Aircraft joint, aircraft component, installation tool, and method for joining structural components

Abstract

The present disclosure provides an aircraft joint. The aircraft joint includes: a first structural member having an attachment hole configured to receive a corresponding attachment fastener provided by a second structural member; a curable composite bushing that is compressed and cured between the outer diameter of the attachment fastener and the attachment hole when the first member is installed to the second member, wherein the curable composite bushing prevents radial displacement of the attachment fastener within the attachment hole when cured. The present disclosure also provides an aircraft assembly including the above aircraft joint, a tool for installing the composite bushing, and a method for joining a first aircraft structural member to a second aircraft structural member.

Description

Technical Field

[0001] The present technology relates to a curable composite bushing for an aircraft joint, an aircraft joint, an aircraft structural component including the aircraft joint, and a method and tool for installing the curable composite bushing. Background Art

[0002] During the assembly of an aircraft structure, it is known to attach two or more components together at a joint by drilling holes and installing one or more fasteners in predetermined hole locations after the two or more components are mounted relative to each other in fixed positions. Once the fasteners are installed, the joint is made and a larger structural component is formed. It is known that solid bushings can be used during the assembly of structural joints and components. Solid bushings are used when there is an undesirable eccentricity or clearance between the fastener and the corresponding hole in which the fastener is to be disposed and this undesirable eccentricity or clearance is to be eliminated to ensure proper fitting of the fastener. Typically, solid bushings are machined from blanks or selected from a series of pre-machined solid bushings and then press-fitted into existing holes by interference fit. The solid bushing can then be drilled at a desired location to provide a proper hole for bolt installation.

[0003] In the assembly of an aircraft structure, machining and drilling a custom solid bushing component to the required precision is an exacting process and thus takes time and requires stops in the component, thereby increasing the total cost of the assembly process.

[0004] Sometimes, due to hidden irregularities in the surface of the hole itself, the prepared solid bushing or hole may not exactly correspond to the size of the hole, which results in improper seating of the fastener when the fastener is installed in the hole. Improper seating of the fastener in the hole results in the fastener not being able to fully transmit the applied load during aircraft operation, thereby causing static fatigue effects. Therefore, the aircraft structure can be designed on the conservative assumption that a certain number of fasteners in the structural component are improperly seated fasteners. This conservative approach ultimately results in a design that requires a greater number of fasteners to transmit a given load, thereby resulting in a higher degree of redundancy. However, due to the increased number of attachment components, the structural component will also cost more to manufacture, will be heavier and will require more maintenance. In addition, in some cases, the hole may be so irregular that the redundancy will not compensate for the degree of improper seating of the fastener, resulting in the hole and the solid bushing needing to be reworked or even resulting in the component being scrapped. Additionally, using solid bushings of various sizes requires storage and asset management, which takes up space on the assembly floor and resources. Further, the process of fitting a solid bushing into an existing hole may itself cause damage to the component, especially in the case of composite materials. This may again require rework, which can be time-consuming and thus may increase costs.

[0005] Accordingly, it is an object of the present technology to provide a bushing that is more effectively constructed, less expensive to use, and less likely to cause damage, and an aircraft joint including such a bushing. Another object is to provide an aircraft structural assembly including one or more bushed joints to reduce the assembly time of the structural assembly.

[0006] Another object of the present technology is to provide a tool configured to install a bushing according to the object of the present technology described above. Finally, an object of the present technology is to provide a standard-sized bushing that can fit holes of multiple sizes without machining, and also to provide a standard bushing that can be installed faster and stored more easily than those previously known. Summary of the Invention

[0007] According to an embodiment of the present technology, there is provided an aircraft structural joint, the aircraft structural joint including: a first structural member provided with an attachment hole configured to receive a corresponding attachment fastener provided by a second structural member; a curable composite bushing that is compressed and cured between the outer diameter of the attachment fastener and the attachment hole when the first member is mounted to the second member, wherein the curable composite bushing prevents radial displacement of the attachment fastener within the attachment hole. The curable composite bushing may include a generally hollow cylindrical body formed of a matrix material pre-impregnated with reinforcing fibers. The reinforcing fibers of the curable composite bushing may be oriented in a generally circumferential direction around the longitudinal axis of the bushing, wherein the body defines a plurality of corrugations extending between the inner diameter and the outer diameter of the bushing, and wherein the corrugations improve the compressibility of the bushing in a direction generally collinear with the longitudinal axis of the bushing. The aircraft joint may further include a third structural member between the first structural member and the second structural member, wherein the third structural member is configured to displace the first structural member relative to the second structural member, and the third structural member is further provided with a hole configured to receive a portion of a protruding fastener element. One or more joints may be provided in the aircraft structural assembly. The aircraft structure may include a first member as a leading edge assembly and a second member as a wing box beam element.

[0008] Another embodiment of the present technology provides a tool for installing a composite bushing. The tool includes a pressing member configured to press a curable composite bushing into a gap between an attachment fastener and an attachment hole such that the bushing generally conforms to the dimensions of the gap. The tool may also include one or more guides configured to engage an outer surface of a first structural member adjacent to the hole or fastener, and the guides are further configured to guide the curable composite bushing into the gap. The guides may also be configured to guide the pressing member to a pressing position. The tool may further include a heating device configured to apply heat energy to the curable composite bushing when the curable composite bushing is being guided or pressed. The heating device may be mounted in the guide or the pressing member. The heating device may be provided by an electric heating element, by a hot gas injection device, by an ultrasonic energy emitting device, or by an electric induction device.

