design aircraft structures and select materials with high resistance against cracking;
Related terms:
A life prediction method for aircraft structure based on enveloping life surface
Y.T. He, ... C.F. Li, in Recent Advances in Structural Integrity Analysis - Proceedings of the International Congress (APCF/SIF-2014), 2014
2 CONCEPT OF AIRCRAFT STRUCTURE ENVELOPING LIFE SURFACE
ASELS describes the safe and reliable life scope for aircraft structures in service. It reflects the interrelationships between stress level (S, in MPa), fatigue life (Nf, in flight hours), and calendar life (Ny, in years). When an aircraft is used so heavily that it exceeds the limits of ASELS, the structural state is considered to be unsafe. As shown in Fig.1, the ASELS can be considered as an extension of the ASELC (Sight B) along the coordinate direction of stress level or an extension of the S-N curve (Sight A) along the coordinate direction of calendar life.
Fig.1. Diagram of an aircraft structure enveloping life surface
Fig.1 shows the ASELS in a typical environment. This can be used to predict the structural residual life of an aircraft grounded in the environment. If an aircraft is grounded in different environments in full life circle (the corrosion properties relevant for high altitudes can be ignored), different ASELSs corresponding to these environments should be used for the prediction of residual life. In Fig.1, both directions of the abscissa axis (Ny) are positive; they are calendar lives under different states of protective coating. The abscissa of Tp1/Tp2 is the effective period of protective coating; structures can be considered to be suffering from pure fatigue damage within this period. The points Np1 and Np2 represent safety lives with high reliability under the stress levels S1 and S2, respectively. These parameters can be obtained through component or full-scale fatigue testing and through reliability analyses. The surface Np1-D1-D2-Np2 reflects change laws of the fatigue lives of structures under different stress levels in the corrosive environment without protective coating. The surface D1-Nc1-Nc2-D2 is a boundary limit designed to prevent unexpected fracture of a structure due to corrosion fatigue damage.
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https://www.sciencedirect.com/science/article/pii/B9780081002032500474
Morphing Technology for Advanced Future Commercial Aircrafts
Miguel Á. Castillo Acero, ... Yasser Essa, in Morphing Wing Technologies, 2018
2.5 Conclusions
SARISTU project proved the economic advantages of implementing morphing structures in-flight. These systems, in fact, ensured an appreciable increase in terms of the overall aircraft efficiency, although the increases of weight and complexity were not negligible. In particular, the manufacture of hundreds of pieces is not sustainable from an industrial point of view and should be strongly limited before such engineering solutions can be actually implemented. The simplification of the proposed devices may be then considered a main driver for future studies. Analogously, the use of lighter and better workable materials, specifically suited for these applications, can be another remarkable field of investigation.
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https://www.sciencedirect.com/science/article/pii/B9780081009642000198
Tailor made blanks for the aerospace industry
J. Sinke, ... R. Benedictus, in Tailor Welded Blanks for Advanced Manufacturing, 2011
Manufacturing principles
Aircraft structures are assembled from many parts (order of 1000 to 10 000 – fasteners not included), which are made from various materials like composites, metal alloys and hybrid materials. A wide range of different production processes are used in order to manufacture these parts. Since TMBs are metallic, further discussions are focused on the manufacturing principles for metals. Metal structures are assembled from sheet metal parts that have been cut and formed into the desired shape. The forming processes are often universal; much of the applied tooling is not product related. These tools are applied to a variety of different products and are very suitable for limited product series. The universal processes perfectly match the needs of the aircraft industry, where the diversity in parts is huge and the production quantities are low (in the order of 1000).
Aircraft structures are also characterized by variation in materials and thicknesses. Each area is dimensioned by several load cases, which result in different materials and material conditions, and a specific distribution of thicknesses over the entire structure. Changes from one material to another, or from one thickness to another, require separate parts that have to be either joined or made as integral parts. The first option is often used for metal structures, which results in joints that add weight, time and costs during assembly. The option of integral parts, often applied to composites, has advantages with respect to weight, assembly time and costs, but is more expensive and often more costly to maintain.
