Curable composite bush

A curable composite bush for an aircraft joint comprising a generally hollow cylindrical body formed from a matrix material impregnated with a reinforcement material substantially composed of fibers, the fibers being oriented in a generally circumferential direction about a longitudinal axis of the bush. The body may define a plurality of corrugations extending between an inner and an outer diameter of the bush to improve the compressibility of the bush in a direction substantially collinear to a longitudinal axis of the bush.

CROSS-REFERENCES TO RELATED APPLICATIONS

This application claims the benefit of the Great Britain Patent Application No. 1 719 791.4 filed on Nov. 28, 2017, the entire disclosure of which is incorporated herein by way of reference.

FIELD OF INVENTION

The present invention relates to a curable composite bushing for an aircraft joint, an aircraft joint, an aircraft structural assembly incorporating such an aircraft joint and a method and tool for installing a curable composite bushing.

BACKGROUND TO THE INVENTION

During aircraft structural assembly, it is known to attach two or more components together at a joint by drilling and installing one or more fasteners in predetermined hole positions after the components are mounted at a fixed position relative one another. Once the fasteners are installed the joint is made and a larger structural assembly is formed. It is known that solid bushings or bushes can be used during the assembly of structural joints and assemblies. They are used when unintended eccentricity or gaps exist between a fastener and a corresponding hole into to which the fastener is placed, needs to be removed in order to ensure correct fit of the fastener. Normally a solid bush is machined from a blank or selected from a range of pre-machined solid bushes and then fitted with an interference fit within the existing bore. The solid bush may then be drilled in the desired position to provide a corrected bore into which the bolt is installed.

In the assembly of aircraft structures, machining and drilling of made to order solid bushing parts to the required accuracy is a precision process and therefore takes time and requires a stop in the assembly, therefore increasing overall costs of the assembly process.

Sometimes the prepared solid bushing or the hole may not correspond to the dimensions of the hole exactly due to hidden irregularities in the surface of the hole itself, which results in a poor seating of the fastener when installed in the hole. Incorrect seating of the fastener in the hole results in the fastener not transferring the applied load fully when the aircraft is in operation, which has static and fatigue implications. Therefore, an aircraft structure may be designed with a conservative assumption that a certain number of fasteners in a structural assembly would be incorrectly seated fasteners. This conservative approach ultimately leads to a design with a higher number of fasteners required to transfer a given load, thus leading to a higher degree of redundancy, however, the structural assembly will also be more costly to manufacture, heavier and require more maintenance due to the increased attachment part count. Furthermore, in some cases the hole may be so irregular that redundancy will not suffice for the degree of incorrect seating of the fastener, leading to the hole and the solid bushing needing to be reworked and or even the component to be scrapped. In addition, the use of solid bushings of various sizes requires storage and asset management, which takes up space of the assembly floor and resources. Furthermore, the process of fitting a solid bush into the existing bore may itself result in damage to the assembly, particularly in the case of composites. This again may require rework which can be time consuming and therefore can increase costs.

An object of the present invention is therefore to provide a bushing and an aircraft joint incorporating the bushing that is more efficient to build, cheaper to use and less likely to cause damage. A further object is to provide an aircraft structure assembly incorporating one or more joints with bushing in order to decrease assembly time of the structural assembly.

Another object of the present invention is to provide a tooling configured to install a bushing according to the object of the technology previously described. Lastly, it is also an object of the present invention to provide a standard sized bushing that is adaptable to fit multiple dimensions of a hole without machining, and furthermore, a standard bushing that is quicker to install and easier to store than those previously known.

SUMMARY OF THE INVENTION

According to at least one embodiment of the present invention, there is provided a curable composite bush for an aircraft joint comprising a generally hollow cylindrical body formed from a matrix material impregnated with a reinforcement material substantially composed of fibers, the fibers being oriented in a generally circumferential direction about a longitudinal axis of the bush. The body of the bush may define a plurality of corrugations extending between an inner and an outer diameter of the bush, and wherein the corrugations improve the compressibility of the bush in a direction substantially collinear to a longitudinal axis of the bush. The corrugations may be substantially chevron shaped in an uncompressed state. The corrugations may be substantially sinusoidal shaped in an uncompressed state. One or more of the corrugations may extend in the form of a helix along the length of the curable composite bush. Furthermore, the cross section of the bush may vary along the length of the bush.

