'i aşmamalıdır. Ayrıca arıza olasılığı ne olursa olsun tek bir parçanın arızasından dolayı felaket durumu oluşmamalıdır, yani tekil arızaya (single failure) imkân verilmemelidir. FHA analizi önce uçak için yapılır, daha sonra buradan çıkan sonuçlara göre alt sistemler için tekrar edilir. FHA analizlerinden sonra uçak güvenliği ön değerlendirmesi (PSSA) yapılmalıdır. Hata ağacı analizi, PSSA analizinin önemli bir öğesidir ve izin verilen arıza sıklıklarının en alt parçaya kadar indirilmesine izin verir. Bir parçanın arıza olasılığı gibi, arıza durumunda üst sisteme ne etki edeceği de önem teşkil eder. FHA, PSSA gibi "yukarıdan-aşağı" (top-down) yapıdaki analizler ile uçaktan en alt parçaya kadar inilerek her bir arızanın farklı biçimleri için olasılıklar, arızaların üst sistemlere etkileri, parça arızalar arası ortalama süresi (MTBF: Mean Time Between Failure) gibi isterler belirlenir. Parçaların veya alt sistemlerin kendi arıza sıklıkları ve arıza biçimleri bağımsız bir çalışma ile kestirilerek; hata ağacındaki ilgili yerlere yerleştirilmeli ve beklentilerin karşılanıp karşılanmadığı, olası arızaların üst sisteme öngörülenden farklı bir olumsuz etkisi olup olmadığı incelenmelidir. En doğru yaklaşım parçaların test edilmesi veya sahadaki kullanımından gelen değerlerin kullanılmasıdır. Ancak henüz geliştirme aşamasında olan, piyasaya sürülmemiş veya yeni sürülmüş parçalar için yeterince saha verisi yoktur, test ise çok maliyetlidir. Bu durumda benzer parçaların arıza sıklıklarından yola çıkılarak hazırlanmış ampirik formüller kullanılır. Elektronik parçalar için çok sayıda kaynak bulunmasına karşın mekanik parçaların arıza sıklıklarının kestirilmesi için en önemli kaynak, A.B.D. Donanması'nın yayımladığı Donanma Su Üstü Savaş Merkezi Carderock Bölümü'nün yayınladığı Mekanik Parçaların Güvenilirlik Kestirimleri için Prosedürler El Kitabı'dır. (NSWC – Naval Surface Warfare Center Carderock Division, Handbook of Reliability Prediction Procedures for Mechanical Equipment). Söz konusu kitapta yay, conta, filtre, rulman, kayış gibi kullanılan birçok mekanik parça için kestirimlerin nasıl yapılacağı tarif edilir. Çalışmada NSWC'nin yöntemleri ve hesaplama mantığı anlatılmıştır. Arıza biçimleri aynı kaynakta isim olarak geçse de arıza biçimlerinin oranları bulunmamaktadır. Arıza biçimleri için, Güvenilirlik Analiz Merkezi'nin yayınladığı CRTA-FMECA [28] belgesindeki bilgiler veya RIAC tarafından yayınlanan "Failure Mode/Mechanism Distributions – 2013 (FMD-2013) kullanılmıştır. Bu tez çalışmasında örnek çalışma olarak bir İnsansız Hava Aracı (İHA) sistemindeki bir servo eyleyici için arıza sıklığı hesaplanmış, arıza biçimleri ve olası üst seviye etkileri tespit edilmiştir. Yapılan çalışmalar, alt sistemin üst sisteme vereceği çıktıları içermektedir. Uçak sistemi seviyesinde yapılması gereken ve tüm uçağı kapsayacak olan FHA, FTA analizlerinin örnekleri, bu tez çalışması kapsamına alınmamıştır. Kullanılan eyleyici tasarımı ve boyutları, askeri alanda kullanılan parçanın gizliliği gözetilerek sadeleştirilmiş ve bir miktar değiştirilmiştir. Dolayısı ile sonuç rakamlar, gerçekte kullanılan örnektekinden farklıdır.
Hipersonik akışta şok-şok ve şok-sınır tabaka etkileşim mekanizmalarının incelenmesi ve ortaya çıkan yüzey isı transferi açısından değerlendirilmesi
Son zamanlarda atmosfer içerisinde hipersonik hızlara ulaşabilen uçak, roket vb. araçların geliştirilmesi için yoğun çalışmalar gerçekleştirilmektedir. Yüksek hızlarda şok dalgası etkileşimleri yapısal zararlar verebilecek yüksek ısınmalara sebep olduğundan bu tür araçların tasarımında karşılaşılan zorlukların en önemlilerinden birisidir. Bu çalışmada, hipersonik hızlardaki akışlarda karşılaşılan şok-şok ve şok-sınır tabaka etkileşimleri, bu etkileşimlerden kaynaklanan yüksek ısı transferi Hesaplamalı Akışkanlar Dinamiği (HAD) yöntemleriyle incelenmiştir. Tez kapsamında literatürde deney sonuçları yer alan ve hipersonik araçların birçok elemanında görülen iki açılı rampa geometrisi kullanılmıştır. İlk olarak, iki rampa açısının sabit olduğu bir durumda sürtünmesiz, laminer ve türbülans modelleriyle analizler gerçekleştirilmiştir. Elde edilen sonuçlar deney ve diğer hesaplamalı çalışmalarla karşılaştırılarak hangi akış modelinin gerçeğe en yakın sonuç verdiği tespit edilmiştir. Daha sonra iki ve üç boyutlu hesaplama ağı kullanılarak laminer bir akış için iki açılı rampa geometrisinin ilk açısı sabit tutulup ikinci açısı parametrik olarak değiştirilmiştir. Farklı ikinci rampa açıları için kısa ve uzun dönemde akışın fiziği anlaşılmaya çalışılmış, rampa yüzeylerinde oluşan yüksek basınç ve sıcaklık değişiklikleri değerlendirilmiştir. Bu çalışmanın literatüre en önemli katkılarından birisi, hipersonik hızlarda ulaşılan deney sürelerinin sonrasında hesaplamalı sonuçların akışın zamana göre periyodik olduğunu göstermesidir. İlk olarak, hesaplama modeli ve çözücünün doğruluğunu tespit edebilmek için iki boyutlu hesaplama alanında iki açılı rampa geometrisi üzerinde sürtünmesiz, laminer ve türbülanslı akış modelleri için zamana bağlı HAD analizleri gerçekleştirilmiştir. Hesaplamalar açık kaynak kodlu bir sonlu hacimler yazılımı olan OpenFOAM ile yapılmıştır. Birçok farklı çözücü barındıran OpenFOAM yazılımı içerisinde yoğunluk temelli bir Navier-Stokes çözücüsü olan rhoCentralFoam kullanılmıştır. rhoCentralFoam çözücüsü zamanda ve uzayda sırasıyla 1. ve 2. mertebe doğruluğa sahip olup Kurganov ve Tadmor'un merkezi upwind şemalarını kullandığı için akış alanındaki şok ve benzeri süreksizlikleri salınımsız temsil edebilmektedir. Gazın termodinamik özelliklerini sıcaklığa bağlı olarak modellemek için 7 katsayılı NASA termodinamik bağıntıları kullanılmıştır. Seçilen iki açılı rampa geometrisi etrafında bu çözücü yardımıyla elde edilen sonuçlar deney ölçümleri ve başka hesaplamalı çalışmalarla karşılaştırılarak hem hesaplama modeli hem de çözücü doğrulanmıştır. Ayrıca bu çalışmanın dışında deney verisi bulunan farklı açılara sahip başka bir rampa geometrisi için de çözücü doğrulama çalışması yapılmıştır. Bu doğrulama çalışması sonucunda laminer akış kabulünün deney ölçümleriyle en tutarlı sonuçları verdiği tespit edilmiştir. İkinci olarak, iki açılı rampa geometrisinin ilk rampa açısı sabit tutulmuş, ikinci rampa açısı θ2=40° ile 60° arasında 5° aralıklarla sistematik olarak değiştirilmiştir. Düşük entalpiye sahip iki boyutlu bir akışta farklı θ2 açıları için ortaya çıkan şok etkileşim mekanizmaları, akışın fiziği ve ısı geçişinde meydana gelen değişimler t=327 µs için Mach konturları ve yüzey üzerindeki ortalama ısı akısı dağılımları yardımıyla incelenmiştir. İkinci rampa açısı arttıkça yay (BS) ve üçlü noktadan ikinci rampa yüzeyine doğru ilerleyen şok (TS) dalgalarının güçlendiği, ayrılma bölgesinin hücum kenarına doğru genişlediği, çok sayıda yeni şok dalgasının oluştuğu ve daha geniş bir alanda etkili ısınmalar olduğu gözlenmiştir. Şok etkileşim mekanizması olarak, sadece Edney Tip V ve onun oluşum aşamaları görülmüş ve θ2=50°, 55° ve 60°'de akımın daimi olmadığı gözlenmiştir. TS'nin çarpma açısı ve şiddetinin, TS'nin çarptığı noktada oluşan ayrılma baloncuğunun boyutlarını belirlediği ve diğer yandan ayrılma baloncuğu büyüklüğünün de BS'nin durma mesafesi ve dolayısıyla TS'nin şiddetini değiştirdiği gözlenmiştir. Yani bu iki mekanizma rampa yüzeyi üzerinde sınır tabaka ve BS arasında iki yönlü bir döngü oluşturmuştur. TS çarptığı noktada ters basınç gradyanına sebep olmuş, bu bozuntuların sınır tabaka içerisinde akımüstüne doğru yayılması ise θ2 açısı arttıkça karmaşık ve daimi olmayan bir akışı tetiklemiştir. Bunlara ek olarak, bu döngünün oluşabilmesi için θ2 açısı için bir eşik değeri olduğu ve bu değerin θ2=45°- 50° arasında yer aldığı tespit edilmiştir. Önceki bölümde t=327 µs için değerlendirilen sonuçlardan sonra aynı problem bu sürenin 50 katına kadar süren uzunluklarda analiz edilerek akış alanı içerisinde periyodik bir değişim olma ihtimali araştırılmıştır. Önceki bölümde tespit edilen iki yönlü döngüdeki eşik değerini bulmak için θ2=45°- 50° arası 1° aralıklarla uzun dönemli çözülmüş ve eşik değerinin θ2=47° olduğu tespit edilmiştir. Bundan sonra ikinci rampa açısı θ2=50° ile 60° arasında 5° aralıklarla anlık ayrılma açısı, ayrılma şoku (SS) ve TS'nin anlık konumu, farklı istasyonlardaki anlık basınç ve ısı akısı takip edilmiş, yoğunluk gradyan ve girdap yapılarını görselleştirmede yardımcı olan Q kriteri konturlarıyla şok etkileşim mekanizması ve akış fiziği tarif edilmiştir. Uzun dönemli sonuçlar, θ2=47° eşik değeri altında iki açılı rampa üzerindeki akışın daimi olduğunu, bu eşik değerinin üzerinde ise akışın zamana göre periyodik olduğunu göstermiştir. θ2 açısı arttıkça akış alanı içerisinde şu değişiklikler gözlenmiştir: rampa köşesi civarında girdap sayısı artar, hücum kenarından ayrılma noktasına olan mesafe kısalır, ayrılma bölgesi kalınlaşarak uzar ve TS'nin rampa yüzeyine çarpma süresi kısalır. Elde edilen sonuçlar θ2=60°'nin periyodik akış için bir limit olduğunu göstermiş ve zamana göre periyodik davranış ile ikinci rampa uzunluğu arasında bir ilişki bulunabileceği belirtilmiştir. Son olarak, akışın iki boyutluluğunu koruyup korumadığı ve üç boyutluluğun getirdiği etkilerin incelenmesi için üç boyutta θ2=55° için analizler gerçekleştirilmiştir. Elde edilen üç boyutlu sonuçlar t=327 µs için deney ve başka hesaplamalı çalışmalarla karşılaştırılmış ve iki boyutlu hesaplama alanından elde edilen sonuçlarla arasındaki farklar değerlendirilmiştir. İki ve üç boyutlu sonuçların ikisinde de akışın daimi olmadığı görülmüştür. Hem iki hem de üç boyutlu sonuçlar için yoğunluk gradyan konturları analiz süresinin başlarında Schlieren görüntüleriyle birebir örtüşmektedir. Ancak ilerleyen zamanlarda üç boyutlu sonuçlardaki yoğunluk gradyan konturları Schlieren görüntülerinden uzaklaştığı gözlenmiştir ve sebebinin üç boyuttaki rahatlama etkileri olduğu gösterilmiştir. İki ve üç boyutlu ısı akısı dağılımları karşılaştırıldığında, üç boyutlu sonuçların deney ölçümleriyle daha uyumlu olduğu fakat TS çarpma noktasının akımaltına doğru kaydığı gözlenmiştir. Üç boyutlu sonuçların yer aldığı bölümde θ2=55° değerlendirmesinden sonra, iki açılı rampa geometrisi önceki bölümlerde olduğu gibi ikinci rampa açısı θ2=45° ile 60° arasında 5° aralıklarla değiştirilmiştir. İki boyutlu sonuçlara benzer olarak, θ2=50°, 55° ve 60°'de akımın daimi olmadığı gözlenmiştir. İki ve üç boyutlu sonuçlar karşılaştırıldığında aralarında şu değişiklikler gözlenmiştir: θ2=45° açısında bütün anlarda hemen hemen aynı şok yapıları görülürken θ2>45° açılarında sadece yaklaşık ilk 100 µs'de aynı şok etkileşim mekanizmaları görülmüş, θ2=50° açısında 150 µs sonrasında SS ve BS'nin konumları iki boyutlu sonuçlara göre akımaltında kalmış, θ2 açısı arttığında bu iki şok dalgasının iki ve üç boyutlu sonuçlardaki konumları arasındaki mesafe daha da artmıştır. Bu sebeple, θ2=55° ve 60°'de TS çarpma noktasının konumu iki boyutlu sonuçlara göre önemli ölçüde farklılaşmıştır. İki boyutlu sonuçlarda ayrılma bölgesi içerisinde ayrılma balonlarından dolayı birçok ek şok dalgası oluşurken üç boyutlu sonuçlarda daha az ayrılma baloncuğu ve bu ayrılma baloncuğu kaynaklı daha az şok dalgası oluştuğu gözlenmiştir. θ2=45° haricinde bütün θ2 durumlarında akış tamamen üç boyutludur. Üç boyutlu rahatlama etkisi sebebiyle bütün θ2 açılarında rampa kenarlarına doğru BS'nin akımaltındaki sesaltı bölgenin küçüldüğü, SS'nin akımaltına doğru çekildiği, ayrılma bölgesinin akımaltına doğru daraldığı ve inceldiği ve TP'nin rampa yüzeyine doğru yaklaştığı gözlenmiştir. Özellikle θ2=55° ve 60°'de TS'nin çarpma hattında ayrık yapılar oluşturan akımaltına doğru uzanan Görtler tipi girdapsı yapılar tespit edilmiştir. En yüksek sıcaklıklar TS'nin çarpma hattında meydana gelirken Görtler tipi girdapsı yapılar θ2=55° ve 60°'de TS çarpma hattındaki yüksek ısınmaların akımaltı veya akımüstüne doğru taşınmasına neden olmaktadır. Özetle, üç boyutlu hesaplama alanından elde edilen sonuçlar, özellikle θ2=55° ve 60°'de görülen akışların daimi olmayan, Görtler tipi girdapsı yapıların yer aldığı üç boyutlu doğasını ortaya çıkarmıştır. Sonuç olarak, ikinci rampa açısı arttıkça karmaşık ve daimi olmayan bir akışı tetikleyen sınır tabaka ile BS arasındaki iki yönlü bir döngünün olduğunu ortaya çıkarması bu çalışmanın literatüre sağladığı ana katkılardandır. Ayrıca bu döngünün oluşabilmesi için bir eşik değerinin olduğu, bu eşik değerinin de θ2=47° olduğu tespit edilmiştir. Literatüre yapılan katkıların en önemlilerinden birisi de hipersonik hızlarda ulaşılan deney sürelerinin sonrasında hesaplamalı sonuçların akışın zamana göre periyodik olduğunu göstermesidir. Görülen zamana bağlı periyodik hareketin hipersonik bir araç tasarımında kısıtlayıcı bir faktör olabileceği ve bu sebeple bu çalışmanın bulgularının hipersonik hava araçlarının tasarımında dikkate alınmasını önemli kılmaktadır. Literatüre yapılan katkıların bir diğeri de üç boyutlu akışlarda tespit edilen Görtler tipi girdapsı yapıların yüksek ısınma bölgelerini akımaltına veya akımüstüne taşıyabilmeleridir. Bu sonuç, yüksek hızlı araçların tasarımında akış içerisindeki girdap yapılarının tespit ve takibinin yapılmasını önemli hale getirmektedir.
