CN114986948B

CN114986948B - Repair process and method for honeycomb sandwich structure of composite material

Info

Publication number: CN114986948B
Application number: CN202110225072.6A
Authority: CN
Inventors: 沈宏; 杨雪岭; 张超; 王向辉; 李东益; 李永恒; 於亮亮; 周丽
Original assignee: AECC Commercial Aircraft Engine Co Ltd
Current assignee: AECC Commercial Aircraft Engine Co Ltd
Priority date: 2021-03-01
Filing date: 2021-03-01
Publication date: 2023-06-02
Anticipated expiration: 2041-03-01
Also published as: CN114986948A

Abstract

One aspect of the present disclosure relates to a method of repairing a composite honeycomb sandwich structure, the method comprising removing damaged skin on a laminate and trimming skin cuts to a regular geometry; removing damaged core material in the honeycomb core to form a honeycomb core notch; grinding each layer of skin on the skin incision into a step with the shape consistent with the geometric shape; installing a replacement honeycomb core into the honeycomb core cutout by means of shims and bonding tools; removing the shim and the bonding tool after curing the replacement honeycomb core; sequentially mounting a layer of material corresponding to each layer of skin of the laminated board from inside to outside and the geometric shape and the size of the layer on the step on the skin incision; and (3) curing the layering. Other aspects are also encompassed by the present disclosure.

Description

Repair process and method for honeycomb sandwich structure of composite material

Technical Field

The present application relates generally to composite repair, and more particularly to repair processes and methods for aerospace composite honeycomb sandwich structures.

Background

In the running and maintenance process of domestic civil aviation engines and the like, damage to a Nacelle (Nacell) structure, such as surface scratches, pits, tears, perforations and the like, is inevitably caused by factors such as foreign object impact, lightning stroke, maintenance error and the like.

About 60% of the nacelle structure is a composite honeycomb sandwich structure. The traditional composite material repairing process needs to develop special repairing process aiming at different damage types of individual structural members, so that manpower and material resources are wasted, and management confusion is easy to cause. When an airline maintenance personnel performs repair work, the maintenance personnel needs to consult related technical manuals to acquire a repair method, which indirectly increases the maintenance time cost and affects the operation efficiency of the aeroengine.

Thus, there is a need for a typical repair process and method for forming a composite material that is highly practical for use in an aircraft.

Disclosure of Invention

An aspect of the present disclosure relates to a repair method of a composite honeycomb sandwich structure including a honeycomb core and a laminate attached to at least one side of the honeycomb core, the repair method comprising: removing damaged skin on the laminate and trimming skin cuts to a regular geometry; removing damaged core materials in the honeycomb core to form honeycomb core cuts; grinding each layer of skin on the skin incision into a step with the shape consistent with the geometric shape; installing a replacement honeycomb core into the honeycomb core cutout by means of a shim and an adhesive tool, wherein the adhesive tool is used for accurately positioning the replacement honeycomb core in the horizontal direction and the shim is used for accurately positioning the replacement honeycomb core in the vertical direction; removing the shim and the bonding tool after curing the bonding material of the replacement honeycomb core; sequentially mounting a layer of material corresponding to each layer of skin of the laminated board from inside to outside and the geometric shape and the size of the layer of material on the step on the skin incision; and curing the adhesive material of the ply.

According to an exemplary embodiment, the repair method further comprises fabricating the replacement honeycomb core after removing the damaged core material in the honeycomb core, including looking up the material and density of the honeycomb core; and cutting the honeycomb core material and processing the replacement honeycomb core according to the size and shape of the honeycomb core cut.

According to an exemplary embodiment, installing a replacement honeycomb core into the honeycomb core cutout further comprises ensuring that the bonding orientation of the replacement honeycomb core in the horizontal direction is consistent with the orientation of the honeycomb core in the composite honeycomb sandwich structure; and avoiding movement and rotation of the replacement honeycomb core in the horizontal direction by means of the bonding tool.

According to an exemplary embodiment, removing damaged skin on the laminate and trimming skin cuts to a regular geometry includes one or more of the following or any combination thereof: removing all loose falling, cracks and layering; ensuring that all fillet radii are greater than or equal to 0.5 inch; ensuring that the skin cut has an oblique angle of less than or equal to 45 °; and ensuring that the skin cuts are smooth.

According to an exemplary embodiment, the composite honeycomb sandwich structure comprises a single-sided laminate plus honeycomb core damage, and the repair method further comprises fabricating bonding tools and shims after each skin of the laminate is ground into a step on the skin kerf that conforms to the geometry of the skin kerf; or the composite honeycomb sandwich structure comprises the honeycomb core and laminated sheets attached to both sides of the honeycomb core, and the composite honeycomb sandwich structure comprises a full penetration lesion, and the repair method further comprises removing the lesion skins on both sides of the laminated sheets and trimming the skin incisions on both sides to a regular geometry, respectively, and after each layer of skin of the laminated sheets is ground on the skin incisions on both sides to a step having a shape conforming to the geometry of the skin incision on the side, respectively, manufacturing a bonding tool and a shim, respectively, on both sides.