The matrix material of the composite bush may substantially comprise a thermoplastic type of material. The thermoplastic matrix material may be polyether ether ketone. Alternatively, the matrix material may comprise a thermoset type of material.

The reinforcement fibers may be substantially continuous or substantially discontinuous and the reinforcement material may comprise fibers of more than one type of reinforcement fiber. In addition, the reinforcement fibers may be evenly distributed throughout the body of the composite bush. The reinforcement fibers may be formed from a glass fiber material, such as E-glass fiber material. The reinforcement fibers may be formed from a graphite fiber material or an aramid fiber material. A composite bush may be provided with a half-length of the corrugation in an uncompressed state between 0.5 mm and 3 mm. The inner diameter of the bush in an uncompressed state may be between 4 mm and 30 mm. The inner diameter of the bush may be approximately 10 mm and the thickness may be approximately 0.5 mm.

Advantages of the present invention will now become apparent from the detailed description with appropriate reference to the accompanying drawings.

DETAILED DESCRIPTION

With reference toFIG. 1, an aircraft101is shown with a wing103(also referred to as an airfoil) extending approximately horizontally through a fuselage105. A further wing107(also referred to as the horizontal tail plane) extends approximately horizontally from either side of a rear portion of the fuselage105. Yet a further wing109(also known as the vertical tail plane) extends vertically from an upper rear portion of the fuselage105.

The wings103,109,107and the fuselage105may be formed of structural assemblies joined together by structural joints which function to restrain at least a pair of assemblies from moving relative to one another once the joints have been fastened, in other words the pair of assemblies are fixedly attached to one another. Each structural assembly may itself be formed of smaller structural assemblies joined together by a further number of joints.

The wings103,107,109each have a leading edge111and a trailing edge113, and a set of movable high-lift or control elements115, such a slats, flaps, ailerons, rudders and elevators, which are moveable (i.e., non-fixed) devices actuatable during operation between a deployed position and a retracted position according to the inputs of a controller. Adjacent, above or underneath each movable control element115and/or in the areas where no movables control elements115are provided, the leading edge and trailing edge structure of the wing103is fixed, i.e., not moving to function as a movable control element115.

Geometrical characteristics of the aircraft101and its structural assemblies may be described with reference to a set of orthogonal principal aircraft axes X, Y and Z. The longitudinal axis (X) has its origin at the center of gravity117of the aircraft101and extends lengthwise through the fuselage105from the nose to the tail in the normal direction of flight. The lateral axis (Y) also has its origin at the center of gravity117and extends substantially crosswise from wingtip to wing tip. The vertical or normal axis (Z) also has its origin at the center of gravity and passes vertically through the center of gravity. A further pair of local axes are defined with reference to aircraft principal axes X, Y, Z for any portion of a given structural assembly. In the present example, a set of local axes X′, Y′, Z′ is defined at a fixed leading edge structural assembly104located in proximity to the leading edge111of the wing103. The first axis Y′ lies in parallel to the principal XY plane at a swept angle (commonly referred to as the wing sweep angle) from the principal Y axis. This is may be referred to as the spanwise axis. The second axis X′ again lies in parallel to the same XY plane and is perpendicular to the Y′ axis at the point of its origin, which can be selected at any spanwise position, but in this case at a lower portion of the fixed leading edge structural assembly104.

With reference toFIG. 2, and according to an embodiment of the present technology, a section view is shown of a fixed leading edge assembly104at a spanwise A-A ofFIG. 1after completion of an assembly process.

The fixed leading edge assembly104comprises a first structural component201that is joined to a second structural component205. The first and second components are spanwise extending in the Y′ dimension over substantially the full span of the wing103and are manufactured from composite CFRP (carbon fiber reinforced polymer) material. At the position shown inFIG. 2the components201and205are joined by a pair of joints200A (upper),200B (lower); however, it should be appreciated that further joints substantially the same as joints200A,200B are similarly provided at spanwise intervals along the spanwise length of the first and second components201,205. In the present example the first component201is a unitary, preassembled, structural module called a modular leading edge assembly and the second component205is a unitary wing box front spar member205, however it should also be appreciated that the components201,205may represent any plurality of structural components to be joined together to form any given structural assembly. It should also be appreciated that only one or more joints may be used for a given structural assembly, depending on the expected load transfer or failure design principle applied in the design of the assembly in question, e.g., Safe-life or fail-safe (multiple load path) design principles. Preferably the types of joints200A,200B are used for highly loaded structural assemblies similar to the fixed leading edge structural assembly104shown, for example, between an attachment component for a movable, or at the joint between the wings103,107,109and the fuselage105.

