• Section 5.2 presents characteristics and properties of typical materials used for the construction of the modern GA aircraft.

• Section 5.3 presents a description of the fabrication and installation of various structural components of an aircraft.

(1) Aerodynamic loads (or airloads for short) refer to forces and moments caused by the dynamic pressure to which the aircraft is subjected. Airloads include forces, such as wing lift and drag, and moments, like wing torsion and bending. Their magnitude depends on the weight of the aircraft, the load factor, its geometry, and, again, dynamic pressure. The total magnitude is defined based on requirements set forth by the aviation authorities – for instance, 14 CFR Part 23 and 25. However, the local values depend on the geometry. Thus, consider two aircraft, A and B, of equal weight and wing area that differ only by the wing aspect ratio (AR) and taper ratio (TR). Assume aircraft A has the higher AR and lower TR. For reasons that will be detailed in Chapter 9, The anatomy of the wing, it will generate substantially higher bending moment than aircraft B.

(2) Inertia loads refer to forces and moments caused by subjecting aircraft components to acceleration. An example is the battery, which does not experience any aerodynamic loads. Its support structure must be capable of reacting the forces that result from the applied load factors. Other components, such as a piston engine, are simultaneously subjected to both aerodynamic and inertia loads.

(3) Operational loads refer to loads other than aerodynamics and inertia that are simply caused by the fact that the airplane is being used. Examples of such loads include door hinge and locking loads, floor loading loads, wing step-on loads and other similar loads. Such loads are often tricky to define, but are usually small compared to, say, the wing loads. Operational loads usually lead to wear and tear.

Casting

Casting is one of the oldest manufacturing methods known to man, dating back to at least 4000 BCE [2, Table 1, p. 6]. The process entails the following steps:

The advantage is that the original model of the part can be shaped from material less strong than the material used in the casting. For instance, it is possible to make the original part from wood, whereas the copies are made from some metal. There are a number of different casting methods in existence that depend on the material used or the desired shape. For this reason, casting takes considerable expertise to do well. Casting of aircraft metals (aluminum or steel) will leave the material fully annealed and thus lacking strength. For this reason casting should never be used for critical aircraft structure.

Molding

The difference between casting and molding is that molding involves the construction of a heat-tolerant mold that is used to make multiple part copies, whereas casting involves one mold per part. Molding has become a very sophisticated manufacturing process that requires considerable expertise. An example of such processes is injection molding, in which material in a liquid form is injected under high pressure into the mold – an operation intended to eliminate air bubbles from the material, which are a source of stress concentrations in the material that can render it far less durable than otherwise.

Sheet Metal Forming

The concept forming refers to the process of forcing the material into a particular shape. Industry has developed a large number of methods to force metals into particular shapes. Presenting all of them is beyond the scope of this book, however, when it comes to aircraft, sheet metal forming and forging are the best known. Forging is presented separately as it is considered by many to be in a class on its own.

Sheet metal is most often formed to introduce flanges to stiffen the material so it can be used for stringers and spars or for joining with other sheet metal parts (see paragraph about joining below). It is sometimes formed to provide the skin curvatures of lifting surfaces, although this is only required when the skin thickness becomes too large for the panel to flex freely. Thicknesses of that magnitude are common in the inboard wing skin panels of commercial jetliners or military aircraft. The thickness of aluminum sheets used for GA aircraft is usually small enough to allow it to flex with ease. The material is usually cut to shape using hand- or hydraulically actuated shears, depending on sheet thickness and cut length. Then, the forming takes place using a special tool called a sheet metal brake. The bending operation requires some planning, as there are limits to how tight the bend radius can be. As a consequence, allowances have to be made for extra material for the bend itself. Another phenomenon, springback, must be considered when working with sheet metals. It requires the operator to bend the sheet to a predetermined angle which is slightly greater than the desired angle. Once removed from the metal brake, the sheet will spring back to the desired angle.

There are two kinds of surface flexing the engineer must be aware of: simple and compound (see Figure 5-1). All metals will readily undergo a simple surface flex (or deformation), which in effect is a simple plate bending. Compound flex, on the other hand, is accompanied by internal twist (shearing) of the material molecules, in addition to bending deformation about two axes. Metals resist this type of deformation so it is practically impossible to form the compound flex unless its internal molecular structure is stretched using specialized forming methods such as hydraulic pressing. This fact is imperative when selecting material for aircraft components. The compound surface is where composite materials shine – but the manufacturing of composites is usually more expensive than for aluminum. If it is imperative that the surface features a compound flex, as is the case for very low-drag aircraft such as sailplanes, consider composites. If it suffices to use a simple surface flex, such as frustum fuselages and simply tapered wings, consider sheet metal.

Extrusions

An extrusion is the process of forcing an ingot of near-molten metal through a die with a specific geometric pattern. This is a common process for aluminum alloys intended for use in airframes, although it is also used to produce structural steel for buildings. The process converts the half-molten ingot into a long and straight column of structural material featuring a constant cross-sectional shape. When made from aluminum alloys it is ideal for use as longerons or stringers in airframes. It is common to find extrusions whose cross-section resembles letters such as ‘H,’ ‘L,’ ‘T,’ ‘U’ (also called ‘C’) and ‘Z,’ and far more complicated shapes are available. Of these, the L-extrusion, usually called an angle extrusion, is of great use as a stringer or a spar cap in aluminum spars. The C-extrusion, usually called a C-channel, is of great use for various brackets and hinges designed to react high structural loads. The use of extrusions in aircraft is extensive and includes not only stringers, but seat-tracks, brackets, wing attachment fittings, and countless other applications. Extrusions have higher material strengths than plates as the formation of it compresses the grain structure.

