GB2583940A

GB2583940A - Aircraft panel assembly

Info

Publication number: GB2583940A
Application number: GB1906794.1A
Authority: GB
Inventors: Heaysman Chris; Gaitonde Martin; Seegel Hauke
Original assignee: Airbus Operations GmbH; Airbus Operations Ltd
Current assignee: Airbus Operations GmbH; Airbus Operations Ltd
Priority date: 2019-05-14
Filing date: 2019-05-14
Publication date: 2020-11-18
Also published as: GB201906794D0

Abstract

An aircraft panel assembly has a panel 22, stiffeners 8, and a rib foot beam 18. Each stiffener has a part (fig.6,25) attached to the panel, and a structural part 26, spaced apart from the panel. The rib foot beam crosses the stiffeners at a series of intersections 23, where the beam passes between the panel and the structural part of the stiffener. Preferably, each stiffener has a pair of feet, with webs (fig.6,28) connecting the feet to a crown; the feet bonded to an inner panel surface, and the beam bonded to the panel and to the stiffeners at the intersections. A layer (fig.9,29) may form the crown of the stiffeners and cover the beam between the stiffeners. An aircraft wing assembly may use the panel as wing covers; with a rib connecting upper and lower covers, and attached to the beam between the intersections. A method of manufacturing an aircraft panel assembly is also claimed, having the steps of providing fibre preforms on a forming tool, which may involve spraying cut fibres onto the forming tool; laying the beam in position; co-infusing the preforms with a matrix material; and curing the matrix material.

Description

AIRCRAFT PANEL ASSEMBLY FIELD OF THE INVENTION

[0001] The present invention relates to an aircraft panel assembly, and a method of manufacturing an aircraft panel assembly.

BACKGROUND OF THE INVENTION

[0002] EP0995673 discloses in Figure 12 a shell construction for an aircraft with stringer webs and ribs which form a grid.

[0003] W02015/015152 discloses in Figure 14 an aircraft structure with rib feet which span the gap between adjacent stringers.

SUMMARY OF THE INVENTION

[0004] According to a first aspect of the invention, there is provided an aircraft panel assembly comprising: a panel; and a plurality of stiffeners on the panel, wherein each stiffener comprises an attachment part attached to the panel and a structural part spaced apart from the panel; and a rib foot beam which crosses the stiffeners at a series of intersections, wherein at each intersection the rib foot beam is located between the panel and the structural part of a respective one of the stiffeners.

[0005] The aircraft panel assembly has a number of advantages compared with the structure of W02015/015152. Firstly, the use of a single rib foot beam rather than a plurality of rib feet makes the assembly less complex to manufacture. Secondly, the use of a single rib foot beam means that it is not necessary to align multiple rib feet with each other. Thirdly, positioning the rib foot beam between the panel and the structural parts of the stiffeners provides improved stress performance because the rib foot beam is held down on the panel by the structural parts of the stiffeners.

[0006] The attachment part of each stiffener may be bonded (for instance by a co-infused joint) to the panel, or attached to the panel in some other way (for instance by fasteners). The attachment part may be in direct contact with the panel, or it may be attached to the panel without being in direct contact with the panel -for instance there may be one or more layers of additional material between the attachment part and the panel.

[0007] The rib foot beam may be bonded (for instance by a co-bonded joint) to the panel, or attached to the panel in some other way (for instance by fasteners). The rib foot beam may be in direct contact with the panel, or it may be attached to the panel without being in direct contact with the panel [0008] The rib foot beam may be bonded (for instance by a co-bonded joint) to the stiffeners at the intersections. This enhances the strength of the structure.

[0009] Preferably the panel, the rib foot beam and/or the structural parts of the stiffeners comprise fibre-reinforced composite material, such as carbon fibre-reinforced epoxy resin. Alternatively the panel, the rib foot beam and/or the structural parts of the stiffeners may be metal [0010] At each intersection the structural part of the respective one of the stiffeners may extend continuously across the intersection. This continuous structure is advantageous because it enhances the bending stiffness of the stiffener and the strength of the assembly.

[0011] The assembly may comprise a layer (typically a layer of fibre-reinforced composite material) which forms the structural parts all of the stiffeners. Optionally the layer forms the structural parts all of the stiffeners and also covers the rib foot beam between the stiffeners.