Effect of chamfered nozzle geometry on impinging jet flow
The widespread usage of impingement cooling makes this technique popular. From simple hand drier to advanced gas turbine, most of the industries use this technique for drying or cooling. Therefore, many studies both experimental and numerical are conducted to study every aspect of this technique. However, for optimum usage of this technique, the affecting parameters should be analyzed and designed accordingly. There are important parameters, which directly influences the heat transfer efficiency of jet impingement. These nozzle parameters are Reynolds number, depending on the flow speed and sizing of the jet nozzle, Mach number, jet-to-target spacing, jet-to-jet spacing and finally jet geometry. So far, variety of geometries have been tried and analyzed to study the flow structures and contribution to the heat transfer. Because, studying the flow structures would bring us at the end to the optimum shape of efficient heat transfer by impinging jet technique. Through this study another jet geometry, which is the chamfered, was studied. The experimental study was conducted by 00, 300 and 450 chamfered nozzles for two different Reynolds number (9,600 and 19,200) and for three different jet-to-target spacings (H/dn = 4, 6, 10) to observe the effect of these parameters. To understand the flow field, hot wire anemometer measurements were conducted and backed up with pressure measuerements.
Experimental investigation of flexibility effects in flapping wing aerodynamics
In the present PhD thesis, flexibility effects in flapping wing aerodynamics are experimentally investigated. For that purpose, one rigid and three chordwise flexible wings are produced and their aerodynamic performances are compared in a water channel. Digital Particle Image Velocimetry (DPIV) system is used to determine instantaneous velocity fields and vortical structures around the wings in conjunction with simultaneous direct force/moment measurements. The imaging system also acquired the instantaneous deformation of the wing. Natural frequency effects and deformation of the flexible wings are linked to the force measurements and efficiencies. Since all natural flyers and swimmers have some degree of flexibility on their wings, it is important to understand the phenomena for a more effective aerial/nautical vehicle design. Main objective is to determine the aerodynamic performances of different flexible wings under harmonic plunging or rapid pitch-up motions. The experiments are performed in the close-circuit, free-surface, large-scale water channel located in the Trisonic Laboratories at the Faculty of Aeronautics and Astronautics of Istanbul Technical University (Istanbul, Turkey). Two distinct motions are executed in the experiments. Perching or pitch-up is one of these motions and based on sudden change in angle of attack and halting the wing at a constant degree of angle. The pitch-up motions of the airfoil were accomplished using a Kollmorgen/Danaher Motion AKM33E servo motor. The models were subjected to performing two types of pitch-up motion: fast and slow. The model started from 0° and attained its final angle of attack of 45° in 1 second for the fast motion and in 6 seconds for the slow motion, corresponding to 1 and 6 convective times, respectively. The motor motion profiles were generated by a signal generator Labview VI (Virtual Instrument) for the given amplitude and duration. Labview VI also triggers both the force data acquisition and the PIV system. The other motion type is the purely plunging motion that is generated by harmonic cosine position function. Plunge motion of the model is accomplished with Kollmorgen/ Danaher Motion AKM54K servo motor. The motor motion profiles were generated by a Labview VI for various amplitudes and frequencies. The experiments are performed on four different wing materials: a rigid wing produced of plexiglass, three flexible wings produced of homogeneous polycarbonate sheet (lexan), an acetate wing, and an elastic model produced in 3D printer. The rigid wing has a thickness of 5 mm, chord length of 100 mm and transparent to illuminate both sides of the wing. Flexible wings have 10 mm metal part on leading edge to avoid spanwise flexibility and the total chord length is 100 mm. The wing's rotation axis is the leading edge, which also coincides with the center of the force sensor. Natural frequencies of the manufactured wings are one of the backbone parameters of the thesis. Motion parameters are determined based on these natural frequencies. The wings are exposed to sudden loading and given free chordwise oscillation to determine their natural frequencies while the wing is in its position, fully submerged in the water channel. Natural frequencies are determined as 7.40 Hz, 2.32 Hz, 1.25 Hz, 0.51 Hz for rigid, lexan, printed and acetate wings respectively. Force data is collected by a six-component ATI NANO-17 IP68 Force/Torque sensor. The sensor was mounted on the rod between the model and the plunge servo motor, oriented with its cylindrical z-axis coincides with the leading edge of the wings. For the perching rigid wings, whether the plate pitches up fast or slow, the flow remains mostly 2D along the span for plates with AR = 4 and above. During the motion, the effect of the AR is to shift the force variation in amplitude without changing its variation character. The effect of AR, hence, the three-dimensionality of the flow becomes pronounced after the motion ends. The formation of a local maximum in the force variations after the motion ends was evident for wings with AR > 2. There exist two peaks for wings with AR ≥ 6. These secondary or even tertiary peaks occur because of the formation of a new LEV. Even if the effective ARs of the wings are the same, the one with a symmetry plane did not yield exactly the same flow topology and hence the same force-time histories as the fully submerged wing. There were slight differences due to the three-dimensionality of the flow. Therefore, an endplate representing the symmetry plane should carefully be placed with special care to ensure that the gap between the wing root and the endplate is approximately 2 mm (i.e., 2% of the chord). Perching motion in flexible wings primarily depend on the deflection of the wings caused by the oncoming freestream. The deflection is high for the most flexible acetate wing, which results in a lower angle of attack. As a result, low CD is obtained for the acetate wing. When flow structures are examined, streamline topology of the acetate wing shows differences due to the reduced angle of attack both in fast and slow motion cases. Continuous LEV and TEV vortex shedding can be observed on the acetate wing. For the plunging rigid wings, it is found that the phase lag between the motion and the cross-stream force component decreases as the reduced frequency increases, and for a given reduced frequency, the phase lag increases as the plunging amplitude increases. The variation of the horizontal spacing between vortex cores with the reduced frequency follows the same characteristics as those of the phase lag for the cross-stream force. On the other hand, the phase lag for the streamwise force component decreases as the plunging amplitude increases for low reduced frequencies. The decrease converges to a phase lag value of 90°, which coincides with the maximum velocity of the wing. After the minimum is reached, the phase lag for the streamwise force component starts to increase. The minimum phase lag coincides with the reduced frequency where the crossover from drag to thrust is observed and maximum flapping efficiency is obtained. When the thrust is produced, the phase lag for the streamwise force component increases. Moreover, according to the vortex structures, the crossover from drag to thrust is observed when the LEV starts to hold around the leading edge and does not convect downstream along the plate. In flexible flapping wings, trailing edge amplitude makes a peak when the oscillations are performed near natural frequency of the wing (f/fn ~1). For the acetate wing, frequencies beyond natural frequency can be tested and after f/fn = 2, trailing edge amplitudes start to increase again. The values are in a continuous increase in the measurement range up to f/fn = 4.4. This finding suggests that, higher trailing edge amplitudes can be reached when the flapping frequencies are beyond the resonant frequency. When Φy phases are considered, it is observed that the flexible wings have completely different phase variations compared to the rigid wings. Natural frequency appears to have an effect on Φy phase variations. Phase values increase up to the near natural frequency, and then exhibit a decreasing followed by an increasing trend at higher flapping frequencies (can be observed for acetate wing). Besides the natural frequency, flapping amplitude also has an effect on phases such as the peak near natural frequency diminishes for higher amplitudes in acetate wing. This implies that natural frequency and flapping amplitude should be interpreted together. An interesting result is obtained from the combined Φy - ΦTE phases graph: trailing edge phases and Φy force phases are intersecting at different plunging frequencies for different wings. The intersecting frequencies coincide with the natural frequency of the wing. In other words, cross-stream force phases and trailing edge phases are equal when the wings are plunged near their natural frequencies. Flexible wings have peak efficiency around natural frequencies where their trailing edge phases and Fy force phases also coincide. Unlike the rigid wing, distribution of the Φx phases have similar characteristics with the Φy phases in flexible wings. That similarity indicates that the flexibility effects are observed not only in cross-stream direction but also in streamwise direction. In flexible wings, elastic forces emerge in both streamwise and cross-stream directions coupled with the trailing edge deflection. The similarity between the streamwise and the cross-stream force phases is a result of these prevailing elastic forces over the other existing forces (circulatory and added mass). The efficiency peaks shift towards high Strouhal numbers when the flapping amplitude increases. The peak efficiencies are obtained in a Strouhal number range of 0.1 – 0.6 for the flexible wings. High efficiency zone in the literature was expressed in a Strouhal number range of 0.2 – 0.4 and for the flexible wings, this range is generally valid. The boundaries of the range are slightly extended for the flexible wings in different flapping amplitudes. Same vortical instances are attained at different stages of the motion cycle for the flexible wings. It is observed that the TE motion determines the vortex shedding timing. As the wing gets more flexible, vortex shedding from the TE delays and similar vortex structures on the TE are observed later in the flapping period. Chordwise flexibility under parameters presented herein, alters the TEV shedding timing and vortex strength but does not have a pronounced effect on LEV shedding. For the rigid wing, the vortices burst after one chord length in the wake. As the wing gets more flexible, the vortices start to burst late. The vortices in the most flexible acetate wing's wake preserve their coherency and can be tracked in full measurement region. Also number of vortices observed in the wake increases when the wing is more flexible. While the flexibility for the wing flapping around its natural frequency exploits and enhances the high velocity zones created around the wing with the TE timing, another low efficiency flapping frequency cases cannot perform that. It can be considered as "catching" the flow and transferring it to the wake through vortices, which is found to be crucial for high efficiency.
Uçak gövdesinde yalıtım malzemesinin taşınımla ısı geçişine etkisi
Halkalar arasında ve silindir yapılar içerisinde doğal taşınım ve zorlamalı taşınım olayı, ısıl enerji depolama sistemleri, yer altı elektrik iletim kabloları, güneş enerjisi kollektörleri, nükleer reaktörler, uçak gövdesi yalıtım malzemesi gibi birçok mühendislik uygulamalarında incelenen konular arasında yer alır. Kullanım amaçlarından biri ısı ve ses yalıtımı olan uçak gövdesi gözenekli yalıtım malzemesi içerisinde uçuş boyunca nem birikimi gerçekleşmekte, kabin ve dış gövde sıcaklık farkından dolayı yalıtım malzemesi içerisindeki su buharı hal değişimi ile sıvı veya katı hale geçebilir. Yalıtım malzemesinde gerçekleşen nem birikimi olayı uçağın ağırlığını arttırır ve malzemenin korozyona uğramasına neden olarak yalıtım özelliğini kaybetmesine yol açabilir. Yapılan çalışmada eş merkezli iç içe geçmiş halkalar arasında doğal taşınım, eş merkezli halkalar arasında gözenekli ortam olan doğal taşınım ve dış duvarın iç kısmında belli bir kalınlıkta gözenekli ortam olan silindir içerisinde zorlamalı taşınım olayı incelenerek bu üç problem için sayısal olarak elde edilen çözümler sunuldu. Yapılan tez çalışmasının bir diğer amacı ise sayısal olarak çözülen bu üç problem yardımıyla uçuş koşullarına göre uçaklarda kabin ve dış gövde arasında gerçekleşen hava akışını, gözenekli yalıtım malzemesinde ortaya çıkan nem birikimi ve su buharının hal değişimi olayını araştırıp inceleyebilmek ve sayısal olarak hesaplayabilmek için altyapı çalışması oluşturmaktır. Sonlu hacim methodu (FVM) ile ayrıklaştırılan akışkan ve gözenekli ortam için yazılan genel korunum denklemleri, basınç ve hız doğrulaması için kullanılan SIMPLE algoritması yardımı ile çözüldü. Eş konumlu ağ yapısının kullanıldığı bu çalışmada yön değiştirmeli kapalı yöntem ve üç bant genişlikli matris algoritması (TDMA), lineer cebir denklem sistemlerini çözmek için kullanıldı. İlk olarak iç ve dış duvarların sabit sıcaklıkta ve iç duvar sıcaklığının yüksek olduğu eş merkezli iç içe geçmiş halkalar arasında doğal taşınım olayı incelendi. Halkalar arasında sıcaklık farkından dolayı oluşan kaldırma kuvveti etkisi ile meydana gelen hava akışının sıcaklık ve akım fonksiyonu profilleri, belli açılardaki sıcaklık ve hız dağılımları, ortalama Nusselt sayıları farklı Rayleigh sayılarında elde edildi. İncelenen ikinci problem ise dış duvarın iç tarafında belli bir kalınlıkta gözenekli yalıtım malzemesi olan eş merkezli halkalar arasında doğal taşınım olayıdır. Duvar sıcaklıkların sabit olduğu halkalar arasındaki ısı geçişi ve hava akışı hesaplamaları farklı Rayleigh sayısı, Darcy sayısı ve gözenekli yalıtım malzemesi kalınlıklarında elde edilerek bu üç parametrenin etkileri irdelendi. Isıl dengesizlik kabulleri yapılarak ikinci problem sayısal olarak çözüldü. Son olarak dış duvar ile hava kanalı arasında gözenekli yalıtım malzemesi olan bir silindir içerisinde zorlamalı taşınım problemi ele alındı. Dikey simetri ekseni üzerinde silindirin alt tarafından belli bir sıcaklık ve hız ile kanaldan giren hava, yan yana bulunan gözenekli ortam ve kanaldan geçerek simetri ekseninin üst tarafından silindirin içerisini dolaşır. Daha sonra hava, kanal girişinin hemen üstünde yer alan çıkış bölgesinden tahliye edilir. Farklı Reynolds sayılarının, Darcy sayılarının ve gözenekli ortam kalınlıklarının zorlamalı taşınım olayı üzerinde etkileri tartışıldı. İncelenen bu problemde, gözenekli yalıtım malzemesinde ısıl dengesizlik kabulü yapılarak akışkan ve gözenekli ortam katı iskelet sıcaklığı profilleri, akım fonksiyonu, vektörel hızlar ve ortalama Nusselt sayısı hesaplamaları hesaplandı. Üç problem için elde edilen sonuçlar ile gözenekli ortamlı ve ortamsız eş merkezli halkalar arasında gerçekleşen doğal taşınım olayı ve dış duvarın iç kısmında belli kalınlıkta gözenekli ortam olan bir silindir için zorlamalı taşınım olayı ifade edildi. Bu sonuçlar, tez çalışması sonrasında incelenecek olan uçak gövdesi ve kabin arasında gerçekleşen ısı ve kütle geçişi olayları için alt yapı ve temel bir bakış açısı oluşturulması konusunda faydalı olacaktır.
