Havacılıkta kullanılan kompozit malzemelerde hasar tespiti ve onarım yöntemleri
Inspection of damage and repair methods of composite materials used in aviation
Yazar: MEHMET FATİH KARAER
Danışman: DOÇ. DR. VEDAT ZİYA DOĞAN
İstanbul Teknik Üniversitesi / Fen Bilimleri Enstitüsü / Uçak ve Uzay Mühendisliği Ana Bilim Dalı / Uçak ve Uzay Mühendisliği Bilim Dalı
Derece: Yüksek Lisans, 2014
Konu: Havacılık Mühendisliği = Aeronautical Engineering ; Sivil Havacılık = Civil Aviation ; Uçak Mühendisliği = Aircraft Engineering

Özet:
Bu çalışmada havacılıkta sık kullanılan kompozit malzemelerde meydana gelen hasarların tespiti ve bu hasarların onarım yöntemleri ile ilgili araştırma yapılmış ve örnek kompozit plaka üzerinde hasar oluşturularak bu hasarın tespiti ve onarım basamakları ayrıntılı bir şekilde aktarılmıştır. Öncelikle kompozit plaka üzerinde çekiç ile hasar oluşturulmuştur. Oluşan bu hasarın tespiti için kompozit malzemelerde hasar tespit yöntemlerinden olan tab testi uygulanarak hasarın sınırları belirlenmiştir. Belirlenen bu sınırlar dikkate alınıp, hasarlı üst yüzey, uygun kesici aletler kullanılarak kaldırılmıştır. 
Sonraki aşamada çekirdekte hasarın olup olmadığı kontrol edilmiştir. Bu çalışmada çekirdek hasarlı olduğu için çekirdek de kaldırılmıştır. Hasarlı bölge uygun kimyasallar kullanılarak temizlenmiştir. Çıkarılan çekirdeğin boyutlarına uygun olacak şekilde yeni çekirdek gerekli kimyasal yapıştırıcılar kullanılarak yerleştirilmiştir. Yerleştirilen çekirdeğin tam yapışması için sıcak yapıştırma makineleri kullanılmıştır. 
Üst yüzeyin kumaşlarla kaplanması aşamasında kullanılacak kumaş türüne göre iki ayrı onarım yöntemi mevcuttur : Önceden rezin emdirilmiş kumaşlar (Prepreg) ve sonradan rezin emdirilen kuru kumaşlar (Wet-layup). Bu çalışmada her iki yöntem de ayrı ayrı uygulanmıştır. Her iki çalışma için de uygun kumaş türleri belirlenip gerekli boyutlarda kesilmiştir. Kesilen kumaşlar, yüzeyden kaldırılan katman sayısı kadar, çekirdek üzerine serilmiştir. 
Son olarak serilen kumaşların yapışmasının sağlanması için kür işlemine geçilmiştir. Kür işlemi için öncelikle kompozit plaka vakuma alınmış ve daha sonra sıcak yapıştırma işlemine geçilmiştir. Sıcak yapıştırma makinelerine uygun sıcaklık artışları ve işlem süresi tanımlanarak onarım sürecinin son basamağı da tamamlanmıştır. Sıcak yapıştırma süresi sona erdikten sonra kompozit plaka vakumdan çıkarılarak yapılan tamir tab testi ile test edilmiştir. 
Tamirin kalitesi, düzenli çalışma ile doğru orantılıdır. Seçilen kumaş türleri, kuru kumaşlar kullanılarak sonradan rezin eklenecekse eklenen rezinin miktarı, vakuma alınma sırasında hava kaçağının olmaması için tamirin yapıldığı ortamın tozdan arındırılmış olmasına dikkat edilmesi tamirin kalitesini etkileyen sebepler arasında sayılabilir. Bu çalışmada yukarıda bahsedilen işlem adımları ayrıntılı olarak aktarılmıştır. Daha açıklayıcı olması bakımından da fotoğraflar kullanılmıştır. Sonuç olarak, uygulanan her iki tamir yöntemi karşılaştırılmıştır.

Abstract:
Composites are used in a wide range of applications in aerospace, marine, automotive, surface transport and sports equipment markets. In this study, inspection of damages in composite materials used in aviation and repair methods of damages have been searched, and a sample composite plate has been damaged then repair steps of this damage have been identified in detail. First of all, composite plate was damaged with a hummer. Then, outline of this impact damage has been identified using the tab test method which is one of the  inspection method in composite plates. Damage to composite components is not always visible to the naked eye and the extent of damage is best determined for structural components by suitable Non Destructive Test (NDT) methods. Alternatively the damaged areas can be located by simply tapping the composite surface and listening to the sound. The damaged areas give a dull response to the tapping, and the boundary between the good and damaged composite can easily be mapped to identify the area for repair.

