10 Aerospace Materials

Introduction

Early airplanes were made of simple, glued wood and cotton fabric-covered structures. While they could be constructed with simple tools, they were not always strong and required considerable maintenance. Advances in construction materials and assembly processes led to much more durable stressed skin designs made of lightweight aluminum alloys, which have been the mainstay of aircraft and spacecraft structures for nearly a century. Aluminum alloys combine the base metal of aluminum with elements like copper, magnesium, and zinc to enhance the mechanical properties and durability of the material.

More recently, streamlined aircraft with ultra-smooth aerodynamic surfaces have been made almost entirely of composite materials such as glass and carbon fiber-reinforced plastic composites. Composite materials are also used extensively in the construction of modern spacecraft. An advantage of such composite materials is that they can be tailored in shape to carry the applied loads more efficiently, thereby reducing structural weight for a given strength and stiffness; the critical metric is the strength-to-weight ratio or structural efficiency, i.e.,

    \[ \frac{ \text{Material strength}}{\text{Structural weight}} \equiv \frac{\text{How strong the material is}}{\text{How heavy the structure is}} \equiv \text{Structural efficiency} \]

However, other material properties, such as fatigue resistance, corrosion resistance, durability, and repairability, may also be important considerations in selecting the most appropriate construction material(s) for a given aerospace application.

Learning Objectives

  • Learn the factors involved in selecting materials for aerospace applications.
  • Appreciate that different materials have different stiffnesses and failure levels.
  • Recognize some of the advantages and disadvantages of composite materials for constructing aerospace structures.
  • Learn about future aerospace materials and production technologies such as 3D printing.

Types of Materials

Many factors must be considered when selecting materials for aerospace applications, including manufacturing methods and costs. Today, the choices often come down to using aluminum alloys, composites, or both. Sometimes, Titanium alloys may be needed for their strength, toughness, and temperature tolerance (they do not creep), although titanium is at least five times more expensive than aluminum.

Each potential construction material has its relative merits, and the choice of construction material depends on the specific requirements of the aircraft and its mission. To keep the options open, it is usually advisable to leave any material selection based on a “clean-sheet” airframe design quite broad. The table below compares the relative merits of different aerospace construction materials. Besides the strength of the materials, factors such as material costs, tooling, manufacturing processes, fatigue resistance, durability, repairability, corrosion resistance, and crashworthiness are also important. Notice that the rankings shown in the table are somewhat subjective and based on a broad comparison; there may also be other factors to consider than those shown in this table.

Comparison of the relative merits of different aerospace construction materials.
Material Cost Tooling Manufacturing Process Fatigue Resistance Durability Repair-ability Corrosion Resistance Crash-worthiness
Wood and Fabric Low Low Simple Low Low Moderate Low Low
Stressed-Skin Aluminum Moderate Moderate Moderate Moderate High High High Moderate
Titanium High High Moderate Good Good Moderate Good Moderate
Modern Composites High High Complex High High Low High High
  1. Wood and Fabric: This material was widely used in early aircraft because of its low cost and ease of construction. However, it was not very durable, which made it unsuitable for modern aircraft.
  2. Stressed-Skin Aluminum: Aluminum alloys have been widely used in aircraft construction since the 1930s. It was a significant improvement over wood and fabric, offering better strength and durability.
  3. Stressed-Skin Titanium: The use of titanium alloys in the airframe is attractive for high-speed flight vehicles because it offers the most balanced choice regarding strength, weight, and heat resistance. For example, the SR-71 Blackbird was primarily made of titanium, comprising over 90% of its structural weight.
  4. Modern Composites: Composite materials such as carbon fiber reinforced polymers are widely used in contemporary aircraft because they are lightweight, highly fatigue-resistant, durable, and corrosion-resistant. They also offer excellent crashworthiness, especially when combined with Kevlar.

Modern aerospace structures may have 50% or more of their structure (by weight) made of composites, with some new airframe designs reaching as much as 90%. However, the material cost is high, and the tooling and manufacturing processes can be complex. The investments needed in tooling for manufacturing aerospace composite parts can be considerable, and airframe manufacturers may often outsource parts to companies that specialize in designing and manufacturing parts made of one or more types of composite materials. As shown in the figure below, advanced materials can comprise a substantial proportion of the airframe and its weight.

 

Using carbon fiber composites and other advanced materials has improved the strength-to-weight and durability of modern aircraft structures. (Airbus image.)

Strength of Materials

The strength of materials is typically established by measuring coupons or other standard specimens using various testing methods. The tests used depend on the material characteristics that need to be established.

  1. Tensile Strength: Tensile strength tests measure the maximum stress a material can withstand while being stretched or pulled before it breaks. This stress is determined by placing a material sample in a machine that produces an axial tension until the sample fractures. The stress at which failure occurs can then be determined.
  2. Compressive Strength: This measures the strength of the material when it is pushed or squeezed. The type of strength is determined by applying an inward force to a material sample until it fails.
  3. Shear Strength: This measures the material’s ability to withstand forces that cause the internal structure to deform under the action of parallel crosswise forces, i.e., the loading carried by a fastener such as a bolt or rivet.
  4. Flexural Strength: The flexural strength is determined by applying a bending moment to a sample until it fractures.
  5. Impact Strength: This measures a material’s ability to resist a sudden impact and how brittle it is. It is usually determined by subjecting a notched sample to a falling weight or pendulum and measuring the energy needed for the material to fracture.

These are just a few examples, and other specialized tests depend on the specific properties and applications of the material being evaluated. A material’s strength and other properties are crucial in engineering design to ensure that structures and components can withstand the loads and forces they will encounter during use, plus a margin of safety.