[0012] At each intersection the structural part of the respective one of the stiffeners may comprise reinforcement fibres which extend continuously across the intersection.

[0013] At each intersection the stiffener may have an opening (such a slot, hole or gap) through which the rib foot beam passes [0014] The aircraft panel assembly may comprise an aircraft cover panel (such as a wing or fuselage cover panel); a wall of an aircraft fuel tank; or any other type of aircraft panel assembly. In the case of a wing cover panel, the stiffeners are commonly referred to as stringers; whereas in the case of a fuselage panel they may be referred to as longerons [0015] Optionally the stiffeners are carried by an inner surface of the panel, and the panel has an outer surface with a curved aerodynamic profile.

[0016] Optionally the attachment part of each stiffener comprises a pair of feet and a pair of webs; and the structural part of each stiffener comprises a crown which joins the pair of webs to each other. For example each stiffener may have an omega-shaped or hat-shaped profile Alternatively each stiffener may have a different profile such as T-shaped profile, an I-shaped profile, a top heavy I-shaped profile, a baseball bat profile, a lollipop profile or a blade-shaped profile.

[0017] Optionally each stiffener comprises a core, such as a foam material, between the webs.

[0018] Optionally the attachment part of each stiffener comprises a pair of feet and a pair of webs; and the structural part of each stiffener comprises a crown which joins the pair of webs to each other and a reinforcement element between the webs.

[0019] The rib foot beam may have a height perpendicular to the panel which is smaller at the intersections than between the intersections. This reduced height at the intersections reduces the size of the slot or other opening which may be formed in the stiffener to enable the rib foot beam to pass through.

[0020] The structural part of each stiffener may be at an apex of the stiffener. This spaces the structural part of the stiffener well away from a neutral bending axis of the assembly so enhances its stiffness.

[0021] A further aspect of the invention provides an aircraft fuel tank comprising a panel assembly according to the first aspect. The aircraft fuel tank may be a wing tank, a centre tank, or any other fuel tank of an aircraft.

[0022] The aircraft panel assembly of the first aspect is particularly well suited to use in an aircraft fuel tank, because the structural parts of the stiffeners can hold the rib foot beam down against the panel and prevent fuel pressure forces from detaching the rib foot beam from the panel.

[0023] A further aspect of the invention provides an aircraft wing comprising an upper wing cover assembly; a lower wing cover assembly; and a rib connecting the upper wing cover assembly to the lower wing cover assembly; wherein at least one of the wing cover assemblies comprises a panel assembly according to the first aspect with its rib foot beam attached to the rib between the intersections.

[0024] The rib foot beam may be attached to the rib by fasteners, or it may be bonded to the rib [0025] A further aspect of the invention provides an aircraft wing comprising an upper cover; a lower cover, ribs connecting the upper cover to the lower cover; and a plurality of stiffeners attached to the upper and lower covers, wherein each rib is joined to each cover by a respective rib/cover connection arrangement, at least one of the rib/cover connection arrangements comprises a rib foot beam which crosses the stiffeners at a series of intersections and is attached to a respective one of the ribs between the intersections, and at each intersection the rib foot beam is located between the cover and at least part of a respective one of the stiffeners.

[0026] Optionally the wing further comprises a fuel tank between the upper and lower covers [0027] A further aspect of the invention provides a method of manufacturing an aircraft panel assembly, the panel assembly comprising a panel; a plurality of stiffeners; and a rib foot beam, the method comprising: providing a fibre preform on a forming tool, wherein the fibre preform comprises a plurality of stiffener preforms; laying the rib foot beam across the stiffener preforms so that it crosses the stiffener preforms at a series of intersections; laying a panel preform over the fibre preform and the rib foot beam so that at each intersection the rib foot beam is located between the panel preform and a structural part of a respective one of the stiffener preforms; co-infusing the panel preform and the fibre preform with a matrix material; and curing the matrix material.

[0028] Optionally the providing step comprises cutting fibres to form cut fibres, and spraying the cut fibres onto the forming tool. The cut fibres may be cut rovings (bundles of fibres), individual fibres, or cut fibres in any other form.

