WO2017081456A1

WO2017081456A1 - Methods and patches for repairing composite laminates

Info

Publication number: WO2017081456A1
Application number: PCT/GB2016/053495
Authority: WO
Inventors: John Alexander ARMSTRONG; Samuel Colin HANNA; Steven David ROBB; Timothy BLOOMER
Original assignee: Short Brothers Plc
Priority date: 2015-11-11
Filing date: 2016-11-09
Publication date: 2017-05-18
Also published as: GB201519877D0

Abstract

Patches (20) and associated methods for repairing composite laminates (10) are disclosed where the composite laminate comprises a plurality of laminate plies (18) oriented according to a stacking sequence of the composite laminate and the patch comprises a plurality of repair plies (22) configured to at least partially fill a cavity (12) formed in the composite laminate and where the repair plies of the patch are oriented according to a stacking sequence different from the stacking sequence of the composite laminate.

Description

METHODS AND PATCHES FOR REPAIRING COMPOSITE LAMINATES

CROSS-REFERENCE TO RELATED APPLICATION(S)

[0000] This patent application claims priority from UK patent application no.

1519877.3 filed 1 1th November 2015.

TECHNICAL FIELD

[0001] The disclosure relates generally to composite materials, and more particularly to repairing composite laminates.

BACKGROUND OF THE ART

[0002] Composite materials are becoming more prevalent in aircraft structures. When structural damage occurs to a composite laminate or a localized area of such composite laminate is otherwise determined not to conform to design specifications, such composite laminate can often be repaired depending on the nature and extent of the damage or irregularity. The repair process involves removing one or more layers (i.e., plies) from the composite laminate in the area to be repaired to form a cavity in the composite laminate and the cavity is then filled by a patch (e.g., plug) comprising a stack of replacement plies. The side walls of the cavity can be tapered to facilitate bonding between the patch and the composite laminate. In traditional repair methods, the plies of the composite patch are selected, stacked and individually oriented to match respective corresponding plies in the composite laminate. Accordingly, the installation of such patch requires knowledge of the number and type of plies in the composite laminate and of the stacking sequence of the plies. As a result, traditional repair methods can be labour intensive and expensive.

SUMMARY

[0003] In one aspect, the disclosure describes a method for repairing a composite laminate comprising an assembly of a plurality of laminate plies comprising fibrous material where each laminate ply is oriented in accordance with a stacking sequence of the composite laminate. The method comprises:

at least partially filling a cavity formed in the composite laminate with a stack of repair plies comprising fibrous material and forming a patch where each repair ply is oriented in accordance with a stacking sequence of the patch, the stacking sequence of the patch being different from the stacking sequence of the composite laminate;

bonding the fibrous material of the repair plies; and

bonding the repair plies to the composite laminate.

[0004] The repair plies may comprise one or more lower plies comprising unidirectional fibrous material.

[0005] The stacking sequence of the patch may comprise a repeating pattern of ply orientations where one of the orientations is more common than the other orientations in the pattern to cause anisotropy in the mechanical properties of the patch. The more common ply orientation of the stacking sequence of the patch may correspond to a reference orientation on the composite laminate and also correspond to an orientation along which the patch has a relatively higher stiffness.

[0006] The stacking sequence of the patch may comprise a repeating pattern of ply orientations for the lower plies comprising the following ply orientations: 45 degrees, 135 degrees, 0 degrees, 90 degrees and 0 degrees.

[0007] The repair plies may be pre-impregnated with matrix material.

[0008] The repair plies may comprise one or more upper plies comprising a woven fibrous material.

[0009] The one or more upper plies may overlap onto an outer surface of the composite laminate.

[0010] The repair plies may comprise one or more upper plies comprising unidirectional fibrous material.

[0011] The cavity in the composite laminate may have a depth extending through a number of laminate plies, and, a number of repair plies in the patch filling the cavity in the composite laminate may differ from the number of laminate plies.

[0012] Bonding the fibrous material of the repair plies may comprise co-curing the patch with the composite laminate.

[0013] In another aspect, the disclosure describes a repaired composite laminate comprising: an assembly of a plurality of laminate plies where each laminate ply is oriented in accordance with a stacking sequence of the composite laminate; and a patch at least partially filling a cavity formed in the laminate plies and being bonded to the laminate plies, the patch comprising:

a stack of plurality of repair plies comprising fibrous material where each repair ply is oriented in accordance with a stacking sequence of the patch, the stacking sequence of the patch being different from the stacking sequence of the composite laminate; and

a matrix material bonding the fibrous material of the repair plies. [0014] The repair plies may comprise one or more lower plies comprising unidirectional fibrous material.

[0015] The stacking sequence of the patch may comprise a repeating pattern of ply orientations where one of the orientations is more common than the other orientations in the pattern to cause anisotropy in the mechanical properties of the patch. The more common ply orientation of the stacking sequence of the patch may correspond to a reference orientation on the composite laminate and also correspond to an orientation along which the patch has a relatively higher stiffness.

