Aircraft structure and construction method thereof

Abstract

An aircraft structure and method of construction thereof that avoids inefficiencies associated with current methods of constructing aircraft structures, reduces the manufacturing time required to construct an aircraft structure, and reduces the costs involved in constructing an aircraft structure. The aircraft structure and method of construction thereof include a geodesic or multi-grid frame of intersecting thermoplastic composite strips joined to the inner surface of a thermoplastic composite fuselage skin panel by overmolding and / or co-consolidation of the grid frame and the skin panel.

Description

Technical Field

[0001] The present disclosure relates to an aircraft structure and method of construction thereof that avoids the inefficiencies involved in current methods of constructing aircraft structures, reduces the manufacturing time required to construct the aircraft structure, and reduces the costs involved in constructing the aircraft structure. More specifically, the aircraft structure and method of construction thereof of the present disclosure includes a geodesic or multi-grid frame of intersecting thermoplastic composite strips that are joined to the inner surface of a thermoplastic composite fuselage skin panel by overmolding and / or co-consolidation of the grid frame and the skin panel. Background Art

[0002] In traditional basic aircraft construction, the skin panels of the aircraft that make up the fuselage of the aircraft are attached and supported on a reinforced frame. The frame usually has a generally rectangular configuration consisting of a frame extending around the periphery of the fuselage and stringers that can be attached to the frame and extend in the longitudinal direction of the fuselage. The skin panels that form the fuselage of the aircraft are attached to the stringers by fasteners, by adhesives, by co-curing or by other equivalent means. The stringers help carry the bending loads in the fuselage and support the skin panels to prevent them from bending. If the stringers are connected to the frames, the stringers can also transfer loads from the skin panels of the aircraft fuselage to the frames of the aircraft frame.

[0003] The cost of constructing an aircraft having an airframe consisting of stringers and frames contributes significantly to the overall cost of the aircraft. The stringers and frames must often be formed on complex contours, which in the case of composite structures may take several manufacturing steps and may require extra attention to avoid wrinkles. Positioning and attaching the stringers and frames relative to each other and to the skin panels of the fuselage contribute significantly to the time required to construct the aircraft, and thus to the overall cost of the aircraft.

[0004] Constructing the frame of an aircraft from frames and stringers also requires that the frames and stringers be constructed in a complex pattern that follows the irregular inner surface of the fuselage skin. This contributes significantly to the pre-flight time and tooling costs required to construct the aircraft. In addition, the fasteners used to connect the stringers, frames and skin contribute significantly to the assembly time of the aircraft and the weight of the aircraft.

[0005] Furthermore, since the frame and stringer frame have a generally rectangular configuration, it is problematic to integrate window openings or other generally circular openings into the frame. It is often necessary to bolt the window frame or circular opening frame to the frame after the frame is constructed. Other penetrations such as for drain masts or antennas require localized skin reinforcement, which is achieved by including plies in the composite layup. This adds more time required to manufacture the aircraft and also increases the weight of the aircraft construction. Summary of the invention

[0006] The disclosed thermoplastic multi-grid overmolded / co-consolidated aircraft fuselage structure and method of construction thereof overcomes many of the disadvantages associated with current practices for constructing aircraft fuselages discussed previously. The disclosed aircraft structure employs an overmolding process that significantly simplifies the aircraft structure and significantly reduces the time required to construct the aircraft structure.

[0007] The aircraft structure and method of construction described herein are part of a small portion of a complete aircraft fuselage. However, the aircraft structure and method described herein can be employed in constructing large sections of an aircraft fuselage that are then assembled to form the overall structure of the aircraft fuselage. Furthermore, the aircraft structure and method of construction described herein can be employed in constructing other parts of an aircraft.

[0008] The aircraft structure consists of a skin panel having an outer surface and an opposite inner surface. The outer surface of the skin panel will form a portion of the outer surface of the aircraft fuselage. According to the method of the present disclosure, the skin panel is positioned in a mold, and the outer surface of the skin panel is in contact with the surface of the mold. In an alternative method, a plurality of strips are positioned in the mold, the skin panel is positioned on the plurality of strips and then the mold is closed.