[0029] A further aspect of the invention provides a method of manufacturing an aircraft panel assembly, the panel assembly comprising a panel and a plurality of stiffeners, the method comprising: providing a fibre preform on a forming tool by cutting fibres to form cut fibres and spraying the cut fibres onto the forming tool, wherein the fibre preform comprises a plurality of stiffener preforms; laying a panel preform over the fibre preform; co-infusing the panel preform and the fibre preform with a matrix material; and curing the matrix material. The cut fibres may be cut rovings (bundles of fibres), individual fibres, or cut fibres in any other form. The panel assembly may be provided with a rib foot beam as in the previous aspect, or it may have no rib foot beam.

[0030] The following comments apply, where appropriate, to either of the above methods.

[0031] Preferably the fibre preform comprises a plurality of stiffener foot preforms, and each stiffener foot preform nms continuously between an adjacent pair of the stiffener preforms.

[0032] Preferably each stiffener preform comprises pair of web preforms and a crown preform which joins the pair of web preforms to each other.

[0033] The panel may be laid in direct contact with the fibre preform (and optionally the rib foot beam) so the panel preform becomes directly bonded to the fibre preform (and optionally the rib foot beam) by the matrix material. Alternatively, additional dry fibre material may be laid over the fibre preform and the rib foot beam before the panel preform is laid over them. In this case the panel preform is not in direct contact with the fibre preform and/or the rib foot beam.

[0034] Optionally the fibre preform comprises a rib foot preform; and the rib foot beam is laid in the rib foot preform [0035] Fibrous reinforcement material may be laid into each stiffener preform so that at each intersection the rib foot beam is located between the panel preform and the fibrous reinforcement material of a respective one of the stiffener preforms.

[0036] Optionally the stiffener preforms are formed on male features of the forming tool, and the method further comprises separating the fibre preform and the forming tool before laying the rib foot beam across the stiffener preforms and/or before laying the panel preform over the fibre preform.

[0037] Alternatively the stiffener preforms may be formed in grooves of the forming tool. In this case it is not necessary to separate the fibre preform and the forming tool before laying the rib foot beam across the stiffener preforms and/or before laying the panel preform over the fibre preform.

[0038] Cores may be fitted into each stiffener preform [0039] The fibre preform may be compacted on the forming tool with a compaction tool before the rib foot beam is laid across the stiffener preforms and/or before laying the panel preform over the fibre preform.

BRIEF DESCRIPTION OF TI-TIE DRAWINGS

[0040] Embodiments of the invention will now be described with reference to the accompanying drawings, in which: [0041] Figure 1 shows an aircraft; [0042] Figure 2 is a schematic view of the wing box of one of the wings; [0043] Figure 3 shows the rib/cover connection arrangements for one of the ribs; [0044] Figure 4 shows an aircraft panel assembly according to a first embodiment of the invention; [0045] Figure 5 is a plan view of the assembly of Figure 4; [0046] Figure 6 is a cross-section across one of the stiffeners taken along line A-A in Figure 5, [0047] Figure 7 is a schematic view of two adjacent stringers; [0048] Figure 8 is a cross-section along one of the stiffeners at an intersection, taken along line B-B in Figure 6; [0049] Figure 9 is a cross-section across one of the rib feet, taken along line D-D in Figure 5; [0050] Figure 10 shows a female forming tool; [0051] Figure 11 shows a fibre preform on the female forming tool of Figure 10; [0052] Figure 12 shows a male compaction tool fitted with the fibre preform between the male compaction tool and the female forming tool; [0053] Figure 13 is a plan view of the apparatus of Figure 13; [0054] Figure 14 a cross-section across one of the stiffener preforms taken along line E-E in Figure 13, [0055] Figure 15 shows a rib foot beam and batten preforms laid onto the fibre preform; [0056] Figure 16 shows form cores and additional material; [0057] Figure 17 shows a cover panel laid onto the other parts; [0058] Figure 18 shows the final preform assembly; [0059] Figure 19 is a plan view of the final preform assembly on an infusion tool, [0060] Figure 20 a cross-section across one of the stiffener preforms taken along line F-F in Figure 19; [0061] Figure 21 shows a male forming tool; [0062] Figure 22 shows a fibre preform on the male forming tool of Figure 21; [0063] Figure 23 is a cross-section across one of the stiffener preforms during compaction; [0064] Figure 24 shows a rib foot beam, spar core and batten preforms laid onto the fibre preform, [0065] Figure 25 shows form cores, additional material and a cover panel laid onto the assembly of Figure 24; [0066] Figure 26 shows the final preform assembly on an infusion tool, and [0067] Figure 27 shows an alternative panel assembly. DETAILED DESCRIPTTON OF EMBODIMENT(S) [0068] Figure 1 shows an aircraft 1 with port and starboard wings 2, 3. Each wing has a cantilevered structure with a length extending in a spanwise direction from a root to a tip, the root being joined to an aircraft fuselage 4. The wings 2, 3 are similar in construction so only the starboard wing 3 will be described in detail with reference to Figures 2 and 3 [0069] The main structural element of the wing is a wing box formed by upper and lower cover panels 21, 22 and front and rear spars 6, 7 shown in cross-section in Figure 3. The cover panels 21, 22 and spars 6, 7 are each Carbon Fibre Reinforced Polymer (CFRP) laminate components. Each cover panel has a curved aerodynamic surface (the upper surface of the upper cover panel 21 and the lower surface of the lower cover panel 22) over which air flows during flight of the aircraft. Each cover panel also has an inner surface carrying a series of stiffeners 8 extending in the spanwise direction. Each cover panel carries a large number of stiffeners 8, only five of which are shown in Figure 2 and only six of which are shown in Figure 3 for purposes of clarity. Each stiffener 8 is joined to one cover panel but not the other. In the case of an aircraft wing cover panel, the stiffeners 8 are commonly referred to as stringers, but the term "stiffeners" will be used below.

