CN114919210A - Forming method of composite material wing framework - Google Patents
Forming method of composite material wing framework Download PDFInfo
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- CN114919210A CN114919210A CN202210429533.6A CN202210429533A CN114919210A CN 114919210 A CN114919210 A CN 114919210A CN 202210429533 A CN202210429533 A CN 202210429533A CN 114919210 A CN114919210 A CN 114919210A
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- 238000000034 method Methods 0.000 title claims abstract description 53
- 239000002131 composite material Substances 0.000 title claims abstract description 30
- 238000007731 hot pressing Methods 0.000 claims abstract description 20
- 238000000465 moulding Methods 0.000 claims abstract description 14
- 230000000149 penetrating effect Effects 0.000 claims description 6
- 239000004744 fabric Substances 0.000 claims description 3
- 238000002955 isolation Methods 0.000 claims description 3
- 238000005507 spraying Methods 0.000 claims description 3
- 238000009417 prefabrication Methods 0.000 claims 1
- 238000004519 manufacturing process Methods 0.000 abstract description 16
- 238000004026 adhesive bonding Methods 0.000 description 5
- 230000000694 effects Effects 0.000 description 3
- 239000003292 glue Substances 0.000 description 3
- 239000000463 material Substances 0.000 description 3
- 239000011347 resin Substances 0.000 description 3
- 229920005989 resin Polymers 0.000 description 3
- 238000007711 solidification Methods 0.000 description 3
- 230000008023 solidification Effects 0.000 description 3
- 229920000049 Carbon (fiber) Polymers 0.000 description 2
- 239000004917 carbon fiber Substances 0.000 description 2
- 239000003795 chemical substances by application Substances 0.000 description 2
- 239000000835 fiber Substances 0.000 description 2
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 description 2
- 230000003014 reinforcing effect Effects 0.000 description 2
- 229920001187 thermosetting polymer Polymers 0.000 description 2
- 239000000853 adhesive Substances 0.000 description 1
- 230000001070 adhesive effect Effects 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 239000000203 mixture Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000004806 packaging method and process Methods 0.000 description 1
- 238000002360 preparation method Methods 0.000 description 1
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/54—Component parts, details or accessories; Auxiliary operations, e.g. feeding or storage of prepregs or SMC after impregnation or during ageing
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/30—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29L—INDEXING SCHEME ASSOCIATED WITH SUBCLASS B29C, RELATING TO PARTICULAR ARTICLES
- B29L2031/00—Other particular articles
- B29L2031/30—Vehicles, e.g. ships or aircraft, or body parts thereof
- B29L2031/3076—Aircrafts
- B29L2031/3085—Wings
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/40—Weight reduction
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- Casting Or Compression Moulding Of Plastics Or The Like (AREA)
Abstract
The invention belongs to the technical field of aircraft manufacturing, and discloses a method for forming a composite material wing framework. The method comprises the following steps: respectively paving and sticking a plurality of layers of prepreg layers on a prefabricated front beam mould and a prefabricated rear beam mould; paving and sticking a plurality of layers of prepreg layers on a prefabricated rib mould; laying a plurality of prefabricated upper limiting pieces on the vacant surfaces of the outermost prepreg layers on the front beam mold at intervals, and simultaneously laying a plurality of prefabricated lower limiting pieces on the vacant surfaces of the outermost prepreg layers on the rear beam mold at intervals; and the two ends of the rib mould are respectively and tightly inserted between the corresponding upper limiting piece and the lower limiting piece; fixing two ends of a prefabricated positioning clamping plate on a front beam mold and a rear beam mold respectively; and demolding after hot-pressing curing molding to obtain the composite material wing framework. The invention not only simplifies the molding process, but also reduces the molding cost; and the integral strength and the stress capacity of the wing framework are also improved.
Description
Technical Field
The invention relates to the technical field of aircraft manufacturing, in particular to a method for forming a composite material wing framework.