Forging

The best-known and probably the oldest forming operation is forging, dating back perhaps as far as 8000 BCE [2, p. 384]. Forging is when a metal is locally subjected to large compressive forces in the form of “hammering” using various dies and tools. This hammering can be done to either cold or hot parts, but it usually increases the strength, toughness, and durability of the material through the process of work hardening. For this reason, it is not unusual to see aircraft components that must react large forces, such as landing gear struts, made from forged metals [2, p. 384]. Work hardening is a consequence of the deformation of the grain structure of the material. Since most forged parts are subjected to secondary machining operations to improve appearance, the work hardening tends to complicate the manufacturing process by making it harder to finish. Forging metals such as iron and steel at elevated temperatures will reduce work hardening and make them easier to post-process. With respect to the economics of forging, the designer must choose carefully, as the cost of forging a non-critical part may be much higher than, say, molding.

Machining

Machining is the fabrication of a part through the removal of excess material. There are a number of ways machining takes place, the most common being sawing, cutting, turning, and milling. Like the other methods above, machining takes a lot of expertise and experience to do well, but an understanding of what can or cannot be machined can open the door of success and close the door of failure. Machining aluminum and low-carbon steels (e.g. AISI 1025) is relatively easy, but this becomes gradually more difficult when the carbon content increases. Hardened steels are very difficult to machine and require sophisticated tools to accomplish this. Machining is much easier to accomplish when the material is in its annealed state. Afterwards, it is necessary to heat-treat the part to acquire adequate strength, even though many parts will undergo further machining post heat-treating.

Welding

Welding is the joining of parts made from identical metals by heating them to a point of surface melting and then bringing them together to allow their molecules to coalesce. A filler material is often used to create a stronger joint. Welding is one of the most common ways to join parts and forms a very strong and durable bond between the parts. This contrasts joining of parts using soldering or brazing, both of which do not melt the working parts. A large number of methods can be used to perform the welding: most notable are a gas flame, an electric arc, a laser, and an electron beam. Low-carbon-grade steels are easily welded or brazed by all techniques and the filler material should be comparable in strength to the base metals. Steels with higher carbon levels will often require stress-relieving after the welding has been completed and sometimes even subsequent heat treating. Welding is commonly used to join parts making up engine mounts, landing gear, and fuselages, demonstrating the method can take a beating if properly done. It is a drawback that the process often leads to warping that may change the intended geometry. The welding of critical structural aircraft parts should always be done by a certified welder. Critical structural parts should not be made from welded aluminum due to a reduction in fatigue life.

Joining

Joining is typically used to assemble a large part from many smaller ones. It includes operations such as riveting, using threaded fasteners, welding, and so on. For instance, consider the fabrication of an engine mount, which requires a number of typically tubular parts to be joined through the process of welding (see above). Also, the fabrication of an aluminum wing is accomplished through the joining of ribs, spars, stringers, and skin through the process of riveting sheet metal. The aspiring aircraft designer should acquire a deep understanding of the two most common riveting techniques used in the industry: bucking and blind riveting.

Of the two, bucking is the primary method used and is employed when two (or more) aluminum sheets are to be joined (or to join a sheet to an extrusion) and there is ample access to both sides of the parts to be joined, for the technician. The standard procedure is shown in Figure 5-2 in four steps. First the sheets are aligned using carefully placed clamps (not shown). Then, holes are drilled at specific intervals depending on the shear stress to be transferred from one sheet to the next through the rivets. Since the drilling operation typically forms sharp edges (or burrs) on the opposite side, these must be removed prior to the insertion of the rivets. Otherwise, the joining will not develop full strength.

The technician usually and temporarily inserts a special tool through selected holes called a Cleco^®. This prevents the sheets from slipping during further drilling or bucking operations. The third step involves inserting the proper rivets into the hole, and the fourth is the actual bucking operation. It often requires two technicians to accomplish, in particular if large sheets are being joined. The technician on the head side of the rivet places an air hammer against the rivet, while the other places a heavy metal block called a bucking-bar against the opposite side of the rivet. When both are ready, the operator of the air hammer presses a trigger on the air hammer to generate a short burst of hammering to the rivet. The inertia of the bucking bar will then deform the rivet such that a solid and strong attachment is formed, as the hammering will cold work the rivet. Bucking takes practice and careless handling of the tools may damage the sheets around the rivet.

Blind-riveting is only used when lack of access to the back side of the sheets prevents the use of an air hammer and bucking bar. It is also used for non-critical structural assembly. Driving a blind rivet is a very simple two-step operation (see Figure 5-3) in which a special tool, a rivet gun, is used to pull out the stem (or spindle) until it snaps at its weak spot, where its diameter has been deliberately reduced. This allows the stem to be pulled up just enough to compress the opposite end of the rivet and lock it in place. Blind rivets are also available as structural rivets and, as stated earlier, are sometimes the only choice. Cherrymax^® is the best known brand for such rivets. Blind riveting, while far easier to perform than bucking, still requires care in installation in order to avoid tilting of the stem, which might misalign the rivet. It is also considerably more expensive that conventional bucking rivet installation.

Finally, there are a number of different types of rivet heads, but as presenting all of them is not appropriate in this text, only the two most common types will be cited. These are: universal and counter-sunk rivet heads (Figure 5-4). The universal head (as shown in Figure 5-2 and Figure 5-3) is typically used for low-performance aircraft (in terms of airspeed), whereas the counter-sunk rivet head will be flush with respect to the surface of the sheet. This reduces the drag of the airplane and is thus implemented on high-speed aircraft. However, this brings additional complication to the table that increases the cost of the riveting: counter-sunk rivets require an indentation to be made for the rivet head, either by a special drilling operation or forming of a dimple using a special tool. Either one increases production costs.

