Bir Taktik İnsansız Hava Aracı Kanadının Yapısal Davranışının Sayısal Ve Deneysel Olarak İncelenmesi

Abstract

Bu çalışmada insansız hava aracının kanadının yükler altındaki davranışı ve titreşim özellikleri sayısal ve deneysel olarak incelenmiştir. Taktik insansız hava aracının tek kanadı 3,5m uzunluğundadır. Kök veteri 840 mm ve uç veteri 630 mm’dir. Kanat tamamiyle kompozit yapıdan oluşmaktadır. Kanat imal edilirken karbon/epoksi, kevlar, cam elyafı ve kopük kullanılmıştır. Kanat üst ve alt parçaları ayrı ayrı imal edildikten sonra yapıştırma yöntemiyle birleştirilmiştir. Kanat kirişi kanat bitiminden itibaren bir metre daha uzatılarak gövdeye montajı sağlanmıştır. Kanat kirişi konik yapıda olup kiriş yuvasına sıkı oturması sağlanarak mümkün olduğunca ankastre mesnetlenmeye çalışılmıştır. Kanada uygulanacak yükler STANAG 4671 kriterlerine göre belirlemiştir. Bu kriterlere göre nihai yükleme durumunda hava aracının ağırlığının üç katı kanatlara yüklenmiştir. Hava aracının seyir uçuşu sırasındaki ağırlığı dikkate alınmıştır. Kanat üzerindeki aerodinamik yüklerin parabolik olarak dağılmaktadır. Uygulama kolaylığı bakımından bu yükler açıklık boyunca kademeli olarak değişen üniform yükler olarak uygulanmıştır. Kanat açıklığı doğrultusunda yükler değişim gösterirken, veter doğrultusunda yüklerin değişmediği kabulü yapılmıştır. Bu şartlar altında kanada yüklemeler yapılarak deneyler gerçekleştirilmiştir. Kanat ANSYS sonlu elemalar programında kabuk elemanlar ile modellenmiştir. Yapılan statik analiz sırasında uçuş yükleri kanada etki ettirilmiş ve kanattaki deformasyonlar gözlemlenmiştir. Titreşim analizi sonucunda kanadın mod şekilleri ve doğal frekansları elde edilmiştir. Deneysel çalışmalarda kanat gövdeye montaj edildiği gibi ankastre mesnetlenerek üzerine statik yüklemeler yapılmıştır. Bu yüklemeler sonucunda kanadın belirlenen yerlerinde gerineçler yardımıyla gerinimler ölçülmüştür. Yapılan titreşim deneyinde kanada anlık darbe ve ani bırakma yöntemleri ile ivmeölçerden alınan zamana bağlı veriler ile kanadın doğal frekansı elde edilmiştir. Elde edilen deneysel veriler ile sayısal veriler değerlendirilmiştir. Elde edilen sonuçlarda yapılan yaklaşıma göre belirli bir hata oranında gerçek değerlere yaklaşılmaya çalışılmıştır. Statik testlerde elde edilen sonuçlara göre analiz sonuçlarında kanat boyunca gerinimler yakın bulunurken, kanat veteri boyunca elde edilen gerinimlerde hata oranının daha fazla olduğu görülmektedir. Titreşim deneyinden elde edilen veriler ile analizlerden elde edilen veriler arasında ikinci ve üçüncü doğal frekanslarda yaklaşımlardan kaynaklı belirli farklar olduğu görülmüştür.