[0070] The wing box also has a plurality of transverse ribs, each rib being joined to the covers 21, 22 and the spars 6, 7. The ribs include an inner-most inboard rib 10 located at the root of the wing box, and a number of further ribs spaced apart from the innermost rib along the length of the wing box. The wing box is divided into two fuel tanks: an inboard wing fuel tank bounded by the inboard rib 10, a mid-span rib 11, the cover panels 21, 22 and the spars 6, 7; and an outboard wing fuel tank bounded by the mid-span rib 11, an outboard rib 12 at the tip of the wing box, the cover panels 21, 22 and the spars 6, 7.

[0071] The inboard rib 10 is an attachment rib which forms the root of the wing box and is joined to a centre wing box 20 within the body of the fuselage 4. Baffle ribs 13 (shown in dashed lines) form internal baffles within the fuel tanks which divide the fuel tanks into bays. The ribs 10, 11, 12 are sealed to prevent the flow of fuel out of the two fuel tanks, but the baffle ribs 13 are not sealed so that fuel can flow across them between the bays. As can be seen in Figure 2, the stiffeners 8 stop short of the inboard rib 10 and the outboard rib 12, but pass through the baffle ribs 13 and the mid-span rib 11.

[0072] Figure 4 shows a panel assembly which includes the cover 22; and four of the stiffeners 8 carried on the inner surface of the cover 22. As shown most clearly in Figure 6, each stiffener 8 comprises a pair of feet 25; a pair of webs 28; and a crown 26 which joins the pair of webs 28 to each other. The feet 25, the webs 28 and the crown 26 together form a top-hat shaped profile in cross-section.

[0073] The feet 25 are directly bonded to the inner surface of the lower cover panel 22 by a co-infused joint, and the crown 26 is spaced apart from the lower cover panel 22 at an apex of the stiffener. Each stiffener also has a foam core 24 positioned between the webs 28 at the base of the stiffener and carried by the lower cover panel 22; and a batten 27 positioned between the webs 28 and in contact with the crown 26 at the apex of the stiffener. The batten 27 is made of fibre-reinforced composite material, such as carbon fibre-reinforced epoxy resin.

[0074] The crown 26 and batten 27 provide structural parts of the stiffener, spaced well away from the neutral bending axis at the apex of the stringer to enhance its bending stiffness. The core 24 acts as a filler, and as a support for the batten 27.

[0075] A rib foot beam 18 crosses the stiffeners at a series of intersections 23, one of which is shown in cross-section in Figure 8. The foam core 24 is formed in segments, and at the intersection 23 the rib foot beam 18 passes through slots in the webs 28 and through the gap between the two adjacent segments of the foam core 24 as shown in Figure 8. As also shown in Figure 8, at each intersection 23 the rib foot beam 18 is located between the lower cover panel 22 and the structural parts (i.e. the crown 26 and batten 27) of a respective one of the stiffeners 8.

