WO2000024553A1

WO2000024553A1 - Composite laminate cutting

Info

Publication number: WO2000024553A1
Application number: PCT/GB1999/003461
Authority: WO
Inventors: Stephen Williams
Original assignee: Bae Systems Plc
Priority date: 1998-10-22
Filing date: 1999-10-20
Publication date: 2000-05-04
Also published as: GB9823018D0; JP2002528278A; AU6351999A

Abstract

The method comprises positioning a die (16) on the laminate (10) with the laminate in an uncured or part-cured condition, applying a covering (24) to the die (16) and laminate (10) and applying fluid, such as gas, under pressure to the covering (24) to load the die sufficiently to cut through the laminate. In one preferred embodiment the die (16) is left in position after cutting the laminate (10) to form a liner for the cut formed by the die (16).

Description

COMPOSITE LAMINATE CUTTING

The invention relates to the cutting of a composite laminate and to the production of an access opening and access panel arrangement. In paπicuiar, but not exclusively, the invention is concerned with a method of cutting a composite laminate used as a skin on an aerofoil and to the mounting of under-wing access panels to wing skins for aircraft fuel bays.

Existing designs for under-wing access panels for giving access to areas such as fuel bays demand, on inboard locations, that no fasteners for the panels penetrate the wing skin. In order to meet that demand, each panel is normally held in place by a clamping arrangement and is located by a flange machined around the access hole periphery to allow the panel to sit aerodvnamicallv flush with the outer skin surface. For carbon fibre wing skin applications, certain problems arise with that type of design. Firstly, it is not desirable to have a machined rebate around the periphery of an access opening, since it may initiate delamination of the skin in the event of impact on the closure panel. Secondly, as is the case with metal skins, it is not acceptable to have fasteners penetrating a carbon fibre wing skin around the access opening periphery. Thirdly, as the rebate needs to be located on the outside surface, there is a requirement to turn the component over during initial panel manufacture to facilitate machining, which turning is undesirable.

A problem which arises when producing openings in a composite laminate such as a carbon fibre wing skin, is that the opening must be made smaller than is required during initial lay-up of the wing skin fabric. As is well known, the wing skin fabrics are normally impregnated with a suitable epoxy matrix and, in their uncured condition, they are covered with an air tight bag which is then subjected to vacuum to pull the bag into contact with the uncured laminate. The assembly is then placed inside an autoclave and is subject to heat and pressure, the differential pressure acting between the pressurised atmosphere of the autoclave and the vacuum bag compressing the fabric as the wing skin cures. Once cured, the opening in the cured fabric is completed by cutting to net size involving an additional manufacturing step. Another method has been proposed which involves laying up carbon fibre fabrics having pre-cut holes around a mandrel, the mandrel projecting through the holes. However, such a method of laying up is time consuming. One object of the present invention is to provide a method of die cutting the composite laminate which will help to overcome such problems.

According to a first aspect of the invention, there is provided a method of die cutting a composite laminate, the method being characterised by positioning a die on the laminate with the laminate in an uncured or part cured condition, applying a covering to the die and laminate and applying fluid under pressure to the covering to load the die sufficiently for it to cut throu 'gBh' the laminate.

By arranging the die beneath the cover and applying pressure, the cutting can take place during the curing process with the die remaining in position in the laminate during at least part of the curing stage. Therefore, an opening produced in the laminate can be made accurately without requiring a subsequent cutting step to form. say. an access opening in the laminate, as is required with prior methods.

Preferably, the method comprises providing a die having a ring-like cutting section.

One particular advantage of the present invention is that the die can be left in situ to form a liner or a surround for the periphery of the cut formed by the die. In that respect, and according to a second aspect of the invention there is provided a method of die cutting a composite laminate comprising positioning a die on the laminate, applying a load to the die which is sufficient for the die to cut through the laminate and using the die to form a liner or a surround for the periphery of the cut formed by the die.

The method may include leaving the die in position around the periphery of the cut after cutting to form the liner or surround. If desired, an adhesive can be used to secure the die in position around the periphery of the cut.