Next to rivets, threaded fasteners (or bolts) are the most commonly used fasteners in aircraft (see Figure 5-5). Such fasteners have superior tensile (and shear) strengths compared to rivets (which are only intended for use in shear), but are far more expensive to use. Like all aircraft hardware, threaded fasteners must be traceable to an approved manufacturing process. Most bolts used for aircraft applications are general-purpose (e.g. AN-3 through AN-20 bolts), internal-wrenching (e.g. MS-20004 through MS-20024), and close tolerance (e.g. the hex-headed AN-173 through AN-186 or NAS-80 through NAS-86).

The shank of these bolts features a smooth section, called the grip, and a threaded section onto which the nut is mounted. The length of the grip must be equal to or slightly exceed the thickness of the material it is intended to hold. The nut must be tightened or torqued to the right amount to preload the fastener. This ensures the joined parts do not slip during service, ensures a more uniform transfer of loads, and increases the fatigue life of the fastener. Nuts are usually self-locking or non-self-locking. Castellated nuts are a type of the latter that are locked in place using special safety-pins called cotter-pins. Such nuts are required for all structurally critical parts, such as engine mounts, landing gear, and wing attachments. The installation of threaded fasteners should always use flat washers (e.g. AN960) so that torquing the nut will not damage the surface of the joining materials.

Types of Composite

There are three common forms of composite used for industrial applications:

In its most basic form, composites consist of layers of fibers in the form of a cloth that are impregnated with some type of plastic matrix (or resin) and then cured to form a rigid structure. An example of this is fiberglass cloth embedded in epoxy resin. This is how the so-called fiberglass-reinforced plastics (FRP) and carbon-reinforced plastics (CRP) are prepared. Sometimes a third constituent material, called a core, is added to fabricate the so-called composite sandwich. The purpose of the core is to separate the plies by a given thickness and that way increase the stiffness of the structure. The resulting panels are light, stiff, and strong and are ideal for use as skin for wing, HT, VT, or fuselage structures. Panels so stiffened allow multiple ribs and frames to be eliminated from the structure, simplifying the airframe.

To better understand the potential the core has in the stiffness of composites, consider Figure 5-7, which shows three 10′′ long cantilevered composite beams. The top one is a simple 4-ply laminate consisting of typical aircraft-grade fiberglass laid up using a [+45°/−45°]_S layup (S stands for symmetrical). The center and bottom ones feature the same fiberglass layup, with the addition of a 0.375′′ and 0.75′′ core, respectively. The resulting thicknesses and normalized densities (the density of the bottom beam is 1.69× that of the top one) can be seen in the figure. Then, some load is applied to the tip (right end) of the laminates such that the top one deflects 1′′. Applying the same load to the center and bottom beams would cause them to deflect 1/110′′ and 1/409′′ respectively. This means that the bending stiffness of the center beam is 110× greater and the bottom one is 409× stiffer than the top laminate. The huge increase in stiffness only costs a very modest increase in weight.

Structural Analysis of Composite Materials

There are two approaches used to perform structural analysis of composite materials; micro- and macromechanics. Micromechanics examines the interaction of the constituent materials (i.e. of the fibers and matrix) on a microscopic level. One of the outputs is the predicted “average” properties (such as strength and stiffness) of a composite laminate in terms of the properties and behavior of the constituent materials. Another one is the prediction of the distributions of stresses and strains in the laminate on a ply-to-ply basis.

An example of the capability of micromechanics is shown in Figure 5-8. A 6-core-6 sandwich with an unsymmetrical layup [+45°/+45°/0°/+45°/0°/+45°]_S with bi-directional cloth is subjected to pure bending. The three left columns show strains in the composite as a whole, whereas the three right columns show strains in each of the plies making up the composite. The darker shaded region on the top and bottom of each column is the fiberglass plies and the lighter and thicker (taller) center region represents the thickness of the core (0.375′′). Considering the composite strains it can be seen that the moment creates strains identical to those predicted by classic solid mechanics theory. The moment (applied about the x-axis) creates strain about both x- and y-axes through Poisson’s ratio. However, when considering the ply strains, it can be seen that the largest strains are picked up by the four 0° plies and the core. The core has a very low modulus of elasticity (Young’s modulus) so it can stretch quite a bit without the formation of large stresses. The plies, on the other hand, have a very high modulus of elasticity, so the four 0° plies will generate substantially larger stresses than the +45° ones. As a consequence, if the applied moment becomes large enough, they are the first plies to fail. The application of micromechanics further allows the structural analyst to evaluate whether, if this happens, the remaining six plies will be capable of reacting the moment, or will fail subsequently. That sort of analysis is called residual strength analysis and is a standard procedure in the development of composite aircraft.

Macromechanics is the study of composite materials assuming they can be approximated as if they were homogeneous and the effects of the constituent materials are detected only as averaged apparent properties of the composite. This way, composite structural members are treated almost as if they were isotropic (except with different properties along each material axis), yielding a convenient analysis workaround for use in finite element analysis software.

The structural analysis of composites is performed using the classic laminate theory, which allows the prediction of stresses and strains in composite laminates. Among others, Tsai [9] and Jones [10] provide a good treatise of the theory. It is based on the following assumptions:

General directions of the stress and strain axes are denoted as shown in Figure 5-9. Then, material properties in the principal material directions 1, 2, and 3 are as follows:

The examples presented in Figure 5-7 and Figure 5-8 are prepared using the theory.

Pros and Cons of Composite Materials

FRPs offer many benefits over traditional materials, among which are high strength, light weight, flexibility in design, ease in the fabrication of compound surfaces, part consolidation, high dielectric strength, dimensional stability, and corrosion resistance. At this time, composites are being used in a seemingly endless number of applications, ranging from recreational boats, where they have practically replaced other traditional building methods, to aircraft. In the aircraft industry, composites are exceptional because of a favorable strength-to-weight ratio and the ease with which compound surfaces can be fabricated. Such surfaces are essential for drag reduction in aircraft. Due to good electrical insulating properties composites are also ideal for use in appliances, tools, and other machinery. Furthermore, they are corrosion-resistant and offer extended service life over metals. The author has had an experimental composite structure exposed to the elements since 1989 to evaluate this aspect of composites. When last inspected, degradation was impossible to detect visually, although admittedly no coupons have been pulled off this structure to validate whether the strength properties have changed. Overall, in addition to high tensile strength, glass fibers offer excellent thermal properties, as well as great impact- and chemical-resistance characteristics.