According to an exemplary embodiment, the bonding tool has a contour conforming to the geometry of the skin cut of the corresponding side and is larger on each side than the honeycomb core cut; and the bonding tool is used for bonding with the replacement honeycomb core through light pressure and fixedly matching on the skin notch so that the replacement honeycomb core is positioned in the horizontal direction.

According to an exemplary embodiment, the shim is shaped and sized to conform to at least a portion of the step on the skin cutout of the corresponding side; and the shim is adapted to be adhered to the bonding tool by light pressure and to be fixedly fitted over the at least partial step on the skin cutout so that the replacement honeycomb core is positioned in the vertical direction.

According to an exemplary embodiment, if the composite honeycomb sandwich structure includes a full penetration lesion, installing a replacement honeycomb core into the honeycomb core cutout using an adhesive material by means of a shim and an adhesive tool further includes first installing the shim and adhesive tool on one side of the composite honeycomb sandwich structure, including: placing an adhesive tool and shim in sequence over the skin cut on the side of the repair area; fixing the shim and bonding tool in place.

According to an exemplary embodiment, curing the bonding material of the replacement honeycomb core further comprises placing a separation membrane over the shim; applying mechanical pressure or weight on the shim; and heating and maintaining until the bonding material is cured.

According to an exemplary embodiment, curing the bonding material of the layup comprises placing a vacuum bag over the layup with the same pressure on each side; heating and holding the layup until the layup of bonding material cures, wherein the layup of bonding material comprises glue, resin, or pre-dipping of the layup; and removing the vacuum bag after cooling.

Other related aspects are also encompassed by the present disclosure.

Drawings

Fig. 1 illustrates an illustration of an example composite honeycomb sandwich structure in accordance with an aspect of the disclosure.

Fig. 2 illustrates an exploded view of an example composite honeycomb sandwich structure in accordance with an aspect of the disclosure.

Fig. 3 illustrates a diagram of a composite honeycomb sandwich structure lesion according to an aspect of the present disclosure.

FIG. 4 illustrates a block diagram of a method of repairing an aerospace composite honeycomb sandwich structure in accordance with an aspect of the disclosure.

Fig. 5 illustrates a schematic diagram of repair effects of an aerospace composite honeycomb sandwich structure according to an aspect of the disclosure.

Fig. 6 illustrates a schematic diagram of a repair structure for an aerospace composite honeycomb sandwich structure according to an aspect of the disclosure.

Fig. 7 illustrates a schematic diagram of a repair material curing scheme for an aerospace composite honeycomb sandwich structure according to an aspect of the disclosure.

Detailed Description

The structural forms of aerospace composites (e.g., aircraft engine composites, etc.) mainly include two broad categories, laminate and honeycomb sandwich forms, respectively. And wherein the repair process in a honeycomb sandwich structure is most complicated.

Fig. 1 illustrates a diagram of an example composite honeycomb sandwich structure 100 in accordance with an aspect of the disclosure. As shown in fig. 1, the composite honeycomb sandwich structure may be comprised of mainly 3 components, and may be generally structured in the form of: laminate(s) 102a on one side + honeycomb core 104+ laminate(s) 102b on the other side. Although the figures show a single-sided laminate as three layers and a certain thickness of laminate and honeycomb core, these are for illustration only and not by way of limitation. Laminates according to the present disclosure may include one or more layers, the laminates and honeycomb core may have various thicknesses, and laminates on both sides are not limited to the same number of layers, thickness, and the same materials and order of arrangement. The present disclosure is equally applicable to structures having laminates on only one side of the honeycomb core.

Fig. 2 illustrates an exploded view of an example composite honeycomb sandwich structure 200 in accordance with an aspect of the disclosure. A top view and a perspective view of a honeycomb core in a composite honeycomb sandwich structure according to an example are shown in fig. 2 (a). As can be seen, the honeycomb core may have a relatively low density three-dimensional (e.g., in the direction of L, W, T) structural arrangement similar to honeycomb that may be primarily subjected to shear stress in the composite honeycomb sandwich structure 200, giving it a high flexural stiffness. Although a hexagonal cell shape is shown in the drawings, the present disclosure is not limited thereto, but may also be applied to honeycomb cores of other cell shapes, such as rectangular, circular, various reinforcing types, and the like.

A cross-sectional view of a single-sided laminate 102 in a composite honeycomb sandwich structure according to an example is shown in fig. 2 (b). As can be seen in the figures, the laminate may comprise one or more layers of the same/different materials, such as being laminated, heat and pressure bonded, etc. to form a whole.