[0076] At each intersection 23 the structural parts 26, 27 of the respective one of the stiffeners extend continuously across the intersection, unlike the foam core 24 which is broken into segments as shown in Figure 8 and the webs 28 which are formed with slots to allow passage of the rib foot beam 18. This continuity of the structural parts 26, 27 is advantageous because it enhances the bending stiffness of the stiffener, and enables carbon fibres in the structural parts 26, 27 to extend continuously across the intersection 23.

[0077] The bottom edge of the rib foot beam 18 is bonded to the lower cover panel 22 panel by a co-bonded joint, and at each intersection 23 the rib foot beam 18 is also co-bonded to the end faces of the foam core segments 24 and the batten 27 as shown in Figure 8 [0078] The feet 25, webs 28 and crowns 26 of all of the stiffeners are formed from a single layer 29 of fibre-reinforced epoxy resin, so that the feet 25 run continuously between adjacent stiffeners as shown in Figure 7 rather than terminating at distinct edges between the adjacent stiffeners The layer 29 also covers the rib foot beam 18 as shown in Figure 9 to form a rib foot 19 between each pair of stiffeners as shown in Figure 4.

[0079] Each rib 10, 11 13 connects the upper cover 21 to the lower cover 22, and Figure 3 shows the upper and lower rib/cover connection arrangements for the rib 11 by way of example. The rib feet 19 (including the rib foot beam 18) are attached to the rib 11 between the stiffeners by fasteners 14 which pass through the rib 11 and the rib feet 19. The stiffeners 8 pass through mouse-hole openings 20 in the rib 11.

[0080] As noted above, the upper and lower cover panels 21, 22 provide the upper and lower walls respectively of a fuel tank. If the fuel tank is over-filled, then large fuel pressure forces can be generated which risk detaching the rib feet 19 from the cover panel. The interlocking rib foot/stringer arrangement enables the structural parts 26, 27 of the stiffeners to hold the rib foot beam 18 down against the cover panel and prevent fuel pressure forces from separating the rib foot beam 18 from the cover panel.

[0081] The use of a single rib foot beam 18 makes the assembly less complex to manufacture It is also not necessary to align multiple rib feet with each other.

[0082] A first method of manufacturing an aircraft panel assembly is shown in Figures 10-20. The panel assembly is similar to the panel assembly of Figure 4, and the same reference numbers will be used to indicate equivalent components. The method involves the use of various dry-fibre preforms which become infused with epoxy resin to form a composite component in the final part. Such preforms are given the same reference numbers as the component they form in the final infused part, appended by the letter "a". So for example the fibre preform 29a becomes the layer 29 in the final infused part.

[0083] A female forming tool 100 shown in Figure 10 has a criss-crossing arrangement of rib foot tool grooves 111 and stiffener tool grooves 110. A fibre preform 29a is first provided on the tool as shown in Figure 11. The fibre preform 29a is made of a dry carbon fibre material which may be provided on the forming tool 100 in a number of different ways dependent on the requirements of the part, such as the required fibre volume or the load carrying areas of the part.

[0084] In one example the fibre preform 29a may be provided on the forming tool 100 by press-forming a carbon fibre veil, a fleece, a long fibre carpet, or a stretch-breaking fabric, which is then stabilised using a binder. The binder is then activated by, for example, heat applied by hot air, friction or some other method. The veil, fleece, carpet or fabric may be laid up as a single layer or as a patchwork of segments which collectively form a single layer.

[0085] Alternatively, the fibre preform 29a may be deposited using the method disclosed in US2015/0273736 Al, by spraying cut rovings onto the female forming tool 100. Reinforced fibre bundles (known as rovings) are pulled from a bobbin and cut to length in a deposition head. The cut rovings are sprayed onto the surface of the female forming tool 100 using the deposition head, which is preferably mounted on a robot to enable a repeatable result. The rovings may be equipped with a binder, or the binder may be applied in-situ to the cut rovings during the spraying process. The female forming tool 100 is preferably porous and under vacuum to pull the cut rovings to the surface of the tool 100 during the spraying process. This process provides a quasi-isotropic preform 29a with the fibres oriented randomly.

[0086] As shown in Figure 11, the fibre preform 29a comprises stiffener preforms 110a in the stiffener tool grooves 110 and stiffener foot preforms 25a running continuously between the stiffener tool grooves 110. Each stiffener preform 110a comprises a pair of web portions 28a and a crown portion. The fibre preform 29a also comprises a rib foot preform Illa in the tool groove 111.