FRPs also come with disadvantages. To begin with, the resin is highly toxic; if not handled with care, it can easily result in serious dermatitis. It is also subject to storage limitations, strength variability, and impact sensitivity, all of which may cause serious strength degradation. Impact may also cause delamination, which is a separation of plies that results in strength, stiffness, and buckling issues. It is a serious flaw that composites tend to fail with limited warning. Metals, in contrast, fail only after an inelastic elongation. The strength of composite structures is vulnerable to fabrication flaws such as wrinkling, bridging, and dry fibers which will compromise its strength. In the professional manufacturing environment, the structure is carefully inspected against these, which adds cost to the production. Composites are notoriously poor in bearing and require careful attention to cleanliness during the construction process. Additionally, they often require specific surface finish requirements. For instance, FRPs and CRPs require light colors, preferably white, on surfaces exposed to sunlight to minimize heat absorption. Heat can be very detrimental to the strength of the resin being used. As a consequence, their operational temperature limits are well below that of aluminum, which is not that great to begin with. Additionally, the fact that FRPs are good electrical insulators makes them very vulnerable to catastrophic failure if struck by lightning. This is particularly critical to airplanes and calls for specific methods to carry electrical current by the introduction of metal conductors that have to be co-cured with the composite. To add insult to injury, these conductors are typically a “one-shot deal.” They have to be replaced upon landing, unless of course a second flash of lightning strikes first.

Fibers

While it is this combination of matrix, fiber and manufacturing process that gives the composite its superior performance, it is helpful to consider these elements separately.

It is primarily the glass fibers that are embedded in the resin that account for the strength advantage FRPs have over unreinforced plastics. Fibers typically come in three forms: uni-directional, bi-directional, and as fiber mats. The first two are shown in Figure 5-9. Fiber mats are chopped strands of fibers that are randomly assembled into a cloth. They are not to be used for primary structures in aircraft as their strength and stiffness properties are unacceptably poor. On the other hand they are acceptable as secondary structures provided the ply thickness is low enough. Fiber mats are commonly used for boats, swimming pools, and jacuzzis.

The fibers play an imperative role in a composite structure reacting loads. Since the fibers are much stiffer than the matrix, the load inevitably is reacted by the fibers. The resin matrix, in contrast, serves to distribute the load among the fibers, besides retaining the intended shape of the structure. Several types of fibers are available commercially, of which the most common are introduced below:

Resin

The purpose of the resin is to bind the fibers together into a single structural unit and, in the process, distribute strains among them while protecting them from the elements. Generally, there are two kinds of resin: thermosets and thermoplastics. The difference depends on the chemistry of the polymers, both of which contain highly complex molecular chains. In the case of thermosets, as the resin cures, molecular chains crosslink to form a rigid structure that cannot be changed through the further application of heat; the final product is irreversible. Thermoplastics, on the other hand, can be processed at higher temperatures; they can be reheated and reshaped more than once; the final product is reversible.

Sandwich Core Materials

The sandwich core can be made from a multitude of materials, although with some constraints. First, the resin must not be a solvent for the core; and second, it has to be resilient enough to not fail before the fiberglass. The following materials are well suited for use in aircraft composite sandwiches, although some are not used for certified aircraft:

Glass Transition Temperature

In terms of FRPs and GRPs, the glass transition temperature, T_G, refers to the temperature at which the resin transitions from a hard and relatively brittle state into a molten (or soft) state. Reaching this temperature in operation could be catastrophic to a primary structure as it renders the laminate incapable of reacting the applied loads. Most FRPs and GRPs used for aviation applications have a TG in excess of 180 °F.

Gelcoat

Gelcoat is what provides the glossy, high-quality finish on the exposed surface of FRPs and GRPs. It is a polyester or epoxy resin specifically prepared with chemicals to control viscosity and cure-time, as well as pigment with the desired color. Gelcoat is sprayed into the mold ahead of the plies being laid up.

Pre-cure

Pre-cure is a term used for flat laminated plates that are cured prior to being used as a supplemental structural material. Think of it as a flat sheet of aluminum alloy, except it is made from FRP. Having these at one’s disposal is priceless, as one can cut them to a desired shape, and then co-cure them with a laminate layup. Pre-cures are frequently used to place hard points in a sandwich laminate, through which metal fasteners may be used. Their thickness is then equal to the thickness of the core. This will form a kind of island of solid laminate in the sandwich panel, which, as stated earlier, is ideal to provide bearing strength and transfer fastener load into the sandwich.

Aircraft Construction Methodologies

There are primarily two methods used to build composite airplanes; moldless composite sandwich construction, and molded composite construction. The former is typically used for homebuilt or kit aircraft and is a method generally thought to have been pioneered by the well-known Burt Rutan to permit customers to fabricate the experimental Rutan VariEze and LongEze kit planes [11]. The method is explained in detail by Lambie [12] and Clarke [13]. The first step in the application of the method is to “preform” the sandwich core using a multitude of methods. Once the core has been prepared, it is covered with fiberglass cloth and subsequently impregnated with resin (or “wetlay”). The impregnation takes place by pouring resin over the cloth and then paintbrushes and squeegees are used to spread it and to wet the entire cloth. This step requires careful attention to prevent too much resin from being used. If more than one ply is required, the second ply is laid on top of the first while it is still wet, and more resin is added, and so on. The part is then allowed to cure.