Young’s Modulus

The modulus of elasticity or Young’s modulus, which, as previously discussed, is given the symbol E, is one of the factors used to calculate a material’s “strength” and deformation under external loading, i.e., the higher the stiffness, the lower the deformation, i.e.,

(1)   \begin{equation*} E = \frac{\sigma}{\gamma} \end{equation*}

Different materials have different values of E and other characteristics, as shown in the figure below.  Stiff materials have a higher Young’s modulus value but tend to be brittle and fracture easily. Soft and easily extendable materials have a low value of Young’s modulus. There are many online resources where quantitative information on material characteristics is readily available. Notice that a stiff material has a high Young’s modulus and will change its shape only slightly under external loading. A stiff material requires high loads to deform it elastically. A flexible material has a low Young’s modulus, and its shape will change considerably under loading.

Representative stress-strain relationships for different materials. It is essential to know the material’s properties, e.g., aluminum, composite, etc.

Eventually, with increasing applied load, the linear relationship between stress and strain changes and becomes nonlinear. This point is called the yield point, and the material now behaves more like a plastic (i.e., the material flows) in that deformations caused by the loads will remain even after the load is removed. Therefore, the material will take on permanent plastic deformation. Eventually, the material will break or rupture if the stresses become too high. A strong material requires high loads to permanently deform it as well as to break it.

The table below gives some reference values of Young’s modulus for aerospace materials. Values of Young’s modulus can vary slightly depending on the specific composition and treatment of the materials. Composite materials will have a range of values due to fiber orientation, resin type, and fiber content variations. Graphene and CNTs are included as they represent the cutting edge of materials science, though they are not yet commonly used in aerospace applications.

Values of Young’s modulus and yield stress values for typical aerospace materials.
Material Young’s Modulus (GPa) Yield Stress (MPa)
Aluminum Alloy (7075-T6) 71.7 503
Titanium Alloy (Ti-6Al-4V) 113 880
Stainless Steel (304) 193 215
Carbon Fiber Reinforced Polymer (CFRP) 70–230 600–1,500
Glass Fiber Reinforced Polymer (GFRP) 70–90 500–700
Inconel 718 200 1,035
Beryllium 287 240
Magnesium Alloy (AZ31B) 45 200
Nickel Alloy (Hastelloy X) 205 340
High-Strength Steel (Maraging Steel) 210 1,400
Aluminum Lithium Alloy (Al-Li) 79 470
Graphene ~ 1,000 ~ 130,000
Carbon Nanotubes (CNTs) ~ 1,000 ~ 40,000

Materials Selection Charts

One way to approach the material selection problem for an aerospace structure is to eliminate possible materials (e.g., wood, steel, etc.) rather than trying to identify the single best material too soon in the design process. Material selection charts are often used when choosing materials for a particular engineering application. In this type of presentation, one material property is plotted on the ordinate of the chart and another property on the abscissae.

A particularly relevant material selection diagram for flight vehicle applications is one with strength on the ordinate and density on the abscissa, as shown below. In this case, the value of Young’s modulus describes how stiff the material is. This type of chart helps expose the relative benefits of particular materials so that the most appropriate material can be selected for a specific structural application.

A materials selection chart as material stiffness (Young’s modulus) versus its density.

In reference to this type of chart, some materials look very attractive for the light, stiff components needed for flight vehicles. Unfortunately, there are few inexpensive, strong materials. Wood is not strong enough as a primary airframe construction material other than for small airplanes. Three typical metallic construction materials for flight vehicle structures are aluminum, steel, and titanium. Steel is by far the heaviest, being approximately three times as heavy as aluminum or titanium, which rules it out as a primary construction material for most aerospace structures. However, steel is used for components where much strength and durability are needed, such as landing gear. Aluminum alloys perform excellently in strength, lightness, durability, cost, etc., and have traditionally been used to construct flight vehicles. Although aluminum is lighter than titanium, titanium is stronger and has better fatigue resistance. Composites offer a good compromise, but they are usually relatively expensive in terms of the material itself and the tooling needed to manufacture the necessary components, such as molds, layup facilities, and autoclaves.

Check Your Understanding #1 – Material selection based on strength and cost

An aerospace engineer must choose between aluminum alloy 7075-T6 and titanium alloy 6Al-4V for machining a critical load-bearing component for a spar on an aircraft wing. The selection criteria include tensile strength, weight, and cost. The component must withstand a tensile load of 900 kN before reaching its yield point. Compare the relative features of the two materials to decide which one to use. Assume that the supplier can cut the material only in square blocks to the nearest cm in linear length.

The material properties and costs are:

      • Aluminum alloy 7075-T6:
        • Density: 2.81 g/cm^3
        • Tensile Strength: 570 MPa
        • Cost: $4 per kg
      • Titanium alloy Ti-6Al-4V:
        • Density: 4.43 g/cm^3
        • Tensile Strength: 950 MPa
        • Cost: $40 per kg
Show solution/hide solution

Consider first the aluminum alloy 7075-T6.

    \[ \text{\small Required cross-sectional area} = \frac{\text{\small Tensile load}}{\text{\small Tensile yield strength}} \]

Therefore, the required area is

    \[ \text{\small Area} = \frac{900 \times 10^3}{570 \times 10^6 } = 0.001579~\text{m}^2 = 15.79~\text{ cm}^2 \]

Therefore, rounding up to the nearest cm means that the volume of material required will be 256 cm^3, so the cost of the block at $4 per kilogram will be

    \[ \text{\small Cost} = 2.81 \times 256.0 \times (4.0 \times 10^{-3}) = \$2.88 \]