[0009] Another embodiment of the present technology provides a method of joining a first aircraft structural member to a second aircraft structural member. The method includes the steps of: holding a first member having an attachment hole in an installation position relative to a second member having an attachment fastener such that the attachment fastener is inserted into the attachment hole, thereby creating a gap between the fastener and the hole; placing a first end of a curable composite bushing in the gap; providing a tool for installing the bushing, the tool including one or more guides and a pressing member; positioning the guides adjacent to the bushing such that the guides enclose the bushing within an inner surface; positioning the pressing member within the guides and pressing the curable composite bushing at the other end of the pressing member into the gap between the attachment fastener and the attachment hole such that the bushing generally conforms to the dimensions of the gap, such that there is no gap thereafter; holding the composite bushing in a compressed state until the composite bushing is cured. The method may further include the step of heating the curable composite bushing with a heating device before, during, or after the pressing member moves toward the fastener. The method may further include the steps of providing a second guide and slidably engaging the second guide with the fastener to create a radial passage between the second guide and the pressing member, and feeding the curable composite bushing into the passage. The method may further include the step of moving the pressing member into and out of engagement with the compressed curable composite bushing to ensure that the curable composite bushing is fully compressed into the gap.

[0010] Now, the advantages of the present technology will become apparent through a detailed description with due reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0011] Embodiments of the present technology will now be described by way of example only with reference to the following drawings, in which:

[0012] Figure 1 A plan view of a commercial aircraft 101 is shown.

[0013] Figure 2 shows a cross-sectional view of the wing structure assembly 103 taken at the indicated cross-section A-A position, and also shows a pair of identical aircraft joints 200A and 200B. Figure 1

[0014] Figure 3 shows an isometric perspective view of the curable composite bushing 207.

[0015] Figure 4A and Figure 4B shows Figure 3 a schematic cross-sectional view of an embodiment of the curable composite bushing 207 taken along the longitudinal axis 301 of the bushing 207, where the bushing 207 is in an uncompressed state.

[0016] Figure 4C shows Figure 3 , Figure 4A and Figure 4B a schematic cross-sectional view of an embodiment of the curable composite bushing 207 taken along the longitudinal axis 301 of the bushing 207, where the bushing 207 is in a compressed state.

[0017] Figure 5A and Figure 5B respectively show the aircraft joint 200 of Figure 2 in the first engagement stage and the second engagement stage using the curable composite bushing 207 and an installation tool.

[0018] Figure 6A and Figure 6B respectively show the aircraft joint 200 of Figure 2 in the first engagement stage and the second engagement stage using another type of curable composite bushing 207 and an installation tool.

[0019] Figure 7 shows a method of forming the aircraft joint 200 of the structural assembly 103. DETAILED DESCRIPTION

[0020] Referring to Figure 1 , an aircraft 101 is shown. The aircraft 101 has a wing 103 (the wing is also referred to as an airfoil) that extends generally horizontally through the fuselage 105. Another wing 107 (also referred to as a horizontal stabilizer) extends generally horizontally from both sides of the rear portion of the fuselage 105. Yet another wing 109 (also referred to as a vertical stabilizer) extends vertically from the upper rear portion of the fuselage 105.

[0021] The wings 103, 109, 107 and the fuselage 105 may be formed of structural components joined together by structural joints that are configured to limit relative movement between at least a pair of components with respect to each other once the joints are fastened, in other words, the pair of components are fixedly attached to each other. Each structural component itself may be formed of smaller structural components joined together by a further number of joints.

[0022] Each of the wings 103, 107, 109 has a leading edge 111 and a trailing edge 113 and a set of movable high-lift or control elements 115 such as slats, flaps, ailerons, rudders, and elevators, which are movable (i.e., non-fixed) devices that are capable of being actuated during operation between a deployed position and a retracted position in response to an input from a controller. In the regions adjacent to, above, or below each movable member 115 and / or in regions where no movable member 115 is provided, the leading-edge structure and the trailing-edge structure of the wing 103 are fixed, i.e., do not move to act as the movable member 115.

[0023] The geometry of the aircraft 101 and its structural components may be described with reference to a set of orthogonal primary aircraft axes X, Y, and Z. The origin of the longitudinal axis (X) is located at the center of gravity of the aircraft 117 and extends longitudinally from the nose through the fuselage 105 to the tail in the normal flight direction. The origin of the lateral axis (Y) is also located at the center of gravity 117 and extends generally laterally from wingtip to wingtip. The origin of the vertical axis or normal axis (Z) is also located at the center of gravity and passes vertically through the center of gravity. For any portion of a given structural component, another pair of local axes is defined with reference to the primary aircraft axes X, Y, Z. In the present example, a set of local axes X', Y', Z' is defined at a fixed leading-edge structural component 104 located near the leading edge 111 of the wing 103. The first axis Y' is parallel to the primary XY plane and is at a sweep angle with respect to the primary Y axis (the sweep angle is commonly referred to as the wing sweep angle). The first axis Y' may be referred to as the spanwise axis. The second axis X' is also parallel to the same XY plane and is perpendicular to the Y' axis at its origin, which may be selected at any position in the spanwise direction, but in this case is selected at the lower portion of the fixed leading-edge structural component 104.