[0009] A plurality of strips are then positioned on the inner surface of the skin panel in the mold. The plurality of strips have an elongated length dimension, a width dimension, and a thickness dimension. The plurality of strips also have a slit that partially extends through the width dimension of the strip. By assembling the plurality of strips together at their slits, the slits enable the plurality of strips to be assembled together into an intersecting, multi-line, or multi-grid configuration. The plurality of strips assembled together into a multi-line or multi-grid configuration are positioned on the inner surface of the skin panel in the mold, and the edges of the plurality of strips are engaged on the inner surface of the skin panel. This positions the width dimension of the plurality of strips in the mold to extend outward from the inner surface of the skin panel.

[0010] The mold is then closed over the multiple strips on the inner surface of the skin panel inside the mold. The heated, liquefied overmolding material is then injected into the mold. The overmolding material flows through the mold and extends between the inner surface of the skin panel and the multiple strips on the inner surface of the skin panel. The heating of the overmolding material causes the contact portion of the inner surface of the skin panel with the multiple strips to melt, fusing the inner surface of the skin panel with the contact portion of the multiple strips. The overmolding material injected into the mold is then allowed to cool, forming a joint between the inner surface of the skin panel and the contact portion of the multiple strips. In this way, the overmolding material fixes the multiple strips to the inner surface of the skin panel and fixes the multiple strips to each other at the connection position with the slots.

[0011] The mold is then opened and the aircraft structure consisting of the skin panel, the plurality of strips on the inner surface of the skin panel, and the overmolding material securing the plurality of strips to the inner surface of the skin panel is removed from the mold.

[0012] According to the above method, an aircraft fuselage structure is constructed consisting of a skin panel, a plurality of intersecting strips secured to an inner surface of the skin panel, and an overmold material securing the plurality of strips to the inner surface of the skin panel.

[0013] Additional cavities in the mold can be filled with overmolding material to form smaller reinforcing elements perpendicular to the skin and / or multiple strips to support the skin, multiple strips, or both and prevent bending or introduction of loads, and increase the thickness of the skin or multiple strips to act as multipliers in the locations where they will penetrate.

[0014] The features, functions, and advantages that have been discussed can be achieved independently in various embodiments, but may also be achieved in combination with other embodiments, further details of which can be understood with reference to the following written description and drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] Figure 1 is an illustration of a cross-sectional elevation view of a thermoplastic multi-grid overmolded / co-consolidated aircraft fuselage constructed in accordance with the method of the present disclosure.

[0016] Figure 2 yes Figure 1 Illustration of a top perspective view of a thermoplastic multi-grid overmolded / co-consolidated fuselage structure of an aircraft fuselage represented in.

[0017] Figure 3 is a flow chart of a method of constructing a thermoplastic multi-grid overmolded / co-consolidated fuselage structure of the present disclosure.

[0018] Figure 4 is with Figure 2 and showing examples of overmolded material formed into reinforcement fillets between intersecting reinforcement strips and formed into reinforcement triangles. DETAILED DESCRIPTION

[0019] Figure 1 is an illustration of a cross-sectional elevation view of an aircraft fuselage 10 consisting of reinforcement strips 12 joined and secured to the inner surface of a skin 14 of the aircraft fuselage 10. The reinforcement strips 12 and the skin 14 of the fuselage are constructed of a fiber reinforced thermoplastic composite material. The reinforcement strips 12 intersect and are secured together in geodesic lines or multiple grids, and the frames 16 are secured to the inner surface of the skin 14. The skin 14 consists of a plurality of aircraft skin panels that form the skin of the fuselage 10.

[0020] Figure 2is a diagrammatic representation of a single skin panel 18 of fiber reinforced thermoplastic composite material comprising multiple skin panels of an aircraft fuselage 10 and a plurality of strips 22 of fiber reinforced thermoplastic composite material secured to an inner surface 24 of the skin panel 18. It should be understood that Figure 2 The skin panel 18 and strip 22 only consist of Figure 1 The entire multi-grid frame 16 and a portion of the fuselage skin 14 of the aircraft fuselage 10 are shown in FIG. Figure 1 The overall construction of the multi-grid frame 16 and the fuselage skin 14 is repeated in similar sub-assemblies as shown in Figure 2 The skin panel 18 and the plurality of strips 22 are constructed.