Now consider the titanium alloy Ti-6Al-4V.

    \[ \text{\small Area} = \frac{900,000}{950 \times 10^6} = 0.000947~\text{m}^2 = 9.47~\text{cm}^2 \]

Therefore, rounding up to the nearest cm means that the volume of material required will be 100 cm^3, so the cost of the block at $40 per kilogram will be

    \[ \text{\small Cost} = 4.43  \times 100.0 \times (40.0 \times 10^{-3})  = \$17.72 \]

Notice also that the relative weight of the titanium alloy compared to the aluminum alloy will be

    \[ \dfrac{100.0 \times 4.43}{256.0 \times 2.81} \approx 0.62 \]

Conclusion: The titanium requires less material volume to carry the required load, and the machined part will be smaller and about 40% lighter. However, making the same part out of titanium will be about six times the cost of materials alone. Titanium is also harder to machine, so it will take longer to make each part. The cost of making individual parts adds up quickly for serial production. If weight is critical, then titanium may be the only choice. What would your choice be?

Composite Materials

Composite materials are increasingly important in the construction of aerospace structures. A composite material is defined as

(2)   \begin{equation*} \text{Composite} = \text{Matrix} + \text{filler(s) (e.g., fibers, particles, etc.)} \end{equation*}

The matrix is usually a polymer or epoxy resin binder. The most common filler materials are glass and carbon fibers.

The advantages of using composite materials for structural design are their high strength-to-weight ratios, optimum load-carrying capabilities, elimination of many rivets and other fasteners, as well as giving fatigue and corrosion resistance. Eliminating rivets alone may save 5% of the structural weight over conventional aluminum construction methods. In the 1960s, secondary airframe structures, such as fairings, spoilers, and flight controls, were developed out of composites for weight savings over aluminum parts. New generations of aircraft are being designed with entire fuselages and wing structures made of advanced composites, although the manufacturing techniques and tooling require considerable financial investments.

The first type of aircraft to use a majority of composite materials were German sailplanes such as the Glasflügel Libelle. The RAH-66 Comanche helicopter was one of the first modern aircraft to use advanced composites in its construction. Advanced composites are typically classified as those made of carbon fibers containing fiber volumes more significant than 50%, which is typical of new airliners such as the Boeing 787 and the Airbus A350.

The aerospace industry constantly strives to improve airframe manufacturing by reducing weight and costs, and composite materials are very attractive to this end. Modern composite airframe construction also improves aerodynamics by eliminating surface joints and rivet heads, providing an exceptionally smooth wing shape and reducing drag. Today, larger, lighter, and more integrated aircraft and spacecraft structures are built from advanced composite materials.

Composite Construction

The construction of a composite structure consists of a core fiber material embedded in a polymer or resin binder, often called a polymer matrix. There are two main classes of advanced composites: thermoset and thermoplastic, but ceramic and metal matrix types also exist.  Each of these substances alone has poor strength, but when appropriately combined in the matrix,  it becomes a very stiff and strong structure. The fibers are the primary load-carrying elements of a composite material. Therefore, the material is only strong and stiff in the direction of the fibers, as shown in the figure below, and different combinations of fibers and orientations must used to give the required properties of the composite material. Carbon fiber is the most common type of core material because of its high stiffness and strength as well as low weight, called a carbon fiber reinforced polymer (CFRP).

 

The fibers are set in a matrix of cured polymer in a composite fiber material. The fibers carry nearly all the loads applied to a composite structure.

Not only are the fiber angles important, but the stacking sequence is essential for obtaining the optimum strength-to-weight ratio of the composite material.  The fibers can be laid down and stacked in almost any combination of angles, ideally in an optimum arrangement to carry the applied loads most structurally efficiently. The stacking sequence and orientation of the fibers affect the overall stiffness and failure strength of the composite material. In a unidirectional composite material, the fibers are aligned in one particular direction and will only carry a load in that direction, as shown in the figure below. For the bidirectional orientation of the fibers, the mechanical properties will be distributed in both longitudinal and traverse directions. To carry loads in other directions, the layers can be configured at ± 45° or whatever angles are needed to carry the imposed loads.

In a composite material, the fibers and layers can be configured as needed to carry the imposed loads.

Airframe components made from composites can be designed using the FEM so that the fiber orientations and stacking sequence can produce the optimum mechanical properties of the component in response to the applied loads. Making an anisotropic material (e.g., using combinations of bidirectional or multidirectional fibers) is critical to obtaining the high strength-to-weight ratio of composite components compared to using directionally isotropic materials such as aluminum.

Types of Fibers

The most common composite fiber reinforcement materials are glass, carbon and Kevlar. They are usually supplied as woven or knitted cloth. However, other forms are also used. Glass fiber (usually called  “Fiberglass” under its trademarked name) has good strength and stiffness, and is inexpensive and easy to work with. Often used for secondary structures like wingtips, fairings, etc. Carbon fiber has excellent strength to weight and, more importantly, great stiffness to weight, so it is very useful in primary structures. It is, however, much more expensive than glass fiber. Kevlar has good strength to weight but low stiffness. Therefore, because of its low modulus, it can more readily deform and absorb energy, which is good for structures where crashworthiness is important. For example, airplane nose cones and wing leading edges are one application of Kevlar, which is much less prone than carbon or glass fiber to fracture, such as from a bird strike.

Types of Polymer Matrix Materials

There are many types of polymer matrix materials. i.e., plastics. Polyepoxides or epoxies comprise a bisphenol resin and a hardener that combine and cure to give a solid material. There are many different epoxy formulations with a wide variety of mechanical and other properties. Polyesters or vinyl-esters use a catalyst to cure them although the catalyst itself is not chemically consumed. Polyesters and vinyl-esters generally have poorer properties than epoxies and shorter working times, which makes them more challenging to use in some manufacturing applications.