Background
The wing is an important composition structure of the airplane, and the pressure difference between the upper side and the lower side of the wing provides the flying lift force for the airplane. The wing mainly comprises wing spars, wing ribs, rib frames, skin and other components, wherein a framework consisting of a front wing spar, a rear wing spar and the wing ribs is a stressed component of the wing. In order to improve the bearing effect of the airplane, the main stream structure of the wing framework at present is a composite material wing framework.
The composite material wing framework has large size and complex structure. Therefore, the common manufacturing method at present is co-cementing molding; the method comprises the following specific steps: firstly, paving a prepreg layer in a prefabricated front beam mould, and performing hot-pressing curing to form a front wing beam; paving a prepreg layer in a prefabricated rear beam mould, and performing hot-pressing curing to form a rear wing beam; paving a prepreg layer in the prefabricated rib mold, and performing hot-pressing curing to form a wing rib; and then placing the front wing spar, the rear wing spar and the wing ribs according to target positions, and forming the wing framework through adhesion of glue materials.
However, the method has the following defects in actual manufacturing: firstly, the preparation process comprises the steps of respectively preparing a front wing beam, a rear wing beam and wing ribs, and then gluing and combining; resulting in a correspondingly complicated manufacturing flow. Secondly, as a plurality of wing ribs are generally arranged, more tools are needed to complete the manufacturing of the wing framework; resulting in higher manufacturing costs. And most importantly, because the connection mode among the front wing beam, the rear wing beam and the wing ribs is glue joint, the wing framework is of a splicing structure in nature, so that the technical bottleneck of improving the overall strength of the wing framework can be caused on one hand, and the risk of disintegration and fracture of the wing framework from the glue joint can be caused on the other hand when the wing framework is stressed for a long time or is stressed abnormally.
Disclosure of Invention
The invention aims to provide a method for molding a composite material wing framework, which aims to solve the problems of complex molding process and high cost in the existing molding process; and the whole strength and the stress capability of the formed wing framework are difficult to improve.
In order to achieve the above purpose, the invention provides the following technical scheme:
a method of forming a composite material wing skeleton, comprising:
respectively paving and sticking a plurality of layers of prepreg layers on a prefabricated front beam mould and a prefabricated rear beam mould;
paving and sticking a plurality of layers of prepreg layers on a prefabricated rib mould;
laying a plurality of prefabricated upper limiting pieces on the vacant surfaces of the outermost prepreg layers on the front beam mold at intervals, and simultaneously laying a plurality of prefabricated lower limiting pieces on the vacant surfaces of the outermost prepreg layers on the rear beam mold at intervals; two ends of the rib die are respectively and tightly inserted between the corresponding upper limiting piece and the lower limiting piece;
fixing two ends of a prefabricated positioning clamping plate on the front beam mold and the rear beam mold respectively;
and demolding after hot-pressing curing molding to obtain the composite material wing framework.
Further, the front beam form and the back beam form are both C-shaped forms, and after the plurality of layers of prepreg layers are respectively laid on the prefabricated front beam form and the prefabricated back beam form, the method includes the following steps:
and arranging the opening directions of the front beam mould and the rear beam mould to face the prefabricated rib mould.
Further, the front beam form is a C-shaped form, and after the plurality of layers of prepreg layers are respectively laid and attached on the prefabricated front beam form and the prefabricated rear beam form, the method includes the following steps:
arranging a prefabricated rib mould deviating from the opening direction of the front beam mould;
paving and sticking a plurality of layers of prepreg layers on a prefabricated front beam side plate;
arranging the front beam side plates on two sides of the front beam mould and enabling the vacant surfaces of the prepreg layers on the outermost layers of the front beam side plates to be attached to the vacant surfaces of the prepreg layers on the outermost layers of the front beam mould;
and paving a plurality of prepreg layers in a structure formed by the empty residual surface of the outermost prepreg layer on the front beam side plate and the rest part of the empty residual surface of the outermost prepreg layer on the front beam mould.