Data-driven aircraft performance models for improving baseline fuel estimations
In this era, modern life in globalized World forces people to live faster compared to previous times. The need which is caused to satisfy this pace incurs some industries a huge responsibility. Aviation industry is one of these main industries. Therefore, there is an enormous cash flow to the industry. As a result of this massive influx of cash even a minor cost reduction in aviation makes a great difference in Global perspective. This expected growth of aviation industry in Globe, force researchers to design improved Air Traffic Management (ATM) technologies and procedures. In order to increase capacity of the system and its efficiency while reducing environmental impact and maintaining safety regulations. To overcome these challenges, new automation and information systems are being developed and this new global operational concept is called Trajectory Based Operations (TBO). In aviation, there are some associations such as, EUROCONTROL, IATA, ICAO, ATCA, which focus on Air Traffic Management (ATM). These associations have partnership with operating airlines and aircraft manufacturers such as, Boeing, Airbus, Cessna on projects to reduce fuel consumption. These partnership relations provide Aircraft Performance Model (APM) documents like BADA to researchers which is an input of TBOs. The BADA APM is designed to simulate and predict the trajectories of aircraft in order to ATM research and operations. The parameters of aircraft performance and trajectories can be calculated using information and data contained in BADA. In order to simulate reality for both operational and research modeling each and every real aircraft needs a corresponding aircraft model and due to experimental study BADA also provides models. Nowadays in every single moment of 24 hours of a day, there are at least 10000 commercial aircraft cruise in Worldwide. This situation causes to a huge fuel cost in aviation industry and every year there are hundreds of studies to reduce this cost in global. Via new techniques in computer science, aircraft can be modeled better compared to past and as a result of these better models we can get more precise results in our studies. Undoubtfully, Earth does not have infinite sources. Therefore, the motivation of this study focuses to determine fuel consumption of airliners more precisely by using artificial neural network techniques due to reduce the cost which is caused by unnecessary additional fuel. In trajectory side both deterministic and probabilistic methods are tried but with the improvements in computation science World turn its face to the machine learning algorithms. The motivation of this thesis is improving baseline estimations of aircraft performance models with data-driven methods. In computer science, python is mostly preferred due to its more understandable nature compared to other coding languages. It is also one of the most effective software language for complicated systems such as aircraft. To be clear, modeling and simulating an aircraft is a very complex interdisciplinary profession. Therefore, python is preferred in this study due to its advantages. In this study Keras, Pandas, Numpy, Scipy, Mxnet, Matplotlib, Bokeh, Glob and Sklearn libraries are used. Firstly, formulas and technical parameters in BADA document are coded into computer environment. Quick Access Recorder (QAR) is used to store all information for flight in aircraft. It is also known as black box data. Secondly, for this thesis QAR data for the given aircraft which are B738 AC#1 (219 flights), B738 AC#2 (563 flights). QAR data for all flights are modified due to necessity and merged according to their tail number and separated to 3 according to flight phases such as, climb, cruise and descent. After that, these 3 modified and merged csv files are converted to binary to be used in machine learning. Secondly, artificial neural network algorithms such as Adam algorithm are applied to these binary files. In order to achieve a better learning a lot of method is used in thesis. Black-box model is created in order to control accuracy of data-driven APM. By the results of Black-box model it is easily seen that Black-box model and actual QAR data matches more than %99. After that Throttle/Auto-Throttle model which is the main aim of the thesis study is created. However, standing alone APM model is not sufficient. Thirdly, to obtain more healthy comparison, these trained models integrated to belonged flight plans and fuel results are compared with each other. In this phase, several data process is applied to the flight plan data. The main reason of data process is originated to the sensor characteristics. On the other hand, it is obvious that, following flight plan completely is impossible due to several reasons such as unexpected weather conditions or high traffic density which is managed by air traffic control (ATC). To overcome these situations an implementation algorithm is used in thesis.
Large eddy simulation of secondary flow structures in low-pressure turbine blade
Increasing the efficiency of a gas turbine engine and reducing fuel consumption depend strongly on the efficiency of gas turbine components. The low pressure turbine (LPT), which is one of the heaviest components in a gas turbine engine, is subjected to very high temperatures and complex flow fields. The efficiency of a turbine is crucial in an engine since the power is generated by the turbine. Turbine efficiency can be increased by reducing the aerodynamic losses. Aerodynamic losses in turbine stages can be classified as profile losses, tip leakage losses, and secondary flow losses. The secondary flow losses are responsible for a significant amount (30% - 50%) of losses in the turbine. Hence, in this study, secondary flow structures, their interactions with each other, reasons for the formation of these structures and effects of the inflow boundary condition on these structures are studied. Simulations are performed with pisoFoam (pressure implicit with splitting of operators) solver in OpenFOAM (Open Field Operation and Manipulation) software which is an open source computational fluid dynamics (CFD) solver. Large eddy simulation (LES) method is used to investigate the unsteady, turbulent flow in a turbine passage. The untwisted T106 turbine blade geometry is used in the analysis. Half span is used for all simulations thanks to symmetry plane boundary condition. The computational domain split into 29 blocks in order to generate structural hexahedral elements. Three different grids are generated in order to study the grid resolution effect on statistics and to identify most appropriate grid. The coarse grid has 3 million grid points, medium grid has 15 million grid points, and fine grid has 30 million grid points. First simulation is performed with the coarse grid in order to validate and investigate the capability of the LES solver in OpenFOAM software. This case has uniform inlet and an inflow angle of 45 degree. After the numeric validation, an inflow angle of 37.7 degree is used for the rest of the simulations presented in this study. Then, the coarse mesh with an endwall is used to investigate the effect of the subgrid model on flow field. In these simulations, the effects of Smagorinsky model and k-equation model on turbulence and secondary flow structures are compared. Results indicate that k-equation model gives more accurate representation of the flow field than the Smagorinsky model. As a result of that, rest of the simulations are performed with k-equation model. After that, the effect of grid resolution (medium and fine grids) on the flow field is investigated. In these simulations, analysis of the energy density spectrum and the instantaneous second invariant velocity gradient tensor show that medium grid is fine enough to capture the relevant secondary flow structures. Hence, other studies on the effects of the inflow boundary condition on secondary flow structures and losses are performed with the medium grid. In order to investigate effects of the inlet boundary conditions on the secondary flow structures and losses, four different inlet boundary conditions are applied. These are uniform inlet, laminar boundary layer (LBL) velocity profile inlet with a shape factor of 2.2, turbulent boundary layer (TBL) velocity profile inlet with a shape factor of 1.5, and unsteady wake with secondary flows (UWwSF) inlet boundary condition, which is obtained from the wake results of a simulation with TBL inlet. Results indicate that inlet boundary conditions cause significant changes on the horseshoe vortex system, the strength of secondary flow structures, and total pressure losses along the flow field.
Design, modelling and control of a reaction wheel actuated system
The stability and control of spacecrafts has been an significant topic, since the beginning of the space age. The spacecrafts (S/C) can be required to be stable in some missions depending on the requirements. The S/C can also be required to rotate with a given speed, to observe a point in space, to keep its orientation. In these cases, depending on the orbit of the S/C, the attitude of the S/C should be changed with an active control method. The attitude control of a S/C can be realized with active and/or passive methods. Spin stabilization is a passive method and used for the control of some S/C in the beginning of the space age. This method utilizes the gyroscopic rigidity of a spinning object to reject the disturbances on the rotating axis. Gravity gradient stabilization is another passive stabilization technique. In this method, the S/C should be longer in one axis comparing it to the other axis. The gradient in the length causes the gradient in the gravity. In this way, the longest axis is aligned with the gravity vector and stability can be maintained through the orbit. However the S/C can rotate freely about the gravity axis, due to no control effort on that axis. Therefore it is not suitable for precise missions. In active methods, the attitude of the S/C can be changed with actuators. Magnetic torquers are used to exert a moment on the S/C by applying a current on an inductor. They can be used to dump the excess momentum of the momentum exchange devices. Thrusters can be used in the orbit transfers due to their high torque/force output. They can be used to apply high angle maneuvers on the S/C, also the momentum dumping operations can be done using thrusters. Momentum exhange devices uses the reaction torque generated when the speed or the rotating axis of a high inertia flywheel is changed. Reaction wheel (RW) is composed of a flywheel, an electric motor to drive it and electronic components to change the voltage/current applied on the motor in order to change the speed of the flywheel. Control moment gyroscopes (CMG) can be considered as a RW attached to a gimbal, and on that gimbal axis there is a high torque capable electric motor. The rotation axis can be changed by this gimbal and high control moments comparing to the RW, can be achieved by CMGs. RWs and CMGs are used in high precise missions. CMGs are more complex than the RWs and they are much more expensive. Depending on the mission requirements an active or a passive method and proper actuators can be chosen. In this thesis, reaction wheels, their electro-mechanical structure is researched. A literature survey is conducted on RWs and the control applications with RWs. In the literature, it is observed that most of the research is done on computer simulation environment. It is important to test the systems both mechanical and electronic, on a realistic test environment before real operations. On a test system, the sensors and actuators can be tested and a possible failure can be prevented beforehand. In this research, for the purposes of modelling an control, a RW and a RW actuated test system is designed on computer aided drafting (CAD) programme and then manufactured. As a result, it become possible to observe the behavior of this dynamic system. A mathematical model is developed and validated with the experimental data. Then, the desinged controller is tested on the platform. In the first section, the current research in the literature is presented. The purpose of the thesis is explained. In the second section, the design process of RW and the test platform is explained with the CAD drawings. The testbed can be rotated in two axis, pitch and yaw. On each axis, there is a RW to control the orientation. There are encoders on each axis to measure the angles. The measured angles are processed with a controller and the control signal is applied on the RWs. The processes of measurement, calculation, data recording, controller implementation are done on a micro-controller. In the third section, the mathematical model for a RW and the testbed is obtained. RW is modelled using DC motor model. The testbed is modelled with the Newton's second law, the friction is taken as viscous friction and embedded in the equation of motion. During tests, it is observed that the simulation results and the real data does not match. An update for the friction model is required. Literature is surveyed on the topic of friction. It is decided that the Stribeck friction model best represent the nonlinear behavior of the testbed. In the fourth section, controllers are designed. The first one is an open loop control law to determine the time when to increase speed of the flywheel and when to decrease it. In order to check if the desired angle is reached a feedback controller is needed. The second controller is A PD based controller designed using parameter space approach. In the last section, the contribution of the thesis is explained briefly and future work is discussed.