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https://www.sciencedirect.com/science/article/pii/B9781845697044500083
Numerical modelling of impact and damage tolerance in aerospace composite structures
A.F. Johnson, ... D. Schueler, in Numerical Modelling of Failure in Advanced Composite Materials, 2015
18.1 Introduction
Aircraft structures are vulnerable to impact damage resulting from impact by hard or soft bodies, such as steel fragments, birds, burst tyre rubber or hail. With the increasing use of polymer composite materials in aircraft structures, the impact response of such materials has been a subject of extensive research that has led to the development of theoretical models able to capture the impact mechanics of materials and structures, required for the derivation of reliable design rules needed by the aircraft industry (Abrate, 1998; Davies and Olsson, 2004). The use of polymer composite materials in commercial aircraft was first limited to secondary structures such as inspection panels, spoilers or air brakes that do not reduce aircraft structural integrity on failure. Improvements in manufacturing technology, the development of innovative material systems and a better understanding of their mechanical behaviour have now led to the use of composites for primary aircraft structures that carry flight, pressurisation or ground loads, which are critical for structural integrity. For primary load-bearing structures, such as the fuselage or wing skin with related safety regulations (FAA AC 23-13A, 2005), validated numerical methods are required to determine impact behaviour. These enable cost savings through ‘right-first-time’ design and reduction of experimental tests in the certification process by use of computational analysis, which was a major objective of the industry-led EU project MAAXIMUS (2013).
Impact incidents are commonly classified according to impact velocity: low velocity, high velocity and ballistic impact (Davies and Zhang, 1995; Olsson, 2000). Under low-velocity impact (LVI), an impacted plate deforms according to plate theory as the impact duration is longer than the time needed for flexural waves to reach the boundaries. For shorter impact durations under high-velocity impact (HVI), the target response is governed by flexural waves. Ballistic impact is characterised by a local impact response from through-thickness dilatational stress waves. Typical scenarios for LVI are ‘tool drop’ where the impactor hits the target accelerated by gravity from up to a few metres in height and the impact of ground equipment, such as stairs or deicing equipment. HVI scenarios usually occur in flight or on take-off and include bird strikes or impact by hail, tyre rubber, engine fragments or runway debris. Ballistic impact cases are low mass with often supersonic impact velocities arising from weapons with application to security protection systems.
In this chapter, the focus is on carbon fibre composite aircraft structures under HVI that are commonly classified further into soft body impact (hail, birds, tyre rubber) and hard body impact (runway debris, engine parts). Soft bodies may disintegrate on impact exhibiting a fluid-like flow behaviour or be highly deformable (rubber), whereas hard bodies usually remain intact after impact. For aircraft structures, the impact scenarios for different categories of projectiles are specified in safety regulations, or are derived from test programmes. Bird strike regulations require the airplane to safely continue its flight after impact with a 4-lb bird (8lb on the empennage) at design cruising speed VC (EASA CS25, FAA 14 CFR 25: §§ 25.571 and 25.631). For hailstones diameters range from 5 to 100mm with velocities for hail on ground 10–50m/s and in flight at VC, see Field et al. (2009). In the case of uncontained engine failure, metal fragments can impact the fuselage or wing with velocities in the range 61–295m/s and masses ranging from 22g to 20.4kg (DOT/FAA/AR-04/16, 2004). Foreign object debris (FOD) such as runway debris, stones or metal fragments could impact aircraft when launched by a tyre or by jet blast from another aircraft, as discussed in Chadwick et al. (2001). Burst tyre fragments may impact the lower wing skin or fuselage with impact velocities dependent on the tangential tyre speed and typically about 100m/s, impactor mass can be up to several kg (Mines et al., 2007; Toso-Pentecôte et al., 2010).
