Fly away from the longitudinal beam end cap

Abstract

The present invention relates to fly-off stringer end caps. Systems and methods for utilizing stringer end caps are provided. One embodiment is a method of forming a stringer. The method includes laying up a stringer preform including a fiber reinforcement material, placing the stringer preform onto a skin panel preform, bonding the end cap to the stringer preform and the skin panel preform, and co-curing the stringer preform and the skin panel preform while the end cap is bonded to the stringer preform and the skin panel preform, producing a composite part including the end cap.

Description

Technical Field

[0001] The present disclosure relates to the field of manufacturing, and more particularly to the manufacturing of stringers for aircraft. Background Art

[0002] The stringers for an aircraft (e.g., the wings of an aircraft) can be made of composite materials. These stringers can be laid up as preforms, machined into stringers, and then co-cured into the wing skin to form a complete wing or a portion thereof. When the stringers are made in this manner, the stringers are laid up and cured so that they terminate within the wing skin. Then after curing is complete, the stringers can be cut to the desired size with the desired slope. The transition from the stringer to the skin is often referred to as a "stringer transition section" and is used to transfer loads in a composite wing structure.

[0003] Manufacturing stringers made of composite materials in this manner may result in out-of-tolerance edge conditions, and out-of-tolerance conditions are undesirable because they may require rework. Also, it may be difficult to lay up and cure the preform at the desired slope due to the complex geometries that may be required.

[0004] It would therefore be desirable to have a method and apparatus that takes into account at least some of the above-mentioned issues as well as other possible issues. Summary of the invention

[0005] Embodiments described herein provide rigid end caps for stringers that present a desired slope and stringer transition section quality for terminating the stringers. These end caps are fixed to preforms for stringers to form an integral part of a composite component. The end caps eliminate the need to shape the preforms to a specific slope, while also eliminating the need to cut the stringers after curing. The end caps form a flyaway tooling that enforces the desired shape to the stringers during curing, while also bearing and transmitting loads within the wing after the wing has been manufactured. One embodiment is a method for forming a stringer. The method includes: laying a stringer preform including a fiber reinforced material, placing the stringer preform on a skin panel preform, bonding the end caps to the stringer preform and the skin panel preform, and co-curing the stringer preform and the skin panel preform while the end caps are bonded to the stringer preform and the skin panel preform, producing a composite component including the end caps.

[0006] Another embodiment is a non-transitory computer readable medium containing programming instructions that, when executed by a processor, are operable to perform a method of forming a stringer. The method includes laying up a stringer preform including a fiber reinforcement material, placing the stringer preform onto a skin panel preform, bonding an end cap to the stringer preform and the skin panel preform, and co-curing the stringer preform and the skin panel preform while the end cap is bonded to the stringer preform and the skin panel preform, producing a composite part including the end cap.

[0007] Another embodiment is an article. The article includes a composite component. The composite component includes a skin panel made of a fiber-reinforced material, a stringer of the fiber-reinforced material co-cured to the skin panel, and an end cap that is jointly bonded to the stringer and the skin panel, and the end cap includes a chamfer that overlaps the fiber-reinforced material from the stringer.

[0008] Another embodiment is an end cap for a longitudinal beam, which includes a flange, an inclined surface extending from the flange and gradually thinning the flange, a web protruding from the flange, and an inclined surface extending from the web and gradually thinning the web.

[0009] Another embodiment is a method for manufacturing a composite component, the method comprising: forming a skin panel preform, bonding a backing plate tool end to the skin panel preform, anchoring a stringer preform at the skin panel preform and the backing plate tool end, and co-curing the skin panel preform and the stringer preform with the backing plate tool end.

[0010] Another embodiment is an apparatus in the form of a flyaway tool for an aircraft. The apparatus includes an end cap including a flange, a ramp extending from the flange and tapering the flange, a web protruding from the flange, and a ramp extending from the web and tapering the web. The end cap supports a preform for a stringer during manufacture and is integrated into the finished stringer as a component of the aircraft.

[0011] According to one aspect of the present disclosure, a method for forming a stringer includes:

[0012] Laying a longitudinal beam preform (202) including a fiber-reinforced material;

[0013] Placing the stringer preform onto the skin panel preform (204);

[0014] bonding end caps to the stringer preform and the skin panel preform (206); and

[0015] The stringer preform and the skin panel preform are co-cured while the end cap is bonded to the stringer preform and the skin panel preform, thereby producing a composite part including the end cap (208).

[0016] Advantageously, the method is a method wherein:

[0017] The end cap is coupled to the outboard end of the longitudinal beam preform, and the method further comprises:

[0018] Another end cap is bonded to the inner end of the longitudinal beam preform.

[0019] Preferably, the method further comprises:

[0020] The inclined surface of the end cap is overlapped with the longitudinal beam preform.

[0021] Preferably, the method further comprises:

[0022] Loads are transferred to the end cap via a sloped surface selected from the group consisting of a step lap, a linear slope, a scarf, and a stagger.

[0023] Preferably, the method further comprises:

[0024] The composite part is assembled into a portion of the fuselage of an aircraft so that the end cap becomes a fly-away tool.

[0025] Preferably, the method further comprises:

[0026] The end cap is manufactured via additive manufacturing.

[0027] Preferably, the method further comprises:

[0028] Prior to the co-curing, the end caps, stringer preforms and skin panel preforms are vacuum bagged.

[0029] Preferably, the method is a method wherein:

[0030] The end caps are bonded to the outboard ends of the stringer preforms and secured to the wing panels.

[0031] Preferably, the method is a method wherein:

[0032] The end cap is bonded to the outboard end of the stringer preform and fastened to the wing panel, and the method further comprises:

[0033] Another end cap is bonded to the inboard end of the stringer preform, and the end cap is fastened to the wing panel.

[0034] Preferably, a portion of an aircraft is assembled according to the method as described above.

[0035] According to one aspect of the present disclosure, a non-transitory computer readable medium containing programming instructions, which when executed by a processor is operable to perform a method for forming a stringer, the method comprising:

[0036] Laying a longitudinal beam preform (202) including a fiber-reinforced material;

[0037] Placing the stringer preform onto the skin panel preform (204);

[0038] bonding end caps to the stringer preform and the skin panel preform (206); and

[0039] The stringer preform and the skin panel preform are co-cured while the end cap is bonded to the stringer preform and the skin panel preform, thereby producing a composite part including the end cap (208).

[0040] Advantageously, the medium is a medium in which:

[0041] The end cap is coupled to the outboard end of the longitudinal beam preform, and the method further comprises:

[0042] Another end cap is bonded to the inner end of the longitudinal beam preform.