A parallel monolithic approach for the numerical simulation of fluid-structure interaction problems
In this study, a novel numerical algorithm has been developed for the simulation of fluid-structure interaction problems in a parallel fully coupled solution approach. For the fluid part of the problem, the Arbitrary Lagrangian-Eulerian (ALE) form of the incompressible, unsteady Navier-Stokes equations are used to model the fluid motion. The ALE based governing equations of the fluid domain are discretized using an unstructured finite volume approach, where the primitive variables of the fluid are defined based on a side-centered arrangement. In this side-centered arrangement, the velocity vector components are defined at the mid-point of each cell face, while the pressure is defined at the element centroid. This arrangement of the primitive variables leads to a stable numerical scheme and it does not require any \textit{ad-hoc} modifications in order to enhance the pressure-velocity coupling. The most appealing feature of the present finite volume approach is that it leads to the classical seven-point Laplace operator on uniform Cartesian meshes for the pressure Poisson equation as in the classical MAC scheme, which is very important for the efficient solution of the large-scale FSI problems in 3D, since the fluid subproblem requires the majority of the computational resources in the coupled system for the most of cases. Another point to be considered in terms of accuracy and stability of the ALE based finite volume approach is the mesh movement strategy. In the ALE method, the mesh follows the interface between the fluid and solid boundary and the governing equations are discretized on a moving mesh, which requires the imposition of special conditions on the mesh movement for the accuracy and stability of the time integration scheme. This condition is imposed by the enforcement of the discrete geometric conservation law (DGCL). In the present study, the fluxes due to mesh motion are computed in a way that the geometric conservation law is satisfied in a discrete level. A special attention is also given in order to satisfy the continuity equation exactly within each element. The summation of the discrete continuity equations can be exactly reduced to the domain boundary, which is important for the global mass conservation. For time integration in the fluid domain second-order backward difference formula is used. The mesh deformation algorithm within the fluid domain is based on an efficient algebraic method, where the interior fluid nodes are deformed based on a exponential function of the displacement of the nearest node on the common fluid-structure interface. The main advantage of the present algebraic method is that it leads to a highly sparse algebraic equation, which is very important for the efficiency of the overall algorithm. The deformation of the solid domain is governed by the constitutive laws for the Saint Venant-Kirchhoff material model, which is applicable to geometrically nonlinear analysis of the structure with large elastic displacements. The discretization of the governing equations of the solid domain is accomplished by the classical Galerkin finite element method in a Lagrangian frame. For three-dimensional finite element discretization, eight-node iso-parametric hexahedral elements are used to model the solid domain, while four-node quadrilateral finite elements are used for two-dimensional discretizations. The numerical integrals related to the structural stiffness matrix, mass matrix and force vector are calculated using two-point Gauss quadrature method. The time integration of equations for the solid domain is based on the generalized$-\alpha$ method, which is a second order accurate, one-step implicit time integration method based on Newmark type approximations for the displacements and velocities of the structure. In this method, the high frequency damping character of the numerical algorithm can be easily controlled by selecting the appropriate generalized$-\alpha$ parameters. The linearization of the structural dynamics equations is based on a quasi-Newton type algorithm, where the Jacobian matrix of the equation system is not exactly calculated at each time level. This approach is also applied for the ALE algorithm for the fluid domain, where the exact Newtons' method leads to extra non-zero blocks in the whole coupled system of equations, which significantly increases the memory demand in three-dimensional computations. The solution of the equations resulting from the discreatization of the fluid and structure domains is based on a monolithic approach, where the fluid and solid equations are assembled into a single equation system and solved in a fully coupled way at each time level. The coupling between the fluid and structure domains is simplified by creating fluid and structure meshes which are conformal along the fluid-structure interface. The present monolithic FSI solver uses a preconditioned Krylov subspace method to solve the resulting fully coupled system. Because of the zero-block diagonal resulting from the divergence-free constraint, it is rather difficult to construct robust preconditioners for the whole coupled system. In the present approach, we use an upper triangular right preconditioner which results in a scaled discrete Laplacian instead of the zero block in the original system. Since this preconditioning leads to a significant increase in the number of non-zero elements due to the matrix-matrix multiplications, we replace the non-zero block of the preconditioner with a computationally less expensive matrix, which contains the contributions from the right and left element only, since the largest contribution for the pressure gradients in the momentum equations comes from the right and left elements that share the common face where the components of the velocity vector are discretized. Although, this approximation does not change the convergence rate of the iterative solver significantly, it leads to a significant reduction in the computing time and memory requirement in 3D. The present one-level iterative solution approach is based on the restricted additive Schwarz method with the flexible GMRES(m) Krylov subspace algorithm due to the non-symmetric nature of the system. A block-incomplete LU (ILU) factorization is used within each partitioned sub-domains. In order to handle the nonlinear nature of the algebraic equations, sub-iterations are performed for each time step until a satisfactory convergence criteria is reached. The implementation of the present preconditioned Krylov subspace algorithm, matrix-matrix multiplications and the restricted additive Schwarz preconditioner are carried out using the PETSc (Portable, Extensible Toolkit for Scientific Computation) library developed at the Argonne National Laboratories, which is a collection of data structures and routines for parallel implementation of linear and nonlinear equation solvers. The unstructured computational mesh for the whole fluid and solid domains is decomposed into a set of sub-domains using the METIS library, which is a set of programs developed for partitioning unstructured graphs, meshes, and producing fill reducing orderings for sparse matrices. In order to validate the accuracy of the developed FSI solver and examine the scaling properties of the proposed algorithm, the present method is initially applied to several FSI benchmark problems that are frequently addressed in the literature. The first validation case is a rather popular two-dimensional FSI test problem. The problem consists of a Newtonian fluid flow interacting with an elastic bar behind a fixed rigid circular obstacle which is placed asymmetrically between parallel lateral walls. In this fluid-structure interaction scenario, shedding vortices induce time dependent displacements and periodic oscillations of the structure. Both steady and unsteady flow solutions are considered for this test case. In order to demonstrate the mesh convergence and scaling properties of the present FSI algorithm, three different mesh resolutions are considered. The effects of the mesh resolution, the number of processors, the level of ILU(k) preconditioner and the amount of overlap for the restricted additive Schwarz preconditioner are presented. The obtained results for this two-dimensional FSI problem are compared to several solutions from the literature and the accuracy of the present algorithm is validated. In the second FSI benchmark case a three-dimensional problem configuration is considered, which is also a well addressed FSI problem in the literature. This problem setting simply simulates the blood flow through elastic arteries. The test configuration consists of an incompressible viscous flow through a flexible circular tube. The considered boundary conditions for this FSI problem produces a wave propagation through the elastic tube with resultant time dependent radial and axial deformations of the structure. A scaling test is also conducted for this 3D benchmark problem for two different mesh resolutions. It is shown that the computed results for radial displacements of the structure are relatively in good agrement with the results in the literature. In the subsequent validation case, a steady FSI benchmark is considered, where an elastic solid is immersed in a rectangular channel, and the deflection of a test point on the elastic structure is taken in account for comparison to the result given in the literature. The computed results for this FSI test case are also found to be in good agrement with the literature. The fourth test case represents a three-dimensional configuration for the vortex-induced vibrations of a flag attached behind a rectangular rigid body in an incompressible viscous flow. The present calculations for this test case are performed on a relatively fine mesh. Although this test case is rather demanding in terms of the required computer power, the FSI algorithm achieves similar scaling without a significant performance loss. The final benchmark problem is motivated to demonstrate the exact mass conservation property of the present FSI algorithm when the fluid domain is fully enclosed in a solid domain. For this purpose, a two dimensional flexible circular ring is placed symmetrically between two parallel rigid plates. This circular ring, which is immersed in a fluid domain, completely encloses a part of the fluid domain itself. This configuration mimics the deformation of a red blood cell inside a capillary wall at a very low Reynolds number. This final test demonstrates that the developed algorithm with its compatible FSI interface condition achieves mass conservation at machine precision. Subsequent to the numerical experiments and FSI solver validation, the present algorithm is applied to a realistic fluid-structure interaction problem frequently encountered in cardiovascular FSI. This fluid-structure interaction problem corresponds to an impulsively accelerated blood flow within a cerebral artery with aneurysm located at the apex of the bifurcation of a branch. The blood is assumed to be a Newtonian fluid, and the artery wall is modeled as a Saint Venant Kirchhoff material. The octree method is used to generate the initial all hexahedral conforming coarse mesh. Various hemodynamic quantities of interest like fluid velocities, blood pressure and wall shear stresses (WSS) are computed as well as the time dependent artery wall deformations. Good scaling character has also been obtained for this FSI application. Finally, the methods used to develop and test the present FSI solver are summarized. Advantages and the drawbacks of the present solver are addressed with possible future applications.
Investigation and development of nontraditional approach for attitude estimation of small satellites
IGRF (International Geomagnetic Reference Field) and Tsyganenko's T89 magnetic field models are the selected models for the comparative study. These models can be one of the input vector separately besides the Sun direction to Singular Value Decomposition (SVD) method. Extraterrestrial effects coming from the Sun and/or interplanetary rays are the main consideration for estimating the magnetic field components close to the actual results. T89 model that uses an index changing with the external effects needs a model (IGRF) simulating the terrestrial magnetic field changes. Proposed models give the results in an acceptable limit with respect to the measurement data. Moreover, T89 model is the best-fit model if the absolute errors for the attitude angles are calculated. The difference between the defined models can be ignored with decreasing altitude but also become important on the higher orbits. This research also describes the development of nontraditional attitude determination system that can rely on sensor measurements which are magnetometer, sun sensor and rate gyros. Vectors coming from the selected sensor data and developed models can be placed in Wahba's problem. The system uses SVD method to minimize the Wahba's loss function and determine the attitude of the satellite. In order to obtain the attitude of the satellite with desired accuracy an extended Kalman filter (EKF) for satellite's angular motion parameter estimation is designed. The SVD and EKF algorithms are integrated for estimation of the attitude angles and angular velocities. The nontraditional method is compared with the traditional approach which uses the same sensors as magnetometer and sun sensor in addition to the gyros. In the eclipse period, algorithm is switched into the prediction phase for better attitude estimation of the satellite; therefore, satellite's attitude angles when no data coming from the sun sensor are improved. Magnetometers are assumed to be calibrated before the nontraditional technique is applied on the satellite. All the necessary algorithms are designed in the MATLAB package program and the simulations are performed using that software. The test platform has been used for the verification of the algorithms presented. The algorithms are implemented on the test platform using the magnetic field sensor data in the single-frame method.
Uydu ile hassas görüntülemede iz düşüm yönlendirme metodu
Gerek askeri gerekse sivil alanlarda görüntü ihtiyacı için, tek seferde büyük alanları kapsayabilmesi ve küresel olarak tüm coğrafi bölgelere ulaşım imkanları sebebiyle yer gözlem uydularına olan rağbet gittikçe artmaktadır. Bununla birlikte, gelişen teknoloji ile elektro-optik görev yükü ekipmanının içindeki aynalar büyümekte ve algılayıcı üzerindeki piksel boyları küçülmekte olup, böylelikle daha yüksek çözünürlüklere ve kaliteye ulaşılması mümkün olmaktadır. Günümüzde metre altı seviyelerde yer örnekleme mesafesine sahip elektro-optik yer gözlem uydularının sayısı gittikçe artmaktadır. Ancak, görüntüyü algılayacak kamera ekipmanının (görev yükü) yüksek çözünürlükte kaliteli görüntü algılama kabiliyeti olsa da, görev yükünün işlevini kalite, çözünürlük ve doğru hedefin görüntülenebilmesi bakımından tam olarak gerçekleştirebilmesi için uydunun da yeterli doğrulukta konum ve yönelim profiline sahip olması gerekmektedir. Push-broom tipinde elektro-optik görev yüküne sahip bir uydunun görüntüyü algılaması için görüş eksenini belirli bir sabit hızda yeryüzünde hareket ettirmesi gerekmektedir. Uydunun bunu sağlayabilmesi için görüntüleme aktivitesinin başında doğru konumda ve doğru yönelimde (ve yönelim hızında) olması gerekmektedir. Bununla birlikte, görüntüleme aktivitesi sırasında görüş ekseninin yeryüzünde sabit hızda hareket edebilmesi için referans yönelim profilini (örneğin yer istasyonundan sağlanan) yeterli doğrulukta gerçekleyebilmesi gerekmektedir. Bu çalışmada, push-broom tipinde görev yüküne sahip yüksek çözünürlüklü yer gözlem uydularının görüntüleme aktiviteleri sırasında takip etmesi gereken yönelim profilini oluşturmada kullanılabilecek bir yöntem olan ve iz düşüm yönlendirme metodu olarak adlandırılan yöntem sunulmaktadır. Çalışmada, tez kapsamında geliştirilen ve yöntemin uygulandığı iz sürücü algoritma tanıtılmaktadır. Bu yöntem ile hesaplanan yönelim profilinin çıktısı polinom katsayılarıdır. Polinom katsayılarının kullanılmasındaki sebepler, uydu bilgisayarında istenen zaman adımı (örneğin kontrol frekansında) ile çözülmesinin kısa zamanda oluşu, yer istasyonundan gönderilen tele-komuttaki parametreleri azaltması ve analitik türevinin alınmasındaki basitliktir. Görüntüleme aktivitelerinin ve hata analizlerinin gerçekleştirilebilmesi için gerekli tüm girdiler, yine tez kapsamında geliştirilen algoritmalar kullanılarak elde edilmiştir. Üst seviyede bu algoritmalar yörünge ilerleticiyi, randevu algoritmasını (görüntüleme olanaklarının hesaplanması), sayısal yükseklik modelinden yükseklik değerlerinin elde edilmesini, yeryüzünde yer izinin ilerletilmesini, görselleme araçlarını, test & hata analizi fonksiyonlarını ve uygulamaya yönelik analizlerin gerçekleştirilmesinde kullanılan fonksiyon ve arayüzleri kapsamaktadır. Yeryüzündeki görüntünün geometrik özelliklerini / gereksinimlerini ve push-broom görüntüleme tekniğinin temel prensiplerini kullanan ve tez çalışması kapsamında geliştirilen iz düşüm yönlendirme metodunun çıktısı, belirli bir görüntüleme aktivitesine ilişkin uydunun takip etmesi gereken yönelim profilidir (polinom formatında). Ancak, polinomun mertebesine bağlı olarak metodun kendisi de bir hata kaynağıdır. Tez çalışması kapsamında polinom mertebesine bağlı olarak görüntünün geometrik özelliklerinde ve yönelim profilinde oluşabilecek göreceli hatalar farklı entegrasyon sürelerine sahip görev yüklerini de içine katacak şekilde analiz edilmiştir. İz sürücü algoritmanın girdilerinde yapılabilecek hatalar da elde edilecek görüntünün özellikleri için birer hata kaynağıdır. Örneğin görüntülenecek bölgedeki yeryüzü yüksekliğinin hatalı olarak belirlenmesi ve/veya uydunun yörüngedeki konumunun hatalı olarak bilinmesi, üretilen yönelim profiline ve dolayısıyla görüntünün özelliklerine etkiyecektir. Bu hatalara ilişkin analizler de yine tez çalışması kapsamında gerçekleştirilmiş olup, sonuçları farklı özellikte yörüngeye veya görev yüküne sahip uydular için incelenmiştir. Referans olarak üretilen yönelim profilinde hiç hata olmasa dahi, uydunun bu referansı takip etmesi (gerçeklemesi) sırasında da hatalar meydana gelebilir. Bu hatalar, yönelim hassasiyeti ve kararlılığı kapsamında incelenmiştir (kararlılık analizlerinde katı cisim modeli kullanılmış olup yüksek frekanslı jitter etkileri incelenmemiştir). Görüntünün geometrik özellikleri ile yönelim hassasiyeti ve kararlılığı arasında ilişki kurulmuş ve Monte Carlo analizleri yardımı ile oluşabilecek hatalar incelenmiştir. Bununla birlikte, görev yüküne ve yörüngeye bağlı kararlılık gereksinimi hesaplanmasında kullanılabilecek bir basitleştirilmiş model elde edilmiştir. İz sürücü algoritma kullanılarak coğrafi olarak her bölgede görüntülemenin gerçekleştirilmesi için gerekli yönelim profili elde edilebilmektedir. Ancak elbette bu profil uydu tarafından gerçekleştirilebilse dahi, bu durum tüm coğrafi bölgelerden kaliteli görüntü alınması için yeterli olmayabilir. Elektro-optik görüntülemede görüntülenecek bölgeye yeterli seviyede Güneş ışını düşmelidir ki görev yükü bu ışığı toplayabilsin. Yörüngenin yükseliş düğümü saat açısına ve mevsime göre görüntülenebilecek coğrafi bölgeler (görev yükü kabiliyeti ile ilişkilendirilerek), Güneş başucu açısına bağlı olarak irdelenmiş ve grafiksel olarak sunulmuştur. Son olarak, görev yükü kabiliyetlerine ve yörünge yüksekliğine bağlı olarak tek sefer algılanabilecek en uzun şerit görüntüleme aktiviteleri analiz edilmiştir. Bu kapsamda gerçekleştirilen analizlerde uydunun hafıza, güç ve ısıl kısıtları kullanılmamış, sadece yer örnekleme mesafesi kısıtlarına göre algılanabilecek mesafeler irdelenmiştir. Bununla birlikte, push-broom tekniği ile uzun şerit görüntüleme aktivitelerinin bir problemi olan; yer örnekleme mesafesinin görüntü boyunca değişken olması durumuna çözüm olarak bir yöntem sunulmuştur. Bu yöntemde, yer örnekleme mesafesinden bir miktar ödün verilerek, görüntü içinde sabite yakın yer örnekleme mesafesi elde edilebilmekte ve ayrıca bu şekilde algılanabilecek görüntü uzunluğunda Güneş'in aydınlatma etkisi ve uydunun hafıza, güç ve ısıl kabiliyetleri haricinde bir kısıt oluşmamaktadır. Not edilmelidir ki, bu tip yöntemlerin uydu üreten firmalarca uygulandığı bilinmekle birlikte, nasıl yapıldığı bilinmemektedir. Tez kapsamında sunulan bu çalışma ise özgün bir çalışmadır. Sonuç olarak, tez çalışması kapsamında geliştirilen iz düşüm yönlendirme metodu, girdilerindeki hataların, metodun getirdiği hataların ve çıktılarının uygulanması sırasında oluşabilecek hataların etkileri incelenerek analiz edilmiştir. Tez çalışması kapsamında gerçekleştirilen hemen hemen tüm hata analizlerinde ve uygulamaya yönelik analizlerde, detaylı analizlerde elde edilen sonuçlara yakın değerlerin elde edilebildiği basitleştirilmiş modeller de oluşturulmuştur. Oluşturulan basit modeller kullanılarak, farklı yörüngelerde ve farklı kabiliyette görev yüküne sahip uydular için sonuçlar irdelenmiştir.
