CN110239126B

CN110239126B - Manufacturing method for aircraft engine fan blade

Info

Publication number: CN110239126B
Application number: CN201910495807.XA
Authority: CN
Inventors: 郑会龙; 康振亚; 姚俊; 杨肖芳; 赵世迁; 张谭
Original assignee: Institute of Engineering Thermophysics of CAS
Current assignee: Institute of Engineering Thermophysics of CAS
Priority date: 2019-06-10
Filing date: 2019-06-10
Publication date: 2021-06-08
Anticipated expiration: 2039-06-10
Also published as: CN110239126A

Abstract

The present disclosure provides a method of manufacturing a fan blade for an aircraft engine, comprising: step A: processing the appearance of the hollow metal framework; and B: performing fiber weaving or seam laying on the hollow metal framework to obtain a composite curing core; and C: paving a prepreg paving layer on the surface of the composite curing core, and adjusting the blade profile and the size of the blade; step D: carrying out strengthening combination on the interlamination of the prepreg paving layer and between the prepreg paving layer and the composite curing core by adopting an interlaminar strengthening method; step E: and (5) carrying out secondary curing by adopting an autoclave to form the shell of the blade. The composite material fan blade manufactured by the method is light in weight, can bear bird impact, is firm in structure, is not easy to generate interlayer separation, and meets the severe requirement of composite materials of the fan blade of the aero-engine.

Description

Manufacturing method for aircraft engine fan blade

Technical Field

The disclosure relates to the field of aircraft engine fans, and in particular relates to a manufacturing method for aircraft engine fan blades.

Background

The wide-chord fan blade made of the resin-based composite material not only has advanced pneumatic performance and vibration resistance, but also can obviously reduce the structural weight, has higher propelling efficiency, and is more and more widely applied to the development and the use of engines. The design and process of the traditional fan blade with the large bypass ratio in China are not mature, and the defects of poor bird impact resistance, easy delamination and cracking and the like exist. The belt-paving fan blade manufactured based on the traditional process method cannot meet the impact requirement and is easy to delaminate; in addition, the strength of the root of the blade manufactured by adopting the integral weaving process method is insufficient, and the root is easy to break.

In the prior art, the integral structure of the blade body of the blade is a grid-shaped frame, the composite blade body is formed based on the grid-shaped frame, the process modes of increasing the resin and metal bonding force such as sand blasting/texturing and the like are not considered in the processing of the blade body, the formed blade does not adopt methods such as Z-PIN and the like, the risk of interlayer staggering exists, and the impact resistance and the delamination resistance of the blade are lower.

Another fabrication technique for coating reinforcement of composite fan blades through the use of a mesh coating layer. The structure in the patent is mainly used for enhancing the impact resistance and delamination resistance of the blade by coating at least one part of the surface of the body of the blade with at least one part of a metal coating layer, and possible effects of the manufacturing process method include: when the blade rotates at a high speed, the extra centrifugal load brought by the metal mesh finally acts on the blade root of the composite material through the composite material layer, and the load of the composite material blade is additionally increased.

In the prior art, all the blades are made of metal materials through additive manufacturing and printing, the blade surface is made of metal materials through additive manufacturing, resin-based composite materials are not taken into consideration as forming materials of the final blade, and weight reduction structures are not designed and manufactured on the surface of the blade, so that the optimal weight reduction effect cannot be achieved. Another hybrid turbine blade and process that includes multiple insert sections, the fan blade being constructed of different materials laid up in the thickness direction in at least one chordwise region, the layers being consolidated by cementing, the process shaping method may not be as strong as a single load and force transfer unit with an intermediate metal skeleton to resist birds and the like, and the difficulty of machining may be higher, and the overall process may be more complex.

Disclosure of Invention

Technical problem to be solved

The present disclosure provides a method of manufacturing a fan blade for an aircraft engine to at least partially solve the technical problems set out above.

(II) technical scheme

According to one aspect of the present disclosure, there is provided a method of manufacturing a fan blade for an aircraft engine, comprising:

step A: processing the appearance of the hollow metal framework;

and B: performing fiber weaving or seam laying on the hollow metal framework to obtain a composite curing core;

and C: paving a prepreg paving layer on the surface of the composite curing core, and adjusting the blade profile and the size of the blade;

step D: arranging interlayer reinforcing devices between the layers of the prepreg paving layer and between the prepreg paving layer and the composite curing core for reinforcing and bonding;

step E: and (5) carrying out secondary curing by adopting an autoclave to form the shell of the blade.

In some embodiments of the present disclosure, step E is followed by: step F: and arranging a metal wrapping edge at the edge of the shell of the blade.

