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Uçak yapılarının burkulması, boşlukların analizi ve bağlayıcı seçimi

Buckling of airframe structures cutout analysis and fastaner selection

  1. Tez No: 66549
  2. Yazar: BİROL UZ
  3. Danışmanlar: PROF. DR. AHMET NURİ YÜKSEL
  4. Tez Türü: Yüksek Lisans
  5. Konular: Uçak Mühendisliği, Aircraft Engineering
  6. Anahtar Kelimeler: Belirtilmemiş.
  7. Yıl: 1997
  8. Dil: Türkçe
  9. Üniversite: İstanbul Teknik Üniversitesi
  10. Enstitü: Fen Bilimleri Enstitüsü
  11. Ana Bilim Dalı: Uçak Mühendisliği Ana Bilim Dalı
  12. Bilim Dalı: Belirtilmemiş.
  13. Sayfa Sayısı: 285

Özet

ÖZET Bir uçağın en önemli üç yapısal bileşenini kanatlar, gövde ve kuyruk takımı oluşturmaktadır. Bu yapısal bileşenler de çubuklar ve profillerden oluşmaktadır. Yapısal dizayn gözönüne alındığında; çubuk ve panellerdeki burkulma kararlılığı, panel elemanlanndaki giriş delikleri ve çubukların bağlantı boşlukları, tasarım mühendislerinin daima başım ağrıtan konular olmuştur. Yapısal hesaplar yapılırken en mukavim yapının en az ağırlıkla sağlanması uçak mühendisliğinin temel prensibidir. Bu çalışmanın amacı da verilmiş yük durumları ve dış boyutlar şartları altında minimum ağırlığa sahip yapının hesaplanmasına dair bir metod sunmaktır. Belki de burkulmanın en çok problem yarattığı endüstri havacılık endüstrisidir. Çünkü sürekli olarak yük dağılımında değişmeler olmakta ve malzeme şok yüklemelere maruz kalmaktadır (iniş takımlarında olduğu gibi). Bu yüzden hem kiriş-kolon yapıların burkulması hem de kabuk-kiriş elemanlardan oluşan panellerin burkulmasına ayrı ayrı değinilmiş ve öndizayn için oldukça iyi sonuçlara ulaşılmıştır. Uçak yapısı; kontrol çubukları veya kabloları, hidrolik hatlar, elektrik kablo hattan gibi elemanlar yada panel ve dokular geçmesi için gerekli boşluklarla birlikte, sürekli bir yüzeydir. Bunun dışında pencereler, kapılar, servis panelleri, kapaklar, bomba bölümleri, kontrol yapmak için giriş yuvalan v.s. herbiri ayn analiz yöntemlerigerektiren irili-ufaklı birçok boşluk vardır. Çeşitli boşluk yapılarının tasarlanması için bazı metodlardan sözedilmiştir. Tek bir parça halinde üretilecek bir yapı ideal uçak yapısıdır. Fakat pratikte bunu uygulamak imkansızdır. Gerçek bir uçak yapısında birçok bağlantı parçasının birleştirilmesi gerekmektedir. Ağırlığın daima kritik problem olduğu uçaklarda ise yüksek mukavemet değerlerinin hafif bağlayıcılarla sağlanması istenmektedir. Bağlayıcılar hakkında geniş bir araştırma Bölüm 3'te sunulmuştur. Bölüm 4'te bir uçağın ağırlığının tamamım ana lonjeronun taşıdığı kabul edilerek, bu lonjeronun boyutlandırması yapılmıştır. Sonuçta ele alman malzemeler içerisinde en hafif yapıyı 7075-T6 alüminyum alaşımının sağladığı görülmüştür.

Özet (Çeviri)