Further, after the front beam side plates are arranged on two sides of the front beam mold and the spare surfaces of the outermost prepreg layers on the front beam side plates are attached to the spare surfaces of the outermost prepreg layers on the front beam mold, the method includes:
and aligning the front beam side plate with a waist-shaped hole prefabricated on the upper limiting piece, and penetrating a pin into the front beam side plate so that the front beam side plate has the freedom degree of moving along the direction of the rib mold.
Further, the back beam mold is a C-shaped mold, and after the plurality of layers of prepreg layers are respectively laid and attached in the prefabricated front beam mold and the prefabricated back beam mold, the method comprises the following steps:
arranging a rib mould with an opening direction deviating from the prefabricated rib mould;
paving and sticking a plurality of layers of prepreg layers on a prefabricated back beam side plate;
arranging the rear beam side plates on two sides of the rear beam mould and enabling the vacant surfaces of the prepreg layers on the outermost layers of the rear beam mould to be attached to the vacant surfaces of the prepreg layers on the outermost layers of the rear beam mould;
and paving and sticking a plurality of layers of prepreg layers in a structure formed by the spare surface of the outermost prepreg layer on the side plate of the back beam and the rest part of the spare surface of the outermost prepreg layer on the back beam mould.
Further, after the rear beam side plates are arranged on two sides of the rear beam mold and the vacant surfaces of the outermost prepreg layers on the rear beam mold are attached to the vacant surfaces of the outermost prepreg layers on the rear beam mold, the method includes:
and aligning the back beam side plate with a waist-shaped hole prefabricated on the lower limiting piece, and penetrating a pin into the waist-shaped hole so that the back beam side plate has the freedom of moving along the direction of the rib mold.
Further, the step of laying and pasting a plurality of layers of prepreg layers on the prefabricated rib mold further comprises the following steps:
and reserving the spare end of the prepreg layer at the outermost layer so that the spare end is not paved on the rib mould.
Further, after the several layers of prepreg are laid on the prefabricated rib mold, the method further comprises the following steps:
and respectively positioning and adhering the spare ends of the outermost prepreg layers on the rib molds to the outermost prepreg layers of the corresponding front beam mold and the rear beam mold.
Further, before paving and pasting a plurality of layers of prepreg layers on the prefabricated front beam mold and the prefabricated rear beam mold respectively, the method comprises the following steps:
and spraying a release agent on the surface of the prefabricated front beam mold, the prefabricated rear beam mold, the prefabricated rib mold, the prefabricated upper limiting piece and the prefabricated lower limiting piece, wherein the prepreg layer is paved and adhered, or adhering release cloth.
Further, before demolding after the hot-press curing molding to obtain the composite material wing framework, the method comprises the following steps:
and sequentially placing an isolation film, an air-permeable felt and a vacuum bag on any side surface of the front beam die, the rear beam die, the upper limiting piece, the lower limiting piece and the rib die in a filling manner.
Has the advantages that:
according to the technical scheme, the invention provides the one-time curing molding method of the composite material wing framework.
In the method, firstly, prepreg layers are paved on a prefabricated front beam mold and a prefabricated rear beam mold to form a front wing beam and a rear wing beam. Several layers of prepreg are then laid on the rib form for forming the rib. At this point, the lay-up of all prepreg layers to make the wing skeleton is complete. Secondly, it is known that both ends of a rib in a wing frame are connected with a front wing beam and a rear wing beam respectively. In particular, in one aspect, the front spar and the rear spar have a fixed mechanical profile; on the other hand, the ribs are generally plural and have fixed positions relative to the front and rear spars. Therefore, a plurality of prefabricated upper limiting pieces and a plurality of prefabricated lower limiting pieces are respectively paved on the vacant surfaces of the outermost prepreg layers on the front beam mold and the rear beam mold at intervals; and the two ends of the prefabricated rib mould are respectively and tightly inserted between the corresponding upper limiting piece and the lower limiting piece. At the moment, the upper limiting piece and the lower limiting piece can be used for carrying out auxiliary limitation on the outline of the front wing beam and the outline of the rear wing beam on one hand; on the other hand, the position of each rib is effectively limited, and the rib is prevented from being deviated. Furthermore, the size of the wing frame is relatively large (generally several meters to ten and several meters long) based on the common knowledge, so that in order to prevent the wing frame from being distorted and deformed, two ends of a prefabricated positioning clamping plate are respectively fixed on the front beam mold and the rear beam mold to limit the whole paved structure. Finally, the composite material wing framework is manufactured after hot-pressing solidification and demoulding.