[0043] Preferably, the medium is a medium wherein the method further comprises the following steps:

[0044] The inclined surface of the end cap is overlapped with the longitudinal beam preform.

[0045] Preferably, the medium is a medium wherein:

[0046] Loads are transferred to the end cap via a sloped surface selected from the group consisting of a step lap, a linear slope, a scarf, and a stagger.

[0047] Preferably, the medium is a medium wherein the method further comprises the following steps:

[0048] The composite part is assembled into a portion of the fuselage of an aircraft so that the end cap becomes a fly-away tool.

[0049] Preferably, the medium is a medium wherein the method further comprises the following steps:

[0050] The end cap is manufactured via additive manufacturing.

[0051] Preferably, the medium is a medium wherein the method further comprises the following steps:

[0052] Prior to said co-curing, said end caps, stringer preforms and skin panel preforms are vacuum bagged.

[0053] Preferably, the medium is a medium wherein:

[0054] The end caps are bonded to the outboard ends of the stringer preforms and secured to the wing panels.

[0055] Preferably, the medium is a medium wherein:

[0056] The end cap is bonded to the outboard end of the stringer preform and fastened to the wing panel, and the method further comprises:

[0057] Another end cap is bonded to the inboard end of the stringer preform, and the end cap is fastened to the wing panel.

[0058] Preferably, a portion of an aircraft is assembled according to a method defined by instructions stored on a computer readable medium.

[0059] According to one aspect of the present disclosure, an article includes:

[0060] A composite component (100), the composite component (100) comprising:

[0061] A skin panel (110) made of a fiber-reinforced material;

[0062] A stringer (120) made of the fiber reinforced material, the stringer (120) being co-cured to the skin panel; and

[0063] An end cap (130) is co-joined to the stringer and includes a chamfer (134) that is spliced ​​with fiber reinforcement material from the stringer.

[0064] Advantageously, the article is an article wherein:

[0065] The bevel includes at least one selected from the group consisting of a step-lap bevel, a recessed light, and a staggered.

[0066] Preferably, the article is an article wherein:

[0067] The end cap has a hat-shaped cross-section that is aligned with the hat-shaped cross-section of the longitudinal beam.

[0068] Preferably, the article is an article wherein:

[0069] The end caps are also commonly bonded to the skin panels.

[0070] Preferably, the article is an article wherein:

[0071] The bevel is spliced ​​with the fiber reinforcement material from the stringer.

[0072] Preferably, the article is an article wherein:

[0073] The end cap has a cross section in the form of a T, Z or J shape aligned with the cross section of the longitudinal beam.

[0074] Preferably, the article is an article wherein:

[0075] The end cap is made of at least one of carbon fiber reinforced polymer (CFRP) or metal.

[0076] Preferably, the article is an article wherein:

[0077] The fiber reinforced material includes carbon fiber reinforced polymer (CFRP).

[0078] A part of an aircraft is manufactured using the above-mentioned article.

[0079] According to one aspect of the present disclosure, there is provided an apparatus in the form of an end cap for a longitudinal beam, the end cap comprising:

[0080] Flange (320);

[0081] an inclined surface (322), the inclined surface (322) extending from the flange and gradually thinning the flange;

[0082] a web (310) protruding from the flange; and

[0083] A slope (312) extends from the web and gradually thins the web.

[0084] Advantageously, the device is one in which:

[0085] The cross-section of the end cap is selected from the group consisting of a "T" shape and a hat shape.

[0086] Preferably, the device is a device wherein:

[0087] The web forms a curved shape.

[0088] Preferably, the device is a device wherein:

[0089] The end cap is made of at least one of carbon fiber reinforced polymer (CFRP) or metal.

[0090] Preferably, the device is a device wherein:

[0091] The end cap includes mounting features in the flange.

[0092] Preferably, the device is a device wherein:

[0093] The end cap includes mounting features in the web.

[0094] A part of an aircraft is manufactured using the above-described apparatus.

[0095] According to one aspect of the present disclosure, a method for manufacturing a composite component comprises:

[0096] forming a skin panel preform (1202);

[0097] bonding a backing plate tool end to the skin panel preform (1204);

[0098] anchoring the stringer preform to the skin panel preform and the backing plate tool end (1206); and

[0099] The skin panel preform and the stringer preform are co-cured with the bolster tool end (1208).

[0100] Advantageously, the method further comprises:

[0101] The stringer preform, the tie plate tool end and the skin panel preform are vacuum bagged.

[0102] Preferably, the method further comprises:

[0103] The stringer preform, the tie plate tool end and the skin panel preform are hardened into a composite part.

[0104] Preferably, the method is a method wherein:

[0105] Curing includes curing the stringer preform, the bolster tool end, and the skin panel preform into a composite component while applying pressure.

[0106] Preferably, the method is a method wherein:

[0107] The tie plate tool end forms a fly-off tool that structurally supports the end of the stringer.

[0108] Preferably, the method is a method wherein:

[0109] The backing plate tool end comprises a material selected from the group consisting of metal and carbon fiber reinforced polymer (CFRP).

[0110] A portion of an aircraft assembled according to the method described above.

[0111] According to one aspect of the present disclosure, a device in the form of a fly-away tool for an aircraft comprises:

[0112] The end cover (300) comprises:

[0113] Flange (320);

[0114] an inclined surface (322), the inclined surface (322) extending from the flange and gradually thinning the flange;

[0115] a web (310) protruding from the flange; and

[0116] an inclined surface (312) extending from the web and gradually thinning the web,

[0117] The end caps support the prefabricated parts for the stringers during manufacture and are integrated into the finished stringers as components of the aircraft.

[0118] Advantageously, the device is one in which:

[0119] The web forms a curved shape.

[0120] Preferably, the device is a device wherein:

[0121] The end cap is made of at least one of carbon fiber reinforced polymer (CFRP) or metal.

[0122] Preferably, the device is a device wherein:

[0123] The end cap has a hat-shaped cross section.

[0124] Preferably, the device is a device wherein:

[0125] The end cap has a T-shaped cross section.

[0126] A part of an aircraft is manufactured using the apparatus described above.

[0127] Other illustrative embodiments (e.g., methods and computer-readable media related to the foregoing embodiments) may be described below. The features, functions, and advantages that have been discussed may be implemented independently in various embodiments, or may be combined in other embodiments, further details of which may be seen with reference to the following description and drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0128] Some embodiments of the present disclosure will now be described by way of example only and with reference to the accompanying drawings. The same reference number represents the same element or the same type of element throughout the drawings.

[0129] Figure 1 is a block diagram of one half of a wing including co-cured stringers and a fly-off tool in an illustrative embodiment.

