Composite fabric with rigid member structure

Abstract

A composite structure is provided including a first fabric and a second fabric. A substantially elongate and substantially rigid first member is spaced apart from and coupled with the first fabric via the second fabric. A resin substantially is infused into the first fabric and the second fabric, and substantially encapsulates the first member to form a unitary structure.

Description

BACKGROUND INFORMATION

1. Field

Embodiments relate to the field of aircraft construction, and, more specifically, to a composite fuselage section, a system and a method including a composite fabric structure with rigid member.

2. Background

Various assemblies and methods of assembly exist for composite structures which require opposing mold halves adapted to form the shape and contour of opposite sides of the composite structure. Many require additional assembly operations to fasten components together. To form an aircraft fuselage using composites, a skin is first formed, which typically requires tooling for both the inner mold line and the outer mold line. Then the skin is fastened to longitudinal stringers and transverse frame members. Joining the skin to the stringers and frames requires multiple fasteners. In addition, the role an aircraft fuselage will play, for example, if used for a bomber aircraft verses an attack aircraft etc., will further dictate individualized tooling requirements.

There exists a need to reduce manufacturing tooling requirements, the number of parts for final assembly, and the number of assembly operations.

SUMMARY

Embodiments may include a method of manufacturing a composite structure. The method may include laying out a first fabric on a tool surface; positioning an elongate first member a distance from the first fabric; covering the first member with a second fabric such that tab-like portions of the second fabric contact the first fabric on opposite sides of the first member; coupling the first member to the first fabric by securing the tab-like portions to the first fabric; and forming a unitary structure with the first fabric, the second fabric, and the first member by infusing a resin into the first fabric and the second fabric and encasing the first member within the resin.

The covering the first member with the second fabric may include positioning the second fabric in a slot of a fixture forming a valley with the second fabric and positioning the rod in the valley.

In various embodiments the method may also include positioning a second member over the first fabric and transverse to the first member; covering the second member with a third fabric such that other tab-like portions of the third fabric contact the first fabric on opposite sides of the second member; coupling the second member to the first fabric by securing the other tab-like portions of the third fabric to the first fabric; and wherein the forming a unitary structure includes infusing the resin into the third fabric and encasing the second member within the resin. The securing the second fabric and the securing the third fabric may be by stitching.

Various methods in accordance with various embodiments may also include positioning a substantially airtight film over the combination of the first member, the second member, the first fabric, the second fabric, and the third fabric; and forming a vacuum between the tool surface and the film, and wherein the infusing includes introducing the resin between the tool surface and the film. The method may also include curing the resin. The first member may be a pultruded rod. The second fabric may be a warp-knit fabric and may further comprise orienting the warp of the warp-knit substantially perpendicular to a longitudinal axis of the elongate first member.

Various embodiments may provide an assembly to form a composite assembly, which may include: a first fabric laid on a surface; a second member laid on the first fabric; a third fabric draped over the second member, having two flaps that extend beyond either side of the second member. The two flaps may each be stitched to the first fabric. An elongate first member may pass through the second member and may be supported by the second member a distance from the first fabric. A second fabric may be draped over the first member which may have two other flaps that extend down from the first member. The two other flaps may each be stitched to the first fabric. The assembly may also include a substantially airtight film placed over the first and second members and the first, second, and third fabrics. The assembly may also be adapted to form a vacuum seal between the film and the surface, and further adapted for a resin to be introduced between the film and the surface. The first and second members may be included in at least a portion of the completed composite assembly.

