CA1177459A - Composite structure for joining intersecting structural members of an airframe and the like - Google Patents
Composite structure for joining intersecting structural members of an airframe and the likeInfo
- Publication number
- CA1177459A CA1177459A CA000375674A CA375674A CA1177459A CA 1177459 A CA1177459 A CA 1177459A CA 000375674 A CA000375674 A CA 000375674A CA 375674 A CA375674 A CA 375674A CA 1177459 A CA1177459 A CA 1177459A
- Authority
- CA
- Canada
- Prior art keywords
- fibers
- bundles
- structural
- bundle
- cruciform structure
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired
Links
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
- B64C1/06—Frames; Stringers; Longerons ; Fuselage sections
- B64C1/061—Frames
- B64C1/062—Frames specially adapted to absorb crash loads
-
- 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
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- Aviation & Aerospace Engineering (AREA)
- Moulding By Coating Moulds (AREA)
- Laminated Bodies (AREA)
- Lining Or Joining Of Plastics Or The Like (AREA)
- Woven Fabrics (AREA)
- Processing And Handling Of Plastics And Other Materials For Molding In General (AREA)
Abstract
ABSTRACT OF THE DISCLOSURE
A composite cruciform structure adapted for securely joining two or more intersecting fiber based composite struc-tural members of an airframe and the like, is disclosed. A
first bundle of graphite or other fibers of high strength is intimately interwoven with a second bundle of like fibers, with each bundle of fibers extending in a different direction and being aligned with the general longitudinal axis of a respec-tive first and second composite structural member. A third bundle of fibers is interwoven with both the first and second bundles of fibers to form the cruciform structure. The bundles of fibers are reinforced with a suitable cured resin. The cruciform structure may comprise an integral part of the respective structural members, or may be attached thereto by splicing. The cruciform allows for substantially even distri-bution of structural loads regardless of the direction of the corresponding load force vectors, and is most advantageously utilized in the construction of high strength and low weight airframes.
A composite cruciform structure adapted for securely joining two or more intersecting fiber based composite struc-tural members of an airframe and the like, is disclosed. A
first bundle of graphite or other fibers of high strength is intimately interwoven with a second bundle of like fibers, with each bundle of fibers extending in a different direction and being aligned with the general longitudinal axis of a respec-tive first and second composite structural member. A third bundle of fibers is interwoven with both the first and second bundles of fibers to form the cruciform structure. The bundles of fibers are reinforced with a suitable cured resin. The cruciform structure may comprise an integral part of the respective structural members, or may be attached thereto by splicing. The cruciform allows for substantially even distri-bution of structural loads regardless of the direction of the corresponding load force vectors, and is most advantageously utilized in the construction of high strength and low weight airframes.
Description
1 ¦CO~IPOSITE STRUCTURE FOR JOINING
IINT~P<SI~CTING STRUCTUR~L M~:MBERS
IINT~P<SI~CTING STRUCTUR~L M~:MBERS
2 IOF }~N ~IRF~ME AND THE LIKE
31 . .
¦ BACKGROUND OF THE INVENTION
5 ¦ . lo Field of the Invention ¦ The present invention relates to a ~tructure joining 71 intersecting fiber based composite structural members of 8¦ an airframe and ~he like, and ~o a process for making said ¦ structure. ' 10¦ 2. escri~tion of the Prior Art , 1~¦ Fiber based composite ,materials have been known in ~21 the prior art for a long time. Briefly, such materials co~prise 13¦ a plurality of relatively thin fibers and a reinforcing cured 1~¦ plastic which substantially covers the fibers and holds them 15¦ togethex. Furthermore, it was,recognized in the prior art 16¦ that the structural strength of fiber based composite materials 17¦ is the greatest in the direction of the fibers. Accordingly, l~¦ composite materials have been prepared in the past wherein all 19¦ of the fibers are disposed in one direction parallel to one `' ~¦ another. These type of composite materials are hereinater 21¦ referred to as unidirectional composite materials. ' 22¦ A principal characteristic of unidirectional fiber 23¦ based composite materials iæ ~heir above mentioned anisotropy~
~41 Thus, these materials exhibit relatively great strength to 25¦ withstand forces which are applied substantially in the direction 2~1 f the fibers. ~owever, the load bearing or force withstanding 271 capability of the unidirertional fiber based composite materials 2E~ ¦
29 ' : .
31 . . .. . ~ .
32 ~ .
I .. " ' ' *
.~
j; ~ ! 77459 ¦ against forces which are applied perpendicularly to the dixection 2 ¦ of fibers, is substantially less. This follo~s frp~ the fact
31 . .
¦ BACKGROUND OF THE INVENTION
5 ¦ . lo Field of the Invention ¦ The present invention relates to a ~tructure joining 71 intersecting fiber based composite structural members of 8¦ an airframe and ~he like, and ~o a process for making said ¦ structure. ' 10¦ 2. escri~tion of the Prior Art , 1~¦ Fiber based composite ,materials have been known in ~21 the prior art for a long time. Briefly, such materials co~prise 13¦ a plurality of relatively thin fibers and a reinforcing cured 1~¦ plastic which substantially covers the fibers and holds them 15¦ togethex. Furthermore, it was,recognized in the prior art 16¦ that the structural strength of fiber based composite materials 17¦ is the greatest in the direction of the fibers. Accordingly, l~¦ composite materials have been prepared in the past wherein all 19¦ of the fibers are disposed in one direction parallel to one `' ~¦ another. These type of composite materials are hereinater 21¦ referred to as unidirectional composite materials. ' 22¦ A principal characteristic of unidirectional fiber 23¦ based composite materials iæ ~heir above mentioned anisotropy~
~41 Thus, these materials exhibit relatively great strength to 25¦ withstand forces which are applied substantially in the direction 2~1 f the fibers. ~owever, the load bearing or force withstanding 271 capability of the unidirertional fiber based composite materials 2E~ ¦
29 ' : .
31 . . .. . ~ .
32 ~ .
I .. " ' ' *
.~
j; ~ ! 77459 ¦ against forces which are applied perpendicularly to the dixection 2 ¦ of fibers, is substantially less. This follo~s frp~ the fact
3 ¦ that the p~rallel disposed fibers of ~he composite materials ~ ¦ are held together only by the cured resin. The prior art 51 composite materials or structures are particularly vulnerable I to forces which tend to separate the fibers from one another 71 in a directi~n perpendicular to the layout of the fibers.
81 Nevertheless, composite materials, and particularly plastic 9 reinforced glass fibers (fiberglass) have found several applications in the prior art where a relatively light weight ll and yet strong structural material was desired~
1~ The relatively recent de~elopment of unidirectional 13 composite materials containing graph~te and other fibers 14 of high strength has rendered possible ~le utilization of fiber based composite materials in airframe construction. More 16 particularly, unidirectional composite materials ComprisiDg 17¦ epoxy resin reinforced graphite or other fibers of high 18 ¦ strength are currently used, at least to a limited extent, .