[0130] Figure 2A is a flow chart showing a method for integrating a fly-off tool into a wing in an illustrative embodiment.

[0131] Figure 2B is a flow chart showing a method for integrating a fly-off tool into a wing in an illustrative embodiment.

[0132] Figure 3 and FIG. 4A to FIG. 4C An end cap is shown for a first stringer shape in an illustrative embodiment.

[0133] Figures 5 and 6 An end cap for a second stringer shape is shown in an illustrative embodiment.

[0134] 7A to 7E and Figure 8 The end caps and stringers are shown joined together in an illustrative embodiment.

[0135] Fig. 9 is a perspective view of a wing including a fly-off tool in an illustrative embodiment.

[0136] Fig.10 An inboard end cap for a longeron is shown in an illustrative embodiment.

[0137] Fig.11 is a block diagram showing an inboard end cap for a stringer integrated into a fuselage in an illustrative embodiment.

[0138] Fig.12 is yet another flow chart showing a method for integrating a fly-off tool into a wing in an illustrative embodiment.

[0139] Fig.13 is a flow chart of an aircraft production and service methodology in an illustrative embodiment.

[0140] Fig.14 is a block diagram of an aircraft in an illustrative embodiment. DETAILED DESCRIPTION

[0141] The accompanying drawings and the following description provide specific illustrative embodiments of the present disclosure. Therefore, it will be understood that those skilled in the art will be able to design various arrangements that, although not explicitly described or shown herein, embody the principles of the present disclosure and are included within the scope of the present disclosure. In addition, any examples described herein are intended to help understand the principles of the present disclosure and are to be interpreted as not being limited to such specifically described examples and conditions. Therefore, the present disclosure is not limited to the specific embodiments or examples described below, but is defined by the claims and their equivalents.

[0142] Composite parts, such as carbon fiber reinforced polymer (CFRP) parts, are initially laid up in multiple layers, which together are called preforms. The individual fibers within each layer of the preform are aligned parallel to each other, but different layers present different fiber orientations in order to increase the strength of the resulting composite part along different dimensions. The preform includes a viscous resin that solidifies to harden the preform into a composite part (for example, for use in an aircraft). Carbon fibers that have been impregnated with uncured thermosetting resins or thermoplastic resins are called "prepregs." Other types of carbon fibers include "dry fibers," which are not impregnated with thermosetting resins, but may include tackifiers or adhesives. Dry fibers are impregnated with resin before curing. For thermosetting resins, hardening is a unidirectional process called curing, while for thermoplastic resins, if the resin is reheated, the resin reaches a viscous form.

[0143] Figure 1 is a block diagram of the lower half of a wing 100 including co-cured stringers and fly-off tooling in an illustrative embodiment. Figure 1 The lower half of the wing 100 shown may comprise a co-cured lower half of the wing and can be fixed or bonded to the upper half of the wing to form a complete wing. In other embodiments, the upper half of the wing is assembled in a similar manner to the lower half of the wing 100. Figure 1 The lower half of the wing 100 shown includes a fiber reinforced composite material and rigid end caps for the stringers. This arrangement eliminates the need to trim or cut the stringers at the wing 100 after hardening.

[0144] In this embodiment, Figure 1 A portion of a wing is shown including a skin panel 110 that includes fiber reinforcement in the form of multiple layers 112 of resin 114 and fibers 116. A stringer 120 is disposed on top of the skin panel 110 and includes fiber reinforcement in the form of multiple layers 112 of fibers 116 and resin 114. The stringer 120 is co-cured to the skin panel 110.

[0145] The end caps 130 are adjacent to the inboard end 122 and the outboard end 124 of the longitudinal beam 120. The end caps 130 are rigid before curing and may include a metal that provides high strength and low weight, such as titanium or aluminum (e.g., insulated from carbon fiber by an insulating layer of glass fiber or other material). The end caps 130 may further include hardened composite materials (e.g., thermosets or thermoplastics) and 3D printed metals. Therefore, in one embodiment, the end caps 130 are manufactured via additive manufacturing techniques, such as 3D printing. In other embodiments, subtractive manufacturing techniques are utilized. Each end cap 130 includes a flange 132 having a chamfer 134 for receiving the longitudinal beam 120. The end caps 130 are bolted and / or bonded to the longitudinal beam 120 and / or the skin panel 110.

[0146] The ramps 134 provide a path for transferring loads from the stringers 120 to the end caps 130. As used herein, "ramp" refers to any physical structure that transitions loads along its length, including step laps, scarfs, staggers, linear ramps, and other features. The stringers 120 themselves include flanges and webs, and these structures are complementary to the ramps 134 of the end caps 130. Each end cap 130 also includes a ramp 136 that runs down to the skin panel 110. The ramps 136 provide a path for transferring loads from the end caps 130 to the skin panel 110.

[0147] In one embodiment, the end cap 130 is bonded to the stringer 120 and the skin panel 110 together, and the chamfer 134 of the end cap 130 overlaps the fiber reinforcement material from the stringer 120. In other embodiments, the chamfers 134 and / or 136 are integrated with the stringer 120 via lap joints, step lap joints, or interlacing of the chamfers with the plies of the stringer and / or the skin panel 110. In some of these embodiments, the transition involves laying the stringer preform on the end cap 130 to accommodate the difference in shape. The terms "skin panel 110" and "stringer 120" are used here to refer to uncured preforms and hardened composite parts. That is, the skin panel 110 may refer to an uncured preform for the skin panel waiting to be cured, or may refer to a hardened skin panel. In a similar manner, the stringer 120 may refer to an uncured preform for the stringer, or may refer to a hardened stringer.

[0148] Figure 1Also depicted are components that can be electronically managed by a controller 140 to manufacture the above structure. In this embodiment, the components include a laying machine 142, such as an automatic fiber placement (AFP) machine or a tape dispensing head, which lays tows of unidirectional fiber reinforcement material to form a layer 112. The components also include a pick and place (PNP) machine 144 (e.g., an end effector, a suction device, a gripper, etc.), which picks up the stringer preform and / or the end cap 130 and places the stringer preform and / or the end cap 130 on the skin panel 110. In other embodiments, the stringer preform and the end cap 130 are manually picked up and placed in place by one or more technicians. The controller 140 directs the operation of these components based on instructions in one or more numerical control (NC) programs stored in a memory, and can be implemented as, for example, a custom circuit system, a hardware processor that executes programmed instructions, or some combination thereof.

[0149] About Figure 2A discuss Figure 1 For this embodiment, it is assumed that a skin panel preform of a wing (eg, skin panel 110 ) has been laid up by layup machine 142 .

