Patent Publication Number: US-2021178706-A1

Title: Flyaway stringer end caps

Description:
FIELD 
     The disclosure relates to the field of fabrication, and in particular, to fabrication of stringers for aircraft. 
     BACKGROUND 
     Stringers for an aircraft (e.g., the wings of an aircraft) may be fabricated from composite materials. These stringers may be laid-up as preforms, processed into stringers, and then co-cured to a wing skin in order to form a completed wing or portion thereof. When stringers are fabricated in this manner, the stringers are laid-up and cured such that they terminate within the wing skin. The stringers can then be cut to desired dimensions at desired ramp rates after curing has completed. The transition from stringer to skin is commonly referred to as a “stringer run-out” and is utilized for transferring load in composite wing structures. 
     Fabricating a stringer made of composite materials in this manner may result in edge conditions which are out of tolerance or, and out of tolerance conditions are undesirable as they may necessitate rework. At the same time, it may be difficult to lay up and cure a preform at desired ramp rates, owing to the complex geometries that may be required. 
     Therefore, it would be desirable to have a method and apparatus that take into account at least some of the issues discussed above, as well as other possible issues. 
     SUMMARY 
     Embodiments described herein provide rigid end caps for stringers that exhibit desired ramp rates and stringer run-out quality for terminating a stringer. These end caps are affixed to preforms for the stringers to form integral portions of a composite part. The end caps eliminate the need to shape the preforms to specified ramp rates, while also eliminating the need to cut the stringers after curing. The end caps form flyaway tooling that enforces desired shapes onto stringers during curing, while also bearing and transferring loads within the wing after the wing has been fabricated. One embodiment is a method of forming a stringer. The method includes laying up a stringer preform comprising fiber-reinforced 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, resulting in a composite part that includes the end cap. 
     One embodiment is a method for fabricating a composite part. The method includes forming a skin panel preform comprising fiber reinforced material, disposing rigid end caps at the skin panel preform at end locations of stringer preforms that will be placed at the skin panel, locating the stringer preforms at the skin panel preform via the rigid end caps, and anchoring the stringer preforms to the skin panel preform. 
     A further embodiment is an apparatus for receiving a stringer preform. The apparatus includes a skin panel preform comprising fiber reinforced material, and rigid end caps that are disposed atop the skin panel preform and are separated by a length of a stringer preform that will be placed onto the skin panel preform. 
     A further embodiment is a non-transitory computer readable medium embodying programmed instructions which, when executed by a processor, are operable for performing a method for forming a stringer. The method includes forming a skin panel preform comprising fiber reinforced material, disposing rigid end caps at the skin panel preform at end locations of stringer preforms that will be placed at the skin panel, locating the stringer preforms at the skin panel preform via the rigid end caps, and anchoring the stringer preforms to the skin panel preform. 
     A further embodiment is a system that forms a portion of an aircraft. The system includes a section of airframe comprising a skin panel comprising fiber reinforced material, stringers that are affixed to the skin panel and that comprise fiber reinforced material, and rigid end caps that are integral with the skin panel and the stringers, the rigid end caps comprising flyaway tooling that supported the stringers during hardening. 
     Other illustrative embodiments (e.g., methods and computer-readable media relating to the foregoing embodiments) may be described below. The features, functions, and advantages that have been discussed can be achieved independently in various embodiments or may be combined in yet other embodiments further details of which can be seen with reference to the following description and drawings. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
       Some embodiments of the present disclosure are now 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 on all drawings. 
         FIG. 1  is a block diagram of one half of a wing that includes co-cured stringers and flyaway tooling in an illustrative embodiment. 
         FIG. 2A  is a flowchart illustrating a method for integrating flyaway tooling into a wing in an illustrative embodiment. 
         FIG. 2B  is a flowchart illustrating a method for integrating flyaway tooling into a wing in an illustrative embodiment. 
         FIGS. 3 and 4A-4C  illustrate an end cap for a first stringer shape in an illustrative embodiment. 
         FIGS. 5-6  illustrate an end cap for a second stringer shape in an illustrative embodiment. 
         FIGS. 7A-7E and 8  illustrate a co-bonded end cap and stringer in an illustrative embodiment. 
         FIG. 9  is a perspective view of a wing that includes flyaway tooling in an illustrative embodiment. 