The features, functions, and advantages can be achieved independently in various embodiments of the present disclosure or may be combined in yet other embodiments in which further details can be seen with reference to the following description and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The novel features believed characteristic of the advantageous embodiments are set forth in the appended claims. The advantageous embodiments, however, as well as a preferred mode of use, further objectives, and advantages thereof, will best be understood by reference to the following detailed description of an advantageous embodiment of the present disclosure when read in conjunction with the accompanying drawings, wherein:

FIG. 1 is a flow diagram of aircraft production and service methodology;

FIG. 2 is a block diagram of an aircraft;

FIG. 3A illustrates a perspective view, and FIG. 3B an exploded perspective view in accordance with various embodiments;

FIG. 4 illustrates a cross-sectional view taken at line 4-4 in FIG. 3A;

FIG. 5 illustrates a cross-sectional view a portion of a flooring panel taken at line 5-5 in FIG. 3A;

FIG. 6 illustrates an end view of a fixture that may be used in the assembly of various embodiments;

FIG. 7 illustrates a side view with portions in cutaway view for illustration in accordance with various embodiments;

FIG. 8 illustrates a somewhat schematic view of portions of an assembly process in accordance with various embodiments;

FIGS. 9A and 9B illustrate cross-sectional views of respectively different embodiments illustrating selected design parameters being varied and respective varied resultant structures;

FIG. 10A illustrates a side view and FIG. 10B illustrates a cross-sectional view taken at the line 10B-10B in accordance with various embodiments;

FIG. 11A is a block diagram and FIG. 11B is a cross-sectional view taken at the line 11B-11B in FIG. 11A;

FIG. 12 is a flow diagram illustrating a method in accordance with various embodiments; and

FIG. 13 is a flow diagram illustrating a method in accordance with various embodiments.

DETAILED DESCRIPTION

In the following detailed description, reference is made to the accompanying drawings which form a part hereof, and in which are shown by way of illustration embodiments which may be practiced. It is to be understood that other embodiments may be utilized and structural or logical changes may be made without departing from the scope of the disclosure. Therefore, the following detailed description is not to be taken in a limiting sense, and the scope of embodiments is defined by the appended claims and their equivalents.

Various operations may be described as multiple discrete operations in turn, in a manner that may be helpful in understanding the embodiments; however, the order of description should not be construed to imply that these operations are order dependent.

The description may use perspective-based descriptions such as up/down, back/front, and top/bottom. Such descriptions are merely used to facilitate the discussion and are not intended to restrict the application of the embodiments.

The terms “coupled” and “connected,” along with their derivatives, may be used. It should be understood that these terms are not intended as synonyms for each other. Rather, in particular embodiments, “connected” may be used to indicate that two or more elements are in direct physical or electrical contact with each other. “Coupled” may mean that two or more elements are in direct physical or electrical contact. However, “coupled” may also mean that two or more elements are not in direct contact with each other, but yet still cooperate or interact with each other.

For the purposes of the description, a phrase in the form “A/B” means A or B. For the purposes of the description, a phrase in the form “A and/or B” means “(A), (B), or (A and B).” For the purposes of the description, a phrase in the form “at least one of A, B, and C” means “(A), (B), (C), (A and B), (A and C), (B and C), or (A, B and C).” For the purposes of the description, a phrase in the form “(A)B” means “(B) or (AB),” that is, A is an optional element.

The description may use the phrases “in an embodiment,” or “in embodiments,” which may each refer to one or more of the same or different embodiments. Furthermore, the terms “comprising,” “including,” “having,” and the like, as used with respect to the embodiments, are synonymous.

Embodiments may provide a very efficient airframe structure having the skin, stringers, and frame elements which may maintain residual load-carrying capabilities under a variety of damage scenarios. Embodiments may enable higher operating strains and ultimately a lighter airframe. In addition to the improved structural performance, embodiments of, for example, dry carbon fiber, pultruded rods, and stitching thread may also provide manufacturing advantages. A self-supporting stitched preform design of various embodiments may include a new out-of-autoclave fabrication approach that may provide orders of magnitude lower building costs than would otherwise be possible using traditional composite design and manufacturing practices.

Embodiments may provide larger one-piece occurred panel designs with seamless transitions and damage-arrest interfaces. Embodiments may also provide performance and cost advantages that are capable of enabling a whole new generation of highly-contoured aircraft designs that may result in substantially improved mission capabilities and lower life cycle costs for future aircraft, including military aircraft.