[0016] The stacking sequence of the patch may comprise a repeating pattern of ply orientations for the lower plies comprising the following ply orientations: 45 degrees, 135 degrees, 0 degrees, 90 degrees and 0 degrees.

[0017] The repair plies may be pre-impregnated with the matrix material.

[0018] The repaired composite laminate may comprise one or more upper plies comprising a woven fibrous material.

[0019] The one or more upper plies may overlap onto an outer surface of the composite laminate.

[0020] The patch may have an outer surface that is substantially flush with an outer surface of the composite laminate.

[0021] The repair plies may comprise one or more upper plies comprising unidirectional fibrous material. [0022] The cavity in the laminate plies may have a depth extending through a number of laminate plies, and, a number of repair plies in the patch filling the cavity in the composite laminate may differ from the number of laminate plies.

[0023] In another aspect, the disclosure describes an aircraft comprising a repaired composite laminate as described herein.

[0024] In another aspect, the disclosure describes a patch for repairing a composite laminate where the composite laminate comprises an assembly of a plurality of laminate plies and where each laminate ply is oriented in accordance with a stacking sequence of the composite laminate. The patch comprises:

a stack of plurality of repair plies configured to at least partially fill a cavity formed in the composite laminate, each repair ply comprising fibrous material and being oriented in accordance with a stacking sequence of the patch, the stacking sequence of the patch being different from the stacking sequence of the composite laminate; and

a matrix material for bonding the fibrous material of the repair plies.

[0025] The repair plies may comprise a one or more lower plies comprising unidirectional fibrous material.

[0026] The stacking sequence of the patch may comprise a repeating pattern of ply orientations where one of the orientations is more common than the other orientations in the pattern to cause anisotropy in the mechanical properties of the patch.

[0027] The stacking sequence of the patch may comprise a repeating pattern of ply orientations for the lower plies comprising the following ply orientations: 45 degrees, 135 degrees, 0 degrees, 90 degrees and 0 degrees.

[0028] The repair plies may be pre-impregnated with the matrix material.

[0029] The repair plies may comprise one or more upper plies comprising a woven fibrous material.

[0030] The one or more upper plies may be configured to overlap onto an outer surface of the composite laminate.

[0031] The repair plies may comprise one or more upper plies comprising unidirectional fibrous material. [0032] In a further aspect, the disclosure describes an aircraft component such as a wing comprising a patch as described herein.

[0033] Further details of these and other aspects of the subject matter of this application will be apparent from the detailed description and drawings included below.

DESCRIPTION OF THE DRAWINGS

[0034] Reference is now made to the accompanying drawings, in which:

[0035] FIG. 1 shows an cross-sectional view of a partial composite laminate in which a cavity has been formed to remove a portion of the composite laminate in need of repair;

[0036] FIG. 2A schematically shows an exemplary cavity profile configured to provide a scarf interface with a repair patch;

[0037] FIG. 2B schematically shows an exemplary cavity profile configured to provide a stepped interface with a repair patch;

[0038] FIG. 3 schematically shows a cross-sectional view of an exemplary repair patch filling a cavity in a composite laminate;

[0039] FIGS. 4A and 4B respectively show first and second portions of a table defining an exemplary stacking sequence for the repair plies of the repair patch of FIG. 3;

[0040] FIG. 5 schematically shows a cross-sectional view of another exemplary repair patch filling a cavity in a composite laminate;

[0041] FIGS. 6A and 6B respectively show first and second portions of a table defining an exemplary stacking sequence for the repair plies of the repair patch of FIG. 5;

[0042] FIG. 7 is a flowchart illustrating an exemplary method for repairing a composite laminate using the repair patch of FIG. 3 or the repair patch of FIG. 5; and

[0043] FIG. 8 schematically illustrates an exemplary setup for curing the repair patch of FIG. 3 or the repair patch of FIG. 5. DETAILED DESCRIPTION

[0044] The present disclosure relates to repairing structures comprising composite materials including, for example, fiber-reinforced composite materials and advanced composite materials also known as advanced polymer matrix composites which generally comprise high strength or intermediate modulus fibers bound together by a matrix material (e.g., polymer) or any known or other composite material(s) suitable for use in aircraft structural parts such as, for example, fuselage skins, (e.g., upper and lower) wing skins and fuel tanks. It is understood that aspects of this disclosure may be equally applicable to other composite structures suitable for other applications such as transport (e.g., trains, busses, ships, watercraft) and automotive for example. Such composite materials may, for example, include fiber reinforcement materials such as carbon, aramid and/or glass fibers embedded into a thermosetting or thermoplastic matrix material. In various embodiments, the present disclosure discloses methods and repair patches (e.g., plugs) for repairing composite laminates comprising assemblies of layers (i.e., plies) of fibrous composite materials which are joined to provide desired properties such as in-plane stiffness, bending stiffness and strength.

