Patent Description:
During a process of assembling an aircraft assembly, structural components are brought together and fastened to each other with fasteners. One such process involves holding the components relative to each other in one or more jigs, undertaking a machining operation to drill holes in the components to receive fasteners, and fastening fasteners through the holes to mount the components to each other.

Following formation of the holes by drilling, the components need to be disassembled to allow for a deburring process to be undertaken. This allows for the removal of chaff from around the holes generated by the machining operation. The components are then repositioned and the fasteners inserted through the aligned holes in a fastening operation.

One such aircraft assembly is a wing box assembly which includes upper and lower covers with other components received between the upper and lower covers, such as spars, ribs and landing gear ribs. Removing the covers to enable the deburring operation to take place is an inefficient and time consuming process.

It has been recognised that it is unnecessary to undertake a deburring operation for components formed from some materials used in aerospace applications, for example aluminium and carbon fibre reinforced plastic. However, it is also recognised that the use of such materials in some applications is unsuitable, for example when a high loading capability is required.

<CIT> describes connecting elements for fastening CFRP components and metal components in an aircraft or spacecraft.

According to an aspect of the invention, there is provided an aircraft assembly according to claim <NUM>.

With such an arrangement, a machining process to form a hole in the first component during an assembly process will act on a material having a lower hardness value than the remainder of the component. As such, the wear on the tool, for example a drill bit or grinding tool is minimised. Furthermore, the need to provide a deburring process may be reduced or eliminated. As such, the need to separate the components during the assembly process following a machining operation and prior to a fastening operation is removed.

The likelihood of swarf formed by the machining process acting on adjacent parts is minimized.

The machining process to form a hole may be a drilling process. The machining process to form a hole may be a grinding process.

The material hardness of the second structural component adjacent to the insert may at least substantially correspond to the material hardness of the insert.

With such an arrangement, the ease of forming a hole through both structural components is maximized.

The second structural component may abut the insert.

The body and the insert may form a one piece component. The insert may be mechanically fixed in the body. The insert may be formed from a cured resin. The insert may be cured in the body during manufacture.

The insert may have a maximum material hardness of <NUM> Vickers Hardness (HV), preferably a maximum material hardness of <NUM> HV, and more preferably a maximum material hardness of <NUM> HV. However, it will be understood that the material hardness of the insert is dependent on the material hardness of the body. The material hardness of the insert may also be dependent on the material hardness of one or more adjacent components in the stack of components.

The insert is formed from at least one of aluminum and carbon fibre reinforced plastic.

The body is formed from one of steel and titanium.

The insert may be in an interference fit with the body. The insert may be, for example, welded in the body, bonded in the body, cured in the body, cold worked in the body, or press fit in the body. Accordingly, the load transfer path between the insert and the body is enhanced in a shear load direction.

The insert may comprise a lip, the lip being engaged in the body to retain the insert in an axial direction of the fastener. Accordingly, the load transfer path between the insert and the body is maximized in a pull through load direction.

A portion of the insert may be retained between the body and the second component.

The aircraft assembly may comprise a key configuration between the insert and the body which is configured to prevent rotation of the insert relative to the body about an axis of the fastener.

The insert may have a central axis. The machined hole may be offset from the central axis.

The insert may be one of an array of inserts in the body.

The fastener may be one of a plurality of fasteners. At least one of a plurality of fasteners may extend through each of the inserts.

The fastener may comprise a blind fastener.

The aircraft assembly may be a landing gear assembly.

The insert may be one of an array of inserts, wherein each of the array of inserts corresponds to a component mounting point.

According to another aspect of the present invention, there is provided an aircraft comprising at least one of the aircraft assembly as set out above and the aircraft component as set out above.

According to another aspect of the present invention, there is provided a method of assembling an aircraft assembly according to claim <NUM>.

The method may comprise, following forming the hole in the insert, without moving the first and second aircraft structural components apart, inserting the fastener to fasten the first and second aircraft structural components together.

According to the method, the material hardness of the second aircraft structural component adjacent to the insert may at least substantially correspond to the material hardness of the insert.

