Mekanik bağlantılı kompozit yapıların ilerlemeli hasar analizi
Progressive damage analysis of mechanically connected composite structures
Yazar: ŞEYMA SALİHA PALA
Danışman: PROF. DR. ATA MUĞAN
İstanbul Teknik Üniversitesi / Fen Bilimleri Enstitüsü / Makine Mühendisliği Ana Bilim Dalı / Katı Cisimlerin Mekaniği Bilim Dalı
Derece: Yüksek Lisans, 2020
Konu: Makine Mühendisliği = Mechanical Engineering

Özet:
Kompozit malzemeler, makroskobik düzeyde birleştirilen ve birbiri içinde homojen bir şekilde çözünmeyen birden fazla bileşenden oluşan, çeşitli dayanım ve ağırlık avantajlarından dolayı havacılık sanayisinde sıklıkla kullanılan bir malzemedir. Hava araçları çeşitli nedenlerle (dolu, kuş çarpması, serviste aletin düşmesi vb.) dinamik hasarlara maruz kalırlar. Özelikle mekanik bağlantılı kompozit yapılarda bağlantı elemanların varlığından dolayı çeşitli hasar tipleri oluşabilmektedir. Bağlantı elemanlarının ve kompozit yapıların tüm uçuş boyunca herhangi bir hasara uğramadan bütünlüğünü koruması gerekmektedir. Çünkü deliğin etrafında hasar meydana gelmesi demek vidanın fonksiyonunu gerçekleştirmemesi, dolayısıyla istenmeyen durumların oluşması anlamına gelmektedir. Bu nedenle delik etrafında ve bağlantı elamanında herhangi bir hasar oluşumuna izin verilmemelidir. Bu amaçla tasarımı yapılan kompozit panellerde belirli bir ezilme hasar dayanım limiti belirlenmelidir. Hava aracı üstüne entegrasyonu gerçekleştirilen kompozit radomların uçuş yükleri altında doğrulama analizleri yapılırken, mekanik bağlantılı kompozit yapılarda yatak dayanım limit değerini elde etmek gereklidir. Bu amaçla havacılık standartına uygun çeşitli testler gerçekleştirilmektedir. Dünyada havacılık standartına uygun bu testleri gerçekleştiren çok az firma olmakla beraber test maliyetleri çok yüksektir. Bu amaçla havacılık standartına uygun gerçekleştirilen testlerin bilgisayar ortamında benzetiminin yapılması ihtiyacı oluşmuştur. Bu amaçla ilk olarak uçak üzeri entegrasyonu gerçekleştirilen kompozit radomların plaka takviye ve dizilim açıları bilgileri kullanılarak çeşitli test kuponları üretilmiştir. Bu test kuponlarının test simülasyonları bilgisayar ortamında gerçekleştirilmiştir.
Kupon çeşitliliğini ve bağlantı tipini çok tutulmasındaki amaç, farklı parametrelerinin etkisini rahat bir şekilde gözlemlemektir. Simülasyon modeli üç boyutlu modellenmiş ve analizler ANSYS programı ile gerçekleştirilmiştir. Tüm kupon analizlerinde Hashin ilerlemeli hasar analizi metodu kullanılmıştır. İlermeli hasar analizi sonlu eleman modeli cıvata, kompozit panel, burç, metal levha üç boyutlu ve elastik olarak test fikstürü ise rijit modellenmiştir. Parçalar arasındaki temas yüzeyleri tanımlanarak malzeme indirgeme kuralı ile ilerlemeli hasar analizleri gerçekleştirilmiştir. Bilgisayar ortamında yapılan ilerlemeli hasar analizi çalışmaları gerçekleştirilen deneysel çalısmalar ile karsılastırılmıs ve %2-%10 oranlarında farklılık elde edilmistir. Gerçekleştirilen deneysel ve sayısal sonuçlarda düşük hata oranı yanında benzer hasar şekilleri de elde edilmiştir. Ayrıca kompozit yapının ezilme hasar dayanımında ön gerilme, hasar teorisi tipi, burç kullanımı, civata delik merkezi ile levha kenarı arasındaki uzaklığın delik çapına oranı (E/D), levha genişliğinin delik çapına oranı (W/D), sıcak-nemli ortam koşulları, sürtünme katsayısı gibi çeşitli parametrelerin etkisi incelenmiştir.

