Repairing method for composite material and composite material using the same

• US 9,993,983 B2
• Filed: 02/25/2011
• Issued: 06/12/2018
• Est. Priority Date: 02/26/2010
• Status: Expired due to Fees

• Current Assignee: Mitsubishi Heavy Industries, Ltd.
• Original Assignee: Mitsubishi Heavy Industries, Ltd.
• Inventors: Nishimura, Toshihiko
• Primary Examiner(s): Afzali, Sarang
• Assistant Examiner(s): FORD, DARRELL CHRISTOPHER
• Application Number:
  US13/519,388
• Publication Number:
  US 20120301702A1
• Time in Patent Office: 2,664 Days
• Field of Search: 2940209, 2940201, 2940202, 2940211, 2940212, 2940218, 2940219, 2952501
• CPC Class Codes: B29C 65/56 using mechanical means or mechanical connections, e.g. form-fits
• CPC Class Codes: B29C 73/04 using preformed elements
• CPC Class Codes: B29C 73/06 using plugs sealing in the hole
• CPC Class Codes: B64F 5/40 Maintaining or repairing aircraft
• CPC Class Codes: Y10T 29/49732 by attaching repair preform, e.g., remaking, restoring, or patching
• CPC Class Codes: Y10T 428/249923 Including interlaminar mechanical fastener

Abstract

A repairing method for a repaired composite material and a composite material using this method, which are capable of readily repairing, in a short time, a gap due to delamination generated between layers of completely hardened composite material. In the repairing method for composite material, a hole (4) is formed through the delamination (3) generated between layers (1a, 1b) of hardened composite material, and a joining member (6) is inserted into the hole (4) so as to join the delamination (3) generated between the layers (1a, 1b) of the hardened composite material.

Claims

1. A repairing method for repairing a delamination formed between layers of a hardened composite material,

wherein the composite material is used for an aircraft, and is a laminated structure including at least a first layer of the composite material and a second layer which is laminated on the first layer and is a part of the composite material, and is hardened,

wherein an object to be repaired by the repairing method is an area define the delamination where the first layer is partially separated from the second layer so as to form a gap between the first layer and the second layer which are cured and bound,

wherein the method comprises:

drilling a hole in both the first layer and the second layer within the area defining the delamination, in a plan view of the area, so as to penetrate through the first layer and the second layer,

joining the partially separated portion by inserting a joining member into the hole and joining the first layer and the second layer by drawing the first layer and the second layer together to eliminate the gap formed by the delamination; and

applying a bending process to an end portion of the joining member that is inserted through the hole such that the first layer and the second layer remain joined together,

wherein a size of a cross section area of the hole is smaller than a size of a cross section area of the partially separated portion, in a plan view of the composite material.

2. The repairing method for composite material according to claim 1, wherein sleeve is inserted into the hole, and the joining member is inserted through the sleeve.

3. The repairing method for composite material according to claim 1,

wherein an end portion of the joining member is inserted through the hole,

wherein the end portion of the joining member that is bent is inserted into a crimping member which is a hollow cylinder having an inner diameter that permits the end portion of the joining member to pass therethrough, and

wherein the end portion of the joining member, which is inserted in the crimping member, is crimped by applying force onto an outer side of the crimping member.

4. The repairing method for composite material according to claim 1, wherein an end portion of the joining member that is inserted through the hole is covered with a sealant.

5. The repairing method for composite material according to claim 1, wherein galvanic corrosion resistance material is provided between the joining member and the layers of the composite material.

6. The repairing method for composite material according to claim 1, wherein the joining member is in a quadrilateral cross sectional shape perpendicular to an axial direction of the joining member, the axial direction being parallel to a direction in which the joining member is inserted into the hole.

7. The repairing method for composite material according to claim 1, wherein the joining member comprises a staple having plural legs.

8. The repairing method for composite material according to claim 7, wherein the staple is made of reinforced fiber.

9. The repairing method for composite material according to claim 1, wherein the joining member is provided with surface coating for galvanic corrosion prevention.

10. The repairing method for composite material according to claim 1, wherein the joining member is made of shape memory alloy.

11. The repairing method for composite material according to claim 1,

wherein end portions of the joining member that are bent are hooked to respective hook portions of a constraining member which is stuck on a surface of the first layer of the hardened composite material,

wherein the end portions, which are hooked to the respective hook portions, are bent to opposite sides relative to respective free ends of the hook portions, and

wherein the free ends of the hook portions are disposed so as to be oriented in an opposite direction from each other.

12. A repairing method for repairing a delamination formed between layers of a hardened composite material,

wherein the composite material is used for an aircraft, and is a laminated structure including at least a first layer of the composite material and a second layer which is laminated on the first layer and is part of the composite material, and is hardened,

wherein an object to be repaired by the repairing method is an area defining the delamination, where the first layer is partially separated from the second layer so as to form a gap between the first layer and the second layer which are cured and bound,

wherein the method comprises:

drilling a plurality of holes through both the first layer and the second layer within the area defining the delamination so as to penetrate through the first layer and the second layer; and

joining the partially separated portion by inserting a joining member into the holes and joining the first layer and the second layer by drawing the first layer and the second layer together to eliminate the gap formed by the delamination, such that the first layer and the second layer remain joined together due to bending of the joining member,

wherein a size of a cross section area of each of the holes is smaller than a size of a cross section area of the partially separated portion, in a plan view of the composite material, and

wherein the joining member is a staple having a plate-like staple body which extends in an extending direction of the first layer and which contacts the first layer, a first pair of legs extending perpendicularly to the extending direction and from a pair of sides of the staple body which oppose each other in a first opposing direction, and a second pair of legs extending perpendicularly to the extending direction and from another pair of sides of the staple body which oppose each other in a second opposing direction which is perpendicular to the first opposing direction, wherein the first pair of legs and the second pair of legs are inserted into the holes, respectively.