[0150] Figure 2A is a flow chart showing a method 200 for integrating a flyaway tool into a wing in an illustrative embodiment. The steps of the method 200 are described with reference to Figure 1 Although the method 200 is described with reference to the lower half of the wing 100 shown, those skilled in the art will appreciate that the method 200 may be performed on other parts of the wing (e.g., the upper half of the wing) in other environments. The steps of the flowchart described herein are not all-inclusive and may include other steps not shown. The steps described herein may also be performed in an alternative order.

[0151] In step 202, the controller 140 directs the layup machine 142 to lay up the stringer preform. In one embodiment, this includes applying multiple layers of unidirectional fiber reinforcement material to a layup mandrel (not shown) or other tooling piece that defines the shape of the stringer preform.

[0152] In one embodiment, the PNP machine 144 places an end cap 130 at each end of the stringer preform. The end cap 130 helps to enforce the desired shape at the stringer preform before and during curing. In addition, the end cap 130 provides a ramp 136 (e.g., for a transition section) and / or other complex geometries in a rigid form, which means that these geometries do not need to be mechanically supported during the curing process. This reduces the complexity of the layup and curing of the wing, which reduces the costs associated with labor and materials.

[0153] In step 204, the controller 140 instructs the PNP machine 144 to place the stringer preform onto the skin panel preform (e.g., skin panel 110). During this operation, the stringer preform has not yet solidified (i.e., is still in a "green state") and therefore remains loose. Therefore, when transporting the stringer preform, the PNP machine 144 can force or maintain the desired curvature at the stringer preform by applying suction (e.g., via a vacuum connection) or using a support structure. In one embodiment, the PNP machine 144 picks up a plurality of stringer preforms and places them at the skin panel preform. This can also include placing the stringer preform on a laying mandrel and then laying the skin against the stringer preform. For the skin panel defining the upper portion of the wing, the stringer may include a preform for a hat-shaped stringer. For the skin panel defining the lower portion of the wing, the stringer may include a preform for a "T"-shaped stringer. Other types of stringers include Z-shaped stringers and stringers of any suitable cross-section. In one embodiment, the placement operation involves overlapping the chamfer 134 of the end cap 130 with one or more layers of the stringer preform. This may include overlapping the chamfer 134 with the chamfer at the stringer preform. In this way, after the stringer preform hardens, the chamfer 134 transfers loads between the hardened stringer and the end cap 130. At the same time, the chamfer 136 transfers loads between the end cap 130 and the skin panel 110.

[0154] Step 206 includes bonding the end caps 130 to the stringer preform and the skin panel preform. In one embodiment, this includes applying an adhesive (e.g., epoxy, glue, or other self-hardening chemical) to the end caps 130 before placing the end caps or stringer preforms on the skin panel preforms, and waiting for the adhesive to harden after the end caps and stringer preforms are placed in place. In one embodiment, an end cap 130 is bonded to the outboard end of a stringer preform, and the other end cap 130 is bonded to the inboard end of a stringer preform.

[0155] After the end caps 130 have been incorporated into place, the position of the end caps 130 relative to the stringer preforms and the skin panel preforms is held in place by the hardened adhesive. This ensures that the vacuum bagging setup and consolidation does not shift the positions of the stringer preforms, the skin panel preforms, and the end caps relative to each other. Therefore, the PNP machine 144, another machine, or a technician can continue to vacuum bag the end caps, the stringer preforms, and the skin panel preforms (i.e., before these elements are co-cured). The end caps both anchor the stringer preforms to the skin panel preforms and position the stringers at the skin panel preforms. In one embodiment, vacuum bags are used to consolidate these components by applying pressure prior to curing.

[0156] Step 208 includes co-curing the stringer preform to the skin panel preform while the end cap 130 is bonded to the stringer preform and the skin panel preform, thereby producing a composite part including the end cap 130 as an integral part. In one embodiment, co-curing includes placing the vacuum-bagged part into an autoclave, applying heat via the autoclave until the resin 114 reaches a curing temperature, and applying pressure via the vacuum bag and / or via the autoclave to consolidate and cure the end cap 130, the stringer preform, and the skin panel preform. These parts are co-cured into an integral composite part including an integral fly-away tool. That is, the end cap 130 operates as a tool to provide support for the preform during curing and vacuum bagging, and also provides mechanical strength when the resulting portion of the wing is assembled into a portion of the fuselage of the aircraft. The tool becomes physically integrated with the stringer after hardening. Therefore, the tool is a fly-away tool because it is integrated into the aircraft and "flies away" as part of the aircraft after manufacturing is completed.

[0157] The method 200 provides a technical benefit over the prior art because it enables a runout to be quickly integrated into a composite stringer, providing support during vacuum bagging and curing. This is because the end caps, which are rigid prior to curing, resist the compressive forces applied by the vacuum bag, which could crush or bend elongated portions, such as the web of the stringer preform. Thus, the end caps help constrain the stringer preform to a desired shape during curing. The method 200 additionally eliminates the need to cut or remove material from the composite stringer after it has been co-cured to the skin panel.

[0158] Figure 2B A method 250 for integrating a fly-off tool into a wing in an illustrative embodiment is shown. The method 250 includes forming (e.g., laying up) a skin panel preform including a fiber reinforcement material (e.g., CFRP) in step 252, providing a rigid end cap at the skin panel preform at an end position of a stringer preform (e.g., laid up from CFRP) to be placed at the skin panel in step 254, positioning the stringer preform at the skin panel preform via the rigid end cap in step 256, and anchoring the stringer preform to the skin panel preform in step 258.

[0159] In another embodiment, the method 250 includes consolidating the stringer preform and the skin panel preform via a vacuum bag covering the stringer preform, the skin panel preform, and the end cap. The method 250 may also include hardening the stringer preform and the skin panel preform to form a section of the wing including the end cap. The end cap is bonded to the stringer preform and the skin panel preform, and in one embodiment, the method 250 also includes co-curing the stringer preform and the skin panel preform while the rigid end cap is bonded to the stringer preform and the skin panel preform, thereby producing a composite part including the end cap.

[0160] Figure 3 and FIG. 4A to FIG. 4C An end cap 300 is shown for a first stringer shape in the illustrative embodiment. Figure 4A and Figure 3 In this embodiment, the end cap 300 is designed as an end cap for a longitudinal beam having a T-shaped cross section. That is, the end cap 300 has a T-shaped cross section aligned with the T-shaped cross section of the longitudinal beam.