In either the first or second aspect, a suitable location means may be provided for positioning the die on the laminate.

The method may include applying a vacuum beneath the covering whereby the load applied to the covering and die is created as a result of differential pressure on the covering. As mentioned above, certain problems need to be overcome when fixing an access panel to an aircraft wing skin of composite laminate form.

The method according to the first or second aspect of the invention may include using the die to form an access opening and using the die to form a liner for the access opening, the die having means thereon for mounting an access panel across the opening.

As the die itself has means thereon for mounting an access panel across the opening, the periphery of the opening can be left completely intact which is highly desirable.

According to a third aspect of the invention, there is provided a method of die cutting a composite laminate to produce an access opening comprising positioning a die on the laminate, applying a load to the die which is sufficient for the die to cut through the laminate and using the die to form a liner for the access opening, the die having means thereon for mounting an access panel across the opening.

Preferably, the method includes forming the die to define a recess into which a clamping member is received for securing the access panel in position. Preferably, the clamping member is a taper fit within the recess.

The method may include placing one surface of the composite laminate on a work surface such as the surface of a tool and positioning the die against the opposite surface of the composite laminate prior to applying pressure to the covering. In such a case, the method may include providing on the work surface a material which lies immediately beneath a cutter section of the die and advancing the cutting section through the composite laminate so that it comes to rest against the material. The material may be in the form of a ring.

According to a fourth aspect of the invention there is provided a die-cut composite laminate made using a method according to the first, second or third aspect of the invention or any one of the consistory clauses relating thereto.

According to a fifth aspect of the invention, there is provided an access opening and access panel arrangement comprising a liner in the opening formed with a recess and a clamping member which fits in the recess for securing the panel in position in the liner. Such an access panel arrangement can be utilised in a wing skin formed from laid-up carbon fibre fabric or from a wing skin made of metal.

Preferably, the clamping member is in the form of a ring which fits into a ring-like recess in the liner. The clamping member may be a taper fit in the liner. In use. the clamping member and the access panel are preferably arranged one each side of a radial wall section of the liner. Fastenings may be provided to extend through aligned openings in the clamping member and radial wall section for fastening the closure panel in position.

Preferably, the fastenings urge the clamping member into the recess to secure the panel.

According to a sixth aspect of the invention there is provided a method of die cutting a composite laminate, the method comprising positioning a die having a cutter section on a completely uncured or part-cured laminate, positioning inner and outer locator means on the laminate, arranging the cutter section to pass between the locator means, applying a covering to the die and laminate, applying a vacuum between the covering and the laminate, increasing gas pressure on the covering sufficient for the die to cut through the laminate, increasing the temperature of the laminate to complete curing thereof, and then removing the die from the laminate.

According to a seventh aspect of the invention there is provided a method of die cutting a wing skin made from a composite laminate, the method comprising positioning the wing skin on a tool, positioning inner and outer locator means on the wing skin, placing a die having a cutter section such that the cutter section is positioned between the locator means, applying a covering to the die and wing skin to form a complete assembly, applying a vacuum between the covering and the wing skin, increasing gas pressure on the covering sufficient for the die to cut through the wing skin, increasing the temperature of the wing skin to complete curing thereof, and then removing the die from the wing skin.

The method may include providing a location assembly such as a jig for positioning the die in the desired position relative to the wing skin.

According to an eighth aspect of the invention there is provided a method of die cutting a composite laminate, the method comprising positioning a die on an uncured or part-cured laminate, applying a load to the die which is sufficient to cut through the laminate, removing a section of laminate from the die to leave an access opening and then using the die as a liner or a surround for the periphery of the opening in the laminate.

The method may include mounting an access panel on the die. once the die forms the liner or surround.

According to a ninth aspect of the invention there is provided a composite laminate made by a method as set out in the immediately preceding paragraph, the access panel being secured to the die by fastener means which passes through a portion of the die clear of the composite laminate. The fastener means may pass through a radial section of the die. The fastener means may pass through a clamping member which clamps against the die, preferably the radial section thereof. The clamping member preferably locates in a recess defined between the die and the access panel and preferably lies flush with an adjacent outer surface of the die.