         FIG. 10  illustrates an inboard end cap for a stringer in an illustrative embodiment. 
         FIG. 11  is a block diagram illustrating inboard end caps for stringers integrated into an airframe in an illustrative embodiment. 
         FIG. 12  is a further flowchart illustrating a method for integrating flyaway tooling into a wing in an illustrative embodiment. 
         FIG. 13  is a flow diagram of aircraft production and service methodology in an illustrative embodiment. 
         FIG. 14  is a block diagram of an aircraft in an illustrative embodiment. 
     
    
    
     DESCRIPTION 
     The figures and the following description provide specific illustrative embodiments of the disclosure. It will thus be appreciated that those skilled in the art will be able to devise various arrangements that, although not explicitly described or shown herein, embody the principles of the disclosure and are included within the scope of the disclosure. Furthermore, any examples described herein are intended to aid in understanding the principles of the disclosure, and are to be construed as being without limitation to such specifically recited examples and conditions. As a result, the disclosure is not limited to the specific embodiments or examples described below, but by the claims and their equivalents. 
     Composite parts, such as Carbon Fiber Reinforced Polymer (CFRP) parts, are initially laid-up in multiple layers that together are referred to as a preform. Individual fibers within each layer of the preform are aligned parallel with each other, but different layers exhibit 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 in order to harden the preform into a composite part (e.g., for use in an aircraft). Carbon fiber that has been impregnated with an uncured thermoset resin or a thermoplastic resin is referred to as “prepreg.” Other types of carbon fiber include “dry fiber” which has not been impregnated with thermoset resin but may include a tackifier or binder. Dry fiber is infused with resin prior to curing. For thermoset resins, the hardening is a one-way process referred to as curing, while for thermoplastic resins, the resin reaches a viscous form if it is re-heated. 
       FIG. 1  is a block diagram of a lower half of a wing  100  that includes co-cured stringers and flyaway tooling in an illustrative embodiment. The lower half of the wing  100  illustrated in  FIG. 1  may comprise a co-cured lower half of a wing, and is capable of being affixed or bonded to an upper half of a wing in order to form a complete wing. In further embodiments, an upper half of a wing is assembled in a similar manner to that of the lower half of the wing  100 . The lower half of a wing  100  depicted in  FIG. 1  includes fiber-reinforced composite materials as well as rigid end caps for stringers. This arrangement eliminates the need for trimming or cutting the stringers at the wing  100  after hardening. 
     In this embodiment,  FIG. 1  depicts a portion of a wing that includes a skin panel  110 , which comprises fiber-reinforced material in the form of multiple layers  112  of resin  114  and fibers  116 . A stringer  120  is disposed atop the skin panel  110 , and comprises fiber-reinforced material in the form of layers  112  of fibers  116  and resin  114 . The stringer  120  is co-cured to the skin panel  110 . 
     End caps  130  abut the inboard end  122  and outboard end  124  of the stringer  120 . The end caps  130  are rigid prior to curing, and may comprise a metal that provides high strength with low weight, such as titanium or aluminum (e.g., isolated from carbon fiber by isolation plies of fiberglass or other material). End caps  130  may further comprise hardened composite materials (e.g., thermoset or thermoplastic), as well as  3 D-printed metals. Thus, in one embodiment the end caps  130  are fabricated via additive manufacturing techniques, such as  3 D printing. In further embodiments, subtractive manufacturing techniques are utilized. Each end cap  130  includes a flange  132  with a ramp  134  for receiving the stringer  120 . The end caps  130  are bolted and/or bonded to the stringer  120  and/or skin panel  110 . 
     The ramp  134  provides a pathway for transferring load from the stringer  120  to the end cap  130 . As used herein, a “ramp” refers to any physical structure that transitions load along its length, including step laps, scarfing, interleaving, linear ramps, and other features. The stringer  120  itself includes a flange and web, and these structures are complementary to the ramp  134  of the end cap  130 . Each end cap  130  also includes a ramp  136  that proceeds down to the skin panel  110 . The ramp  136  provides a pathway for transferring load from the end cap  130  to the skin panel  110 . 