Takviyeli kompozit bir plakanın dinamik analizi
Takviyeli kompozit plakalar, yüksek mukavemet ve göreli hafiflik özellikleri sebebiyle kullanım alanı gün geçtikçe genişleyen yapılardır. Kompozit malzeme, tanım gereği farklı özellikteki malzemeleri, istenilen mukavemet özelliklerinde oluşturmaktır. Örneğin, katmanlı kompozit plakalar için katmanların dizilişi plakanın statik ve dinamik karakteristiğini etkileyen en önemli parametredir. Bu parametreye dikkat edilerek aynı ağırlıkta olmasına rağmen statik ve dinamik olarak çok daha mukavemetli yapılar oluşturulabilir. Bu çalışmada, katmanlı kompozitler için, farklı boyutlarda ve dizilişlerde takviyesiz ve takviyeli plakalar oluşturularak statik ve dinamik analizleri ANSYS 16.2 paket programı kullanılarak gerçekleştirildi. Kalınlığı kenar uzunluklarına göre oldukça küçük olan ince plakalar kullanıldığı için analitik çözümler kayma deformasyonunu ihmal edecek şekilde klasik laminasyon teorisi kullanılarak elde edildi. Tasarım değişkenlerinin numunelerin çökme, gerilme ve doğal frekans değerlerine etkisi incelendi. Bu kapsamda çalışma takviyeli ve takviyesiz olarak iki bölümde gerçekleştirildi. İlk bölümde takviyesiz plakalar için incelemeler yapıldı. Kare bir plaka oluşturularak farklı sınır şartlarında, 15'er derecelik dizilim değişimiyle, aynı yayılı yük altında çökme ve gerilme değerlerinin değişimi incelendi. Serbest titreşim analizi yapılarak açı değişiminin mod şekilleri ve frekanslarına etkisi gözlemlendi. Ardından açıklık oranı azaltılarak bahsedilen parametrelerin farklı açıklık oranlarında değişimi incelendi. ANSYS kullanılarak gerçekleştirilen bu analizler, literatürden edinilen analitik çözümlerle kıyaslandı. İkinci bölümde ise takviyeli kompozit plakalar için incelemeler yapıldı. ½ açıklık oranında katmanlı kompozit plaka oluşturuldu. Ankastre sınır şartlarında bu plakanın ilk olarak takviyesiz analizi gerçekleştirildi. Ardından kompozit katmanlar ile oluşturulan takviye eklenerek takviye kullanımının çökme, gerilme ve doğal frekans değerlerinde meydana getirdiği değişim incelendi. Takviyedeki katman diziliş değişiminin etkisi hesaplanarak optimum diziliş elde edildi. Aynı kütlede takviyesiz ve takviyeli plakalar oluşturularak, takviyenin etkinliği hakkında yorumlar yapıldı. Farklı kesit alanına sahip, aynı kütlede ve yükseklikte takviyeler kullanılarak, kesit alanının analiz sonuçlarına etkisi gözlemlendi. Hacim ve kesit alanı sabit tutularak metal malzemelerden oluşturulmuş takviyeler kullanıldı. Ağırlık göz önünde bulundurularak, takviye için malzeme seçiminin çökme, gerilme ve doğal frekans değerlerine etkisi incelendi. Son olaraksa takviyenin doğrultusu değiştirilerek aynı ağırlıkta çift takviye kullanımının ortaya çıkardığı değişim yorumlandı.
Experimental investigation of an oscillating tandem-wing power generator
With the rapid increase in energy demand and along with considering the environmental and sustainable aspects, clean and renewable energy is a necessity for human beings. Considering the potential possibility of energy production inspired by nature, the concept of a flapping wing power generator is investigated. In this study the feasibility and advantages of using flapping wings to generate energy are studied and documented in the literature by researchers. However, previous studies lack some aspects to develop a high performing flapping wing power generator. In the scope of this study various approaches are investigated in order to define the parameters for a flapping wing power generator for optimum performance. Firstly three dimensionality effects are studied for a single flapping wing. Secondly, flapping wing power generators operating inside solid side walls are investigated for 2D and 3D cases. And lastly, an oscillating tandem wing power generator is studied with different phase angles between fore and hindwings. In the experimental study, the flow structures and the forces acting on a finite flat plate that performs sinusoidal and non-sinusoidal pitching and plunging motions are investigated via detailed quantitative flow visualization technique with simultaneous direct force measurements. Reynolds number, pitch pivot position, pitch and plunge amplitudes are kept constant for all experiments. Two different phase angles of = 90° and 110° between pitch and plunge are studied for single free and constrained wings. However, only = 90° is considered during the tandem wing cases because this phase angle results in higher performance compared with 110°. Five different motion types from rapid stroke reversal (TR = 0.1) to sinusoidal reversal (TR = 0.5) are chosen as the flapping motion kinematics. Experiments are conducted in the large scale water channel located at the Trisonic laboratory in İTU. Oscillation of the airfoil is provided by servo motors controlled by computers. All particle image velocimetry (PIV) data are acquired via 2 CCD cameras by illumination of the examined flow plane with a Nd: Yag Laser. Force and torque data are acquired simultaneously by using a six axis water resistant sensor rotating with the wing. Some major results could be concluded from the experiments are described briefly in this section. Early shedding of higher strength vortices occur when the pitch reversal rate is increased. In 3D case leading edge vortex (LEV) and tip vortex interactions are observed near the tip of the finite flat plate where vortical structures shrink in size and stretch out to the wake losing their strength to roll up. As a result, three dimensionality reduces the overall efficiency as compared with 2D cases. For rapid reversals (TR = 0.1), none of the test cases were capable to produce a positive mean power coefficient. Performance of the flapping foil power generator is highly influenced by side-walls. Constrained flow has smaller effect on non-sinusoidal flapping as compared to sinusoidal flapping. The sinusoidal flapping flat plate can achieve higher increase in efficiency in constrained flow compared to the free flow. However, if the minimum distance between the trailing edge and the side wall is too small results in degraded performance on mean power generation. The results imply that there's an optimum wall distance using side-walls in flapping foil power generation that might be a way to compensate for the loss from stroke reversals at maximum and minimum plunge positions predominantly for sinusoidal flapping. In tandem configurations, the power extraction performance of the hindwing are greatly affected by its interaction with the wake of the forewing. Adjusting the phase angle between the two wings can change the mode of vortex interaction. Formation, growth and shedding process of hindwing LEV could be distributed by the shed vortex of the forewing. The timing of interaction and direction of resulted force could be controlled effectively via phase angle. As a consequence the performance loss in hindwing for tandem configuration could be reduced to minimum level and thus energy production can be maximized. Some of the cases studied exhibit constructive vortex interaction that reinforce the LEV and results in higher power production.
Modelling, processing and mechanical properties of aspartic polyurea
Nowadays, polymers are used widely in almost every up-to-date industry. They show high levels of strength-to-weight ratios. Polyurea is a kind of polymer that made of reaction between pre-polymers and specific hardener (diisocyanate). Today's industries are particularly interested in replacing components of their vehicles and parts with reinforced polymers like polyurea and polyurethan. Polyurea has attracted both academic and industrial interests because of capacity to create a dramatic enhancement in properties at various of compounds of diisocyanate and polyamines. Generally, in current study we tried to derive mechanical properties of aspartic polyurea and then attempted to build a proper mechanical model for fabricated material properties. Polyurea has many applications such as coating. For instance, at sport saloons the floor coating must have maximum stiffness beside flexibility. Somewhere else may be there is a need to have a good strength beside sufficient softness against impact loads. During this research the major problem was the creation of gas foams. By studying the stoichiometry and chemometric observations of the reaction it was obtained that these foams were made of carbon dioxide. There were many methods to omit these foams such as using anti-foam additives, carbon dioxide absorber additives, decreasing viscosity and using bubble evacuation system (vacuum). According to our available facilities and tools, we were just able to use anti-foam additives and unfortunately we were unable to omit this malfunction completely, but by controlling process conditions we reduced the amount of bubbles to the lower level. We performed two kinds of experimental test on various polyurea samples. Also different fabrication methods and composition ratio have been used to fabricate modified aspartic polyurea. Uniaxial tension and uniaxial compression tests were performed to characterize the behavior of the materials. In order to choose the appropriate manufacturing protocol and the suitable enhancement for industrial usage a parametric analysis utilized to choose best material in aspect of strength and mechanical properties. Then the extracted data were imported into ABAQUS to capture a suitable hyperelastic model. The evaluated model gave us the ability of future and complex analysis of polyurea in the several mechanical conditions. We were able to produce different kind of polyurea by interfering and changing initial chemical and manufacturing parameters. We aimed on exchanging three parameters of polyurea. These parameters were composition, polymerization mixing speed and polymerization temperature. We attempted to produce different mechanical behavior and measured these parameters to select and compare them with their manufacturing parameters. The best ones contained suitable property of flexibility, strength and stiffness. To achieve these features, after manufacturing several unique samples, the standard mechanical tests carried out. After analyzing the experimental procedures in the result chapter the engineering stress vs, engineering strain curves represented for each sample. By performing data curve fitting process the material coefficients extracted. These coefficients can be recorded as material properties and they would be used instead of raw experimental data. For instance, to analyze a composite plate made of two layers of steel and polyurea there won't be need to insert uniaxial test data to our finite element analyzer machine and these elicited material coefficients would be adequate. As a discussion, we debated about selecting best manufacturing parameters among outputs. By comparing differences between each sample we chose the appropriate manufacturing protocol. It was resulted that polyurea was mechanically strong, specifically in terms of strength to weight. The results showed that the materials have interesting behavior when they uniaxialy compressed, in other word polyurea had very high compressive strength and after unloading the material showed its viscoelastic nature. Polyurea was able to support million times its own weight. we concluded that the polyurea can be utilized to applications that requires high compressive strength and high strain values. The results expressed that the Yeoh and Ogden hyperelastic models can predict the elastic mechanical behaviour of polyurea samples even under large strain compression. Also it is notable that the Observations through the tests showed that the polyurea was strong and useful for static load aspects, also whole tests carried out in room temperature and constant strain rates, so then for those applications which are under dynamic loads or in intense load conditions (impact, explosion and thermal stress-strain problems), extended tests are required. In other words, dynamic mechanical analysis should take place for understanding strength and strain at various strain levels and temperatures. In the next studies the viscoelastic behaviors also should be considered and relative parameters should be obtained then joint with available hyperelastic parameters in this study to build up a very constitutive hyper-viscoelastic model.
A design of a test bed for cubesat attitude determination and control system
Owing to the recent developments in miniaturization and integration technologies, CubeSat's can handle more complex practical missions. Such missions require 3-axis control of the satellite along with a miniaturized 3-axis attitude determination and control systems (ADCS). With the QB0 project, the CubeSat's developed in ITU-SSDTL have also started to use such ADCS systems. Whether they are developed or procured, an ADCS system usually requires a suitable test bed to test the behavior and the performance of it. An ADCS system used for LEO missions usually operates within the Earth's magnetic environments. Therefore magnetic field sensors such as magnetometers are employed to measure the mediums' magnetic field or the change in the magnetic field to determine the orientation and motion of a satellite. Magnetic actuators such as magneto torquers are used to align with the magnetic field of the earth or to damp the tumbling motion of a satellite resulting from unbalanced torque distribution on it. A major goal of the SSDTL is to have such a test bed available in the lab. In addition to related software, the major components are a Helmholtz cage and a suitable air bearing table. With this in mind, the purpose of the present thesis is to aid the development of such an ADCS test bed designing first a suitable Helmholtz Cage system. The sizing of the test system depends on requirements such as the maximum mass of the satellite to be tested, coordinates of center of mass and disturbance levels to be counteracted. In the present thesis, first a comprehensive nanosatellites literature survey is conducted. The success rates and mission failure reasons are also investigated. Based on this preliminary study it is observed that 3U CubeSat's are generally adopted for most missions. Therefore a Helmholtz Cage that will house a 3U CubeSat is considered. A circular Helmholtz cage that will fit a 3U CubeSat is designed and analyzed. 1 axis and 3 axis Helmholtz cage cases are considered. Magnetic field lines present within the cage are demonstrated. Since the size of the Cage considered was just large enough to house 3U the magnetic field lines were not uniform enough. Therefore similar analysis is carried out for larger cage of double the size of the small one. Both are compared for uniformity of the magnetic fields. Then a small square cage is designed and analyzed. Again, 1 axis and 3 axis Helmholtz cage cases are considered. Magnetic field lines present within the cage are demonstrated. The large square cage was found to be the best choice for testing a 3U CubeSat. As a result, a square Helmholtz cage which can be used to test a 3U CubeSat at LEO from 250 km to 1500 km.