Referring more particularly to the drawings, embodiments of the disclosure may be described in the context of an aircraft manufacturing and service method 100 as shown in FIG. 1 and an aircraft 102 as shown in FIG. 2. During preproduction, exemplary method 100 may include specification and design 104 of the aircraft 102 and material procurement 106. During production, component and subassembly manufacturing 108 and system integration 110 of the aircraft 102 takes place. Thereafter, the aircraft 102 may go through certification and delivery 112 in order to be placed in service 114. While in service by a customer, the aircraft 102 is scheduled for routine maintenance and service 116 (which may also include modification, reconfiguration, refurbishment, and so on).

Each of the processes of method 100 may be performed or carried out by a system integrator, a third party, and/or an operator (e.g., a customer). For the purposes of this description, a system integrator may include without limitation any number of aircraft manufacturers and major-system subcontractors; a third party may include without limitation any number of venders, subcontractors, and suppliers; and an operator may be an airline, leasing company, military entity, service organization, and so on.

As shown in FIG. 2, the aircraft 102 produced by exemplary method 100 may include an airframe 118 with a plurality of systems 120 and an interior 122. Examples of high-level systems 120 include one or more of a propulsion system 124, an electrical system 126, a hydraulic system 128, and an environmental system 130. Any number of other systems may be included. Although an aerospace example is shown, the principles of the invention may be applied to other industries, such as the automotive industry.

Apparatus and methods embodied herein may be employed during any one or more of the stages of the production and service method 100. For example, components or subassemblies corresponding to production process 108 may be fabricated or manufactured in a manner similar to components or subassemblies produced while the aircraft 102 is in service. Also, one or more apparatus embodiments, method embodiments, or a combination thereof may be utilized during the production stages 108 and 110, for example, by substantially expediting assembly of or reducing the cost of an aircraft 102. Similarly, one or more of apparatus embodiments, method embodiments, or a combination thereof may be utilized while the aircraft 102 is in service, for example and without limitation, to maintenance and service 116.

FIG. 3A is a perspective view and FIG. 3B is an exploded perspective view illustrating various embodiments. A composite structure 10 may include a first fabric 12, and a second fabric 14. The embodiment illustrated may include a substantially elongate and substantially rigid first member 16 spaced apart from and coupled with the first fabric 12 via the second fabric 14. The first member 16 may be spaced from the first fabric 12 by a distance 18. A resin may be infused into the first fabric 12 and the second fabric 14, and may also encapsulate the first member 16 to form a unitary structure 11. Examples of structure and methods for applying the resin in accordance with various embodiments are discussed below.

FIG. 4 is a cross-sectional view taken through the line 4-4 in FIG. 3A. The second fabric 14 may be stitched to the first fabric 12 along two seams 20. Each of the seams 20 may be substantially parallel with the first member 16 such that a portion 22 of the first fabric between the two seams 20 may be disposed to form a loop 21 over the first member 16 to couple the first member 16 with the first fabric 12. While first member 16 is shown as a rod for illustration and without limitation any structural shape may be used. While first member 16 is shown with a uniform cross section over its length, the cross section shape and/or size of first member 16 may vary over its length. A tightening stitch 24 may be disposed through two locations 26 on the second fabric 14 adjacent the first member 16. A wrapping portion 28 may be defined on the second fabric 14 between the two locations 26. The tightening stitch 24 may be adapted to tighten the wrapping portion 28 over the first member 16. Additional stitching rows 30 may also serve to couple the second fabric 14 to the first fabric 12. A tear strap 32 may be disposed between the first fabric 12 and the second fabric 14.

FIG. 5 is a cross-sectional view taken through the line 5-5 in FIG. 3A. Referring now to FIG. 5 and also to FIGS. 3A and 3B, a second member 34 may be disposed transverse to the first member 16 and may be coupled with the first fabric 12 via a third fabric 36. The resin may also be infused into the third fabric 36 and may encapsulate the second member 34. The second member 34 and the third fabric 36 may form part of the unitary structure 11. First and second members could be graphite and/or foam or some other material. Both could be any structural shape and/or have a cross section that varies over its length.

In various embodiments the first member 16 may pass through the second member 34. The second member 34 may support the first member 16 at the distance 18 (FIG. 3A). The first member 16 may be a pultruded rod. At least part of the second member 34 may be made substantially from foam core.

