Лако-красочные материалы — производство

Bonding is a surface-to-surface and material joining technique that makes it possible to join any combination of similar or dissimilar materials. In aircraft manufacture, bonding offers the following benefits compared to riveting, joining by screws, soldering and welding:

• High aerodynamic surface quality of the bonded components owing to smooth, high-precision contours.

• Excellent properties with regards to stiffness, fatigue resistance and damage — tolerance owing to a flow of forces that is evenly distributed over the surface and is free of notches. This is combined with good damping properties of the adhesive layers in terms of vibrations and even ultrasonic fatigue that may be induced by the engines.

• A possibility to bond different materials without impairing the adherents via the influence of temperature, introduced stresses or damages of the surfaces.

• Easy construction of sandwich structures using lightweight core materials and outer sheets, and the possibility to manufacture multilayer structures in the form of metal laminates with good damping properties and high resistance against crack propagation through the adhesive layers.

• Gas-proof and liquid-tight properties, as well as electrochemical insulation of the adherents via the adhesive.

As mentioned above, in connection with the long-term resistance in humid environments, there are some disadvantages which must be considered when designing adhesively bonded joints. It is possible, however, to limit these disadvan­tages to a foreseeable extent. It is important to note (and will shortly be illustrated) that the very high-quality standards which must be applied during the production process are prerequisites for high-quality bonds. Furthermore, nondestructive testing of the strength characteristics of a joint is only possible to a certain extent. With today’s common nondestructive testing methods, it is only possible to detect a lack of adhesive, a macroscopic delamination, and sometimes — especially with phenolic resins — the results of foaming that takes place during the course of curing if the pressure is not uniformly applied. Finally, with regards to long-term aspects, adhesive joints are characterized by some well-known degradation reactions which must be carefully investigated again and again with regards to their actual hazard potential.

This must be done as soon as new adhesive systems are developed, although they do not represent any exclusion criterion with regards to high-strength structures intended for a long service life. Within the adherents, degradation mechanisms

may also occur which represent a well-calculated risk, particularly with regards to long-term resistance.

Bonding technology is a must, especially in the manufacture of large metal aircraft structures. In the skin of a wide-bodied aircraft with a fuselage diameter of (commonly) 5.5 m and more, the fuselage shell must bear stresses of around 110-120 N mm~2, which is easily calculated from a cabin pressure of 0.7-0.8bar (0.7-0.8 x 105 N m~2) that is needed to ensure passenger survival, and a common sheet wall thickness of 1.6 mm; gust loadings must also be added to this. The fatigue resistance of the aluminum alloys used, determined using polished round bar samples, is approximately 130 N mm~2 in terms of their fatigue limit during near­unlimited alternations of load; this does not leave much room for structural purposes, especially with regards to the weight. The same is true when considering that these types of stress only occur during flights, and that the number of flights will correspond to 60 000-70 000 loading periods (in terms of the outer cover of the fuselage and for the inner pressure at high altitude) for short-haul aircraft, and to a maximum of 40 000 periods for long-haul aircraft. Consequently, the aircraft are not sized in the area of their fatigue limits, but rather in the area of their fatigue strengths. It must also be taken into account that aircraft structures are extremely notched structures; apertures for cables, windows and doors and, in the case of a wide-body aircraft up to 3.5 million rivet holes, will weaken these metallic structures. In aircraft manufacture, bonding technology allows reduced stress concentrations to be achieved in the bond-lines. Furthermore, it is also possible to reinforce the common wall thickness of 1.6 mm by a doubling of the skin by using bonded sheets of 1.6 mm thickness or very often not more than 0.6 mm, in the areas of notches, windows, doors and longitudinal joints, as well as transverse joints. Bonding also has economic advantages when compared to the so-called ‘integral construction’ method, as the latter involves a reduction in structural surfaces to minimum thickness by means of chemical surface removal in areas where there is no need for a greater thickness. The service life may also be improved by the use of ‘bonded doublers’, most likely as a result of the good damping properties of adhesive layers (notably in the areas of longitudinal and transverse joints), although these joints are generally riveted. This situation can be demon­strated by the following example, where a typical longitudinal j oint in a wide-bodied aircraft is created in the form of a simple lap joint that is then riveted. If, for the purpose of stiffening, doubler sheets are merely inserted, only one-fourth of the service life is achieved compared to a lap joint where doublers are bonded to both skin sheets to be joined. It must be noted that the riveting joint between both panels is not bonded. The fatigue resistance of the longitudinal and transverse joints, respectively, may again be increased by a factor of 2 if an adhesive layer is introduced into the joint, in addition to the rivets. This method was adopted long ago by Fokker, who used two-part epoxy resin adhesives.