According to a tenth aspect of the invention there is provided a method of die cutting a composite laminate, the method comprising mounting inner and outer locator means on an uncured or part-cured laminate, positioning a die having a cutter section on the laminate, with the cutter section of the die .arranged between the locator means, placing a stripper plate between the die and the composite laminate, and increasing pressure on the die so that the cutter section suddenly breaks through the stripper plate to cut the laminate.

Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which:

Figure 1 is a diagrammatic cross section through a carbon fibre laminate with a die in position ready to cut the laminate;

Figure 2 is a view similar to Figure 1 showing the laminate penetrated by the die;

Figure 3 is a plan view of part of a tool on which is positioned a wing skin formed from a carbon fibre laminate and having a die in position thereon; Figure 4 is a cross section through the tool shown in Figure 3 on the line IV - IV in Figure 3;

Figure 5 is a cross section through the tool shown in Figure 3 on the line V - V in Figure 3;

Figure 6 is a diagrammatic cross section of a wing skin penetrated by a die;

Figure 7 shows the die as an insert within the wall of an aperture formed by the die in the wing-skin;

Figure 8 is an exploded view which shows the way in which a permanently inserted die can support an access panel;

Figure 9 is a view similar to Figure 8 but showing the various components to a smaller scale when assembled; and

Figure 10 shows the way in which the die may cooperate with a stripper plate.

Referring to Figure 1, a carbon fibre laminate 10 comprising a lay-up of carbon fibre fabrics impregnated with, for example, a suitable epoxy matrix has thereon inner and outer locators 12, 14 arranged to receive a cutting die 16.

It will be noted that the cutting die 16 is of T-shaped cross section having a vertical limb 18 formed as a knife edge 20 at its lower end and a horizontal limb 22 at its upper end. The knife edge 20 defines a cutting section of the die 16. The assembly of laminate 10, locators 12, 14 and die 16 is covered by an impermeable bag 24 and vacuum is applied to a space 26 beneath the bag 24 by withdrawing air through a passageway indicated diagrammatically at 25. At this point, the carbon fibre laminate 10 is in an uncured condition or only partially cured and the assembly 10, 12, 14, 16, 24 is then placed in an autoclave. The air pressure in the autoclave is increased along with the air temperature and differential pressure on the vacuum bag 24 in particular over the top of the horizontal limb 22 causes the die 16 to penetrate the laminate completely as shown in Figure 2. Simultaneously, heat in the autoclave causes the laminate 10 to cure with the die in the Figure 2 position. Once the laminate 10 has cured the assembly 10, 12, 14, 16, 24 can be removed from the autoclave. The die 16 can then be withdrawn and the supports 12, 14 removed. In that way the die 16 cuts a net-size opening in the laminate 10 thereby avoiding the need to cut the opening to net size after curing. To assist withdrawal a release agent may be applied to the die 16.

Alternatively, the die 16 may be constructed so that the effects of temperature are minimised, for example, through the use of low thermal coefficient of expansion materials.

Referring now to Figures 3 to 5, a wing skin 28 made from a carbon fibre laminate 10, as described already with reference to Fig 1, is to be formed with an access opening. The skin 28 is placed on a tool 30 and a location assembly 32 is positioned on the tool 30. The location assembly comprises two cross members 34 interconnected by a rectangular frame 36. Each of the cross members 34, has two side members 38 which bolt to the side of the tool 30. The cross members 34 are bolted to an outer locator 14 which defines an inner surface having straight sides 40 and semicircular ends 42. A die 16 of complementary shape is then placed inside the outer locator 14 along with an inner locator 12 of complementary shape to the inside surface of the die 16. The location assembly 32 is then unbolted from the tool 30 and from the outer locator 14 and the assembly is then removed. It will be noted from Figure 5 that the die 16 has a radial flange 46 at its upper end which is spaced from the upper surface 47 of the inner locator 12. It will be noted that a ring of material 45 such as a soft aluminium alloy is set in the upper surface of the tool 30 immediately beneath the knife edge of the die 16. A vacuum bag 24 is then placed over the tool as shown in broken lines in Figure 5 and vacuum is applied to the space 26 between the bag 24 and the tool 30. The assembly is then placed in an autoclave as before and pressure is applied as described with respect to Figures 1 and 2. The difference in pressure then causes the die 16 to penetrate the skin 10, the knife edge 20 of the die coming to rest against the ring of material 45, the material 45 helping to preserve the knife edge 20.