19 ¦ to provide stringer and frame type structural members in 20 ¦aircraft fuselages, and rib and spar type structural members 21 ¦in aircraft wings. - `-`` ~`
22 ¦ A substantial disadvantage of conventional type `23~ ¦airframe construction is tha~ wherever" two str~tural-members -24 ¦intersect one another, it is necessary to provide a cut-away I , . ................. ~ . .. , .......... ... , 25¦ portion in one of the structural members so as to accommodate 26¦ the other structural memberl Cutting away a portion of a ~7¦ structural me~ber, of course, diminishes its load bearing 28~ ` `
`' ~91 . . ` ~ - ` ` ` ```" ` `
30 1 ~ ` `- ` `:
31 1 ~ ` ` ` ` ` `` `
. 321 . ~ ~
I
' ~ 77~59 capacity. Consequently, in conventional airframe construction it is necessary to provide additional reinforcing members to fasten the two structural members to one another at their point of intersection.
Although the state-of-the art application of high strength, relatively light wei~ht unidirectional composite materials has offered certain advantages, it has not, up to the present invention, resulted in an altogether different highly advantageous method or structure for joining two intersection structural members to one another.
Accordingly, there is a substantial need for the novel and unique high strength structure and method of the present invention which provides for a high strength junction of two or more substantially intersection composite structural members.
SUM~RY OF THE INVENTIO~
It is an object of the present invention to provide a high strength interconnection or junction of two or more structural members in an airframe construction and the like wherein each structural member comprises composite material.
In accordance with the present invention there is provided a composite cruciform structure for joining together intersecting structural members of an aircraft construction and the like, said cruciform structure comprising a first bundle of fibers interwoven with a second bundle of fibers and additional fibers interwoven with each of said first and second bundles of fibers in a direction substantially perpendicular to the fibers of said first and second bundles to form said cruciform structure ~. .
~ - ' ' ~ . . , ~ ~77~59 the interwoven fibers of said first and second bundles intersecting one another, said fibers having a reinforcing resin on the surface thereof to provide a strong structural joinder of said structural members without a substantial break in the continuity of the fibers and to provide substantially even load bearing and force transmitting capability in a plurality of directions with respect to said intersection of said structural members.
Also in accordance with the invention there is provided a composite cruciform structure for joining together intersecting structural membes of an aircraft construction and the like, said cruciform structure comprising a first bundle of fibers interwoven with a second bundle of fibers, and additional bundles of fibers disposed substantially perpendicular to said .
first and second bundles of fibers and interwoven with both said first and second bundles to form said cruciform structure, the interwoven fibers of said first and second bundles intersecting one another, said fibers being selected from the group consisting of carbon, graphite, Kevlar, glass and boron fibers, each of said bundles of fibers comprising a plurality of layers of fibers, said bundles being provided with a reinforcing resin on the surface thereof, to provide a strong structural joinder of said structural members without a substantial break in the continuity of said fibers and to provide substantially even load bearing and force transmitting capability in a plurality of directions with respect to said intersection of said structural members.
: `~
. ~
i .! 77~59 :.
~ urther in accordance with the invention there is provided a process for strongly joining a first substantially elongated composite st~uctural member and a second substantially elongated composite structural member, said structural members extendin~ in different directions relative to one another, the process comprising the steps of:
interweaving a first bundle of fibers comprising a portion of the first structural member with a second bundle of fibers comprising a portion of the second structural member, the interwoven fibers of said first and second bundles intersecting one another, said first and second bundles of fibers substantially extending in the same direction as the respective elongated structural members;
interweaving additional fibers with said first and second bundles disposed substantially perpendicular relative to said first and second bundles of fibers;
applying a suitable reinforcing resin to said fibers including the area wherein the bundles are interwoven; and curing said resin.
The objects and features of the present invention are set forth in the appended claims. The present invention may be best understood by reference to the following description, ~ , , ::
.
¦ takcn in connection with the drawings in which like numerals 21 indic~te like parts.
¦ BRI~F DESCRIPTION OF T~E DI~AWI21GS
Fi~ure 1 is a perspective view of a portion of a fuselage of an aircraft wherein a stringer and a frame member ... ... _ - . . , . .... . . .. , _ .... , . . . . ~ _ .
6 are joined to oneanother in accordance with the prior art;
7 Figure 2 is a perspective view of a portion of a 81 wing of an aircraft wherein a spar and a ri~ member are joined gl to one another in accordance with the prior art;
10¦ Figure 3 is a schematic perspecti~e~ view illustxating 11¦ the principle of joining two intersecting fiber based uni-1~¦ directional composite structural members to one another in 13 I accordance with the present invention;
1~ ¦ Figure 4 is a schematic, perspectîve view further 15 ¦ illustrating the principle of joining two intersecting fiber 16 ¦ based unidirectional composite structural members to one another 17 ¦ in accordance with the present invention;
18 ¦ Figure 5 is a schematic, partial perspective vie~
19 ¦ of an airplane fu~elage with .par~:of ~he ski~ bein~ibroken.away, 20 ¦ the view.schematically illustrating frame and stringer members 21 ¦ of the fuselage ~eing joined ta one another in accordance .
22¦ with the present invention;
231 Figure 6 is a schematic vi~w of a wing of an airplane ~41 with part of the skin of the wing being broken away, the view 2~¦ schematically illus~ratîng st~uctur.a~ members o~ the wing 26¦ being joined to one another in accordance with the pxesent 2271 invention, and 291 . '`' 301 . . .
31 . .
32 . -6-~ ' 77~59 Figure 7 is a schematic perspective view showing six composite cruciform structural members being spliced to one another, each composite structural member including a substantially cruciform shaped junction with another struct-ural member, said junctions being formed in accordance withthe present invention.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
The following specification taken in conjunction with the drawings sets forth the preferred embodiment of the present invention. The embodiments of the invention disclosed herein are the best modes contemplated by the inventor for carrying out his invention in a commercial environment, although it should be understood that various modifications can be accomplished within the parameters of the present ; 15 invention.
Figures 1 and 2 of the drawing figures respectively depict aircarft fuselage and wing construction in accordance with the prior art. In Figure 1 a structural member incor-porating a cut-away portion illustrates, by way of example, a fore-and-aft positioned stringer of an aircraft fuselage, while another structural member illustrates a laterally extending frame member of the fuselage. The necessity of providing a cut-away portion in one of the structural members whenever two structural members "intersect" one another, or "attempt to occupy the same space", is not however limited to airframe construction and a similar problem is encountered in the construction of boats, vehicle frames, buildings, etc.
~ - 7 ~ 77459 It is readily apparent to those skilled in the art that providing appropriately positioned cut-away portions and mounting the necessary reinforcing members or clips is time consuming and si~nificantly contributes to the overall construction cost.
With particular reference to Figure 1, it is noted, that in the state-of-the-art airframe and the like composite structures, a cut-away portion is provided in one of the structural members as in conventional metal structures.
Attachment of the additional reinforcing members or "clips"
may be accomplished, however, by using a structural adhesive resin instead of welding, rivets, screws or bolts and nuts of a conventional metal construction. These figures are explained further below in comparison with fuselage and wing construction in accordance with the present invention, which are shown on the rest of the drawing figures. The schematic view of Figure 3 discloses a three dimensional cruciform shaped structure or joining member 12 which comprises a junction of two structural members in accordance with the present invention. Each structural member is made of a fiber based unidirectional composite material and the manner of joining the fiber based composite structural members to one another comprises a principal novel feature of the present invention.