The pre-forming of the core is done by a multitude of methods. It can be something as simple (but crude) as carving or sanding the foam to shape, although for parts that require greater accuracy (note that accuracy is a relative term) the foam is cut to shape using an electrically heated wire (or “hot-wiring”). While satisfactory for homebuilt aircraft, this method is never used for certified aircraft because of its inherent flaw of quality irregularity.

Instead, molded composite construction is used for certified aircraft. The method uses “female” or cavity molds that have been accurately shaped to form the outside mold line (OML) of the part. Then fiberglass cloth called “pre-preg” (because it is already impregnated with resin) is laid inside the mold. If more than one ply is required, another pre-preg is laid on top of the previous one, and so on. Once the layup has been completed, some strands of sticky putty are laid around the part and then a plastic sheet is draped over it and tacked to the putty. This encloses the part in a hermetically sealed environment (“vacuum-bagging”). Then a vacuum pump is connected to the plastic and turned on to form a vacuum under the plastic veil. This is a part of a production process to be explained in a moment. Then, the part sitting in the mold is rolled into a warming room (perhaps some 150–180 °F) where it is allowed to cure for a specific number of hours.

So, why is all of this preparation necessary?

The pre-preg is a special fiberglass (or graphite) cloth impregnated with resin under controlled circumstances. This ensures repeatability. In order to certify an airplane, the material qualities have to be repeatable. This means that the strength of the composite laminate should not vary from day to day – it should be the same no matter the time of month or position of the moon. The warming room ensures the resin cures at an optimum temperature, but this maximizes the strength of the laminate. It also lowers the viscosity of the impregnated resin, something taken advantage of through the application of pressure. With the formation of a vacuum on the part side of the plastic veil, atmospheric pressure squeezes air-bubbles out of the pre-preg and helps spread the resin uniformly throughout the laminate. Both improve the quality of the laminate, ensuring the proper fiber/resin ratio. Sometimes, rather than using vacuum, which only applies a 1 atmospheric pressure to the laminate, the laminate is brought into a pressurized container, called an “auto clave.” There it is subjected to as much as 5–10 times the atmospheric pressure.

Fabrication Methods

There are a few fabrication methods used to manufacture FRPs and GRPs that are worth presenting in this context:

The above, which barely scratches the surface, should demonstrate that the field of composites and plastics is booming. Selected properties of typical composite materials are shown in Table 5-8. Note that there is a large variation in properties between fiber brands, fiber volume, resin system, layup process, and other factors. The table should not be used for structural analyses – it is only presented to give ballpark values. Figure 5-10 compares the density, cost, strength, and stiffness of several composite materials (and polyethylene plastic), using E-glass as a baseline. Such a comparison matrix is helpful when selecting material for an application.

The reader wanting to learn more about composite materials and their use and certification in the aviation industry is directed to MIL-HDBK-17 [14], AC-20-107B [15], and AC-21-26 [16].

• A stiffener is a longitudinal or transverse member intended to reinforce a structure by increasing its stiffness.

• A flange is a longitudinal stiffener that runs along the edge of a flexible shell (or sheet), whose purpose is to increase the stiffness of the shell.

• A stringer is a longitudinal stiffener that is not a flange.

• A boom is a beam in the shape of a shell.

• A rib is a transverse stiffener in an open shell, or the end of a closed shell.

• A bulkhead is a transverse member in a closed shell. Strictly speaking the term includes members such as wing ribs and fuselage hoop-frames, although a bulkhead is often used for transverse members that are more substantial than frames and to which other structures attach, for instance, wing, horizontal tail, and engines.

• A primary structure is one that reacts primary loads applied to a structure. Airworthiness Circular 25.1529-1 defines it as structure that significantly contributes to the carrying of flight, ground, or pressure loads.

• A secondary structure carries significantly lower operational loads than a primary structure and is usually used for fairings and doors for unpressurized containers, for example.

• An allowable is a maximum allowable stress value of some specific material property. For instance, the ultimate tensile stress allowable for a 2024-T3 aluminum sheet of 0.125 inch thickness is 64,000 psi.

• A notched allowable is an allowable assuming it has notches in it. This always results in a reduction in the allowable, sometimes as much as 50%.

• Fail-safe means that should the primary load path in a structure fail during operation, an alternative load path exists that prevents a catastrophic failure of the structure as a whole.

Monocoque and Semi-monocoque Structure

The word monocoque comes from the Greek word mono (single) and the French word coque (shell). Monocoque is a structural technique in which stresses are reacted by a thin membrane or a shell of material, rather than a collection of beams. Such structures are stiff in bending, and light, and are therefore ideal for weight-sensitive vehicles such as airplanes. A good way to visualize a monocoque structure is to fold a sheet of paper into a cylinder and tape the free edge using something like Scotch tape. Although the resulting structure is stiff in bending this will reveal its greatest weakness – structural instability. A monocoque structure has a great tendency to fail in buckling or crippling, something easily demonstrated by a person standing on top of an empty aluminum beverage can. The empty container can support a grown man, but push the side in with a pencil with the person standing on it and the can will be compressed in the blink of an eye. This instability necessitates the addition of an internal support structure that resists these failure modes. Such a support structure is generally an assembly of frames, bulkheads, stringers, and longerons (Figure 5-11). The combination is referred to as semi-monocoque construction. Although the addition of the support structure adds to the overall weight of the configuration, it retains its light and stiff characteristics while reacting the applied loads. The advent of the monocoque structure was a major breakthrough in the development of aircraft structures. As stated earlier, such structures react the applied loads in the skin (hence the name “stressed-skin” construction).