The FAA route to certification adopted by civil aircraft manufacturers is based on the well-known test pyramid for aircraft structures, as set out for composite aircraft in FAA AC 20-107B (2010), which foresees five levels of tests from material characterisation test specimens, through structural elements of increasing complexity up to full aircraft structures. It is clear from the range of impact threats and the size of the regulation test programmes that certification of composite structures subjected to impact loads is no longer feasible by extensive test campaigns alone. The way forward as discussed by Hachenberg (2002) is to use design analysis to support a limited number of tests at each level of the test pyramid. This requires the development and validation of computational methods to support the design of a composite aircraft under the full range of flight and service loads defined in the airworthiness specifications. This is a ‘building block’ approach, as each level strongly relies on the validity of design analysis results obtained on simpler structures at the levels below. For the structural impact behaviour, improved composites damage and failure models with appropriate finite element (FE) codes are needed and should be validated by structural impact tests at each level of the test pyramid.
This chapter presents HVI test data from gas gun impact tests on advanced composite structures and discusses numerical methods to predict observed impact damage at the structural element level of the test pyramid, which are validated by the tests. Section 18.2 presents meso-scale composite ply damage and failure models and energy-based delamination models suitable for use in explicit FE codes for prediction of impact damage. The application of the FE models in Section 18.3 is used to predict tyre rubber impact damage on a rib-stiffened composite panel structure. Section 18.4 discusses a test programme on the influence of tensile and compressive pre-stress on HVI damage in composite plate structures, followed by the extension of the FE modelling procedures to pre-stressed panels under impact and prediction of residual strengths and damage tolerance (DT). Concluding remarks in Section 18.5 are followed by future trends, sources of further information and references.
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https://www.sciencedirect.com/science/article/pii/B9780081003329000189
Wing problems
T.H.G. Megson, in Aircraft Structures for Engineering Students (Fifth Edition), 2013
Publisher Summary
Aircraft structures, being extremely flexible, are prone to distortion under load. When these loads are caused by aerodynamic forces, which themselves depend on the geometry of the structure and the orientation of the various structural components to the surrounding airflow, structural distortion results in changes in aerodynamic load, leading to further distortion and so on. The interaction of aerodynamic and elastic forces is known as aeroelasticity. Two distinct types of aeroelastic problem occur. One involves the interaction of aerodynamic and elastic forces of the type described above. Such interactions may exhibit divergent tendencies in a too flexible structure, leading to failure or, in an adequately stiff structure, converge until a condition of stable equilibrium is reached. In this type of problem, static or steady state systems of aerodynamic and elastic forces produce such aeroelastic phenomena as divergence and control reversal. The second class of problem involves the inertia of the structure, as well as aerodynamic and elastic forces. Redistribution of aerodynamic loads and divergence are closely related aeroelastic phenomena; they are, therefore, simultaneously considered. The flexibility of the major aerodynamic surfaces (wings, vertical and horizontal tails) adversely affects the effectiveness of the corresponding control surfaces (ailerons, rudder, and elevators).
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https://www.sciencedirect.com/science/article/pii/B9780080969053000371
Introduction and overview
Chun H. Wang, Cong N. Duong, in Bonded Joints and Repairs to Composite Airframe Structures, 2016
1.2 Criticality of Structure and Damage
Aircraft structures are generally classified as follows in terms of criticality of the structure:
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critical structure, whose integrity is essential in maintaining the overall flight safety of the aircraft (e.g., principal structural elements in transport category aircraft);
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primary structure carries flight, ground, or pressurization loads, and whose failure would reduce the aircraft’s structural integrity;
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secondary structure that, if it was to fail, would affect the operation of the aircraft but not lead to its loss; and
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tertiary structure, in which failure would not significantly affect operation of the aircraft.
Inspection, damage assessment, and repair requirements differ significantly between these classifications. However, even within a single component, the allowable damage type and size (and consequently acceptable repair actions) will vary according to the criticality of the damaged region. The original equipment manufacturer (OEM) generally zones an aircraft component in terms of these regions, and specifies repair limits and the pertinent repair procedures in the structural repair manual (SRM). Damages outside the scope of the SRM, particularly to critical regions of primary structure, require engineering design disposition and approval by the OEM (or its delegate); this book describes some new design options demonstrated by recent research results.