Adhesive joints have another advantage: if cracks develop in the skin sheets, they will only run up to the adjacent stringers and frames, where they are then redirected by the adhesive layers and run back to the skin sheets (Figure 8.5). This effect is most likely due to the excellent damping properties of the adhesives. Although cracks

cannot be avoided completely, in this way it is possible to control the risk with regards to the overall structure by specifically arranging the stringers and frames, thus preventing crack propagation. The thickness of the sheets may be reduced by using sandwich structures with lightweight cores and thin metallic outer sheets. This will save about 15% of the structural weight which, for an average wide-bodied aircraft, is approximately 351, as clearly demonstrated by investigations performed in the United States [7].

It is clear, however, that even today there are limits to the degree of weight reduction possible in metallic aircraft structures, and this point must be considered when building ever-larger, more economic aircrafts. The result has been the initiation of some revolutionary developments in aircraft manufacture. The most important factor here is the strength of the construction material, and in particular its modulus of elasticity in relation to its specific weight. High-strength aluminum alloys, for instance, have a tensile strength in a range of 350-400 N mm~2 and an elastic modulus of approximately 80 000 N mm~2, the specific weight (density) of alumi­num being 2.7 gcm~3. Glass fiber-reinforced plastics, which have been used for many years, have similar values, their fatigue strength reaching more favorable values as compared to the initial strength, and their density being only around 2 g cm~3. In recent years, considerable progress has been made when using carbon fibers (which today are available worldwide) for the production of fiber-reinforced plastics. This has resulted in an increase in the strength and elastic modulus by factors of 2, and a density of 2 g cm~3. When building large aircraft structures, improving the elastic modulus leads to a considerable reduction in the problems of stable dimension and stiffness. Compared to aluminum structures, fiber-reinforced plastics also have a considerably lower risk of fire — an advantage which is often

neglected. In time, therefore, fiber-reinforced plastics — and especially carbon fiber — reinforced plastics — should replace metallic structures far beyond the areas of application where they have been used previously in aircraft manufacture. Moreover, the high costs associated with fiber-reinforced components may be reduced by replacing complex, difficult-to-laminate monolithic structures with plates or profiles that are of a simple design and can be adhesively bonded. Bonding is the best joining technique for those inhomogeneous material structures that have a high notch sensitivity. For example, when using epoxy resin matrices (as is common practice today), bonded carbon fiber-reinforced materials have a good long-term resistance. It must be noted, however, that it is still necessary to provide a mechanically preformed surface preparation. For that purpose, fabrics are laminated onto the outer sheets of the areas to be bonded and then, after curing, peeled offbefore adhesive application; in this way a rough surface is obtained. Another possible approach would be to roughen the surface slightly, using mechanical means. The consequence of this is that there is no impairment of age resistance due to degradation of the oxide layers, as may occur on metal surfaces (see Section 7.5.4). It is very likely, therefore, that this approach will open up a new and very large area of applications for bonding technology in future aircraft manufacture. It is also likely that the hot-setting film adhesives which have been used to date will be replaced by two-part, cold-setting adhesive systems.