[0045] In various embodiments, the repair patches disclosed herein may be tailored for structural performance, aerodynamic performance and/or for appearance (i.e., cosmetics) of the repaired composite laminate. For example, as described below, repair patches of the present disclosure may comprise repair plies stacked according to a stacking sequence that is different from the stacking sequence of the composite laminate but yet that provide adequate structural performance for some applications. Alternatively or in addition repair patches of the present disclosure may comprise repair plies that are stacked and sized to substantially restore the original thickness and/or match the curvature of an outer surface of the composite laminate to provide adequate aerodynamic performance and/or appearance of the repaired composite laminate. In some applications, knowledge of the stacking sequence of the plies of the composite laminate to be repaired may not be required. In some situations, the use of repair patches of the present disclosure for repairing composite laminates may require relatively simplified and/or cost-reducing manufacturing methods in comparison with those required for traditional repair procedures.

[0046] Aspects of various embodiments are described through reference to the drawings. [0047] FIG. 1 shows a cross-sectional view of a partial composite laminate 10 in which cavity 12 has been formed to remove a portion of composite laminate 10 in need of repair. For example, cavity 12 may have been formed to remove a portion of composite laminate 10 damaged during operation or to remove some irregularity(ies) in composite laminate 10 that may have been introduced during manufacturing of composite laminate 10 and that may have been detected by, for example, known or other suitable non-destructive inspection methods. Accordingly, such damage or irregularity in need of repair may have been at or below outer surface 14 of composite laminate 10.

[0048] Composite laminate 10 may, as explained above, be part of an aircraft such as a fuselage skin, upper or lower wing skin, fuel tank, or, other composite structure. Composite laminate 10 may have previously been manufactured using known or other type of resin transfer infusion process including, for example, resin transfer infusion (RTI), or, using some other composite manufacturing process. Composite laminate 10 may comprise outer surface 14, which may also be referred to as an outer mold line, and inner surface 16, which may also be referred to as an inner mold line. Cavity 12 may have any suitable shape and size based on the amount and location of material of composite laminate 10 to be removed from composite laminate 10 in order to remove the damaged/irregular portion(s) of composite laminate 10. In some embodiments, the shape of cavity 12 may be at least partially standardized in order to facilitate the repair procedure. For example, as shown in FIG. 1 , cavity 12 may have a shape defined by a profile (e.g., lines 12A, 12B) that has been revolved about axis CL. In some embodiments, cavity 12 may have the shape of a truncated cone having a substantially circular cross-section and extending into composite laminate. The removal of material from composite laminate 10 to form cavity 12 may be conducted using any suitable material removal process including machining (e.g., milling, grinding, abrading, drilling, laser ablation, water jet cutting) either in a manual or automated manner. Tools of known or other types suitable for machining composite materials may be used to form cavity 12. For example, in some situations, tools of the LESLIE series sold under the trade name GMI-AERO may be suitable for forming cavity 12.

[0049] In reference to FIG. 1 , the region of composite laminate 10 to be repaired may have: a bottom radius A of cavity 12; a scarf length B in a scarfed, stepped or tapered portion of cavity 12; a repair ply overlap length C representing a distance overlapping outer surface 14; an electromagnetic compatibility (EMC) overlap length D representing an overlap distance between an existing EMC foil of composite laminate 10 (if applicable) and an EMC foil of the repair patch used to fill cavity 12; a total (overall) repair radius E and a cavity depth F.

[0050] FIG. 2A schematically shows an exemplary profile of cavity 12 in composite laminate 10 configured to provide a scarf joint (interface) with a repair patch (see FIGS. 3 and 5). As shown in FIG. 1 , cavity 12 may generally have the shape of a truncated cone extending into composite laminate 10. Alternatively, cavity 12 may have any other suitable symmetrical or non-symmetrical shape. Cavity 12 may have substantially smooth (e.g., non-stepped) side walls 12A for interfacing with the repair patch. Cavity 12 may have a depth F that extends through one or more plies 18 of composite laminate 10 where laminate plies 18 comprise fibrous composite material (e.g., carbon fibers). Laminate plies 18 may comprise unidirectional or woven fibrous material where the orientation of the fibers in each laminate ply 18 is in accordance with a stacking sequence of the composite laminate 10. In some embodiments, it may be desirable to abrade side wall(s) 12A to remove surface irregularities.