Material Properties

Composites generally have high strength-to-weight ratios, making them ideal for applications where high strength and low weight are critical. The stiffness or rigidity of composites can be tailored by adjusting the type and orientation of the fibers. Every type of composite material can be uniquely different in terms of its stiffness and failure stress.

Stiffness

To estimate the stiffness or effective Young’s modulus for a composite material like CFRP, a method called the rule of mixtures is used. This stiffness value is a form of weighted average of the individual stiffnesses of the materials comprising the composite. It considers the mechanical properties of the individual components (in this case, the fibers and the polymer matrix) weighted by their volume fractions in the composite.

The rule of mixtures for calculating the stiffness modulus of a composite material is given by

(3)   \begin{equation*} E_{\rm c} = V_{\rm f} \, E_{\rm f} + V_{\rm m}\, E_{\rm m} \end{equation*}

where E_{\rm c} is the Young’s modulus of the composite material, V_{\rm f} and V_{\rm m} are the volume fractions of the fibers and matrix, respectively, and E_{\rm f} and E_{\rm m} are the Young’s moduli of the fibers and matrix, respectively. This value is known as the upper-bound stiffness modulus and corresponds to loading applied in a direction parallel to the fibers, which is the primary load-carrying direction. Typically, E_{\rm f} \gg E_{\rm m}, so that the stiffness of the unidirectional composite is primarily determined by the stiffness of the fibers, as shown in the figure below. Bidirectional and multidirectional fiber orientations can carry loads in other directions so that the fibers can be arranged accordingly for a given loading application.

Stress-strain behavior of a composite material with the fibers aligned in different orientations and arrangements. The best strength and stiffness are obtained by optimally loading the fibers.

Notice that Eq. 3 can also be written as

(4)   \begin{equation*} E_{\rm c} = f \, E_{\rm f} + ( 1 - f) \, E_{\rm m} \end{equation*}

where the volume fraction of the fibers is

(5)   \begin{equation*} f = \dfrac{V_{\rm f}}{V_{\rm f} + V_{\rm m}} \end{equation*}

The inverse rule of mixtures states that in the direction perpendicular to the fibers, the stiffness modulus of a composite is

(6)   \begin{equation*} E_{\rm c} = \left( \dfrac{f}{E_{\rm f}} + \dfrac{( 1 - f)}{E_{\rm m}} \right)^{\!-1} \end{equation*}

This latter value is called the lower-bound stiffness modulus and corresponds to a case of transverse loading applied to the composite material, as also shown in the figure. The validity of Eqs. 3 and 6 are generally limited to composites with fiber volume fractions of no less than 70%.

Maximum Stress Capability

To determine the strength or failure stress \sigma_c, of a composite material, the law of mixtures can again be used. In this case, then

(7)   \begin{equation*} \sigma_{\rm cu} = V_{\rm f} \, \sigma_{\rm fu}  + V_{\rm m}\, \sigma_{\rm mu} \end{equation*}

where \sigma_{\rm fu} and \sigma_{\rm mu} are the ultimate (failing) stresses of the fibers and the matrix, respectively. Because \sigma_{\rm fu} \gg \sigma_{\rm mu}, then the ultimate strength of the composite is primarily determined by the strength of the fibers. The matrix, however, increases the strain at which the fibers will fail.

Indeed, the figure above shows that the fibers will fail at a strain significantly less than what the plastic matrix can sustain, which is denoted by \sigma^{\prime}_{\rm m}, and so Eq. 7 will overestimate the ultimate stress. Therefore, \sigma_{\rm mu} can be replaced by \sigma^{\prime}_{\rm m}, which is the stress in the matrix in which the fibers will fail. Therefore, instead of using Eq. 7 for the strength of the composite, it can be replaced by

(8)   \begin{equation*} \sigma_{\rm cu} = V_{\rm f} \, \sigma_{\rm fu}  + V_{\rm m} \, \sigma^{\prime}_{\rm m} \end{equation*}

which will give a more reasonable estimate of the failure stress of a composite material. Normal production variations in the mechanical properties of the fibers and the matrix material(s) suggest that such calculated values are about 20% higher than those achieved based on statistical testing.

Check Your Understanding #2 – Estimating the stiffness of a composite material

Given that the individual values of Young’s modulus of carbon fiber is 300 GPa and for the matrix material it is 3 GPa, then determine the upper and lower bound moduli of a composite material where the volume fraction of the fibers is 70%, and the volume fraction of the matrix is 30%.

Show solution/hide solution

Using the rule of mixtures for calculating Young’s modulus of a composite material where the loading is applied parallel to the fibers gives

    \[ E_c = V_{\rm f } \, E_{\rm f } + V_{\rm m}\, E_{\rm m} \]

In this case, the modulus of the composite material will be

    \[ E_{\rm c} = 0.70 \times 300.0 + 0.30 \times 3.0 = 210.0~\mbox{GPa} \]

The fiber volume fraction is

    \[ f = \dfrac{V_{\rm f}}{V_{\rm f} + V_{\rm m}} = \dfrac{0.70}{0.70 + 0.3} = 0.7 \]

Therefore, in the direction transverse to the fibers, the inverse rule of mixtures gives the modulus as

    \[ E_c = \left( \dfrac{f}{E_{\rm f}} + \dfrac{( 1 - f)}{E_{\rm m}} \right)^{-1} = \left( \dfrac{0.70}{300.0} + \dfrac{0.3}{3.0} \right)^{-1} = 10.0~\mbox{GPa} \]

Manufacturing Processes

The manufacturing processes for composite materials include hand lay-up, automated fiber placement (AFP), out-of-autoclave (OOA) methods such as resin transfer molding (RTM), vacuum bagging, filament winding, pultrusion, and additive manufacturing (3D printing). Hand lay-up is a traditional, manual method suitable for flexible designs but is labor-intensive, inconsistent in quality, wasteful in resources, and unsuitable for mass production. AFP utilizes robotic systems for precise and repeatable fiber placement, making it ideal for complex shapes, though it requires a significant initial investment in tooling. RTM involves injecting resin into a mold containing dry fiber preforms, offering good surface finishes and suitability for large, complex parts, albeit also with high tooling costs.