[0161] The end cap 300 includes a web 310 that is narrowed via a bevel 312 that extends from the web 310 and tapers the web 310. In this embodiment, the web 310 forms a vertical plane. However, in other embodiments, the web forms a curved shape (e.g., Figure 5 ). The web 310 protrudes from the lower flange 320. The lower flange 320 is narrowed via a chamfer 322 that extends from and tapers the lower flange 320. These chamfers 312 and 322 overlap via corresponding chamfers in the longeron preform so that the combined total thickness of the flange and longeron preform remains constant along the flange. In other embodiments, the chamfers present a stepped pattern that accommodates a stepped overlap or other type of interface / transition between the longeron and the end cap. In other embodiments, staggering of composite plies with metal structure is performed to narrow the structure rather than narrowing in a linear, inclined manner. The chamfer 330 terminates the longeron according to the desired structural constraints and is shaped and tapered to help transition the longeron loads through the end cap to the panel. The chamfer 340 transfers forces through the end cap 300 into the skin panel (not shown). At the same time, ramps 312 running into web 310 and ramps 322 running into lower flange 320 transfer loads from the stringer. The rate of transition of loads through the end cap is established by the geometry of the ramps and the pattern of the ramps (eg, linear, stepped, etc.).

[0162] Figure 4B Depicted after the stringer 400 has been bonded to the end caps Figure 3 A cross-sectional cut side view of the end cap 300 . Figure 4B and Figure 4A The observation arrow 4B corresponds to. Figure 4B , it can be seen that the longitudinal beam 400 includes a bevel 410 and a bevel 420 that cooperate with the bevels 312 and 322, respectively. Figure 4C is with Figure 4A The cross-sectional cut end view corresponding to the observation arrow 4C. Figure 4C, the chamfers 312 and 322 of the end cap 300 are visible, as are the chamfers 410 and 420 of the stringer 400. In a further embodiment, the end cap is bonded to the outboard end of the stringer preform and secured to the wing panel, and the other end cap is bonded to the inboard end of the stringer preform and secured to the wing panel.

[0163] Figures 5 and 6 An end cap 500 is shown for a second stringer shape in the illustrative embodiment. Figure 6 and Figure 5 The end cap 500 corresponds to the observation arrow 6. In this embodiment, the end cap 500 is designed as an end cap for a longitudinal beam having a "hat" shaped cross section. That is, the end cap 500 has a hat-shaped cross section aligned with the hat-shaped cross section of the longitudinal beam.

[0164] The end cap 500 includes an upper arch 510 that narrows via a chamfer 512, and also includes a lower flange 520 that narrows via a chamfer 522. These flanges overlap via corresponding chamfers in the stringer preform so that the total thickness of the combination of the flange and the stringer preform remains constant along the flange. In other embodiments, the chamfers present a step pattern as they narrow, rather than narrowing in a linear manner. The chamfers 530 terminate the stringer according to the desired structural constraints, and the lip 540 distributes the forces borne by the end cap 500 into the skin panel (not shown). In this embodiment, the upper arch 510 forms a void 550 that corresponds to the void in the hat stringer.

[0165] 7A to 7E and Figure 8 The end cap 500 is shown in the illustrative embodiment having been co-bonded to a longeron 700 . Figure 8 is with Fig. 7A 8. In this embodiment, the chamfers 512 and 522 are covered by the fiber reinforcement material from the stringer 700, which forms a corresponding instance of the chamfer 710 to maintain the desired combined thickness of the stringer and the end cap.

[0166] Figure 7B is with Fig. 7A The cross-sectional cut side view corresponding to the viewing arrow 7B and showing the chamfer 710 at the scarf joint 720 Fig. 7A The inclined plane 522 intersects. Figure 7B In the embodiment, the end cap has been bonded to the inclined surface of the longitudinal beam 700, which structurally integrates the longitudinal beam 700 with the end cap 500 to form a single component made of the longitudinal beam and the end cap. Figure 7B The scarf-type joint between the inclined surface of the end cap and the composite sheet is shown. However, a step-type joint can also be achieved, such as Figure 7C In another embodiment, the end caps are interlaced with the composite sheets, such as Fig.7D 740. Other implementations of staggering may also or alternatively be utilized.

[0167] Fig. 7E is with Fig. 7A The cross-sectional cut end view corresponding to the viewing arrow 7E is shown. This view further depicts the intersection between the inclined surface for the end cap and the inclined surface for the longitudinal beam. Fig. 7E As shown, a portion of the stringer 700 overlaps the inclined surface 512 and the inclined surface 522 of the end cap 500 .

[0168] Figure 8 Depicted with Fig. 7A The side view corresponding to the observation arrow 8. Figure 8 , it can be observed that most of the join (ie, the page entering between bevel 710 and bevel 522) is hidden and not visible.

[0169] Fig. 9 is a perspective view of a wing 900 including a fly-off tool in an illustrative embodiment. The tool includes end caps that facilitate the hardening of the stringers by supporting the stringers during curing and become physically integral to the stringers after hardening. Thus, the tool is a fly-off tool in that it is integrated into the aircraft and "flies off" as part of the aircraft after manufacturing is complete. Thus, each end cap serves a dual purpose as a tool to support the preform prior to and during curing, and also as a component that transfers loads from the stringers into the panels when integrated into the aircraft.

[0170] Fig. 9 The wing 900 is shown to include a leading edge 902, a trailing edge 904, and a hanger 906. The wing 900 also includes a longeron 920 and an inboard end cap 910 and an outboard end cap 912 disposed at the inboard and outboard ends of the longeron 920. For example, the inboard end cap 910 may terminate the longeron 920 at the intersection of the sides of the body to form the inboard end of the longeron 920, and the outboard end cap 912 may terminate the outboard end of the longeron 920. In one embodiment, the outboard end caps are smaller than the inboard end caps because they match the tapered portion of the longeron 920. In such an embodiment, the inboard end caps do not taper into the skin panel, but are instead bonded or fastened to the wing box. However, such end caps still taper into the longeron itself. Because Fig. 9 A simplified view is provided so that it can be understood that there may be many more stringers than shown, and that the stringers may be much narrower relative to the size of the wing. For example, the stringers may taper from the inboard end to the outboard end. In addition, some of the stringers may continue to the end of the wing, while others terminate earlier.

[0171] Fig.10An inboard end cap 1000 for a longitudinal beam in an illustrative embodiment is shown. In this embodiment, the inboard end cap 1000 is designed as an end cap for a longitudinal beam having a T-shaped cross section. That is, the end cap 1000 has a T-shaped cross section that is aligned with the T-shaped cross section of the inboard end of the longitudinal beam.