[0051] FIG. 2B schematically shows an exemplary profile of cavity 12 in composite laminate 10 configured to provide lap joints (stepped interface) between laminate plies 18 and repair plies of a repair patch (see FIGS. 3 and 5). Accordingly, cavity 12 may have stepped side walls 12A for interfacing with the repair patch. For example, the steps in side walls 12A may be configured to provide lap joints between one or more laminate plies 18 and corresponding one or more repair plies of the repair patch. In some embodiments, the number of laminate plies 18 and the number of repair plies in the repair patch may correspond in a one-to-one relationship for at least some of the depth F of cavity 12. However, in some embodiments, the number of laminate plies 18 and the number of repair plies in the repair patch may not necessarily correspond in a one-to-one relationship. In some embodiments, each step in side walls 12A may have a height/thickness equivalent to that of one laminate ply 18. In some embodiments where laminate plies 18 are made up of two distinct layers (e.g., biaxial plies), it may be desirable to have steps of a height/thickness equivalent to that of one distinct layer of such multi-layered laminate plies 18. Accordingly, it may be desirable in some situations to partially abrade multi-layered laminate plies 18 to obtain steps of heights/thicknesses that are less than that of one laminate ply 18. [0052] FIG. 3 schematically shows a cross-sectional view of an exemplary repair patch 20 filling cavity 12 in composite laminate 10. Even though outer surface 14 of composite laminate 10 is shown as being substantially planar, it is understood that patch 20 may be used to repair a region of composite laminate 10 having a curved (e.g., including doubly curved) outer surface 14. Patch 20 may be used for repairing composite laminate 10 comprising an assembly of a plurality of laminate plies 18 comprising fibrous material where each laminate ply 18 is oriented in accordance with a stacking sequence of composite laminate 10. Patch 20 may comprise one or more (e.g., a stack) of repair plies 22 (e.g., 22A, 22B) comprising fibrous material where each repair ply 22 is oriented in accordance with a stacking sequence of patch 20. The stacking sequence of patch 20 may be predetermined irrespective of the stacking sequence or composite laminate 10. Accordingly, the stacking sequence of repair plies 22 in patch 20 may be different from the stacking sequence of laminate plies 18 of composite laminate 10. For example, the orientation of the fibers in individual repair plies 22 of patch 20 may not be matched to the orientation of the fibers in laminate plies 18 of composite laminate 10. Accordingly, prior knowledge of the stacking sequence of composite laminate 10 may not be required prior to conducting the repair of composite laminate 10 using patch 20.

[0053] Also, in some embodiments, the number of laminate plies 18 exposed to cavity 12 (see FIGS. 2A and 2B) and the number of repair plies 22 in repair patch 20 may correspond in a one-to-one relationship for at least some of the depth F of cavity 12. Alternatively, in some embodiments, the number of laminate plies 18 exposed to cavity 12 and the number of repair plies 22 in repair patch 20 may not necessarily correspond in a one-to-one relationship. Accordingly, prior knowledge of the number of laminate plies 18 along the depth F (see FIG. 1) of cavity 12 may not be required prior to conducting the repair of composite laminate 10 using patch 20. In some embodiments, patch 20 may comprise a metallic foil or other conductor configured to be in electrical contact with a metallic foil or conductor of composite laminate 10 in order to provide at least partial electrical continuity of the repaired composite laminate 10 across patch 20 to provide some electromagnetic shielding.

[0054] The number and type(s) of repair plies 22 in patch 20 and also the stacking sequence defining the orientation of repair plies 22 in patch 20 may be predetermined based on the anticipated use of the repaired composite laminate 10 so as to provide acceptable mechanical and other properties of the repaired composite laminate 10. For example, patch 20 may be designed to provide a desired stiffness that is compatible for use with the intended composite laminate 10 so that the use of patch 20 may not result in undesirable stress concentrations being formed in composite laminate 10. In some embodiments, patch 20 may be designed to have a stiffness within an acceptable range of the stiffness of composite laminate 10. Such design may be achieved via selection and orientation of repair plies 22.

[0055] In some embodiments, the number and type(s) of repair plies 22 in patch 20 as well as the sequence of repair plies 22 may be selected based on the position of patch on composite laminate 10. For example, in case of composite laminate 10 being a part of an aircraft wing skin, the number, type(s) and sequence of repair plies 22 in patch 20 may depend on whether patch 20 will be applied between ribs of the wing structure or overlap a rib of the wing structure for example.

[0056] Patch 20 may comprise a matrix material for bonding the fibrous material of repair plies 22. For example, one or more of repair plies 22 may comprise a reinforcing fabric which has been pre-impregnated (pre-preg) with a resin system (e.g., epoxy) already including the proper curing agent. As a result, the prepreg repair plies 22 may be laid into cavity 12 and cured without the addition of any more resin or other matrix material. In some embodiments, repair plies 22 may be of the type known as M20 PREPREG sold under the trade name HEXPLY®. One or more film adhesive layers 23 may be disposed between patch 20 and composite laminate 10 in order to bond patch 20 to composite laminate 10. Alternatively, repair plies 22 may be of the self-adhesive pre-preg type in which case film adhesive layers 23 may not be required between patch 20 and composite laminate 10. In some embodiments, one or more other film adhesive layers 23 may be disposed on top of patch 20 to serve as a surfacing ply to improve surface finish of the repaired composite laminate 10.

[0057] In some embodiments, repair plies 22 of patch 20 may comprise one or more lower plies 22A and one or more upper plies 22B. Lower plies 22A may be used to partially or completely fill cavity 12 and may be disposed mainly at an elevation that is below outer surface 14 of composite laminate 10. In some embodiments, one or more of lower plies 22A may comprise unidirectional fibrous material. Upper plies 22B may be applied over lower plies 22A and may be disposed at an elevation that is at, below or above outer surface 14 of composite laminate 10 depending on the properties desired from patch 20. For example, in the case where structural performance is the principal design criteria for patch 20 as opposed to aerodynamic performance or cosmetics, it may be desirable to have upper plies 22B of patch 20 overlap outer surface 14 and be disposed at an elevation above outer surface 14 so that an outer surface 21 of patch 20 may form a protrusion extending from outer surface 14. In some embodiments, one or more of upper plies 22B may comprise unidirectional fibrous material. Alternatively or in addition, one or more of upper plies 22B may comprise woven (e.g., plain weave) fibrous material.