Fig. 3 illustrates a diagram of a composite honeycomb sandwich structure lesion 300 in accordance with an aspect of the present disclosure. The composite honeycomb sandwich structure damage 300 can be divided into various forms according to the penetration level, such as a single-sided laminate damage, a penetrating single-sided laminate+honeycomb core damage, a completely penetrating damage (i.e., penetrating double-sided laminate+honeycomb core damage), etc., as shown in (a), (b), and (c) of fig. 3, respectively.

As can be appreciated, the various forms of composite honeycomb sandwich damage shown in fig. 3 are merely exemplary. Depending on the actual situation, there may be various other forms of composite honeycomb sandwich damage such as, but not limited to, loosening, peeling, cracking, delamination, penetration, etc. of the laminate; and/or breakage, cracking, deformation, penetration, etc. of the honeycomb core.

The repair process and method for each type of damage varies, but is most typical in the form of "through a single-sided laminate + honeycomb core" because it extends to, or even replaces, some other types of damage repair processes. Therefore, the present disclosure provides technical support for the development of repair processes for all types of composite damage (e.g., lightning erosion, honeycomb water intake, layering, degluing, etc.) during the operation of an aircraft engine by building a set of typical composite sandwich repair process models.

FIG. 4 illustrates a block diagram of a repair method 400 for an aerospace composite honeycomb sandwich structure in accordance with an aspect of the disclosure. As shown in fig. 4, a method 400 of repairing an aerospace composite honeycomb sandwich structure may include performing a honeycomb sandwich structure pretreatment 402 at block 402. According to an exemplary embodiment, the honeycomb sandwich pretreatment 402 may include damage region pre-inspection, damage removal, machining a sanded region, inspecting a pre-treated damage region, damage region protection, repair region surface cleaning, and the like, as described in more detail below.

The repair method 400 of an aerospace composite honeycomb sandwich structure may include preparing repair materials and repair tools at block 404. According to an exemplary embodiment, preparing the repair material may include making repair plies and/or replacing the honeycomb core. According to an exemplary embodiment, preparing the repair tool may include making a bonding tool and a shim for temporarily replacing the layup to fill the void between the bonding tool and the external pressure applicator to accurately position the replacement honeycomb core in a vertical (e.g., T) direction; while the bonding tool is used to avoid movement and/or rotation of the replacement honeycomb core during repair, thereby allowing accurate positioning of the replacement honeycomb core in the horizontal (e.g., L and W) directions.

The repair method 400 of an aerospace composite honeycomb sandwich structure may further include installing a replacement honeycomb core at block 406. According to an exemplary embodiment, installing the replacement honeycomb core may include adhering the replacement honeycomb core to a corresponding location in the honeycomb core cutout using an adhering tool and a shim. According to a further exemplary embodiment, installing the replacement honeycomb core may further include curing the bonding material of the replacement honeycomb core and removing the bonding tool and shim. According to a further exemplary embodiment, installing the replacement honeycomb core may further comprise cleaning the repair area.

The method 400 of repairing an aerospace composite honeycomb sandwich structure may further include installing a finishing ply at block 408. According to an exemplary embodiment, installing the repair mat may include applying the mat layer by using an adhesive material. According to some exemplary embodiments, the adhesive material may be an adhesive layer or a resin. According to other exemplary embodiments, at least some of the plies may be resin pre-preg plies, such that no additional adhesive layer may be required.

The repair method 400 of the aerospace composite honeycomb sandwich structure may further include performing repair zone curing at block 410. According to an exemplary embodiment, curing the repair area may include providing a uniform pressure to the repair area, heating the repair area, cooling the repair area, and so forth.

The repair method 400 of the aerospace composite honeycomb sandwich structure may further include performing a repair area test at block 412. According to an exemplary embodiment, the repair area test may include checking whether the repair area is delaminated or degummed, etc. Performing repair area testing may include, for example, using ultrasonic inspection and/or tapping testing, etc.

The repair method 400 of the aerospace composite honeycomb sandwich structure may further include post-treating the repair area at block 414. According to an exemplary embodiment, post-treating the repair area may include, for example, planarizing the repair area, cleaning the repair area, repairing the surface coating, and the like.

The repair method 400 of the aerospace composite honeycomb sandwich structure of fig. 4 is primarily directed to repair of damage patterns penetrating a single-sided laminate panel + honeycomb core, and may also be generalized to repair processes of single-sided laminate damage and/or full penetration damage, and the like.

For example, for a single-sided laminate damage (i.e., no damage to the honeycomb core), steps 402, 404, 408, 410, 412, etc. in method 400 may be employed accordingly, and the foregoing steps may be adapted accordingly. For example, for step 402, the damaged area may involve only a single-sided laminate, not a honeycomb core. As another example, for step 404, preparation of repair material may not require fabrication of replacement honeycomb cores, while preparation of repair tools may not require fabrication of bonding tools, shims, and the like.