According to the steps, after the prepreg layers are laid on the front beam mold, the rear beam mold and the rib mold, the prepreg layers on the rib mold and the prepreg layers in the front beam mold and the rear beam mold are in a mutually adhered state, so that the prefabricated wing framework forms an integral structure. And at the moment, the composite material wing framework can be formed through one-time hot-pressing solidification. Not only simplifies the manufacturing process; the number of moulds required during manufacturing is reduced, corresponding gluing equipment is not needed, and the manufacturing cost is reduced.
And besides paving and pasting corresponding prepreg layers on the front beam mold, the rear beam mold and the rib mold, using a prefabricated upper limiting piece and a prefabricated lower limiting piece to assist in limiting the outline of the front wing beam and the rear wing beam, and using the upper limiting piece and the lower limiting piece to limit the position of the rib mold. And simultaneously, before hot-pressing and curing, reinforcing the front beam mold and the rear beam mold by using prefabricated fixing pieces so as to limit the overall shape of the wing framework to be hot-pressed. At this time, the wing skeleton after hot-pressing curing has a good outline. Compared with the wing framework formed by firstly curing and then gluing, the wing framework has better consistency in structure and does not have a stressed section; thereby having better bearing and stress effects.
It should be understood that all combinations of the foregoing concepts and additional concepts described in greater detail below can be considered as part of the inventive subject matter of this disclosure unless such concepts are mutually inconsistent.
The foregoing and other aspects, embodiments and features of the present teachings can be more fully understood from the following description taken in conjunction with the accompanying drawings. Additional aspects of the present invention, such as features and/or advantages of exemplary embodiments, will be apparent from the description which follows, or may be learned by practice of the specific embodiments according to the teachings of the present invention.
Drawings
The figures are not intended to be drawn to scale. In the drawings, each identical or nearly identical component that is illustrated in various figures may be represented by a like numeral. For purposes of clarity, not every component may be labeled in every drawing. Embodiments of various aspects of the present invention will now be described, by way of example, with reference to the accompanying drawings, in which:
FIG. 1 is a flow chart of a method for forming a composite material airfoil skeleton according to an embodiment of the invention;
FIG. 2 is a view of another method of FIG. 1 during prepreg layup;
FIG. 3 is a flow chart of a method for pre-treating the front beam form, the rear beam form, the rib form, the upper limiting member and the lower limiting member prefabricated in FIG. 1;
FIG. 4 is a flow chart of a method of packaging before curing by hot pressing of FIG. 1;
FIG. 5 is a schematic view of a composite wing frame according to an embodiment of the present invention;
FIG. 6 is a schematic structural view of the rear spar of FIG. 5 when laid down;
FIG. 7 is a flowchart of a method of applying the rear spar of FIG. 5;
fig. 8 is a flowchart of a method of laying the front spar in fig. 5.
The reference numbers in the figures are: 1 is a front wing beam, 2 is a rear wing beam, 3 is a wing rib, 4 is a rear beam mould, 5 is a rear beam side plate, and 6 is a prepreg layer.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the drawings of the embodiments of the present invention. It should be apparent that the described embodiments are only some of the embodiments of the present invention, and not all of them. All other embodiments, which can be derived by a person skilled in the art from the described embodiments of the invention without inventive step, are within the scope of protection of the invention. Unless defined otherwise, technical or scientific terms used herein shall have the ordinary meaning as understood by one of ordinary skill in the art to which this invention belongs.