[0058] Upper plies 22B may be provided to compensate for the lack of high pressure compaction done during the repair process described below as opposed to during the original manufacturing of the original composite laminate 10. During repair, patch 20 is compacted using a vacuum source (i.e., debulking) as opposed to autoclave pressure which may have been used during the original manufacturing of composite laminate 10. In order to compensate for the reduced compaction pressure, additional upper plies 22B may be added to enhance structural performance of patch 20 of FIG. 3.

[0059] The use of woven fibrous material as one or more upper plies 22B may provide improved impact resistance in comparison with unidirectional fibrous material. Woven fibrous material may also provide more resistance to breakout or splintering during drilling. The number of upper plies 22B made of woven fibrous material may be selected based on the properties of patch 20 required for the specific application.

[0060] In some embodiments, patch 20 may be configured to have anisotropic properties depending on its intended application. For example, the stacking sequence of patch 20 may comprise a repeating pattern of ply orientations where one of the orientations is more common than the other orientations in the pattern to cause anisotropy in the mechanical properties of patch 20. For example, in the case of patch 20 being used to repair a wing of an aircraft, patch 20 may be configured to have a preferred orientation in which the stiffness of patch 20 may be higher. Such preferred orientation may be referred as the "0 degree" orientation and may, for example, correspond to a direction that is parallel to a stringer or spar that is part of the structure of the wing. It is understood that various different types, number and orientation stacking sequences of repair plies 22 may be used to achieve the desired properties of patch 20. The use of a standard repeating pattern of ply orientations may provide mechanical properties suitable for some repair situations and may eliminate the need for having to match the orientation of repair plies 22 to the plies of composite laminate 10, which may have a complex layup configuration. [0061] FIGS. 4A and 4B respectively show first and second portions of a table defining an exemplary stacking sequence for repair plies 22 of repair patch 20 that may be suitable and provide adequate properties for some applications. FIGS. 4A and 4B together form a table with values representing a stacking sequence of patch 22 of different cavity depth F and associated scarf lengths B. FIG. 4A contains values for cavity depths ranging from 0.005 inch to 0.07 inch and FIG. 4B contains values for cavity depths ranging from 0.075 inch to 0.14 inch. The top four values in each column of the table are shown as being shaded and represent the stacking sequence of upper plies 22B. The un-shaded values below the shaded values represent the stacking sequence of lower plies 22A. As shown in FIGS. 4A and 4B, the stacking sequence of patch 20 may comprise a repeating pattern of ply orientations and ply types for lower plies 22A and for upper plies 22B. For example, the stacking sequence of a patch 20 made in accordance with FIGS. 4A and 4B may comprise one or more plies comprising plain weave fibrous material having orientations of 0 degrees and 90 degrees disposed over plies of unidirectional fibrous material having a stacking sequence of orientations (e.g., I, II, III, IV and V) in a repeating pattern. In some embodiments, the orientations may comprise the following sequence of five exemplary ply orientations: 45 degrees, 135 degrees, 0 degrees, 90 degrees and 0 degrees in a downward direction into (i.e., from the top to the bottom of) cavity 12 as shown in FIGS. 2A and 2B. Accordingly, orientations III and V may be the same. Upper plies 22B may, in some embodiments, comprise two plies comprising unidirectional fibrous material and two plies comprising plain weave fibrous material. The stacking sequence of repair plies 22 of patch 20 of FIG. 3 is not intended to be limited to the examples shown herein.

[0062] FIG. 5 schematically shows a cross-sectional view of another exemplary repair patch 20 at least partially filling cavity 12 in composite laminate 10. The repair patch 20 comprises elements previously described above in relation to the repair patch 20 of FIG. 3 and therefor will not be repeated below. Like elements have been labelled using like reference numerals. The repair patch 20 of FIG. 5 may be tailored for aerodynamic performance and/or appearance instead of or in addition to structural performance. Accordingly, patch 20 may be configured to substantially restore the original thickness of composite laminate 10 and/or match a curvature of outer surface 14 so that the region of patch 20 may be substantially continuous with outer surface 14 and that no significant protrusion or depression in outer surface 14 may be formed as a result of patch 20. For example, in some embodiments, upper plies 22B may be sized and positioned so that outer surface 21 of patch 20 is substantially flush with outer surface 14 of composite laminate 10.

[0063] In various embodiments, outer surface 14 of composite laminate 10 may be substantially planar or may be (e.g., doubly) curved where such curvature of outer surface 14 is achieved by ply drop-offs for example. In some embodiments, plies 22 of patch 20 may be sized and positioned to comprise ply drop-offs in order to approximate the curvature of outer surface 14 in order to substantially restore the aerodynamic performance or appearance of outer surface 14 of composite laminate 10.