Anlık basınç yüküne maruz ince cidarlı takviyeli sandviç kompozit plağın dinamik davranışı
Bu çalışmada, anlık basınç yüküne maruz katmanlı kompozit plağın dinamik davranışı incelenmiştir. Çalışma kapsamında kompozit malzemelerin üretim teknikleri ve malzeme özellikleri araştırılmıştır. Ayrıca sandviç kompozit plak üzerine belli doğrultularda takviye elemanları yerleştirilmiş ve bu elemanların yapının dinamik davranışına etkisi incelenmiştir. Kompozit plağın anlık basınç yüküne vereceği dinamik cevabın belirlenebilmesi için anlık basınç yükü deneyleri gerçekleştirilmiştir. Deneylerde vakum infüzyon yönetimiyle üretilmiş olan karbon epoksi kompozit plak kullanılmıştır. Takviye elemanı olarak ise polipropilen bal peteği çekirdek ve karbon yüzey katmanlarından oluşan sandviç yapının kullanılmasına karar verilmiştir. Takviye elemanındaki karbon yüzey katmanlarının, polipropilen peteğin altına ve üstüne çeşitli sayılarda yerleştirilmesi ile iki farklı tipte takviye elemanı oluşturulmuştur. Takviye elemanlarının üretimi için elle yatırma yöntemi kullanılmıştır. Anlık basınç yükü deneyleri İTÜ Uçak ve Uzay Bilimleri Fakültesi Kompozit Yapı Laboratuvarı'nda yer alan test düzeneği kullanılarak gerçekleştirilmiştir. Test sistemi basınç depolama tankı, kompresör, regülatör, basınca dayanıklı hortum, şok tüpü, veri toplama sistemi, bağlantılar ve güç kaynaklarından oluşmaktadır. Deney kapsamında anlık basınç yükü, şok tüpü içerisindeki basınçlı havanın membranı yırtarak delinmesi sonucunda elde edilmiştir. Oluşan basınç yükü, test plağı üzerinde farklı konumlarda bulunan sensörler yardımıyla ölçülmüş ve basıncın zamanla değişim grafikleri elde edilmiştir. Plak üzerinde farklı konumlara, farklı sayılarda takviye elemanı yapıştırılarak, takviyeli plağın basınç yüklerine cevabı incelenmiştir. Bu deneylerden elde edilen sonuçlar grafik şeklinde sunulmuştur. Daha sonra veri toplama sisteminin filtreleme seçeneği kullanılarak yeni deneyler gerçekleştirilmiştir. Bu deneylerden elde edilen veriler ve uygun bir fonksiyon yaklaşımı kullanılarak anlık basınç yükü dağılımları elde edilmiştir. Deneysel veriler, ANSYS ticari yazılımında gerçekleştirilen sayısal analiz verileriyle karşılaştırılmış ve sonuçlar yorumlanmıştır. Ayrıca takviye elemanının değişik konfigrasyonlarda kullanılması ile oluşabilecek değişik gerinim değerlerinin incelenmesi için ANSYS yazılımında parametrik çalışma da yapılmıştır. İlerleyen çalışmalarda kullanılmak üzere yeni bir basınç fonksiyonu üzerine çalışmalar yapılmıştır. Bu kapsamda rijit bir plak üzerindeki basınç değerlerinin deneysel verilerine Lagrange interpolasyon polinomu uygulanması sonucu yeni bir yaklaşım elde edilmiştir.
Chemical reactor network modelling of a 1000 HP turboshaft engine reverse flow combustor
Turbulence and combustion are two important phenomena's that are very difficult to solve for scientists and engineers. These two phenomena are playing the biggest role in the all fossil-fuelled systems which have been using for over 150 years. By means of the rapid computer technology developments, it has become possible to solve the turbulent combustion flow problems with computational fluid dynamics (CFD) simulations and to predict the physical results such as fluid flow and temperature, as well the major chemical species with high accuracy. However, by using the same method, it is still a difficult and time consuming process to calculate the minor chemical species such as NOx and CO which are released during combustion process. For this reason, the prediction of this pollutants is very important in the design phase of gas turbine engines. On the other hand, due to environmental factors and legal limitations, development efforts to reduce pollutants such as NOx and CO produced by fossil-fuelled systems such as gas turbine engines have become a necessity. The realization of these efforts in conceptual and detailed design phase's before production is very important in terms of reducing the time and costs of the projects. In this thesis study, a new methodology was developed and applied to predict the emissions of a reverse flow combustor of a 1000 horsepower turbo-shaft helicopter engine in the design process. This methodology includes the following respectively: (1) Three dimensional CFD-combustion simulation for 1 / 13th part of the combustion chamber by using FLUENT v.14 software and obtaining temperature values and velocity vectors at each solution point. (2) Converting the three dimensional CFD results and transferring to a two dimensional plane by using two and a half dimension (2.5 D) approach. (3) Developing a MATLAB based tool to create Chemical Reactor Network (CRN) model on two dimensional plane. (4) Solving the CRN model with using n-heptane fuel mechanism by using CHEMKIN PRO software and obtaining emission results. In this methodology, the temperature and mass flow values are obtained by CFD simulations and the CRN model is constructed by using these values. In this way, complex chemical mechanisms can be calculated quickly and small chemical species can be obtained precisely. The temperature, which is the most important parameter affecting the amount of NOx and CO generated as a result of the combustion reaction, is determinative at the creation stage of the CRN model in this study. Most of the previous similar studies were made for combustion chambers in the form of axially symmetric and simple geometric tubes. Therefore, CFD simulations for a two dimensional plane passing through the centre of the injector, and the creation of a chemical reactor network from these simulation results, are become easier than in this type of combustors. In case of the geometry in xxi this study is an annular and non-symmetrical combustion chamber which are consisting of a sequential fuel injector system. Building the reactor network model using the three dimensional CFD results and the methods used in the previous studies is quite difficult, complicated and long process for the geometry in this work. To overcome this challenge, the 2.5 D approach, which is generally using in image processing studies, has been utilized in this thesis. In this method, nine sections were taken at equal intervals over the three dimensional volume, and the average of CFD results in these sections were transferred to a created two dimensional plane and the CRN model was constructed on this plane. A MATLAB-based tool has been developed to determine reactor volumes and mass fluxes between these volumes. With this tool, the chemical reactor network for three different operating conditions of the combustion chamber (idle, cruise, take-off) was modelled. These models were transferred to the CHEMKIN PRO software and detailed chemical reaction analysis was carried out by using reduced n-heptane fuel mechanism which consist of 159 chemical species and 770 reaction steps. In this study, reactors are assumed as ideal and adiabatic.
Fatigue failure analysis of a turbine blade
As a part of aviation, gas turbines need insurances but they actually face to many failures in different stages such as vibration, fatigue, foreign object damage, corrosion, erosion, sulphidation. Turbine blades are the most component that face to failures and affect on the lifetime and the performance of turbines. Several reasons cause failures on blades, several hostile environment factors, the amount of fatigue stress a blade can tolerate, mechanical design problems, temperature and it's cycling, the amplitudes and frequency of vibratory stress, quality control problems within blades or vanes and delivery system problems for the cooling air and turbine section domestic object damage. Although experimental test and analysis of these failures for a real gas turbine in real work condions are very hard and expensive, many research and methods has been improved to analysis fatigure failures and increase turbine life. In This Study, the main goal is analysing HCF and LCF failures in different operate condition for a model based on real designs with Inconel 718 material and investigate failure reasons, to achieve a solution on design of turbine blade with higher life and decreased failure possibility. 3D model of blade and its assembley to disk has been created and using ANSYS software, a finite element model for LCF and HCF analysis prepared. Using this model and discussing on results and comparing to each other, best design parameters would be defined. In General, three loads are acting on turbine blade that cause in high stresses. These loads are hot air pressure, thermal loads caused by hot air, and resulting centrifugal force from rotation of turbine. I get these loads for real gas turbine performance conditions from GasTurb12 software for different conditions. Also to inspect effect of each load on turbine blade life, other loading cases has defined as changing just one load and taking the other two loads constant. Analysis performed in ANSYS and results for these conditions obtained. By studying on results it can be concluded that fir-tree region of blade, which is connection part of blade to disk, is the source of failure caused by high temperature on that part. Blade surface can resist up to 1000C without any fatigue failure, but dovetail of blade can stand up to 500C. In order to improve design life in high loaded conditions with very high temperature, different materials or very efficient cooling system can be used.
Optimization of a centrifugal compressor impeller using genetic algorithm coupled with artificial neural networks
Nowadays, obtaining a state of art design for any kind of turbo machinery is clearly an issue of optimization and efficient methodology, since the difference between a decent and unacceptable design is determined by very small changes to the design. It can be very though or demanding task to fulfill the performance criteria such as range and efficiency with trial and error. Consequently, automatization of the design process redounds significant amount of time and effort alongside ensuring a certain level of confidence. Taking into consideration the aforementioned, a design methodology for the centrifugal compressors has been developed using on the shelf commercial tools and put to use. In the scope of this work, impeller of an experimental compressor of a micro turbojet engine will be utilized as baseline for optimization. Only impeller part of the compressor is the concern for this study, however the same methodology can be applied to any part of a turbomachinery, given the adequate amount of time. Results are compared to demonstrate the power of the methodology over a trial and error design. Baseline impeller has a pressure ratio of 1:4 with approximate isentropic efficiency of %85. Target for the optimization was at least %1 increase in isentropic efficiency, at the design point, without any change to pressure ratio, mass flow. Optimized impeller is going to replace the baseline and must work with the same baseline diffuser. This means that any increase in the performance parameters of the compressor is due to the geometrical changes in impeller only. Geometry parametrization is the first step and performed using Bezier curves. Any other geometrical parameters are frozen throughout this thesis. The range of geometrical parameters is determined mostly by manufacturing reasons or practical purposes. Rest of the process includes setting up a database automatically with the help of ANSYS Workbench, training an artificial neural network which has been presented by MATLAB as a fast performance prediction tool and a genetic algorithm code for creating new individuals (designs) until optimization criteria are satisfied. Thanks to the developed methodology, baseline impeller's total to total isentropic efficiency has been improved by %1.5 at the end of the study without any significant change to the mass flow ratio and pressure ratio. Furthermore, only internal geometry of the impeller (blade angle distribution) has been modified, meaning that optimized impeller can be retrofitted to the baseline compressor without any inconvenience.
Kanat yapısal deformasyonunun rüzgar türbini rotor performansı üzerine etkisi
Bu çalışma kapsamında çeşitli yükler (aerodinamik, merkezkaç, yerçekimi, vb.) altında çalışan rüzgar türbini kanadının davranışı incelenmiştir. Aerodinamik ve yapısal disiplinlerin etkileşimi göz önüne alınarak, kanat tasarımı sırasında ön tasarımdan performans hesabına kadar olan aşamalarda kullanılabilecek metodoloji geliştirilmiştir ve örnek bir kanat üzerinde parametrik çalışmalar yapılmıştır. Pala Eleman Teorisi ve Klasik Laminasyon Teorisi tez çalışması sırasında faydalanılan temel teorilerdir. Her iki yöntemin ve aralarındaki etkileşimin gerçekleştirilmesi için MATLAB ortamında yazılımlar geliştirilmiştir. Geliştirilen yazılımların çıktıları NREL tarafından gerçekleştirilen çalışmalarla karşılaştırılmış ve sonuçların birbirine yakın olduğu görülmüştür. Bu tez çalışması sonucunda istenilen özelliklerde bir kanat tasarımının gerçekleştirilmesi için kanadın yapısal ve aerodinamik davranışının ve aralarındaki ilişkinin anlaşılması için gerekli disiplinler arası çalışmalar gerçekleştirilmiştir. Tezin yukarıda bahsedilen söz konusu farklı disiplinleri kullanarak ve aralarındaki ilişkiyi inceleyerek yüksek performanslı kanat tasarımı için tasarımcılara ve araştırmacılara bir yol gösterici olması hedeflenmiştir. Ön tasarım aşamasında malzeme seçimi, serim açısı, profil seçimi ve boyutlandırma gibi parametrik çalışmaları, bu parametrelerin performans üzerine etkilerinin hesabını hızlıca gerçekleştirebilecek bir altyapı sunulmaktadır. Aerodinamik ve yapısal etkileşimin hesaplanması tezin kritik noktalarından birine işaret etmektedir. Rüzgar türbininin enerji üretimi rotor ve rüzgar arasındaki etkileşime bağlıdır. Rüzgar hızı, bir ortalama hız ve bunun etrafında salınım yapan, türbülans içeren hız değişikliklerinin bir kombinasyonu olarak kabul edilebilir. Görülmüştür ki rüzgar türbini performansının ana parametreleri (ortalama güç üretimi ve ortalama yükler) ortalama hız nedeniyle oluşan aerodinamik kuvvetler tarafından belirlenmektedir. Rüzgar kayması, yan rüzgarlar, rotor dönüşü nedeniyle oluşan periyodik aerodinamik kuvvetler, türbülans ve dinamik etkiler yüzünden ortaya çıkan rastgele salınım gösteren kuvvetler yorulma yüklerinin kaynağını oluşturmaktadır ve rüzgar türbinin maruz kalabileceği en yüksek yüksek yük değerinin belirlenmesi için kullanılır. Önem taşıyan bu faktörlerin anlaşılması için kararlı durumdaki rüzgar türbini aerodinamiğinin anlaşılması gereklidir. Aerodinamik ile ilgili çalışmaların yer aldığı bölümde öncelikle ideal bir rüzgar türbininin analizine yer verilmiştir. Rotorun genel davranışı, etrafındaki akış ve önemli kavramlar hakkında bilgi verilmiştir. Analizler rüzgar türbinleri için teorik performans limitlerinin hesaplanması için de