- 7a ~: ;
- "
-- I ' ! 77459 1 ~s it was briefly described in the introductory 2 sections of the present application, a fiber based unidixectional 3 composite material includes a plurality of relatively thin
81 Nevertheless, composite materials, and particularly plastic 9 reinforced glass fibers (fiberglass) have found several applications in the prior art where a relatively light weight ll and yet strong structural material was desired~
1~ The relatively recent de~elopment of unidirectional 13 composite materials containing graph~te and other fibers 14 of high strength has rendered possible ~le utilization of fiber based composite materials in airframe construction. More 16 particularly, unidirectional composite materials ComprisiDg 17¦ epoxy resin reinforced graphite or other fibers of high 18 ¦ strength are currently used, at least to a limited extent, .
19 ¦ to provide stringer and frame type structural members in 20 ¦aircraft fuselages, and rib and spar type structural members 21 ¦in aircraft wings. - `-`` ~`
22 ¦ A substantial disadvantage of conventional type `23~ ¦airframe construction is tha~ wherever" two str~tural-members -24 ¦intersect one another, it is necessary to provide a cut-away I , . ................. ~ . .. , .......... ... , 25¦ portion in one of the structural members so as to accommodate 26¦ the other structural memberl Cutting away a portion of a ~7¦ structural me~ber, of course, diminishes its load bearing 28~ ` `
`' ~91 . . ` ~ - ` ` ` ```" ` `
30 1 ~ ` `- ` `:
31 1 ~ ` ` ` ` ` `` `
. 321 . ~ ~
I
' ~ 77~59 capacity. Consequently, in conventional airframe construction it is necessary to provide additional reinforcing members to fasten the two structural members to one another at their point of intersection.
Although the state-of-the art application of high strength, relatively light wei~ht unidirectional composite materials has offered certain advantages, it has not, up to the present invention, resulted in an altogether different highly advantageous method or structure for joining two intersection structural members to one another.
Accordingly, there is a substantial need for the novel and unique high strength structure and method of the present invention which provides for a high strength junction of two or more substantially intersection composite structural members.
SUM~RY OF THE INVENTIO~
It is an object of the present invention to provide a high strength interconnection or junction of two or more structural members in an airframe construction and the like wherein each structural member comprises composite material.
In accordance with the present invention there is provided a composite cruciform structure for joining together intersecting structural members of an aircraft construction and the like, said cruciform structure comprising a first bundle of fibers interwoven with a second bundle of fibers and additional fibers interwoven with each of said first and second bundles of fibers in a direction substantially perpendicular to the fibers of said first and second bundles to form said cruciform structure ~. .
~ - ' ' ~ . . , ~ ~77~59 the interwoven fibers of said first and second bundles intersecting one another, said fibers having a reinforcing resin on the surface thereof to provide a strong structural joinder of said structural members without a substantial break in the continuity of the fibers and to provide substantially even load bearing and force transmitting capability in a plurality of directions with respect to said intersection of said structural members.
Also in accordance with the invention there is provided a composite cruciform structure for joining together intersecting structural membes of an aircraft construction and the like, said cruciform structure comprising a first bundle of fibers interwoven with a second bundle of fibers, and additional bundles of fibers disposed substantially perpendicular to said .
first and second bundles of fibers and interwoven with both said first and second bundles to form said cruciform structure, the interwoven fibers of said first and second bundles intersecting one another, said fibers being selected from the group consisting of carbon, graphite, Kevlar, glass and boron fibers, each of said bundles of fibers comprising a plurality of layers of fibers, said bundles being provided with a reinforcing resin on the surface thereof, to provide a strong structural joinder of said structural members without a substantial break in the continuity of said fibers and to provide substantially even load bearing and force transmitting capability in a plurality of directions with respect to said intersection of said structural members.
: `~
. ~
i .! 77~59 :.
~ urther in accordance with the invention there is provided a process for strongly joining a first substantially elongated composite st~uctural member and a second substantially elongated composite structural member, said structural members extendin~ in different directions relative to one another, the process comprising the steps of:
interweaving a first bundle of fibers comprising a portion of the first structural member with a second bundle of fibers comprising a portion of the second structural member, the interwoven fibers of said first and second bundles intersecting one another, said first and second bundles of fibers substantially extending in the same direction as the respective elongated structural members;
interweaving additional fibers with said first and second bundles disposed substantially perpendicular relative to said first and second bundles of fibers;
applying a suitable reinforcing resin to said fibers including the area wherein the bundles are interwoven; and curing said resin.
The objects and features of the present invention are set forth in the appended claims. The present invention may be best understood by reference to the following description, ~ , , ::
.
¦ takcn in connection with the drawings in which like numerals 21 indic~te like parts.
¦ BRI~F DESCRIPTION OF T~E DI~AWI21GS
Fi~ure 1 is a perspective view of a portion of a fuselage of an aircraft wherein a stringer and a frame member ... ... _ - . . , . .... . . .. , _ .... , . . . . ~ _ .
6 are joined to oneanother in accordance with the prior art;
7 Figure 2 is a perspective view of a portion of a 81 wing of an aircraft wherein a spar and a ri~ member are joined gl to one another in accordance with the prior art;
10¦ Figure 3 is a schematic perspecti~e~ view illustxating 11¦ the principle of joining two intersecting fiber based uni-1~¦ directional composite structural members to one another in 13 I accordance with the present invention;
1~ ¦ Figure 4 is a schematic, perspectîve view further 15 ¦ illustrating the principle of joining two intersecting fiber 16 ¦ based unidirectional composite structural members to one another 17 ¦ in accordance with the present invention;
18 ¦ Figure 5 is a schematic, partial perspective vie~
19 ¦ of an airplane fu~elage with .par~:of ~he ski~ bein~ibroken.away, 20 ¦ the view.schematically illustrating frame and stringer members 21 ¦ of the fuselage ~eing joined ta one another in accordance .
22¦ with the present invention;
231 Figure 6 is a schematic vi~w of a wing of an airplane ~41 with part of the skin of the wing being broken away, the view 2~¦ schematically illus~ratîng st~uctur.a~ members o~ the wing 26¦ being joined to one another in accordance with the pxesent 2271 invention, and 291 . '`' 301 . . .
31 . .
32 . -6-~ ' 77~59 Figure 7 is a schematic perspective view showing six composite cruciform structural members being spliced to one another, each composite structural member including a substantially cruciform shaped junction with another struct-ural member, said junctions being formed in accordance withthe present invention.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
The following specification taken in conjunction with the drawings sets forth the preferred embodiment of the present invention. The embodiments of the invention disclosed herein are the best modes contemplated by the inventor for carrying out his invention in a commercial environment, although it should be understood that various modifications can be accomplished within the parameters of the present ; 15 invention.
Figures 1 and 2 of the drawing figures respectively depict aircarft fuselage and wing construction in accordance with the prior art. In Figure 1 a structural member incor-porating a cut-away portion illustrates, by way of example, a fore-and-aft positioned stringer of an aircraft fuselage, while another structural member illustrates a laterally extending frame member of the fuselage. The necessity of providing a cut-away portion in one of the structural members whenever two structural members "intersect" one another, or "attempt to occupy the same space", is not however limited to airframe construction and a similar problem is encountered in the construction of boats, vehicle frames, buildings, etc.