As another example, for a fully penetrating lesion (i.e., penetrating a double sided laminate + honeycomb core lesion), the various steps 402-414 of the method 400 may be adaptively modified to correspondingly include a treatment on both sides. For example, for step 402, the damaged area may relate to a double sided laminate as well as a honeycomb core. Whereby the lesion field pre-examination may comprise a bilateral lesion field pre-examination; the lesion removal may include double-sided lesion removal; machining the lapping region may include machining of a double sided lapping region; examining the pretreated lesion area may include examining the pretreated lesion area on both sides; the damaged area protection may include damaged area protection on both sides, repair area surface cleaning may include repair area surface cleaning on both sides, and the like, as described in more detail below.

Some specific examples of each step of the repair method 400 of the aerospace composite honeycomb sandwich structure of fig. 4 are given below. According to a particular embodiment, a specific procedure for repairing a typical damage pattern of a composite honeycomb sandwich structure may include:

A. inspection of structural damage, for example:

(1) Performing ultrasonic examination around the damaged area to determine the size of the damage and delamination of the skin; or (b)

(2) A tapping test was performed around the lesion to determine lesion size and skin delamination.

B. Removing damaged skin, for example:

(1) And removing all loose falling, cracks and layering. For example, the cutting tool rotational speed may be 18,000-23,000 rpm. Skin cuts may be trimmed to a regular geometry, e.g., circular, oval, rectangular, triangular, etc.);

(2) Ensuring that all fillet radii are greater than or equal to, for example, 0.5 inch (12.7 mm);

(3) The skin cut of the outer surface ensures an oblique angle (gentle slope) of less than or equal to, for example, 45 °;

(4) Ensuring that all skin incisions are smooth; for example, it may be sanded with 80 mesh sandpaper.

C. Find out how many layers are in the damaged skin, for example:

(1) Measuring the thickness of the damaged skin (note: the thickness of the primer size or other coating should be removed); or (b)

(2) The number of layers is counted by visual inspection, for example:

a) Polishing the edge of a small piece of skin by using abrasive paper;

b) The number of layers in the skin is counted.

D. Removal of damaged core material with a hole digger or other tool, for example:

(1) And removing all loose falling, cracks and layering. For example, the cutting tool rotational speed may be 18,000-23,000 rpm.

Removal of the damaged core typically involves removal of a cylindrical damaged core in order to force the honeycomb kerf uniformly. The present disclosure is not limited thereto but may also include other shapes such as triangular prisms, quadrangular prisms, hexagonal prisms, etc. or other shapes.

E. Machining the sanding area, for example sanding a step on the outer skin with a sanding tool, for example:

the sanding tool may be rotated at 18,000-23,000rpm to maintain the step profile consistent with the skin incision profile. As can be appreciated, the ground area may include other ground areas besides the preferred steps, such as wedge-shaped structures, curved steps, or other curved structures.

F. Inspection of the structural members ensures complete removal of all layered skins, e.g

(1) Performing an ultrasonic examination around the lesion area; or (b)

(2) A tapping test is performed around the damaged area.

G. Protective measures are taken around the damaged area, for example:

(1) A teflon tape, for example 2.0 inches (50.8 mm) wide, is adhered around the repair area for protection.

a) In the region of, for example, 2 layers, the tape is applied such that the inner edge of the tape is spaced from the skin cut by, for example, 1.5

Inches (38.1 mm); or (b)

b) In the region of, for example, 4 layers, the tape is applied such that the inner edge of the tape is spaced from the skin cut by, for example, 2.5

Inches (63.5 mm).

H. The surface coating and ECS outer paint layer (outer skin only) of the repair area are removed, for example:

(1) The surface coating of the repair area on the outer skin and the ECS outer paint layer are removed with abrasive sand paper.

I. Cleaning repair areas, for example:

the repair area is cleaned with clean fleece and a cleaning agent. And (3) wiping with different clean cashmere cloths before volatilizing and drying the solvent.

J. And drying the repair area.

K. Bonding tools and shims are made, for example:

(1) The bonding tool is made, depending on whether the single-sided laminate + honeycomb core is damaged or completely damaged, one or two bonding tools can be made accordingly, for example:

a) The bonding tool was made as follows:

i. ensuring that the bonding tool is larger on each side than the honeycomb core cutout, e.g., about 2.0 inches (50.8 mm) or so, than the honeycomb core cutout.

Ensuring that the appearance of the bonding tool conforms to the skins (e.g., inner skin and outer skin) of the corresponding sides.

Ensuring that the bonding tool is securely mated in place by light pressure (e.g., with a finger or other tool) (i.e., by bonding the bonding material to the honeycomb core and securely mating it over the skin cutout such that the honeycomb core is positioned in a horizontal direction).

(2) For each bonding tool, a corresponding shim is made, for example:

a) The shim is made as follows:

i. ensuring that the shape and size of the shim conforms to at least a portion (e.g., the surrounding) of the stepped region of the skin cutout.

Ensuring that the shim is securely mated in place by light pressure (e.g., with a finger or other tool) (i.e., by an adhesive material adhered to the adhesive tool and securely mated over at least a portion of the step on the skin cutout such that the honeycomb core is oriented in a vertical direction).