Wood Construction

Not too many aircraft are built from wood in modern times, although a few are still being operated today. The most prominent is arguably the de Havilland DH-98 Mosquito, a twin-engine, multi-role combat aircraft, made famous during World War II (see Figure 5-12). The fuselage of the Mosquito was made from a composite consisting of sheets of balsawood bonded to sheets of birch. The wing was a one-piece, all-wood construction. It featured two spars made from spruce and plywood, and the skin was a plywood sheet. A cutaway of the Mosquito can be seen in Figure 5-13, showing details of how ribs, spars, bulkheads, and skins were assembled to make this historic airplane.

The largest flying boat ever built, the Hughes H-4 Hercules, also known as the Spruce Goose, was built from plywood and to this day boasts the largest wingspan of any aircraft in the history of aviation. Its wingspan of 97.5 m (320 ft) is larger even that of the world’s current largest aircraft, the Antonov An-225 Mriya (with 88.4 m or 290-ft wingspan). Well-known examples of GA aircraft made from wood are the Bellanca Viking (designed in 1960) and various types of aircraft made by Jodel (originating in 1946) and the Robin DR400. Additionally, there are a number of kit planes for amateur builders that feature wooden construction.

As with everything, constructing aircraft out of wood has pros and cons. Among the advantages of wood are availability, affordability, comparable strength and good impact resistance, and it is usually easy to work with. Among the disadvantages are inconsistent material properties; crack growth (splitting); low Young’s moduli; possibility of rotting and even termite infestation; flammability; water absorbability (moisture variation); reduction in strength if moisture content exceeds fiber-saturation; and sensitivity to grain direction (anisotropy).

Plywood is an excellent structural material for use in wooden aircraft. It is usually made from an odd number of thin sheets (plies), each oriented at a 90° angle to the adjacent ply. The primary advantage is it offers bi-directional material properties, in addition to offering greater resistance to splitting and much-improved dimensional stability with moisture content. Plywood used in aircraft must comply with standards set by MIL-P-6070, which requires it to be tested for dimensional discrepancies, glue strength, strength properties, and others. Plywood is typically used for wing skin, fuselage skin, ribs, and frames. Common types of plywood for use in aircraft are made from birch, poplar, fir, maple, and mahogany.

Parts made from wood are primarily joined by two means: bonding and mechanical joints. The use of joining shapes (such as lap-joints; tongue and grooves; tenon and mortise, etc.) is not recommended as these invariably lead to stress concentrations that may ultimately lead to failure. Milling or routing parts is acceptable if corners are rounded. Wooden parts require special protection internally, as well as externally.

The adhesive typically used is called Aerodux-500 Resorcinol Adhesive. It is a water- and boil-proof resorcinol/formaldehyde adhesive designed for use in structural wood beams. It requires a hardener to cure, mixed in the ratio 1:1, and can cure at temperatures as low as 7 °C (45 °F).

The reader interested in designing an aircraft constructed from wood is directed toward the documents ANC-18 Design Of Wood Aircraft Structures [19], a classic text on best practices and structural analysis of wooden structures, and NACA R-354 [20], a 34-page report with tips regarding selection and the properties of wood.

Steel Truss Covered with Fabric

Many aircraft feature a fuselage consisting of a truss structure made from steel tubes welded to form a stiff, strong, and light structure. Normally, the truss (see Figure 5-14) is then covered with fabric and dope. Such fuselages are usually made from straight sections of steel tubing (nowadays from 4130 chromium-molybdenum steel) and, more often than not, are rectangular in shape. While robust, such a structure is not exactly ideal for low-drag fuselages. This is not to say that there are not low-drag aircraft around that feature such a structure. As an example, the German Scheibe SF-25 Motor Falke motor glider features a steel truss fuselage structure, although it is an exception to the rule. The method is commonly used for aerobatic and agricultural airplanes, which take a severe beating operationally and for which the truss structure serves well.

Aluminum Construction

Aluminum remains the most common aircraft construction method at the time of writing. Stressed-skin construction has revealed itself as a very efficient means of producing aircraft, thanks to a sheet metal skin riveted to sheet metal frames and bulkheads. Such a structure is very light and stiff, and the industry has developed a large number of impressive tools and techniques to put together aircraft in a short time. The stressed skin means that shear, torsion, and bending loads are reacted by the skin, which is stabilized (made less susceptible to buckling and crippling) using frames and stringers. Of course, a part of the load is reacted by this extra structure, but the difference is that the skins actually transfer loads in the form of stresses, unlike the truss-and-fabric construction of the past.

Aluminum also offers an important benefit – damage repair is relatively straightforward. As such it is surprisingly forgiving of damage, although some would argue it is relatively susceptible to damage. However, the number of parts required to assemble such aircraft is usually significant. This is due to the fact that the aluminum is in the form of thin sheets that are folded and cut to specific shapes before being riveted together. The thickness of the aluminum sheets renders it very flexible. Therefore, it has very limited buckling and crippling resistance, until it is stabilized using stiffeners – which are long sections of aluminum made either by folding other aluminum sheets or from extrusions.

Figure 5-15 shows a cutaway of the famous Supermarine Spitfire and reveals a large number of parts required to make this typical high-performance aluminum aircraft.

Composite Sandwich Construction

Composites and composite sandwich construction have already been discussed. The advent of FAA-certified aircraft such as the Cirrus SR20, SR22, Cessna Corvalis, and a series of aircraft produced by Diamond Aircraft reveals the advantage of such constructions. All feature modern tadpole fuselages and NLF airfoils, making them very efficient. For instance, the SR22 and Corvalis, both of which have fixed landing gear and wide fuselages (50 inches), offer cruising speeds that are right up there with that of rival aircraft such as the aluminum Mooney Ovation, which has retractable landing gear and a narrow fuselage (43.5- inch internal width).