     The end caps  130  are co-bonded to the stringers  120  and the skin panel  110 , and the ramps  134  of the end caps  130  are overlapped with fiber-reinforced material from the stringers  120  in one embodiment. In further embodiments, the ramps  134  and/or  136  integrate with stringers  120  via laps, step laps, or scarf interleaving of the ramps with the plies of the stringer and/or skin panel  110 . In some of these embodiments, the transitions involve laying up a stinger preform upon an end cap  130  to accommodate differences in shape. The terms of “skin panel  110 ” and “stringer  120 ” are utilized herein to refer to both uncured preforms as well as hardened composite parts. That is, a skin panel  110  may refer to an unhardened preform for a skin panel awaiting curing, or may refer to a hardened skin panel. In a similar fashion, a stringer  120  may refer to an unhardened preform for a stringer, or to a hardened stringer. 
       FIG. 1  further depicts components which may be electronically managed by controller  140  to fabricate the structure discussed above. In this embodiment, the components include a layup machine  142 , such as an Automated Fiber Placement (AFP) machine or tape dispensing head that lays up tows of unidirectional fiber-reinforced material to form the layers  112 . The components further include a Pick and Place (PNP) machine  144  (e.g., an end effector, suction device, gripper, etc.), which picks up and places stringer preforms and/or end caps  130  onto skin panel  110 . In further embodiments, the stringer preforms and end caps  130  are manually picked and placed into position by one or more technicians. Controller  140  directs the operations of these components based on instructions stored in one or more Numerical Control (NC) programs in memory, and may be implemented, for example, as custom circuitry, as a hardware processor executing programmed instructions, or some combination thereof 
     Illustrative details of the operation of the components of  FIG. 1  will be discussed with regard to  FIG. 2A . Assume, for this embodiment, that a skin panel preform (e.g., skin panel  110 ) for a wing has been laid-up by layup machine  142 . 
       FIG. 2A  is a flowchart illustrating a method  200  for integrating flyaway tooling into a wing in an illustrative embodiment. The steps of method  200  are described with reference to the lower half of a wing  100  shown in  FIG. 1 , but those skilled in the art will appreciate that method  200  may be performed for other portions of wings (e.g., upper halves of wings) in other environments. The steps of the flowcharts 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. 
     In step  202 , controller  140  directs layup machine  142  to lay up the stringer preform. In one embodiment, this comprises applying multiple layers of unidirectional fiber-reinforced material to a layup mandrel (not shown) or other piece of tooling that defines a shape for the stringer preform. 
     In one embodiment, the PNP machine  144  places an end cap  130  at each end of the stringer preform. The end caps  130  help to enforce a desired shape at the stringer preform before and during curing. Furthermore, the end caps  130  provide ramps  136  (e.g., for runouts) 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 layup and curing for the wing, which reduces expenses related to labor and materials. 
     In step  204 , controller  140  directs the PNP machine  144  to place the stringer preform onto a skin panel preform (e.g., skin panel  110 ). During this operation, the stringer preform has not yet been cured (i.e., is still in the “green state”) and therefore remains flaccid. The PNP machine  144  may therefore enforce or retain a desired curvature at the stringer preform via the application of suction (e.g., via a vacuum connection) or use of supporting structure while the stringer preform is being transported. In one embodiment, the PNP machine  144  picks up and places multiple stringer preforms at the skin panel preform. This may also include placing a stringer preform onto a layup mandrel and then laying up the skin against the stringer preform. For a skin panel that defines an upper portion of a wing, the stringers may comprise preforms for hat stringers. For a skin panel that defines a lower portion of a wing, the stringers may comprise preforms for “T” stringers. Further types of stringers include Z stringers, and stringers of any suitable cross-section. In one embodiment, the placement operation involves overlapping a ramp  134  of the end cap  130  with one or more layers of the stringer preform. This may comprise overlapping the ramp  134  with a ramp at the stringer preform. In this manner, after the stringer preform hardens, the ramp  134  transfers loads between the hardened stringer and the end cap  130 . Meanwhile the ramp  136  transfers loads between the end cap  130  and the skin panel  110 . 
     Step  206  comprises bonding an end cap  130  to the stringer preform and the skin panel preform. In one embodiment, this comprises applying an adhesive (e.g., an epoxy, glue, or other self-hardening chemical) to the end caps  130  prior to placing the end caps or the stringer preforms onto the skin panel preform, and waiting for the adhesive to harden after placing the end caps and stringer preform into position. In one embodiment, an end cap  130  is bonded to an outboard end of the stringer preform, and another end cap  130  is bonded to an inboard end of the stringer preform. 