(3) The sharp edges of the bonding tool and shim are removed with abrasive paper.

(4) Cleaning bonding tools and shims, such as:

clean fleece and cleaning agent are used to clean the bonding tool and shim. And wiping the surface with other clean cashmere cloth before volatilizing and drying the cleaning agent.

L. making a replacement honeycomb core, for example:

(1) Searching the material and density of the core material;

(2) The honeycomb core is cut and the replacement honeycomb core is machined to the correct size and shape according to the size and shape of the removed damaged core.

Cutting the repair layup to ensure that the repair shim layup conforms to the skin cut, for example:

(1) The shims are cut from, for example, a M20/40%/G904 graphite fiber prepreg.

a) 2 shims were cut on each fabric. Each shim ensures a minimum fillet radius of 0.5 inches

(12.7mm)。

b) Each unidirectional tape is cut with a shim. Each shim ensures a minimum corner radius of 0.5 inch (12.7 mm).

(2) Two patches are cut from, for example, a M20/40%/G904 graphite fiber prepreg. Ensuring that the patch shape conforms to the skin incision and ground step shape. Ensuring a minimum fillet radius of, for example, 0.5 inch (12.7 mm).

(3) The adhesive layer is cut from, for example, FM300-2K or-2M. Ensuring that the appearance of the adhesive layer is consistent with the skin notch and the polished step. Ensuring a minimum fillet radius of, for example, 0.5 inch (12.7 mm).

(4) The buff layer is cut from, for example, a M20/39%/120 fiberglass prepreg. Ensuring that the patch shape conforms to the skin incision and ground step shape. Ensuring a minimum fillet radius of, for example, 0.5 inch (12.7 mm).

N. cutting foam, for example:

(1) The foam is cut from, for example, FM4980A or MA-562.

(2) And cutting enough foaming glue until the core material can be wrapped.

If the damage is complete penetration, one side of the installation repair material is prepared, for example:

(1) Placing a layer of non-porous separation membrane over the shim so that external repair material does not adhere to the shim;

(2) Placing the bonding tool and shim in sequence on one side of the repair area (i.e., the bonding tool is closer to the honeycomb core than the shim);

(3) The shim and bonding tool are held in place with, for example, teflon tape.

If the single-sided laminate sheet+honeycomb core is damaged, step O is skipped.

P. installation of replacement honeycomb cores, for example:

(1) Coating and wrapping the foaming glue or resin for replacing the honeycomb core for a circle (for example, comprising side surfaces, a top surface and a bottom surface);

(2) The replacement honeycomb core is placed into the honeycomb core cutout, ensuring that the bonding orientation of the replacement honeycomb core in the horizontal direction (e.g., L and W) is consistent with the orientation of the original honeycomb core. The method comprises the steps of carrying out a first treatment on the surface of the

(3) Sequentially placing the bonding tool and the shim over the honeycomb core (i.e., the bonding tool is closer to the honeycomb core than the shim);

(4) Fixing the shim and bonding tool in place with Teflon tape;

(5) The foaming glue is cured, for example:

a) The nonporous separation membrane is placed in a repair area.

b) Three thermocouples were placed in the repair area and secured in place with teflon tape.

c) Mechanical pressure or weight is applied to the shim and bonding tool to hold the foam in place.

d) The repair area is heated with a thermocouple or a heating lamp.

i. Heating is performed at a rate of, for example, 1-5F (1-4℃)/min, until each thermocouple exhibits, for example, 250+ -10F (121+ -5℃). This temperature is maintained for 60 minutes, for example.

e) The separation membrane, shim and bonding tool are removed.

f) And removing foaming adhesive in the skin polishing gradient area by using the eye mask sand paper.

g) The repair area is cleaned with clean fleece and a cleaning agent. And (5) wiping the surface of the repair area with other clean fleece before the cleaning agent is dried.

And q. installing repair plies in the repair area (one or both sides, depending on the type of damage).

(1) An adhesive layer of, for example, FM300-2K or-2M is placed on top of the core.

(2) A prepreg sheet of, for example, M20/40%/G904 graphite fibers is placed over the adhesive layer.

(3) A repair lay-up of, for example, M20/40%/G904 graphite fibers is placed over the shim.

(4) A glass fiber prepreg, for example, M20/39%/120 is placed over the repair ply.

(5) A vacuum bag is placed in the repair area on each side. The same evacuation machine is used so that the repair area on each side has the same pressure.

(6) Curing repair areas, e.g.

a) Heating repair areas, e.g.

i. Heating is performed at a rate of, for example, 1-5F (1-4℃)/min, until each thermocouple exhibits, for example, 250+ -10F (121+ -5℃). This temperature is maintained for, for example, 120 minutes.

b) Cooling repair areas

i. The repair area is cooled, for example, to 5°f (3 ℃) per minute at a cooling rate, for example

140℉(60℃)。

c) The vacuum bag apparatus and material are removed.