<FIG> shows an aircraft <NUM>. The aircraft <NUM> has a fuselage <NUM>, and starboard and port fixed wings <NUM>, <NUM>. An engine <NUM> is mounted to each wing <NUM>, <NUM>. The aircraft <NUM> is a typical jet passenger transport aircraft but the invention is applicable to a wide variety of fixed wing aircraft types, including commercial, military, passenger, cargo, jet, propeller, general aviation, etc. with any number of engines attached to the wings or fuselage. The invention is also applicable to other aircraft, such as helicopters.

Each wing has a cantilevered structure with a length extending in a span-wise direction from a root <NUM> to a tip <NUM>, with the root <NUM> being joined to the aircraft fuselage <NUM>. The wings <NUM>, <NUM> are similar in construction and so only the starboard wing <NUM> will be described in detail. The wing <NUM> has a leading edge <NUM> and a trailing edge <NUM>. The leading edge <NUM> is at the forward end of the wing and the trailing edge <NUM> is at the rearward end of the wing.

The wing <NUM> comprises a wing box <NUM>. The wing box <NUM> forms a structural assembly including forward and rear spars (part of the rear spar shown in <FIG>), ribs extending between the forward and rear spars, upper and lower covers, <NUM>, <NUM>, and other components.

The wing <NUM> has a span-wise axis which extends in a direction from the wing root <NUM> to the wing tip <NUM>, and a chord-wise axis which extends in the direction from the leading edge <NUM> to the trailing edge <NUM>.

The aircraft <NUM> has landing gear assemblies (not shown). A starboard landing gear is selectively extendable from the starboard wing <NUM>, a port landing gear is selectively extendable from the port wing <NUM>, and a nose landing gear is selectively extendable from the fuselage <NUM>. The starboard and port landing gears are mounted on the wing boxes <NUM> of the wings <NUM>, <NUM>.

Referring to <FIG>, a section of the wing box <NUM> is shown. The section of the wing box <NUM> shown includes part of a rear spar <NUM>. A landing gear rib <NUM>, also known as a gear rib, is mounted on the rear spar <NUM>. The gear rib <NUM> acts as part of the mount for the landing gear assembly. The gear rib <NUM> is fixedly mounted to the rear spar <NUM>.

The upper and lower covers, <NUM>, <NUM> are omitted from view in <FIG>. The upper cover is positioned on the upper side of the rear spar <NUM> and the lower cover is positioned on the lower side of the rear spar <NUM>. The gear rib <NUM> extends between the upper and lower covers. The gear rib <NUM> is mounted to the upper cover and the lower cover when the wing box <NUM> is assembled.

The gear rib <NUM> includes a body <NUM>. The body <NUM> includes an array of component mounting points <NUM>. The component mounting points <NUM> enable other components to be fastened with the gear rib <NUM>. The body <NUM> includes an upper cover mounting flange <NUM> and a lower cover mounting flange <NUM>. Component mounting points <NUM> are formed in each of the upper and lower cover mounting flanges <NUM>, <NUM>.

As described herein, the gear rib <NUM> acts as a first component of an aircraft assembly. The present invention is described herein with reference to mounting the gear rib <NUM> with each of the upper and lower covers, each acting as a second component of the aircraft assembly, however it will be understood that each of the first and second components may be different components, and the arrangement of the aircraft assembly may differ.

As will become apparent hereinafter, the gear rib <NUM> is shown part way through an assembly process in which the upper and lower covers <NUM>, <NUM> have already been positioned with respect to the gear rib <NUM> (although the upper and lower covers are omitted from view for clarity in <FIG>) and with fastening bores <NUM> formed through the upper and lower covers and the gear rib <NUM> but prior to fasteners being inserted.

Referring now to <FIG>, a process for assembling an aircraft assembly <NUM>, for example the wing box <NUM>, will now be described. The assembly process will be described with reference to first and second components <NUM>, <NUM>. The first and second components <NUM>, <NUM> are described above as a gear rib and a cover respectively, however it will be understood that the first and second components and the assembly process may relate to alternative components of an aircraft. Furthermore, it will be understood that the assembly process may be applied to more than two components, for example three components having parts in a stacked configuration.