While drawing the outline of damage, markers which cause corrosion must not be used. There is special markers for drawing on composite materials due to prevent corrosion damages. Sharpie is a suitable marker as an example. After mapping out the damage, we need to remove the damaged material. We will normally use an air powered router, grinder or saw to remove the damage. It is important to remove only the damaged material. We should make the cut-out have a smooth shape. Corners are to have a 1.0 inch or larger radius. It is easy to find how many plies used in facesheet. The thickness of face sheet can be measured then divided by thickness of one composite ply. This is a pratical way. In addition number of plies can be identified by using engineering drawings or related structural repair manuels. In the next step, core has been inspected whether it is damaged or not. There are two ways to remove the core regarding the damage. If the core has not been damaged completely, then there is no need to remove full depth core. In this case, partial core should be removed. In this study, due to full depth damage, core was removed completely. Damaged area has been cleaned with suitable chemicals. 

Most repairs require that you taper sand the repair area. Taper sanding is required to provide an area for the repair plies to bond to the original plies. There are three important items to taper sanding: proper taper ratio, ply boundary identification, complete removal of damage.The area that you taper is a function of the number of plies or the skin thickness. The repairs require a 50:1 ratio for skins with less than six plies, and a 30:1 taper ratio for skins with six or more plies.To identify the number of plies in the area of damage after we taper sand, we must be able to see the ply boundaries. That is why a good even taper sanding is necessary. If the taper does not follow the contour, or is not straight for flat panels, we will not be able to identify the ply boundaries. After you taper sand the repair area, use close visual inspection to make sure that all the damage was removed.

There are several detrimental effects from water on composite parts. Resins naturally absorb some water during service. This absorbed water lowers the glass transition temperature of the resin. This makes the part weaker and less stiff at high temperatures. Absorbed water can also boil during the cure process. This can disbond repair patches and facesheets or simply degrade the repair bond. When water enters the honeycomb core of sandwich structures, this water freezes and expands at high altitude temperatures, damaging the core. The freeze-thaw cycles can cause significant damage over long periods. Water in honeycomb also increases part weight. In addition to performance reduction, the weight can affect the balance of flight control surfaces, and retraction of flaps and landing gear doors.

There are several methods to dry the repair area. One method is to dry the part in an oven. We can also use a vacuum bag and heat blanket or heat lamp. The oven method is preferred, but requires that the part be removed from the aircraft and small enough to fit into an oven. The other methods can be used regardless of location or size of the part.There are two basic methods to restore the core. We can fill the core with potting compound or bond in a new piece of core. The replacement core should be the same material as the core it replaces or an approved substitute. For general substitutions use the same core material with a greater density and/or smaller cell size. In the light of these informations core has been replaced with a new one, in the same size of removed core, using suitable chemical adhesives. Hot bonder has been used to get a quality bonding. 

There are two kind of repair methods related with using which kind of fabric  for the upper facesheet. In prepreg repair method, fabrics preimpregnated  with resin are used, but in wet-layup repair, fabrics are impregnated with resin by hand in the repair time. Both of these repair methods have been performed independently. For both repair types, suitable fabrics have been determined and have been cut with required dimensions.  After that,  same number of plies wtih the removed upper facesheet  have been laid on the core.  In the last step, all plies have been cured with a suitable cure pocess to have a quaility bonding. Before the curing, vacuum bag procedure has been applied to composite plate, and then hot bonder has been applied. 
The vacuum bag provides the seal between the repair and the pressure source. For most of our repairs the pressure source is atmospheric air pressure. That is the pressure caused by the weight of the air around us. With this system, the pressure is applied by removing air from inside the vacuum bag. This is different from an autoclave which uses pressurized gas to provide the pressure. The vacuum bag provides the barrier between the repair and the pressure source. There are two types of vacuum bags : surface bag and envelope bag.

A surface bag is normally easier to make for most repairs. However, it requires the part have a clean solid surface in order to maintain vacuum integrity. If the part has a hole in it, then you may be able to make two surface bags, one on each side of the part. An envelope bag can be used on parts where a surface bag would be difficult to install. If the part is small, then it can be easier to put the part into an envelope bag. An envelope bag can also be used where some part surfaces are not smooth and can be very difficult to surface bag.

The most common method to heat your repair is to use a heat blanket with a computer controlled hot bonder. The thermocouples provide temperature inputs to the controller. For this study, suitable heat up rates and cure times have been identified to hot bonder. After curing process, composite plate has been transfered to out of vacuum bag and the repair has been tested with tab test method.  There are some mechanic tests to perform composite materials to get material characteristics. For example, tension test, compression test, shear test, bending test, etc. In this study tension test has been applied to three sandwich composite plates. One of the parts is damaged, the other one is repaired, and the last one is original part without any damage. Tension loads have been applied all these three parts individually. The tension rate was 2 mm per minute. As it is expected, the repaired part has been cracked from repaired area. The damaged parts has been failed from damaged area, and the original undamaged part has been  failed from the middle of part. The quality of repair is proportional with right working. Types of fabrics, the quantity of resin if fabric will be impregnated by hand in the repair time,  working in a clean room to not have a vacuum bag leak, all this reasons affect the quality of repair. In this study,  all steps of repair process have been identified in detail. Moreover, digital photographs have been used to make it more understandable. In  conclusion,  performed two type of repairs have been compared.