Patent Description:
Aircraft typically comprise an airframe to which an aircraft external skin is fastened. Many aircraft airframes comprise a plurality of frames (or formers) and longerons and/or shear webs. The frames are typically laterally spaced from one another and arranged perpendicular to the longitudinal axis of the aircraft. The primary purpose of frames is to establish the shape of the fuselage. The longerons and/or shear webs are typically structural members which are attached between pairs of frames and are arranged parallel to the longitudinal axis of the aircraft. The longerons and/or shear webs support the aircraft skin and, in use, transfer aerodynamic loads acting on the skin onto the frames.

It is desirable that aircraft airframes are produced to be within very tight tolerance bounds.

Packers may be applied to the airframe and subsequently machined to provide an airframe having a desired shape.

Patent application <CIT> is related to rotary wing aircraft with a fuselage that comprises at least one structural stiffened panel, the structural stiffened panel comprising a stressed skin and a stiffening framework that is rigidly attached to the stressed skin. The stressed skin comprises an inner and outer skin, and a core element assembly that is arranged between the inner skin and outer skin. The core element assembly comprises at least one viscoelastic core element and at least one intermediate core element that are tessellated, wherein the at least one viscoelastic core element is provided for noise and vibration damping.

Patent application <CIT> relates to a skin-stiffened composite panel comprising at least two essentially longitudinal stiffeners arranged distant to each other in a plane with a skin of a planar outer continuous composite layer and an inner continuous composite layer. The at least two essentially longitudinal stiffeners have a height "a" perpendicular to the skin. At least two sandwich assemblies in said plane are along each one of the at least two essentially longitudinal stiffeners, said at least two sandwich assemblies comprising each the planar outer and the inner continuous composite layers and a grid-type foam module with openings and with a height "b", with the foam modules on the same side of the planar outer continuous composite layer as the essentially longitudinal stiffeners between the inner continuous composite layer and the planar outer continuous layer.

Patent application <CIT> discloses a structural element, in particular for an aircraft and spacecraft, comprising a core, the rigidity of which varies at least in portions for optimising the aeroelastic characteristics of the structural element. It also discusses a method for producing a structural element, in particular for an aircraft and spacecraft, which comprises the following steps: provision of a structural element comprising a core; determination of the aeroelastic behaviour of the structural element; and variation, at least in portions, of the rigidity of the core of the structural element such that the aeroelastic behaviour of the structural element is optimised.

Patent application <CIT> discloses an airframe assembly for an aircraft. The airframe assembly includes a first airframe member having a first skin, a second skin, a large cell core joined between the first and second skins and a solid insert having a side surface. The solid insert is joined between the first and second skins such that at least a portion of the side surface is adjacent to the large cell core. The first skin has a first surface disposed opposite the solid insert. The airframe assembly also includes a second airframe member having a second surface. An adhesive joint is disposed between the first and second surfaces structurally bonding the first airframe member to the second airframe member such that the second airframe member is positioned opposite the solid insert.

Patent application <CIT> discloses an aircraft fuselage section having a curved shape with at least a vertical symmetry plane (A-A) and a central longitudinal axis and comprising a skin and a plurality of frames arranged perpendicularly to said longitudinal axis. The aircraft fuselage section also comprises at least an inner reticular structure mounted on a supporting structure comprising longitudinal beams attached to the skin and interconnected with said frames, said inner reticular structure being arranged for creating at least one closed cell with the skin for improving its resistance and its damage tolerance to said impacts. Said inner reticular structure can be formed by panels, rods, cables or belts.

It will be appreciated that relative terms such as horizontal and vertical, top and bottom, above and below, front and back, upper and lower, and so on, are used herein merely for ease of reference to the Figures, and these terms are not limiting as such, and any two differing directions or positions and so on may be implemented rather than truly horizontal and vertical, top and bottom, and so on.

<FIG> is a schematic illustration (not to scale) of a portion of an airframe <NUM> of an aircraft, this portion hereinafter being referred to as the "airframe <NUM>". An aircraft external skin is to be fastened to the airframe <NUM>.

The airframe <NUM> comprises three laterally spaced-apart frames or "formers" <NUM> and three longerons <NUM>. It will be appreciated by those skilled in the art that the airframe <NUM> may comprise a different number of frames <NUM> and/or longerons <NUM>.

The frames <NUM> are made of aluminium or titanium. The frames <NUM> define the shape of the aircraft fuselage and, in use, provide stability to the aircraft by preventing or opposing deflection of the longerons <NUM>. When the aircraft is fully assembled, the frames <NUM> are arranged substantially perpendicularly to the longitudinal axis of the aircraft.