The assembly is then removed from the autoclave and the die 16 and inner locator 12 are removed along with the cut piece of carbon fibre laminate inside the die to leave an access opening. It will be appreciated that the access opening is accurately formed in the autoclave and the presence of the die in the laminate during the curing process minimises the risk of dimensional changes causing the access opening to be inaccurately formed.

As mentioned above, the method in accordance with the invention avoids the need to machine away cured laminate to bring the access opening to the net size which leads to a significant saving of time and cost in the case of a full size wing where several access openings may be required.

In summary, the preferred operational sequence is as follows:

1. Apply wing skin 28 to the tool 30;

2. Position the location assembly 32 along with the die 16 and the inner and outer locators 12,14. Remove location assembly 32 and apply vacuum bag 24 to form a complete assembly;

3. Load the assembly into the autoclave;

4. Run a programmed cycle to compact and heat the assembly in the autoclave such that the die 16 will penetrate the laminate 10 and the laminate will cure;

5. Remove the assembly from the autoclave.

Referring now to Figures 6 and 7, the die in Figure 6 is shown penetrating an aircraft wing-skin 28 fully. Once the die 16 has reached the Figure 6 position a section 28a can be removed from the interior of the die 16 to leave an access opening 29. The die 16 forms a liner for the opening 29 and is left in place permanently. As an alternative, the die 16 can be removed and a layer of adhesive applied thereto over surfaces 16a, 16b. The die can then be reinserted with the adhesive providing extra security against the die 16 becoming detached from the wing-skin 28.

Figure 8 shows the way in which a permanently inserted die can be used to support an access panel for an aircraft wing skin. Looking at Figure 8. a die 16 is provided and may be of similar shape in plan to the die 16 shown in Figure 3. The die 16 is generally of T-shaped cross-section having a vertical limb 18 and a horizontal upper limb 22. The horizontal limb 22 has a tapered radially outer section 48 and a radially inner section 50 formed with a number of circumterentially spaced-apart holes 52. The vertical limb 18 has a frusto-conical radially inner surface 54 and. as shown in broken lines, the limb I S is initially formed with a knife edge 20 at its lower end. The die 16 is made to penetrate the carbon fibre wing-skin 28 in the manner described above with the knife edge 20 projecting beyond an outer surface 28a of the wing skin 28. The die 16 is then removed and the knife edge 20 is cut away so as to leave the vertical limb 18 with a flat lower edge 56 flush with the outer surface 28a of the wing skin 28 when the die 16 is re-inserted. An access panel 58 has a peripheral mounting flange 60 formed with a plurality of peripherally spaced apart holes 62 which align with the respective holes 52. The mounting flange 60 houses a seal 64 for making sealing contact with the upper surface of the horizontal limb 22 when the Figure 8 components are assembled as shown in Figure 9. The access panel 58 has an outer surface 66 and a peripheral surface 68 which faces the frusto-conical surface 54. The frusto-conical surface 54 and the radially inner section 50 of the die 16 and the peripheral surface 68 of the access panel 58 define between them a recess 70 for receiving a clamping member 72. The clamping member 72 has a shape in plan similar to that of the recess 70 and has a frusto-conical outer peripheral surface 74 which co-operates with the frusto-conical surface 54 of the die 16. The clamping member 72 also has a vertical inner surface 76 of the access panel 58. The clamping member 76 has a plurality of peripherally spaced apart holes 78 therein which align with respective holes 62, 52 and which receive fixing screws 80.