A body <NUM> of the first component <NUM> is shown schematically in <FIG>. The body <NUM> may include a flange. The first component <NUM> is formed as a one piece component. The body <NUM> has a first side <NUM> and a second side <NUM>. Although the first and second sides <NUM>, <NUM> are shown parallel to each other, it will be understood that they may be formed at an incline to each other.

An insert <NUM> is in the body <NUM>. The insert <NUM> is accommodated extending across the body <NUM>. The insert <NUM> forms an interference fit with the body <NUM>. The insert <NUM> may be in the flange. It will be understood the insert may be accommodated in the body <NUM> in different configurations. The fit between the insert <NUM> and the body <NUM> is sufficient to allow for a seamless load transfer between the insert <NUM> and the body <NUM> in a shear load direction. The insert <NUM> and the body <NUM> are pre-assembled. The insert <NUM> is pre-formed with the body <NUM>.

The insert <NUM> is a solid part. That is, the insert <NUM> is formed without one or more holes extending through the insert through which a fastener may be received. The insert <NUM> is a disc in an aperture <NUM> in the body <NUM>. The insert <NUM> is cylindrical, however it will be understood that the insert <NUM> may have alternative configurations. For example, the insert <NUM> may have a non-circular cross-section and may have one or more protrusions and/or recesses formed in the insert <NUM>.

The body <NUM> of the first component <NUM> is formed from a titanium alloy. Titanium alloys typically have a material hardness of at least <NUM> HV, although some alloys, for example dependent on treatment, may have a lower hardness. Alternative materials may be used. For example, the body <NUM> of the first component <NUM> may be formed from steel. The material hardness of the material forming the body <NUM> of the first component <NUM> has a material hardness value of at least <NUM> HV. Such materials typically require deburring following the machining of a hole through the material, for example through use of a drill bit or grinding tool.

The insert <NUM> is formed from a different material to the body <NUM>. The insert <NUM> is formed from aluminium. The insert <NUM> may be formed from an alternative material such as carbon fibre reinforced plastic (CFRP). The material forming the insert <NUM> is a softer material than the material forming the body <NUM>. That is, the material hardness of the insert <NUM> is lower than the material hardness of the body <NUM>. The material forming the insert <NUM> has a material hardness of less than <NUM> HV. However, it will be understood that this is dependent on the relative material hardness of the body <NUM>. That is, the material hardness of the insert <NUM> is less than the material hardness of the body <NUM>. The insert <NUM> has sufficient outer dimensions to accommodate a hole for receiving a fastener therethrough. The size of the hole required to be formed through the insert should be sufficient to accommodate the required fastener for fastening the components <NUM>, <NUM> at the component mounting point <NUM>. The insert <NUM> is at a predetermined one of the component mounting points <NUM>. The insert <NUM> is configured to be sized to accommodate any tolerance build up at the component mounting point <NUM> as predetermined for the assembly of the aircraft assembly <NUM>.

Hardness is described herein by reference to Vickers hardness (HV) as a measure of material hardness, although it will be understood that other methods are used to determine material hardness. Examples of Vickers hardness values are provided below:.

Referring to <FIG>, the first component <NUM> is aligned with the second component <NUM>. The second component <NUM> is moved into abutment with the first side <NUM> of the body <NUM>. The first and second components <NUM>, <NUM> are aligned to be fastener together in a predetermined stacked arrangement.

The second component <NUM> includes a body <NUM>. The body <NUM> may form the whole or part of the second component <NUM>. The body <NUM> may include a flange. The second component <NUM> is formed from carbon fibre reinforced plastic. It will be understood that the second component <NUM> may be formed from an alternative material such as aluminium, titanium, or steel.

In the present configuration, the second component <NUM> is shown with a pre-formed hole <NUM>. The pre-formed hole <NUM> extends through the body <NUM>. The hole <NUM> may be preformed prior to bringing the first and second components <NUM>, <NUM> together. The hole <NUM> may be formed during the assembly process. It will be recognised that in an embodiment in which the second component is formed from a material having a material hardness substantially corresponding to that of the insert then any hole formed during the assembly process can be formed without a requirement for a subsequent deburring operation.