In the present embodiment, at the location of each joint200A200B, the first component201is provided with an attachment hole202that is configured to receive a corresponding attachment fastener203provided by the second component205, the attachment fastener203having a longitudinal fastener axis208. Typically, the outer diameter Dfo of attachment fasteners203in such structures varies between 4 and 30 mm, however in the present embodiment of a modular leading edge assembly104, the outer diameter Dfo is approximately 10 mm.

In the design of the assembly104and joint200A/200B shown, the outer diameter Dfo is smaller in dimension than the diameter Dh of the hole202, such that a first joining stage of the assembly process, when the first component201is first mounted to the second component205, a gap501(seeFIG. 5Afor more detail) exists between the fastener203and the hole202.

Due to the design of the components201,205but also due to the design of the manufacturing process, dimensions of the gap501may be exacerbated due eccentricity (e.g., the degree of non-circular dimension) of the hole202as a result of incorrect forming of the hole202during manufacture of the first component201or as a result of the hole202being formed an incorrect position in the first component201relative to the position of the fastener203. This is commonly referred to as an “out of tolerance” hole202. Typically, the cross-sectional dimension of the gap can be between 0.5 mm and 6 mm, depending on the size of the hole202and its eccentricity.

InFIG. 2, the gap501is not apparent because it is occupied completely by a curable composite bush207. The curable composite bush207is installed in a substantially pliable, uncured state between the hole202and the fastener203when the first and second components201,205are in a desired fixed installation position relative to one another, such that the bush207is able to comply with the surfaces of the hole202and the fastener203(regardless of how irregular their shape is) and fill the gap501.

Once this is achieved, the curable composite bush207is then cured to an extent such that its material stiffness is increased so as to prevent radial displacement of the attachment fastener203within the attachment hole202and to provide optimum seating and load transfer between the fastener203and the first component201and second component205. A nut215and a washer213and are installed on the fastener203at each joint200A,200B such that the washer213engages the surface of the first component201and the cured bush207. The nut215is threadably engaged and torqued so that a load bearing joint is formed between the first component201and the second component205.

Use of a joint according to the present technology that incorporates a curable composite bush207in the way described allows for a structural assembly, and a manufacturing or assembly process that does not require the installation and machining of solid bushings, which is advantageous. It also ensures that each structural fastener is fully load carrying, which may also permit less conservative static load and fatigue load design assumptions in the design of the assembly, leading to a lighter structural assembly design of lower part count or of lower manufacturing and maintenance cost.

Use of such joints200A,200B, may further enable structural design and assembly philosophies that are more efficient and more tolerant for gaps501that may exist between the attachment elements202,203, of structural components201,205. It should be appreciated that the structural design or assembly104may incorporate a joint200A,200B, according to the present technology by design, rather than as a remedy to incidental out of tolerance holes. This may particularly beneficial to achieve a high rate manufacture of structural assemblies104particularly those where the type of material used for the components201,205, the dimensions of components201,205, or the number of components201,205to be joined inherently leads to a high variance and incidents of gaps501that need to be tolerated in the design of the joint200A,200B. For example, this is particularly preferred in the assembly of large modular leading edge assemblies104, as described with reference toFIG. 2.

Each joint200A,200B further comprises a third structural component209in the form of a solid donut shaped spacer of equal thickness in the X direction, and manufactured from CFRP composite material. The third component209may be manufactured using any suitably alternative material such as GFRP (glass fiber reinforced polymer), polymer, or metallic alloy material.

