Impact behavior of aircraft composite structures

Aircraft composite structures, especially the composite fuselage, are prone to impact damage [55]. Contrary to metal structures, CFRP absorbs impact energy mainly through material damage and elastic deformation, instead of plastic deformation as with metals [56,57]. The nature of the impact damage is numerous for aircraft structures and has been classified according to the impact velocity [58,59]: tool drop during maintenance operations are classified as low velocity (4 to 10) m/s impact events with an impact energy of up to 35 J; runway debris impact has an intermediate velocity and impact energies in the order of 50 J, while bird strike and accidental damage such as the impact of

integrated parts, are classified as having medium velocities of (100 to 150) m/s with impact energies of over 100 J. In accordance with this classification, Airbus has recently reported an impact damage screening on the composite fuselage of A350XWB [47], as shown in Figure 3.2. The low-velocity impact events with an energy of up to 35 J had a higher probability of occurrence throughout the structure, being highly localized around the passenger and cargo doors.

Figure 3.2 Impact damage screening of A350XWB (by permission of Airbus GmbH).

To assess the low-velocity impact damage, drop weight impact testing has been widely used to study the impact behavior of composite structures [5,56,60–63], adhesive bonded composites [64–

66], and metal-composite laminates [67,68]. Some investigation has also been performed on impacted bolted composite structures [14]. Ochôa et al. [14] reported the detection of internal delamination in impacted stringers bolted to CF-PEKK panels of a horizontal stabilizer torsion box by an ultrasonic guided wave monitoring system. However, no investigation of impact damage evolution and residual strength was addressed.

Under drop weight impact testing, the material or joint can present a pure elastic response known as rebounding, or penetration instead, where the impact energy is totally absorbed by the material and released by the damage process, or perforation [62]. Such responses are dependent on the impact test parameters (impact velocity, indenter shape, and impact mass) [48,55,61,67], the temperature of the test [55,61], and material properties [48]. As reported by Gao et al. [48], CF-PEEK manufactured with a fast cooling rate presented the ability to resist damage initiation upon impact, owing to the low level of crystallinity of the PEEK matrix and consequently higher matrix ductility.

According to the EASA CS 25.571 [69] regulations for impact damage threat assessment of large aircraft, low-energy impact damage of composites is categorized as visible impact damage (VID) or barely visible impact damage (BVID). Typically, VID presents a damage area easily detected by a residual dent on the material surface, while BVID is characterized by internal damage, including matrix cracking, fiber fracture under tension and compression, and shear-driven delamination [49]. As reported by Polimeno and Meo [70], the most critical BVID is delamination, owing to its unstable crack propagation, diminished detectability by non-destructive tests, and up to 40 % decrease of the composite strength and stiffness. To define a threshold for BVID, a residual

dent depth of 0.5 mm is frequently used as the criterion, owing to the high measurement accuracy and probability of detection (90 % according to AIRBUS [47]), leading to a reasonable level of robustness for aircraft structural design [5,55,69]. Moreover, for damage tolerant designs for impacts, VID and BVID must be tested under quasi-static and cyclic loading to define their sensitivity to growth and the residual strength compared to allowable values for the composite structure [49].

Vieille, Casado, and Bouvet [5] compared the impact damage resistance and tolerance of CF-PEEK, CF-PPS, and CF-epoxy by the size of the impact damage and post-impact compression behavior. The authors reported that the BVID threshold was reached at 16 J for CF-PEEK, at 13 J for PPS and at 11 J for epoxy, showing a higher resistance to impact damage initiation of CF-PEEK. Additionally, as the toughness of a PEEK matrix is reported to be higher than PPS and epoxy, the propagation of the impact damage is shown to be relatively slow in CF-PEEK and the residual compressive strength 30 % higher than CF-epoxy and 12 % higher than CF-PPS, leading to better damage tolerance. Tai, Yi, and Tseng et al. [71] reported a decrease in the fatigue life of CF-PEEK impacted by an energy of up to 25 J and cyclic loading under a tension-tension regime. From a C-scan inspection, the authors concluded that the delamination induced by the impact did not grow during fatigue life, leading to no apparent stiffness degradation over fatigue cycles, and final failure only due to the fatigue damage.

With the presence of a metal interlayer in CFRP, additional mechanisms to release the impact energy have been reported [67,68]. Aside from delamination of fiber-matrix interface, adhesion failure of the metal-composite interface along with plastic deformation of the metal layer seems to increase the absorption of impact energy considerably in comparison to CFRP, and so represents a promising strategy for improving the impact resistance of composites [67,68].

Although a finite element model (FEM) has been used to simulate the dynamic response of impacted aircraft composite structures and hence to predict critical impact scenarios and maintenance strategies [17,72,73], there are also simple analytical models that provide valuable estimations on the occurrence of critical delamination in impacted composites. Such models are mainly used in the screening stage of aircraft design [49]. Davis and Robinson [58] have shown that a single central delamination in the mid-plane of a composite plate propagates with an in-plane shear mode (fracture mode II) upon reaching a critical force. Based on these assumptions, the authors proposed an equation for the critical impact force, which is independent of the delamination size and plate’s boundary conditions, as well as which neglects bending strains upon the impact test [74,75]. The accuracy of the equation has been proven for a large number of plate geometries and CFRP [49] and is used in this PhD to predict the occurrence of delamination with impacted friction riveted joints. These results will be discussed in Appendix J.