The longerons <NUM> are made of aluminium or titanium. The longerons <NUM> are elongate members to which the skin of the aircraft is fastened. When the aircraft is fully assembled, the longerons <NUM> run substantially parallel to the longitudinal axis of the aircraft. In this embodiment, the longerons <NUM> are fastened to the frames by a plurality of aircraft fasteners, e.g. bolts.

The aircraft <NUM> may further comprise additional structural elements not shown in <FIG> for reasons or clarity and conciseness. Such example structural elements may attach together frames <NUM> and/or longerons <NUM>. Examples of such structural elements may include, but are not limited to, shear webs, beams, keels, and fuel floors.

To produce the aircraft fuselage, a composite aircraft skin is fastened to the airframe <NUM>. The outer shape of the assembled fuselage (i.e. the outer shape of the fuselage produced by fastening the composite skin to the airframe <NUM>) is referred to as the Outer Mould Line (OML) of the fuselage. In this embodiment, the OML of the fuselage is to be within a pre-specified tolerance. The OML of the fuselage having the required tolerance is facilitated by the Inner Mould Line (IML) of the fuselage being within a pre-specified tolerance. The IML of the fuselage is the surface at which the airframe <NUM> and the aircraft skin abut, i.e. an outer surface of the airframe <NUM> and inner surface of the aircraft skin.

<FIG> is a schematic illustration (not to scale) showing an embodiment of a packer <NUM>.

In this embodiment, the packer <NUM> is for application to the outer surface the airframe <NUM> prior to the aircraft skin being fastened to the airframe <NUM>. Thus, in the assembled aircraft, the packer <NUM> is sandwiched between the airframe <NUM> and the aircraft external skin.

The packer <NUM> may be considered to be a sacrificial layer of material that, after being attached to the airframe <NUM>, may be machined to provide a desired IML for the airframe <NUM>. In this way, attachment of the aircraft skin to the airframe <NUM> may be improved.

In this embodiment, the packer <NUM> is made of or comprises a carbon fibre composite (CFC) material.

The packer <NUM> is elongate. The packer <NUM> may have any appropriate length L. The length L of the packer <NUM> may be application dependent. The packer <NUM> may have any appropriate width W. The width W of the packer <NUM> may be application dependent. The packer <NUM> may have any appropriate thickness T, for example between about <NUM> and about <NUM>, or about <NUM>. The thickness T of the packer <NUM> may be application dependent. Preferably, the packer <NUM> has substantially uniform thickness T prior to being attached to the airframe <NUM>. The thickness T of the packer <NUM> is a distance between an upper surface and a lower surface of the packer <NUM>.

In this embodiment, the packer <NUM> comprises a frame <NUM>, a plurality of pads <NUM>, and a plurality of struts <NUM>. The frame <NUM>, the pads <NUM>, and the struts <NUM> are integrally formed. Preferably, the frame <NUM>, the pads <NUM>, and the struts <NUM> are formed from a single piece of material, i.e. as a single, integral entity.

The frame <NUM> is a substantially rectangular frame. The frame <NUM> comprises four elongate, substantially straight members that are arranged to define a rectangle. More specifically, the frame <NUM> comprises two elongate and opposing side members (namely a first side member 202a and a second side member 202b) that define the sides of the packer <NUM>, and two elongate and opposing end members 202c, 202d disposed between the side members 202a, 202b. The end members 202c, 202d define the ends of the packer <NUM>.

A first plurality of pads 204a (each of which is hereinafter referred to as a "first pad 204a") is arranged along an inside edge of the first side member 202a. A second plurality of pads 204b (each of which is hereinafter referred to as a "second pad 204b") is arranged along an inside edge of the second side member 202b. Each first pad 204a is positioned opposite to a respective one of the second pads 204b. Thus, in this embodiment, there is an equal number of first and second pads 204a, 204b.

The pads 204a, 204b may be approximately circular in shape (e.g. when the packer <NUM> is viewed from above or below). Thus, each pad 204a, 204b may define approximately a disc. The pads 204a, 204b may have any appropriate diameters. The diameters of the pads 204a, 204b may be application dependent. By way of example, the diameters of the pads 204a, 204b may be at least double the diameters of the fasteners that are to be placed through those ds 204a, 204b.

The struts <NUM> are elongate members. The struts may have any appropriate widths, for example between about <NUM> and about <NUM>, or between about <NUM> and <NUM>, e.g. about <NUM>. The widths of the struts <NUM> may be application dependent.