[0024] Referring Figure 2 to Figure 1 and in accordance with an embodiment of the present technology, a cross-sectional view of the fixed leading-edge assembly 104 at A-A in the spanwise direction after completion of the assembly process is shown.

[0025] The fixed leading-edge assembly 104 includes a first structural member 201 that is joined to a second structural member 205. The first and second members extend in the Y' dimension in the spanwise direction across substantially the entire span of the wing 103, and the first and second members are made of a composite CFRP material. In Figure 2At the locations shown, components 201 and 205 are joined by a pair of joints 200A (upper) and 200B (lower); however, it should be understood that along the spanwise length of the first component 201 and the second component 205, other joints substantially the same as joints 200A and 200B are similarly provided at certain intervals in the spanwise direction. In this example, the first component 201 is a one-piece and pre-assembled structural module, also referred to as a modular leading edge assembly, and the second component 205 is a one-piece wing box front spar member 205; however, it should also be understood that components 201, 205 can represent any number of structural components to be joined together to form any given structural assembly. It should also be understood that only one or more joints can be used for a given structural assembly, depending on the intended load transfer or failure design principles applied in the design of the assembly under discussion, e.g., safe-life or fail-safe (multiple load path) design principles. Preferably, this type of joint 200A, 200B is used for highly loaded structural assemblies similar to the fixed leading edge structural assembly 104 shown, e.g., at the joints between attachment components for movable parts or between wings 103, 107, 109 and the fuselage 105.

[0026] In the present embodiment, at the location of each joint 200A, 200B, the first component 201 is provided with attachment holes 202 configured to receive corresponding attachment fasteners 203 provided by the second component 205, the attachment fasteners 203 having a longitudinal fastener axis 208. Generally, the outer diameter Dfo of the attachment fasteners 203 in such a structure varies between 4 mm and 30 mm; however, in the present embodiment of the modular leading edge assembly 104, the diameter Dfo is approximately 10 mm.

[0027] In the design of the shown assembly 104 and joints 200A / 200B, the size of Dfo is smaller than the diameter Dh of the hole 202 such that at the first joining stage of the assembly process when the first component 201 is initially installed to the second component 205, there is a clearance 501 between the fastener 203 and the hole 202 (see in more detail Figure 5A ).

[0028] Due to the design of components 201, 205 but also due to the design of the manufacturing process, the size of the clearance 501 can vary due to the following reasons: eccentricity (e.g., the degree of non-circular dimension) of the hole 202 due to incorrect formation of the hole 202 during the manufacture of the first component 201 or due to the position of the hole 202 relative to the fastener 203 being formed in an incorrect position in the first component 201. This is generally referred to as a hole 202 that is "out of tolerance". Generally, depending on the size of the hole 202 and its eccentricity, the cross-sectional dimension of the clearance can be between 0.5 mm and 6 mm.

[0029] In Figure 2 it, the gap 501 is not apparent because the gap is completely occupied by the curable composite bushing 207. The curable composite bushing 207 is installed in an approximately flexible uncured state between the hole 202 and the fastener 203 when the first member 201 and the second member 205 are in the desired fixed mounting positions relative to each other, such that the bushing 207 can conform to the surfaces of the hole 202 and the fastener 203 (however irregular their shapes may be) and fill the gap 501.

[0030] Once this is achieved, the curable composite bushing 207 is then cured to such an extent that its material stiffness increases to prevent the attached fastener 203 from radially displacing within the attachment hole 202 and to provide optimal seating and load transfer between the fastener 203 and the first member 201 and the second member 205. Nuts 215 and washers 213 are installed on the fasteners 203 at each joint 200A and 200B such that the washers 213 engage the surfaces of the first member 201 and the cured bushing 207. The nuts 215 are threadedly engaged and torqued such that a load-bearing joint is formed between the first member 201 and the second member 205.

[0031] The use of joints incorporating the curable composite bushing 207 in the manner described according to the present technology enables a structural assembly, manufacturing process, or assembly process that does not require the installation and machining of solid bushings, which is advantageous. This also ensures that each structural fastener is fully load-bearing, which can also permit less conservative static and fatigue load design assumptions in the component design, resulting in a lighter structural assembly design with fewer component parts or lower manufacturing and maintenance costs.

[0032] The use of such joints 200A, 200B can also enable a more effective and greater tolerance structural design and assembly concept for the gap 501 that may exist between the attachment elements 202, 203 of the structural members 201, 205. It should be understood that the structural design or assembly 104 can incorporate the joints 200A, 200B by design according to the present technology, rather than as a remedy for accidentally out-of-tolerance holes. This is particularly beneficial for achieving the high-speed manufacturing of structural assemblies 104, especially the following structural assemblies: in which the type of material of the members 201, 205 to be internally joined, the dimensions of the members 201, 205, or the number of members 201, 205 result in a large difference and incidence of the gap 501 that needs to be tolerated in the design of the joints 200A, 200B. For example, this is particularly preferred in the assembly of large modular leading edge assemblies 104 as described with reference to Figure 2 as described.

