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The assessment of the service durability of aerospace components and assemblies has become an important segment of design. In order to meet strict safety requirements, a number of complex and long experiments are carried out. The use of finite element method (FEM) and extended finite element method (XFEM) for the estimation of fatigue life and fatigue crack growth predictions has been proved as a good alternative to the expensive experimental methods. In this paper, both experimental and numerical analyses of 2024-T3 aluminum spar of a light aircraft under variable amplitude loading are presented. FEM has been used for estimation of the spar life to crack initiation, whereas XFEM has been used for fatigue crack growth predictions and fatigue life estimation of damaged spar. The values of stress intensity factors were extracted from the XFEM solution in MorfeoCrack for Abaqus software.

Metal matrix composites (MMCs) exhibit superior characteristics such as low weight, high stiffness, and high mechanical and physical properties. Inheriting such an outstanding combination of specifications, they are nowadays considered as the promising materials in the aerospace and biomedical industries. However, the presence of high abrasive reinforcing particles in MMCs leads to severe manufacturing issues. Due to the tool-particle interactions which occur during the machining of MMCs, high tool wear and poor surface finish are induced and those elements are considered as the main drawbacks of cutting MMCs. In this study, dry turning experiments were conducted for two different inserts and coated carbide on a bar of titanium metal matrix composite (Ti-MMC). Semi-finishing machining is operated with cutting parameters based on the tool supplier's recommendations which were not fully optimized. The maximum flank wear length (VBBmax) was selected as the tool wear criteria.

We applied ultrasonic propagation imaging (UPI) system for rapid and reliable quality control of fuel tanks for a space launcher. The fuel tank is an aluminum-lined CFRP propellant tank. The UPI system uses Q-switched laser (QL) to generate ultrasonic wave on the test specimen, and laser mirror scanner (LMS) to control the laser impinging point that scans the area of interest with high speed. Each ultrasonic wave generated by laser impinging was received by a piezoelectric sensor with coordinate information of the scanned area. After ultrasonic propagation image processing, results with impact damage and manufacturing defect information of the fuel tank were presented.

Implementing Design for Environment (DfE) into the design process requires a strategic integration. Furthermore, as DfE is continuously evolving, flexible processes need to be implemented. This article focuses on the integration of DfE into an optimization framework with the objective of influencing next-generation aircraft. For this purpose, DfE and Structures groups are developing together a set of new environmental indicators covering all life cycle stages of the product by coupling a list of yes/no questions with an environmental matrix. The following indicators are calculated: Regulation risk, Impact of manufacturing the part, CO2 emissions and Recyclability potential. These indicators will be used as constraints in the multi-disciplinary design optimization (MDO) framework, meaning that the structure will be designed while complying with environmental targets and anticipating future regulation changes.

New Refill Friction Stir Spot Welding (Refill FSSW) was performed for the aluminum alloys used for aircrafts. Joints were examined to evaluate its performances, such as tensile shear strength and strength of heat affected zone, with using various surface conditions. As a result of examination, significant lap shear strength by using shoulder plunging method was achieved with shorter process time. Even with the alodine coated aluminum alloys, strong and good quality joints can be made. And finally, actual prototype components were experimentally produced with Refill FSSW robot. Fine preproduction parts with no distortion and sufficient quality were obtained. Refill FSSW will be able to be regarded as suitable for fabrication of aircraft components.

Since the 1960's, lightning protection of aircraft has been an important design aspect, a concern for the flying public, aircraft manufacturers and the Federal Aviation Administration (FAA). With the implementation of major aircraft structures fabricated from carbon fiber reinforced plastic (CFRP) materials, lightning protection has become a more complicated issue to solve. One widely used material for lightning strike protection of CFRP structures within the aerospace industry is expanded metal foil (EMF). EMF is currently used in both military and commercial passenger aircraft. An issue that has historically been an area of concern with EMF is micro cracking of paint on the composite structure which can result in corrosion of the metal foil and subsequent loss of conductivity. This paper addresses the issues of stress and displacement in the composite structure layup which contribute to paint cracking caused by aircraft thermal cycling.

This technical paper deals with design and manufacture of axial and orbital cutters for drilling large diameters holes in Carbon Fiber Reinforced Plastics (CFRP) / Titanium (TA6V) thick stack by means of an Automated Drilling Unit fixed on a drilling template. Creating tools that drill such stacked holes in a single operation is particularly difficult. The common strategies for tool designs designated to cutting composites stacked with metals as titanium include uncoated carbide, tools with a diamond coating applied by chemical vapor deposition (CVD). It is also question of what kind of drilling process should we use to achieve larger holes in minimum time. Therefore, axial drilling process with pilot, drill and ream steps find a competitor drilling process named orbital drilling which can achieve both operations in one step allowing then, burrless, free delamination, small and easy removable chips all with one tool being able to achieve different hole sizes.

Manufacturing operations introduce unreliability into hardware that is not ordinarily accounted for by reliability design engineering efforts. Inspections and test procedures normally interwoven into fabrication processes are imperfect, and allow defects to escape which later result in field failures. Therefore, if the reliability that is designed and developed into an equipment/system is to be achieved, efforts must be applied during production to insure that reliability is built into the hardware. There are various ways to improve the reliability of a product. These include: Simplification Stress reduction/strength enhancement Design Improvement Using higher quality components Environmental Stress Screening before shipment Process Improvements, etc. This paper concentrates on ‘Manufacturing Process Improvement’ effort through the use of design of experiments, (DOE). Hence, improved levels of reliability can be achieved.