[0021] Figure 2 The skin panel 18 shown in FIG. 1 is a thermoplastic composite laminate. The thickness of the skin panel 18 is selected to best adapt the skin panel 18 to the loads to which the aircraft fuselage 10 will be subjected in use. Figure 2 The perimeter edges of the skin panel 18, represented by the four sides 26, 28, 32, 34, are cut and shaped into the desired configuration. Figure 2 1 has an overall rectangular configuration, but the perimeter edges of the skin panel 18 may be cut and shaped into any other desired configuration. Figure 2 Also shown is an opening 38 through the skin panel 18 . Figure 2 The opening 38 is relatively small, but represents any type of opening, such as a window opening or system penetration that would be required through the skin panel 18 .

[0022] exist Figure 2 FIG. 4 shows a padup strip 42 on the inner surface 24 of the skin panel 18. The padup 42 strengthens the skin panel 18 in the area where the padup 42 is located. Figure 2 4 is shown as having a generally elongated rectangular configuration. However, the pads 42 may have any other configuration to reinforce the skin panel 18 in the area where the pads 42 are located. The pads 42 also have a stepped cross-sectional configuration that provides greater thickness and greater reinforcement to the skin panel 18 at the thicker center portion of the pad. The pads 42 are also thermoplastic composite laminates and may be integrated into the laminates that make up the skin.

[0023] A plurality of reinforcing strips 22 are positioned on the inner surface 24 of the skin panel 18. Each of the plurality of reinforcing strips 22 may be cut to shape and cut from a blank of the composite laminate, or they may be laid up with a strip or tape of composite material. Figure 2, represented as flat laminate strips of composite material. However, each of the plurality of strips 22 can be pre-bent by laying the strips on a curved surface or by bending a flat laminate under heated pressure. Each of the plurality of reinforcements 22 can be formed from a multilayer laminate including reinforcements having fibers oriented in different directions (e.g., +60°, -60° and / or 0° or any other orientation). Each of the plurality of reinforcements 22 can also be formed from a fitting, a bracket, or other similar geometric shape. Each of the plurality of reinforcement strips 22 can also be formed from a reinforcing flange or other similar reinforcing element. Figure 2 Five reinforcing strips 46, 48, 52, 54, 56 are shown in FIG. However, the number of reinforcing strips 22 may be greater. Figure 2 The five reinforcement strips 46, 48, 52, 54, 56 shown in the figure are more or less. Each of the reinforcement strips 46, 48, 52, 54, 56 is a thermoplastic composite laminate. Each of the reinforcement strips 46, 48, 52, 54, 56 is cut from the laminate in a desired shape or configuration. Figure 2 Each of the reinforcement strips 46, 48, 52, 54, 56 shown in FIG. 1 has a generally elongated rectangular configuration. However, the reinforcement strip 22 may be cut and formed in other configurations. Figure 2 2 are shown as having substantially the same length, width, and thickness dimensions. However, the reinforcement strips 22 may have different lengths, widths, and thickness dimensions. Figure 2 A reinforcement strip 56 formed in the center of the skin panel 18 is shown as having a center portion 58 having a greater thickness dimension and a greater width dimension than the other reinforcement strips 46, 48, 52, 54. Each of the reinforcement strips 46, 48, 52, 54, 56 has a bottom edge 62, 64, 66, 68, 72, respectively. The bottom edges 62, 64, 66, 68, 72 of the respective strips 46, 48, 52, 54, 56 are trimmed, cut or shaped so that they closely match the contour of the inner surface 24 and conform to and integrally engage the inner surface 24 of the skin panel 18 when the reinforcement strips 22 are assembled to the inner surface 24 of the skin panel 18. The first strip 46 and the second strip 48 of the plurality of strips 22 are formed with bottom edge cutouts 74, 76, respectively. When the first strip 46 and the second strip 48 are positioned Figure 2 When positioned on the inner surface 24 of the skin panel 18 shown in FIG. 8 , the bottom edge cutouts 74 , 76 are shaped so that they conform to and engage the upper surface of the standoff 42 .