~ - 7 ~ 77459 It is readily apparent to those skilled in the art that providing appropriately positioned cut-away portions and mounting the necessary reinforcing members or clips is time consuming and si~nificantly contributes to the overall construction cost.
With particular reference to Figure 1, it is noted, that in the state-of-the-art airframe and the like composite structures, a cut-away portion is provided in one of the structural members as in conventional metal structures.
Attachment of the additional reinforcing members or "clips"
may be accomplished, however, by using a structural adhesive resin instead of welding, rivets, screws or bolts and nuts of a conventional metal construction. These figures are explained further below in comparison with fuselage and wing construction in accordance with the present invention, which are shown on the rest of the drawing figures. The schematic view of Figure 3 discloses a three dimensional cruciform shaped structure or joining member 12 which comprises a junction of two structural members in accordance with the present invention. Each structural member is made of a fiber based unidirectional composite material and the manner of joining the fiber based composite structural members to one another comprises a principal novel feature of the present invention.
- 7a ~: ;
- "
-- I ' ! 77459 1 ~s it was briefly described in the introductory 2 sections of the present application, a fiber based unidixectional 3 composite material includes a plurality of relatively thin
4 fibers which are disposed lengthwise, i.e. parallel relative to one another, and a suitable cured resin which substantially . _ . . . . .. . . . .. ... .. . . . ..... . ... . _ . .. . . . _ . _ . ._ ~ .
6 covers the fibers and holds them ~ogether. The state-of-the-71 art in the design ana manufacture of fiber based composite 81 materials is relatively advanced a~ the present. Therefore, 9 the following concise, general description of fiber based composite materials and the process of their manufacture is 11 intended solely ~or the purpose of facilitating the understanding 12 I of the present invention and for emphasizing and illuminating 13 ¦ the novel features thereof.
14 ¦ Briefly, suitable fibers for the cons-truction of 15 ¦ strong unidirectional com~osite materials are glass, graphite, 16 ¦ carbon, Xevlar and boron fibers. (Kevlar is a trademark of the 17 ¦ E.I. Dupont Company and is used to designate the source of 18 ¦ certain fiber material). The present învention may be practiced 19 ¦ with either one of the above mentio~ed and with other fiber 20 ¦ materials, although the use of glass fibers for the construc~ion 2~ ¦ of heavy duty aircraft components is generally not preferred.
22¦ Generally speaking,~fibers and particularly graphite fibers 23¦ used for construction of various structural members have a 2~1 diameter of approximately 3 mils, and each of said fibers is 251 itself a combination of a plurality of thinner subfibers.
2~1 In order to form a structural member of a predetexmined 271 dimension a bundle of fiber~ is positioned in such a manner that 281 .
~ ~77~5~
¦ the lon~itudinal axes of the fibers are disposed parallel to 2¦ one another. A suitable organic resin, which has not yet 31 reached its fully polymerized or fùlly cured state, is then 4 applied to the fibers. Subsequently, the resin is fully poly-merized or cured under exposure to heat and in some instances 6 high pressure. An important fac~or determining the selection _ 7 of the proper resin for a given application is the nature of the 8 fibers themselves. A person possessing average skill in ~he 9 fiber based composite materials manufacturing arts is able to select the proper resin for a fiber o a given composition.
11 Generally speaking, epoxy based resins are utilized in conjunction 12 with graphite and other high strength fibers, although the 13 present invention may be practiced with any type of fiber and 14 resin combinaticn. Usually the final polymerization or curing step of the composite material is conducted at 2S0-350F ~or 16 1/~ - 3 hours. The exact parameters of the aforementioned curing 17 step are, of course, dependent on the exact nature of the fibers 18 and on the chemical properties of the resinous binding material.
19 Again, the scope of the present invention is not limited in any way by the physical parameters of the curing step.
21 Often, the bundle of the parallel disposed ~ibers 22 is formed in the shape of a relatively thin band or tape~ and 23 several of the bands or tapes may be joined together in the 24 curing step to form a structura~ membex~ An impoxtant characteristic and major advantage ~f the fiber based composite 26 materials~ and particularly of graphite fiber based composite 278 materials is that they provide very high str~ngth in ~he direction 29 `' 32 _9_ -~ l 1¦ of the fibers at a relatively low weight.
2 The graphite fiber based composite materials are 3 particularly preferred in the present invention for the 41 construction of aircraft structural components, because these materials provide a structural integrity as high or higher _.
6¦ than that of steel while ~he weight of these materials is con-71 siderably less than that o~ steel. On the other hand, a s~rious 81 disadvantage of fiber based composite materials lies in their 91 anisotropic behavior; in other words these materials exhibit 10¦ much less structural integrity against forces which are not applied in the direction of fibers. As it is described below, 12¦ this disadvantage is overcome by ~he present invention precisely 13¦ at the points of intersection of two or more composite structural 14¦ members wherein the disadvantage created by the anisotro~y is ~5¦ the least tolerable~
16¦ Because of the well established importance of light 17¦ weight and great structural integrity of materials utilized 18¦ in airframe canstruction, the ensuing description is principally 1~ ¦ directed towards a description of the applic~ion of the present 2n ¦ invention in airframe construction. Furthermore, the fibers ~1¦ utilized in the practice of the present in~entio~ will generally 22¦ be referred to as ~raphlte- fibers. Nevertheless, i~ should ~31 be expressly understood that the scope of the present invention 2~ is not limited either to its application in airframe ~ construction nor to the use of graphite f~bers only. -26 Referring again to ~igure 3, the basic principle ~7 of the novel structure 12 of the present in~ention is explained 2~
32 ~l0-.1 .. ' ' '.
I ~ in detal . On Figure 3 the arrows respectively marked X, Y, and 2 Z indicate three mutually perpendicular axes situated similarly 3 ¦ to the axes of a three dimensional coordinate system. In , ~ ¦ accordance with the present invention, a first substantially ,
6 covers the fibers and holds them ~ogether. The state-of-the-71 art in the design ana manufacture of fiber based composite 81 materials is relatively advanced a~ the present. Therefore, 9 the following concise, general description of fiber based composite materials and the process of their manufacture is 11 intended solely ~or the purpose of facilitating the understanding 12 I of the present invention and for emphasizing and illuminating 13 ¦ the novel features thereof.
14 ¦ Briefly, suitable fibers for the cons-truction of 15 ¦ strong unidirectional com~osite materials are glass, graphite, 16 ¦ carbon, Xevlar and boron fibers. (Kevlar is a trademark of the 17 ¦ E.I. Dupont Company and is used to designate the source of 18 ¦ certain fiber material). The present învention may be practiced 19 ¦ with either one of the above mentio~ed and with other fiber 20 ¦ materials, although the use of glass fibers for the construc~ion 2~ ¦ of heavy duty aircraft components is generally not preferred.
22¦ Generally speaking,~fibers and particularly graphite fibers 23¦ used for construction of various structural members have a 2~1 diameter of approximately 3 mils, and each of said fibers is 251 itself a combination of a plurality of thinner subfibers.