[0087] Next, a male compaction tool 101 with complementary male ridges is fitted as shown in Figure 12 to clamp the fibre preform 29a between the male and female tools 101, 100 as shown in Figure 14. The fibre preform 29a is then compacted between the tools 101, 100 to de-bulk it and mould it into its desired shape.

[0088] Next the male tool 101 is removed, and batten preforms 27a are laid in the stiffener preforms 110a as shown in Figure 15. The batten preforms 27a are made of a dry carbon fibrous reinforcement material with continuous fibres running along their length. The continuous fibres may be deposited by automated fibre placement.

[0089] Next the rib foot beam 18 is laid in the rib foot preform 111a as shown in Figure 11, across the batten preforms 27a The rib foot beam 18 is fitted as a single cured piece of laminated prepreg composite material, and is in contact with the fibrous reinforcement material of the batten preforms 27a at each intersection.

[0090] Next the foam core segments 24 are laid on top of the batten preforms 27a as shown in Figure 16, and additional dry fibre noodles 114 (shown in Figure 16) and 128 (shown in Figure 20) are inserted to fill the gaps [0091] Next a lower cover panel preform 22a is laid on top as shown in Figure 17, in contact with the fibre preform 29a, the rib foot beam 18 and the noodles 114, 128. The assembly is then inverted and the female forming tool 100 is removed to leave the preform assembly shown in Figure 18. The preform assembly is placed on an infusion tool 402 shown in Figures 19 and 20, and covered by a vacuum bag 400. A vacuum is then applied to a vacuum port 403 shown in Figure 20, and epoxy resin is injected via an inlet 401. The resin co-infuses the dry fibre components of the preform assembly (the fibre preform 29a, the batten preforms 27a, the cover panel preform 22a and the noodles 114, 128) and the resin is then cured.

[0092] Co-infusing the preforms 29a, 27a, 22a and noodles 114, 128 with epoxy resin matrix material, and then curing the matrix material, forms various co-infused and co-bonded joints between the elements of the preform assembly.

[0093] In the example of Figures 10-20 the fibre preform 29a is formed on a female forming tool 100 and then compacted between the female forming tool 100 and a male compaction tool 101. In the alternative example shown in Figures 21-26, the opposite is the case: a fibre preform 229a is formed on a male forming tool 200 and then compacted between the male forming tool 200 and a female compaction tool 201. So in the case of Figures 10-20 the stiffener preforms are formed in grooves of the female forming tool 100, whereas in the case of Figure 22 the stiffener preforms are formed on male features of the male forming tool 200.

[0094] The panel assembly formed by the method of Figures 21 to 26 is similar to the panel assembly previously described, and the same reference numbers will be used to indicate equivalent components, incremented by 200. The method involves the use of various dry-fibre preforms which become infused with epoxy resin to form a composite component in the final part Such preforms are given the same reference numbers as the component they form in the final part, appended by the letter "a". So for example the fibre preform 229a becomes the layer 229 in the final part.

[0095] The male forming tool 200 shown in Figure 21 is provided with a criss-crossing arrangement of rib foot tool ridges 211 and stiffener tool ridges 210. The male forming tool also has a spar tool ridge 230 and a gutter 232 at one edge.

[0096] A fibre preform 229a is formed on the male forming tool 200 as shown in Figure 22. The fibre preform 229a is made of a dry carbon fibre material which may be provided on the male forming tool 200 by the same process that the fibre preform 29a is provided on the female forming tool 100, i.e. by press-forming or spraying.

[0097] As shown in Figure 22, the fibre preform 229a comprises stiffener preforms 210a on the stiffener tool ridges 210, stiffener foot preforms 225a between the ridges 210a and a spar preform 231a on the spar tool ridge 230. Each stiffener preform 210a comprises a pair of web portions 228a and a crown portion 226a. The fibre preform 229a also comprises rib foot preforms 211a laid on the rib foot tool ridges 211.

[0098] Each stiffener foot preform 225a runs continuously between an adjacent pair of the stiffener preforms 210a; and each stiffener preform 210a comprises a pair of web preforms 228a and a crown preform 226a which joins the pair of web preforms to each other.

[0099] The spar preform 231a has a high aspect ratio (height/width) which would be difficult to form inside a groove of a female forming tool, so the male profile of the male forming tool 200 is preferred.