The most obvious difference between composite aircraft construction and conventional wood or aluminum construction is the number of parts. For instance, the wing spar of a composite aircraft is typically a one-piece component, tip-to-tip. An aluminum structure, in contrast, consists of multiple parts: spar caps, shear webs, stiffeners, all of which are assembled using rivets. A composite wing also contains far fewer ribs, as the skins are stiffened and, thus, do not need the same number of ribs, and they are devoid of stringers. The typical composite airplane is bonded together, using an actual adhesive, something it has in common with wooden aircraft. From a certain point of view, assembling a modern composite (certified) aircraft is not unlike putting together a plastic model. The shells of components are bonded together, effectively using“glue.” Of course the analogy ends there, but the process requires far fewer parts than do aluminum aircraft.

Fabrication and Installation of Control Surfaces

The HT is usually attached to the fuselage using specific hardpoints, which are fundamentally analogous to wing hardpoints, albeit simpler. Many composite aircraft bond the skin of the HT to a wide flange designed into the fuselage. This ensures the horizontal loads are transferred through a wide bond area. The VT is often an integral part of the fuselage airframe – in particular in composite aircraft. In aluminum aircraft the VT spar is sometimes integral to the aft-most bulkhead, whose purpose may be twofold; to provide a load path for the stabilizing surfaces and a means to anchor the control-system pulleys or bellcranks, which often react substantial loads.

The stabilizing surfaces typically feature symmetric airfoils, such as NACA 0008 through 0012, whose thickness ranges from 8% to 12% of the chord (as indicated by their designations). Such airfoils are low drag and yet offer volume to accommodate control system cables, pushrods, pulleys, and bellcranks. This thickness also results in a stiff structure that is free of flutter within the operating envelope of the aircraft. However, selecting a symmetric airfoil is not a rule. Some propeller-powered aircraft use either a cambered airfoil or a symmetrical airfoil at an angle-of-incidence with respect to the fuselage to reduce propeller effects (see Section 14.2, Propeller effects). Many aircraft feature symmetric airfoils such as NACA 63-008 through NACA 63-012, whose maximum thickness is farther aft than that of the double-o series, allowing the stabilizing surfaces to sustain NLF more extensively. When possible, the designer of efficient aircraft should consider such NLF airfoils for stabilizing surfaces. However, it must be emphasized this is not always practical.

It is important to bring up two issues that may lead to difficulties in the production of control surfaces. The first is the selection of airfoils that have very narrow trailing edges. Often, this is an unfortunate consequence of selecting NLF airfoils. Such airfoils often feature a cusp that makes the space between the upper and lower surfaces very small at the trailing edge. This calls for ribs so small they cannot be installed without being shortened and this may lead to a partially unsupported trailing edge, susceptible to flexing under loads. Changing from one airfoil to another on a tapered wing may require the wing skin to be stretched. If built from aluminum, this may call for costly manufacturing methods in a mass-produced aircraft. A solution often utilized is to simply ignore the cusp and replace it with a flat section (see Figure 5-23). Of course the resulting airfoil is not the one the designer initially intended. It is important to be aware of such problems and (1) avoid design solutions that are impractical from a manufacturing standpoint; and (2) ensure the production airfoils are those that are represented in the design analysis work and not the theoretical ones that are impossible to fabricate.

The other issue also has to do with cusped NLF airfoils. The high pressure generated by the lower surface of the cusp results in hinge moments that deflect the surface trailing edge up at higher AOAs, for instance, during climb (see Figure 5-24). This invalidates the drag coefficient, modifies the airfoil and reduces the performance of the airfoil. The designer should insist that the engineers designing the flight control system are aware of such detrimental tendency.

Unconventional Tails: T-tail, V-tail, and H-tail

In addition to the normal loads that a specific stabilizing assembly must react, the T-tail must react higher asymmetric loads than a conventional tail. When an aircraft featuring a T-tail yaws at a high AOA, each side of the HT generates very different normal loads. The windward side will see a high-pressure region form between the fin and the HT. The leeward side of the HT will generate much less load. This creates asymmetry in the loading on the HT that the fuselage must react in addition to the moment generated by the VT. As a consequence, the fuselage and the VT both end up being heavier than a conventional tail. Additionally, aeroelastically, the mass of the HT placed at the tip of the VT reduces its natural frequency, which, in turn, lowers its flutter airspeed. The remedy is to stiffen the VT, which again adds weight to the airframe.

Similar concerns can be raised about a V-tail. In order to achieve static and dynamic stability in a V-tail aircraft, relatively large surfaces are required. The rudder functionality of the tail requires the ruddervators on the left and right sides to deflect opposite to each other, similar to ailerons. The deflection can result in substantial forces on the two tail surfaces – one acting up and the other down. As a consequence, a large torsion is generated that the fuselage must react. The designer of V-tail aircraft must be cognizant of this effect and ensure the cross-sectional area of the fuselage in the area of the V-tail is large enough to bring the shear flow down to acceptable levels.

The H-tail brings similar complications to the empennage loads as do the T- and V-tails. Additionally, as in the case of the T-tail, the two fins placed at the tip of the HT can be considered point masses at the end of a cantilevered beam, which brings down its natural frequency and, therefore, the flutter speed. The Lockheed Electra, a twin-engine, 10-passenger military transport, designed in the 1930s, suffered catastrophic flutter of its H-tail. The designer must select thick enough airfoils to reduce the likelihood of such events.

The reader is directed to Section 11.3, On the pros and cons of tail configurations, for additional information about these and other tail configurations.

Fuselage Structural Assembly – Conventional Aluminum Construction

A conventional aluminum semi-monocoque fuselage structure is shown in Figure 5-25 and Figure 5-26. This structure typically consists of a row of hoop-frames that are joined with stringers and floor beams and floor frames by rivets or metal bonding. A schematic of a typical cylindrical center structure is shown in Figure 5-26. The aluminum sheets that form the fuselage skin are then riveted to the frames and stringers, forming a very stiff but light structure. In this structure, cutouts must have a generous radius rather than sharp corners, as these will generate substantial stress concentrations that will significantly reduce the life of the structure.