[0172] The inner end cap 1000 includes a web 1010 that is narrowed via a chamfer 1012 that extends from and tapers the web 1010. In this embodiment, the web 1010 forms a vertical plane. The web 1010 protrudes from a lower flange 1020. The lower flange 1020 is narrowed via a chamfer 1022 that extends from and tapers the lower flange 1020. These chamfers 1012 and 1022 overlap / join with corresponding chamfers in the stringer preform so that the total thickness of the combination of the flange and the stringer preform remains constant along the flange at the end 1080. In other embodiments, the chamfers present a stepped pattern that accommodates a step overlap or other type of interface / transition between the stringer and the end cap. In other embodiments, the staggering of the composite plies with the metal structure is performed to narrow the structure, rather than narrowing in a linear, inclined manner. The web extension 1030 terminates the stringer by abutting against the side intersection of the body at end 1080 to transfer the load from the wing box. At the same time, the ramp 1012 running into the web 1010 and the ramp 1022 running into the lower flange 1020 transfer the load from the stringer to the center of the wing box. The rate of transition of the load through the end cap is established by the geometry of the ramp and the pattern of the ramp and web extension (e.g., linear, stepped, etc.).

[0173] Fig.10 Also shown are mounting features 1050 in the web and 1060 in the flange at the inboard end cap 1000. These mounting features (e.g., bolt holes for fasteners, protrusions, etc.) facilitate the process of securing the inboard end cap 1000 to the center wing box.

[0174] Fig.11 is a block diagram showing an inboard end cap of a stringer integrated into a fuselage in an illustrative embodiment. Fig.11 11, the sides of the main body 1110 of the fuselage are adjacent to the inner end cap 1112. The stringer 1120 is attached to the inner end cap 1112 and is also attached to the outer end cap 1114. The stringer 1120 and the end cap are also fixed to the skin panel 1104.

[0175] Fig.121 is another flow chart showing a method for integrating a flyaway tool into a wing in an illustrative embodiment. Step 1202 includes forming a skin panel preform via layup or other techniques. Step 1204 includes bonding a pad tool end (e.g., an end cap) to the skin panel preform. Step 1206 includes anchoring a stringer preform at the skin panel preform and the pad tool end. Step 1208 includes hardening the skin panel preform and the stringer preform with the pad tool end. That is, when the skin panel preform and the stringer preform are supported by the pad tool end, they are hardened into a composite component. In a further embodiment, vacuum bagging is performed before hardening, and the process includes vacuum bagging the stringer preform, the pad tool end, and the skin panel preform. In one embodiment, hardening includes curing the stringer preform, the pad tool end, and the skin panel preform into a composite component while applying pressure. In a further embodiment, the tie plate tool end forms a fly-off tool that structurally supports the end of the stringer. In a further embodiment, the tie plate tool end comprises a material selected from the group consisting of metal and carbon fiber reinforced polymer (CFRP).

[0176] Example

[0177] In the following embodiments, additional processes, systems, and methods are described in the context of end caps used as fly-off tools for stringers.

[0178] With reference to the accompanying drawings in more detail, embodiments of the present disclosure may be Fig.12 The method 1200 shown and Fig.13 1202 is depicted in the context of aircraft manufacturing and service. During pre-production, method 1200 may include specification and design 1204 of aircraft 1202 and material procurement 1206. During production, component and subassembly manufacturing 1208 and system integration 1210 of aircraft 1202 occurs. Thereafter, aircraft 1202 may undergo certification and delivery 1212 for entry into service 1214. While in service by a customer, aircraft 1202 is placed on a routine basis for maintenance and service 1216 (which may also include modification, rebuild, refurbishment, etc.). The apparatus and methods presented herein may be employed during any one or more appropriate stages of production and service described in method 1200 (e.g., specification and design 1204, material procurement 1206, component and subassembly manufacturing 1208, system integration 1210, certification and delivery 1212, service 1214, maintenance and repair 1216) and / or any appropriate component of aircraft 1202 (e.g., airframe 1218, systems 1220, interior 1222, propulsion system 1224, electrical system 1226, hydraulic system 1228, environmental system 1230).

[0179] Each process of method 1200 may be performed or executed by a system integrator, a third party, and / or an operator (e.g., a customer). For purposes of this specification, a system integrator may include, but is not limited to, any number of aircraft manufacturers and main system subcontractors; a third party may include, but is not limited to, any number of suppliers, subcontractors, and vendors; and an operator may be an airline, a leasing company, a military entity, a maintenance organization, etc.

[0180] like Fig.12 As shown, an aircraft 1202 produced by method 1200 may include a fuselage 1218 and an interior 1222 having a plurality of systems 1220. Examples of systems 1220 include one or more of a propulsion system 1224, an electrical system 1226, a hydraulic system 1228, and an environmental system 1230. Any number of other systems may be included. Although an aerospace example is shown, the principles of the present invention may be applied to other industries, such as the automotive industry.

[0181] As described above, the apparatus and methods presented herein may be employed during any one or more stages of production and maintenance described in method 1200. For example, components or subassemblies corresponding to component and subassembly manufacturing 1208 may be manufactured or produced in a manner similar to components or subassemblies produced while aircraft 1202 is in service. Moreover, during subassembly manufacturing 1208 and system integration 1212, one or more apparatus embodiments, method embodiments, or a combination thereof may be utilized, for example, by substantially expediting assembly of aircraft 1202 or reducing its cost. Similarly, one or more apparatus embodiments, method embodiments, or a combination thereof may be utilized while aircraft 1202 is in service, for example and without limitation, during maintenance and repair 1216. Thus, the present invention may be used in any of the stages discussed herein, or any combination thereof, such as specification and design 1204, material procurement 1206, component and subassembly manufacturing 1208, system integration 1210, certification and delivery 1212, service 1214, maintenance and repair 1216, and / or any suitable component of aircraft 1202 (e.g., airframe 1218, systems 1220, interior 1222, propulsion system 1224, electrical system 1226, hydraulic system 1228, and / or environmental system 1230).

[0182] In one embodiment, the component comprises a portion of an airframe 1218 and is manufactured during component and subassembly manufacturing 1208. The component may then be assembled into an aircraft in system integration 1210 and then used in service 1214 until wear renders the component unusable. The component may then be discarded and replaced with a newly manufactured component in maintenance and repair 1216. The components and methods of the present invention may be used throughout component and subassembly manufacturing 1208 to manufacture new components.

[0183] Any of the various control elements (e.g., electrical or electronic components) shown in the drawings or described herein may be implemented as hardware, a processor implementing software, a processor implementing firmware, or some combination thereof. For example, an element may be implemented as dedicated hardware. A dedicated hardware element may be referred to as a "processor," "controller," or some similar terms. When provided by a processor, the function may be provided by a single dedicated processor, a single shared processor, or a plurality of separate processors, some of which may be shared. In addition, the explicit use of the term "processor" or "controller" should not be interpreted as exclusively referring to hardware capable of executing software, and may implicitly include, but is not limited to, digital signal processor (DSP) hardware, network processor, application specific integrated circuit (ASIC) or other circuit system, field programmable gate array (FPGA), read-only memory (ROM), random access memory (RAM), non-volatile storage device, logic, or some other physical hardware components or modules for storing software.