The components of Figure 8 are assembled as shown in Figure 9. Initially, the access panel 55 is moved into position through the opening defined by the die 16 so as to lie above the hoπzontal limb member 22 of the die 16. The clamping member 76 is then placed in position in the recess 70 and the fixing screws 80 are passed through the respective holes 78. 52 and 62 and are screwed into respective anchor nuts 61 on the mounting flange 60 of the access panel 58. The tightening of the fixing screws 80 pull the flange 60 of the access panel 58 downwards into contact with the horizontal limb 22 of the die 16 so that the seal 64 makes sealing contact therewith. Also, the tightening of fixing screws 80 causes the frusto-conical surface 74 of the clamping member 72 to wedge against the frusto-conical surface 54 of the die 16. Therefore, there is a taper-fit between the clamping member 72 and the die 16 whereby the die 16 and clamping member 72 grip each other firmly. In that way the access panel 58 is held securely in its closed position. The surface 76 of the clamping member passes over the surface 68 of the access panel 58 with slight clearance.

In the Figure 9 position, the outer surface 66 of the access panel 58 and outer surface 79 of the clamping member 76 are generally flush with the surface 56 of the die 16. Typical!}', the access panel 58 will be an under-wing access panel for a fuel bay.

If desired an access opening can be formed in a metal wing-skin. e.g. by machining, and a liner similar in shape to the die 16 in Figure 8 inserted to support the access panel 58. In that way, fixing of the access panel is simplified and meets the requirement that fasteners must not penetrate the skin around the access opening. The liner can be secured in place in the opening using a suitable adhesive.

As well as using the die 16 or a member of similar shape thereto to form a liner for an opening, it could form a peripheral surround for a section of material cut from a panel. However, particular advantages are obtained when using the die or member of similar shape to form a liner especially from the point of view of mounting under-wing access panels.

Looking at Figure 10, a stripper plate 82 may be placed beneath the die 16 and the inner and outer locators 12, 14. The stripper plate 82 is formed with an area of weakness 84 which receives the knife edge 20 of the die 16. As pressure builds up on the die 16 in the autoclave, the knife edge 20 is urged with increasing force against the area of weakness 84. Eventually the load causes the knife edge 20 to break through the stripper plate 82 and impact the laminate 10 so as to cut through the laminate suddenly. The use of the stripper plate is particularly advantageous in that the gradual build up load on the knife edge 20 will not cause local deformation of the carbon fibres immediately beneath the knife edge prior to cutting. The stripper plate 82 can be designed such that it breaks when the die 16 is at a specific temperature, the autoclave applying pressure at that temperature. As the die 16 cools it will shrink away from the cut surface of the laminate 10 thereby assisting removal of the die.

Claims

1. A method of die cutting a composite laminate (10), the method being characterised by positioning a die (16) on the composite laminate (10) with the laminate in an uncured or pan-cured condition, applying a covering (24) to the die (16) and laminate (10) and applying fluid under pressure to the covering (24) to load the die (16) sufficiently for it to cut through the laminate (10).

2. A method according to claim 1. characterised in that the method comprises providing a die (16) having a ring-like cutting section.

3. A method according to claim 1 or 2. characterised by using the die (16) to form a liner or a surround for the periphery of the cut formed by the die (16).

4. A method of die cutting a composite laminate, the method comprising positioning a die (16) on the composite laminate (10), applying a load to the die (16) which is sufficient for the die to cut through the laminate (10), characterised by using the die (16) to form a liner or a surround for the periphery of the cut formed by the die.

5. A method according to claim 2 or 3 or 4, characterised by leaving the die (16) in position in or around the periphery of the cut after cutting to form the liner or surround.

6. A method according to any one of claims 2 to 5, characterised by using an adhesive to secure the die (16) in position in or around the periphery of the cut.

7. A method according to any one preceding claim, characterised by providing locator means (12, 14) for positioning the die (16) on the laminate (10).

8. A method according to claim 7, characterised in that the step of providing locator means comprises providing an inner and/or outer section (12, 14) arranged one each side of a cutter section (20) of the die (16).