2~1 In order to form a structural member of a predetexmined 271 dimension a bundle of fiber~ is positioned in such a manner that 281 .
~ ~77~5~
¦ the lon~itudinal axes of the fibers are disposed parallel to 2¦ one another. A suitable organic resin, which has not yet 31 reached its fully polymerized or fùlly cured state, is then 4 applied to the fibers. Subsequently, the resin is fully poly-merized or cured under exposure to heat and in some instances 6 high pressure. An important fac~or determining the selection _ 7 of the proper resin for a given application is the nature of the 8 fibers themselves. A person possessing average skill in ~he 9 fiber based composite materials manufacturing arts is able to select the proper resin for a fiber o a given composition.
11 Generally speaking, epoxy based resins are utilized in conjunction 12 with graphite and other high strength fibers, although the 13 present invention may be practiced with any type of fiber and 14 resin combinaticn. Usually the final polymerization or curing step of the composite material is conducted at 2S0-350F ~or 16 1/~ - 3 hours. The exact parameters of the aforementioned curing 17 step are, of course, dependent on the exact nature of the fibers 18 and on the chemical properties of the resinous binding material.
19 Again, the scope of the present invention is not limited in any way by the physical parameters of the curing step.
21 Often, the bundle of the parallel disposed ~ibers 22 is formed in the shape of a relatively thin band or tape~ and 23 several of the bands or tapes may be joined together in the 24 curing step to form a structura~ membex~ An impoxtant characteristic and major advantage ~f the fiber based composite 26 materials~ and particularly of graphite fiber based composite 278 materials is that they provide very high str~ngth in ~he direction 29 `' 32 _9_ -~ l 1¦ of the fibers at a relatively low weight.
2 The graphite fiber based composite materials are 3 particularly preferred in the present invention for the 41 construction of aircraft structural components, because these materials provide a structural integrity as high or higher _.
6¦ than that of steel while ~he weight of these materials is con-71 siderably less than that o~ steel. On the other hand, a s~rious 81 disadvantage of fiber based composite materials lies in their 91 anisotropic behavior; in other words these materials exhibit 10¦ much less structural integrity against forces which are not applied in the direction of fibers. As it is described below, 12¦ this disadvantage is overcome by ~he present invention precisely 13¦ at the points of intersection of two or more composite structural 14¦ members wherein the disadvantage created by the anisotro~y is ~5¦ the least tolerable~
16¦ Because of the well established importance of light 17¦ weight and great structural integrity of materials utilized 18¦ in airframe canstruction, the ensuing description is principally 1~ ¦ directed towards a description of the applic~ion of the present 2n ¦ invention in airframe construction. Furthermore, the fibers ~1¦ utilized in the practice of the present in~entio~ will generally 22¦ be referred to as ~raphlte- fibers. Nevertheless, i~ should ~31 be expressly understood that the scope of the present invention 2~ is not limited either to its application in airframe ~ construction nor to the use of graphite f~bers only. -26 Referring again to ~igure 3, the basic principle ~7 of the novel structure 12 of the present in~ention is explained 2~
32 ~l0-.1 .. ' ' '.
I ~ in detal . On Figure 3 the arrows respectively marked X, Y, and 2 Z indicate three mutually perpendicular axes situated similarly 3 ¦ to the axes of a three dimensional coordinate system. In , ~ ¦ accordance with the present invention, a first substantially ,
5¦ elongated structural me~ber 14 is disposed substantially along ., --~ .. I . . . . . , . ................ . .. .. . ... . , .. , .... . .. ~ _... ..
1 61 the X axis, and a second substantially elongated structural member 16 is disposed substantially along the Y axis. Although 8 the first and second structural members 14 and 16 per se are 9 not shown on Figure 3, a first bundle of fibers 18 corresponding 10 ,to the first structural member 14 and a second bundle of fibers 11 19 corresponding to ~he second structural member 16, are clearly 12 illustrated on this figure. The first structural member 14 13 may be a stringer in a fuselage 20 of an aircraft 22, and 14 the second structural member 16 may be a frame member in khe fuselage 20 of the aircraft 22 as is illustrated in Figure 5 16 A stringer 24 and a frame member 26 of an aircraft 17 fuselage is shown on Figure 1 which depicts the prior art. It 13 is readily discernible on Figure 1 that where the stringer ;
19 24 and frame members ~6 intersect one another, a cut-away portion has been provided in the stringer 24 so as to accommodate , 21 the frame member 26; In ord,er not to lose structural strength ' 22 and to` provide for transmission of various forces from the 23 stringer 24 and frame m~mbers 26 to one another, a plurality of 24 reinforcing or clip members 28 were provided in'the prior art.
~5 These were attached to the stringer 24 and to the frame member _, 26 26 by welding or by other conventional modes of attachment.
27 Skin attached to the s~ringer 24 and frame member~ 26 is indi-2~ cated by the reference n~meral 30 on Figure 1 3~ ' , ,31 ' ' ', . ' ,''.
3 ~ 77A~59 In the novel structure shown in Figure 3, the first bundles of fibers 18, corresponding to the first structural member 14 and hence to frame member 26 of Figure 1, is interwoven with a second bundle of fibers 19 corresponding to 5 the second structural member 16 (and stringer 24) of Figure 1.
To form the cruciform structure of the present invention, fibers 32 of third or additional bundles are intimately interwoven with fibers of both the first and second bundles 18 and 19. The fibers of bundle 32 are disposed in the direction of the z-axis 10 and thus are substantially perpendicular to the general longitudinal axes of the fibers of bundles 18 and 19. It will thus be appreciated that the interwoven bundles of fibers are in the form of a sheet or layer of woven fibers laying in the x and z axes and an intersecting sheet or layer of woven fibers laying 15 in the y and z axes, thereby forming the desired cruciform structure.
In a preferred embodiment of the present invention, interweaving of the fibers is accomplished in such a manner that each fiber of bundle 18 running in the direction of the x-axis 20 is positioned between two fibers of bundle 32 extending in the directon of the z-axis. This is also true with regard to the fibers of bundle 19, that is, each fiber of bundle 19 is likewise positioned between two fibers of bundle 32.
1~
~J ~ ld~' -....
:
- . , ,:~
.
' ~
~ ~ ~77~59 1 ¦ A~ter having interwoven ~he fibers of the first, I s~cond and third bundles 1~, 19 and 32, a suitable pre-3 polymcrized resin (not shown) is applied to the structure 12.
41 Su~sequently the resin is cured by heat according to standard 51 practice in the ax~. In this regard it is noted that for the I sake of clear illustration of the spatial arrangemen~ of the ~ ~;
71 fibers, the resin applied to the ~ibers has been omitted from ~ 81 the drawing figures. Furthermore, it is emphasized ~hat the 91 drawing figures and particularl~ Figures 3 and 4 are merely 10¦ schematic, and the actual ~umber o~ fibers in each of the bundles 11¦ 18, 19 and 32 is very large, as is in fiher based composite 12¦ structural members of the prior art.