In this embodiment, the struts <NUM> arranged across the width of the packer <NUM>. Each strut <NUM> is coupled between a first pad 204a and a second pad 204b. More specifically, each strut <NUM> is coupled between a first pad 204a and a second pad 204b that is adjacent to the second pad 204b that is opposite to that first pad 204a. Also, each strut <NUM> is coupled between a second pad 204b and a first pad 204a that is adjacent to the first pad 204a that is opposite to that second pad 204b. In this way, two struts <NUM> extend from each pad 204a,b (except the pads at the ends of the packer, from which only a single struct extends). The struts <NUM> extend across the width W of the packer <NUM> oblique to the members 202a-d of the frame <NUM>. The angles of struts to the members 202a-b may be varied by changing the spacing of the pads 204a,b.

The struts <NUM> that couple together adjacent pairs of pads 204a,b cross each other substantially along a longitudinal axis of the packer <NUM> (i.e. i.e. along a centreline of the packer <NUM> running along its length L). Struts <NUM> are integrally formed where they cross. Preferably, the struts <NUM> have uniform thickness along their lengths.

In this embodiment, the frame <NUM>, the pads <NUM>, and the struts <NUM> define a plurality of voids <NUM>. The voids <NUM> may be considered to be through holes through the packer <NUM> from the upper surface of the packer <NUM> to the lower surface of the packer <NUM>. In this embodiment, the widths of the voids <NUM> are greater than the widths of the struts <NUM>.

Advantageously, the voids <NUM> tend to allow the packer <NUM> to bend or flex to any desired shape. The packer <NUM> tends to be more flexible than conventional CFC packers.

<FIG> is a process flow chart showing certain steps of a process of producing the packer <NUM> and using the packer to secure an aircraft external skin to the airframe <NUM>.

At step s2, a substantially flat sheet of CFC is provided. Preferably, the sheet has substantially uniform thickness. Preferably, the thickness of the sheet of CFC material is equal to the desired thickness of the packer <NUM>.

At step s4, the packer <NUM> is cut from the provided sheet. In this embodiment, computer-controlled cutting is implemented to cut the packer shape from the provided sheet thereby to provide the packer <NUM>. By way of example, laser cutting or water-jet cutting may be implemented to cut the packer <NUM> from the sheet.

Advantageously, production of the packer <NUM> by cutting the packer <NUM> from a CFC sheet is low-cost, fast, and simple compared to production of conventional packers. Typically, complex geometry tools (e.g. a lay-up tool and a profile tool) are used to produce conventional CFC packers. Use of such tools tends to be avoided in the production of the packer <NUM>.

At step s5, a plurality of holes is drilled through the packer <NUM>. Each hole is drilled through a respective pad 202a,b of the packer <NUM>. These drilled holes are for receiving aircraft fasteners for fastening the aircraft external skin to the airframe <NUM>. Thus, in this embodiment, the pads 202a,b, act as landings or seats for receiving aircraft fasteners.

At step s6, an adhesive is applied to the external surface of the airframe <NUM> where the packer <NUM> is to be fixed. Any appropriate type of adhesive may be used. The adhesive may be application dependent. The adhesive may be a film adhesive or a paste adhesive. The adhesive may be applied in any appropriate way, for example by spraying or spreading liquid adhesive onto the airframe.

At step s8, the packer <NUM> is pressed onto the portion of the airframe <NUM> to which the adhesive was applied at step s6. The packer <NUM> is pressed onto the airframe <NUM> such that the packer <NUM> is in intimate contact with the airframe <NUM>. The packer <NUM> may be pressed or applied onto the airframe <NUM> is any appropriate way, for example using a roller or by-hand.

Advantageously, the voids <NUM> of the packer <NUM> tend to allow for the packer <NUM> to be pressed through the adhesive, e.g. such that the adhesive is forced by the packer material into the voids <NUM>. This tends to facilitate accurate positioning of the packer on to the airframe <NUM> and provide that the packer <NUM> is held in position by the adhesive without a need for clamping.

Advantageously, use of the flexible packer <NUM> tends to reduce or eliminate a need for clamping devices (e.g. G-clamps) to be used to hold the packer <NUM> onto the airframe <NUM> compared to conventional, stiffer packers. The flexible packer <NUM> tends to bend such that it conforms to the shape of the airframe <NUM> to which it is applied, and may be fixed to the airframe <NUM> without need for clamps or with a reduced clamping requirement. Furthermore, it tends to be possible to apply the flexible packer <NUM> to parts having different geometries. The flexible packer <NUM> tends to account for manufacturing tolerances in the parts to which it is applied.