     After the end caps  130  have been bonded into place, the position of end caps  130  with regard to the stringer preform and the skin panel preform is held in place by the hardened adhesive. This ensures that vacuum bagging setup and consolidation will not shift the position of the stringer preforms, skin panel preform, and end caps with respect to each other. Thus, PNP machine  144 , another machine, or a technician, may proceed to vacuum bag the end cap, stringer preform, and skin panel preform (i.e., prior to co-curing these elements together). The end caps both anchor the stringer preforms to the skin panel preform, and locate the stringers at the skin panel preform. In one embodiment, the vacuum bag is utilized to consolidate these components via the application of pressure, prior to curing. 
     Step  208  comprises co-curing the stringer preform to the skin panel preform while the end caps  130  are bonded to the stringer preform and the skin panel preform, resulting in a composite part that includes the end caps  130  as integral components. In one embodiment, co-curing comprises placing the vacuum-bagged components into an autoclave, applying heat via the autoclave until resin  114  reaches a curing temperature, and applying pressure via the vacuum bag and/or via the autoclave in order to consolidate and cure the end caps  130 , stringer preforms, and skin panel preform. These components are co-cured into an integral composite part that includes integral flyaway tooling. That is, the end caps  130  operate as tooling to provide support for the preforms during curing and vacuum bagging, and also provide mechanical strength when the resulting portion of wing is assembled into a portion of an airframe of an aircraft. The tooling becomes physically integral with stringers after hardening. Therefore, the tooling is flyaway tooling because it is integrated into an aircraft and “flies away” as part of the aircraft after fabrication has completed. 
     Method  200  provides a technical benefit over prior techniques, because it enables a runout to be rapidly integrated into a composite stringer, provides support during vacuum bagging and curing. This is because the end caps, being rigid prior to curing, resist compaction forces applied by a vacuum bag that could crush or bend elongated portions such as a web of a stringer preform. Hence, the end caps help to constrain stringer preforms to desired shapes during curing. Method  200  additionally eliminates the need to cut or remove material from a composite stringer after the composite stringer has been co-cured to a skin panel. 
       FIG. 2B  illustrates a method  250  for integrating flyaway tooling into a wing in an illustrative embodiment. Method  250  comprises forming (e.g., laying up) a skin panel preform comprising fiber reinforced material (e.g., CFRP) in step  252 , disposing rigid end caps at the skin panel preform at end locations of stringer preforms (e.g., laid-up from CFRP) that will be placed at the skin panel in step  254 , locating the stringer preforms at the skin panel preform via the rigid end caps in step  256 , and anchoring the stringer preforms to the skin panel preform in step  258 . 
     In a further embodiment, method  250  includes consolidating the stringer preforms and the skin panel preform via a vacuum bag that covers the stringer preforms, skin panel preform, and end caps. Method  250  may further comprise hardening the stringer preforms and the skin panel preform to form a section of wing that includes the end caps. The end caps are bonded to the stringer preforms and the skin panel preform, and in one embodiment the method  250  further comprises co-curing the stringer preform and the skin panel preform while the rigid end caps are bonded to the stringer preform and the skin panel preform, resulting in a composite part that includes the end cap. 
       FIGS. 3 and 4A-4C  illustrate an end cap  300  for a first stringer shape in an illustrative embodiment.  FIG. 4A  corresponds with view arrows  4 A of  FIG. 3 . In this embodiment, the end cap  300  is designed as an end cap for a stringer having a T-shaped cross-section. That is, the end cap  300  has a T-shaped cross-section that aligns with a T-shaped cross section of a stringer. 