Foreign object impact is usually the main type of damage concerning composite aircraft structures. To ensure continuing airworthiness, it is necessary to identify damage severity and detectability as part of the ongoing maintenance process. Current airworthiness regulations (FAA, 2010) classify various damage types into five categories, as indicated in Figure 1.2 that illustrates the relationship between design strength and damage size:
Figure 1.2. Allowable strength versus damage size.
- •
Category 1: Allowable damage or allowable manufacturing defects that do not degrade structural integrity, and hence may go undetected by scheduled inspections. Structures containing this type of damage are capable of sustaining the ultimate load for the life of the aircraft structure. Some examples include barely visible impact damage (BVID), small delamination, porosity, small scratches, and so forth. No repairs are needed.
- •
Category 2: Damage that can be reliably detected at scheduled inspection intervals. This type of damage should not grow or, if slow or arrested growth occurs, the residual strength of the damaged structure during the inspection internal is sufficiently above the limit load capability. Some examples include visible impact damage, deep gouges or debonding, and major local overheating damage. Repairs are needed to restore the design ultimate load capability.
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Category 3: Damage that can be readily detected, within a few flights, by operations or maintenance personnel without special skills in composite inspection. The structure can still maintain limit or near limit load capability. Repairs are required immediately to restore design ultimate load capability.
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Category 4: Discrete source damage that will reduce the structural strength to below the design limit load such that flight maneuvers become limited (i.e., structure can maintain safe flight at reduced levels). Examples include rotor burst, bird strikes, tire burst, and severe in-flight hail. Repairs are needed immediately after flight.
- •
Category 5: Severe damage outside design but is self-evident and known to operations, such as anomalous ground collision with service vehicles, flight overload conditions, abnormally hard landings, and so forth. Immediate repair is required.
Analytical methods for assessing the residual strength of damaged composite components are needed to ensure that only necessarily required repairs are undertaken. Essentially, one of the following decisions must be made:
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No repair action—damage is negligible.
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Only needed correction is cosmetic or sealing repair because damage is minor.
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Structural repair is required (if feasible) because strength is reduced below ultimate design allowable, or has the potential to be reduced in subsequent service.
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Replacement is required as repair is not economically or technically feasible and component must be replaced.
For BVID, quite large areas of damage (typically 25mm diameter) can be tolerated for older generation carbon/epoxy systems (and brittle high-temperature systems) without failures occurring below the ultimate design strain allowable, generally around 5000 microstrain for quasi-isotropic laminates made of unidirectional (tape) lamina. Recently, advanced computational modeling techniques have been shown to be able to accurately predict the residual strength of composite laminates containing holes of various sizes and shapes (Wang et al., 2011a; Ridha et al., 2014). Thus, the residual strength assessment of a structure following impact damage can be performed similarly by using these advanced computational methods.
Fatigue studies have also shown that BVID will not grow under realistic cyclic strain levels for typical carbon/epoxy laminates. This is an important point because BVID will often not be detected until a 100% nondestructive inspection is undertaken. Even though there is a possibility of damage growth and residual strength degradation under hygrothermal cycling conditions, this appears to be a serious concern only under severe cycling conditions. This possibly catastrophic flaw growth under severe hygrothermal cycling may result from expansion of entrapped moisture due to freezing or steam formation on heating during supersonic flight.
For safety-critical structures, coupons, structural details, elements, and subcomponents are required to be tested under fatigue loading to determine the sensitivity of structure to damage growth and to demonstrate their compliance with either no-growth or slow-growth requirements. This is to ensure that a damaged structure should not be exposed to an excessive period of time when its residual strength is less than the ultimate. Once the damage (greater than the allowable damage size under category 1) is detected, the component is either repaired to restore ultimate load capability or replaced.
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Nondestructive inspection and structural health monitoring of aerospace materials
In Introduction to Aerospace Materials, 2012
23.4 Summary
Aircraft structures and engine components must be nondestructively inspected after manufacturing and throughout their operational life for the presence of defects and damage. Most inspections are currently performed using NDT methods such as ultrasonics, radiography and thermography. Structural health monitoring (SHM) is emerging as an alternative to conventional NDI, in which sensor systems are used with little or no human invention to monitor aircraft for damage.