In some embodiments of the present disclosure, the step a comprises:

substep A1: machining or 3D printing to obtain a metal framework;

substep A2: on the metal framework, obtaining a hollow metal framework by adopting mechanical processing or electrochemical processing;

substep A3: and carrying out surface sand blasting or texturing treatment on the hollow metal framework.

In some embodiments of the present disclosure, the step a comprises:

substep A1: processing through 3D printing to obtain a hollow metal framework;

substep A2: and carrying out surface sand blasting or texturing treatment on the hollow metal framework.

In some embodiments of the present disclosure, the step B includes:

substep B1: paving seams on the composite material respectively based on the hollow parts on the two curved surfaces of the leaf basin and the leaf back on the hollow metal framework;

substep B2: and carrying out resin curing or RTM treatment to obtain the composite curing core.

In some embodiments of the present disclosure, step F further comprises profiling the metal border.

In some embodiments of the present disclosure, the prepreg layup in step C is wrapped around the blade in the radial direction from the root of the blade to the tip of the blade.

In some embodiments of the present disclosure, in step D, the interlayer reinforcing device is made of Z-Pin interlayer reinforcing material.

In some embodiments of the present disclosure, the hollowed metal skeleton is a double-arrow type or a fishbone type.

(III) advantageous effects

According to the technical scheme, the manufacturing method for the fan blade of the aircraft engine disclosed by the invention has at least one or part of the following beneficial effects:

(1) the structure of the hollow metal framework is convenient for fixing the composite curing core, and meanwhile, the self weight of the framework is effectively reduced.

(2) The interlayer reinforcing devices are distributed in the blades, so that the blades can meet the strength requirement.

(3) The metal edge-wrapping structure is beneficial to strengthening the strength of the front edge of the blade, and the front edge cracking and layering phenomena caused by bird collision are avoided.

Drawings

FIG. 1 is a block diagram of a method of manufacturing a fan blade for an aircraft engine according to an embodiment of the disclosure.

FIG. 2a is a schematic structural view of a fan blade for an aircraft engine manufactured according to an embodiment of the disclosure.

Fig. 2b is a schematic cross-sectional view taken along line a-a of fig. 2 a.

FIG. 3a is a schematic view of a blade profile of a blade for an aircraft engine manufactured according to an embodiment of the disclosure.

FIG. 3B is a schematic cross-sectional view B-B of FIG. 3 a.

FIG. 4 is a schematic view of a hollow metal skeleton structure of a fan blade for an aircraft engine manufactured according to an embodiment of the disclosure.

FIG. 5a is a schematic view of a composite cured core structure of a blade for an aircraft engine fan blade made according to an embodiment of the disclosure.

FIG. 5b is a schematic cross-sectional view of C-C of FIG. 5 a.

[ description of main reference numerals in the drawings ] of the embodiments of the present disclosure

1-laying a prepreg;

2-composite curing core;

3-hollowing out the metal framework;

4-an interlaminar reinforcement means;

and 5-metal edge wrapping.

Detailed Description

For the purpose of promoting a better understanding of the objects, aspects and advantages of the present disclosure, reference is made to the following detailed description taken in conjunction with the accompanying drawings.

Certain embodiments of the present disclosure will now be described more fully hereinafter with reference to the accompanying drawings, in which some, but not all embodiments of the disclosure are shown. Indeed, various embodiments of the disclosure may be embodied in many different forms and should not be construed as limited to the embodiments set forth herein; rather, these embodiments are provided so that this disclosure will satisfy applicable legal requirements.

In a first exemplary embodiment of the present disclosure, a method of manufacturing a fan blade for an aircraft engine is provided. FIG. 1 is a block diagram of a method of manufacturing a fan blade for an aircraft engine according to an embodiment of the disclosure. As shown in fig. 1, the present disclosure is a method of manufacturing a fan blade for an aircraft engine, comprising: step A: and (5) processing the appearance of the hollow metal framework. And B: and (4) performing fiber weaving or seam paving on the hollow metal framework to obtain the composite curing core. And C: and (3) paving a prepreg paving layer on the surface of the composite curing core, and adjusting the blade profile and the size of the blade. Specifically, the prepreg paving layer is wrapped and wound from the tenon root of the blade to the blade top of the blade along the radial direction of the blade. It should be noted that wrapping and winding need to ensure the continuity of the material, and the deformation of the fan blade can be reduced to the maximum extent, so that the deformation of the hollow metal framework of the blade and the deformation of the internal composite material are close to each other, and the magnitude of the interlayer stress is reduced. Step D: and arranging interlayer reinforcing devices between the layers of the prepreg paving layer and between the prepreg paving layer and the composite curing core for reinforcing and bonding. Z-Pin interlayer reinforcing materials can be selected specifically. Step E: and (5) carrying out secondary curing by adopting an autoclave to form the shell of the blade. Step F: and arranging a metal wrapping edge at the edge of the shell of the blade, and carrying out profiling machining on the metal wrapping edge.