Figure 5-25 shows a very important feature of such a structure: all cutouts feature generously rounded corners or are elliptical or circular in some airplanes. The purpose of this is to reduce stress concentrations that result from the removal of material and, thus, increase the durability (and safety) the structure.

Fuselage Structural Assembly – Composite Construction

A semi-monocoque fuselage made from composites features a significantly different philosophy. The skin is stiffened using a composite sandwich that consists of a layer of material like honeycomb or foam sandwiched between layers of fiberglass or graphite or similar. This way, the total number of frames can be reduced substantially. Such fuselages are typically made by first bonding the internal structure together and allowing it to cure. Then, this assembly is bonded as a single piece to the stiffened fuselage skin halves. The fuselage skins have joggles to which the adhesive is applied. Joining the skins' halves together properly is easier said, than done.

A common difficulty in the assembly of semi-monocoque composite fuselages is to control the thickness of the bonds (adhesive). Certified composite aircraft have to demonstrate the repeatability of the bondline strengths. This is done by the manufacturer, which assesses a range of bondline thicknesses expected to be seen during production (e.g. 0.040′′ to 0.125′′). Then, repeated strength tests of specimens using those thicknesses are performed, and this allows the strength of the bondline thicknesses to be established. Then, during production, the bondline thickness of all bonds is inspected and if found outside these limits a repair must be designed, basically using the assumption that the bondline strength is zero. Then, a technician performs the repair. It can be seen that poor production quality can cost the corresponding manufacturer a fortune. For this reason, established manufacturers perform the bonding operations using special machinery that substantially cuts down on such deviations (or non-conformances as they are called in industry).

Special Considerations: Pressurization

Tens of thousands of passenger and business aircraft operate every day at altitudes ranging from 25,000 to 51,000 ft. At those altitudes, especially altitudes above 40,000 ft, people quickly die if exposed to the outside atmosphere. For this reason, any aircraft designed to routinely operate at such altitudes must provide oxygen to the crew and passengers. Effectively, such aircraft are pressure vessels and are capable of maintaining higher pressure inside the cabin than that of the ambient atmosphere. Generally, people will begin to suffer from oxygen deficiency at altitudes as low as 14,000 ft. The individual capability varies, with some individuals capable of climbing mountains as high as 28,600 ft,³ and as low as 6000–8000 ft for people who suffer from heart problems. Requirements in 14 CFR Part 121 to supply oxygen to the occupants when operating an aircraft are stipulated in the following paragraphs:

However, it is paragraphs 14 CFR 23.841, Pressurized cabins, and 14 CFR 25.841, Pressurized cabins, that stipulate what capability the airframe must possess in order to sustain cabin pressure in the case of a system failure. Design guidelines are also given in SAE ARP1270, Aircraft Cabin Pressurization Control Criteria.

With respect to aircraft, requiring pressurizations inflicts serious design, manufacturing, maintenance, and operational limitations on the aircraft. It is essential that the designer becomes aware of some of these requirements as they expose a number of challenges, ranging from fuselage deformation to system and equipment installation.

Figure 5-27 shows a cabin pressurization schedule for a typical commercial jetliner. As the airplane climbs, the pressurization will immediately begin to delay the altitude reduction in the cabin. As the airplane reaches its intended cruise altitude (here 39,000 ft) the pressure difference between the atmosphere at that altitude and the cabin is about 8 psi (atmospheric pressure at S-L is 14.7 psi), but this is equivalent to an atmospheric pressure at 8000 ft, enough for all except the weakest of us to survive.

From a structural standpoint, the most efficient pressure vessel is one that reacts lateral or out-of-plane stresses as tensile stresses [22]. This implies that for a given volume, a spherical pressure vessel is the most efficient one (see Figure 5-28). However, when it comes to transporting passengers and maintaining acceptable performance, and stability and control, this geometry is not practical. The next best shape is a cylinder, which is of course a sphere that has been split along a meridional and the two halves attached to a cylinder (again see Figure 5-28). From a structural standpoint, this is the ideal shape for a pressurized aircraft, and this explains why this form prevails in the aviation industry.

As shown in Figure 5-28, the sphere reacts the out-of-plane pressure load as hoop stresses only. However, the cylinder reacts it as hoop and tangential stresses. When used for fuselages, this requires an especially reinforced structure to be placed at either end of the fuselage. This structure is called a pressure bulkhead, and it must react a substantial pressure force. The layperson is often oblivious to the forces the fuselage must support which are solely attributed to the pressure differential. For instance, the typical passenger door in a commercial aircraft is 42 × 72 inches (type A door). When exposed to an 8 psi pressure differential (see Figure 5-27) the total out-of-plane force acting on it amounts to 42 × 72 × 8 = 24,192 lb_f. This explains why it is so amusing when angry passengers threaten to open the doors in flight.

In all seriousness, this number depicts the robustness of the reinforcement required to hold the doors in place. A common method is to use doors that are shaped like a plug (see Figure 5-27). Such doors help distribute the pressure load around the door frame. It is sad but true that doors in pressurized aircraft are overlooked by almost anyone except the engineers who designed them. It is sad because these doors are truly a marvel of modern engineering. Not only do they have to react some 12 tons of load, when opened, most types swing to the outside of the airplane by a simply operated door handle. Think about that next time you board an airplane. How would you get a plug type door to do this feat? In addition to an immense pressurization load, cutouts for windows and doors will further cause stress concentration requiring a yet greater amount of material to sustain. The aspiring designer insisting on super-large entry doors and windows for a pressurized airplane should be very cognizant of the structural challenges this may cause. While large openings can be implemented technically, it will most certainly reduce the useful load of the aircraft.