Vacuum bagging improves fiber-to-resin ratios and uniformity by compacting fiber lay-ups under a vacuum and is particularly useful for “one-off” parts. In contrast, filament winding is perfect for creating high-strength cylindrical structures by winding fibers around a rotating mandrel. Pultrusion allows continuous production of constant cross-sectional profiles with consistent quality for mass production. OOA processes, including vacuum-assisted methods, cure composites without autoclaves, reducing costs for large structures but usefully at the expense of lesser mechanical properties and lower strength-to-weight ratios. Additive manufacturing, or 3D printing, is rapidly emerging in composite materials (e.g., carbon-impregnated) for rapid prototyping and complex designs, though it currently offers lower strength and less predictable manufacturing consistency.

These advanced manufacturing techniques are becoming increasingly crucial in aerospace for producing components like fuselage sections and interior parts, where high strength-to-weight ratios, corrosion resistance, and fatigue resistance are paramount. The choice of manufacturing process depends on factors such as part complexity, production volume, and specific performance requirements. By enabling the creation of high-performance composite materials, these techniques continue to reduce weight and improve the overall structural efficiency in aerospace applications.

The photograph below shows a carbon-fiber composite wing spar laid up for autoclave curing. The fibers are laid down in a mold as a tape impregnated with uncured resin; this material is often called “prepreg.” Prepreg composite materials can be arranged anisotropically, with the fibers best aligned at the angles needed to carry the applied bending, shear, and torsional loads. The resulting layup is cured under pressure and temperature in an autoclave to produce the final structure. The manufacturing processes are designed to create high-quality composite parts at the standards required for the airworthiness certification of aerospace products.

Tooling and tape-laying machines being used to build up a carbon-fiber spar.

Concerns with Composites

One concern in the manufacture of composite materials is the predictability and repeatability of their properties, which will be influenced by the material constituents, manufacturing processes, and the accuracy of modeling techniques. While advanced manufacturing methods and sophisticated simulations have significantly improved the predictability, ongoing testing and validation are crucial to ensure that the composites meet the desired performance criteria in practical applications. Mechanical testing (tensile, compressive, and shear tests) and non-destructive testing (ultrasonic, X-ray) are used to confirm the predicted performance and identify any discrepancies.

Another concern for aircraft made of composite materials is damage detection and repairability. While damage in metallic structures may be relatively easy to detect (e.g., a surface dent or a crack), damage in composite parts is often subsurface as a delamination or internal fiber cracking. Composites can also be more difficult to repair after damage than metallic construction. New repair methods are constantly being developed for composite materials that restore the structure’s original design strength. Nevertheless, composite structural repair needs specialized tools and materials and must often be accomplished at specialized repair facilities, which will only exist at some airports. In this regard, damaged aircraft may need to be ferried to these unique facilities, the damage repaired, and the aircraft flown back before returning to service.

Sandwich Materials

A sandwich structure is built from a core of a lightweight core material laminated or sandwiched between thin outer skin sheets using an adhesive, as shown in the photograph below. The paper honeycomb material is called Nomex, but aluminum honeycomb or foam like Rohacell may also be used. The skins may be made of aluminum sheets or composite materials. In some high-temperature applications, stainless steel could be used.

An example of a sandwich construction using a paper-honeycomb “Nomex” core structure with a bonded glass-fiber laminate skin.

Material Fatigue

Fatigue is a material failure resulting from repetitive loading cycles, usually at high applied stress levels. In a fatigue failure, a structural component will fail at an applied load well below the average failure load. Aluminum airframe structures are prone to fatigue issues because they are highly stressed and subjected to repetitive loading cycles, e.g., during takeoff, routine and maneuvering flight, rough air, landing, etc. Composite materials tend to have much better fatigue characteristics. Regardless, sharp corners and regions of stress concentration are always potential sources of fatigue cracks, so round holes and smooth transitions are used to increase fatigue resistance. Assessing the fatigue life of an aircraft is an integral part of the design process.

Metal Fatigue

Metal fatigue is the weakening of a metal caused by repeatedly applied cyclic loads, typical of those found in flight structures, resulting in the formation of cracks. The tragic crashes of the de Havilland Comet brought attention to the issue of metal fatigue, which was found to have originated at stress concentrations at improperly prepared rivet holes. The pressurized fuselage, coupled with the thin skin and higher stresses around window cutouts, caused the fuselage of the Comet to suffer explosive decompression, resulting in the loss of two aircraft; the original report explains the details. The subsequent research and understanding of the metal fatigue problem led to significant improvements in the strength and durability of lightweight aluminum structures that benefited all airframe manufacturers.

The mechanism of metal fatigue starts with the development of small micro-cracks in the structure. These grow slowly over many repeated loading cycles to become notable cracks until, eventually, the surrounding material catastrophically fails from a local overstress, as shown in the schematic below. Such cracks often occur in areas of highly loaded airframe structures but are usually difficult to inspect, at least using visual methods. As a result, they can go unnoticed until structural failure occurs, which can be a safety-of-flight issue.