Said step a, in some embodiments, comprises: substep A1: machining or 3D printing to obtain a metal framework; substep A2: on the metal framework, obtaining a hollow metal framework by adopting mechanical processing or electrochemical processing; substep A3: and carrying out surface sand blasting or texturing treatment on the hollow metal framework. In other embodiments, the step a includes: substep A1: processing through 3D printing to obtain a hollow metal framework; substep A2: and carrying out surface sand blasting or texturing treatment on the hollow metal framework.

The step B comprises the following steps: substep B1: and (3) paving and sewing the composite material on the basis of the hollow parts on the two curved surfaces of the leaf basin and the leaf back on the hollow metal framework respectively, so that the weaving structure of the composite curing core 2 is tightly attached to the hollow metal framework 3 to form a metal-fiber composite weaving body. Substep B2: and carrying out resin curing or RTM treatment to obtain the composite curing core. The cured composite curing core 2 is embedded in the hollow metal framework 3 to form the outer surface of the blade with a smooth blade basin, a blade back and a front edge and a rear edge.

FIG. 2a is a schematic structural view of a fan blade for an aircraft engine manufactured according to an embodiment of the disclosure. Fig. 2b is a schematic cross-sectional view taken along line a-a of fig. 2 a. As shown in fig. 2a and 2b, the blade structure includes: the prepreg paving layer 1, the composite curing core 2, the hollow metal framework 3, the interlayer reinforcing device 4 and the metal wrapping 5.

Wherein fretwork metal skeleton 3 structure is double arrow shape or for the fishbone shape, and this structure is located the core of whole aeroengine fan blade structure, and fretwork metal skeleton 3 goes up through machining or 3D prints the hollow out construction that forms, and the fretwork part is the lightening hole of fan blade, when lightening skeleton weight, also conveniently is used for the fibre when metal and carbon fiber weave to alternate with the Z to the fibre when the fibre is spread the seam. After the surface of the hollow metal framework 3 is subjected to sand blasting or texturing treatment, the roughness of the surface is increased, and the surface is convenient to combine with the interface of the composite material.

The composite curing core 2 is a blade core structure formed by a composite material woven or sewn on the basis of the hollow metal framework 3 and cured and formed, wherein the composite material can be one or more of fiber and resin.

The prepreg paving layer 1 is wrapped outside the composite curing core 2 and forms the outer surface of the blade, the interlaminar strength is strengthened by the interlaminar strengthening device 4 in the prepreg paving layer 1, the interlaminar strengthening device 4 can also be directly arranged in the composite curing core 2, and the bonding strength between the composite curing core 2 and the prepreg paving layer 1 is strengthened.

A metal wrapping edge 5 is adhered to the front end of the prepreg paving layer 1, and the structure can be used for strengthening the strength of the front edge and avoiding the phenomena of front edge cracking and delamination caused by bird collision.

FIG. 3a is a schematic view of a blade profile of a blade for an aircraft engine manufactured according to an embodiment of the disclosure. FIG. 3B is a schematic cross-sectional view B-B of FIG. 3 a. As shown in fig. 3a and 3b, the prepreg layup 1 is a prepreg layer wrapped outside the composite curing core 2, the prepreg layer wraps the composite curing core 2 and forms the outer surface of the blade, a metal wrapping 5 is adhered to the front end of the prepreg layup 1, the metal wrapping 5 is mainly used for bearing the impact force of the blade on a bird during high-speed operation, and the metal wrapping 5 and the hollow metal framework of the composite curing core 2 are overlapped. The interlaminar reinforcing devices 4 are distributed in the blade, can be used for strengthening the interlaminar strength of each layer in the prepreg layer 1, and can also strengthen the interlaminar bonding between the prepreg layer 1 and the composite curing core 2, so that the blade can meet the strength requirement.

FIG. 4 is a schematic view of a hollow metal skeleton structure of a fan blade for an aircraft engine manufactured according to an embodiment of the disclosure. As shown in fig. 4, the hollowed metal framework 3 may be in a double-arrow shape or a fishbone shape, and the hollowed parts are reasonably distributed on the surface of the hollowed metal framework 3 according to strength design, so that the overall weight of the framework is reduced, the mechanical load on the root of the blade during operation is reduced, the weight reduction of the blade can be effectively realized, and the related performance of the engine is improved.