The use of "first," "second," and similar terms in the description and in the claims of the present application does not denote any order, quantity, or importance, but rather the terms are used to distinguish one element from another. Similarly, the singular forms "a," "an," or "the" do not denote a limitation of quantity, but rather denote the presence of at least one, unless the context clearly dictates otherwise. The terms "comprises" or "comprising," and the like, mean that the elements or components listed in the preceding list of elements or components include the features, integers, steps, operations, elements and/or components listed in the following list of elements or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, elements, components and/or groups thereof. "upper", "lower", "left", "right", and the like are used only to indicate relative positional relationships, and when the absolute position of the object to be described is changed, the relative positional relationships may also be changed accordingly.
The embodiment of the invention provides a method for forming a composite material wing framework. The molding method comprises the following steps: paving and pasting prepreg layers on the front beam mold and the rear beam mold; paving a prepreg layer on the rib mold; limiting the position of the rib die and the shapes of the prefabricated front wing beam and the prefabricated rear wing beam; fixing the overall shape of the integrally prefabricated composite material wing framework; and carrying out hot-pressing curing and demoulding to form the required wing framework. Therefore, the forming method can form the wing framework through one-time curing forming, so that the manufacturing cost is reduced, and the manufacturing process is simplified. The appearance profile of the wing framework is ensured in the manufacturing process, so that the internal structural strength is improved, and the bearing and stress capacity is improved.
The disclosed method of forming a composite wing frame is described in further detail below with reference to the embodiments shown in the drawings.
Example 1
As shown in fig. 1, the molding method includes the steps of:
s102, paving and sticking a plurality of layers of prepreg layers on a prefabricated front beam mould and a prefabricated rear beam mould respectively;
the prepreg layer is a material having adhesive surface formed by impregnating the resin material with the composite material fiber. In this embodiment, the composite material fiber is specifically a carbon fiber, and the resin for impregnating the carbon fiber is a thermosetting resin.
In this step, the prepreg layer is applied to the front spar mold to form a front spar, and the prepreg layer is applied to the rear spar mold to form a rear spar.
S104, paving and sticking a plurality of layers of prepreg layers on a prefabricated rib mould;
in this step, the prepreg layer is laid on the rib form for the purpose of forming each rib.
S106, laying a plurality of prefabricated upper limiting pieces on the vacant surfaces of the outermost prepreg layers on the front beam mold at intervals, and simultaneously laying a plurality of prefabricated lower limiting pieces on the vacant surfaces of the outermost prepreg layers on the rear beam mold at intervals; two ends of the rib die are respectively and tightly inserted between the corresponding upper limiting piece and the lower limiting piece;
it is known that the two ends of a rib in a wing frame are respectively connected with a front wing beam and a rear wing beam. Wherein the front spar and the rear spar have a fixed mechanical profile; and the ribs are typically plural and each rib has a fixed position relative to both the front spar and the rear spar.
Therefore, in this step, on one hand, the upper limiting piece and the lower limiting piece are used for auxiliary limiting of the outline of the front wing beam and the outline of the rear wing beam; on the other hand, the position of each rib is effectively limited, and the rib is prevented from being deviated.
S108, fixing two ends of a prefabricated positioning clamping plate on the front beam mold and the rear beam mold respectively;
the dimensions of the wing frame are relatively large (typically several meters to tens of meters long) based on common knowledge. Therefore, in order to prevent the wing framework from being distorted, in the step, two ends of the prefabricated positioning clamping plate are respectively fixed on the front beam mold and the rear beam mold so as to limit the whole paved wing framework structure.
And S110, demolding after hot-pressing curing molding to obtain the composite material wing framework.
In the step, the temperature is generally 120-180 ℃ during hot pressing and curing, the pressure is generally 0.3-0.7 MPa, and the temperature is kept for 1-3 h.