Fatigue cracks can be initiated at stress risers, such as a damaged rivet hole. If undetected, the cracks may slowly propagate until the part fails.

Fatigue failures generally occur in localized regions called stress risers, which can be caused by damage, scratches, or corrosion. Rivet holes are a common source of fatigue cracks, and holes must be adequately prepared after careful drilling to remove any burrs and rough edges before driving home a rivet. After many loading cycles, micro-cracks can be initiated in the metal at any stress riser. These cracks grow in size and length with each loading cycle, although the initial process is typically prolonged.

Eventually, the enlarged cracks can reach a size where the stresses at the tip are high enough to cause more rapid crack propagation. The signature of the crack growth in the material is the curved striations emanating radially from the stress riser, which can be seen in the photograph below. The higher stresses acting over the remainder of the (uncracked) material become enough to cause sudden ductile failure from stress overload. However, it may take tens or hundreds of thousands of loading cycles for this process to occur, assuming the cracked part is not noticed during routine inspection and then repaired or replaced.

A classic fatigue failure, where cracks initiated at a stress riser slowly propagate until sudden failure occurs from an overstress.

Fatigue characteristics are measured by testing material coupons, sub-structures, and even complete airframes. By subjecting the structure to the simulated cyclic magnitudes and stress levels encountered during flight, areas prone to fatigue cracking can be identified and remedied. Remediation includes replacing the part or reinforcing the structure with a patch or a doubler. The importance of FEM at the design stage cannot be underestimated, which can predict areas of locally high stresses and those likely to be subjected to fatigue issues before the structure is even built.

Despite best design practices to eliminate stress risers, fatigue cracks will inevitably appear in an aircraft structure over decades of use, e.g., especially on so-called aging aircraft. Therefore, detecting and repairing such fatigue cracks is essential before they result in severe structural failure. In simple, unpressurized structures, the crack growth can be stopped by drilling a suitably radiused hole at the end of the crack to relieve the stresses. In pressurized structures, a complete repair may be required by adding an appropriately sealed reinforcing structure. Aging aircraft may also suffer from corrosion issues and corrosion-induced fatigue cracking, resulting in serious structural problems that may be difficult and expensive to repair. For these reasons, older aircraft are often retired from service when they experience significant downtime and/or costly repairs.

Wöhler Curves

Fatigue properties of materials are often described using an SN curve or Wöhler (Woehler) curve, an example being shown in the figure below. The SN curve describes the relationship between cyclic stress amplitude and the number of cycles to failure. On the abscissa, the number of cycles to failure is given on a logarithmic scale, and the ordinate (either linear or logarithmic), the stress amplitude (or sometimes the maximum stress) of the cycle is used.

Wöhler curves for different materials will have different characteristics. Fatigue tests are performed by applying cyclic stresses with constant amplitude until failure.

Ferrous materials, typical of Material A, tend to have good fatigue properties. They generally exhibit an endurance limit below a certain stress amplitude, effectively having infinite fatigue life. Aluminum, typical of Material B, has no such behavior and will continue to show a decreasing fatigue resistance with increased cycles. In most cases, aluminum parts subjected to high cyclic stresses must be assigned a safe life in that the part must be withdrawn from service at a certain number of cycles, often measured in takeoff and landings or flight hours. Composite materials, typical of Material C, have good fatigue properties and are much better than most metals. However, their fatigue characteristics tend to have significant variability during testing despite being generally excellent and having much longer fatigue lives than aluminum.

These types of SN curves are obtained from fatigue tests done under controlled conditions in the laboratory. Tests are performed by applying cyclic stresses with constant stress amplitude on various types of specimens until the final fatigue failure occurs. Sometimes, the testing may be stopped after many cycles (N > 10^7), the results being interpreted as an infinite service life. Fatigue curves are often obtained for material specimen coupons, which are designed to give the fatigue properties of the material itself rather than a specific part.

Actual flight structures of multiple components will have more specific SN curves. Sometimes, a critical flight structure being tested, such as a wing spar or wing carry-through structure, will have notches cut to initiate fatigue cracking so that the fatigue life of the aircraft is not over-estimated. Complete airframes may also be tested on the ground in a test rig where the cyclic flight loads can be simulated much quicker, and potential issues can be identified before the aircraft ever develops any fatigue cracks in regular use.

Miner’s Rule

An aircraft will be subjected to many types and amplitudes of stresses during normal flight. Unexpected loads, such as the occasional hard landing, may have to be factored into the fatigue life estimate. It is possible to estimate the fatigue life using the method known as Miner’s Rule. This rule states that the cumulative fatigue damage is given by

(9)   \begin{equation*} D = \sum_{k = 1}^{K} \frac{n_k}{N_k} \end{equation*}

where n_k is the number of load cycles to which the component has been subjected in a certain stress range, and N_k is the number of load cycles in that stress range to reach fatigue failure. The index k varies over K discrete stress ranges. The ratio n_k/N_k is referred to as the partial fatigue damage, and of all partial damage cycles over all occurring stress ranges, it is then a measure of the total fatigue damage. Failure can be expected to occur when D = 1.