13¦ Figure 4 represents a schematic view of an embodi-14¦ ment of the cruci~orm shaped structure 34 of the present in-vention wherein each bundle of fibers la, 19 and 32 comprises 16 ~o layers of fibers. A first and a second layer of fibers in 17 the bundle are provided for the sake of illustration with the 18 respective reference numerals 36 and 38. The layers of the fibers 19 o~ the several bundles are interwoven with one another in a 20¦ manner similar to the interweaving of ~he single layers of fibers `21¦ as is shown on Figure 3. Each bundle 18, 19 and 32 is respectively 22¦ disposed substantially in the direction of the respective X, Y
231 an~ z axes, although it should be agai~understood:that~thè novel 2~¦ structure of the present invention may also be constructed 251 in such a manner that the respective bundles of fibers and 26¦ h~nce the respective structural members 14 and 16 are not at 27~ a 90 angle relative to one another.
2g I .
~ ! 77a.5 9 1 In the actual practice of the present invention i~ is 2 often necessary`to provide multiple layers of fibers in each 3 bundle in order t~ obtain a junction of the structural members 4 which is su~ficiently strong for incorporation in an aircraft.
51 Each layer of fiber ig relatively thin as compared to its length . _ _ .. . . . . ....................... . . . . . . ...... .... ... . . . .... _ _
1 61 the X axis, and a second substantially elongated structural member 16 is disposed substantially along the Y axis. Although 8 the first and second structural members 14 and 16 per se are 9 not shown on Figure 3, a first bundle of fibers 18 corresponding 10 ,to the first structural member 14 and a second bundle of fibers 11 19 corresponding to ~he second structural member 16, are clearly 12 illustrated on this figure. The first structural member 14 13 may be a stringer in a fuselage 20 of an aircraft 22, and 14 the second structural member 16 may be a frame member in khe fuselage 20 of the aircraft 22 as is illustrated in Figure 5 16 A stringer 24 and a frame member 26 of an aircraft 17 fuselage is shown on Figure 1 which depicts the prior art. It 13 is readily discernible on Figure 1 that where the stringer ;
19 24 and frame members ~6 intersect one another, a cut-away portion has been provided in the stringer 24 so as to accommodate , 21 the frame member 26; In ord,er not to lose structural strength ' 22 and to` provide for transmission of various forces from the 23 stringer 24 and frame m~mbers 26 to one another, a plurality of 24 reinforcing or clip members 28 were provided in'the prior art.
~5 These were attached to the stringer 24 and to the frame member _, 26 26 by welding or by other conventional modes of attachment.
27 Skin attached to the s~ringer 24 and frame member~ 26 is indi-2~ cated by the reference n~meral 30 on Figure 1 3~ ' , ,31 ' ' ', . ' ,''.
3 ~ 77A~59 In the novel structure shown in Figure 3, the first bundles of fibers 18, corresponding to the first structural member 14 and hence to frame member 26 of Figure 1, is interwoven with a second bundle of fibers 19 corresponding to 5 the second structural member 16 (and stringer 24) of Figure 1.
To form the cruciform structure of the present invention, fibers 32 of third or additional bundles are intimately interwoven with fibers of both the first and second bundles 18 and 19. The fibers of bundle 32 are disposed in the direction of the z-axis 10 and thus are substantially perpendicular to the general longitudinal axes of the fibers of bundles 18 and 19. It will thus be appreciated that the interwoven bundles of fibers are in the form of a sheet or layer of woven fibers laying in the x and z axes and an intersecting sheet or layer of woven fibers laying 15 in the y and z axes, thereby forming the desired cruciform structure.
In a preferred embodiment of the present invention, interweaving of the fibers is accomplished in such a manner that each fiber of bundle 18 running in the direction of the x-axis 20 is positioned between two fibers of bundle 32 extending in the directon of the z-axis. This is also true with regard to the fibers of bundle 19, that is, each fiber of bundle 19 is likewise positioned between two fibers of bundle 32.
1~
~J ~ ld~' -....
:
- . , ,:~
.
' ~
~ ~ ~77~59 1 ¦ A~ter having interwoven ~he fibers of the first, I s~cond and third bundles 1~, 19 and 32, a suitable pre-3 polymcrized resin (not shown) is applied to the structure 12.
41 Su~sequently the resin is cured by heat according to standard 51 practice in the ax~. In this regard it is noted that for the I sake of clear illustration of the spatial arrangemen~ of the ~ ~;
71 fibers, the resin applied to the ~ibers has been omitted from ~ 81 the drawing figures. Furthermore, it is emphasized ~hat the 91 drawing figures and particularl~ Figures 3 and 4 are merely 10¦ schematic, and the actual ~umber o~ fibers in each of the bundles 11¦ 18, 19 and 32 is very large, as is in fiher based composite 12¦ structural members of the prior art.
13¦ Figure 4 represents a schematic view of an embodi-14¦ ment of the cruci~orm shaped structure 34 of the present in-vention wherein each bundle of fibers la, 19 and 32 comprises 16 ~o layers of fibers. A first and a second layer of fibers in 17 the bundle are provided for the sake of illustration with the 18 respective reference numerals 36 and 38. The layers of the fibers 19 o~ the several bundles are interwoven with one another in a 20¦ manner similar to the interweaving of ~he single layers of fibers `21¦ as is shown on Figure 3. Each bundle 18, 19 and 32 is respectively 22¦ disposed substantially in the direction of the respective X, Y
231 an~ z axes, although it should be agai~understood:that~thè novel 2~¦ structure of the present invention may also be constructed 251 in such a manner that the respective bundles of fibers and 26¦ h~nce the respective structural members 14 and 16 are not at 27~ a 90 angle relative to one another.
2g I .
~ ! 77a.5 9 1 In the actual practice of the present invention i~ is 2 often necessary`to provide multiple layers of fibers in each 3 bundle in order t~ obtain a junction of the structural members 4 which is su~ficiently strong for incorporation in an aircraft.
51 Each layer of fiber ig relatively thin as compared to its length . _ _ .. . . . . ....................... . . . . . . ...... .... ... . . . .... _ _
6 and width. Therefore, the layers are referred to as two dimen-
7 ¦sional layers. Actual dimensions of the structural members are ~ ¦detennined by the par~icular engineering requriements of the 9 ¦aircraft frame or other structure. In certain embodiments 10 ¦one structural member may comprise a ~ anti~ lesser number 11 ¦of layers of fibers than a second structural member which is 12 ¦interwoven therewith. -13¦ Referring now to Figure ~ which schematically depicts '41 a wing 4~ c~nstruction~in accordance with the prior art, it is 15¦ noted that the first and secon~ structural members 14 and 16 16 of Figures 3 and 4 may also correspond respectively to a wing 17 spar member 42 and to a wing rib member 44O Skin ~6 of the wing 18 ~h; may be attached to the novel composite wing Structure by 19 conventional means or ~y the use of a structural adhesive 20 plastic. The use of structural adhesive plastic is well 21 ~stablished in the arts and need not be described here in detail~
22 Referring now to Figures 5 and 6, an aircraft fuselage 23 20 and a wing 46 are schematically shown wherein intersecting 24 frame 26 and stringer 24 members and intersecting wing spar 2~ ~2 and r~b n~mbers 44 are respectively constructed in accordance 2~ with the present invention. The schematic drawin~ of Figure 5 27 and 6 re~eal that due t~ the novel mode of construction, no 28 cut-away portion is provided where these members intersect each 2~ other. ~his is, of coùrse in sharp contrast wi~h th~ pr~or art ~2 -14-~ '' .
