GB2074117A - Composite structure for joining intersecting structural members of an airframe - Google Patents
Composite structure for joining intersecting structural members of an airframe Download PDFInfo
- Publication number
- GB2074117A GB2074117A GB8112305A GB8112305A GB2074117A GB 2074117 A GB2074117 A GB 2074117A GB 8112305 A GB8112305 A GB 8112305A GB 8112305 A GB8112305 A GB 8112305A GB 2074117 A GB2074117 A GB 2074117A
- Authority
- GB
- United Kingdom
- Prior art keywords
- fibers
- bundle
- structural
- bundles
- structural member
- 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.)
- Granted
Links
- 239000002131 composite material Substances 0.000 title claims abstract description 59
- 238000005304 joining Methods 0.000 title claims description 14
- 239000000835 fiber Substances 0.000 claims abstract description 168
- 229920005989 resin Polymers 0.000 claims abstract description 30
- 239000011347 resin Substances 0.000 claims abstract description 30
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 claims abstract description 21
- 229910002804 graphite Inorganic materials 0.000 claims abstract description 16
- 239000010439 graphite Substances 0.000 claims abstract description 16
- 238000010276 construction Methods 0.000 claims description 25
- 238000000034 method Methods 0.000 claims description 13
- 230000003014 reinforcing effect Effects 0.000 claims description 13
- 229920000271 Kevlar® Polymers 0.000 claims description 6
- 239000004761 kevlar Substances 0.000 claims description 6
- 239000004840 adhesive resin Substances 0.000 claims description 5
- 229920006223 adhesive resin Polymers 0.000 claims description 5
- 229910052799 carbon Inorganic materials 0.000 claims description 5
- 239000011521 glass Substances 0.000 claims description 5
- ZOXJGFHDIHLPTG-UHFFFAOYSA-N Boron Chemical compound [B] ZOXJGFHDIHLPTG-UHFFFAOYSA-N 0.000 claims description 4
- 229910052796 boron Inorganic materials 0.000 claims description 4
- 239000012779 reinforcing material Substances 0.000 claims description 4
- 230000000694 effects Effects 0.000 claims description 2
- 230000002787 reinforcement Effects 0.000 claims 1
- 239000000463 material Substances 0.000 abstract description 12
- 239000010410 layer Substances 0.000 description 7
- 238000004519 manufacturing process Methods 0.000 description 5
- 239000004033 plastic Substances 0.000 description 5
- 239000000853 adhesive Substances 0.000 description 3
- 230000001070 adhesive effect Effects 0.000 description 3
- 229910000831 Steel Inorganic materials 0.000 description 2
- 239000002657 fibrous material Substances 0.000 description 2
- 239000003365 glass fiber Substances 0.000 description 2
- 238000010348 incorporation Methods 0.000 description 2
- 239000002184 metal Substances 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 239000010959 steel Substances 0.000 description 2
- 238000003466 welding Methods 0.000 description 2
- 239000004593 Epoxy Substances 0.000 description 1
- 230000005540 biological transmission Effects 0.000 description 1
- 238000005520 cutting process Methods 0.000 description 1
- 230000001419 dependent effect Effects 0.000 description 1
- 238000013461 design Methods 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 230000003292 diminished effect Effects 0.000 description 1
- 239000003822 epoxy resin Substances 0.000 description 1
- 239000011152 fibreglass Substances 0.000 description 1
- 239000000203 mixture Substances 0.000 description 1
- 229920000647 polyepoxide Polymers 0.000 description 1
- 238000006116 polymerization reaction Methods 0.000 description 1
- 230000002035 prolonged effect Effects 0.000 description 1
- 239000002356 single layer Substances 0.000 description 1
- 239000000126 substance Substances 0.000 description 1
- 239000002759 woven fabric Substances 0.000 description 1
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
A joint between two structural members of composite graphite fibre or other fibre-reinforced material, which may for example form part of the frame of an aircraft, is provided by interdigitated fibres of two bundles 18 and 19 each of which eventually forms one of the structural members 14 and 16. The bundles 18 and 19 are positioned so that they intersect one another at an angle and the graphite fibres are interdigitated at the point of intersection. The fibres are bonded together with a curable resin so as to form a rigid joint. A joint of this kind permits even distribution of structural loads irrespective of the direction of the load force. The joint may be further reinforced by the addition of a third bundle of fibres 32 which extends at an angle to both bundles 18 and 19 and which are interwoven with the fibres of both members 14 and 16 at the intersection. <IMAGE>
Description
SPECIFICATION
Composite structure for joining intersecting structural members of an airframe and the like
The present invention relates to a structure joining intersecting fiber based composite structural members of an airframe and the like, and to a process for making said structure.