NDI methods have the capability to detect certain (but not all) types of damage in metals and composites. Ultrasonics, thermography and eddy current inspections are capable of detecting damage and cracks aligned parallel with the material surface whereas radiography is better suited to detecting cracks normal to the surface. It is often necessary to use two or more inspection methods to obtain a complete description of the type, amount and location of the damage.
Some NDI techniques can be used to inspect metals but not fibre–polymer composites. Of the NDI methods described in this chapter, damage in composites is difficult to detect using eddy current and magnetic particle owing to their low electromagnetic properties, and using liquid dye penetrant because most damage is internal (e.g. delaminations) and does not break the surface.
SHM has the potential to reduce aircraft downtime for routine inspections and reduce design safety factors for damage tolerance because of the early detection of damage. It is often only necessary to locate SHM sensors in components prone to damage (e.g. heavily-loaded parts, parts susceptible to impact damage), rather than covering the entire aircraft with a complex, integrated sensor network system.
SHM techniques are classified as local or global (wide-area). Examples of local health monitoring include Bragg grating optical fibre sensors and comparative vacuum monitoring, whereas wide-area monitoring techniques are acoustic emission and acousto-ultrasonics.
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https://www.sciencedirect.com/science/article/pii/B9781855739468500236
Corrosion prediction in the aerospace industry
J. Ullett, in Corrosion Control in the Aerospace Industry, 2009
6.2.3 Surface finishes and coatings
Aluminum aircraft structure undergoes one or more surface treatments to prevent the onset of corrosion. These treatments include: anodization, chromate conversion coating, and primer coating. Internal structure that is safety critical (e.g., wing box) or that is subjected to harsh environments (e.g., stone-spray from landing and take-offs) may receive a topcoat of glossy polyurethane in addition to a primer coating. The outer mold-line of military aircraft is typically coated with a low-gloss topcoat. These highly-filled topcoats are more prone to breakdown of barrier properties and do not provide the same degree of corrosion protection as do glossy topcoats used on commercial airliners.
The corrosion of the aluminum substrate will not occur until the protective coatings are compromised. Unfortunately, local defects in the coatings system may occur soon after depot repaint or field touch-up due to removal of access panels and other routine maintenance activities. Differences in modulus and thermal expansion coefficients between aluminum structure and steel or titanium fasteners are another cause of localized protective-coating failure. Thus, while the coating system on-the-whole may provide excellent barrier properties for decades (particularly for interior structure), many opportunities exist for localized mechanical or chemical degradation of the protective coatings.
URL:
https://www.sciencedirect.com/science/article/pii/B9781845693459500062
Sustainable bio composites for aircraft components
Naveen Jesu Arockiam, ... Naheed Saba, in Sustainable Composites for Aerospace Applications, 2018
6.4.2 Bio composites for aircraft wing boxes
In aircraft structures the reduction of structural weight by using high performance fibers has evolved, but the existing carbon fiber-reinforced polymer composites are nonrenewable. Boegler et al. have utilized natural fiber composites for load-bearing structures. They have made a model of civil transport aircraft (Airbus A320-200). The wing box weight has been calculated by using the material properties of the natural fiber composites, and they have compared the results with a reference wing box made of an aluminum alloy. While using ramie fiber composites, it was found that the decrease in weight was around 12%–14% [9].
In the first step, the mechanical properties, such as Young’s modulus (E), yield strength (σy), and density (ρ) of raw natural fibers were calculated. After initial screening, hemp, sisal, flax, and remie fibers were selected. Fibers from hemp, sisal, and flax are frequently used in automobile interiors. Hemie fiber has been chosen because of its high Young’s modulus and yield strength. Table 6.2 shows the mechanical properties of the selected fiber and matrix.