The hole <NUM> is aligned with the insert <NUM>. That is, the hole <NUM> fully overlaps the insert <NUM>. The hole <NUM> does not overlap the body <NUM>. In an arrangement in which the hole <NUM> is formed during the assembly process, then the position of the hole is predefined as a component mounting point <NUM>. The insert <NUM> is comparatively sized with the preformed hole <NUM> to accommodate any pre-determined tolerance build ups during assembly of the components <NUM>, <NUM>.

The preformed hole <NUM> has a second component hole axis <NUM>. It will be noted that the second component hole axis <NUM> is offset from a central axis <NUM> of the insert <NUM>. In the event of no misalignment or tolerance build-up, then the second component hole axis <NUM> and central axis <NUM> of the insert <NUM> may be coaxial.

Upon alignment of the first and second components <NUM>, <NUM> in an arrangement for assembly, a machine operation is performed. The machine operation bores a hole. A drill bit <NUM> is used to bore a through hole <NUM> in the insert <NUM>. The drill bit <NUM> is a boring tool. A grinding tool may be used to bore the through hole <NUM>. The drill bit <NUM> is aligned at the component mounting point <NUM>. In an embodiment in which the hole <NUM> in the second component <NUM> is preformed, then the drill bit <NUM> may be aligned with the axis <NUM> of the preformed hole <NUM>. Alternatively, the component mounting point <NUM> is determined and the drill bit <NUM> is used to form the hole through both the first and second components <NUM>, <NUM>. In <FIG>, the drill bit <NUM> is shown during the machining operation partially engaged with the insert <NUM>. The drill bit <NUM> is acting in a direction through the second component <NUM> and into the first component <NUM>. In embodiments, the opposite direction may be used.

The machining operation forming the machined hole ensures alignment of the holes <NUM>, <NUM> through both the first and second components <NUM>, <NUM>. The holes <NUM>, <NUM> form a fastening bore <NUM>. The axis <NUM> of the hole <NUM> in the second component is therefore coaxial with the axis of the through hole <NUM> in the first component <NUM>. The through hole <NUM> is formed fully through the insert <NUM>. The insert <NUM> forms a collar around the through hole <NUM>.

Once the machining operation is complete, a fastening operation is performed. A fastener <NUM> is inserted through the fastening bore <NUM>. The fastener <NUM> is fastened in an engaged position to mount the first and second components <NUM>, <NUM> with each other. It will be recognised that following the machining operation there is no need to deburr either of the first or second components <NUM>, <NUM>, in particular as the machining process acts on a softer material. The material hardness of the insert is less than the corresponding material hardness of the material surrounding the insert.

It will be understood that other material properties may contribute to aid the machining operation. For example, in embodiments at least one of the material toughness, the material abrasiveness and the material ductility of the insert is less than the corresponding material toughness, material abrasiveness and material ductility of the body.

The fastener <NUM> is shown as a bolt <NUM> and a nut <NUM> arrangement. However, it will be appreciated that the fastener <NUM> may be a blind fastener. That is a fastener that is inserted through the fastening bore <NUM> and engaged with both of the first and second components <NUM>, <NUM> from one side of the assembly only. An advantage of this arrangement is that the machining operation and the fastening operation may be performed from the second component side of the assembly <NUM> only.

The interference fit between the insert <NUM> and the body <NUM> provides for shear loads to be sufficiently transferred between the first component <NUM> and the fastener <NUM> to the second component <NUM>. In <FIG>, the first component side of the fastener is shown in contact with the insert <NUM> only, however it will be understood that the end <NUM> of the fastener <NUM> may be configured to extend over at least part of the body <NUM>. Such a configuration would aid the transfer of a pull through load on the first component <NUM>.