Fiber based composite materials have been known in the prior art for a long time. Briefly, such materials comprise a plurality of relatively thin fibers and a reinforcing cured plastic which substantially covers the fibers and holds them together. Furthermore, it was recognized in the prior art that the structural strength of fiber based composite materials is the greatest in the direction of the fibers. Accordingiy, composite materials have been prepared in the past wherein all of the fibers are disposed in one direction parallel to one another. These type of composite materials are hereinafter referred to as unidirectional composite materials.
A principal characteristic of unidirectional fiber based composite materials is their above mentioned anisotropy. Thus, these materials exhibit relatively great strength to withstand forces which are applied substantially in the direction of the fibers. However, the load bearing or force withstanding capability of the unidirectional fiber based composite materials against forces which are applied perpendicularly to the direction of fibers, is substantially less. This follows from the fact that the parallel disposed fibers of the composite materials are held together only by the cured resin. The prior art composite materials or structures are particularly vulnerable to forces which tend to separate the fibers from one another in a direction perpendicular to the layout of the fibers.Nevertheless, composite materials, and particularly plastic reinforced glass fibers (fiberglass) have found several applications in the prior art where a relatively light weight and yet strong structural material was desired.
The relatively recent development of unidirectional composite materials containing graphite and other fibers of high strength has rendered possible the utilization of fiber based composite materials in airframe construction.
More particularly, unidirectional composite materials comprising epoxy resin reinforced graphite or other fibers of high strength are currently used, at least to a limited extent, to provide stringer and frame type structural members in aircraft fuselages, and rib and spar type structural members in aircraft wings.
A substantial disadvantage of conventional type airframe construction is that wherever two structural members intersect one another, it is necessary to provide a cut-away portion in one of the structural members so as to accommodate the other structural member. Cutting away a portion of a structural member, of course, diminishes its load bearing 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.
The principle of prior art airframe construction utilizing the aforementioned reinforcing members is schematically shown on Figures 1 and 2 of the drawings. On Figure 1 a structural member incorporating a cut-away portion illustrates, e.g. a fore-and-aft positioned stringer of an aircraft fuselage, and another structural member illustrates a laterally extending frame member of the fuselage. The necessity of providing a cutaway portion in one of the structural members whenever two structural members "intersect" one another, or "attempt to occupy the same space", is not limited to airframe construction. A similar problem is encountered in the construction of boats, vehicle frames, buildings, etc.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 significantly contributes to the overall construction cost.
Although the state-of-the-art application of high strength, relatively light weight 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 intersecting structural members to one another.
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 of "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.
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 intersecting composite structural members.
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.
It is another object of the present invention to provide a high strength interconnection or junction of two or more composite structural members in an airframe construction and the like wherein the strength of neither structural member is diminished by a cut-away portion.
It is still another object of the present invention to provide a high strength interconnection or junction of two or more composite structural members in an airframe construction and the like which allows for substantially even load bearing and force transmitting capability in a plurality of directions relative to the interconnection.
It is yet another object of the present invention to provide a high strength joining member
interconnecting two or more composite structural
members in an airframe construction and the like which is readily fastened to and becomes a
portion of the joined composite structural
members.
These and other objects and advantages are
attained by a first bundle of fibers which is
interwoven with a second bundle of fibers
substantially in a space wherein a first and a
second structural member intersect one another.
Each bundle comprises a plurality of substantially
parallel disposed graphite or like high strength fibers suitable for incorporation in a high strength
composite structural material. A third bundle of
like fibers which are substantially perpendicular
relative to the fibers of both the first and second
bundle may be interwoven with the first and with
the second bundle in the space, if desired. The third bundle of fibers may be part of a third
structural member.All the fibers are reinforced
with a suitable cured resin so as to provide a high
strength composite joining member wherein the first bundle of fibers extends in the general
direction of the first structural member and the second bundle of fibers extends in the general
direction of the second structural members The joining member may be spliced by a suitable structural adhesive resin to the first and second structural members or alternatively may comprise integral parts thereof.