Miner’s Rule assesses the fraction of the fatigue life consumed at each stress level and then adds the fractions for all the levels to estimate the fatigue damage in aggregate. The stress level damage can be defined as the product of the stress, \sigma_k, and the number of cycles, n_k, operated under that stress, i.e.,

(10)   \begin{equation*} W_k = n_k \, \sigma_k \end{equation*}

Miner’s Rule assumes that the stress level damage to failure is the same for all applied stress levels, so

(11)   \begin{equation*} W_{\rm fail} = N_k \, \sigma_k~~\mbox{for all values of $k$} \end{equation*}

For example, if W_{\rm fail} = 5,000 for a component, it will fail after 5,000 cycles at a stress level of 1, 1,000 cycles at a stress level of 5, or 2,500 cycles at a stress level of 2, etc. Therefore, the cumulative damage can be written as

(12)   \begin{equation*} D = \sum_{k = 1}^{K} \frac{n_k \, \sigma_k}{N_k \, \sigma_k} = \frac{1}{W_{\rm fail}} \sum_{k = 1}^{K} n_k \, \sigma_k \end{equation*}

While Miner’s Rule is imperfect, partly because it assumes that the fatigue damage in each stress range is linearly additive, it gives reasonable estimates of structural fatigue life.

Check Your Understanding #3 – Estimating fatigue life

It is required to estimate the “cycles to failure” of a thin aluminum rod. Preliminary experiments suggest that the angle between the two arms of the bent rod affects the number of cycles to failure. Therefore, the “angle” can be treated as a surrogate for the “stress.” The first test suggests that it takes 63 repeated bends to an angle of 30 degrees for the rod to suffer a fatigue failure. A subsequent test on a new rod subjects it to the following repeated bends and loading cycles.

Bend angle (deg)  Cycles
15 40
40 15
45 10

Based on Miner’s rule, how many cycles with a 30-degree bend can be expected to break the rod?

Show solution/hide solution

It is necessary to calculate the cumulative damage for each stress cycle, i.e.,

    \[ D = \frac{1}{W_{\rm fail}} \sum_{k = 1}^{K} n_k \, \sigma_k \]

In this case, W_{\rm fail} = 63 \times 30 = 1,890, so the cumulative fatigue damage, in this case, can be expressed as

    \[ D = \frac{(15 \times 40) (30 \times 15) 45 \times 10) 30 \times N}{1,890} = 1 \]

where N is the remaining number of cycles. Performing the arithmetic gives

    \[ D = \frac{600 \times 450 \times 450 \times 30 \times N}{1,890} = 1 \]

Therefore, 30 N = 390, and so N = 13.

Thermal Protection Materials

Thermal protection materials are essential for safeguarding spacecraft during the extreme conditions of atmospheric re-entry and deep space missions. These materials are designed to endure high thermal loads, ensuring the integrity and safety of the vehicle and its contents.

Ablative materials, like phenolic-impregnated carbon ablator (PICA) and SLA-561V, absorb and dissipate heat by eroding in a controlled manner, protecting the spacecraft from intense temperatures. Reinforced Carbon-Carbon (RCC) composites, used in critical areas such as the Space Shuttle’s leading edges, can withstand temperatures up to 1,650 °C. Ceramic tiles, like those used on the Space Shuttle, and flexible insulation blankets like Advanced Flexible Reusable Surface Insulation (AFRSI), provide additional layers of thermal protection through high-temperature resistance and lightweight, conformable insulation.

Each type of thermal protection material offers unique advantages suited to specific mission needs. For example, the Mars Science Laboratory’s Curiosity Rover used a PICA heat shield to endure the Martian atmosphere’s entry heat. Multi-layer insulation (MLI), composed of reflective layers separated by spacers, is crucial for maintaining temperature control in the vacuum of space, often used in satellites and other spacecraft. The Orion spacecraft employs an advanced AVCOAT system, as shown in the photograph below, which combines ablative properties with a honeycomb structure for enhanced performance. These materials and systems are fundamental to the success of space missions, providing the necessary protection against the harsh thermal environments encountered during launch, transit, and re-entry.

The heat shield is one of the most critical elements protecting a spacecraft during re-entry through Earth’s atmosphere. The shield for Orion is made of tiles of ablative material called AVCOAT.

Additive Manufacturing & 3D Printing

Additive manufacturing, sometimes called 3D printing, involves the construction of a part manufactured directly from a CAD model. The basic principle is that an appropriate material is deposited, joined, and solidified, typically layer by layer, to make a part. Fused filament fabrication or fused deposition modeling is the most common method for making 3D-printed parts. This process uses a continuous filament of a thermoplastic material, which is melted at a printer nozzle and deposited on the part in incremental layers. Techniques such as direct metal laser sintering and direct metal laser melting can also produce 3D-printed metal parts.

In this case, parts are 3D printed using PLA plastic in a process called fused deposition. The parts are typically strong and light but may be rough or textured.

An advantage of 3D printing is the ability to produce complex shapes that would be difficult, time-consuming, and expensive to construct by conventional manufacturing methods. An advantage for aerospace applications is that 3D-printed parts can also be made with internal geodesic structures to reduce weight. The attractiveness of 3D printing technologies is such that the range of available materials has increased exponentially over the last decade. Some materials are now of the strength, surface finish, and overall quality for which the industry uses 3D printing for production parts.

One of the concerns with 3D-printed parts for aircraft applications is to ensure their airworthiness, i.e., can they be safely used? In the U.S., the FAA requires that all aircraft parts be manufactured to comply with airworthiness standards under FAA production approvals. To this end, aerospace manufacturers using 3D printing will need to continue to develop processes that comply with regulatory requirements and can prove their airworthiness. Recently, the FAA approved the production of a 3D-printed fuel nozzle for the GE LEAP engine. This part was previously made using 20 pieces welded together, but 3D printing technology allowed it to be made from a single piece, which also cut its weight by 25%.