     The end cap  300  includes a web  310  which narrows via ramp  312 , which extends from the web  310  and tapers the web  310 . In this embodiment, the web  310  forms a vertical plane. However, in further embodiments the web forms a curved shape (e.g., as shown in  FIG. 5 ). The web  310  protrudes from a lower flange  320 . Lower flange  320  narrows via ramp  322 , which extends from the lower flange  320  and tapers the lower flange  320 . These ramps  312  and  322  are overlapped by corresponding ramps in the stringer preform, such that an overall thickness of the combination of flange and stringer preform remains constant along the flange. In further embodiments, the ramps exhibit a stairstep pattern that accommodates a step lap or other type of interface/transition between the stringer and the end cap. In still further embodiments, interleaving of composite plies with metallic structure is performed to narrow the structure instead of narrowing in a linear ramped fashion. Ramp  330  terminates the stringer in accordance with desired structural constraints, and is shaped and tapered to help transition the stringer load through the end cap to the panel. Ramp  340  transfers forces through the end cap  300  into a skin panel (not shown). Meanwhile, the ramp  312  proceeding into the web  310 , and the ramp  322  proceeding into the lower flange  320  transfer load from the stringer. The rate of transition of load through the end cap is established by the geometry and of the ramps and the pattern (e.g., linear, step, etc.) of the ramps. 
       FIG. 4B  depicts a section-cut side view of the end cap  300  of  FIG. 3 , after a stringer  400  has been bonded thereto.  FIG. 4B  corresponds with view arrows  4 B of  FIG. 4A . In  FIG. 4B , it can be seen that stringer  400  includes ramp  410  and ramp  420 , which mate with ramps  312  and  322 , respectively.  FIG. 4C  is a section cut end view that corresponds with view arrows  4 C of  FIG. 4A . In  FIG. 4C  ramps  312  and  322  of end cap  300  are visible, as are ramps  410  and  420  of stringer  400 . In further embodiments the end cap is bonded to an outboard end of the stringer preform and fastened to the wing panel, and another end cap is bonded to an inboard end of the stringer preform and fastened to the wing panel. 
       FIGS. 5-6  illustrate an end cap  500  for a second stringer shape in an illustrative embodiment.  FIG. 6  corresponds with view arrows  6  of  FIG. 5 . In this embodiment, the end cap  500  is designed as an end cap for a stringer having a “hat” shaped cross-section. That is, the end cap  500  has a hat-shaped cross-section that aligns with a hat-shaped cross section of a stringer. 
     The end cap  500  includes an upper arch  510  which narrows via ramp  512 , and also includes a lower flange  520 , which narrows via ramp  522 . These flanges are overlapped by corresponding ramps in the stringer preform, such that an overall thickness of the combination of flange and stringer preform remains constant along the flange. In further embodiments, the ramps exhibit a stairstep pattern as they narrow, instead of narrowing in a linear fashion. Ramp  530  terminates the stringer in accordance with desired structural constraints, and lip  540  distributes forces borne through the end cap  500  into a skin panel (not shown). In this embodiment, the upper arch  510  forms a void  550  which corresponds with a void in the hat stringer. 
       FIGS. 7A-7E and 8  illustrate an end cap  500  that has been co-bonded to a stringer  700  in an illustrative embodiment.  FIG. 8  is a side view corresponding with view arrows  8  of  FIG. 7A . In this embodiment, ramp  512  and ramp  522  are covered by fiber-reinforced material from the stringer  700 , which forms a corresponding instance of a ramp  710  to maintain a desired combined thickness of the stringer and end cap. 
       FIG. 7B  is a section-cut side view that corresponds with view arrows  7 B of  FIG. 7A , and depicts the ramp  710  intersecting with ramp  522  of  FIG. 7A  at scarf joint  720 . In FIG.  7 B, the end cap has been bonded to a ramp of a stringer  700 , which structurally unites the stringer  700  with the end cap  500 , forming a single component made from the stringer as well as the end cap. Specifically,  FIG. 7B  depicts a scarf type of join between the ramp of the end cap and the composite plies. However, a step type of join can also be implemented, as depicted by step joint  730  of  FIG. 7C . In further embodiments, the end cap is interleaved with composite plies, as shown by interleaved joint  740  of  FIG. 7D . Still further implementations of interleaving may also or alternatively be utilized. 
       FIG. 7E  is a section-cut end view that corresponds with view arrows  7 E of  FIG. 7A . This view further depicts and intersection between ramps for an end cap and ramps for a stringer. As shown in  FIG. 7E , a portion of the stringer  700  overlaps with the ramp  512  and ramp  522  of the end cap  500 . 
       FIG. 8  depicts a side view that corresponds with view arrows  8  of  FIG. 7A . In  FIG. 8 , it can be observed that a majority of the join (i.e., proceeding into the page between the ramp  710  and the ramp  522 ) is hidden from view. 