FIG. 5a is a schematic view of a composite cured core structure of a blade for an aircraft engine fan blade made according to an embodiment of the disclosure. FIG. 5b is a schematic cross-sectional view of C-C of FIG. 5 a. As shown in fig. 5a and 5b, the inner side of the arc is the basin of the fan blade of the aircraft engine, the outer side of the arc is the back of the blade, the basin and the back of the blade are both finished by a seam paving method, the seam paving is performed to be close to the basin and the back of the blade, fiber puncture and sewing are realized between the basin and the back of the blade through lightening holes, the paving layers of the basin and the back of the blade are tightly wrapped on the hollow metal framework 3, and the final woven body is formed by resin curing to form the final appearance characteristics of the basin and the back of the blade consisting of the composite curing core 2. The composite curing core 2 after resin curing can be embedded in the longitudinal groove of the hollow metal framework 3 to finally form a smooth appearance curved surface, so that the prepreg paving layer 1 can be conveniently wrapped, paved and cured.

So far, the embodiments of the present disclosure have been described in detail with reference to the accompanying drawings. It is to be noted that, in the attached drawings or in the description, the implementation modes not shown or described are all the modes known by the ordinary skilled person in the field of technology, and are not described in detail. Further, the above definitions of the various elements and methods are not limited to the various specific structures, shapes or arrangements of parts mentioned in the examples, which may be easily modified or substituted by those of ordinary skill in the art.

From the above description, those skilled in the art should clearly recognize that the present disclosure provides a method of manufacturing an aircraft engine fan blade.

In summary, the present disclosure provides a manufacturing method for an aircraft engine fan blade, which is light in weight, capable of withstanding bird impact, firm in structure, and less prone to interlayer separation, and meets the composite material requirement of the aircraft engine fan blade, and is suitable for being widely used in the field of aircraft engine manufacturing.

Furthermore, the word "comprising" does not exclude the presence of elements or steps not listed in a claim. The word "a" or "an" preceding an element does not exclude the presence of a plurality of such elements.

In addition, unless steps are specifically described or must occur in sequence, the order of the steps is not limited to that listed above and may be changed or rearranged as desired by the desired design. The embodiments described above may be mixed and matched with each other or with other embodiments based on design and reliability considerations, i.e., technical features in different embodiments may be freely combined to form further embodiments.

Similarly, it should be appreciated that in the foregoing description of exemplary embodiments of the disclosure, various features of the disclosure are sometimes grouped together in a single embodiment, figure, or description thereof for the purpose of streamlining the disclosure and aiding in the understanding of one or more of the various disclosed aspects. However, the disclosed method should not be interpreted as reflecting an intention that: that is, the claimed disclosure requires more features than are expressly recited in each claim. Rather, as the following claims reflect, disclosed aspects lie in less than all features of a single foregoing disclosed embodiment. Thus, the claims following the detailed description are hereby expressly incorporated into this detailed description, with each claim standing on its own as a separate embodiment of this disclosure.

The above-mentioned embodiments are intended to illustrate the objects, aspects and advantages of the present disclosure in further detail, and it should be understood that the above-mentioned embodiments are only illustrative of the present disclosure and are not intended to limit the present disclosure, and any modifications, equivalents, improvements and the like made within the spirit and principle of the present disclosure should be included in the scope of the present disclosure.

Claims

1. A method of manufacturing a fan blade for an aircraft engine, comprising:

step A: processing the appearance of the hollow metal framework; carrying out surface sand blasting or texturing treatment on the hollow metal framework;

and B: performing fiber weaving or seam laying on the hollow metal framework to obtain a composite curing core; wherein, the step B comprises the following steps:

substep B2: carrying out resin curing or RTM treatment to obtain a composite curing core;

step D: arranging interlayer reinforcing devices between the layers of the prepreg paving layer and between the prepreg paving layer and the composite curing core for reinforcing and bonding; the interlayer reinforcing devices are distributed in the blade;

2. The method of manufacturing a fan blade for an aircraft engine as defined in claim 1, further comprising, after said step E:

step F: and arranging a metal wrapping edge at the edge of the shell of the blade.

3. The method of manufacturing a fan blade for an aircraft engine as defined in claim 1, wherein the step a comprises:

substep A1: machining or 3D printing to obtain a metal framework;

substep A2: on the metal framework, the hollow metal framework is obtained by adopting mechanical processing or electrochemical processing.

4. The method of manufacturing a fan blade for an aircraft engine as defined in claim 1, wherein the step a comprises:

substep A1: and (4) printing and processing through 3D to obtain the hollow metal framework.

5. The method of manufacturing a fan blade for an aircraft engine of claim 2, wherein step F further comprises profiling the metal hem.

6. The method for manufacturing a fan blade for an aircraft engine of claim 1, wherein the prepreg layup in step C is wrapped and wound in the radial direction of the blade from the tenon root to the tip of the blade.

7. The method for manufacturing an aircraft engine fan blade according to claim 1, wherein in step D, the interlayer reinforcing device is made of Z-Pin interlayer reinforcing material.

8. The method for manufacturing an aircraft engine fan blade according to claim 1, wherein the hollowed-out metal skeleton is of a double-arrow head type or a fishbone type.