[0033] Each joint 200A, 200B also includes a third structural member 209, which is a solid annular spacer having equal thickness in the X direction and is made of CFRP composite material. The third member 209 can be manufactured by using any suitable alternative material, such as GFRP, polymer or metal alloy material.

[0034] At each joint 200A, 200B, each third structural member 209 is configured to displace the first structural member 201 relative to the second structural member 205 in the + / -X' direction. The third structural member 209 can also be configured to reduce the contact surface area between the first member 201 and the second member 205, such that the load transfer is provided substantially by the attachment fasteners 203. Preferably, each member 209 can be machined or trimmed to correct any misalignment between the first member 201 and the second member 205 in the mentioned direction. Alternatively, the third members 209 of each joint 200A and 200B can have different sizes, and the third members 209 of each joint 200A, 200B can be machined in different ways to achieve the required correction of the alignment between the first member 201 and the second member 205. As shown, the third member 209 is provided with a hole 210, which is configured to receive a portion of the attachment fastener 203, such that the attachment fastener 203 is supported in the radial direction, i.e., the direction perpendicular to the fastener axis 208, but such support may not be necessary.

[0035] Refer to Figure 3 , an exemplary curable composite bushing 207 for either of the joints 200A and 200B for Figure 2 is shown. The bushing 207 is shown in an uncured state. The bushing 207 is formed by an end-open, elongated, corrugated and generally cylindrical body 301. The corrugated portion 307 of the present embodiment is generally V-shaped. The body 301 also has a central longitudinal axis 304, and the longitudinal axis 304 defines the center point of the cross-section of the body 301 taken perpendicular to the longitudinal axis 304. The body 301 is formed of a composite material, which consists of a reinforcing material 303 pre-impregnated with a matrix material 303.

[0036] Refer to Figure 4A , a cross-section of the body 301 is shown. The cross-section lies in a plane formed by the longitudinal axis 304 and a line perpendicular to the axis 304.

[0037] The body 301 has a constant thickness t along its length L1. The thickness of the present embodiment is approximately 0.5 mm. The reinforcing material 303 is formed of continuous fibers, and the continuous fibers are uniformly distributed along the length L1 of the body 301 in the thickness t of the body 301. The fibers are oriented substantially concentrically about the longitudinal axis 304 of the body 301, and the fibers are aligned generally parallel to the circumferential direction 305 of the body 301 at any cross-section of the body 301 observed in a plane perpendicular to the axis 304.

[0038] A pair of exemplary enlarged views V1 and V2 are provided to show the distribution of the matrix material 302 and the reinforcing material 303 along the length and thickness t of the body 301 and the orientation of the reinforcing material 302. The ends of the fibers can be regarded as points 303, which represent their orientation.

[0039] The body 301 has a continuous cross-section, and the continuous cross-section correspondingly has constant dimension values Do and Di. However, the cross-sectional dimensions of the body 301 can vary along its length. For example, the body 301 can taper from one end to the other end. As needed, the cross-section used can alternatively be elliptical, square, triangular, or any combination thereof to accommodate the characteristics of the attachment hole.

[0040] The V-shaped corrugations 307 define the inner diameter Di and the outer diameter Do of the bushing 207 and are configured to allow the bushing 207 to be compressed and allow the body 301 to be telescoped in the direction of the longitudinal axis 304 such that the corrugations 307 overlap in series when the bushing 207 is compressed along the axis 304. Di is sized to be substantially the same as the size of Dfo of the attachment fastener 203 to which the bushing 207 is to be applied. The V-shaped corrugations are preferred because the V-shaped corrugations allow for the maximum compaction density and the most uniform fiber distribution of the fibers in both the radial direction relative to the axis 304 and the direction parallel to the axis 304 when the bushing 207 enters the compressed state, thereby ensuring more uniform mechanical stiffness characteristics of the bushing 207 when the bushing 207 is installed in the attachment hole and cured.

[0041] Referring to Figure 4B , an alternative embodiment of the present technology is shown. The bushing 207 that is substantially consistent with the foregoing embodiment has the same length L1, thickness t, material type, and inner diameter Di and outer diameter Do. However, in the present embodiment, the corrugations 307 have a sinusoidal profile. This may be preferred compared to alternative shapes such as the V-shaped corrugations 307 because the sinusoidal corrugations 307 may be less susceptible to point damage and may be easier to manufacture, although the sinusoidal corrugations will result in a less uniform distribution of the fibers when the bushing 207 is brought into the compressed state.

[0042] In addition, the corrugated portion 307 is continuously defined at an angle Φ that is non-perpendicular to the longitudinal axis such that the corrugated portion 307 extends in a helical form along the length L1 of the curable composite bushing 207. Such a helix can equally provide for the previously described V-shaped corrugations or any suitable alternative corrugated portion that can be provided in a helical 3D form. The helical form may be preferred such that the bushing 207 can be manufactured using a continuous extrusion process.

[0043] Reference is made Figure 4C to a cross-section of the curable composite bushing 207 of Figure 4A and Figure 4B in a compressed and cured state, where the compressed and cured state is the state of the bushing 207 filling the gap as shown in Figure 2 . In the cured state of Figure 2 , the axis of the fastener 208 is concentrically aligned with the axis 304 of the bushing 207 such that the values of Dfo and Di are substantially the same. The corrugated portion 307 substantially continuously overlaps when the bushing 207 is compressed to the compressed length L2 such that Do and Dh are also substantially the same size. The half-length l defines the radial dimension of the overlap portion relative to the longitudinal axis 304 of the bushing, and the half-length l is 3 mm in this example.