9. A method according to claim 7 or 8, characterised by providing positioning means (32. 34. 36, 38) for positioning one of the locator means (12, 14) relative to the laminate (10).

10. A method according to any one preceding claim, characterised by applying a vacuum beneath the covering (24) whereby the load applied to the covering (24) is created as a result of differential pressure on the covering.

11. A method according to any one preceding claim, characterised by placing the covered laminate (10) and a tool (30) in an autoclave to provide a pressurised atmosphere.

12. A method according to any one preceding claim, characterised by placing a stripper plate (82) between the die (16) and the laminate (10) to be cut and loading the die (16) sufficiently to break through the stripper plate (82) and impact the laminate (10) and cut through the laminate suddenly.

13. A method according to any one preceding claim, characterised by placing the covered laminate (10) and tool (30) in an autoclave to provide heat to complete curing of the laminate (10).

14. A method according to any preceding claim, characterised by placing one surface of the composite laminate on a support such as the surface of a tool (30) and positioning the die (16) against the opposite surface of the composite laminate prior to applying pressure to the covering (24).

15. A method according to claim 14 characterised by providing on the support a material (45) which lies immediately beneath a cutter section (20) of the die (16) and advancing the cutting section (20) through the composite laminate (10) so that it comes to rest against the material (45).

16. A method according to claim 15, characterised in that the material (45) is in the form of a ring.

17. A die-cut composite laminate (10) such as an aircraft skin made using a method as claimed in any preceding claim.

18. A method according to any of claims 1 to 16, characterised by using the die (16) to form an access opening (29) and using the die (16) to form a liner for the access opening, the die (16) having means (60) thereon for mounting an access panel (58) across the opening (29).

19. A method of die cutting a composite laminate to produce an access opening (29), comprising positioning a die (16) on the laminate (10), applying a load to the die (16) which is sufficient for the die to cut through the laminate (10) characterised by using the die (16) to form a liner for the access opening (29), the die having means thereon for mounting an access panel (58) across the opening.

20. A method according to claim 19, characterised by forming the die (16) to define a recess (70) and into which a clamping member (72) for securing the access panel (58) in position.

21. A method according to claim 20, characterised by forming at least one of the recess (70) and clamping member (72) with a tapered surface (74) such that the clamping member is a taper fit within the recess (70).

22. .An access opening and access panel aπangement made by a method according to any one of claims 18 to 21.

23. An access opening and access panel arrangement, characterised in that a liner ( 16) is provided in the opening, the liner (16) is formed with a recess (70) and a clamping member (72) fits in the recess when securing the panel (58) in position in the liner.

24. An access opening and access panel arrangement according to claim 23, characterised in that the liner (16) is in the form a ring.

25. An access opening and access panel arrangement according to claim 23 or 24, characterised in that the clamping member (72) is in the form of a ring which fits into the recess (70) of the liner (16).

26. An access opening and access panel arrangement according to claim 23, 24 or 25, characterised in that the clamping member (72) is a taper fit in the liner (16).

27. An access opening and access panel arrangement according to claim 26, characterised in that the clamping member (72) has a tapered surface (74) which co-operates with a surface (54) defining the recess (70) in the liner (16).

28. An access opening and access panel aπangement according to claim 26 or 27, characterised in that the liner (16) has a tapered surface (54) which co-operates with a surface (74) of the clamping member (72).

29. An access opening and access panel arrangement according to any one of claims 23 to 28, characterised in that the clamping member (72) and the access panel (58) are, in use arranged one each side of a radial wall section (50) of the liner (16).

30. An access opening and access panel arrangement according to claim 29, characterised in that fastener means (80) is provided to extend through aligned openings (52, 62, 78) in the clamping member (72) and radial wall (50) for fastening the access panel (58) in position.

31. An access opening and access panel arrangement according to claim 30, characterised in that the fastener means (80) urges the clamping member (72) into the recess (70) to secure the panel (58).

32. An access opening and access panel aπangement according to any one of claims 19 to 28, formed on part of an aircraft such as an aircraft wing.