~ ~ !77459 1 ¦ fuselage and wing constructions which are illustrated in ~igure~
2 ~ 1 and 2.
3 ¦ It is an additional aspect and add;t~onal advant~ge 4 ¦ of the present invention that the cruoiform shaped structures ¦ cxempli~ied in Figures 3 and 4 as 12 and 34 may be provided _ ___ _ 6 ¦ in a preformed shape prior to assembly into an alr~rame or like 71 str~c~ure. ~ is standard practice in the art to manufacture 81 fiber based composite materials in the shape of a woven fabric 9¦ or a unidirectional tape which already contains the binding 10¦ organic resin in a suitable prepolymerized form. These materials 11¦ are routinely referred to in the ar~ as pre-impregnated or 12¦ "prepreg'~ materials. Because final curing of the binding resin 13¦ does not occur unless the resin is subjected to heat, the pre-14¦ impregnated composite materials usually maintain their uncured 15¦ state for a prolonged period of time particularly if they are 16¦ kept at lower than ambient temperature.
17¦ Thus, it is possible, in accordance with the present 18¦ in~ention to manufac~ure several portions of an airframe and 19¦ the like from composite materials in a preimpregnated state,.
201 As an example,`~igure 7 schematically illustrates three portions 21¦ 48 of a stringer 24 of a fuselage 20, with each portion 48 22¦ already ha~iny two respective portions 52 of a ~rame member 26 231 attached th~reto, by the above described interwoven composite 241 structure. Each of ~hese portions of the stringer 24 and of the frame 26 are in a preimpregnated state.
26 They are splice~ to one anot~er duxing the a sembly o the 27 airframe by the use o~ s~ructural adhesive plastic (not 28 shown) an y conventlonal splice plates 54 sr~ratirally illrst~ated 31~ -15-.
_ . .,, ~
3 ~ 77~L5~
~- , 1 Fi~ure 7. Subsequent to splicing, the ent~re airframe or a 2 suitabl~ selected part thereof is finally cured ~n an autoclave 3 or oven (not shown)~ In ~his final curing step, requisite 4 curincJ o~ the structurial adhesive resin may also occur.
What has been described above is a no~el structure _ ____ ~ for strongly joining intersecting structural members made of 7 fiber based unidirectional composite materials. The novel
22 Referring now to Figures 5 and 6, an aircraft fuselage 23 20 and a wing 46 are schematically shown wherein intersecting 24 frame 26 and stringer 24 members and intersecting wing spar 2~ ~2 and r~b n~mbers 44 are respectively constructed in accordance 2~ with the present invention. The schematic drawin~ of Figure 5 27 and 6 re~eal that due t~ the novel mode of construction, no 28 cut-away portion is provided where these members intersect each 2~ other. ~his is, of coùrse in sharp contrast wi~h th~ pr~or art ~2 -14-~ '' .
~ ~ !77459 1 ¦ fuselage and wing constructions which are illustrated in ~igure~
2 ~ 1 and 2.
3 ¦ It is an additional aspect and add;t~onal advant~ge 4 ¦ of the present invention that the cruoiform shaped structures ¦ cxempli~ied in Figures 3 and 4 as 12 and 34 may be provided _ ___ _ 6 ¦ in a preformed shape prior to assembly into an alr~rame or like 71 str~c~ure. ~ is standard practice in the art to manufacture 81 fiber based composite materials in the shape of a woven fabric 9¦ or a unidirectional tape which already contains the binding 10¦ organic resin in a suitable prepolymerized form. These materials 11¦ are routinely referred to in the ar~ as pre-impregnated or 12¦ "prepreg'~ materials. Because final curing of the binding resin 13¦ does not occur unless the resin is subjected to heat, the pre-14¦ impregnated composite materials usually maintain their uncured 15¦ state for a prolonged period of time particularly if they are 16¦ kept at lower than ambient temperature.
17¦ Thus, it is possible, in accordance with the present 18¦ in~ention to manufac~ure several portions of an airframe and 19¦ the like from composite materials in a preimpregnated state,.
201 As an example,`~igure 7 schematically illustrates three portions 21¦ 48 of a stringer 24 of a fuselage 20, with each portion 48 22¦ already ha~iny two respective portions 52 of a ~rame member 26 231 attached th~reto, by the above described interwoven composite 241 structure. Each of ~hese portions of the stringer 24 and of the frame 26 are in a preimpregnated state.
26 They are splice~ to one anot~er duxing the a sembly o the 27 airframe by the use o~ s~ructural adhesive plastic (not 28 shown) an y conventlonal splice plates 54 sr~ratirally illrst~ated 31~ -15-.
_ . .,, ~
3 ~ 77~L5~
~- , 1 Fi~ure 7. Subsequent to splicing, the ent~re airframe or a 2 suitabl~ selected part thereof is finally cured ~n an autoclave 3 or oven (not shown)~ In ~his final curing step, requisite 4 curincJ o~ the structurial adhesive resin may also occur.
What has been described above is a no~el structure _ ____ ~ for strongly joining intersecting structural members made of 7 fiber based unidirectional composite materials. The novel
8 structure is capable of overcoming prior art constraints of
9 load absorbing continuity. Various modifications o~ the present
10 invention may become readily apparent to those skilled in the
11 art~ Consequently, the scope of the present invention should
12 e interpreted solely from the following claims,
13 //
154 /i 17 .
19 `
~ ~1 . ' . ' ' ' , , ' , .
` 2g . ~". ,', .
~; 30 , ~ 31 ~2 ~16-.'
154 /i 17 .
19 `
~ ~1 . ' . ' ' ' , , ' , .
` 2g . ~". ,', .
~; 30 , ~ 31 ~2 ~16-.'
Claims (17)
1. A composite cruciform structure for joining together intersecting structural members of an aircraft construction and the like, said cruciform structure comprising a first bundle of fibers interwoven with a second bundle of fibers and additional fibers interwoven with each of said first and second bundles of fibers in a direction substantially perpendicular to the fibers of said first and second bundles to form said cruciform structure, the interwoven fibers of said first and second bundles intersecting one another, said fibers having a reinforcing resin on the surface thereof to provide a strong structural joinder of said structural members without a substantial break in the continuity of the fibers and to provide substantially even load bearing and force transmitting capability in a plurality of directions with respect to said intersection of said structural members.
2. The cruciform structure of claim 1 wherein each bundle of fibers comprises a plurality of layers of fibers.
3. The cruciform structure of claim 1 wherein the fibers are selected from the group consisting of carbon, graphite, Kevlar, glass and boron fibers.
4. The cruciform structure of claim 2 wherein the fibers are selected from the group consisting of carbon, graphite, Kevlar, glass and boron fibers.
5. The cruciform structure of claim 3 or 4 wherein each fiber is approximately 0.001 inch in diameter, and each of said fibers comprises a plurality of thinner subfibers.
6. The cruciform structure of any one of claims 1 to 3 in which said first and second bundles of fibers are substantially perpendicular to one another.