In accordance with a first aspect of the invention there is provided a member made of a composite material for providing a strongly jointed connection of at least a first structural member and of a second structural member, the member comprising a first bundle of the fibers having a
plurality of parallel disposed fibers extending in a first direction, and a second bundle of fibers
having a plurality of parallel disposed fibers, the
second bundle of fibers extending in a second direction and substantially intersecting the first bundle of fibers, the fibers of the first bundle and of the second bundle being interwoven substantially where the respective fibers of the first and second bundle intersect one another, said bundles of fibers being provided substantially along their entire length with a suitable reinforcing resin, whereby the first bundle of fibers bonded with said resin comprising the first structural member and the second bundle of fibers bonded with said resin comprising the second structural member are strongly bonded to one another and in effect provide the strongly jointed connection of the first and second structural members.
In accordance with a second aspect of the invention there is provided a method of joining two members of fiber-based composite material comprising providing a first bundle of generally parallel fibers to form a first member and a second bundle of generally parallel fibers to form a second member, placing the first and second bundles adjacent to and at an angle to one another, interweaving the fibers of the first and second bundles where the bundles intersect, and providing a hardenable reinforcing material along the lengths of both members whereby, when the reinforcing material is hardened, the fibers of the first and second bundles are bonded together to form a rigid joint.
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, taken in connection with the drawings in which like numerals indicate like parts.
Figure 1 is a perspective view of a portion of a fuselage of an aircraft wherein a stringer and a frame member are joined to one another in accordance with the prior art;
Figure 2 is a perspective view of a portion of a wing of an aircraft wherein a spar and a rib member are joined to one another in accordance with the prior art;
Figure 3 is a schematic perspective view illustrating the principle of joining two intersecting fiber based unidirectional composite structural members to one another in accordance with the present invention;
Figure 4 is a schematic, perspective view further illustrating the principle of joining two intersecting fiber based unidirectional composite structural members to one another in accordance with the present invention;;
Figure 5 is a schematic, partial perspective view of an airplane fuselage with part of the skin being broken away, the view schematically illustrating frame and stringer members of the fuselage being joined to one another in accordance with the present invention;
Figure 6 is a schematic view of a wing of an airplane with part of the skin of the wing being broken away, the view schematically illustrating structural members of the wing being joined to one another in accordance with the present invention, and
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 structural member, said junctions being formed in accordance with the present invention.
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 invention.
Figures 1 and 2 of the drawing figures respectively depict aircraft fuselage and wing construction in accordance with the prior art.
These figures are explained below in comparison with fuselage and wing construction in accordance with the present invention, 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.
As it was briefly described in the introductory sections of the present application, a fiber based unidirectional composite material includes a plurality of relatively thin fibers which are disposed lengthwise, i.e. parallel relative to one another, and a suitable cured resin which substantially covers the fibers and holds them together. The state-of-the-art in the design and manufacture of fiber based composite materials is relatively advanced at the present. Therefore, the following concise, general description of fiber based composite materials and the process of their manufacture is intended solely for the purpose of facilitating the understanding of the present invention and for emphasizing and illuminating the novel features thereof.
Briefly, suitable fibers for the construction of strong unidirectional composite materials are glass, graphite, carbon, Kevlar and boron fibers.
(Kevlar is a trademark of the E.l. Dupont Company and is used to designate the source of certain fiber material.) The present invention may be practiced with either one of the above mentioned and with other fiber materials, although the use of glass fibers for the construction of heavy duty aircraft components is generally not preferred. Generally speaking, fibers and particularly graphite fibers used for construction of various structural members have a diameter of approximately 3 mils, and each of said fibers is itself a combination of a plurality of thinner subfibers.
In order to form a structural member of a predetermined dimension a bundle of fibers is positioned in such a manner that the longitudinal axes of the fibers are disposed parallel to one another. A suitable organic resin, which has not yet reached its fully polymerized or fully cured state, is then applied to the fibers. Subsequently, the resin is fully polymerized or cured under exposure to heat and in some instances high pressure. An important factor determining the selection of the proper resin for a given application is the nature of the fibers themselves. A person possessing average skill in the fiber based composite materials manufacturing arts is able to select the proper resin for a fiber of a given composition.Generally speaking, epoxy based resins are utilized in conjunction with graphite and other high strength fibers, although the present invention may be practiced with any type of fiber and resin combination. Usually the final polymerization or curing step of the composite material is conducted at 250-3500F for ±3 hours. The exact parameters of the aforementioned curing step are, of course, dependent on the exact nature of the fibers and on the chemical properties of the resinous binding material. Again, the scope of the present invention
is not limited in any way by the physical
parameters of the curing step.