[0024] The reinforcement strips 22 are positioned on the inner surface 24 of the skin panel 18 in an intersecting grid pattern or an intersecting short-distance line or a multi-grid structure. The pattern formed by the reinforcement strips 22 can be locally tailored to accommodate changes in load levels and directions. With the intersecting short-distance line or multi-grid structure of the reinforcement strips 22 positioned on the inner surface 24 of the skin panel 18, the width dimension of the plurality of reinforcement strips 22 extends outward from the inner surface 24 of the skin panel 18. The intersecting reinforcement strips 22 are connected together by providing matching slots in the intersecting strips and then overmolding the material to be described. Other equivalent types of connections between the strips 22 can be used. The adjacent intersecting first strips 46 and the second strips 48 have corresponding slots 82, 84 formed partially through the strips. The slot 82 of the first strip 46 is formed from the bottom edge 62 of the strip to partially pass through the first strip 46. The slot 84 passing through the second strip 48 is formed from the top edge 86 of the second strip to partially pass through the second strip 48. The second strip 48 is inserted into the partial slot 82 of the first strip 46 at the partial slot 84 of the second strip 48, and the first strip 46 is inserted into the partial slot 84 of the second strip 48 at the partial slot 82 of the first strip 46. This forms the intersection joint between the first strip 46 and the second strip 48. The intersections of other strips of the plurality of strips 22 may be joined together in the same manner as the intersection between the first strip 46 and the second strip 48.

[0025] An overmolding composite or overmolding material 92 is disposed on the inner surface 24 of the skin panel 18. The overmolding material 92 may be any type of thermoplastic resin, or any type of thermoplastic resin impregnated with reinforcements (e.g., short fiber reinforcements). The overmolding composite or overmolding material 92 may also be a neat resin, an injection / molding material, or a type of reinforced resin material that is compatible with the resin material of the skin panel 18 and the plurality of reinforcement strips 22. The overmolding material 92 extends over portions of the inner surface 24 of the skin panel 18, over the pads 42, and over portions of the plurality of reinforcement strips 22. The overmolding material 92 secures the plurality of reinforcement strips 22 to the inner surface 24 of the skin panel 18 and the pads 42 and it secures the reinforcement strips to their respective adjacent reinforcement strips. The overmold material 92 may be formed as a reinforcement flange for the plurality of stiffeners 22 or as other types of reinforcement on the inner surface 24 of the skin panel 18 , such as a thicker reinforced perimeter portion surrounding the opening 38 through the skin panel 18 . Figure 4 is an illustration of overmold material 92 formed into a reinforcement fillet 92 ′ at the intersection of the first reinforcement strip 46 and the second reinforcement strip 48 and into a reinforcement triangle 92 ″ between the skin panel 18 and the second reinforcement strip 48 .

[0026] Figure 3 Representation structure Figure 2The method steps involved in the aircraft structure in the embodiment of the present invention are as follows. Initially, the skin panel 18 is positioned in the mold 94. The skin panel 18 is positioned in the mold 94, wherein the outer surface 96 of the skin panel 18 contacts and is supported on the molding surface 98 of the mold 94. The outer surface 96 of the skin panel 18 will form a part of the outer surface of the aircraft fuselage 10.

[0027] The pads 42 and the plurality of reinforcing strips 22 are then positioned on the inner surface 24 of the skin panel 18. The plurality of strips 22 are assembled into a geodesic or multi-grid configuration when positioned in the mold 94 and positioned on the inner surface 24 of the skin panel 18. The plurality of strips 22 have a width dimension extending outwardly from the inner surface 24 of the skin panel 18.

[0028] The mold 94 is then closed over the skin panel 18 and the plurality of reinforcing strips 22 on the inner surface 24 of the skin panel 18. The mold 94 may then be heated to heat the plurality of strips 22 and the inner surface 24 of the skin panel 18. The heat in the mold 94 may partially melt the pads 42, the plurality of reinforcing strips 22, and the inner surface 24 of the skin panel 18, thereby fusing the pads 42, the plurality of reinforcing strips 22, and the inner surface 24 of the skin panel 18 together. Figure 2 With the type of thermoplastic composite materials used in constructing aircraft structures, mold 94 may not need to be heated.