       FIG. 9  is a perspective view of a wing  900  that includes flyaway tooling in an illustrative embodiment. The tooling comprises end caps that facilitated hardening of the stringers by supporting the stringers during curing, and that became physically integral with the stringers after hardening. Therefore, the tooling is flyaway tooling because it is integrated into an aircraft and “flies away” as part of the aircraft after fabrication has completed. Thus, each end cap plays a dual role as tooling that supports preforms prior to and during curing, and also as a component that transfers load from the stringer into the panel when integrated into an aircraft. 
       FIG. 9  illustrates that the wing  900  includes leading edge  902 , trailing edge  904 , and pylon  906 . Wing  900  further includes stringers  920 , as well as inboard end caps  910  and outboard end caps  912  disposed at the inboard and outboard ends of the stringers  920 . For example, the inboard end caps  910  may terminate a stringer  920  at a side-of-body intersection to form an inboard end of a stringer  920 , and the outboard end caps  912  may terminate an outboard end of a stringer  920 . In one embodiment, outboard end caps are smaller than inboard end caps, because they mate with tapered portions of the stringers  920 . In such an embodiment inboard end caps do not taper into a skin panel, but rather are bonded or fastened to a wing box. Such end caps still taper into the stringers themselves, however. Because  FIG. 9  provides a simplified view, it will be understood that there can be many more stringers than illustrated, and stringers can be much narrower relative to wing size. For example, a stringer can taper from inboard end to outboard end. Also, some of the stringers may continue to the tip of the wing while others terminate earlier. 
       FIG. 10  illustrates an inboard end cap  1000  for a stringer in an illustrative embodiment. In this embodiment, the inboard end cap  1000  is designed as an end cap for a stringer having a T-shaped cross-section. That is, the end cap  1000  has a T-shaped cross-section that aligns with a T-shaped cross section of an inboard end of a stringer. 
     The inboard end cap  1000  includes a web  1010  which narrows via ramp  1012 , which extends from the web  1010  and tapers the web  1010 . In this embodiment, the web  1010  forms a vertical plane. The web  1010  protrudes from a lower flange  1020 . Lower flange  1020  narrows via ramp  1022 , which extends from the lower flange  1020  and tapers the lower flange  1020 . These ramps  1012  and  1022  are overlapped by/spliced with corresponding ramps in the stringer preform, such that an overall thickness of the combination of flange and stringer preform remains constant along the flange at end  1080 . In further embodiments, the ramps exhibit a stairstep pattern that accommodates a step lap or other type of interface/transition between the stringer and the end cap. In still further embodiments, interleaving of composite plies with metallic structure is performed to narrow the structure instead of narrowing in a linear ramped fashion. Web extension  1030  terminates the stringer by abutting against a side of body intersection at end  1080  to transfer load from a wing box. Meanwhile, the ramp  1012  proceeding into the web  1010 , and the ramp  1022  proceeding into the lower flange  1020  transfer load from the stringer to a center of the wing box. The rate of transition of load through the end cap is established by the geometry and of the ramps and the pattern (e.g., linear, step, etc.) of the ramps and web extension. 
       FIG. 10  further depicts mounting features  1050  in the web and mounting feature  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 affixing the inboard end cap  1000  to a center wing box. 
       FIG. 11  is a block diagram illustrating inboard end caps for stringers integrated into an airframe in an illustrative embodiment. In  FIG. 11 , a side of body  1110  of an airframe is abutted by inboard end caps  1112 . Stringers  1120  are attached to the inboard end caps  1112 , and are also attached to outboard end caps  1114 . The stringers  1120  and end caps are also affixed to a skin panel  1104 . 
       FIG. 12  is a further flowchart illustrating a method for integrating flyaway tooling into a wing in an illustrative embodiment. Step  1202  comprises forming a skin panel preform via layup or other techniques. Step  1204  comprises bonding a caul tooling end (e.g., an end cap) onto the skin panel preform. Step  1206  comprises anchoring a stringer preform at the skin panel preform and the caul tooling end. Step  1208  comprises hardening the skin panel preform and the stringer preform with the caul tooling end. That is, while the skin panel preform and the stringer preform are supported by the caul tooling end, they are hardened into a composite part. In a further embodiment, vacuum bagging is performed prior to hardening, and this process includes vacuum bagging the stringer preform, the caul tooling end, and the skin panel preform. In one embodiment, hardening comprises curing the stringer preform, the caul tooling end, and the skin panel preform into a composite part while applying pressure. In a further embodiment, the caul tooling end forms flyaway tooling that structurally supports ends of a stringer. In a further embodiment, the caul tooling end comprises a material selected from the group consisting of metal and Carbon Fiber Reinforced Polymer (CFRP). 