Often, the bundle of the parallel disposed fibers is formed in the shape of a relatively thin band or tape, and several of the bands or tapes may be joined together in the curing step to form a structural member. An important characteristic and major advantage of the fiber based composite materials, and particularly of graphite fiber based composite materials is that they provide very high strength in the direction of the fibers at a relatively low weight.
The graphite fiber based composite materials are particularly preferred in the present invention for the construction of aircraft structural components, because these materials provide a structural integrity as high or higher than that of steel while the weight of these materials is considerably less than that of steel. On the other hand, a serious disadvantage of fiber based composite materials lies in their anisotropic behavior; in other words these materials exhibit much less structural integrity against forces which are not applied in the direction of fibers. As it is described below, this disadvantage is overcome by the present invention precisely at the points of intersection of two or more composite structural members wherein the disadvantage created by the anisotropy is the least tolerable.
Because of the well established importance of light weight and great structural integrity of materials utilized in airframe construction, the ensuing description is principally directed towards a description of the application of the present invention in airframe construction. Furthermore, the fibers utilized in the practice of the present invention will generally be referred to as graphite fibers. Nevertheless, it should be expressly understood that the scope of the present invention is not limited either to its application in airframe construction nor to the use of graphite fibers only.
Referring again to Figure 3, the basic principle of the novel structure 1 2 of the present invention is explained in detail. On Figure 3 the arrows respectively marked X, Y, and Z indicate three mutually perpendicular axes situated similarly to the axes of a three dimensional coordinate system.
In accordance with the present invention, a first substantially elongated structural member 14 is disposed substantially along the X axis, and a second substantially elongated structural member 1 6 is disposed substantially along the Y axis.
Although the first and second structural members 14 and 16 per se are not shown on Figure 3, a first bundle of fibers 1 8 corresponding to the first structural member 14 and a second bundle of fibers 1 9 corresponding to the second structural member 16, are clearly illustrated on this figure.
The first structural member 14 may be a stringer in a fuselage 20 of an aircraft 22, and the second structural member 16 may be a frame member in the fuselage 20 of the aircraft 22 as is illustrated in Figure 5.
A stringer 24 and a frame member 26 of an aircraft fuselage is shown on Figure 1 which depicts the prior art. It is readily discernible on
Figure 1 that where the stringer 24 and frame members 26 intersect one another, a cut-away portion has been provided in the stringer 24 so as to accommodate the frame member 26. In order not to lose structural strength and to provide for transmission of various forces from the stringer 24 and frame members 26 to one another, a plurality of reinforcing or clip members 28 were provided in the prior art. These were attached to the stringer 24 and to the frame member 26 by welding or by other conventional modes of attachment. Skin attached to the stringer 24 and frame members 26 is indicated by the reference numeral 30 on
Figure 1.
In the novel structure shown in Figure 3, the first bundle of fibers 1 8 corresponding to the first structural member 14 and hence to the frame member 26 of Figure 1, are intimately interwoven with the second bundle of fibers 19. In a preferred embodiment of the present invention, the interweaving is accomplished in such a manner that each fiber of the first bundle 18 running in the direction of axis X is positioned between two fibers of the second bundle 1 9 which run in the direction of axis Y. This is also true with regards to the fibers of the second bundle 19; i.e. each fiber of the second bundle 1 9 is positioned between two fibers of the first bundle 18. The second bundle of fibers 19 corresponds to the second structural member 1 6 and therefore to the stringer 24 of Figure 1.
In order to provide for further structural strength to the assembled first and second structural members 14 and 1 6 a third bundle 32 of fibers is interwoven with the first and second
bundles 1 8 and 1 9. The fibers of the third bundle
32 are disposed in the direction of axis Z. Thus, these fibers are perpendicular to the general
longitudinal axes of the fibers of the first 18 and of the second bundle 19. The fibers of the third bundle 32 may comprise a part of a third structural member (not shown), or may be, and
usually are, provided merely to lend additional structural integrity to the cruciform shaped structure 12. This additional structural integrity is particularly important to guard against forces acting along the Z axis, and eliminates certain prior art structural constraints of load bearing and transmitting continuity.
