Damage tolerance of CFRP airframe bolted joints in bearing, following bolt pull-through failure

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Abstract

The experimental study presented herein, investigated the residual strength of bolted joints on Carbon Fiber Reinforced Polymer (CFRP) airframe structures within the context of structural damage tolerance and airworthiness regulations. The damage scenario assumed, subjected a series of bolted joint CFRP laminate specimens to quasi-static bearing loading, following bolt pull-through failure events of different magnitude. Representative CFRP laminate specimens manufactured from AS7/8552 carbon fiber/epoxy matrix system were artificially damaged under bolt pull-through loading, following the herein proposed modifications to the current pull-through ASTM testing procedure. The specimens were subsequently tested in static bearing loading for examining the specimen residual bearing strength. The residual joint bearing strength was related to the displacement travelled passed the initial failure stage in pull-through mode and was measured up to a maximum of a 13% decrease for the tested samples and the maximum damage imposed. The study explored the safe utilization of bolted joints at higher operating loading levels, within the context of the current airworthiness regulations. The inherent damage arrest features of the joints were highlighted. The study concluded with comments and suggestions on the expansion of the current utilization spectrum of damaged bolted joints from pull-through loading in airframe design, bound by the current airworthiness certification requirements.

Introduction

The aerospace vehicle design sector has embraced the application of Fiber Reinforced Polymer materials (FRP) for some decades now. These materials provided the FRP airframe structural designs with enhanced strength and stiffness regarding the most traditional ones made of aluminum alloys and/or other metallic materials. For most designs, a reduction in the structural weight was achieved especially for the structures with defined load paths, benefiting from the directional property tailoring of the unidirectional laminate FRP materials.

The production method for classic aluminum alloy airframe structures is to assemble forged, machine formed and thin sheet of aluminum alloy components together, in forming larger structural assemblies fastened by bolts and rivets. It has been recognized that the fewer the joints in such structures, the stiffer and the lighter the design is, the longer the life of the structure generally and the fewer the problems encountered in service [1,2]. There are nevertheless limitations to the size of the structural components that could be produced as single structural items prior to assembly, for avoiding excessive fastening and riveting.

With the introduction of FRPs, new manufacturing concepts have emerged. The FRP material itself is generated while the structural component is being manufactured, by curing of the composite matrix within or outside oven/autoclave chambers. The idea of generating complete airframes during a single manufacturing stage, started to be quite tempting since the conception in the application of the FRPs in airframe manufacturing. Generally, larger structural parts are generated from FRPs but there are limitations to this single stage production process as well. The current status for the majority of FRP airframe components in the aviation industry that have not been cured or co-cured with their mating part, is to assemble them by the use of bolts. Riveting is not a preferred method for FRP laminate assembly mainly due to the high transversal loading incurred during the fastening/forming process, loading which most of the FRP laminates are relatively intolerant [[1], [3], [4]]. Adhesive bonding can be regarded as another means of assembling, but currently poses a lot of certifications issues especially to the certification of civil aircraft.

Bolted joints on FPRs have been criticized as not being very efficient means of assembling structural components. Joint efficiency is judged by the ratio of the actual loading transferred through the joined structure versus the load transferred had the structure been uninterrupted, that is without the presence of any joining means. In that respect, FRP structures can reach a ratio of 0.4 while similar arrangement in aluminum structures can reach the level of 0.65 [1,4]. On many occasions though, when comparing the weight per unit length of assembled structures for carrying the same amount of loading, FRP bolted structures, as a complete assembly, weight less than a supposed substitute made of aluminum alloys.

Amongst the possible bolted joint failure modes [1,2,4], airframe design favors bearing failure [2,4]. Design recommendations regarding edge distances, fastener spacing, thickness and layup of the laminate, will safeguard the bolted structure from failing under different failure modes than bolt bearing. Bolted joints on CFRP structural elements are not designed in general to receive transverse loading along the laminate thickness. Transverse loading will be a component of the loading eccentricity generated in single lap joints; it can result from manufacturing and assembly tolerances mismatches or defects; it can also be generated from thermal expansion and mismatch between mated components; be the result of non-anticipated structural overloading; finally it could occur from foreign object impact. On many occasions, such failure events have been reported from in service Non Destructive structural Inspections (NDI), hence the question of the residual joint strength and safe joint utilization has arisen.