SUMMARY BUCKLING OF AIRFRAME STRUCTURES, CUTOUT ANALYSIS and FASTENER SELECTION The most important three structural components of an aircraft, are wings, fuselage and empennage. These include beams and panels. When structural design regarded stability of buckling in beams and panels, service openings and cutouts of beam in panel elements have been to subjects which caused troubles to design engineers. When structural calculations are done, it's the basic principle of aeronautics to have the strongest structure with the lightest one. The purpose of this study is under the given loads and outer dimensions conditions to present a method related to the calculation of the structure which has the minimum weight. In the application part, the analysis of cutouts which are used to make the spar light, fasteners and dimensions (chord length and profile known) concerning to the main spar attached to wing profile have been made. Buckling is the general term frequently used in aircraft analysis to describe the failure of a structural element when a portion of the element moves normal to direction of primary load applications. The equilibrium related to the strength of buckling is defined by using the given equation below: F“=KE '^' ^gJ (Eq.l) Where:gi and gz are geometry term and K is a parameter which is a function Xlllof geometry, boundary conditions and loading condition. The problem buckling in aircrafts occur İn different cases. In a beam- column buckling the main feature is the lateral bending of the main element This lateral bending occurs in the axes in which inertia moment is, as shown Fig.l..^m Axis of mintanım moment of inertia 9eotlra A”A A- Figure 1 Primary failure or bucling of a column member The Euler equation (Eq. 1) is universally to describe flexural stability of elastic or long columns. However, in the short and intermediate length range the column behaviour can be considerably considerably more complex. In this range the column strength is highly dependent upon the column cross-sectional details and the material inelastic properties. A beam column is a compression member which is also subject to bending. The bending may be caused by eccentric application of the compressive load, initial curvature of the member, lateral loading, or any combination of these loadings, (see Fig. 2) If the bending moment in a beam-column with a compressive load P is compared to the bending moment in a simple beam with no axial load, the following approximate relationship is obtained: M0 JKn«= - 7 1- \P“; ef XIV^^^^max is the beading moment ia the beam-column? M> fwL^ \ ö y bending moment in a simple beam resisting the same lateral loading, and Pa - is the is the Euler column load. P ? ?*-.”^--~-- *f- T < - p support support (a) Locate eccentricity at supports support support i «maks support“f L (”inch") (b) Initial eccentricity due to manufacturing tolerance Pi P2 P3 P4 P5 Pö P7 Pi P p i £ *A y y y y y P U- (c) with lateral loads Figure 2 Typical beam-column cases For thin plate buckling, the above equation becomes, 2 T7^2 knlEt F -. I2(l- n£)V (Eq.2) or F^mB (Eq,3) XVTK&ı sküı-strihger panel construction is a tögfcat development of the necessity of providing a continuous surface for an aeroplane, combined with the requirement that the weight of structure should be as smalt as possible. The components of sldn- stringer panel may be classified as follows:. Longitudinal reinforcing member. Skin. Transverse reinforcing members Inter-rivet buckling is of special importance on compression panels with skin attached by rivets. stringer Figure 3 Inter-rivet buckling Normally the skin-stringer construction will be designed so that the rivet spacing is derived from the crippling stress of the stringer. However, when the inter- rivet buckling stress of the skin is reached before the crippling stress. The high performance levels in machines and equipment continue to pkce more exacting demands on the design of structural components. In aircraft, where weight is always critical problem, integrally stiffened structural sections as shown Fig. 4 have proved particularly effective as a lightweight, high strength construction. XVI(crcsüed ft Vtfbmgedsec&x Figure 4 Three popular sections of integrally stiffened panels. The aircraft structure is continually faced with requirements for openings at webs and panels to provide access or to let other member such as control rods or cables, hydraulic lines, electrical wire bundles, etc. pass through. Perhaps the most noticeable feature of cutouts is the rounding of the corners; sharp corners cause excessively high stress concentrations. Cutouts in aircraft wing and empennage structures constitute one of the most troublesome problems confronting the aircraft designer. Because the stress concentrations caused by cutouts are localised, a number of valuable partial solutions of the problem can be obtained by analysing the behaviour, under load (axial or shear load) of simple skin-stringer panels or integrally stiffened panels. For reduction of the number of stringers, three stringer method can used. Figure 5 _.-.- Coaming stringers Skin-stringer panel with rectangular cutout (actual condition) strtn&cts xvuRib static» Figure 6 Three stringer application A complete airplane structure is manufactured from many parts. These parts are made from sheets, extruded sections, forgings, castings, tubes or machined shapes, which must be coined together to form subassemblies. The subassemblies must then be joined together to form larger assemblies and then finally assembled into a completed airplane. The ideal aircraft structure structure would be a single complete unit of the same material involving one manufacturing operation. Unfortunately the present day types of materials and their method of working dictates a composite structure. Furthermore, general requirements of repair, maintenance and stowage dictate' a structure of several main units held to other units by main or primary fittings or connections, with each unit incorporating many primary and secondary connections involving fittings, bolts, rivets, welding etc. For structural economy, engineers in the initial layout of the aircraft should strive to use a minimum number of fittings, particularly those fitting connecting units which carry large loads. Thus, in wing structure, splicing the main beam flanges or introducing fittings near the center line of the airplane are far more costly then splices or fittings placed farther outboard where member sizes and loads are considerably smaller. Avoid changes in direction of heavy members such as wing beams and fuselage longerons as these involve heavy fittings. If joints are necessary in XV1Ucontinuous beam place them near points of inflection in order that the bending moments to be transferred through the joint to be kept to a small magnitude. In column design with end fittings, avoid introducing eccentricities on the beam; on the other hand make use of fitting to increase column end fixity. In making a fastener selection, the designer must write down alt the conditions to be en countered by the overall design. These are not specific requirements which are determined for each part but are the general ranges over which the entire aircraft is expected to operate. After the designer determines the type of joint to be used, he than determines whether the fastener is loaded axially or in shear. For static tension loading, the same enlarged root radius are not essential For joints with incidental tension loadings, nearly any fastener can be considered, including rivets, blind rivets and shear head fasteners. Shear loading requires consideration of fasteners such as bolts or swaged collar fasteners with conventional or shear heads. Shear heads offer the higher strength -weight ratios. For shear fatigue loadings, the designer selects fasteners capable of interference fit or of developing high residual compression stresses in the structural material around the holes. In the design process, the designer considers the relationship between cost, weight anf function in selecting from the previously chosen fasteners, the optimum combination of these factors should be carefully considered. In some military applications, ground support equipment and facilities may be required, the design case limits the choice to bolts or screws. Inaccessibility, when complete, requires the use of blind fasteners unless previously installed nut plates can be used with bolts or screws. XIX

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