Still referring to FIGS. 3A and 3B, various embodiments can be used to form all or part of an aircraft fuselage. The fuselage may include a fuselage skin 12′. The fuselage skin 12′ may include a first fabric 12. Two or more stringer members 15 may be oriented longitudinally with a longitudinal axis 40 of the fuselage. Each stringer members 15 may include an elongate rigid first member 16 disposed a distance 18 from the fuselage skin 12′. Each elongate rigid and first member 16 may be coupled with the fuselage skin 12′ via a second fabric 14. Two or more frame members 34′ may be oriented transverse to the fuselage and may be of sufficient rigidity to provide support for the first members 16 at the distance 18 from the skin 12′. The frame members 34′ may be coupled with the skin 12′ via a third fabric 36. A resin may be infused into the first 12, second 14, and third fabrics 36 and may encapsulate the first 16 and second members 34. The second fabric 14 may be looped over the first member 16 and may be long enough to form at least one tab-like portion 70 that contacts the first fabric 12. The at least one tab-like portion 70 may be stitched to the first fabric 12.

Frame members 34′ may run perpendicular to the stringers 16′ in a radial direction 41 and may be spaced at, for example, roughly 20-inch intervals and may be constructed, for example and without limitation, from a foam-core design wrapped within warp-knit fabric. The second members 34 may include additional material buildups at the top 31 and base 33 to provide bending stiffness. For example, frame caps 35 (FIG. 5) may be used and may be continuous annularly around the fuselage. Slots 37 (FIG. 3B) may be cut in frame member 34′ webs 43 to allow stringer members 15′ to pass through. The excess frame web material at each stringer slot 37 may drape around the stringer members 15′ to provide a fillet flange that would strengthen the intersection. Other prominent airframe features like window and door assemblies may also be included in an integrated design.

FIG. 5 illustrates the third fabric 36 as having two plies or layers of fabric. Various embodiments may use various numbers of layers, various fabric types and technologies for any and all of the first, second, and third fabrics. For example, a dry warp-knit fabric may be used. Warp-knitting technology may provide the flexibility of mixing low concentrations and material types. For example, carbon and glass of various tensile strengths and moduli, at various orientations and locations may be utilized to create highly optimized multiaxial fabrics. These fabrics can be specifically tailored for different applications and in the case of aircraft fuselage construction for different aircraft types and loading conditions to optimize cost and weight. Embodiments may enable a multirole design because it may offer an inexpensive way to optimize the weight and cost performance of the basic material for each application or aircraft type. For example, multiple aircraft differing roles, i.e., bomber, passenger, etc. may be assembled using common tooling and common parts but having differing features.

FIG. 6 is an end view illustrating various embodiments wherein a fixture 50 may be used as an aide to assemble various elements. Multiple, for example six, second fabrics 14 may be positioned in multiple slots 52 in the fixture 50 forming multiple valleys, or pockets 54 with the second fabrics 14. Multiple first members 16 in the form of elongated rods 16 may be positioned in the valleys, or pockets 54. One or more second member 34 may be disposed in respective one or more transverse slots 53 in the fixture 50. The one or more second members may include slots 37 through which the first members 16 and second fabrics may pass. The slots may be keyhole shaped as illustrated, and may be adopted to support the first members 16 during fabrication. The first fabric 12 may then be coupled with each of the second fabrics 14.

FIG. 7 is a side view with a portion cutaway for illustration of an assembly in accordance with various embodiments. The assembly illustrated in FIG. 4 may be removed from the fixture and placed on a surface 56 which may be on an outer mold line 58. A substantially airtight film 60 may be placed over the first and second members 16, 34 and the first, second and third fabrics 12, 14, 36, and adapted to form a vacuum seal between the film 60 and the surface 56. The first and second members 16, 34 may be included in at least a portion of the completed composite assembly.