kullanılmıştır. Genel aerodinamik kavramlar ve kesitlerin (airfoil) çalışma ilkeleri anlatılmıştır. Rüzgar türbinin klasik aerodinamik analizi ilk olarak Betz ve Glauert tarafından 1930'lu yıllarda gerçekleştirilmiştir. Daha sonra teori bilgisayarlarda uygulanabilecek şekilde genişletilmiştir. Tüm bu metotlarda, momentum teorisi ve pala eleman teorisi birleştirilerek rotorun dairesel bir bölümünün performans karakteristiklerinin hesaplanmasını sağlayan bir bant teorisine dönüştürülmüştür. Rotorun tamamının performansı ise tüm dairesel bölümler için elde edilen değerler toplanarak (entegre edilerek) bulunur. Öncelikle momentum teorisi ve pala eleman teorisi açıklanmış ve ideal çalışma şartlarında optimum bir kanat tasarımı için kullanılmıştır. Günümüzde kullanılan rüzgar türbini kanatlarının ana hatlarını belirleyen kavramlar gösterilmiştir. Daha sonra iki yaklaşımın bir kombinasyonu olan bant teorisi veya diğer bir adıyla pala elemanı momentum (PEM) teorisi, bir rüzgar türbini rotorunun aerodinamik tasarımı ve performans analizine yönelik bir prosedür yaratmak için kullanılmıştır. Ayrıca, aerodinamik kayıplar ve gerçekci bir akış alanı göz önüne alınarak optimum kanat tasarımı için bir başlangıç noktasına varılmıştır. Tasarımcı, operasyon sırasında meydana çıkan aerodinamik, yerçekimi, atalet ve kontrol kaynaklı kuvvetler nedeniyle oluşan yapısal deformasyon kaynaklı indüklenmiş burulma açısını kanat tasarımı sırasında dikkate almalıdır. Bu nedenle, hedeflenmiş çalışma koşullarında en iyi performansı gösterecek şekilde kanatların tasarlanması için kompozit rüzgar türbini kanatlarının yapısal davranışının iyi anlaşılması gereklidir. Böyle bir analiz kompozit mekaniği ve aerodinamik arasında bir etkileşim kurulmasını gerektirmektedir. Bu konuda ilk akla gelen üç boyutlu çözümler olan sonlu elemanlar ve hesaplamalı akışkanlar dinamiği ile etkileşimin sağlanması bir opsiyondur ancak zaman alıcıdır ve ön tasarım aşamasında hızlı çalışmaların yapılabilmesi için yeteri kadar esnek değildir. Alternatif yöntem hem aerodinamik hem de yapısal tarafta indirgenmiş modeller kullanarak tasarımcıya sistematik bir yaklaşım ve kanat davranışı hakkında bilgi sağlamaktır. Daha önce bahsedilen bölümde aerodinamik ile ilgili kısım Pala Eleman Momentum Teorisi'ne dayandırılarak gerçekleştirilmiştir. Sonraki bölümde ise yapısal kısma odaklanılmıştır, Klasik Laminasyon (KLT) ve İnce Cidarlı Yapılar teorisi kullanılarak uygulamaya yönelik ve hızlı yöntem bir geliştirilmiştir. Bu yöntem ön tasarım aşamasında parametrik çalışmalar yapmaya, malzeme ve laminasyon parametrelerini seçmeye yarayan bir aracın temelini oluşturmaktadır. Türbin kanatları kiriş gibi modellemeye uygun uzun ve narin yapılardır, bu nedenle kesit burulma ve eğilme katılığı gibi yapısal özelliklerini hesaplayacak şekilde bir modelleme gerçekleştirmek iyi bir yaklaşımdır. Bu yaklaşım için iki ana teori kullanılmaktadır. Birincisi kompozit yapının ana karateristiklerini hesaplamak için gerekli olan Klasik Laminasyon Teorisidir, ikincisi ise KLT ile profil geometrisi arasındaki ilişkiyi Kuran İnce Cidarlı Yapılar teorisidir. Tezin ilgili bölümünde yukarıda bahsedilen teorilerin detayları verilmiş ve NACA profili tabanlı kompozit kanat geometrilerinin mekanik özellikleri hesaplanmıştır. Tezin Aerodinamik ve Yapısal Etkileşimin Kurulması adlı bölümünde aerodinamik ve yapısal etkileşimin kurulması için gerekli teorik altyapı her iki disiplin için de oluşturulmuştur. Bu bölümde aerodinamik yüklerin kanattaki deformasyon sonucunda değişimi, yükler ve deformasyonun birbiri ile bağıntısı iteratif bir etkileşim kullanılarak ifade edilmiştir. Bunun için öncelikle aerodinamik yükler ve bu yükler kaynaklı deformasyon hesaplanmış, devamında ise bu deformasyon yeni aerodinamik yükler hesaplamak için kullanılmıştır. Bu yaklaşım biribirini takip eden iki iterasyonarasında kabul edilebilir bir hata seviyesine ulaşılana kadar devam etmiş ve final yük ve deformasyon durumuna ulaşılmıştır. Tez kapsamında yapılan çalışmaların uygulanması için farklı çalışma koşulları dikkate alınarak 12 m boyunda örnek bir kanat için parametrik hesaplamalar yapılmıştır. Bu hesaplamalar sonucunda kanadın rijit veya elastik kabul edilmesinin enerji üretim performansına etkileri ortaya konmuştur. Elastik kanat için önemli yapısal parametreler geliştirilen yazılım yardımıyla hesaplanmıştır. Örnek kanat üzerinde yapılan çalışmalar sonucunda temel olarak aşağıdaki farklar gözlenmiştir. • Kanat uç istasyonlarında oluşan yüklerdeki farklılık %2,5 seviyesindedir. • Türbin güç eğrisindeki farklılık rüzgar hızının artışı ile doğru orantılı olarak %0,1 - %12 arasında değişmektedir. Beklendiği üzere yüksek hızlar daha fazla deformasyona neden olmakta ve performans düşmektedir. • Türkiye'nin üç farklı ilindeki rüzgar ölçüm verileri kullanılarak yapılan çalışmada rijit ve elastik kanat kabulü arasında %1,3-%1,5 arasında değişen enerji üretim farklılıkları hesaplanmıştır. Geliştirilen yöntemler ile eğilme-burulma etkileşiminin performans üzerine olumlu yönde katkı sağlaması ihtimali de incelenmiştir. Amaç kanadın aerodinamik performansını en yüksek seviyeye çıkaracak optimum burulma açısının, etkileşim yardımıyla ortaya çıkarılması, serim açısı, veter uzunluğu ve et kalınlığı gibi parametrelerin optimum değerlerinin bu doğrultuda tespit edilmesidir. Tezin son bölümünde örnek kanat üzerinde bu doğrultuda bir çalışma yapılmıştır. Etkileri küçük olmakla beraber, eğilme-burulma etkileşimi sayesinde performansın artırılması potansiyel bir çalışma alanıdır.
Hibrit kompozit gövde profilli otobüsün ECE R66-02 yönetmeliğine uygun devrilme analizi
Son yıllarda, araç güvenliğine olan duyarlılığın artması, her geçen gün daha dayanıklı ve yolcu güvenliğini ön planda tutan araçların geliştirilmesinin önünü açmıştır. Kural koyucu kurumlar, son yıllarda araç üreticilerinin üzerindeki baskıyı arttırarak, özellikle otobüs sektöründeki devrilme kazalarında yaşanan can kayıplarını yasal düzenlemelerle azaltmaya çalışmaktadırlar. Bu nedenle büyük yolcu taşıtlarının devrilme kazalarında yapısal bütünlüğünü test etmek ve dayanımını arttırmak için ECE-R66 yönetmeliği Birleşmiş Milletler çatısı altında oluşturularak birçok ülkede yasalarla zorunlu hale getirilmiştir. Öte yandan, araç güvenliğinin artması araçlarda birtakım ağırlık artışına neden olmakta ve bu da paralı yükten feragat etmenin yanında, yakıt tüketimini de arttırarak çevresel birtakım etkilere sebep olmaktadır. Araç ağırlığındaki artış, üreticileri hafif ve sağlam alternatif malzemeleri kullanmaya itmektedir. Bunlardan bazıları yüksek dayanımlı çelikler, magnezyum alaşımlar ve bu çalışmanın da konusu olan kompozit malzemelerdir. Bu çalışmada, hibrit kompozit gövde profili kullanılarak, ECE R66-02 yönetmeliğine uygun, otobüs gövde barı tasarımı yapılacaktır. Kompozit malzeme olarak Toray Composite America (TCA) tarafından üretilen ve AGATE Programı (Advanced General Aviation Transport Experiment) tarafından testleri yapılan T700/2510 karbon elyaf/epoksi kompozit malzemesi kullanılacaktır. Analizlerde doğrusal olmayan açık kodlu LS-DYNA sonlu elemanlar programı kullanılmaktadır. Çalışmanın ilk bölümünde otobüs sınıfları, kaza özetleri ve kompozitlerle ilgili genel bilgiler yanında çarpışma dayanımları ve araç hafifletme konuları işlenmektedir. Bu çalışmanın ilham kaynağı olan, literatürde yapılmış kompozit otobüs çalışmalarına birçok örnek verilerek bu çalışma için gerekli zemin oluşturulmuştur. İkinci bölümde ECE-R66 yönetmeliği genel hatları ile anlatılmış, yaşam alanının modellenmesi, ağırlık merkezinin ölçümü gibi birtakım test koşulları hakkında bilgiler verilmiştir. Çalışmanın üçüncü bölümünde, kompozit malzemelerin sınıfları, mekaniği, hasarlanma şekilleri, çarpışma dayanımları, enerji emilim özellikleri ve üretim yöntemleri gibi birtakım bilgiler verilerek, kompozit malzemelerin çarpışma durumlarındaki birtakım avantajları ön planda tutulmuştur. Dördüncü bölümde sonlu elemanlar teorisi yanında, doğrusal ve doğrusal olmayan analizler hakkında bilgi verilerek bu çalışmada kullanılan yazılımlar hakkında genel bilgiler verilmiştir. Sonraki bölümde çelik yapı için TEMSA'da yapılan doğrulama ve onaylama testleri anlatılmaktadır. Doğrulama testleri kapsamında belirli bir profil bağlantısı üzerinde 3 nokta eğilme testi yapılırken, onaylaması testinde belirli bir otobüs kesiti ECE-R66 yönetmeliğine uygun şekilde belirli bir platformdan devrilmiş ve sonuçlar SEM analizleri ile doğrulanmıştır. Bu çalışmada bahsi geçen kompozit yapı için fiziki testler yapılamamış sadece analiz sonuçları verilmiştir. Altıncı bölümde LS-DYNA'da devrileme analizine uygun olacak şekilde otobüsün sayısal modellenmesi anlatılmıştır. Malzeme tipleri hakkında bilgiler yanı sıra yaşam alanı, temas tipi ve enerji değerleri hakkıda bilgiler verilerek analize hazırlık yapılmıştır. Son bölümde devrilme analizinde başarı sağlamış bir hibrit kompozit yapıdaki otobüsün yapısı incelenmiş, enerji denklemleri yanı sıra çelik malzeme ile yapılan otobüsle arasındaki farklar incelenmiştir.
Towards to the direct numerical simulation of insect flight
The parallel large-scale unstructured finite volume method based on an Arbitrary Lagrangian-Eulerian (ALE) formulation has been employed in order to investigate the near wake structure of Drosophila in hover and forward flight. The numerical algorithm is based on side-centered finite volume method where the velocity vector components are defined at the mid-point of each cell face while the pressure is defined at the element centroid. DISTENE MeshGems-Hexa algorithm based on the octree method is used to generate the all hexahedral mesh for the wing-body combination. The mesh deformation algorithm is based on the indirect radial basis function (RBF) method at each time level while avoiding remeshing in order to enhance numerical robustness. For the algebraic solution of the resulting large-scale equations, a matrix factorization is introduced similar to that of the projection method for the whole coupled system and we use two-cycle of Boomer AMG solver for the scaled discrete Laplacian provided by the HYPRE library, a high performance preconditioning package developed at Lawrence Livermore National Laboratory, which we access through the PETSc library. The present numerical algorithm is applied to the numerical simulation of flow field around a flapping Drosophila in hover and forward flight. The time variation of the Eulerian coherent structures in the near wake is analyzed with the lambda-2 criterion. In addition, the Lagrangian coherent structures is also investigated using finite-time Lyapunov exponents (FTLE) fields. For the symmetrical wing rotation with respect to stroke reversal, the flapping frequency is determined to balance the weight of the Drosophila. The body effect on the Drosophila flight is investigated by comparing the numerical results for the wing-body combination and no-body configuration. The visualization results indicate that the both cases are very similar to each other. Two different wing kinematic are applied for the Drosophila in forward flight in order to balance drag force due to forward motion. In the first case, the wing tip-path plane is tilted forward to generate forward force (as in helicopter rotor). In the second case, paddling wing motion is used to generate the forward force. The Eulerian coherent structures are analyzed with the lambda-2 criterion. In addition, a new monolithic approach is proposed to solve the fluid-structure interaction (FSI) problem with Lagrange multipliers in order to model free swimming/flying. The solid domain is modeled by the constitutive laws for the nonlinear Saint Venant-Kirchhoff material and the classical Galerkin finite element method is used to discretize the governing equations in a Lagrangian frame. In order to impose the body motion/deformation, the distance between each pair of nodes is imposed using the Lagrange multiplier approach, which is independent from the frame of reference. The resulting algebraic linear equations are solved in a fully coupled manner using a dual approach (null space method). The present numerical algorithm is validated for the two-dimensional three-link swimmer proposed by Eldredge (2008). In the future, the present method will be applied to Drosophila in free flight in three-dimensions.
Design of a star tracker for nano/micro satellites
With increasing interest to the small satellites, the capabilities demanding from them also increasing. For several years now, the missions can be performed with about thousand kilogram satellites, can be accomplished with micro and nano satellites nowadays. Especially the number of nano satellites that send to space is increased 7 times in 2012. [5] Primary reason of this is the cost of the nano satellites are decreased to let them be developed in a university. Low cost of a small satellites lead up to many universities to develop their own Cubesat and stimulate the development of subsystems for nano satellites. Thus the capabilities of Cubesats begin to increase. Parallel to this advance, more accurate attitude determination and control required at Cubesats. To compensate this increasing attitude determination requirements, sun sensors, nadir sensors, magnetometers and gyros are used many times. However, the star trackers was not a part of this sensor set for a while. Because of their high cost, high mass and high power consumption rates it was not possible to use them on Cubesats. Nowadays it is available a few different start trackers systems can be used in nano satellites. Their cost is already high but with increasing availability of the components used in a star tracker they are becoming cheaper. In this thesis, with taking into account of this situation , a preliminary work has been performed to develop a star tracker with a low power consumption, low cost and low mass for Cubesats. Primarily a used friendly simulation environment developed to simulate the images taken with a star tracker. With using this simulation environment, images with different camera properties and various noise ratios are generated and projected to a screen. The algorithms that will be used are run on this images and the performance of the algorithms monitored. With the help of this simulation environment an efficient star detection algorithm is developed and a preliminary work for star tracking algorithm is performed. As Star tracker camera Raspberry Pi NOIR Camera V2 is used in tests. The reason of choosing this camera for test is its easy integration with Raspberry Pi development board. Although the cameras properties are different from a star tracker camera, it is considered sufficient for tests. The algorithms are developed and tested in Python language on Raspberry Pi board and on a desktop computer. Desired results have been taken in test environment and some improvements in algorithms foreseen as future work.