[0064] FIGS. 6A and 6B respectively show first and second portions of a table defining an exemplary stacking sequence for repair plies 22 of repair patch 20 of FIG. 5 that may be suitable and provide adequate properties for some applications. FIGS. 6A and 6B together form a table with values representing a stacking sequence of patch 22 of different cavity depth F and associated scarf lengths B. FIG. 6A contains values for cavity depths ranging from 0.005 inch to 0.07 inch and FIG. 6B contains values for cavity depths ranging from 0.075 inch to 0.14 inch. The top value in each column of the table is shown as being shaded and represents the configuration of upper ply(ies) 22B. It is understood that the patch 20 of FIG. 5 could comprise more than one upper ply 22B in some embodiments. The un-shaded values below the shaded values represent the stacking sequence of lower plies 22A. As shown in FIGS. 6A and 6B, the stacking sequence of patch 20 may comprise a repeating pattern of ply orientations (e.g., I, II, III, IV and V) and ply types for lower plies 22A and for upper plies 22B. For example, lower plies 22A may comprise unidirectional fibrous material and may comprise the following five exemplary ply orientations: 45 degree, 135 degrees, 0 degrees, 90 degrees and 0 degrees in a downward direction into (i.e., from the top to the bottom of) cavity 12 as shown in FIGS. 2A and 2B. For example, upper plies 22B may comprise, in some embodiments, a single ply comprising plain weave fibrous material having orientations of 0 degrees and 90 degrees. Plain weave upper ply 22B may be disposed over unidirectional upper plies 22B. The stacking sequence of repair plies 22 of patch 20 of FIG. 5 is not intended to be limited to the examples shown herein.

[0065] FIG. 7 is a flowchart illustrating a method 700 for repairing composite laminate 10 using the repair patch 20 of the types shown in FIGS. 3 and/or the repair patch 20 of the type shown in FIG. 5. As explained above, composite laminate 10 may comprise an assembly of a plurality of laminate plies 18 where each laminate ply 18 comprises fibrous material that is oriented in accordance with a stacking sequence of composite laminate 10. Method 700 may comprise: at least partially filling cavity 12 formed in composite laminate 10 with a stack of repair plies 22 comprising fibrous material and forming patch 20 where each repair ply 22 is oriented in accordance with a stacking sequence of patch 20 and the stacking sequence of patch 20 is different from the stacking sequence of composite laminate (see block 702); bonding the fibrous material of repair plies 22 (see block 704); and bonding repair plies 22 to composite laminate 10 (see block 706).

[0066] In various embodiments, bonding the fibrous material of the repair plies

22 of patch 20 may comprise co-curing patch 20 (e.g., in situ) with composite laminate 10, or, patch 20 may comprise a pre-cured patch 20 that is subsequently bonded to composite laminate 10. Patch 20 may be bonded to composite laminate 10 via one or more film adhesive layers 23 (see FIGS. 3 and 5).

[0067] As explained above, repair plies 22 may comprise one or more lower plies 22A comprising unidirectional fibrous material. Repair plies 22 may also comprise one or more upper plies 22B comprising a woven fibrous material. Alternatively or in addition, one or more upper plies 22B may comprise unidirectional fibrous material. One or more of upper plies 22B may overlap onto outer surface 14 of composite laminate 10. One or more repair plies 22 may be pre-impregnated with matrix material that may serve to bond the fibrous material of repair plies 22 together. One or more film adhesive layers 23 may be disposed between patch 20 and composite laminate 10 in order to bond patch 20 to composite laminate 10. In some embodiments, such film adhesive layers 23 may comprise product number FM300-2 sold by CYTEC INDUSTRIES.

[0068] In some embodiments of method 700, the stacking sequence of patch

20 may comprise a repeating pattern of ply orientations where one of the orientations may be more common than the other orientations in the pattern to cause anisotropy in the mechanical properties of patch 20. The more common ply orientation of the stacking sequence of patch 20 may correspond to a reference orientation on composite laminate 10 and also correspond to an orientation along which patch 20 has a relatively higher stiffness than along other orientations. In some embodiments, the stacking sequence of patch 20 comprises a repeating pattern of ply orientations for lower plies 22A comprising the following ply orientations: 45 degrees, 135 degrees, 0 degrees, 90 degrees and 0 degrees. Alternatively, patch 20 may comprise a repeating pattern of ply orientations where the same number of plies 22 are disposed at each orientation of the stacking sequence so that the stiffness of patch 20 may be substantially the same in each of the orientations in the pattern. The use of a standard repeating pattern of ply orientations may provide mechanical properties suitable for some repair situations and may eliminate the need for having to match the orientation of repair plies 22 of the plies of composite laminate 10, which may have a complex layup configuration. [0069] In some embodiments of method 700, the number of laminate plies 18 exposed to cavity 12 (see FIGS. 2A and 2B) and the number of repair plies 22 in repair patch 20 may correspond in a one-to-one relationship for at least some of the depth F of cavity 12. Alternatively, in some embodiments of method 700, the number of laminate plies 18 exposed to cavity 12 and the number of repair plies 22 in repair patch 20 may not necessarily correspond in a one-to-one relationship. Accordingly, prior knowledge of the number of laminate plies 18 along the depth F (see FIG. 1) of cavity 12 may not be required prior to conducting the repair of composite laminate 10 using patch 20. Accordingly, cavity 12 may have a depth extending through a number of laminate plies 18 and a number of repair plies 22 in patch 20 filling cavity 12 may differ from the number of laminate plies 18.