At each joint200A,200B, each third structural component209is configured to displace the first structural component201from the second structural component205in the +/−X′ direction. It may also be configured to reduce the contact surface area between first component201and second component205such that load transfer is substantially provided through the attachment fasteners203. It may be preferable that each component209may be machined or fettled to correct any misalignment between the first component201and the second component205in the direction mentioned. The third component209of each joint200A,200B may alternatively not be of the same dimension and that the third component209of each joint200A,200B may be machined differently, in order to achieve a desired correction in alignment between the first component201and the second component205. As shown the third component209is provided with a hole210configured to receive a portion of the attachment fastener203, such that it is supported in a radial direction i.e., perpendicular to the fastener axis208, however such support may not be necessary.

With reference toFIG. 3, an exemplary curable composite bush207for use in either of the joints200A,200B ofFIG. 2is shown. The bush207is shown in an uncured state. The bush207is formed by an open ended, elongate, corrugated and generally cylindrical body301. The corrugations307of the present embodiment are substantially chevron shaped. The body301further has a central longitudinal axis304that defines a center point for cross sections of the body301taken perpendicularly along the longitudinal axis304. The body301is formed from a composite material composed of reinforcement material303pre-impregnated with a matrix material302.

With reference toFIG. 4A, a cross-section of the body301is shown. The cross-section lies on a plane formed by the longitudinal axis304and a line perpendicular to the axis304.

The body301has a thickness t that is constant along its length L1. The thickness t of the present embodiment is approximately 0.5 mm. The reinforcement material303is formed of continuous fibers that are evenly distributed through the thickness t of the body301and along the length L1of the body301. The fibers are orientated substantially concentrically about the longitudinal axis304of the body301, and are generally aligned in parallel with a circumference305of the body301at any cross-section of the body301viewed on a plane that is perpendicular to the axis304.

A pair of exemplary magnified views V1, V2is provided in order to demonstrate the distribution of the matrix material302and reinforcement material303along the length and thickness t of the body301, as well as the orientation of the reinforcement material302. The, ends of the fibers can be seen as dots, which is representative of their orientation.

The body301is of continuous cross-section with constant dimension values of an outer diameter Do and an inner diameter Di, respectively. However, the body301may vary in cross-sectional dimensions along its length L1. For example, the body301may be tapered from one end to the other end. The cross-section used may alternatively be elliptical, square, triangular or any combination thereof, as required to suit the characteristics of the attachment hole.

Chevron shaped corrugations307define the inner diameter Di and the outer diameter Do of the bush207and are configured allow the bush207to be compressed and allow the body301to collapse in the direction of the longitudinal axis304such that the corrugations307overlap in series, when the bush207is compressed along the axis304. The inner diameter Di is dimensioned to substantially the same dimension of the outer diameter Dfo of the attachment fastener203, to which the bush207is to be applied. Chevron shaped corrugations are preferable as they permit the highest packing density and most even distribution of fibers both radially from, and in parallel to, the axis304when the bush207is brought into a compressed state, thus ensuring more even mechanical stiffness properties through the bush207, when it is installed and cured in an attachment hole.

With reference toFIG. 4B, an alternative embodiment of the present technology is shown. A bush207substantially in accordance with the previous embodiment comprises the same length L1, thickness t, material types and inner and outer diameters, Di, and Do. However, in the present embodiment, the corrugations307have a sinusoidal profile. This may be preferable over alternative shapes such as chevron shaped corrugations307, because sinusoidal corrugations307may be less prone to point damages and may be easier to manufacture, even though they may result in a less evenly distributed fibers when the bush207is brought into a compressed state.

Furthermore, the corrugations307are defined continuously at a non-perpendicular angle1to the longitudinal axis, such that the corrugations307extend in the form of a helix along the length L1of the curable composite bush207. This helix may equally be provided for the previously described chevron shaped corrugation or for any suitable alternative corrugation shape that is capable of being provided in a helical 3D form. The helical form may be preferable to enable the bush207to be manufactured using a continuous extrusion process.

With reference toFIG. 4C; a cross-section is shown of the curable composite bush207ofFIGS. 4A and 4Bin a compressed and cured state, which is the state of the bush207that fills the gap shown inFIG. 2. In the cured state ofFIG. 2, the axis208of the fastener203aligns concentrically to the axis304of the bush207, such that the values of the outer diameter Dfo and the inner diameter Di are substantially the same. The corrugations307substantially overlap in series when the bush207is compressed to a compressed length L2, such that the outer diameter Do of the bush207and the diameter Dh of the hole202are also the substantially same dimension. A half-length l, defines the radial dimension of overlap with respect to the longitudinal axis304of the bush, which in the present example is 3 mm.