The widths of the frame <NUM>, struts <NUM>, and pads <NUM> are relatively small, e.g. compared to conventional packers. Thus, advantageously, air entrapment under the packer <NUM> when the packer <NUM> is adhered to the airframe <NUM> tends to be reduced or eliminated. Furthermore, it tends to be easier to remove any trapped air between the packer <NUM> and airframe <NUM> prior to curing of the adhesive. Thus, the packer <NUM> tends to provide for an improved bond between the packer <NUM> and the airframe <NUM>, and an improved peel strength. Furthermore, the reduction in entrapped air tends to reduce the likelihood of disbonds, moisture ingress, and corrosion.

At step s10, excess adhesive is removed from the airframe <NUM> and packer <NUM>. This is performed prior to the curing of the adhesive. Removal of the excess adhesive may be performed in any appropriate way, for example by wiping or scraping using an appropriate tool. For example, a scraper may be scraped over the upper surface of the packer <NUM> (i.e. flush with the upper surface) thereby to remove any excess adhesive that extends above the upper surface of the packer <NUM>.

Conventionally, removal of the uncured excess adhesive when using conventional packer tends to be difficult, for example since the clamping devices (e.g. G-clamps) that are conventionally used to press the conventional packers onto the airframe may obstruct the process. As such, conventionally, excess adhesive is removed after curing, which tends to be more difficult and risks damage to the airframe/packer. Use of the above described flexible packer <NUM> advantageously tends to avoid or reduce the use of such clamping device, thereby facilitating the removal of unwanted adhesive before it is cured.

At step s12, the adhesive is cured. Curing of the adhesive may be done in any appropriate way. Curing the adhesive may comprise leaving the adhesive to cure over time, or applying electromagnetic radiation (e.g. heat and/or light) to the adhesive.

Thus, the packer <NUM> is fixed to the airframe <NUM>.

<FIG> is a schematic illustration showing the aircraft airframe <NUM> having multiple packers <NUM> attached thereto. <FIG> shows packers <NUM> attached to a longeron <NUM> of the airframe <NUM>. However, it will be appreciated by those skilled in the art that one or more packers <NUM> may be attached to any appropriate surface of the airframe <NUM>.

<FIG> further shows a plurality of holes <NUM> that were drilled through the pads 202a,b of the packer at step s5 above. In this embodiment, these holes <NUM> through the packer <NUM> align with respective through holes through the longeron <NUM>. These aligned pairs of holes are for receiving aircraft fasteners, as described in more detail later below with reference to <FIG>.

At step s14, the external surface of the packer <NUM> (i.e. the surface opposite to the surface adhered to the airframe <NUM>) is machined to define a desired IML for the airframe <NUM>. This machining of the packer <NUM> may comprise measuring an external surface of the assembly (i.e. the airframe <NUM> with the packer <NUM> applied thereto), and, using the measurements, controlling a machining apparatus (e.g. a computer numerical control (CNC) machining apparatus) to machine the packer <NUM> to define a desired IML.

Thus, an airframe assembly defining an IML within a specified tolerance tends to be provided.

At step s16, the aircraft external skin is positioned against the outer surface of the packer <NUM>. The aircraft external skin is positioned such that through holes through the aircraft external skin are aligned with the holes <NUM> through the packer <NUM> and longeron <NUM>.

Advantageously, the IML being within the specified tolerance tends to facilitate the locating of the aircraft skin against the airframe assembly.

At step s18, the aircraft external skin is fastened to the airframe <NUM> using a plurality of aircraft fasteners.

<FIG> is a schematic illustration (not to scale) showing a cross section through a portion of the structure <NUM> formed by fastening the aircraft external skin <NUM> to the airframe <NUM>. <FIG> illustrates attachment of the aircraft external skin <NUM> to the airframe <NUM> by a fastener <NUM>.

In this embodiment, the fastener <NUM> comprises a threaded bolt <NUM> and a nut <NUM>.

The bolt <NUM> passes, in turn, through the aircraft external skin <NUM>, a pad 202a/b of the packer <NUM>, a longeron <NUM> of the airframe <NUM>, and the nut <NUM> which secures the bolt <NUM>.

Thus, the pad 202a/b is a landing for the fastener <NUM>. Advantageously, in use, the pads 202a/b of the packers <NUM> bear the loads from the fasteners <NUM> passing therethrough. In this embodiment, the packer <NUM> is a load-bearing structure. The material of the packer <NUM> (i.e. one or more of the frame <NUM>, one or more pads <NUM>, and one or more struts <NUM>) is located at load-critical positions on the airframe assembly (preferably at all load-critical locations). An example load-critical position may be a position at which a fastener is located. Also, in this embodiment, the voids <NUM> are not located at load-critical positions on the airframe assembly.