In the past aircrafts are made of metallic parts. Nowadays aircrafts are began to built from composites. Composites are stronger than metals against to corrosion, fatigue and they are lighter then metals. Barely it is important to check the strength of the part made of composites, because they are not isotropic materials like metals. Composites are combinations of two or more organic or inorganic components. One of those components is matrix which is the material that holds everything together while the other materials serves as a reinforcement. The most commen matrix materials were thermosetting materials such as epoxy. The reinforcing materials can be glass fiber or carbon fiber. Making composite structures is more complex than manufacturing aluminum structures. To make a composite structure, the composite material, in tape or fabric form, is laid out on molds and put in under heat and pressure. The resin matrix material flows and when the heat is removed. It is easy to make different shapes with composites. to increase strength, the fibers can be wound tightly to increase strength. Composites can be layered, with the fibers in each layer running in a different direction. This feature clears the way of designing structure that behave in certain ways. The most important value of composite materials is that they can be both lightweight and strong. The heavier an aircraft weighs, the more fuel it burns, so reducing weight is critical to aeronautical engineers. Controling of the structures of unmanned aerial vehicals made from composite materials affects life time and performance of aircraft. Not to increase cost of uav and not to harm civilians due to accident of uav, structures of aircraft must be built strong enough. Despite their strength and low weight, composites have not been a complete solution for aircraft structures. Composites are difficult to inspect for flaws. Some of composites absorb moisture. The most known problem is money about composites. They can be expensive, primarily because they are labor intensive and often require complex and expensive fabrication machines. Against to composites, aluminum is easy to manufacture and repair. Anyone who has ever gotten into a minor car accident has learned that dented metal can be hammered back into shape, but a crunched composite bumper has to be completely replaced. The same is true for many composite materials. In aviation, it is more critical, because producting composite for aircraft is more expensive than the other industries. The second potentially major cost adventages of composite structures is that they can eliminate mechanical fasteners nedded to assemble the structure. Assembly can account for up to half the recurring cost of an aluminum structure. The structure can be bonded using specially designed joints. Some of these assembly steps can be combined with curing, so that parts of the structure are cured and bonded in same operations. Composites can be made to absorb impacts, for instance, or the blast from an explosion. Because of this feature, composites are used in bulletproof vests and panels, and to shield airplanes, buildings, and military vehicles from explosions. Composites are strong to damage from the weather and from harsh chemicals that can eat away at other materials. Composites are good choices where chemicals are handled. Outdoors, they stand up to wide changes in temperature. In this project, it is investigated that the structural behavior and vibration features of a tactical unmanned aerial vehicle wing. The wing of tactical aerial vehicle has 3,5 meter lengths. Root chord is 840 mm end the tip chord is 630 mm. The whole wing is composite. During manufacturing carbon/epoksi, kevlar, fiberglass and foam were used. The wing was built from two main parts. These two parts are fixed to each other with glue. Spar was producted 1 meter longer than wing, so the wing could be fixed to the fuselage via using this part. Spar was designed as conical shape. Conical spar gave an advantage of clamped montage to the fuselage. To fix the wing to the fuselage a pin was used. The loads applied to the wing were calculated according to load criterias in STANAG 4671. Three times of the weight of the UAV were applied to the wing as maximum load. To determine the maximum load, the loiter weight of the UAV was considered. The load distribution was considered as elliptical but to apply the loads more practical the loads are used uniform in specified areas. Through the wing span the loads were changed because of the elliptical loading. It is assumed that the loads didn’t change in chord direction.according to this loads and boundary conditions, tests and analysis were made . In static analysis, on the downside surface of the wing the strains was close to the experimental datas. In critical places, such as at the root of wing, and on the spar, there were not extreme stresses. In vibration analysis, second and third modes were not match to the results of vibration tests. The wing was modeled in ANSYS finite element programe as shell structure wing. While analyzing the static case some loads are acted on the wing and deformations are viewed. At the end of the vibration analysis, mode types and natural frequencies were taken considering the weight of the wing. In tests, wing was clamped as installing to the UAV body, and some static loads were applied. Some strains values were taken from the specific locations on the surface of the wing after loads were applied by using straingages. At vibration tests natural frequencies were reached by method of applying impuls and instant removing loads to the wing by. Experimental datas and analysis results were compared. It was tried to have close results to experimental datas with a difference. According to results of static tests, it has been reached close results in analysises throughout wingspan but the results throughout cord were not close to test data. According to vibration tests, second and third natural frequencies are not close to test datas because of the assumptions.