[0029] The heated and liquefied overmold material 92 is then injected into a mold 94. The mold 94 is configured so that it directs the liquefied overmold material 92 through the mold to the contact portions of the pads 42 and the plurality of strips 22 on the inner surface 24 of the skin panel 18 and to the intersections or connections between the plurality of strips 22. Again, depending on the configuration, Figure 2 The heating of the liquefied overmold material 92 causes the pads 42 and the plurality of strips 22 in contact with the inner surface 24 of the skin panel 18 to partially melt and fuse together. The liquefied overmold material 92 is directed through the mold 94 to the mold relative to the inner surface 24 of the skin panel 18. Figure 2 18. The positions of the support pads 42, the plurality of reinforcement strips 22, and the inner surface 24 of the skin panel 18 are shown in FIG. The liquefied overmold material 92 may also form reinforcements around the openings 38 through the skin panel 18 and form reinforcing elements for the inner surface 24 of the skin panel 18 and / or the plurality of reinforcement strips 22. The overmold material 92 injected into the mold 94 is then allowed to cool, thereby forming a joint between the contact portions of the support pads 42 and the plurality of reinforcement strips 22 with the inner surface 24 of the skin panel 18, securing the support pads 42 and the plurality of reinforcement strips 22 to the inner surface 24 of the skin panel 18.

[0030] The mold is then opened and the aircraft structure consisting of the skin panels 18 , the pads 42 and the plurality of reinforcement strips 22 is removed from the mold.

[0031] According to the above method, Figure 2 The aircraft fuselage structure shown in FIG. 1 is constructed to include a skin panel 18, a pad 42 on the inner surface 24 of the skin panel 18, a plurality of intersecting reinforcement strips secured to the inner surface 24 of the skin panel 18, and an overmold material 92 securing the pad 42 and the plurality of reinforcement strips 22 to the inner surface 24 of the skin panel 18 and to each other.

[0032] Although the overmolded structure is described herein as an aircraft structure, the overmolded structure may also be applied to space launch vehicle structures, bulkhead structures, or other such structural applications.

[0033] Since various modifications can be made to the aircraft structure and method of construction thereof described and illustrated herein without departing from the scope of the invention, it is intended that all matter contained in the above description or shown in the accompanying drawings shall be interpreted as illustrative and not restrictive. Therefore, the breadth and scope of the present disclosure should not be limited by any of the above-described exemplary embodiments, but should be defined only in accordance with the following claims appended hereto and their equivalents.

Claims

1. An aircraft structure (10), comprising: a panel (18) having an outer surface (96) and an inner surface (24) on opposite sides of the panel (18); a plurality of strips (22) located on the inner surface (24) of the panel (18), the plurality of strips (22) intersecting, the bottom edges of the plurality of strips (22) being joined to the inner surface of the panel (18) such that the width dimension of the plurality of strips (22) extends in a direction away from the inner surface of the panel (18); an overmold material (92) extending between the plurality of strips (22), the overmold material (92) securing the plurality of strips (22) together, The overmold material (92) extends between an inner surface (24) of the panel (18) and the plurality of strips (22), the overmold material (92) securing the plurality of strips (22) to the inner surface (24) of the panel (18); The plurality of strips (22) include a first strip (46) and a second strip (48), the first strip (46) having a slot (82), and the second strip (48) extending into the slot (82) of the first strip (46); The second strip (48) has a slot (84), and the first strip (46) extends into the slot (84) of the second strip (48); The inner surface (24) of the panel (18) and the contacting portions of the plurality of strips (22) are fused together by melting; and, The inner surface (24) of the panel (18) has a support strip (42) thereon, the first strip (46) and the second strip (48) of the plurality of strips (22) being respectively formed with bottom edge cutouts (74, 76), the bottom edge cutouts (74, 76) being shaped so that they conform to and engage the upper surface of the support strip (42), The support strip (42) has a trapezoidal cross-sectional structure including a thicker central portion and a thinner edge portion, and is arranged so that a lower surface of the support strip (42) in contact with an inner surface (24) of the panel (18) is larger than an upper surface engaged with bottom edge cutouts (74, 76) of the first strip (46) and the second strip (48).

2. The aircraft structure according to claim 1, further comprising: The plurality of strips have edges that are secured to the inner surface of the panel by the overmold material.

3. The aircraft structure according to claim 1, further comprising: said plurality of strips having a length extending over said inner surface of said panel; and The plurality of strips have a width extending outwardly from the inner surface of the panel.

4. The aircraft structure according to claim 3, further comprising: Some of the plurality of strips have different width dimensions.

5. The aircraft structure according to claim 1, further comprising: Some of the plurality of strips have different thickness dimensions.

6. The aircraft structure according to claim 1, further comprising: The plurality of strips are flat laminate strips of composite material.

7. The aircraft structure according to claim 1, further comprising: The overmoulding material is a fibre-reinforced thermoplastic.