Future Aerospace Materials

The aerospace industry is constantly evolving, with new materials and technologies being developed to improve vehicle performance, reduce weight, and enhance safety. These innovations are helping to push the boundaries in the constant quest for more efficient, durable, and capable aircraft and spacecraft. Future aerospace materials may include:

    1. Ceramic Matrix Composites (CMCs): These types of materials are known for their high-temperature resistance and low weight. Ceramics are increasingly used in jet engines and other components exposed to extreme heat, such as turbine blades, where they can operate at higher temperatures than traditional metal alloys, thereby improving engine efficiency and reducing fuel consumption.
    2. Graphene and Carbon Nanotubes: These nanomaterials offer extraordinary strength-to-weight ratios and electrical conductivity. Carbon nanotubes (CNTs) are manufactured cylindrical molecules composed of carbon atoms arranged in a hexagonal lattice, as shine in the figure below, which is similar to the structure of graphene. They come in various forms, including single-walled carbon nanotubes (SWCNTs) and multi-walled carbon nanotubes (MWCNTs), each with unique properties and applications. Their applications in aerospace include lightweight structural components, advanced sensors, and improved batteries.
      A single-walled carbon nanotube.
    3. Self-Healing Materials: These materials can automatically repair damage, which is particularly valuable for maintaining the integrity of aerospace structures over time. Polymers embedded with microcapsules of healing agents or materials that respond to environmental changes (e.g., temperature, pressure) or physical damage, are being researched for potential use in aircraft skins and other structural components.
    4. Shape Memory Alloys (SMAs): Shape memory alloys are made from nickel and titanium, called nitinol. These materials change their shape when heated above a specific temperature, e.g., by passing an electrical current through them, and then they will return to their original shape when they cool; the concept is shown in the figure below. This effect occurs because of a phase change in the metal’s molecular lattice structure. Notice that SMAs must be cooled below the transition temperature when heated, thereby exhibiting a hysteresis in their strain behavior. Therefore, when heated and cooled, such materials can extract energy, produce forces, and do work. SMAs can potentially be used in various aerospace applications, including actuators and adaptive structures.
      Shape memory alloys (SMAs) contract when the temperature is increased and then they relax again when cooled.
    5. Hybrid Materials: Combining different materials at the micro or nanoscale can result in hybrids with tailored properties for specific applications. For example, metal-polymer hybrids can offer the strength of metals with the lightweight properties of polymers.
    6. Multifunctional materials: Combining different materials can also allow them to be used for structural purposes as well as, for example, having electrical conductivity. By using the structure to carry electrical power and signals, multifunctional materials could eliminate miles of wiring and tons of weight can be potentially eliminated on a commercial airliner.
    7. Metamaterials: These are engineered materials with properties not found in naturally occurring substances. Metamaterials can manipulate electromagnetic waves, including light and radar. In aerospace, they are being explored for applications in stealth technology, antennas, and improving communication systems.

Summary & Closure

Aerospace structures are a key discipline within the engineering field, which is dedicated to designing and analyzing aircraft and spacecraft components. This discipline ensures that these structures are lightweight, durable, and capable of withstanding extreme conditions, such as high speeds, pressure differentials, and temperature variations. Using advanced lightweight materials and optimized structures is crucial to achieving lighter airframe components. The manufacturability of the structures from the selected materials is also important because it affects the feasibility and costs of mass-producing the components. The development of advanced technologies, such as additive manufacturing, has the potential to help reduce airframe weight even further. For example, using 3D printing materials and techniques, complex structures can be created with minimal material waste, potentially yielding significant weight savings. As the aerospace industry continues to evolve, the ongoing advancements in structural materials and manufacturing technologies will play a pivotal role in shaping the future of air and space flight.

5-Question Self-Assessment Quickquiz

For Further Thought or Discussion

  • Consider the expected costs of making airplane structures from molded composites versus traditional riveted metallic construction. Which method is likely to be more expensive, and why?
  • Eliminating many fasteners in composite construction also improves fatigue resistance. Explain.
  • What role do aerospace materials play in making the industry more sustainable? Discuss efforts to develop recyclable materials and reduce the environmental impact of material production and disposal.
  • How do aerospace engineers address the challenges of material fatigue and durability? Discuss the importance of testing and modeling in predicting material behavior over the lifespan of an aircraft.
  • What potential do nanomaterials have in the aerospace industry? Explore applications where nanomaterials can provide significant benefits and the challenges involved in their implementation.
  • How are new materials being developed for space exploration, such as those used in the Mars rovers or the Artemis program? Discuss the specific requirements for materials used in space missions.

Other Useful Online Resources

To learn more about aerospace materials, try some of these online resources:

  • Good video about the advantages of composite materials.
  • NASA 360 Composite Materials AeClips.
  • From Metals to Composites – Decoding aircraft materials.
  • MAPTIS. An extensive database of materials properties and data used by NASA for aerospace applications.
  • MatWeb. A comprehensive database of material properties, including many aerospace materials.
  • ASM International offers a wide range of resources, including materials databases, handbooks, and journals focused on materials science and engineering.
  • AZO Materials. Provides news, articles, and data on materials, including those used in aerospace applications.
  • SAMPE. An organization dedicated to the promotion of technical excellence in materials and processes.
  • ASTM International. Provides standards and technical papers on materials testing and specifications, including aerospace materials.
  • CW CompositiesWorld. It focuses on composite materials, which are extensively used in aerospace applications.

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Introduction to Aerospace Flight Vehicles Copyright © 2022 – 2024 by J. Gordon Leishman is licensed under a Creative Commons Attribution-NonCommercial-NoDerivatives 4.0 International License, except where otherwise noted.