In this embodiment, the hole in the aircraft external skin <NUM> through which the bolt <NUM> passes is countersunk thereby to provide that the head of the bolt <NUM> is substantially flush with the external surface 502a of the aircraft external skin <NUM>.

Advantageously, the IML being within the specified tolerance tends to provide that the OML, which is defined by the external surface 502a of the aircraft external skin <NUM>, is within a pre-specified tolerance.

Thus, the process of producing the packer <NUM> and using the packer <NUM> to secure an aircraft external skin <NUM> to the airframe <NUM> is provided.

It should be noted that certain of the process steps depicted in the flowchart of <FIG> and described above may be omitted or such process steps may be performed in differing order to that presented above and shown in that Figure. Furthermore, although all the process steps have, for convenience and ease of understanding, been depicted as discrete temporally-sequential steps, nevertheless some of the process steps may in fact be performed simultaneously or at least overlapping to some extent temporally.

In the above embodiments, an airframe of an aircraft fuselage is produced. However, in other embodiments, a different type of structure is produced. For example, an airframe of a different part of the aircraft, e.g. a wing or the empennage, may be produced.

In the above embodiments, the packer is made of or comprises CFC. However, in other embodiments, the packer is made of one or more different material instead of or in additional to CFC. Examples of appropriate materials include but are not limited to glass-fibre reinforced composite materials, e.g. glass-reinforced plastic (GRP), aluminium and alloys thereof, plastics, steel, titanium and alloys thereof.

In the above embodiments, the packer is elongate and rectangular. However, in other embodiments, the packer has a different appropriate shape. The shape of the packer may be application dependent.

In the above embodiments, the shape of the packer and the shapes of the voids therethrough are defined by the frame, the plurality of pads, and the plurality of struts. In other embodiments, the packer has a different shape, for example a frame, struts and/or pads may be arranged differently, such that the shape of the voids are different. The voids may have any appropriate shape including, but not limited to, triangular, square, rhombus, pentagonal, hexagonal, heptagonal, octagonal, etc. For example, the packer may comprise voids defining a honeycomb pattern.

In the above embodiments, the pads are substantially disc-shaped structures. However, in other embodiments, one or more of the pads is a different shape, e.g. polygonal such as square, triangular, hexagonal, etc..

In the above embodiments, the packer is produced by cutting the packer from a CFC sheet. However, in other embodiments, the packer is formed in a different way, for example using additive manufacturing (AM) such as fused deposition modelling (FDM).

In the above embodiments, the pads of the packer provide landings for the fasteners that fasten the aircraft skin to the airframe. The pads are load-bearing structures that bear the loads from the fasteners. The packer is a load-bearing structure, with the packer material located at load-critical positions on the airframe assembly and the voids located at (e.g. only) non-load-critical positions.

However, in other embodiments the packer is non-load-bearing. In such embodiments, packer material may be located at (e.g. only) the non-load-critical positions on the airframe assembly, while the voids are located at (e.g. all of) the load-critical positions. In such embodiments, after the packer <NUM> is applied to the airframe <NUM> (e.g. at step s8 or s10 in the method of <FIG> described in more detail earlier above), the cavities or recesses that are defined by the voids <NUM> may be filled with a load-bearing adhesive (for example Hysol, such as Loctite EA9394S or Loctite EA9395). The packer <NUM> acts as a "dam" that retains the liquid/uncured load-bearing adhesive in desired positions against the airframe <NUM>. The aircraft skin <NUM> may than be disposed against the airframe <NUM> (with the packer <NUM> and load-bearing adhesive thereon) and the adhesive cured. Fasteners <NUM> may be used to fasten the aircraft skin <NUM> to the airframe <NUM>. The fasteners <NUM> pass through the cured load-bearing adhesive, and not through the packer material.

Claim 1:
A method of attaching an aircraft external skin to an aircraft airframe, the method comprising:
providing a packer, wherein the packer is for application to the aircraft airframe, and wherein the packer comprises:
a sheet of material having a first surface and a second surface opposite to the first surface; and
a plurality of voids defined through the sheet of material from the first surface to the second surface, thereby to provide that the sheet of material is flexible; and wherein the method further comprises:
attaching the packer to an external surface of the airframe;
after attaching the packer to the external surface of the airframe, filling one or more of the voids with an adhesive; thereafter
machining the packer to provide a desired Inner Mould Line (IML) for the airframe; and
attaching the aircraft external skin to an external surface of the airframe and packer assembly using a plurality of fasteners, each fastener passing through the aircraft external skin, a respective region of cured adhesive located in a void, and at least part of the airframe.