8. An aircraft structure (10), comprising: A panel (18) having an outer surface (96) and an inner surface (24) opposite the outer surface; a plurality of strips (22) located on the inner surface (24) of the panel (18), the plurality of strips (22) intersecting each other, the plurality of strips (22) having bottom edges connected to the inner surface (24) of the panel (18), the bottom edges of the plurality of strips (22) being joined to the inner surface of the panel (18) such that the width dimension of the plurality of strips (22) extends in a direction away from the inner surface of the panel (18); an overmold material (92) disposed on the inner surface (24) of the panel (18), the overmold material (92) extending between the inner surface (24) of the panel (18) and the plurality of strips (22), the overmold material (92) securing the plurality of strips (22) to the inner surface (24) of the panel (18), The plurality of strips (22) include a first strip (46) and a second strip (48), the first strip (46) having a slot (82), and the second strip (48) extending into the slot (82) of the first strip (46); The second strip (48) has a slot (84), and the first strip (46) extends into the slot (84) of the second strip (48); and, The inner surface (24) of the panel (18) and the contacting portions of the plurality of strips (22) are fused together by melting; and, The inner surface (24) of the panel (18) has a support strip (42) thereon, the first strip (46) and the second strip (48) of the plurality of strips (22) being respectively formed with bottom edge cutouts (74, 76), the bottom edge cutouts (74, 76) being shaped so that they conform to and engage the upper surface of the support strip (42), The support strip (42) has a trapezoidal cross-sectional structure including a thicker central portion and a thinner edge portion, and is arranged so that a lower surface of the support strip (42) in contact with an inner surface (24) of the panel (18) is larger than an upper surface engaged with bottom edge cutouts (74, 76) of the first strip (46) and the second strip (48).

9. The aircraft structure according to claim 8, further comprising: said plurality of strips having a length dimension extending over said inner surface of said panel; and The plurality of strips have a width dimension extending outwardly from the inner surface of the panel.

10. The aircraft structure according to claim 8, further comprising: Some of the plurality of strips have different width dimensions; and Some of the plurality of strips have different thickness dimensions.

11. The aircraft structure according to claim 8, further comprising: The overmold material extends over the first strip and the second strip at the slot of the first strip.

12. The aircraft structure of claim 8, further comprising: Each strip of the plurality of strips is a flat laminate of composite material.

13. The aircraft structure of claim 8, further comprising: The overmolding material is a thermoplastic.

14. A method of constructing an aircraft structure (10), the method comprising: positioning a panel (18) having an outer surface (96) and an inner surface (24) opposite the outer surface (96) in a mold (94), wherein the outer surface (96) of the panel (18) contacts the mold (94); Positioning a plurality of strips (22) on the inner surface (24) of the panel (18), with bottom edges of the plurality of strips (22) engaged on the inner surface of the panel (18) such that a width dimension of the plurality of strips (22) extends in a direction away from the inner surface of the panel (18), wherein the plurality of strips (22) are arranged in an intersecting grid pattern; Closing the mold (94); injecting an overmolding material (92) into the mold (94) and overmolding the inner surface (24) of the panel (18) and the plurality of strips (22) with the overmolding material (92), The overmold material (92) extends between an inner surface (24) of the panel (18) and the plurality of strips (22), the overmold material (92) securing the plurality of strips (22) to the inner surface (24) of the panel (18); The plurality of strips (22) include a first strip (46) and a second strip (48), the first strip (46) having a slot (82), and the second strip (48) extending into the slot (82) of the first strip (46); The second strip (48) has a slot (84), and the first strip (46) extends into the slot (84) of the second strip (48); The inner surface (24) of the panel (18) and the contacting portions of the plurality of strips (22) are fused together by melting; and, The inner surface (24) of the panel (18) has a support strip (42) thereon, the first strip (46) and the second strip (48) of the plurality of strips (22) being respectively formed with bottom edge cutouts (74, 76), the bottom edge cutouts (74, 76) being shaped so that they conform to and engage the upper surface of the support strip (42), The support strip (42) has a trapezoidal cross-sectional structure including a thicker central portion and a thinner edge portion, and is arranged so that a lower surface of the support strip (42) in contact with an inner surface (24) of the panel (18) is larger than an upper surface engaged with bottom edge cutouts (74, 76) of the first strip (46) and the second strip (48).

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