[0044] In Figure 3 , Figure 4A and Figure 4B , the matrix material 302 used in the bushing 207 is polyetheretherketone (PEEK), which is a thermoplastic matrix material; however, any other suitable thermoplastic matrix material, such as polyethersulfone (PES), polyetherimide (PEI), or polyphenylene sulfide (PPS), may alternatively be used. For the purposes of this description, the curable composite bushing 207 that is partially formed of the thermoplastic matrix material 302 can be referred to as being in a "cured" state when the temperature of the thermoplastic matrix material 302 is below its applicable melting temperature and the bushing 207 is in a compressed state. The composite bushing 207 can be referred to as being in an "uncured" state when the temperature of the thermoplastic matrix material 302 reaches or exceeds its applicable melting temperature and / or the bushing 207 is in a decompressed state.

[0045] Alternatively, the body 301 can be formed from a reinforcing material 303 pre-impregnated with a partially polymerized or non-polymerized thermosetting matrix material 302 selected from among matrix materials commonly used in aerospace, such as polyester, epoxy resin, vinyl ester, bismaleimide, phenolic, or polyimide. For the purposes of this description, a curable composite bushing 207 partially formed from the thermosetting matrix material 302 is said to be in a "cured" state when the bushing 207 is deformed into a compressed state and the matrix material 302 polymerizes such that the bushing 207 is irreversibly deformed.

[0046] The reinforcing material 303 in this embodiment consists of continuous glass fibers of alumina borosilicate glass, also known as "E-glass", but any other suitable continuous glass fiber reinforcing material, such as S-glass, can also be used. Additionally, a reinforcing material 303 using other material types, such as graphite / carbon fibers or aramid fibers, can be employed. Continuous fibers are preferred because during compression of the bushing 207, their length ensures that they remain oriented generally concentrically about the longitudinal axis 304 of the body 301 and parallel to the circumferential direction 305 of the body 301, which would not be the case for short fibers, and this will reorient and ultimately impede the nesting of the bushing 207 in a desired manner. Furthermore, once compressed within the gap 501, the continuous fibers will remain generally aligned and will interact such that the composite material is evenly distributed between the hole 202 and the fastener 203 when compressed.

[0047] The use of a thermoplastic matrix material 302 may be preferred because the thermoplastic matrix material 302 allows the curable composite bushing 207 to be more easily handled and stored during assembly operations when the matrix material of the bushing 207 is below its applicable melting temperature as shown in Figure 3 and Figure 4A and the thermoplastic matrix material 302 is elastically deformable. In this state, for example, the bushing 207 is easier to handle and less susceptible to accidental damage compared to a curable composite bushing 207 having a body 301 formed from a reinforced, partially cured thermosetting resin matrix material 302.

[0048] The use of a thermoplastic matrix material 302 would also be advantageous because the thermoplastic matrix material does not require special storage and storage period considerations, which are necessary for a body 301 to be formed using a partially cured thermosetting resin matrix material 302, which has a predetermined storage period and may require cooling to provide a storage period that can be used on the assembly line.

[0049] Finally, it may be necessary to disassemble the joints 200A, 200B at some point during the life cycle of the aircraft, and it is foreseeable that it will be easier to remove the bushing 207 by applying heat than by applying machining, and machining will be required to remove a thermosetting type of bushing 207.

[0050] That is, in certain cases, particularly in cases where a matrix material with higher mechanical property characteristics is required or in cases where a matrix material 303 with irreversible characteristics is required due to the high temperature environment of the joint 200, it may be preferable to use a thermosetting matrix material 302.

[0051] It may be advantageous to use E-glass fiber as the matrix material 302 because E-glass fiber has higher compressive strength characteristics when compared to alternative high-performance fibers, such as carbon fiber or aramid fiber, which can alternatively be used in aircraft structures. Additionally, glass fiber is compatible in terms of plating with a wider range of structural materials commonly used in aircraft joints, such as titanium or aluminum alloy.

[0052] It should be understood that the volume ratio of the reinforcement material 303 to the matrix material 302 can vary depending on the specific mechanical stiffness and strength characteristics required for the bushing 207 when the bushing 207 is compressed and cured in the joints 200A, 200B. Additionally, mixtures of different types of reinforcement materials 303 containing fibers can be used as needed. Reinforcement materials 303 with non-uniform distribution can also be used and the non-uniform distribution of the reinforcement materials 303 can be customized to accommodate the direction and level of the primary load between the attachment fastener 203 and the attachment hole 202.

[0053] Referring to Figure 5A and Figure 5B , an attachment of a first component 201 to a second component 205 using a curable composite bushing 207 is shown.

[0054] In Figure 5A , a first joining stage shows the first component 201 being held in an installed position relative to the second component 205 such that the attachment fastener 203 is inserted into the attachment hole 202, which creates a gap 501 between the fastener 203 and the hole 202. Then, a first end 303 of the curable composite bushing 207 consistent with the bushing described in Figure 3 , Figure 4A and Figure 4B is introduced into the gap 501 between the outer diameter Dfo of the fastener and the diameter Dh of the hole, i.e., the first end 303 is inserted onto the exposed end of the fastener 203.