Although Figure 3 shows the first second and third bundles 18, 19 and 32 of fibers being perpendicularly disposed to one another, it is important to understand that the present invention is not limited in this manner. In other embodiments (not shown) the several bundles of fibers may be disposed at other than 900 angles relative to one another.
After having interwoven the fibers of the first, second and third bundles 1 8, 19 and 32, a suitable prepolymerized resin (not shown) is applied to the structure 12. Subsequently the resin is cured by heat according to standard practice in the art. In this regard it is noted that for the sake of clear illustration of the spatial arrangement of the fibers, the resin applied to the fibers has been omitted from the drawing figures.
Furthermore, it is emphasized that the drawing figures and particularly Figures 3 and 4 are merely schematic, and the actual number of fibers in each of the bundles 18, 19 and 32 is very iarge, as is in fiber based composite structural members of the prior art.
Figure 4 represents a schematic view of an embodiment of the cruciform shaped structure 34 of the present invention wherein each bundle of fibers 18, 19 and 32 comprises two layers of fibers. A first and a second layer of fibers in the bundle are provided for the sake of illustration with the respective reference numerals 36 and 38.
The layers of the fibers of the several bundles are interwoven with one another in a manner similar to the interweaving of the single layers of fibers as is shown on Figure 3. Each bundle 18, 19 and 32 is respectively disposed substantially in the direction of the respective X, Y and Z axes, although it should be again understood that the novel structure of the present invention may also be constructed in such a manner that the respective bundles of fibers and hence the respective structural members 1 4 and 1 6 are not at a 900 angle relative to one another.
In the actual practice of the present invention it is often necessary to provide multiple layers of fibers in each bundle in order to obtain a junction of the structural members which is sufficiently strong for incorporation in an aircraft. Each layer of fiber is relatively thin as compared to its length and width. Therefore, the layers are referred to as two dimensional layers. Actual dimensions of the structural members are determined by the particular engineering requirements of the aircraft frame or other structure. In certain embodiments one structural member may comprise a substantially lesser number of layers of fibers than a second structural member which is interwoven therewith.
Referring now to Figure 2 which schematically depicts a wing 40 construction in accordance with the prior art, it is noted that the first and second structural members 14 and 16 of Figures 3 and 4 may also correspond respectively to a wing spar member 42 and to a wing rib member 44. Skin 46 of the wing 46 may be attached to the novel composite wing structure by conventional means or by the use of aEstructural adhesive plastic. The use of structural adhesive plastic is well established in the arts and need not be described here in detail.
Referring now to Figures 5 and 6, an aircraft fuselage 20 and a wing 46 are schematically shown wherein intersecting frame 26 and stringer 24 members and intersecting wing spar 42 and rib members 44 are respectively constructed in accordance with the present invention. The schematic drawing of Figures 5 and 6 reveal that due to the novel mode of construction, no cutaway portion is provided where these members intersect each other. This is, of course in sharp contrast with the prior art fuselage and wing constructions which are illustrated in Figures 1 and 2.
It is an additional aspect and additional advantage of the present invention that the cruciform shaped structures exemplified in Figures 3 and 4 as 12 and 34 may be provided in a preformed shape prior to assembly into an airframe or like structure. It is standard practice in the art to manufacture fiber based composite materials in the shape of a woven fabric or a unidirectional tape which already contains the binding organic resin in a suitable prepolymerized form. These materials are routinely referred to in the art as pre-impregnated or "prepreg" materials.
Because final curing of the binding resin does not occur unless the resin is subjected to heat, the pre-impregnated composite materials usually maintain their uncured state for a prolonged period of time particularly if they are kept at lower than ambient temperature.
Thus, it is possible, in accordance with the present invention to manufacture several portions of an airframe and the like from composite materials in a pre-impregnated state. As an example, Figure 7 schematically illustrates three portions 48 of a stringer 24 of a fuselage 20, with each portion 48 already having two respective portions 52 of a frame member 26 attached thereto, by the above described interwoven composite structure. Each of these portions of the stringer 24 and of the frame 26 are in a preimpregnated state. They are spliced to one another during the assembly of the airframe by the use of structural adhesive plastic (not shown) and by conventional splice plates 54 schematically illustrated in Figure 7. Subsequent to splicing, the entire airframe or a suitably selected part thereof is finally cured in an autoclave or oven (not shown). In this final curing step, requisite curing of the structural adhesive resin may also occur.