The use of prepregs to repair one or more of the plies is described herein. The present disclosure may also include embodiments that use glue or resin to bond at least a portion of the plies.

Checking if the repair area has delaminated or degummed, for example:

(1) Performing ultrasonic detection on the repair area; or (b)

(2) Tapping test is carried out on the repair area.

S. ensure smooth and even repair areas, for example:

(1) Sharp edges are removed and smoothed to surfaces near the repair area using, for example, 180-320 mesh ophthalmic membrane sandpaper. The fibers of the sanding layer are sanded without damage.

And T, cleaning the repair area by using clean fleece and a cleaning agent. And (5) wiping the surface of the repair area with other clean fleece before the cleaning agent is dried.

And U, repairing the surface coating.

And V. repairing the ECS outer paint layer (only the outer skin).

Some specific embodiments of each step of the repair method 400 of the aerospace composite honeycomb sandwich structure of fig. 4 are given above. For single-sided laminate damage, step D, K, L, N, O, P, etc., may be omitted, for example.

As can be appreciated, while specific preferred embodiments of materials, processes (e.g., cutting tool speed, grinding tool type), dimensions, angles, temperatures, times, etc. are given above, one of ordinary skill in the art could make minor modifications to the specific preferred parameters or ranges of parameters described above while still remaining within the scope of the present disclosure.

On the other hand, the specific embodiments of each of the steps described above can be changed in order, omitted, or other additional steps inserted before, after, during, etc., re-combined, etc., or any combination thereof, within the reasonable scope of the present disclosure. For example, step K, step L, step M, etc. can be performed simultaneously with step (I, J), etc. or sequentially in other order.

Through the development of the typical composite material repairing process, the development cost of other types of composite material repairing processes can be reduced, and repetitive work is avoided.

Fig. 5 illustrates a schematic diagram of repair effects 500 of damage removal, replacement honeycomb core installation, and layup installation for an aerospace composite honeycomb sandwich structure according to an aspect of the disclosure.

Fig. 5 (a) shows damage removal in the case of damage to the single-sided laminate. As can be seen, only the damage on the single-sided laminate is removed in this case to form the skin cut and the step is ground on the skin cut.

Fig. 5 (b) shows damage removal in the case of penetrating a single-sided laminate + honeycomb core damage. As can be seen, in this example, the damaged single-sided laminate is removed to form the skin cut, and the damaged honeycomb core is removed to form the honeycomb core cut, and the step is polished over the skin cut.

Fig. 5 (c) shows the lesion removal in the case of a completely penetrating lesion. As can be seen, in this case, the laminate on both sides of the damage needs to be removed to form skin incisions on both sides, respectively, and the damaged honeycomb core needs to be removed to form honeycomb core incisions, and steps need to be polished on the skin incisions on both sides, respectively.

Fig. 5 (a') shows a case where the single-sided laminate is damaged. As can be seen, there is no need to replace the honeycomb core, bonding tool and shim in this example.

Fig. 5 (b') shows an alternative honeycomb core installation penetrating a single-sided laminate + honeycomb core damage. As can be seen, the honeycomb core in this example has been installed in the honeycomb core cutout. The bonding tool 502 bonds with the replacement honeycomb core and is fixedly fitted over the skin cutout such that the replacement honeycomb core is positioned in the horizontal direction. The shim 504 is bonded to the bonding tool and is fixedly fitted over at least a portion of the step in the skin cutout so that the replacement honeycomb core is oriented in the vertical direction.

Fig. 5 (c') shows an alternate honeycomb core installation with complete penetration damage. As can be seen, in this example, the bonding tool 502 and shim 504 on one side (e.g., below) are first secured in place, and then the replacement honeycomb 508 is installed in the honeycomb core cutout. The bonding tool 502 on the other side (e.g., above) is then bonded and securely mated to the replacement honeycomb core 508 over the skin cutout, and the shim 504 is bonded and securely mated to the bonding tool 502 over at least a portion of the step over the skin cutout. The double sided bonding tool 502 allows the replacement honeycomb core 508 to be positioned in a horizontal direction. The double sided shim 504 allows the replacement honeycomb core 508 to be positioned in a vertical direction.

Fig. 5 (a ") shows a ply installation of a single sided laminate damage. As can be seen, single sided laminate damage is repaired by sequentially installing a layup 506 conforming to the material of each skin of the laminate from inside to outside and its geometry and dimensions on the steps, as shown, over the skin cuts.

Fig. 5 (b ") shows a lay-up installation penetrating a single-sided laminate + honeycomb core damage. As can be seen, the shim 504 and bonding tool 502 are removed after curing the bonding material of the honeycomb core, and then a layer 506 of material corresponding to each layer of skin from the inside to the outside of the laminate and its geometry and dimensions on the steps is sequentially installed over the skin cut as shown. In this way, the penetration of the single-sided laminate + honeycomb core damage is repaired.