[0055] A tool 500 for installing a bushing 207, including a guide 503 and a pressing member 505, is positioned near the bushing 207. The guide 503 and the pressing member 505 each include a body: the body is made of steel and formed as an end-open cylindrical member. An end surface 504 defined by the guide 503 is substantially flat and has an inner diameter substantially the same as the diameter Dh of the hole 202. The end surface 504 is configured to follow the outer surface of the first component 201 near the hole 202, so that the guide 503 can be stably held in place by the user.

[0056] The inner surface 506 of the guide 503 is configured to be substantially smooth and polished and may include a non-stick treatment. The inner surface 506 is configured to contact the bushing 207 to ensure that it remains substantially cylindrical and is oriented in the direction of its longitudinal axis 304 when the bushing 207 is compressed into the gap 501. The inner surface 506 of the guide 503 is also configured to guide the pressing member 505 in a direction substantially parallel to the longitudinal axis 304 of the bushing 207.

[0057] Another end surface 508 defined by the pressing member 505 is substantially flat and has an inner diameter substantially the same as the diameter Dfo of the fastener 203 and an outer diameter substantially the same as the diameter Dh of the hole 202. The pressing member 505 is configured to slide within the guide 503 and press the curable composite bushing 207 into the gap 501 between the attachment fastener 202 and the attachment hole 203 at the other end 508, such that the bushing 207 substantially follows the dimensions of the gap 502. The use of the pressing member 505 helps to ensure that the compression pressure is evenly applied to the bushing 207 by the pressing member 505 at its other end surface 508. The combination of the pressing member 505 and the guide 503 described herein is also advantageous because the relative displacement of the pressing member 505 relative to the guide 503 can be used to determine the degree of compression of the bushing 207 within the gap 501. Therefore, the amount of bushing material compressed within the gap 501 can be derived from such a measurement.

[0058] In this example, the curable composite bushing 207 is partially formed of a thermoplastic material 302, and thus a heating device in the form of an electric heating element 507 that is nested with the guide 503 and separated from the pressing member 505 is provided. The electric heating element 507 may be preferred because it is easier to control than other means, such as liquid heating. Embedding the heating element 507 avoids any interference between the smooth inner surface 506 of the guide 503 and the bushing 207, and reduces the possibility of dirt accumulating between the tool 500 and the bushing 207. These dirt are undesirable because they may affect the curing of the bushing 207 and may also pose a fire risk when heat is applied.

[0059] Alternatively, it may also be sufficient to attach the heating element 507 to the surface portion of the pressing member 505 or the guiding member 503. Before the pressing of the bushing 207 begins, once the bushing 207 is positioned within the gap 501 as shown, the heating device 507 can be enabled. The heating device 507 may only need to be provided in the guiding member 503 or the pressing member 505.

[0060] Once the bushing 207, the guiding member 503, and the pressing member 505 are arranged in the Figure 5A position shown, the pressing member 505 moves towards the fastener 203 to compress the bushing 207 into the gap 501 during the second engagement stage, such that there is no longer a gap 501 thereafter, as Figure 5B shown. The curable composite bushing 207 is then held in the Figure 5B compressed state shown until the curable composite bushing 207 is cured. The tool 500 is then removed. In the cured state, the bushing 207 prevents the attached fastener 203 from radially shifting within the attachment hole 202. In the case of installing a bushing 207 partially formed of a thermoplastic matrix material 302, before the pressing member 505 moves towards the fastener 203, the bushing 207 is heated by the heating device 507 to bring the curable composite bushing 207 into a molten and flexible state. Then, the pressing member 505 can be moved to compress the bushing 207 into the gap 501. In the case of installing a bushing 207 partially formed of a thermosetting matrix material 302, the heating device 507 can also be applied to accelerate the polymerization of the matrix material, which may be desirable in high engagement rate applications.

[0061] It should be understood that for certain applications, the corrugations 307 may not be required in the previously described embodiments, and the material properties of the bushing in response to heat application, particularly its flexibility, may be sufficient to feed the bushing into the gap 501. Referring to Figure 6A and 6B , an exemplary curable composite bushing 601 without corrugations and a tool 600 for installing the curable composite bushing 601 are provided. Figure 6A and 6B The joints 200A / 200B and the tool 600 shown in Figure 5A and 5B are substantially the same as the previous embodiment shown in

[0062] The curable composite bushing 601 includes a continuous tube having a generally circular cross-section, and the curable composite bushing 601 has a body 301 in the form of a cylindrical body without a corrugated portion. The body 301 of the bushing 601 includes a thermoplastic matrix material 302 pre-impregnated with continuous fibers of a reinforcing material 303, the reinforcing material 303 being oriented about a longitudinal axis 304 of the body 301 and arranged generally parallel to the circumferential direction 305 of the body 301 at any cross-section of the body 301 when viewed in a plane perpendicular to the axis 304.