[00100] Next, a female compaction tool 201 with complementary female grooves is fitted to clamp the fibre preform 229a between the male and female tools as shown in Figure 23. The fibre preform 229a is then compacted between the tools to de-bulk it and mould it into its desired shape.

[00101] Next the male forming tool 200 is removed, batten preforms 227a shown in Figure 24 are laid inside the stiffener preforms 210a, and then the rib foot beam 218 is laid inside the rib foot preform 211a across the batten preforms 227a. The rib foot beam 218 is fitted as a single cured piece of laminated prepreg composite material, and is in contact with the fibrous reinforcement material of the batten preforms 227a at each intersection.

[00102] A spar core 214 is also fitted inside the spar preform 231a. The spar core 214 may be fitted as a single cured piece of laminated prepreg composite material, and made of a dry carbon fibrous reinforcement material, including continuous fibres running along the length of the spar core 214.

[00103] Next the foam core segments 224 are laid on top of the batten preforms 227a as shown in Figure 25, and additional dry fibre material inserted to fill the gaps This additional dry fibre material includes noodles (not shown), a first fibre preform sheet 240 with a flange 241 fitted into a gap on one side of the spar core 214, and a second fibre preform sheet 242 with a flange 243 fitted into a gap on the other side of the spar core 214 [00104] Next a panel preform 222a is laid on top as shown in Figure 25, in contact with the fibre preform sheets 240, 242 and the spar core 214. The assembly is then inverted and placed on an infusion tool 202 shown in Figure 26 The female compaction tool 201 is then removed and replaced by a vacuum bag (not shown) and the panel assembly is infused using the same process as previously described to produce the panel assembly shown in Figure 26 on the infusion tool 202. The resin co-infuses the dry fibre components of the assembly (the fibre preform 229a, the batten preforms 227a, the panel preform 222a and the additional material including the preform sheets 240, 242) and is then cured to form the final panel assembly shown in Figure 26 [00105] In Figure 17 the panel preform 22a is laid in direct contact with the fibre preform 29a and the rib foot beam 18 so the panel preform 22a becomes directly bonded to the fibre preform 29a and the rib foot beam 18 by the resin matrix material. In the case of Figure 25, the additional dry fibre material 240, 242 has been laid over the fibre preform 229a and the rib foot beam 218 before the panel preform 222a is laid over them.

So in this case the panel preform 222a is not in direct contact with the fibre preform 229a or the rib foot beam 218 [00106] The panel assembly of Figure 26 includes a cover 222; and a spar 231 and stiffeners 208 which are carried on the inner surface of the cover 222. Each stiffener 208 comprises a pair of feet 225; a pair of webs 228; and a crown 226, all formed from a single layer 229 which also covers the rib foot beam between the stiffeners. Each stiffener also has a foam core 224 and a batten 227.

[00107] Figure 27 shows an alternative panel assembly which is similar to the assembly of Figure 4, and the same reference numbers are used to indicate equivalent components. The rib foot beam has a height perpendicular to the cover panel which is smaller at the intersections 23 (height h2) than between the intersections (height h1). The reduced height h2 at the intersections 23 reduces the size of the slot or other opening which is formed in the stiffener to enable the rib foot beam to pass through. The increased height hl of the rib feet between the intersections provides more surface area to accommodate bolts or other fasteners.

[00108] Where the word 'or' appears this is to be construed to mean 'and/or' such that items referred to are not necessarily mutually exclusive and may be used in any appropriate combination.

[00109] Although the invention has been described above with reference to one or more preferred embodiments, it will be appreciated that various changes or modifications may be made without departing from the scope of the invention as defined in the appended claims.