Table 6.2. Characteristics of materials used in wing box
| Material | Density (g/cm3) | Yield strength (MPa) | Young’s modulus (GPa) |
|---|---|---|---|
| Polylactic acid | 1.21–1.25 | 48–60 | 3.45–3.83 |
| Epoxy resin | 1.1–1.4 | 36–71 | 2.35–3.08 |
| Ramie | 1.5 | 500 | 44 |
| Flax | 1.4 | 800–1500 | 60–80 |
| Hemp | 1.48 | 550–900 | 70 |
| Sisal | 1.33 | 600–700 | 38 |
Epoxy resin is considered for calculation, but it is not biodegradable. In order to make a 100% renewable and biodegradable composite material, poly lactic acid (PLA) is considered as a polymer matrix. For the wing mass calculation, the following parameters are essential such as density (ρ), Young’s modulus (E), yield strength (σy), and shear modulus (G). Further, the assumption is made that the material is isotropic (Table 6.3). According to the rule of mixture ρ, E, and σy can be calculated by considering individual fiber and matrix.
Table 6.3. Input parameters required for wing box mass calculation
| Material | Density (kg/m2) | Young’s modulus (GPa) | Shear modulus (GPa) | Yield strength (MPa) |
|---|---|---|---|---|
| Hemp/polylactic acid | 1366 | 14.15 | 8.98 | 85.80 |
| Flax/polylactic acid | 1336 | 18.05 | 11.45 | 73.71 |
| Sisal/polylactic acid | 1324 | 5.37 | 3.41 | 146.45 |
| Ramie/polylactic acid | 1354 | 9.08 | 5.76 | 127.14 |
| Hemp/epoxy resin | 1326 | 14.05 | 8.92 | 87.23 |
| Flax/epoxy resin | 1296 | 17.95 | 11.39 | 75.14 |
| Sisal/epoxy resin | 1284 | 5.28 | 3.35 | 147.88 |
| Ramie/epoxy resin | 1314 | 8.98 | 5.70 | 128.57 |
Rule of mixture
where X can be any property (ρ, E, σy); Vf, volume fraction of the fiber; Vm, volume fraction of the matrix; f, fiber; m, matrix.
The Shear modulus (G) can be derived from the following equation.
where Poisson’s ratio can be assumed to be 0.33.
Table 6.4 shows the calculated ρ, E, σy, and G values.
Table 6.4. Calculated wing mass of different composites with respect to aluminum
| Material | Wing mass (m) (kg) |
|---|---|
| Aluminum 7000 | 8829 |
| Hemp/epoxy resin | 10503 |
| Flax/epoxy resin | 11355 |
| Sisal/epoxy resin | – |
| Ramie/epoxy resin | 7576 |
| Hemp/polylactic acid | 10815 |
| Flax/polylactic acid | 11784 |
| Sisal/polylactic acid | – |
| Ramie/polylactic acid | 7758 |
A Matlab program has been developed by using the vortex lattice method for the estimation of wing mass, namely the aerodynamic code of TORNADO to calculate the aerodynamic lift forces on a given wing configuration.
Table 6.4 shows the calculated wing masses for different natural fiber composites compared with a reference wing box aluminum alloy of the 7000 series. For the sisal fiber composites, the model could not finish the finite element iteration that can be owed to the lowest value of E. Hemp and flax fiber-based composites drastically increase the weight of the wing box, whereas Ramie fiber-based composites show the reduction in weight of the wing box without compromising structural integrity [9].
URL:
https://www.sciencedirect.com/science/article/pii/B9780081021316000062
Repair Tolerance for Composite Structures Using Probabilistic Methodologies
He Ren, ... Yong Chen, in Reliability Based Aircraft Maintenance Optimization and Applications, 2017
8.4.1 Load Case
For civil aircraft structures, gust load is mainly considered as the critical load case [73]:
where ɛ is the actual load, ɛLL is the limit load, and ɛUL is the ultimate load, ɛUL=1.5ɛLL.
Since the probability value under different exceeding conditions changes significantly by the power of 10, a log-linear model is used to describe the load occurrence probability. The load exceedance curve is shown in Fig.8.7.
Figure 8.7. Load exceedance curve.
(8.4)
URL:
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