FIG. 8 is a cross-sectional view illustrating, somewhat schematically, embodiments wherein the resin may be introduced and/or infused into the assembly. The resin may be pushed through the assembly which may be a dry carbon perform by a vacuum-induced pressure differential, that may not only promote resin flow but which may also create a pressure differential across the film 60 vacuum bag surface. A resin inlet 62 may introduce the resin under the film 60 while, or after, or just before a vacuum line 64 creates the pressure differential. This pressure may compact the preform and may enable laminate fiber volumes of nearly 60%. Even relatively low-pressure processing techniques may achieve higher fiber volume fractions and mechanical properties. The process may be superior to state-of-the-art prepreg material systems that typically require autoclave processing environments.

The edges 66 of the film 60 may be sealed to a surface 56 of a mold 58 with a seal 68, such as a sealant tape. As illustrated the seal 68 may pass over, for example, the vacuum line 64. A peel ply 70 may be used to, without limitation, leave a clean and contaminant free surface, and/or leave a textured surface. The peel ply 70 may be applied as the last material in the composite laminate sequence, and may be removed following a curing operation. A breather or bleeder fabric 72 may be used to, for example, to ensure that air or other gas sealed under the vacuum bag can be easily extracted and for more uniform vacuum distribution. The breather fabric 72 may also absorb excess resin that may be bled from the laminate.

FIGS. 9A and 9B illustrate cross-sectional views similar to FIG. 4 in accordance with various embodiments. Each view illustrates a first member 16 disposed at a distance 18 or 18′ from the first fabric 12, and coupled to the first fabric 12 with the second fabric 14. Each arrangement is illustrated between a substantially air tight film 60 and an outer mold line 58 surface 56. A comparison of FIGS. 9A and 9B may illustrate the versatility of various embodiments, as the same outer mold line 58 may be used for various end products by simply varying selected parameters of the design with little or no additional investment in tooling or training. The parameters that may be varied and may include but may not be limited to: the length and thickness of the fabrics 12, 14; the number of layers of fabric 12, 14; the distance 18 or 18′ of the first member 16 from the first fabric 12; the amount of resin used; and the pressure at which the resin is introduced. It will be understood that similar versatility can be demonstrated regarding the second member 34.

FIG. 10A is a front view and FIG. 10B is a side view taken through the line 10B-10B of FIG. 10A in accordance with various embodiments. A substantially elongate and substantially rigid first member 216 may be coupled with the first fabric 212 via a second fabric 214, and spaced a first distance 218 from the first fabric 212. In addition, a substantially elongate and substantially rigid second member 316 may be coupled with the first fabric 212 via a third fabric 314, and spaced a second distance 318 from the first fabric 212. A substantially rigid insert 334 may be disposed within the third fabric 314. The insert 334 may be adopted to support the first member 216, and may also be adopted to support and/or maintain the second member 316 at the distance 318. The insert 334 may be foam core 334. The third fabric 314 may include a hole 380 located to enable something to pass through, for example and without limitation, a pipe, wiring or conduit. The hole 380 may also be adapted as a window, or an opening for inspection and/or maintenance.

A resin may be infused into the first fabric 212, the second fabric 214 and third fabric 314, and may also encapsulate the first member 216 and the second member 316, to form a unitary structure 211.

FIG. 11A is a block diagram and FIG. 11B is a cross-sectional view taken at the line 11B-11B in FIG. 11A illustrating various embodiments. A composite structure 1010 may include a first fabric 1012, a second fabric 1014, and a substantially elongate and substantially rigid first member 1016 spaced apart from and coupled with the first fabric 1012 via the second fabric 1014. The composite structure 1010 may further include a second member 1034 disposed transverse to the first member 1016 and coupled with the first fabric 1012 via a third fabric 1036.

FIG. 12 illustrates a method 400 in accordance with various embodiments. The method may include a method of manufacturing a composite structure comprising: laying out a first fabric on a tool surface, 402; positioning an elongate first member a distance from the first fabric, 404; covering the first member with a second fabric such that tab-like portions of the second fabric contact the first fabric on opposite sides of the first member, 406; coupling the first member to the first fabric by securing the tab-like portions to the first fabric, 408; and forming a unitary structure with the first fabric, the second fabric, and the first member by infusing a resin into the first fabric and the second fabric and impregnating the first member within the resin, 410. The covering the first member with the second fabric may include positioning the second fabric in a slot of a fixture forming a valley with the second fabric and positioning the rod in the valley.