The matrix material302used in the bush207ofFIGS. 3, 4A and 4Bis Polyetheretherketone (PEEK); a thermoplastic matrix material, however any other suitable thermoplastic matrix material may alternatively be used such as polyethersulfide (PES), polyetherimide (PEI), or polyphenylenesulfide (PPS). For the purpose of this description, a curable composite bush207formed in part from a thermoplastic matrix material302is said to be in a “cured” state, when the temperature of the thermoplastic matrix material302is brought below its applicable melting temperature and the bush207is in a compressed state. It is said to be in an “uncured” state when the temperature of the thermoplastic matrix material302has reached or is above its applicable melting temperature and/or the bush207is in a decompressed or noncompressed state.

The body301may alternatively be formed from a reinforcement material303pre-impregnated with a partially or non-polymerized thermosetting matrix material302, chosen from one of matrix materials commonly used in aerospace such as polyester, epoxy, vinylester, bismaleimide, phenolic or polyimide. For the purpose of this description, a curable composite bush207formed in part from a thermoset matrix material302is said to be in a “cured” state, when the bush207is deformed in a compressed state and the matrix material302polymerized such that bush207is irreversibly deformed.

The reinforcement material303in the present embodiment is composed of continuous glass fibers of alumina borosilicate glass otherwise known as ‘E-glass’, however any other suitable continuous glass fiber reinforcement may be used, for example S-Glass. In addition, a reinforcement material303using other material types may be used such as graphite/carbon type fibers or aramid type of fibers. Continuous fibers are preferable because during compression of the bush207their length ensures that they remain oriented substantially concentrically about the longitudinal axis304of the body301in parallel with a circumference305of the body301, which would not be the case for short fibers, which may re-orientate and ultimately hinder the collapse of the bush207in a desired way. Furthermore, once compressed within the gap501, the continuous fibers will remain generally aligned and will interact so that composite material is evenly distributed between the hole202and fastener203when compressed.

Use of a thermoplastic matrix material302may be preferable as it allows the curable composite bush207to be handled, stored more easily during assembly operations when the matrix material of the bush207is below its applicable melting temperature, as shown inFIGS. 3 and 4A, and is resiliently deformable. In this state, the bush207is less prone to unintentional damage, and can be handled more easily, than a curable composite bush207with a body301formed from a reinforced, partially cured thermoset resin matrix material302, for example.

It may also be advantageous to use a thermoplastic matrix material302as it does not require special storage and shelf life considerations that could be necessary were the body301to be formed using a partially cured thermoset resin matrix material302, which have a pre-determined shelf life and may require cooling to provide a usable shelf-life in an assembly line

Lastly, disassembly of the joint200A,200B may be required at some point in the aircraft's life cycle, and it is foreseen that removal of the bush207by application of heat is easier than application of machining, which would be required for removing a thermosetting type of bush207.

That said, use of a thermosetting matrix material302may in some instances be preferable, particularly where a matrix material with higher mechanical performance properties is required or where a matrix material302with irreversible properties is required due to a high temperature environment of the joint200.

The use of E-glass fibers as a matrix material302may be advantageous as they have higher compression strength properties when compared to alternative high performance fibers, which may alternatively be used in aircraft structure, for example carbon fiber or aramid fiber. Furthermore, glass fibers are galvanically compatible with a wider range of structural materials commonly used in in aircraft joints, for example, titanium or aluminum alloy.

It should be appreciated, that the volume ratio of reinforcement material303to matrix material302may be varied depending on the specific mechanical stiffness and strength properties required from the bush207when it is compressed and cured in a joint200A,200B. Furthermore, mixtures of different type of reinforcement material303including fibers may be used, if required. An uneven distribution of the reinforcement material303may also be used and tailored to suit the principal load direction and levels between the attachment fastener203and attachment hole202.

With reference toFIGS. 5A and 5B, attachment of a first component201to a second component205is shown using a curable composite bush207.