[0184] In addition, the control element can be implemented as an instruction that can be executed by a processor or computer to perform the function of the element. Some examples of instructions are software, program code, and firmware. When executed by a processor, the instructions are operable to direct the processor to perform the function of the element. The instructions can be stored on a storage device that can be read by the processor. Some examples of storage devices are digital or solid-state memory, magnetic storage media such as disks and tapes, hard drives, or optically readable digital data storage media.

[0185] This article also provides the following terms, which relate to:

[0186] 1. A method for forming a longitudinal beam, the method comprising:

[0187] Laying a longitudinal beam preform (202) including a fiber-reinforced material;

[0188] Placing the stringer preform onto the skin panel preform (204);

[0189] bonding end caps to the stringer preform and the skin panel preform (206); and

[0190] The stringer preform and the skin panel preform are co-cured while the end cap is bonded to the stringer preform and the skin panel preform, thereby producing a composite part including the end cap (208).

[0191] 2. The method of clause 1, wherein:

[0192] The end cap is coupled to the outboard end of the longitudinal beam preform, and the method further comprises:

[0193] Another end cap is bonded to the inner end of the longitudinal beam preform.

[0194] 3. The method of clause 1 or 2, further comprising:

[0195] The inclined surface of the end cap is overlapped with the longitudinal beam preform.

[0196] 4. The method of any one of clauses 1 to 3, further comprising:

[0197] Loads are transferred to the end cap via a sloped surface selected from the group consisting of a step lap, a linear slope, a scarf, and a stagger.

[0198] 5. The method of any one of clauses 1 to 4, further comprising:

[0199] The composite part is assembled into a portion of the fuselage of an aircraft so that the end cap becomes a fly-away tool.

[0200] 6. The method of any one of clauses 1 to 5, further comprising:

[0201] The end cap is manufactured via additive manufacturing.

[0202] 7. The method of any one of clauses 1 to 6, further comprising:

[0203] Prior to said co-curing, said end caps, stringer preforms and skin panel preforms are vacuum bagged.

[0204] 8. The method of any one of clauses 1 to 7, wherein:

[0205] The end caps are bonded to the outboard ends of the stringer preforms and secured to the wing panels.

[0206] 9. The method of any one of clauses 1 to 8, wherein:

[0207] The end cap is bonded to the outboard end of the stringer preform and fastened to the wing panel, and the method further comprises:

[0208] Another end cap is bonded to the inboard end of the stringer preform, and the end cap is fastened to the wing panel.

[0209] 10. A part of an aircraft assembled according to the method of any of clauses 1 to 9.

[0210] 11. A non-transitory computer readable medium containing programming instructions, the programming instructions being operable when executed by a processor to perform a method of forming a stringer, the method comprising:

[0211] Laying a longitudinal beam preform (202) including a fiber-reinforced material;

[0212] Placing the stringer preform onto the skin panel preform (204);

[0213] bonding end caps to the stringer preform and the skin panel preform (206); and

[0214] The stringer preform and the skin panel preform are co-cured while the end cap is bonded to the stringer preform and the skin panel preform, thereby producing a composite part including the end cap (208).

[0215] 12. The medium of clause 11, wherein:

[0216] The end cap is coupled to the outboard end of the longitudinal beam preform, and the method further comprises:

[0217] Another end cap is bonded to the inner end of the longitudinal beam preform.

[0218] 13. The medium of clauses 11 to 12, wherein the method further comprises:

[0219] The inclined surface of the end cap is overlapped with the longitudinal beam preform.

[0220] 14. The medium of any one of clauses 11 to 13, wherein:

[0221] Loads are transferred to the end cap via a sloped surface selected from the group consisting of a step lap, a linear slope, a scarf, and a stagger.

[0222] 15. The medium of any of clauses 11 to 14, wherein the method further comprises:

[0223] The composite part is assembled into a portion of the fuselage of an aircraft so that the end cap becomes a fly-away tool.

[0224] 16. The medium of any of clauses 11 to 15, wherein the method further comprises:

[0225] The end cap is manufactured via additive manufacturing.

[0226] 17. The medium of any of clauses 11 to 16, wherein the method further comprises:

[0227] Prior to said co-curing, said end caps, stringer preforms and skin panel preforms are vacuum bagged.

[0228] 18. The medium of any one of clauses 11 to 17, wherein:

[0229] The end caps are bonded to the outboard ends of the stringer preforms and secured to the wing panels.

[0230] 19. The medium of any one of clauses 11 to 18, wherein:

[0231] The end cap is bonded to the outboard end of the stringer preform and fastened to the wing panel, and the method further comprises:

[0232] Another end cap is bonded to the inboard end of the stringer preform, and the end cap is fastened to the wing panel.

[0233] 20. A portion of an aircraft assembled according to a method defined by instructions stored on a computer-readable medium as described in any of clauses 11 to 19.

[0234] 21. An article, comprising:

[0235] A composite component (100), the composite component (100) comprising:

[0236] A skin panel (110) of fiber-reinforced material;

[0237] A stringer (120) of fiber reinforced material, the stringer (120) being co-cured to the skin panel; and

[0238] An end cap (130) is co-joined to the stringer and includes a chamfer (134) that is spliced ​​with fiber reinforcement material from the stringer.

[0239] 22. The article of manufacture of clause 21, wherein:

[0240] The bevel includes at least one selected from the group consisting of a step-lap bevel, a recessed light, and a staggered.

[0241] 23. The article of manufacture of clause 22 or 23, wherein:

[0242] The end cap has a hat-shaped cross-section that is aligned with the hat-shaped cross-section of the longitudinal beam.

[0243] 24. The article of any one of clauses 21 to 23, wherein:

[0244] The end caps are also commonly bonded to the skin panels.

[0245] 25. The article of any one of clauses 21 to 24, wherein:

[0246] The bevel is spliced ​​with the fiber reinforcement material from the stringer.

[0247] 26. The article of any one of clauses 21 to 25, wherein:

[0248] The end cap has a cross section in the form of a T-shape, a Z-shape or a J-shape aligned with the cross section of the longitudinal beam.

[0249] 27. The article of any one of clauses 21 to 26, wherein:

[0250] The end cap is made of at least one of carbon fiber reinforced polymer (CFRP) or metal.