Claims

CLAIMS1 An aircraft panel assembly comprising: a panel; and a plurality of stiffeners on the panel, wherein each stiffener comprises an attachment part attached to the panel arid a structural part spaced apart from the panel; and a rib foot beam which crosses the stiffeners at a series of intersections, wherein at each intersection the rib foot beam is located between the panel and the structural part of a respective one of the stiffeners.
2 An aircraft panel assembly according to claim 1 wherein the attachment part of each stiffener is bonded to the panel.
3 An aircraft panel assembly according to any preceding claim wherein the rib foot beam is bonded to the panel.
4 An aircraft panel assembly according to any preceding claim wherein the rib foot beam is bonded to the stiffeners at the intersections.
An aircraft panel assembly according to any preceding claim wherein at each intersection the structural part of the respective one of the stiffeners extends continuously across the intersection.
6 An aircraft panel assembly according to any preceding claim wherein at each intersection the structural part of the respective one of the stiffeners comprises reinforcement fibres which extend continuously across the intersection 7 An aircraft panel assembly according to any preceding claim further comprising a layer which forms the structural parts all of the stiffeners and covers the rib foot beam between the stiffeners.8 An aircraft panel assembly according to any preceding claim wherein the stiffeners are bonded to an inner surface of the panel, and the panel has an outer surface with a curved aerodynamic profile.9 An aircraft panel assembly according to any preceding claim wherein the attachment part of each stiffener comprises a pair of feet and a pair of webs; and the structural part of each stiffener comprises a crown which joins the pair of webs to each other.An aircraft panel assembly according to any preceding claim wherein the attachment part of each stiffener comprises a pair of feet and a pair of webs; and the structural part of each stiffener comprises a crown which joins the pair of webs to each other and a reinforcement element between the webs 11 An aircraft fuel tank comprising an aircraft panel assembly according to any preceding claim.12 An aircraft wing comprising an upper wing cover assembly; a lower wing cover assembly; and a rib connecting the upper wing cover assembly to the lower wing cover assembly, wherein at least one of the wing cover assemblies comprises an aircraft panel assembly according to any of claims 1 to 10 with its rib foot beam attached to the rib between the intersections.13 An aircraft wing comprising an upper cover; a lower cover; ribs connecting the upper cover to the lower cover; and a plurality of stiffeners attached to the upper and lower covers, wherein each rib is joined to each cover by a respective rib/cover connection arrangement, at least one of the rib/cover connection arrangements comprises a rib foot beam which crosses the stiffeners at a series of intersections and is attached to a respective one of the ribs between the intersections, and at each intersection the rib foot beam is located between the cover and at least part of a respective one of the stiffeners 14 A method of manufacturing an aircraft panel assembly, the panel assembly comprising a panel; a plurality of stiffeners; and a rib foot beam, the method comprising: a, providing a fibre preform on a forming tool, wherein the fibre preform comprises a plurality of stiffener preforms; b. laying the rib foot beam across the stiffener preforms so that it crosses the stiffener preforms at a series of intersections; c. laying a panel preform over the fibre preform and the rib foot beam so that at each intersection the rib foot beam is located between the panel preform and a structural part of a respective one of the stiffener preforms; d. co-infusing the panel preform and the fibre preform with a matrix material; and e, curing the matrix material A method according to claim 14 wherein the fibre preform comprises a rib foot preform, and step b comprises laying the rib foot beam in the rib foot preform 16 A method according to claim 15 further comprising fitting fibrous reinforcement material into each stiffener preform after step a and before step b so that at each intersection the rib foot beam is located between the panel preform and the fibrous reinforcement material of a respective one of the stiffener preforms.17 A method according to claim 14, 15 or 16 wherein the stiffener preforms are formed on male features of the forming tool, and the method further comprises separating the fibre preform and the forming tool after step a, and before step b.18 A method according to any of claims 14 to 17 wherein the stiffener preforms are formed in grooves of the forming tool.19 A method according to any of claims 14 to 18 wherein step a. comprises cutting fibres to form cut fibres, and spraying the cut fibres onto the forming tool.A method according to any of claims 14 to 19 further comprising after step a. and before step b., compacting the fibre preform on the forming tool with a compaction tool.21 A method of manufacturing an aircraft panel assembly, the panel assembly comprising a panel and a plurality of stiffeners, the method comprising: a providing a fibre preform on a forming tool by cutting fibres to form cut fibres and spraying the cut fibres onto the forming tool, wherein the fibre preform comprises a plurality of stiffener preforms, b. laying a panel preform over the fibre preform; c, co-infusing the panel preform and the fibre preform with a matrix material; and d. curing the matrix material 22 A method according to claim 21 wherein the stiffener preforms are formed on male features of the forming tool, and the method further comprises separating the fibre preform and the forming tool after step a and before step b 23 A method according to claim 21 wherein the stiffener preforms are formed in grooves of the forming tool 24 A method according to any of claims 21 to 23 further comprising after step a. and before step b., compacting the fibre preform on the forming tool with a compaction tool