InFIG. 5A, a first joining stage shows the first component201held in an installation position relative to the second component205, such that the attachment fastener203is inserted in the attachment hole202, which creates a gap501between the fastener203and the hole202. A first end303of a curable composite bush207in accordance with the bush described inFIGS. 3, 4A and 4B, is then introduced into the gap501between the outer diameter Dfo of the fastener and the diameter Dh of the hole202, i.e., it is inserted over the exposed end of the fastener203.

A tool500for installing the bush207comprising a guide503and a compactor505is positioned in proximity to the bush207. The guide503and compactor505each comprise a body formed from steel and shaped as an open-ended cylinder. An end surface504, defined by the guide503, is substantially planar and has an inner diameter substantially the same as the diameter Dh of the hole202. The end surface504is configured to comply with an exterior surface of the first component201in proximity to the hole202such that the guide503can be steadily held in position by a user.

An inner surface506of the guide503is configured to be substantially smooth and polished and may comprise a non-stick treatment. The inner surface506is configured so as to contact the bush207in order to ensure that it stays substantially cylindrical in form and orientated in the direction of its longitudinal axis304as the bush207is compressed into the gap501. The inner surface506of the guide503is further configured to guide the compactor505in a direction substantially parallel to the longitudinal axis304of the bush207.

A further end surface508, defined by the compactor505, is substantially planar and has an inner diameter substantially the same as the outer diameter Dfo of the fastener203and an outer diameter substantially the same as the diameter Dh of the hole202. The compactor505is configured to slide within the guide503and to compact the curable composite bush207, at the further end508, into the gap501between the attachment fastener203and an attachment hole202, such that the bush207substantially conforms to the dimensions of the gap501. The use of the compactor505helps to ensure that compression pressure is evenly applied by the compactor505at its further end surface508to the bush207. The combination of the presently described compactor505and guide503is also advantageous as the relative displacement of the compactor505relative to the guide503may be used to determine the degree of compression of the bush207within the gap501. As such the amount of bush material compacted within the gap501can be derived from such measurements.

In the present example, the curable composite bush207is formed partially of thermoplastic material302, therefore heating means are provided in the form of electrical heating elements507embedded with the guide503and separately with the compactor505. Electrical heating elements507may be preferable as they are easier to control than other means, such as liquid heating. Embedding the heating element507avoids any interference between the smooth inner surface506of the guide503and the bush207and reduces the likelihood of collecting contaminants that may be transferred between the tool500and the bush207. Such contaminants are undesirable as they may affect the curing of the bush207and may also pose a fire risk when heating is applied.

It may alternatively be sufficient to attach the heating elements507to a surface portion of the guide503or compactor505. The heating means507may be activated once the bush207is positioned as shown within the gap501, before compacting of the bush207is started. Heating means507may only be needed to be provided in the guide503or the compactor505.

Once the bush207, guide503and compactor505are put into the position shown inFIG. 5A, the compactor505is moved towards the fastener203to compress the bush207into the gap501in a second joining stage, so that no gap501exists thereafter, as shown inFIG. 5B. The curable composite bush207is then held in the compressed state shown inFIG. 5Buntil it is cured. The tooling500is then removed. In the cured state, the bush207prevents radial displacement of the attachment fastener203within the attachment hole202. In the case of installing a bush207partially formed from a thermoplastic matrix material302, the bush207is heated by the heating means507before the compactor505is moved towards the fastener203into order to bring the curable composite bush207into a melted pliable state, after which the compactor505can be moved to compress the bush207into the gap501. In the case of installing a bush207partially formed from a thermosetting matrix material302, heating means507may also be applied in order to accelerate the polymerization of the matrix material, which may be desirable in high joining rate applications.

It should be appreciated for the embodiments so far described that, for certain applications the corrugations307, may not be required and that the material properties of the bush, particularly its pliability, in response to heat application may suffice for feeding a bush into a gap501. With reference toFIGS. 6A and 6B, an exemplary curable composite bush601without corrugations is provided as well as a tool600for installing it. The joint200A/200B and tool600shown inFIGS. 6A and 6Bare substantially in accordance with the previous embodiment shown inFIGS. 5A and 5B, however the bush601is different in form than previously provided and the tool600has an additional second guide603fitted with a deployable radial cutter605.