As can be seen from the above steps S102, S104, and S110, after the prepreg layers are laid on the front beam mold, the back beam mold, and the rib mold, the prepreg layers on the rib mold and the prepreg layers in the front beam mold and the back beam mold are in a state of being bonded to each other, and thus an integral structure of the prefabricated wing framework is formed. And at the moment, the composite material wing framework can be formed through one-time hot-pressing solidification. Not only simplifies the manufacturing process; the number of moulds required during manufacturing is reduced, corresponding gluing equipment is not needed, and the manufacturing cost is reduced.
As can be seen from the above steps S106 and S108, in addition to the application of the prepreg layers to the front beam mold, the rear beam mold, and the rib mold, the outer profiles of the front spar and the rear spar are assisted by the upper and lower stoppers, which are prefabricated, and the position of the rib mold is restricted by the upper and lower stoppers. And simultaneously, before hot-pressing and curing, reinforcing the front beam mold and the rear beam mold by using prefabricated fixing pieces so as to limit the overall shape of the wing framework to be hot-pressed. At this time, the wing skeleton after hot-pressing curing has a good outline. Compared with the wing framework formed by firstly curing and then gluing, the wing framework has better consistency in structure and does not have a stressed section; thereby having better bearing and stress effects.
As shown in fig. 2, in order to increase the stability of the connection between the wing rib and the front wing spar and the rear wing spar on the wing skeleton after the completion of the hot press curing, when performing step S104, the method further includes:
step S104', reserving the spare end of the outermost prepreg layer so that the spare end is not paved on the rib mold;
after step S104, the method further includes:
and S105, respectively positioning and adhering the spare ends of the outermost prepreg layers on the rib molds to the outermost prepreg layers of the corresponding front beam molds and the corresponding rear beam molds.
In the steps S104' and S105, on the other hand, the positions of the ribs are preliminarily fixed. On the other hand, the contact area between the rib and the front wing beam and between the rib and the rear wing beam is greatly increased. In this case, the joint between the front spar and the rib is formed by hot pressing the same prepreg layer, and the joint between the rear spar and the rib is also formed by hot pressing the same prepreg layer. Therefore, the stability of connection among the wing ribs, the front wing beam and the rear wing beam is improved, and the bearing and stress capacity of the whole wing framework is further improved.
As shown in fig. 3, in order to facilitate demolding after thermosetting, each mold is pretreated before step S102, including:
step S101, spraying a release agent on the surface of the prefabricated front beam mold, rear beam mold, rib mold, upper limiting piece and lower limiting piece for paving the prepreg layer, or pasting release cloth.
As shown in fig. 4, when the thermocompression curing is performed using the thermocompression tube, a package pretreatment is performed before step S110, including:
s109, sequentially placing an isolation film, an air-permeable felt and a vacuum bag on any side face of the front beam die, the rear beam die, the upper limiting piece, the lower limiting piece and the rib die in a filling mode.
Example 2
For a composite wing frame, the front and rear spars may be C-shaped.
Based on this, when the wing skeleton is formed by the forming method described in embodiment 1, both the front beam mold and the back beam mold are provided as C-shaped molds. At this time, after step S102, the method further includes:
and S103, arranging the opening directions of the front beam mold and the rear beam mold to face the prefabricated rib mold.
Example 3
As shown in fig. 5, the composite wing frame is composed of a plurality of ribs 3, and a front wing spar 1 and a rear wing spar 2 arranged at two ends of the ribs, wherein the front wing spar 1 and the rear wing spar 2 are in an i-shaped structure.