Various embodiments may provide that during any or all of the operations identified as 404, 406, and 408, any or all of the first member, second member, and first and second fabrics, may be tacked together using a tackifier. This may allow formation of more complex preforms in the steps prior to debalking, resin injection, and curing. Tacking together may be used along with, or instead of, stitching.

In various embodiments a method 500, shown in FIG. 13, may further comprise positioning a substantially airtight film or bag over the combination of the first member, the first fabric, and the second fabric, 502; and forming a vacuum between the tool surface and the film, and wherein the infusing includes introducing the resin between the tool surface and the film, 504.

Various embodiments may use operations that may be the same as or substantially similar to operations in methods described herein, or may be included with devices the same, or similar to, those described herein, that may utilize a pre-impregnated material already impregnated with a resin, also known as “prepreg”.

Although certain embodiments have been illustrated and described herein for purposes of description of the preferred embodiment, it will be appreciated by those of ordinary skill in the art that a wide variety of alternate and/or equivalent embodiments or implementations calculated to achieve the same purposes may be substituted for the embodiments shown and described. Those with skill in the art will readily appreciate that embodiments may be implemented in a very wide variety of ways. This application is intended to cover any adaptations or variations of the embodiments discussed herein. Therefore, it is manifestly intended that embodiments be limited only by the claims and the equivalents thereof.

Claims

1. A method of manufacturing a composite structure comprising: laying out a first fabric on a tool surface, the first fabric being a dry fabric;positioning an elongate and substantially rigid first member a distance from the first fabric;coupling the first member to the first fabric using a second fabric, the second fabric being a dry fabric, to create a preform, the preform being a self-supporting dry preform assembly that is configured to maintain a substantially semi-rigid state of equilibrium without the support of interior tooling features; andforming a unitary structure with the first fabric and the second fabric.

2. The method of claim 1 further comprising: covering the first member with the second fabric such that first tab-like portions of the second fabric contact the first fabric on opposite sides of the first member.

3. The method of claim 2 wherein the coupling the first member to the first fabric includes securing the first tab-like portions of the second fabric to the first fabric.

4. The method of claim 2 wherein the forming a unitary structure includes infusing a resin into the first fabric and the second fabric and encasing the first member within the resin.

5. The method of claim 2 wherein the covering the first member with the second fabric includes positioning the second fabric in a slot of a fixture forming a valley with the second fabric and subsequently positioning the first member in the valley.

6. The method of claim 3 further comprising: positioning a second member over the first fabric and substantially transverse to the first member;adding to the preform by coupling the second member to the first fabric using a third fabric, the third fabric being a dry fabric; andwherein forming the unitary structure further includes forming the unitary structure with the second member and the third fabric.

7. The method of claim 6 further comprising: covering the second member with the third fabric such that second tab-like portions of the third fabric contact the first fabric on opposite sides of the second member.

8. The method of claim 7 wherein the coupling the second member to the first fabric includes securing the second tab-like portions of the third fabric to the first fabric.

9. The method of claim 6 wherein the forming the unitary structure further includes infusing resin into the third fabric and encasing the second member within the resin.

10. The method of claim 3 wherein the securing the second fabric is by stitching.

11. The method of claim 8 wherein the securing the second fabric and the securing the third fabric is by stitching.

12. The method of claim 9 further comprising positioning an airtight film over the combination of the first member, the second member, the first fabric, the second fabric, and the third fabric; sealing edges of the airtight film to the tool surface; andforming a vacuum between the tool surface and the film, and wherein the infusing includes introducing the resin between the tool surface and the film.

13. The method of claim 9 further comprising curing the resin.

14. The method of claim 1 wherein the first member is a pultruded rod.

15. The method of claim 6 wherein the second member is foam core.

16. The method of claim 3 wherein the second fabric is a warp-knit fabric and further comprising orienting the warp of the warp-knit fabric substantially perpendicular to a longitudinal axis of the elongate first member.