[0070] FIG. 8 schematically illustrates an apparatus 24 (setup) for curing repair patch 20 of FIG. 3 or repair patch 20 of FIG. 5. Apparatus 24 and associated procedure for curing patch 20 and bonding patch 20 to composite laminate 10 may be of any known or other type(s) suitable for use with curing of pre-preg composite materials. For example, repair patch 20 may be cured (e.g., co-cured with composite laminate 10) using a suitable autoclave or out-of-autoclave process. Accordingly, apparatus 24 may comprise vacuum bag 26 (barrier) substantially sealed to composite laminate 10 via a suitable adhesive such as double-sided sealant tape 28, which may be selected to withstand temperatures at which apparatus 24 is subjected to during use. Vacuum bag 26, sealant tape 28 and composite laminate 10 may cooperate to define a substantially enclosed volume 30 including patch 20 and to which vacuum source 32 may be connected in order to produce a vacuum condition inside volume 30 during curing. Apparatus 24 may comprise heater blanket 34 disposed between two release films 36 or other suitable heating means. Heater blanket 34 may be electrically powered and used to heat patch 20 in order to activate and cause curing of the matrix material part of the prepreg repair plies 22 of patch 20. The operation of heater blanket 34 may be controlled by a suitable temperature controller (not shown) using one or more thermocouples 35A and 35B to provide feedback signal(s) for the purpose of controlling a temperature to which patch 20 is exposed. Thermocouples 35A and 35B may be disposed near or at periphery of patch 20. Thermocouple 35B may be disposed at a distance from thermocouple 35A so that the temperatures at two different regions of volume 30 may be monitored.

[0071] In some embodiments, release films 36 may each comprise a cohesively formed plastic that does not readily adhere to other polymers or other type of known or other release medium. For example, solid release films 36 may be configured to not chemically bond to the patch 20 or to composite laminate 10 so that it may be easily removed by peeling after curing.

[0072] Apparatus 24 may comprise breather 38 of known or other type disposed in volume 30 between vacuum bag 26 and the upper release film 36 disposed above heater blanket 34. Breather 38 may provide passage space for gas/air drawn under vacuum from different regions of volume 30 toward vacuum source 32. The application of a vacuum using vacuum source 32 will tend to collapse vacuum bag 26 and pressurize repair plies 22 of patch 20. Accordingly, a debulking process may be carried out before curing (i.e., without heat) during the placement of repair plies 22 into cavity 12 and/or after all repair plies 22 have been put in place. Debulking may comprise removing air from prepreg repair plies 22 in order to increase the density of patch 20. Debulking may be conducted at intervals during the layup of patch 20.

[0073] Apparatus 24 may also comprise tow 40 that may be attached to composite laminate 10 via a suitable adhesive. In some embodiments, tow 40 may be secured to composite laminate 10 via tape 42. Such tape 42 may be green composite bonding tape of the type sold under the trade name FLASHBREAKER 1 by AIRTECH. Tow 40 may comprise an untwisted bundle of filaments made of glass or carbon. The relatively small thickness of tow 40 may not significantly interfere with heater blanket 34 or other components of apparatus 24 however it may permit the evacuation of air from patch 20 using vacuum source 32. [0074] The above description is meant to be exemplary only, and one skilled in the relevant arts will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. For example, the blocks and/or operations in the flowcharts and drawings described herein are for purposes of example only. There may be many variations to these blocks and/or operations without departing from the teachings of the present disclosure. For instance, the blocks may be performed in a differing order, or blocks may be added, deleted, or modified. The present disclosure may be embodied in other specific forms without departing from the subject matter of the claims. Also, one skilled in the relevant arts will appreciate that while the patches, repaired composite laminates and methods disclosed and shown herein may comprise a specific number of elements/components, the patches, repaired composite laminates and methods could be modified to include additional or fewer of such elements/components. The present disclosure is also intended to cover and embrace all suitable changes in technology. Modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims. Also, the scope of the claims should not be limited by the preferred embodiments set forth in the examples, but should be given the broadest interpretation consistent with the description as a whole.

Claims

WHAT IS CLAIMED IS:

1. A method for repairing a composite laminate comprising an assembly of a plurality of laminate plies comprising fibrous material where each laminate ply is oriented in accordance with a stacking sequence of the composite laminate, the method comprising:

bonding the fibrous material of the repair plies; and

bonding the repair plies to the composite laminate.

2. The method as defined in claim 1 , wherein the repair plies comprise one or more lower plies comprising unidirectional fibrous material.

3. The method as defined in any one of claims 1 and 2, wherein the stacking sequence of the patch comprises a repeating pattern of ply orientations where one of the orientations is more common than the other orientations in the pattern to cause anisotropy in the mechanical properties of the patch.