For this reason, when the wing frame is formed by the forming method described in example 1, the back beam mold 4 is provided as a C-type mold. After step S102, as shown in fig. 6 to 7, the laying of the prepreg layer on the rear spar further includes:
step S103.2, setting the opening direction of the back beam mold 4 to deviate from the prefabricated rib mold;
s103.4, paving and sticking a plurality of layers of prepreg layers 6 on a prefabricated back beam side plate 5;
step S103.6, arranging the rear beam side plates 5 on two sides of the rear beam mold 4 and enabling the spare surfaces of the outermost prepreg layers on the rear beam side plates to be attached to the spare surfaces of the outermost prepreg layers on the rear beam mold 4;
and S103.8, paving and adhering a plurality of layers of prepreg layers 6 in a structure formed by the vacant surfaces of the outermost prepreg layers on the back beam side plates 5 and the rest parts of the vacant surfaces of the outermost prepreg layers on the back beam molds.
The h-shaped rear wing beam is formed by laying two C-shaped structures, as obtained in the step S103.2 to the step S103.8. Specifically, the first C-shaped structure is formed by laying on the C-shaped back beam mold 4, and the second C-shaped structure is formed by laying on the structure formed by the back beam mold and the back beam side plate. And in the paving and pasting process, the bottom surfaces of the two C-shaped structures are directly pasted with each other, so that the formed I shape is an integral structure. And in order to increase the structural integrity of the I-shaped structure and improve the stability of the formed I-shaped rear wing beam, prepreg layers are laid on the rear wing beam side plates.
In order to prevent the profile of the back beam wing from being affected by the underpressure generated in the hot-pressing process after the back beam side plate is added, after step S103.6, the method further comprises the following steps:
step S103.7, aligning the back beam side plate 5 with a waist-shaped hole prefabricated on the lower limiting piece, and penetrating a pin into the waist-shaped hole so that the back beam side plate 5 has the freedom degree of moving along the direction of the rib mold.
And similarly, setting the front beam mold as a C-shaped mold. As shown in fig. 8, for the laying of the prepreg layer on the front spar, further comprising:
step S103.2', arranging the opening direction of the front beam mould to deviate from the prefabricated rib mould;
s103.4', paving and sticking a plurality of layers of prepreg layers on a prefabricated front beam side plate;
step S103.6', arranging the front beam side plates on two sides of the front beam mould and attaching the vacant surfaces of the outermost prepreg layers on the front beam side plates to the vacant surfaces of the outermost prepreg layers on the front beam mould;
step S103.7', aligning the front beam side plate with a waist-shaped hole prefabricated on the upper limiting piece, and penetrating a pin into the front beam side plate to enable the front beam side plate to have the freedom degree of moving along the rib mold direction;
and S103.8', paving and pasting a plurality of layers of prepreg layers in a structure formed by the empty residual surface of the outermost prepreg layer on the front beam side plate and the residual part of the empty residual surface of the outermost prepreg layer on the front beam mould.
In order to solve the process flow, when the front wing beam and the rear wing beam are both in an I shape, the step S103.2 to the step S103.8 and the step S103.2 '-the step S103.8' can be synchronously performed.
Although the present invention has been described with reference to the preferred embodiments, it is not intended to be limited thereto. Those skilled in the art can make various changes and modifications without departing from the spirit and scope of the invention. Therefore, the protection scope of the present invention should be defined by the appended claims.
Claims (10)
1. A method of forming a composite material aerofoil framework, comprising:
respectively paving and sticking a plurality of layers of prepreg layers on a prefabricated front beam mould and a prefabricated rear beam mould;
paving and sticking a plurality of layers of prepreg layers on a prefabricated rib mould;
laying a plurality of prefabricated upper limiting pieces on the vacant surfaces of the outermost prepreg layers on the front beam mold at intervals, and simultaneously laying a plurality of prefabricated lower limiting pieces on the vacant surfaces of the outermost prepreg layers on the rear beam mold at intervals; two ends of the rib die are respectively and tightly inserted between the corresponding upper limiting piece and the lower limiting piece;
fixing two ends of a prefabricated positioning clamping plate on the front beam mold and the rear beam mold respectively;
and demolding after hot-pressing curing molding to obtain the composite material wing framework.