Fig. 5 (c ") shows a fully penetrating damaged layup installation. As can be seen, after curing the bonding material of the honeycomb core, the two-sided shims 504 and bonding tool 502 are removed, and then a layer 506 conforming to the material of each layer of skin of the laminate from inside to outside and its geometry and dimensions on the steps is sequentially installed onto the skin cut of that side, respectively, as shown. In this way, the completely penetrating lesion is repaired.

Fig. 6 illustrates a schematic diagram of a repair structure 600 for an aerospace composite honeycomb sandwich structure according to an aspect of the disclosure. As shown in fig. 6, repair structure 600 of an aerospace composite honeycomb sandwich structure includes repair area 602, where damaged skin and damaged core material have been removed after inspection of the structure for damage as described in connection with block 402 of fig. 4 and/or step A, B, D, E of fig. 5, etc., and a sanded area 604 (e.g., sanded out of a step on the outer skin) has been machined.

Thereafter, the replacement honeycomb core 608 is installed into the honeycomb cutout by the adhesive layer 606 or other bonding material using repair tools (including, for example, bonding tools and shims, not shown) for temporarily replacing the lay-up to fill the void between the bonding tools and the external pressure applicator, thereby accurately positioning the replacement honeycomb core 608 in the vertical (e.g., T) direction; while the bonding tool is used to avoid movement and/or rotation of the replacement honeycomb core 608 during repair, thereby allowing for accurate positioning of the replacement honeycomb core 608 in the horizontal (e.g., L and W) directions.

Finally, after removing the repair tools (including, for example, bonding tools and shims), a repair ply 612 corresponding to the laminate plate is installed.

Although a repair structure 600 is shown in fig. 6 penetrating a single-sided laminate + honeycomb core damage, the present application may correspondingly include repair of single-sided laminate damage as well as repair of full penetration damage.

Fig. 7 illustrates a schematic diagram of a repair material curing scheme 700 for an aerospace composite honeycomb sandwich structure according to an aspect of the disclosure.

As can be seen, one side of the composite honeycomb sandwich to be repaired is placed on an electric blanket 702 and the electric blanket is sized to ensure that the repair area is at least 2 inches beyond each side to ensure that the repair area is sufficiently and uniformly heated. The temperature measurement on this side is performed by thermocouple 704.

Above repair ply 708 are placed in order a hole separating membrane 710, a surface air permeable cloth 712, a solid separating membrane 714, and an electric blanket 718. The apertured separation membrane 710 may be sized to extend beyond the repair area, for example, by at least 1 inch, on each side. The surface breather cloth 712 may be sized to extend beyond the apertured separating film, for example, by at least 2 inches, on each side. The dimensions of the solid separation membrane 714 may be substantially the same as the dimensions of the electric blanket 718 and may be slightly smaller than the dimensions of the surface air permeable cloth 712, for example, to ensure that the dimensions of the surface air permeable cloth 712 exceed the dimensions of the solid separation membrane 714 and the electric blanket 718 on each side by about 0.5 inch, for example, and to ensure that the dimensions of the solid separation membrane 714 and the electric blanket 718 exceed the dimensions of the apertured separation membrane on each side by about 1.5 inch, for example. The temperature measurement on this side is performed by thermocouple 716. The dimensions of the ventilation cloth 720 may remain substantially the same as the dimensions of the surface ventilation cloth 712.

A vacuum bag 724 is placed over the repair area on each side and sealed using seals 722. The same vacuum machine is used to draw vacuum through vacuum line 728 via vacuum probe 730 so that the repair area on each side has the same pressure. The pressure may be monitored using a vacuum gauge 726.

During curing, the repair area may be heated first using electric blankets 702 and 718. For example, heating may be performed at a rate of 1-5F (1-4℃)/min until each thermocouple (704, 716) exhibits an temperature of, for example, 250+ -10F (121+ -5℃). This temperature was maintained for 120 minutes.

Subsequently, the repair area may be cooled. For example, the repair area may be cooled to, for example, 140F (60℃) at a cooling rate of 5F (3℃)/min.

Finally, vacuum bag equipment and material, etc. are removed. At this point the curing is complete. Repair area testing, repair area post-processing, etc., may be performed further afterwards, as previously described.

Through the development of the typical composite material repair process of the patent, the research and development cost of other types of composite material repair processes can be reduced, repeated work is avoided, other types of repair processes can directly refer to or partially refer to the scheme, and thus, the repair process is more convenient for line maintenance personnel to review the technical manual, the maintenance working time of the line maintenance personnel is reduced, and the engine running efficiency is improved. In addition, the typical repair process can also be used for training and practical operation exercises of maintenance personnel, and the capability growth of the maintenance personnel is promoted.

What has been described above is merely an illustrative embodiment of the present invention. The scope of the invention is not limited in this respect. Any changes or substitutions that would be easily recognized by those skilled in the art within the technical scope of the present disclosure are intended to be covered by the present invention.