[0063] In Figure 6A In the first joining stage, the first component 201 is shown held in the installed position relative to the second component 205 such that the attachment fastener 203 is inserted into the attachment hole 202, which forms a gap 501 between the fastener 203 and the hole 202. The second guide 603 is slidably engaged to the fastener 202. The second guide 603 is cylindrical and has an outer diameter at its end furthest from the fastener 202 that is smaller than the inner diameter of the compression member 505, such that there is a radial passage 607 between the second guide 603 and the compression member 505. The outer diameter of the second guide 603 decreases towards the end attached to the fastener 202, such that a taper is provided. The second guide is formed of steel and has a smooth, polished and non-stick surface, and the second guide is provided with a heating device 507 in the form of a single embedded heating element 507. Once the second guide 603 is attached to the fastener, the guide 503 and the compression member 505 are positioned. Then, the first end of the curable composite bushing 601 as described above is fed around the second guide 603 and into the passage 607.

[0064] Referring to Figure 6B Once the bushing 601, the guide 503, the compression member 505, and the second guide 603 are arranged in the positions shown in Figure 6A a second joining stage is then carried out in which the bushing 601 is heated to a temperature approximately equal to or higher than the melting temperature of the thermoplastic material. Simultaneously or afterwards, the bushing 601 is fed into the gap 501 in the passage 607 towards the fastener 203 between the outer diameter Dfo of the fastener and the diameter Dh of the hole until sufficient material is introduced to eliminate the gap 501 thereafter. The compression member 505 can then be moved into and out of engagement with the compressed curable composite bushing 601 to ensure that the curable composite bushing 601 is fully compressed into the gap 501. The bushing 601 then waits in the compressed state for "curing", for example cooling below its melting point, such that in Figure 6BIn the cured state shown, the bushing 601 prevents radial displacement of the attachment fastener 203 within the attachment hole 202. Once the bushing 601 is cured, before the tool 600 is removed, the pressing member 505 is withdrawn from the fastener 202 and the radial cutter is deployed to cut the unused portion of the bushing 601.

[0065] Referring Figure 5A , Figure 5B , Figure 6A and Figure 6B , it should be understood that the heating device 507 for the tools 500, 600 can be provided by other suitable means. The heating device 507 can be replaced by a hot gas jet device attached to the other end of the guiding device 503. The hot gas jet device may be preferred as it may use less energy and provide a better heat distribution for the curable composite bushing 207. Alternatively, the heating device 507 can be provided by an ultrasonic energy emitting device attached to the tool. When the material of the bushing 207 is in a more flexible state, the ultrasonic energy emitting device can be used to direct ultrasonic energy to heat the material of the bushing 207 and remove defects within the bushing 207.

[0066] Alternatively, the heating device 507 can be provided by using the electrical inductivity of the fastener 203 or the fibers of the reinforcement 303 with an electrical induction device. This alternative embodiment may be preferred in cases where access to the bushing 207 is particularly problematic.

[0067] Referring Figure 7 , a method of joining a first aircraft structural component to a second aircraft structural component is provided, the method comprising the following steps: 701 - holding a first component 201 having an attachment hole 202 in an installation position relative to a second component 205 having an attachment fastener 203 such that the attachment fastener 203 is inserted into the attachment hole 202, which forms a gap 501 between the fastener 203 and the hole 202; 703 - setting a first end of the curable composite bushing 207 into the gap 501; 705 - providing a tool 500 for installing the bushing 207, the tool 500 including one or more guides 503, 603 and a pressing member 505; 707 - positioning the guide 503 close to the bushing 207 such that the guide 503 encloses the bushing 207 within an inner surface 506; 709 - positioning the pressing member 505 within the guide 503 and pressing the curable composite bushing 207 into the gap 501 between the attachment fastener 202 and the attachment hole 203 at the other end 508 of the pressing member 505 such that the bushing 207 generally conforms to the dimensions of the gap 501, such that there is no longer a gap 501 thereafter; 711 - holding the composite bushing 207 in a compressed state until the composite bushing 207 is cured.

[0068] The method may further include the following steps: 713 - heating the curable composite bushing 207 with a heating device 507 before, during, or after the pressing member 505 moves towards the fastener 203; 715 - moving the pressing member 505 into and out of engagement with the compressed curable composite bushing 207 to ensure that the curable composite bushing 601 is fully compressed into the gap 501; 717 - providing a second guide 603 and slidably engaging the second guide 603 with the fastener 202 to form a radial passage 607 between the second guide 603 and the pressing member 505, into which the curable composite bushing 601 can be fed; and 719 - engaging a radial cutter to remove any excess curable composite bushing 601.

[0069] If, in the foregoing description, reference is made to an integer or component having known, obvious, or foreseeable equivalents, such equivalents are incorporated herein as if set forth individually. The true scope of the present technology should be determined with reference to the claims, which scope should be construed to include any such equivalents. The reader will also understand that the integers or features of the present technology described as preferred, advantageous, convenient, etc. are optional and do not limit the scope of the independent claims. Furthermore, it should be understood that these optional integers or features, although they may be beneficial in some embodiments of the present technology, may be undesirable in other embodiments and may therefore be absent. If the term "or" is used in the foregoing description, unless otherwise expressly stated, the term should be understood to mean "and / or".