2. The method of claim 1, wherein the front spar mold and the back spar mold are both C-shaped molds, and after the plurality of prepreg layers are respectively laid on the prefabricated front spar mold and the prefabricated back spar mold, the method comprises:
and arranging the opening directions of the front beam mould and the rear beam mould to face the prefabricated rib mould.
3. The method of claim 1, wherein the front spar form is a C-type form, and the step of applying the plurality of prepreg layers to the pre-fabricated front spar form and the pre-fabricated back spar form comprises:
arranging a rib mould with the opening direction of the front beam mould deviating from the prefabrication direction;
paving and sticking a plurality of layers of prepreg layers on a prefabricated front beam side plate;
arranging the front beam side plates on two sides of the front beam mould and enabling the vacant surfaces of the prepreg layers on the outermost layers of the front beam side plates to be attached to the vacant surfaces of the prepreg layers on the outermost layers of the front beam mould;
and paving and pasting a plurality of layers of prepreg layers in a structure formed by the empty residual surface of the outermost prepreg layer on the front beam side plate and the residual part of the empty residual surface of the outermost prepreg layer on the front beam mould.
4. The method of claim 3, wherein the step of positioning the side panels on opposite sides of the front spar form and attaching the margins of the outermost prepreg layer to the margins of the outermost prepreg layer comprises:
and aligning the front beam side plate with a waist-shaped hole prefabricated on the upper limiting piece, and penetrating a pin into the front beam side plate so that the front beam side plate has the freedom degree of moving along the direction of the rib mold.
5. The method of claim 1, wherein the back spar form is a C-type form, and the step of applying the plurality of prepreg layers to the pre-fabricated front spar form and back spar form comprises:
arranging the opening direction of the back beam mould to deviate from the prefabricated rib mould;
paving and sticking a plurality of layers of prepreg layers on a prefabricated back beam side plate;
arranging the rear beam side plates on two sides of the rear beam mould and enabling the vacant surfaces of the prepreg layers on the outermost layers of the rear beam mould to be attached to the vacant surfaces of the prepreg layers on the outermost layers of the rear beam mould;
and paving and sticking a plurality of layers of prepreg layers in a structure formed by the spare surface of the outermost prepreg layer on the side plate of the back beam and the rest part of the spare surface of the outermost prepreg layer on the back beam mould.
6. The method of claim 5, wherein the step of attaching the trailing side panels to the trailing side forms with the free surfaces of the outermost prepreg layers of the trailing side forms attached to the free surfaces of the outermost prepreg layers of the trailing side forms comprises:
and aligning the back beam side plate with a waist-shaped hole prefabricated on the lower limiting piece, and penetrating a pin into the waist-shaped hole so that the back beam side plate has the freedom of moving along the direction of the rib mold.
7. A method of forming a composite airframe as defined in claim 1, wherein the applying a plurality of prepreg layers to a prefabricated rib form further comprises:
and reserving the vacant ends of the outermost prepreg layer to prevent the vacant ends from being paved on the rib mould.
8. A method of forming a composite airframe as defined in claim 7, wherein the step of applying the plurality of prepreg layers to the prefabricated rib form further comprises:
and respectively positioning and adhering the spare ends of the outermost prepreg layers on the rib molds to the outermost prepreg layers of the corresponding front beam molds and the rear beam molds.
9. The method of claim 1, wherein the step of applying the plurality of prepreg layers to the prefabricated front and back spar forms comprises:
and spraying a release agent on the surface of the prefabricated front beam mold, the prefabricated rear beam mold, the prefabricated rib mold, the prefabricated upper limiting piece and the prefabricated lower limiting piece, wherein the prepreg layer is paved and adhered, or adhering release cloth.
10. The method of claim 1, wherein the step of demolding after the thermocompression curing to obtain the composite wing frame comprises:
and sequentially placing an isolation film, an air-permeable felt and a vacuum bag on any side face of the front beam mold, the rear beam mold, the upper limiting piece, the lower limiting piece and the rib mold in a filling manner.
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