The concept of “damage tolerant design” has been conceptualized for some decades. This engineering design and sizing approach started to be applied to metallic structure fatigue critical locations, following the realization that a structure cannot be free of defects and regular inspection should occur at scheduled intervals during the structure lifecycle [4]. In structures made of metallic materials, damage tolerance design in essence, is accepting the existence of supposed cracks at critical locations that propagate under cyclic loading. In structures currently made from FRP materials and under current approved design and certification specifications [[4], [5], [6]], damage tolerance design has more of a meaning of meeting the imposed structural performance requirements, with the structure being damaged to a certain extent by containing various damage types, at sizes representative of manufacturing errors and/or damage from usage under the expected service environment.

The aerospace industry has been very careful in the sizing of structures made of FRP materials, making sure that enough reserve strength is always present even in the event of adverse conditions. Currently, the industry is pushing towards exploring the safe utilization boundaries of FRPs with greater confidence, making use of the inherent damage tolerance features of the CFRP materials wherever applicable. Succeeding in understanding and designing for predictable response of FRP structures at their damaged state, will eventually lead to the design of lighter and safer aircraft vehicles.

The experimental study presented herein, aimed at exploring the damage tolerance of bolted joints on FRP laminates, by experimentally measuring the residual strength in bolt bearing loading following an assumed bolt pull-through failure event scenario. The material for specimen manufacture used was unidirectional HexPly pre-impregnated (pre-preg) AS7/8552 carbon fiber/epoxy matrix tape. This material is currently used on a number of civil transport airframes. The study places the damage tolerance scenario within the context of airframe design and civil airworthiness certification regulations. It aimed at investigating the possible expansion of the current bolted joint utilization spectrum, by showcasing the results of the current study. It also proposed the application of a modified testing procedure for exploring the damage tolerance characteristics of bolted joint on FRP laminates, under the specified damage sequence scenario.

Section snippets

Experimental procedures

The aim of this part of the study was to artificially inflict representative damage resulting from bolt pull-through loading at various loading and damage levels. This loading, which is along the thickness of the laminated CFRP plate, would then be removed, to give its place to a subsequent bearing tensile load along the specimen major longitudinal direction. This two-step experimental process is sketched in Fig. 1. The crosshatched area underneath the bolt depicted in Fig. 1a, signifies the

Experimental testing

For meeting the requirements of the study and the quick turnaround time required for drafting the joint structural response at a preliminary stage, a total of 16 specimens were manufactured and tested. Specimen material was AS7/8552 HexPly® prepreg unidirectional tape, with nominal laminate thickness of 0.145 mm. Twenty four layers quasi-isotropic layup configuration [45 90 -45 0]3s was used, leading to specimen thickness of approximately 3.5 mm. The bearing loading testing procedure ASTM D5961

Results discussion within the context of airworthiness and safety

Civil and military aircraft airworthiness certification specifications form a part in the means of controlling and ensuring the safety of flight. Within such specifications, as for large civil aircraft for example [5,6], the expected performance criteria for airframe structures are captured. According to the specifications clauses relevant to the investigation, airframe structures must (*):

  • Function properly under the application of the maximum service loading applied quasi-statically without

Conclusions

A modified testing procedure was proposed for investigating the static bearing strength reduction of FRP laminate bolted joints that have prior been damaged by a pull through event or by events generating similar types of damage caused by transverse to the laminate thickness loading. It was found that for the material and the specimen design used and for the level of damage inflicted on the specimens, the reduction in the joint bearing strength was up to 13% from its original undamaged state.

CRediT authorship contribution statement

Ioannis K. Giannopoulos: Writing - original draft, Conceptualization, Formal analysis, Investigation, Methodology, Project administration, Resources, Supervision, Validation, Visualization, Writing - review & editing. Kaelan Grafton: Writing - original draft, Data curation, Formal analysis, Visualization. Shijun Guo: Supervision. Howard Smith: Supervision.

Declaration of competing interest

The authors declare that they have no known competing financial interests or personal relationships that could have appeared to influence the work reported in this paper.

References (19)

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