The methods disclosed herein comprise one or more steps or actions for achieving the described method. These method steps and/or actions may be interchanged with one another without departing from the scope of the claims. In other words, unless a specific order of steps or actions is specified, the order and/or use of specific steps and/or actions may be modified without departing from the scope of the claims.

It is to be understood that the claims are not limited to the precise configurations and components illustrated above. Various modifications, substitutions and alterations can be made in the arrangement, operation and details of the methods and apparatus described above without departing from the scope of the claims.

Claims

1. A method of repairing a composite honeycomb sandwich structure comprising a honeycomb core and a laminate attached to at least one side of the honeycomb core, the method comprising:

removing damaged skin on the laminate and trimming skin cuts to a regular geometry;

removing damaged core materials in the honeycomb core to form honeycomb core cuts;

grinding each layer of skin on the skin incision into a step with the shape consistent with the geometric shape;

installing a replacement honeycomb core into the honeycomb core cutout by means of a shim and an adhesive tool, wherein the adhesive tool is used for accurately positioning the replacement honeycomb core in a horizontal direction and the shim is used for accurately positioning the replacement honeycomb core in a vertical direction;

removing the shim and the bonding tool after curing the bonding material of the replacement honeycomb core;

sequentially mounting a layer of material consistent with the geometry and dimensions of each layer of skin of the laminate from inside to outside on the skin incision; and

and curing the adhesive material of the ply.

2. The repair method of claim 1 further comprising, after removing damaged core material in the honeycomb core, fabricating the replacement honeycomb core comprising:

searching the material and density of the honeycomb core; and

cutting the honeycomb core material and processing the replacement honeycomb core according to the size and shape of the honeycomb core cut.

3. The repair method of claim 2 wherein installing a replacement honeycomb core into the honeycomb core cutout further comprises:

ensuring that the bonding orientation of the replacement honeycomb core in the horizontal direction is consistent with the orientation of the honeycomb core in the composite honeycomb sandwich structure; and

the displacement and rotation of the replacement honeycomb core in the horizontal direction is avoided by means of the bonding tool.

4. The repair method of claim 1, wherein removing the damaged skin on the laminate and trimming skin cuts to a regular geometry comprises one or more of the following or any combination thereof:

removing all loose falling, cracks and layering;

ensuring that all fillet radii are greater than or equal to 0.5 inch;

ensuring that the skin cut has an oblique angle of less than or equal to 45 °; and

ensuring that the skin cuts are smooth.

5. The repair method of claim 1 wherein the composite honeycomb sandwich structure comprises a single-sided laminate plus honeycomb core damage, and the repair method further comprises creating bonding tools and shims after each skin of the laminate is ground into a step on the skin kerf that conforms to the geometry of the skin kerf; or alternatively

The composite honeycomb sandwich structure comprises the honeycomb core and laminated sheets attached to both sides of the honeycomb core, and the composite honeycomb sandwich structure comprises a full penetration lesion, and the repair method further comprises removing the lesion skins on both sides of the laminated sheets and trimming the skin incisions on both sides to a regular geometry, respectively, and after each layer of skin of the laminated sheets is ground on the skin incisions on both sides to a step having a shape conforming to the geometry of the skin incision on the side, respectively, manufacturing a bonding tool and a shim, respectively, on both sides.

6. The method of claim 5, wherein the bonding tool has a contour conforming to the geometry of the skin cutout of the corresponding side and is larger on each side than the honeycomb core cutout; and is also provided with

The bonding tool is used for bonding with the replacement honeycomb core through light pressure and fixedly matching with the skin notch so that the replacement honeycomb core is positioned in the horizontal direction.

7. The repair method of claim 5, wherein the shim is shaped and sized to conform to at least a portion of the step on the skin cutout of the corresponding side; and is also provided with

The shim is adapted to be bonded to the bonding tool by light pressure and to be fixedly fitted over the at least partial step on the skin cutout so that the replacement honeycomb core is positioned in a vertical direction.

8. The repair method of claim 5 wherein, if the composite honeycomb sandwich structure includes a full penetration damage, installing a replacement honeycomb core into the honeycomb core cutout using an adhesive material with a shim and an adhesive tool further comprises first installing the shim and adhesive tool on one side of the composite honeycomb sandwich structure, comprising:

placing an adhesive tool and shim in sequence over the skin cut on the side of the repair area; and

the shim and bonding tool are secured in place.

9. The repair method of claim 1 wherein curing the bonding material of the replacement honeycomb core further comprises:

placing a separation membrane over the shim;

applying mechanical pressure or weight on the shim;

and heating and maintaining until the bonding material is cured.

10. The repair method of claim 1 wherein curing the layup of adhesive material comprises:

placing a vacuum bag over the layup with the same pressure on each side;

heating and holding the layup until the layup of bonding material cures, wherein the layup of bonding material comprises glue, resin, or prepreg of the layup; and

the vacuum bag is removed after cooling.