4. The method as defined in claim 3, wherein the more common ply orientation of the stacking sequence of the patch corresponds to a reference orientation on the composite laminate and also corresponds to an orientation along which the patch has a relatively higher stiffness.

5. The method as defined in claim 2, wherein the stacking sequence of the patch comprises a repeating pattern of ply orientations for the lower plies comprising the following ply orientations: 45 degrees, 135 degrees, 0 degrees, 90 degrees and 0 degrees.

6. The method as defined in any one of claims 1 to 5, wherein the repair plies are pre-impregnated with matrix material.

7. The method as defined in any one of claims 1 to 8, wherein the repair plies comprise one or more upper plies comprising a woven fibrous material.

8. The method as defined in claim 7, wherein the one or more upper plies overlap onto an outer surface of the composite laminate.

9. The method as defined in any one of claims 1 to 8, wherein the repair plies comprise one or more upper plies comprising unidirectional fibrous material.

10. The method as defined in any one of claims 1 to 9, wherein:

the cavity in the composite laminate has a depth extending through a number of laminate plies; and

a number of repair plies in the patch filling the cavity in the composite laminate differs from the number of laminate plies.

1 1. The method as defined in any one of claims 1 to 10, wherein bonding the fibrous material of the repair plies comprises co-curing the patch with the composite laminate.

12. A repaired composite laminate comprising:

an assembly of a plurality of laminate plies where each laminate ply is oriented in accordance with a stacking sequence of the composite laminate; and a patch at least partially filling a cavity formed in the laminate plies and being bonded to the laminate plies, the patch comprising:

a matrix material bonding the fibrous material of the repair plies.

13. The repaired composite laminate as defined in claim 12, wherein the repair plies comprise one or more lower plies comprising unidirectional fibrous material.

14. The repaired composite laminate as defined in any one of claims 12 and 13, wherein the stacking sequence of the patch comprises a repeating pattern of ply orientations where one of the orientations is more common than the other orientations in the pattern to cause anisotropy in the mechanical properties of the patch.

15. The repaired composite laminate as defined in claim 14, wherein the more common ply orientation of the stacking sequence of the patch corresponds to a reference orientation on the composite laminate and also corresponds to an orientation along which the patch has a relatively higher stiffness.

16. The repaired composite laminate as defined in claim 13, wherein the stacking sequence of the patch comprise a repeating pattern of ply orientations for the lower plies comprising the following ply orientations: 45 degrees, 135 degrees, 0 degrees, 90 degrees and 0 degrees.

17. The repaired composite laminate as defined in any one of claims 12 to 16, wherein the repair plies are pre-impregnated with the matrix material.

18. The repaired composite laminate as defined in any one of claims 12 to 17, comprising one or more upper plies comprising a woven fibrous material.

19. The repaired composite laminate as defined in claim 18, wherein the one or more upper plies overlap onto an outer surface of the composite laminate.

20. The repaired composite laminate as defined in any one of claims 12 to 18, wherein the patch has an outer surface that is substantially flush with an outer surface of the composite laminate.

21. The repaired composite laminate as defined in any one of claims 12 to 20, wherein the repair plies comprise one or more upper plies comprising unidirectional fibrous material.

22. The repaired composite laminate as defined in any one of claims 12 to 21 , wherein:

the cavity in the laminate plies has a depth extending through a number of laminate plies; and a number of repair plies in the patch filling the cavity in the composite laminate differs from the number of laminate plies.

23. An aircraft component comprising the repaired composite laminate as defined in any one of claims 12 to 22.

24. A patch for repairing a composite laminate where the composite laminate comprises an assembly of a plurality of laminate plies, each laminate ply being oriented in accordance with a stacking sequence of the composite laminate, the patch comprising:

a matrix material for bonding the fibrous material of the repair plies.

25. The patch as defined in claim 24, the repair plies comprise a one or more lower plies comprising unidirectional fibrous material.

26. The patch as defined in any one of claims 24 and 25, wherein the stacking sequence of the patch comprise a repeating pattern of ply orientations where one of the orientations is more common than the other orientations in the pattern to cause anisotropy in the mechanical properties of the patch.

27. The patch as defined in claim 26, wherein the stacking sequence of the patch comprise a repeating pattern of ply orientations for the lower plies comprising the following ply orientations: 45 degrees, 135 degrees, 0 degrees, 90 degrees and 0 degrees.

28. The patch as defined in any one of claims 24 to 27, wherein the repair plies are pre-impregnated with the matrix material.

29. The patch as defined in any one of claims 24 to 28, wherein the repair plies comprise one or more upper plies comprising a woven fibrous material.

30. The patch as defined in claim 29, wherein the one or more upper plies are configured to overlap onto an outer surface of the composite laminate.

31. The patch as defined in any one of claims 24 to 30, wherein the repair plies comprise one or more upper plies comprising unidirectional fibrous material.

32. An aircraft component comprising the patch as defined in any one of claims 24 to 31.

33. Any, some, or all features of novelty described, suggested, referred to, exemplified, or shown herein, and corresponding patches, laminates and associated methods.