Another embodiment is shown in <FIG>. The embodiment in <FIG> is generally the same as described above and the assembly process is generally the same and so a detailed description will be omitted herein. However, in this embodiment the configuration of the insert differs. <FIG> shows a partially assembled aircraft assembly <NUM> with first and second components <NUM>, <NUM>. The partially formed aircraft assembly <NUM> is shown following the machining operation and prior to the fastening operation. As such, a through hole <NUM> is formed through an insert <NUM>. The insert <NUM> is received in the body <NUM> of the first component <NUM>. The insert <NUM> is generally the same as the insert <NUM> described above, however in this embodiment the insert <NUM> includes a lip <NUM>. The lip <NUM> is a circumferentially extending flange. The lip <NUM> may have a different configuration, and extend only partially around the insert <NUM>. The lip <NUM> protrudes outwardly. The lip <NUM> is a protrusion. The lip <NUM> is received on a shoulder <NUM> of the body <NUM>. The lip <NUM> is received between the shoulder <NUM> and the first side <NUM> of the body <NUM>. The lip <NUM> aids retention of the insert <NUM> in the body <NUM>. When assembled, the lip <NUM> is received between the shoulder <NUM> of the body <NUM> of the first component <NUM> and the second component <NUM>. As such, the insert <NUM> is able to handle greater pull through loads acting on the aircraft assembly <NUM>. In embodiments the lip <NUM> is a countersink.

Referring to <FIG>, another embodiment is shown. The arrangement of this embodiment is generally the same as the embodiments shown above. <FIG> shows the first component <NUM> prior to assembly with the second component <NUM>, and prior to the machining operation. As such, no hole is formed through the insert. An insert <NUM> is shown in the body <NUM> of the first component <NUM>. The insert <NUM> has a key configuration <NUM>. The key configuration includes a key <NUM> and a key slot <NUM>. The key <NUM> protrudes from a main part of the insert. The key <NUM> protrudes radially outwardly in the present embodiment. The key <NUM> is received in a corresponding key slot <NUM> in the body <NUM>. The key <NUM> may have differing configurations and may comprise two or more key features. The key configuration <NUM> aids prevention of any relative rotation of the insert <NUM> and the body <NUM>, for example, such as may be applied during the machining process.

In each of the embodiments described above, it will be appreciated that the insert and the body together form the first component <NUM> as a one piece component. The first component <NUM> includes a plurality of inserts preassembled with the body <NUM>. The location of the insert <NUM> corresponds to the position of predetermined component mounting points <NUM>. The inserts are preformed without any through holes formed therein through which fasteners may be engaged, and therefore the fastener receiving holes are formed during the assembly process. It has been recognised that by using a relatively softer material than that of the body of the component, that it is possibly to remove the need for subsequent machining processes following the forming of the hole in the insert and therefore reducing the assembly time. It will be recognised that in some embodiments two or more through holes arranged to receive fasteners may be formed in a single insert.

In the embodiment shown in <FIG> in which the first component is a landing gear rib <NUM> and the second component is one of the covers <NUM>, <NUM>, it will be appreciated that a component that is required to carry a significant load transfer may lead to the cover having to be removed in order to deburr holes machined in the component. However, with the arrangements described above it has been recognised that inserts may be used to allow the holes to be formed in a relatively softer material to remove the further machining requirement and so remove the need to remove the cover. As such, the assembly time and complexity of the assembly process may be reduced. Furthermore, as the tools, for example the drill bits used during the assembly process are required to act on a softer material hardness only, then the wear on these tools is minimised.

Claim 1:
An aircraft assembly comprising:
a first structural component (<NUM>);
a second structural component (<NUM>);
a fastener (<NUM>) fastening the first structural component (<NUM>) to the second structural component (<NUM>);
wherein the first structural component (<NUM>) comprises a body (<NUM>) and an insert (<NUM>) in the body, the insert (<NUM>) having a machined hole through which the fastener (<NUM>) extends; and
wherein the material hardness of the insert (<NUM>) is less than the material hardness of the body (<NUM>), the body (<NUM>) formed from one of steel and titanium, and characterised in that the insert (<NUM>) is formed from one of aluminium and carbon fibre reinforced plastic.