[0251] 28. The article of any one of clauses 21 to 27, wherein:

[0252] The fiber reinforced material includes carbon fiber reinforced polymer (CFRP).

[0253] 29. Use of an article according to any one of clauses 21 to 28 to manufacture a part of an aircraft.

[0254] 30. An apparatus in the form of an end cap for a longitudinal beam, the end cap comprising:

[0255] Flange (320);

[0256] an inclined surface (322), the inclined surface (322) extending from the flange and gradually thinning the flange;

[0257] a web (310) protruding from the flange; and

[0258] A slope (312) extends from the web and gradually thins the web.

[0259] 31. The apparatus of clause 30, wherein:

[0260] The cross-section of the end cap is selected from the group consisting of a "T" shape and a hat shape.

[0261] 32. The apparatus of clause 30 or 31, wherein:

[0262] The web forms a curved shape.

[0263] 33. The apparatus of any one of clauses 30 to 32, wherein:

[0264] The end cap is made of at least one of carbon fiber reinforced polymer (CFRP) or metal.

[0265] 34. The apparatus of any of clauses 30 to 33, wherein:

[0266] The end cap includes mounting features in the flange.

[0267] 35. The apparatus of any one of clauses 30 to 34, wherein:

[0268] The end cap includes mounting features in the web.

[0269] 36. Use of an apparatus according to any of clauses 30 to 35 to manufacture a part of an aircraft.

[0270] 37. A method for manufacturing a composite component, the method comprising:

[0271] forming a skin panel preform (1202);

[0272] bonding a backing plate tool end to the skin panel preform (1204);

[0273] anchoring the stringer preform to the skin panel preform and the backing plate tool end (1206); and

[0274] The skin panel preform and the stringer preform are co-cured with the bolster tool end (1208).

[0275] 38. The method of clause 37, further comprising:

[0276] The stringer preform, the tie plate tool end and the skin panel preform are vacuum bagged.

[0277] 39. The method of clause 37 or 38, further comprising:

[0278] The stringer preform, the tie plate tool end and the skin panel preform are hardened into a composite part.

[0279] 40. The method of any one of clauses 37 to 39, wherein:

[0280] Curing includes curing the stringer preform, the bolster tool end, and the skin panel preform into a composite component while applying pressure.

[0281] 41. The method of any one of clauses 37 to 40, wherein:

[0282] The tie plate tool end forms a fly-off tool that structurally supports the end of the stringer.

[0283] 42. The method of any one of clauses 37 to 41, wherein:

[0284] The backing plate tool end comprises a material selected from the group consisting of metal and carbon fiber reinforced polymer (CFRP).

[0285] 43. A part of an aircraft assembled according to the method of any of clauses 37 to 42.

[0286] 44. An apparatus in the form of a fly-away tool for an aircraft, the apparatus comprising:

[0287] The end cover (300) comprises:

[0288] Flange (320);

[0289] an inclined surface (322), the inclined surface (322) extending from the flange and gradually thinning the flange;

[0290] a web (310) protruding from the flange; and

[0291] A slope (312) extends from the web and gradually thins the web.

[0292] The end caps support the prefabricated part of the stringer during manufacture and are integrated into the finished stringer as a component of the aircraft.

[0293] 45. The apparatus of clause 44, wherein:

[0294] The web forms a curved shape.

[0295] 46. ​​The apparatus of clause 44 or 45, wherein:

[0296] The end cap is made of at least one of carbon fiber reinforced polymer (CFRP) or metal.

[0297] 47. The apparatus of any of clauses 44 to 46, wherein:

[0298] The end cap has a hat-shaped cross section.

[0299] 48. The apparatus of any of clauses 44 to 47, wherein:

[0300] The end cap has a T-shaped cross section.

[0301] 49. Use of the apparatus of any one of clauses 44 to 48 to manufacture a part of an aircraft.

[0302] Although specific embodiments are described herein, the scope of the present disclosure is not limited to those specific embodiments.The scope of the present disclosure is defined by the appended claims and any equivalents thereof.

Claims

1. A method for forming a longitudinal beam, the method comprising: Laying down of prefabricated longitudinal beams including fiber reinforcement materials; placing the longitudinal beam prefabricated part on the skin panel prefabricated part; bonding end caps to the stringer preform and the skin panel preform; as well as co-curing the stringer preform and the skin panel preform while the end cap is bonded to the stringer preform and the skin panel preform, thereby producing a composite part including the end cap; The method further comprises causing the inclined surface of the end cover to overlap with the longitudinal beam preform.

2. The method according to claim 1, wherein: The end cap is coupled to the outboard end of the longitudinal beam preform, and the method further comprises: Another end cap is bonded to the inner end of the longitudinal beam preform.

3. The method according to claim 1 or 2, further comprising: Loads are transferred to the end cap via a sloped surface selected from the group consisting of a step lap, a linear slope, a scarf, and a stagger.

4. The method according to claim 1 or 2, further comprising: The composite part is assembled into a portion of the fuselage of an aircraft so that the end cap becomes a fly-away tool.

5. The method according to claim 1 or 2, further comprising: The end cap is manufactured via additive manufacturing.

6. The method according to claim 1 or 2, further comprising: Prior to the co-curing, the end caps, stringer preforms and skin panel preforms are vacuum bagged.

7. The method according to claim 1 or 2, wherein: The end cap is bonded to the outboard end of the stringer preform and fastened to the wing panel, and the method further comprises: Another end cap is bonded to the inboard end of the stringer preform, and the end cap is fastened to the wing panel.

8. A non-transitory computer readable medium containing programming instructions operable, when executed by a processor, to perform the method of claim 1.

9. A product, comprising: A composite component (100), the composite component (100) comprising: Skin panels (110) made of fiber-reinforced material; A longitudinal beam (120) made of the fiber reinforced material, the longitudinal beam (120) being co-cured to the skin panel; and An end cap (130) is commonly bonded to the stringer and includes a chamfer (134) that is spliced ​​with fiber reinforcement material from the stringer.

10. The article of claim 9, wherein: The bevel includes at least one selected from the group consisting of a step-lap bevel, a recessed light, and a staggered.

11. The article according to claim 9 or 10, wherein: The end cap has a hat-shaped cross-section that is aligned with the hat-shaped cross-section of the longitudinal beam.

12. The article according to claim 9 or 10, wherein: The end caps are also commonly bonded to the skin panels.

13. The article according to claim 9 or 10, wherein: The end cap has a cross section in the form of a T-shape, a Z-shape or a J-shape aligned with the cross section of the longitudinal beam.

14. The article of claim 9 or 10, wherein: The end cap is made of at least one of carbon fiber reinforced polymer (CFRP) or metal.

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