What has been described above is a novel structure for strongly joining intersecting structural members made of fiber based unidirectional composite materials. The novel structure is capable of overcoming prior art constraints of load absorbing continuity. Various modifications of the present invention may become readily apparent to those skilled in the art.
Consequently, the scope of the present invention should be interpreted solely from the following
Claims (27)
1. A member made of a composite material for providing a strongly jointed connection of at least a first structural member and of a second structural member, the member comprising:
a first bundle of fibers having a plurality of parallel disposed fibers extending in a first direction, and
a second bundle of fibers having a plurality of parallel disposed fibers, the second bundle of fibers extending in a second direction and substantially intersecting the first bundle of fibers, the fibers of the first bundle and of the second
bundle being interwoven substantially where the
respective fibers of the first and second bundle
intersect one another, said bundles of fibers being
provided substantially along their entire length
with a suitable reinforcing resin, whereby the first
bundle of fibers bonded with said resin comprising
the first structural member and the second bundle
of fibers bonded with said resin comprising the
second structural member are strongly bonded to
one another and in effect provide the strongly
jointed connection of the first and second
structural members.
2. The invention of Claim 1 wherein the first
bundle of fibers and the second bundle of fibers
intersect one another substantially at a 90 angle.
3. The invention of Claim 2 wherein the
member further comprises a third bundle of fibers
reinforced with the suitable resin and being
disposed substantially at a 90 angle relative to
said first and to said second direction, said third
bundle of fibers providing further structural
reinforcement to the member.
4. The invention of Claim 1 wherein each
bundle of fibers comprises a plurality of layers of
fibers.
5. The invention of Claim 1 wherein the fibers
are selected from a group consisting of carbon,
graphite, Kevlar, glass and boron fibers.
6. The invention of Claim 5 wherein each fiber
is approximately 0.001 inch in diameter, and each
of said fibers comprises a plurality of thinner
subfibers.
7. In an aircraft frame construction and the like
wherein at least a first structural member and a
second structural member intersect one another,
and are attached to one another, the combination
which comprises:
a first bundle of substantially parallel disposed
fibers comprising a first portion of the first
structural member, and
a second bundle of substantially parallel
disposed fibers comprising a first portion of the
second structural member, the first and second
bundle of fibers being interwoven relative to one
another where the first and second bundles
substantially intersect one another, each of the
first and second bundles being provided with a
suitable reinforcing cured resin, whereby a strong
structural attachment of the first and second
structural members occurs without a substantial
break in the continuity of the fibers of the first and
second structural members.
8. The invention of Claim 7 further comprising a
third bundle of fibers reinforced with the suitable
cured resin, said third bundle being interwoven
with the first and second bundles at least
substantially in the area wherein the first and
second bundles intersect one another.
9. The invention of Claim 8 wherein said third bundle of fibers is disposed substantially at a
900C angle relative to said first bundle and said
second bundle.
10. The invention of Claim 8 wherein the first
and second bundles of fibers intersect one another
at a substantially 900 angle.
11. The invention of Claim 8 wherein the fibers are selected from a group consisting of carbon, graphite, Kevlar, glass andlboron fibers.
12. The invention of Claim 8 wherein the first structural member comprises a stringer member in a fuselage of an aircraft and wherein the second structural member comprises a frame member in the fuselage of the aircraft.
13. The invention of Claim 12 wherein the first and second bundles of fibers are respectively spliced to fibers comprising second portions of the respective stringer and frame members.
14. The invention of Claim 8 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 or empennage of the aircraft.
15. The invention of Claim 8 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.
1 6. In a junction of a first composite structural member having a plurality of substantially parallel fibers reinforced with a suitable cured resin, and a second composite structural member having a plurality of substantially parallel fibers reinforced with a suitable cured resin, the improvement comprising::
a reinforcing bundle of substantially parallel fibers disposed substantially at a 900 angle relative to the fibers of the first structural member and to the fibers of the second structural member, the fibers of the first structural member, of the second structural member and of the reinforcing bundle being mutually interwoven in the junction, the fibers of said reinforcing bundle also being provided with the suitable cured resin whereby the junction of the first and second structural members becomes capable of bearing substantially large structural loads applied in several directions relative to the structural members.