The curable composite bush601comprises a continuous tube of substantially circular cross-section and has a body301in the form of an un-corrugated cylinder. The body301of the bush601is composed of a thermoplastic matrix material302pre-impregnated with continuous fibers of a reinforcement material303, that are orientated about the longitudinal axis304of the body301, and are generally aligned in parallel with a circumference305of the body301at any cross-section of the body301when viewed on a plane that is perpendicular to the axis304.

InFIG. 6A, in a first joining stage the first component201is shown held in an installation position relative to the second component205, such that the attachment fastener203is inserted in the attachment hole202, which creates a gap501between the fastener203and the hole202. The second guide603is slideably engaged to the fastener hole202. The second guide603is cylindrical with an outer diameter at an end furthest from the fastener202that has a diameter smaller than inner diameter of the compactor505, such that a radial channel607exists between the second guide603and the compactor505. The outer diameter of the second guide603reduces towards the attachment end to the fastener203such that a taper is provided. The second guide is formed from steel, has a smooth, polished and non-stick surface, and is provided with heating means507in the form of a single embedded heating element507. Once the second guide603is attached to the fastener, the guide503and compactor505are then positioned. Then, a first end of a curable composite bush601as described is fed around the second guide603and into the channel607.

With reference toFIG. 6B, once the bush601, guide503and compactor505, and second guide603are put into position shown inFIG. 6A, a second joining stage ensues wherein the bush601is heated to a temperature approximately equal to or above the melting temperature of the thermoplastic material. Simultaneously or thereafter, the bush601is fed between the outer diameter Dfo of the fastener and the diameter Dh of the hole in the channel607towards the fastener203and into the gap501, until sufficient material is introduced to remove the gap501. The compactor505may then be moved into and out of engagement with the compressed curable composite bush601to ensure it has been sufficiently compressed into the gap501. The bush601is then left to “cure,” e.g., cool below its melting point in the compressed state, such that in the cured state shown inFIG. 6B, the bush601prevents radial displacement of the attachment fastener203within the attachment hole202. Once the bush601is cured, the compactor505is withdrawn away from the fastener203and the radial cutter is deployed to cut the unused portion of bush601, before the tool600is then removed.

With reference toFIGS. 5A, 5B, 6A, and 6B, it should be appreciated that the heating means507for the tooling500,600may be provided by other suitable means. A heating means507may instead be provided by a heated gas jet device attached at the othermost end of the guide503. A heated gas jet device may be preferable as it may use less energy and provide a better heat distribution to the curable composite bush207. Alternatively, heating means507may be provided by an ultrasonic energy emitting device attached to the tool. Use of an ultrasonic energy emitting device may be used to direct ultrasonic wave energy to heat the material of the bush207and remove imperfections within it, when the material of the bush207is in a more pliable state.

Alternatively, heating means507may be provided by electrical induction of the fastener203or fibers of the reinforcement material303, using an electrical inducting device. Such an alternative embodiment may be preferable where access to the bush207is particularly problematic.

With reference toFIG. 7, a method700of joining a first aircraft structural component to a second aircraft structural component is provided, the method comprising the steps of:701—holding first component201provided with an attachment hole202in an installation position relative to the second component205provided with an attachment fastener203, such that the attachment fastener203is inserted in the attachment hole202, which creates a gap501between the fastener203and the hole202;703—placing a first end of a curable composite bush207into the gap501;705—providing a tool500for installing the bush207, the tool500comprising one or more guides503,603and a compactor505;707—positioning a guide503in proximity to the bush207so that it encloses the bush207within an inner surface506;709—positioning a compactor505within the guide503and compacting the curable composite bush207, at a further end508of the compactor505, into the gap501between the attachment fastener203and an attachment hole202such that the bush207substantially conforms to the dimensions of the gap501so that no gap501exists thereafter;711—holding the composite bush207in the compressed state until it is cured.

The method may further comprise the steps of:713—heating the curable composite bush207with heating means507before, or during or after the compactor505is moved towards the fastener203;715—moving the compactor505into and out of engagement with the compressed curable composite bush207to ensure the curable composite bush601is sufficiently compressed into the gap501;717—providing a second guide603and slideably engaging it a fastener203so as to create a radial channel607between the second guide603and the compactor505, into which a curable composite bush601may be fed; and719—engaging a radial cutter to remove any excess curable composite bush601.