Design, analysis and verification of conventional and non-conventional cylindrical cfrp composite shell with optimized cutout under the effect of combined loading
The rate of the usage of fiber reinforced composite structures both in aerospace and automotive industries has been increasing very rapidly over the past few decades. In addition to conventional fiber orientation laminate composites, the steered fiber layup composites have also become in the center of interests. Having higher specific strength and stiffness and being light-weight compare to metallic structures give polymer composites an important advantage over metals. Traditional fiber-reinforced composite laminates are composed of 0, 90 and 45 degree plies and are called as constant stiffness (CS) composites. Those specific fiber orientation selections are due to the ease of manufacturing, since at the initial times of composite manufacturing the only way was to manufacture manually. On the other hand, with the development of automated fiber placement machines, recently designed steered fiber composites have been able to put into service. This manufacturing technique allows fibers to be oriented other than 0, 90 and 45 degrees such that the curved fiber orientation within a ply changes continuously. Due to spatially oriented fibers within a ply, the laminate has spatially varying stiffness and therefore, those fiber composites are called as variable-stiffness (VS) composites. The purpose of having VS is to improve the mechanical properties of the composite. There have been various researches on the VS composites including flat panels, conical and cylindrical shells with some of the works including cutouts as well. The work presented in this dissertation includes the effect of cutouts varying circumferentially and geometrically on both CS and VS cylindrical shells under pure bending and combined loading. The geometry and cutout configurations of the CFRP cylindrical composite shell was selected to simulate the fuselage of the airframe. The structure under consideration is a circular cylinder with a radius R of 304.5 mm (12 in), length L of 813 mm (32 in), and an overall shell thickness H of 4.392 mm (0.173 in). Loading is introduced at each end of the cylinder and for this study is confined to a pure bending, applied in a direction so that the bottom of the cylinder is under compression and the top in tension. In the combined loading study, in addition to bending moment, vertical shear force applied from one end of the cylinder while applying encastre boundary condition to the other end. The results of a parametric study, concerning on the effects of a pair of cutouts symmetrically located with respect to the vertical axis perpendicular to the bending axis, varying both the location and size of the cutouts and evaluating their effects on the maximum load levels and predominant failure modes of a circular cylindrical shell were presented. Several baseline cases were presented to illustrate the type of results that were encountered. Overall failure trends from the study were introduced, and a supplemental study to highlight an interesting aspect of the results was demonstrated. The analysis methods that were used for the evaluation of the buckling and failure characteristics of shells include linear buckling analysis, nonlinear analysis using Riks method and nonlinear progressive failure analysis using Riks method. During the nonlinear analysis, the imperfections coming from the buckling analysis were also included at a rate of 10% of the shell thickness. Analyses were performed using the commercial FEM solution ABAQUS using S4R element. In that model, the stacking sequence was defined using the COMPOSITE option of the SHELL SECTION designator to define the fiber orientation angle of each layer. For the VS case, additional computations were required due to the spatial variation of the fiber orientation angle within the plies. Subroutines were generated with FORTRAN coding to produce fiber orientation angles for each ply. The fiber angle values were introduced into the model through the DISTRIBUTION TABLE designator within ABAQUS. This feature allows the user to construct tables to list specified values for each element or node. In this case, one table was used for each distinct ply and default values were used for the CS plies. During finite element analysis S4R shell element which possesses one integration point at the centroid of the element was used. The centroid was calculated from the four connected nodes of each element and the fiber orientation angle values were calculated there based on the parametric formulation of the circumferential variation. A Python script was used to construct both the input file and the tables based on the model parameters (geometry, mesh size, material properties, stacking sequence, and cutout configuration), where the tables used the FORTRAN subroutines interactively to calculate the fiber orientation angles. The tables were written to external files so they could be easily changed without having to alter the original input file. During nonlinear progressive failure analysis, Helius PFA commercial tool was used. It is composed of a set of software modules and a composite material library that integrates with the ABAQUS finite element analysis system. Helius PFA utilizes a form of material modeling that is based on Multi-continuum Theory in addition to other common composite failure criteria like Tsi-wu, Hashin, Larc etc. In the analysis of cylindrical shell model with cutouts, failure modes were defined and tables of failure modes were generated for the prescribed cutout size and locations. It was discovered that, there are generally three types of failure modes, which appears under pure bending for cylindrical shells with cutouts. To classify them; Failure Mode I may be defined as strength failure driven by global buckling of the cylinder. Failure Mode II can be defined as; strength failure around the cutout driven by local cutout edge buckling. As third mode, in Failure Mode III, before global buckling, material strength failure starts due to high stress concentration and therefore, it was defined as the strength failure initiated around the cutout without buckling. Finally, comparison of the finite element predictions with the experimental results were realized by matching the end rotations and strains. Although, the VS cylinder was designed to be stiffer than the CS cylinder and expected to have higher failure load in bending, after placing cutouts with certain size and locations on the cylinder, the effect of steered fiber laminae reduced dramatically. The effect of steered fiber is reduced due to the change in the failure mode as a result of applied cutouts. When the cutouts are below the neutral axis, that is the compression side of the cylinder, failures are usually driven by local buckling around the holes in contrast to global buckling from the bottom of the cylinder. The steered fiber laminate were designed to increase the failure load under pure bending and it can be effective in global buckling since the VS shell was optimized without cutouts. Therefore under local buckling conditions failure load were dropped for both CS and VS laminates. After completing the nonlinear analyses and finalizing the cutout size and locations, the cylinders without cutouts were machined with the 0.3R/120 cutout configurations. Following the machining, real imperfections were measured with Digital Image Correlation method. A final improvement of the finite element predictions was achieved by including real geometric imperfections in the model and by performing nonlinear progressive failure analysis using Riks method. Testing of cylindrical shells with cutouts was performed in TU Delft using a fixture that was designed to test the constant-stiffness and the variable-stiffness cylinders in pure bending. Strains and displacements were measured using strain gauges, digital image correlation, LVDT's and lasers. Two cylinders were tested: one with conventional layup laminate baseline cylinder (CS) and one with circumferentially varying laminate stiffness (VS). A comparison of the experimental response of the two cylindrical shells with the finite element predictions showed that the experimental boundary conditions has almost the same flexibility as modeled in the finite element model and resulted in good agreement between the experimental and the numerical results. The ultimate failure loads of testing also showed very good correlations with the finite element predictions. The results of this study show that, during the numerical analyses of composite structures it is important to apply nonlinear progressive failure analyses in order to obtain realistic results close to the test results. It was also proved that there are basically three failure modes of the cylindrical shells with cutouts. Although, the failure resulting from the global buckling of the cylinder has the highest failure load, generally the cylinders with moderate cutout size will tend to either buckle or tear up from the cutout edge. Therefore, in the design of cylindrical aerospace structures the window or door cutouts, the sizes and locations of the cutouts should be adjusted so that the buckling will tend to start from cutout edge to increase the post buckling strength of the cylindrical shells .
Hardware in the loop system development for modeling and control of multirotor vehicles
Unmanned air vehicles were designed to accomplish specific missions thats are dangerous to be fulfilled by man power. Nowadays the big companies want to reduce worker payments in companies budgets, so they tended to robotic applications. Not only big cargo but also restaurants or fast food companies are planning to carry out courier services via quadrotors for delivering cargo to costumers. Main reason for this changing is development of electronic devices, especially sensors, reduction of the prizes, and easily arrival UAVs parts. So, UAVs are going to be modern flying robots in the near future in all of the areas in mind. The topic of the thesis is development of hardware in the loop system to model and control of multirotor vehicles. Other words, we want to model and control multirotor vehicles on a testbed. Why testbed? The testbed gives some advantages. Reducing time consuming such as, eliminating the necessity of going to flight area, cheap workable conditions, and reduction the material and part breakage are most of them. This works is a sample of Model Based Design Process. Model Based Design (MDB) is a process that accelerating prototyping with cost-effective in a short time. MBD process phases are definition of requirement, mathematical modeling of the system(simulation), rapid prototyping, code generation,software in the loop (SIL), processor in the loop (PIL), hardware in the loop (HIL) and validation. First phase is definition of the system or product properties and requirements. The generating mathematical model of the system under boundary conditions effecting the system directly or indirectly in the simulation environment. Regarding to the mathematical model, rapid prototyping and code generation phase is coming to test the system on the simulations carry out on the computer environment that is the most important point to reduce cost and time consume. Software in the lop simulations are carried out to analyzing generated codes on the host computer. Processor in the loop is analyzing the system on target CPU on the embedded controller to reduction of the processor errors. Hardware-in-the-loop (HIL) simulation is a type of real-time simulation to test designed controller with real plant or system input out put via embedded hardware. HIL simulation shows how controller responds, in real time, to realistic virtual simulations. HIL can be use to determine if real physical system (plant) model is valid or not. The Gyroscopic testbed was designed for HIL tests of the quadrotor (IRIS+) for validating and implementing the designed different controller coefficients in real time environments. The HIL testbed is designed in order to how a quadrotor attitude control test easily. Before preliminary design, literature search is done and different test beds are found for both quadrotor testbed and motor propeller testbed. Then working principle of gyroscope is gave inspiration to design testbed. All of the designing processes are carried out on the computer aided drawing environment. Then production and assembly phase are done. Next problem is fixing the quadrotor on the testbench. A support part is designed and produced, then quadrotor vehicle is assembled on the testbench. The motor-propeller testbed is designed with same procedure of HIL testbench. Mathematical modeling of a quadrotor is second chapter of the thesis. Before the implementation of the controller, the simulation of the system give foreknowledge about behavior of the system. The mathematical model of the quadrotor is needed to simulate. Newton-Euler and Euler-Lagrange methods are used to derivate nonlinear model of the system. In the mathematical model, thrust and drag coefficients are used for refer propellers reaction to angular velocities. Some assumptions are taken into account during the modeling. The third chapter of the thesis is related with quadrotor and testbeds parameters estimation. IRIS+ quadrotor vehicle has four motor, one pair rotate clockwise, other rotate counterclockwise with propellers which are self locked type. IRIS+ quadrotor has cross type frame architecture that angle between front arms is one hundred twenty degree but back angle between back arms is one hundred forty degree. Cross type quadrotor has basically two advantages over plus configuration. first, motor need less power for pitch and roll motion and remaining power can be used for increasing translational motion. Second advantage is providing clear vision for camera with obtuse angle in between front propellers. The reference coordinate frames are defined related to cross type frame quadrotors. The quadrotor moments of inertias are founded by experimental studies.Blade element theory is used for calculating the propellers thrust and drag coefficients. Mathematical model of the HIL testbed is derived via Euler Lagrange energy method. The inner and outer circles affect the motion of the quadrotor seriously so differences of angels of the circles and quadrotor are taken into account to effect on moment of inertias. Designing flight controller for simulation and implementation is the last section of the thesis. Main aim of the section is comparison PID and feedback linearization methods outputs both in simulation and real time environments. In simulation, PID and Feedback Linearization Controller is applied on the both nonlinear model of quadrotor and HIL testbench nonlinear model in SIMULINK. PID responses are figured as phi, theta, psi, and Z parameters of the systems outputs. Feedback Linearization controller designed to eliminate nonlinear terms to stabilize the quadrotor vehicle attitude motion with different lambda coefficients. The last part of the simulations is designing feedback linearization controller for HIL testbed with optimum lambda coefficients of phi, theta, and psi. Lambda of Z is't needed because there isn't any motion in Z direction. Implementation of the PID and Feedback Linearization controllers are verified on the Pixhawk open-source autopilot hardware. Pixhawk code is redesigned to implement controllers in real time. The flight data is graphed in MATLAB to visualize motion with changes of angles for both PID and Feedback Linearization controller. Main difference between the results of the PID and Feedback Linearization controller is seen on the response of outputs. PID controller is tracking the signal with higher overshoots and bigger oscillations then Feedback Linearization controller response. But PID is faster than Feedback Linearization Controller. Both of the controller are stayed in effect of moments of inertias of HIL testbed circles. It is easily seen in long time delay on the time response especially in psi angle. Future works are planned to continue in this works line. Especially different nonlinear controller methods are to implement on airvehicles. Moreover, code generation on the autopilot hardware are planned on the UAVs.
Eğik şok kaynaklı laminer sınır tabaka ayrılmasının hesaplamalı yöntemlerle incelenmesi
Bu çalışmanın temel amacı düz plaka üzerindeki şok dalgası ile sınır tabaka etkileşiminin fiziğini kavramaktır. Bu çalışma sayesinde etkileşim bölgesinin hangi durumlarda daimi ve daimi olmayan akış meydana getirdiği ortaya çıkarılmıştır. Hava araçlarına olan zararı göz önüne alındığında, akış yapısının zamana bağlı etkilerini önlemek için plaka yüzey sıcaklığı değiştirilmiştir. Gerçekleştirilen çalışmalarda parametrik bir analiz serisi hazırlanmıştır. Gelen şokun açısı ile yatay düzlem arasındaki açı β her bir analizde değiştirilmiş ve yataydaki düz plaka üzerinde gelişmiş sınır tabaka ile etkileşimi incelenmiştir. Mach sayısı da parametrik olarak değiştirilmiş, bu değişimin şok dalgası sınır tabaka etkileşimi üzerindeki etkisi gösterilmiştir. Etkileşim esnasından oluşan şok dalgalarının, sınır tabakadaki ayrılma / yüzeye yeniden tutunma noktalarının ve ayrılma bölgesindeki baloncukların nicelik ve nitelik olarak değerlendirilmesi yapılmıştır. Gözlenen farklı akış yapıları zamana bağlı ya da zamandan bağımsız olarak sınıflandırılmış, ayrıca akışın 2-boyutluluk etkileri incelenmiştir. Analizler için çözüm ağı üretilirken dörtgen elemanlara sahip bir ağ kullanılmıştır. Bunun yanında akış fiziğini uygun modellemek, şok dalgalarının çarpma / yansıma noktalarını hassas olarak belirlemek için gelen / yansıyan eğik şok dalgaları civarına, sınır tabaka ve etkileşim bölgelerine özel düzenleme ile çözüm ağı sıklaştırılması uygulanmıştır. Bu sayede şok dalgası arkası ve önündeki ani değişen akış özellikleri ve etkileşim bölgesinde kalınlaşan sınır tabaka içindeki akış yapısı bu özel düzenleme ile gerçek fiziğe uygun çözüm alınmıştır. Parametrik analizler gerçekleştirilmeden önce, çözüm ağından ve zaman adımından bağımsız çözümler elde edildiği gösterilmiştir. Tüm analizler sıkıştırılabilir, zamana bağlı bünye denklemlerini Sonlu Hacimler Yöntemi (Finite Volume Method) ile çözmek için geliştirilen OpenFOAM çözücüsü ile gerçekleştirilmiştir. Çalışmaların açık kaynak kodlu yazılım ile gerçekleştirilmesi araştırmacıya bu alandaki bilimsel çalışmaları anlama ve yeni özgün çalışmalar yapma fırsatı tanımıştır. OpenFOAM yazılımı içinde bulunan birkaç sıkıştırılabilir akış çözücüsünden biri olan rhoCentralFoam yöntemi ile sıkıştırılabilir akış fiziğini modellemek mümkündür. rhoCentralFoam yoğunluk temelli sıkıştırılabilir akış çözücüsüdür ve Kurganov & Tadmor'un central-upwind şeması ile kullanılmaktadır. Kullanılan yöntem zamanda 1, uzayda ise 2. dereceden hassasiyet sağlamaktadır. Bu analizler bahsedilen özellikler ile beraber kullanılarak gerçekleştirilmiş, alınan sonuçlar kaynaklardan ulaşılması mümkün deneysel ve nümerik veriler ile karşılaştırılmıştır. Bu karşılaştırma ile kullanılan çözücünün ve yöntemlerinin modelleme kapasitesi de anlaşılmaya çalışılmıştır. Analiz sonuçları etkileşim bölgesinde yer alan baloncukların sayısı , alt duvar üzerindeki yüzey kayma gerilmesi ve basınç dağılımı sayısal olarak incelenmiş; ayrıca vorteks yapıları Q-kriteri metodu kullanılarak görselleştirilmiştir