17. The invention of Claim 1 6 wherein the fiist and second structural members respectively comprise a stringer member and a frame member of the fuselage of an aircraft.
18. The invention of Claim 1 6 wherein the first and second structural members respectively comprise a rib member and a spar member of a wing of an aircraft.
19. The invention of Claim 16 wherein the fibers are selected from a group consisting of carbon, graphite, glass, Kevlar and boron fibers.
20. 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 substantially parallel fibers comprising a portion of the first structural member with a second bundle of substantially parallel fibers comprising a portion of the second structural member, said first and second bundles of parallel fibers substantially extending in the same direction as the respective elongated structural members;
applying a suitable reinforcing resin to said first and second bundles of fibers including the area wherein the bundles are interwoven, and curing said resin.
21. The process of Claim 20 further comprising the step of interweaving a third bundle of substantially parallel fibers with the first and second bundles, the fibers of the third bundle
being disposed at a substantially 900 angle
relative to the first and to the second bundle of fibers.
22. The process of Claim 20 wherein the fibers
are selected from a group consisting of carbon, graphite, glass, Kevlar and boron fibers.
23. The process of Claim 2Q further comprising the step of splicing at least one of the first and second bundle of fibers to another portion of the
respective first or second structural member.
24. The process of Claim 23 wherein the step of splicing includes the steps of applying a suitable adhesive resin to at least one of the respective
bundles of fibers and to the other portion of the
corresponding structural member, and curing said adhesive resin.
25. The invention of Claim 20 wherein the first
and second structural members comprise a portion of the airframe of an aircraft.
26. A method of pining two members of fibre
based composite material comprising providing a first bundle of generally parallel fibers to form a first member and a second bundle of generally
parallel fibers to form a second member, placing the first and second bundles adjacent to and at an angle to one another, interweaving the fibers of the first and second bundles where the bundles intersect, and providing a hardenable reinforcing
material along the lengths of both members whereby, when the reinforcing material is
hardened, the fibers of the first and second bundles are bonded together to form a rigid joint.
27. A method according to Claim 26 substantially as hereinbefore described.
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US14211880A | 1980-04-21 | 1980-04-21 |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| GB2074117A true GB2074117A (en) | 1981-10-28 |
| GB2074117B GB2074117B (en) | 1984-07-25 |
Family
ID=22498614
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB8112305A Expired GB2074117B (en) | 1980-04-21 | 1981-04-21 | Composite structure for joining intersecting structural members of an airframe |
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) |
Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2603249A1 (en) * | 1985-07-15 | 1988-03-04 | Beech Aircraft Corp | CONNECTION OF AIRCRAFT COATING PANEL TO A CARRIER AND METHOD FOR MAKING SAME |
| US4962904A (en) * | 1984-06-07 | 1990-10-16 | The Boeing Company | Transition fitting for high strength composite |
| ES2131479A1 (en) * | 1997-11-10 | 1999-07-16 | Torres Martinez M | Assembly tool and process for laser welding |
| WO2009062712A1 (en) * | 2007-11-13 | 2009-05-22 | Airbus Deutschland Gmbh | Coupling element for interconnecting two longitudinal reinforcing elements |
| US8453975B2 (en) | 2007-11-20 | 2013-06-04 | Airbus Operations Gmbh | Coupling device for coupling fuselage sections; combination of a coupling device and at least one fuselage section; and method for producing the coupling device |
| EP2307271A4 (en) * | 2008-06-30 | 2015-09-02 | Embraer Aeronautica Sa | Monolithic integrated structural panels especially useful for aircraft structures and methods of making the same |
| US10220578B2 (en) | 2014-11-11 | 2019-03-05 | Bayerische Motoren Werke Aktiengesellschaft | Fiber composite material component, and method for producing a fiber composite material component |
Families Citing this family (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| 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 |