Claims

1. An aircraft joint, comprising: A first structural member having an attachment hole configured to receive a corresponding attachment fastener provided by a second structural member; A curable composite bushing that is compressed and cured between the outer diameter of the attachment fastener and the attachment hole when the first structural member is installed to the second structural member, wherein the curable composite bushing prevents radial displacement of the attachment fastener within the attachment hole; Wherein the curable composite bushing includes a hollow cylindrical body formed of a matrix material pre-impregnated with reinforcing fibers, and Wherein the reinforcing fibers of the curable composite bushing are circumferentially oriented about the longitudinal axis of the curable composite bushing, wherein the hollow cylindrical body defines a plurality of corrugations extending between the inner diameter and the outer diameter of the curable composite bushing, and wherein the corrugations enhance the compressibility of the curable composite bushing in a direction collinear with the longitudinal axis of the curable composite bushing.

2. The aircraft joint according to claim 1 further includes a third structural member located between the first structural member and the second structural member, wherein, A third structural member configured to displace the first structural member relative to the second structural member, and the third structural member is further provided with a hole configured to receive a portion of the attachment fastener protruding from the second structural member.

3. An aircraft assembly, comprising one or more aircraft joints according to claim 1 or 2.

4. The aircraft component according to claim 3, wherein, The first structural member is a leading edge assembly, and the second structural member is a wing box beam element.

5. A tool for installing a curable composite bushing in an aircraft joint according to claim 1 or 2, the tool comprising a pressing member configured to press the curable composite bushing into the gap between the attachment fastener and the attachment hole such that the curable composite bushing generally conforms to the dimensions of the gap.

6. The tool according to claim 5, further comprising one or more guides configured to engage the outer surface of the first structural member adjacent to the hole or the fastener, and the guides are further configured to guide the curable composite bushing into the gap.

7. The tool according to claim 6, wherein, The guide is further configured to guide the pressing member to a pressing position.

8. The tool according to claim 6 or 7, further comprising a heating device configured to apply heat energy to the curable composite bushing when the curable composite bushing is being guided or pressed.

9. The tool according to claim 8, wherein The heating device is mounted in the guide or the pressing member.

10. The tool according to claim 8, wherein, The heating device is provided by an electric heating element.

11. The tool according to claim 8, wherein, The heating device is provided by a hot gas injection device.

12. The tool according to claim 8, wherein, The heating device is provided by an ultrasonic energy emitting device.

13. The tool according to claim 8, wherein, The heating device is provided by an electric induction device.

14. A method of joining a first aircraft structural member to a second aircraft structural member, the method comprising the steps of: Holding a first member having an attachment hole in an installation position relative to a second aircraft structural member having an attachment fastener such that the attachment fastener is inserted into the attachment hole, thereby forming a gap between the fastener and the hole; Place the first end of the curable composite bushing in the gap, wherein, The curable composite bushing includes a hollow cylindrical body formed of a matrix material pre-impregnated with reinforcing fibers, and wherein the reinforcing fibers of the curable composite bushing are oriented circumferentially about the longitudinal axis of the curable composite bushing, wherein the hollow cylindrical body defines a plurality of corrugated portions extending between the inner diameter and the outer diameter of the curable composite bushing, and wherein the corrugated portions enhance the compressibility of the curable composite bushing in a direction collinear with the longitudinal axis of the curable composite bushing; A tool for installing the curable composite bushing is provided, the tool including one or more guides and a pressing member; Position the guide adjacent to the curable composite bushing such that the guide encloses the curable composite bushing within an inner surface; Position the pressing member within the guide and press the curable composite bushing at the other end of the pressing member into the gap between the attachment fastener and the attachment hole such that the curable composite bushing generally conforms to the dimensions of the gap, so that there is no gap thereafter; Hold the composite bushing in a compressed state until the composite bushing is cured.

15. The method according to claim 14, further comprising the step of heating the curable composite bushing with a heating device before, during, or after the pressing member moves towards the fastener.

16. The method according to claim 14 or 15, further comprising the steps of providing a second guide and slidably engaging the second guide with the fastener to form a radial channel between the second guide and the pressing member, and feeding the curable composite bushing into the channel.

17. The method according to claim 14 or 15, further comprising the step of moving the pressing member into and out of engagement with the compressed curable composite bushing to ensure that the curable composite bushing is fully compressed into the gap.

18. A tool for installing the curable composite bushing in the aircraft assembly according to claim 3 or 4, the tool including a pressing member configured to press the curable composite bushing into the gap between the attachment fastener and the attachment hole such that the curable composite bushing generally conforms to the dimensions of the gap.

19. The tool according to claim 18, further comprising one or more guides configured to engage an outer surface of the first structural member adjacent to the hole or the fastener, and the guides are further configured to guide the curable composite bushing into the gap.

20. The tool according to claim 19, wherein The guide is further configured to guide the pressing member to a pressing position.

21. The tool according to claim 19 or 20, further comprising a heating device configured to apply heat energy to the curable composite bushing when the curable composite bushing is being guided or pressed.

22. The tool according to claim 21, wherein, The heating device is mounted in the guide or the pressing member.

23. The tool according to claim 21, wherein The heating device is provided by an electric heating element.

24. The tool according to claim 21, wherein, The heating device is provided by a hot gas jet device.

25. The tool according to claim 21, wherein The heating device is provided by an ultrasonic energy emitting device.

26. The tool according to claim 21, wherein The heating device is provided by an electric induction device.

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