     EXAMPLES 
     In the following examples, additional processes, systems, and methods are described in the context of end caps used as flyaway tooling for stringers. 
     Referring more particularly to the drawings, embodiments of the disclosure may be described in the context of aircraft manufacturing and service in method  1200  as shown in  FIG. 12  and an aircraft  1202  as shown in  FIG. 13 . During pre-production, method  1200  may include specification and design  1204  of the aircraft  1202  and material procurement  1206 . During production, component and subassembly manufacturing  1208  and system integration  1210  of the aircraft  1202  takes place. Thereafter, the aircraft  1202  may go through certification and delivery  1212  in order to be placed in service  1214 . While in service by a customer, the aircraft  1202  is scheduled for routine work in maintenance and service  1216  (which may also include modification, reconfiguration, refurbishment, and so on). Apparatus and methods embodied herein may be employed during any one or more suitable stages of the 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 service  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 , environmental  1230 ). 
     Each of the processes of method  1200  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 vendors, subcontractors, and suppliers; and an operator may be an airline, leasing company, military entity, service organization, and so on. 
     As shown in  FIG. 12 , the aircraft  1202  produced by method  1200  may include an airframe  1218  with a plurality of systems  1220  and an interior  1222 . 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 invention may be applied to other industries, such as the automotive industry. 
     As already mentioned above, apparatus and methods embodied herein may be employed during any one or more of the stages of the production and service described in method  1200 . For example, components or subassemblies corresponding to component and subassembly manufacturing  1208  may be fabricated or manufactured in a manner similar to components or subassemblies produced while the aircraft  1202  is in service. Also, one or more apparatus embodiments, method embodiments, or a combination thereof may be utilized during the subassembly manufacturing  1208  and system integration  1212 , for example, by substantially expediting assembly of or reducing the cost of an aircraft  1202 . Similarly, one or more of apparatus embodiments, method embodiments, or a combination thereof may be utilized while the aircraft  1202  is in service, for example and without limitation during the maintenance and service  1216 . Thus, the invention may be used in any 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 service  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  1230 .) 
     In one embodiment, a part comprises a portion of airframe  1218 , and is manufactured during component and subassembly manufacturing  1208 . The part may then be assembled into an aircraft in system integration  1210 , and then be utilized in service  1214  until wear renders the part unusable. Then, in maintenance and service  1216 , the part may be discarded and replaced with a newly manufactured part. Inventive components and methods may be utilized throughout component and subassembly manufacturing  1208  in order to manufacture new parts. 
     Any of the various control elements (e.g., electrical or electronic components) shown in the figures or described herein may be implemented as hardware, a processor implementing software, a processor implementing firmware, or some combination of these. For example, an element may be implemented as dedicated hardware. Dedicated hardware elements may be referred to as “processors”, “controllers”, or some similar terminology. When provided by a processor, the functions may be provided by a single dedicated processor, by a single shared processor, or by a plurality of individual processors, some of which may be shared. Moreover, explicit use of the term “processor” or “controller” should not be construed to refer exclusively to hardware capable of executing software, and may implicitly include, without limitation, digital signal processor (DSP) hardware, a network processor, application specific integrated circuit (ASIC) or other circuitry, field programmable gate array (FPGA), read only memory (ROM) for storing software, random access memory (RAM), non-volatile storage, logic, or some other physical hardware component or module. 
     Also, a control element may be implemented as instructions executable by a processor or a computer to perform the functions of the element. Some examples of instructions are software, program code, and firmware. The instructions are operational when executed by the processor to direct the processor to perform the functions of the element. The instructions may be stored on storage devices that are readable by the processor. Some examples of the storage devices are digital or solid-state memories, magnetic storage media such as a magnetic disks and magnetic tapes, hard drives, or optically readable digital data storage media. 
     Although specific embodiments are described herein, the scope of the disclosure is not limited to those specific embodiments. The scope of the disclosure is defined by the following claims and any equivalents thereof.