| 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 |
-
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 JP JP6109781A patent/JPS5734944A/en active Pending
- 1981-04-21 GB GB8112305A patent/GB2074117B/en not_active Expired
Cited By (11)
| 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 |
| FR2603249A1 (en) * | 1985-07-15 | 1988-03-04 | Beech Aircraft Corp | CONNECTION OF AIRCRAFT COATING PANEL TO A CARRIER AND METHOD FOR MAKING SAME |
| ES2131479A1 (en) * | 1997-11-10 | 1999-07-16 | Torres Martinez M | Assembly tool and process for laser welding |
| WO2009062712A1 (en) * | 2007-11-13 | 2009-05-22 | Airbus Deutschland Gmbh | Coupling element for interconnecting two longitudinal reinforcing elements |
| CN101883717A (en) * | 2007-11-13 | 2010-11-10 | 空中客车营运有限公司 | Coupling element for connecting two longitudinal stiffening elements |
| US8353479B2 (en) | 2007-11-13 | 2013-01-15 | Airbus Operations Gmbh | Arrangement of two fuselage sections of an aircraft and a connecting structure for connecting fuselage skins |
| RU2479466C2 (en) * | 2007-11-13 | 2013-04-20 | Эрбус Оперейшнс Гмбх | System of aircraft fuselage two sections and connection structure for fuselage skin connection |
| CN101883717B (en) * | 2007-11-13 | 2014-08-06 | 空中客车营运有限公司 | System of aircraft fuselage two sections and connection structure for fuselage skin connection |
| US8453975B2 (en) | 2007-11-20 | 2013-06-04 | Airbus Operations Gmbh | Coupling device for coupling fuselage sections; combination of a coupling device and at least one fuselage section; and method for producing the coupling device |
| EP2307271A4 (en) * | 2008-06-30 | 2015-09-02 | Embraer Aeronautica Sa | Monolithic integrated structural panels especially useful for aircraft structures and methods of making the same |
| US10220578B2 (en) | 2014-11-11 | 2019-03-05 | Bayerische Motoren Werke Aktiengesellschaft | Fiber composite material component, and method for producing a fiber composite material component |
Also Published As
| Publication number | Publication date |
|---|---|
| GB2074117B (en) | 1984-07-25 |
| IT1143497B (en) | 1986-10-22 |
| IT8167534A0 (en) | 1981-04-17 |
| CA1177459A (en) | 1984-11-06 |
| DE3115791A1 (en) | 1982-08-12 |
| JPS5734944A (en) | 1982-02-25 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| US4715560A (en) | Composite cruciform structure for joining intersecting structural members of an airframe and the like | |
| AU2011253392B2 (en) | Composite beam chord between reinforcement plates and related method of manufacturing | |
| US4086378A (en) | Stiffened composite structural member and method of fabrication | |
| US9511571B2 (en) | Composite laminate having a damping interlayer and method of making the same | |
| EP1707702B1 (en) | Composite structural members and methods for forming the same | |
| US4120998A (en) | Composite structure | |
| US4113910A (en) | Composite load coupler for reinforcing composite structural joints | |
| US6187411B1 (en) | Stitch-reinforced sandwich panel and method of making same | |
| US4749155A (en) | Method of making wing box cover panel | |
| US4786343A (en) | Method of making delamination resistant composites | |
| US20100038030A1 (en) | Advanced composite aerostructure article having a braided co-cured fly away hollow mandrel and method for fabrication | |
| US6110567A (en) | Composite structural panel having a face sheet reinforced with a channel stiffener grid | |
| EP1868795B1 (en) | Multi-axial laminate composite structures | |
| EP1819591B1 (en) | A trussed structure | |
| EP2799329A1 (en) | Composite skin and stringer structure and method for forming the same | |
| US20100320320A1 (en) | Aeropspace structure including composite beam chord clamped between reinforcement plates | |
| GB2074117A (en) | Composite structure for joining intersecting structural members of an airframe | |
| GB2096530A (en) | A tubular hollow member and a method for its manufacture and a device for carrying out the method | |
| US5527584A (en) | High thermal conductivity triaxial non-metallic honeycomb | |
| EP0888235B1 (en) | Composite cuff structure for helicopter rotors | |
| US9731807B2 (en) | Joints in fibre metal laminates | |
| US5554430A (en) | Low CTE boron/carbon fiber laminate | |
| EP3608090B1 (en) | Composite connector and method of manufacturing the same | |
| US9682519B2 (en) | Integral composite bushing system and method | |
| WO2005049933A1 (en) | Sandwich panel and a method of producing a sandwich panel |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PCNP | Patent ceased through non-payment of renewal fee |