7. The composite cruciform structure of any one of claims 1 to 3 wherein the first structural member comprises a stringer member in a fuselage of an aircraft and the second structural member comprises a frame member in the fuselage of the aircraft.
8. The composite cruciform structure of any one of claims 1 to 3 wherein the first structural member comprises a stringer member in a fuselage of an aircraft and the second structural member comprises a frame member in the fuselage of the aircraft, and wherein the first and second bundles of fibers are respectively spliced to fibers comprising portions of the respective stringer and frame members.
9. The composite cruciform structure of any one of claims 1 to 3 wherein the first structural member comprises a rib member of a wing or empennage of the aircraft and wherein the second structural member comprises a spar member of the wing of empennage of the aircraft.
10. The composite cruciform structure of any one of claims 1 to 3 wherein each bundle of fibers comprises a plurality of substantially two-dimensional layers of fibers, said layers in each bundle being disposed substantially parallel to one another.
11. The composite cruciform structure of any one of claims 1 to 3 in which fibers of said first and second bundles are interwoven between two of said additonal substantially perpendicular fibers.
12. A composite cruciform structure for joining together intersecting structural membes of an aircraft construction and the like, said cruciform structure comprising a first bundle of fibers interwoven with a second bundle of fibers, and additional bundles of fibers disposed substantially perpendicular to said first and second bundles of fibers and interwoven with both said first and second bundles to form said cruciform structure, the interwoven fibers of said first and second bundles intersecting one another, said fibers being selected from the group consisting of carbon, graphite, Kevlar, glass and boron fibers, each of said bundles of fibers comprising a plurality of layers of fibers, said bundles being provided with a reinforcing resin on the surface thereof, to provide a strong structural joinder of said structural members without a substantial break in the continuity of said fibers and to provide substantially even load bearing and force transmitting capability in a plurality of directions with respect to said intersection of said structural members.
13. The cruciform structure of claim 12 in which said first and second bundles of fibers are substantially perpendicular to one another.
14. A process for strongly joining a first substantially elongated composite structural member and a second substantially elongated composite structural member, said structural members extending in different directions relative to one another, the process comprising the steps of:
interweaving a first bundle of fibers comprising a portion of the first structural member with a second bundle of fibers comprising a portion of the second structural member, the interwoven fibers of said first and second bundles intersecting one another, said first and second bundles of fibers substantially extending in the same direction as the respective elongated structural members;
interweaving additional fibers with said first and second bundles disposed substantially perpendicular relative to said first and second bundles of fibers;
applying a suitable reinforcing resin to said fibers including the area wherein the bundles are interwoven; and curing said resin.
interweaving a first bundle of fibers comprising a portion of the first structural member with a second bundle of fibers comprising a portion of the second structural member, the interwoven fibers of said first and second bundles intersecting one another, said first and second bundles of fibers substantially extending in the same direction as the respective elongated structural members;
interweaving additional fibers with said first and second bundles disposed substantially perpendicular relative to said first and second bundles of fibers;
applying a suitable reinforcing resin to said fibers including the area wherein the bundles are interwoven; and curing said resin.
15. The process of claim 14 wherein the fibers are selected from the group consisting of carbon, graphite, glass, Kevlar and boron fibers.
16. The process of claim 14 or 15 further comprising the step of splicing at least one of the first and second bundles of fibers to another portion of the respective first or second structural member.
17 The process of claim 14 or 15 further comprising the step of splicing at least one of the first and second bundles of fibers to another portion of the respective first or second structural member, wherein the step of splicing includes applying an adhesive resin to at least one of the respective bundles of fibers and of the corresponding structural member, and curing said adhesive resin.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US14211880A | 1980-04-21 | 1980-04-21 | |
| US142,118 | 1980-04-21 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1177459A true CA1177459A (en) | 1984-11-06 |
Family
ID=22498614
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000375674A Expired CA1177459A (en) | 1980-04-21 | 1981-04-16 | Composite structure for joining intersecting structural members of an airframe and the like |
Country Status (5)
| Country | Link |
|---|---|
| JP (1) | JPS5734944A (en) |
| CA (1) | CA1177459A (en) |
| DE (1) | DE3115791A1 (en) |
| GB (1) | GB2074117B (en) |
| IT (1) | IT1143497B (en) |
Families Citing this family (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4962904A (en) * | 1984-06-07 | 1990-10-16 | The Boeing Company | Transition fitting for high strength composite |
| US4671470A (en) * | 1985-07-15 | 1987-06-09 | Beech Aircraft Corporation | Method for fastening aircraft frame elements to sandwich skin panels covering same using woven fiber connectors |
| US5223067A (en) * | 1990-02-28 | 1993-06-29 | Fuji Jukogyo Kabushiki Kaisha | Method of fabricating aircraft fuselage structure |
| JP2935722B2 (en) * | 1990-02-28 | 1999-08-16 | 富士重工業株式会社 | Aircraft fuselage structure and molding method thereof |
| ES2131479B1 (en) * | 1997-11-10 | 2000-03-01 | Torres Martinez M | TOOL AND ASSEMBLY PROCESS FOR LASER WELDING. |
| DE102007054053A1 (en) * | 2007-11-13 | 2009-05-20 | Airbus Deutschland Gmbh | Coupling element for connecting two longitudinal stiffening elements |
| DE102007055233A1 (en) | 2007-11-20 | 2009-05-28 | Airbus Deutschland Gmbh | Coupling device for joining fuselage sections, combination of a coupling device and at least one fuselage section and method for producing the coupling device |
| US8079549B2 (en) * | 2008-06-30 | 2011-12-20 | EMBRAER—Empresa Brasileira de Aeronautica S.A. | Monolithic integrated structural panels especially useful for aircraft structures |
| US8973871B2 (en) * | 2013-01-26 | 2015-03-10 | The Boeing Company | Box structures for carrying loads and methods of making the same |
| DE102013219820A1 (en) * | 2013-09-30 | 2015-04-02 | Bayerische Motoren Werke Aktiengesellschaft | Fiber composite component, method for producing a fiber composite component and use of fiber bundles and bracing means for producing a fiber composite component |
| DE102014222933B4 (en) | 2014-11-11 | 2021-09-09 | Bayerische Motoren Werke Aktiengesellschaft | Fiber composite component and method for producing a fiber composite component |
| JP2025167342A (en) | 2024-04-25 | 2025-11-07 | 株式会社シマノ | Fishing Line Guide |
-
1981
- 1981-04-16 CA CA000375674A patent/CA1177459A/en not_active Expired
- 1981-04-17 IT IT67534/81A patent/IT1143497B/en active
- 1981-04-18 DE DE19813115791 patent/DE3115791A1/en not_active Withdrawn
- 1981-04-21 GB GB8112305A patent/GB2074117B/en not_active Expired
- 1981-04-21 JP JP6109781A patent/JPS5734944A/en active Pending
Also Published As
| Publication number | Publication date |
|---|---|
| JPS5734944A (en) | 1982-02-25 |
| IT8167534A0 (en) | 1981-04-17 |
| DE3115791A1 (en) | 1982-08-12 |
| IT1143497B (en) | 1986-10-22 |
| GB2074117A (en) | 1981-10-28 |
| GB2074117B (en) | 1984-07-25 |
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