Abstract:
Composite materials are a material that is frequently used in the aviation industry due to its various strength and weight advantages, consisting of multiple components that are macroscopically combined and are not homogeneously dissolved in each other. Aircraft are exposed to dynamic damages for various reasons (hail, bird strikes, falling of the tool in service, etc.). Especially in mechanically connected composite structures, various types of damage can occur due to the presence of fasteners. Fasteners and composite structures must maintain their integrity throughout the entire flight, without any damage. Because damage to the hole around the hole means that the screw does not perform its function, thus causing undesirable situations. For this reason, any damage should not be allowed around the hole and on the connecting element. For this purpose, a certain crush damage resistance limit should be determined for the composite panels designed. While verification analysis under flight loads of composite radomes integrated on the aircraft, it is necessary to obtain the bearing and pull through strength limit value in mechanically connected composite structures. For this purpose, various tests are carried out in accordance with aviation standards. Although there are very few companies in the world that perform these tests in accordance with aviation standards, test costs are very high. For this purpose, it is necessary to simulate the tests performed in accordance with aviation standards in computer environment. For this purpose, various test coupons were produced by using the plate reinforcement and alignment angles information of composite radomes, which were first integrated on the aircraft. Test simulations of these test coupons were carried out in computer simulation.
The thesis deals with the strenght limits of screws in two different analysis models: pull bearing and pull through.
Pull bearing analysis approach is specified in the next senteces.The purpose of keeping coupon diversity and connection type very much is to easily observe the effect of different parameters. The simulation model was modeled in three dimensions and analyzes were carried out with the ANSYS program. Puck progressive damage analysis method was used in all coupon analysis. three dimensional progressive damage modelling program has been developed to investigate the behavior of bolted laminated composite plates, to determine the failure modes and maximum load carrying capacities under tensile loading. Both stress and failure analyses have been performed using this program in conjunction with a general finite element code, ANSYS. In the progressive damage model, Puck failure criteria have been applied and according to the damage type, material properties of composite plate have been degraded. Specimens manufactured from out-of-autoclave carbon-epoxy prepregs were transversely loaded in an electromechanical testing machine. The load-displacement response was measured. Numerical analysis was conducted to predict joint failure by using a simplified symmetric 3-D finite element approach. Contact elements were used under the region of the fastener head. Progressive damage analysis finite element model bolt,composite panel, bush, metal plate are three-dimensional and the test fixture is rigidly modeled. Progressive damage analyzes were carried out with the material reduction rule by defining the contact surfaces between the parts. Progressive damage analysis studies performed in computer environment have been compared with the experimental studies carried out and a difference of 2-10% has been obtained. In the experimental and numerical results, similar damage types were obtained besides low error rate. In addition, pre-stress in the crush damage strength of the composite structure, damage theory type, bushing, the ratio of the distance between the bolt hole center and the plate edge to the hole diameter (E / D), the ratio of the plate width to the hole diameter (W / D), hot-humid environment conditions, The effects of various parameters such as the friction coefficient have been investigated.
Pull through analysis approach is specified in the next senteces. The design and certification of composite aircraft structures requires extensive and costly testing of bolted joints since no robust predictive tools are available. The majority of investigations in this field focus on the shear loading of bolted carbon fibre-reinforced polymer matrix composites while studies on the behaviour of pull-through are limited. Thus, this work presents an extensive investigation of the factors affecting bolted composite panels subjected to out-of-plane loading with the objective to develop a semi-empirical predictive model.
The purpose of keeping coupon diversity and connection type very much is to easily observe the effect of different parameters. The simulation model was modeled in three dimensions and analyzes were carried out with the ANSYS program. Puck progressive damage analysis method was used in all coupon analysis. three dimensional progressive damage modelling program has been developed to investigate the behavior of bolted laminated composite plates, to determine the failure modes and maximum load carrying capacities under tensile loading. Both stress and failure analyses have been performed using this program in conjunction with a general finite element code, ANSYS. Specimens manufactured from out-of-autoclave carbon-epoxy prepregs were transversely loaded in an electromechanical testing machine. The load-displacement response was measured. Numerical analysis was conducted to predict joint failure by using a simplified symmetric 3-D finite element approach. Contact elements were used under the region of the fastener head.
The experimental results showed that failures were dominated by inter-ply delamination and through-thickness shear failure of the laminate. A conical damage zone was found under the fastener head with damage radially dispersed from the fastener hole. The finite element analysis showed that a concentrated shear stress area under the fastener head caused the cone damage initiation. Progressive damage analyzes were carried out with the material reduction rule by defining the contact surfaces between the parts. Progressive damage analysis studies performed in computer environment have been compared with the experimental studies carried out and a difference of 2-10% has been obtained. In the experimental and numerical results, similar damage types were obtained besides low error rate.
As a result, certain crush damage resistance limit was be determined for the composite panels designed. While verification analysis under flight loads of composite radomes integrated on the aircraft, it is necessary to obtain the bearing and pull through strength limit value in mechanically connected composite structures. For this purpose, various tests were applied in accordance with aviation standards. Although there are very few companies in the world that perform these tests in accordance with aviation standards, test costs are very high. For this purpose, it was carried out to simulate the tests performed in accordance with aviation standards in computer environment.