Bombardier faced new manufacturing process challenges drilling and fastening CSeries* aircraft panels with multi-material stacks of composite (CFRP), titanium and aluminum in which Gemcor responded with a unique, flexible CNC Drivmatic® automatic fastening system, now in production at Bombardier. This joint technical paper is presented by Bombardier, expounding on manufacturing process challenges with the C Series aircraft design requirements and Gemcor presenting a unique solution to automatically fasten CFRP aft fuselage panels and aluminum lithium (Al Li) cockpit panels with the same CNC Drivmatic® system. After installation and preliminary acceptance at Bombardier, the CNC system was further enhanced to automatically fasten the carbon fiber pressure bulkhead dome assembly.

A frame-clip riveting end effector has been developed for installing 3.97mm (5/32) and 4.6mm (3/16) universal head aluminum rivets. The end effector can be mounted on the end of a robot arm. The end effector provides 35.6 kNt (8000 lbs) of rivet upset. Rivets can be installed fifteen millimeters from the IML. The clearance allowed to rivet centerline is 150 millimeters. The riveting process features a unique style of rivet fingers for the universal head rivet. These fingers allow the rivet to be brought in with the ram. This differentiates from some styles of frame-clip end effectors in which the rivet is blown into the hole. The paper shows the technical components of the end effector in sequence: the pneumatic clamp, rivet insert and upset. The end effector will be used for riveting shear ties to frames on the IML of fuselage panels.

Airbus commercial wings are assembled manually in dedicated steel structures. The lead time to design, manufacture and commission these fixtures is often in excess of 24 months. Due to the nature of these fixtures, manufacturing is slow in responding to changes in demand. There is underused capacity in some areas and insufficient ramp-up speed where increased production rate is needed. Reconfigurable Manufacturing Systems and Reconfigurable Assembly Systems (RAS) provide an approach to system design that provides appropriate capacity when needed. The aim of the paper is to review RAS technologies that are suitable for cost-effective wing structure assembly and what knowledge gaps exist for a RAS to be achieved. The paper examines successful cases of RAS and reviews relevant system design approaches. Cost savings are acknowledged and tabularised where demonstrated in research. The research gaps to realising a RAS for wing assembly are identified and different approaches are considered.

Ever since the advent of fiber reinforced polymer materials in the field of Aerospace, Metal-FRP stacks started to gain importance due to their superior fatigue performance, phenomenal low weight and good specific strength. However the machining, specially drilling these multi stack materials has always proved to be a challenge for the field of manufacturing and assembly. Drilling holes in only metal with a drill (metal drill), the material removal is through a process of clear shearing since the tool is much harder and sharper than the base material. The tools hence wears at a much slower and gradual rate, also the malleable properties of the machined metal compensate to the reduced cutting capability of the worn out drill. These properties of the machined metal act like a ‘FAIL SAFE’ mechanism during the machining process assuring a trouble free fail safe environment during the drilling process. However, drilling FRP composites is altogether a different story.

Winter maintenance is based on the intervention of operating services, as well as the use of deicers. Each year, in France, thousands of tons of deicers are spread through runways and taxiways. On the airport sector, the main deicers are sodium or potassium acetates and formates. All these deicers aim to prevent ice formation (preventive strategy) and/or improve the ice melting of snow residual film (curative strategy) at temperatures below 0°C. The operating principle of these compounds is based on the lowering of the solution's freezing point once dissolved in water. The phase diagram's knowledge is predominant to determine the deicer's amount to be applied on the surface. It provides a way to optimize their amounts applied with respect to weather conditions, present or forecasted. The Center for Technical Studies of Equipment in East of France (CETE de l'Est) developed and implemented a method based on Raman spectroscopy to characterize aqueous solutions of airport de-icers.

In order to guarantee systems immunity, lightning induced electromagnetic energy has to be lower than the system's susceptibility threshold. This can be achieved, if the aircraft structure provides a good protection against lightning current as well as against its electromagnetic induced field. Moreover such a structure is also required to constitute a ground plane that guarantees very low common mode impedance between all grounded systems in order to keep them at the same electrical potential. The interaction of lightning with aircraft structure, and the coupling of induced energy with harnesses and systems inside the airframe, is a complex phenomenon, mainly for composite aircraft. Composite structures are either not conductive at all (e.g., fiberglass) or are significantly less conductive than metals (e.g., carbon fiber).

The interference-fit processing and split-sleeve cold expansion processing were employed on the 7050-T7451 aluminum alloy joints and fatigue test of joints were carried out. The effects of the processing on fatigue life were compared and characteristics of fatigue fracture appearance were studied by scanning electron microscope. The results show that the interference-fit bolted joints obtain longer fatigue life with interference of 0.08∼0.14mm, the longest with 0.11mm, and the fatigue life of split-sleeve cold expanded bolted joints increase with hole cold expansion of 3-5%'s growth. The maximum fatigue life is increased to 3 times that of the non- cold worked joints. The fatigue crack fractography present the characteristics of mixed transgranular fracture and material of 7050-T7451 has toughness. Interference-fit and split-sleeve cold expansion processing can also strengthen the wall of hole and make fatigue crack source from the surface of structure, not the wall of the hole.

In a fastening system when there is a small misalignment of the holes, the holes are enlarged to align the axes and a next size fastener is used to fit the joint. But when the misalignment is large then the enlargement need to be proportionally large. In this case a bushing is press fit onto the hole to handle the fastening. If we press fit a bushing, it generates residual stresses in the panel. These residual stresses reduce the damage life of the components on which the bushings were press fit. In the aircraft engine nacelle components the damage life is very critical in various failure conditions such as fan blade out condition, wind milling and bird strike. It increases the flight time in these events. Here four different case studies were considered to study the damage life of the aircraft components made of Aluminum or composite material.