Patent Application
20230078701
This disclosure relates generally to composite structures, and more particularly to composite structures for aircraft and methods for forming composite structures for aircraft.
Composite materials are frequently used in the aerospace industry for a diverse array of structural and dynamic aerostructural applications because of the strength-to-weight advantage provided by composite materials. Various types of molding techniques may be used to construct composite structures or components for an aircraft. For example, resin pressure molding (RPM) techniques and Same Qualified Resin Transfer Molding (SQRTM) techniques may be used to form composite structures for aerospace applications. However, composite structures formed by certain molding techniques may require component thicknesses to be greater than desired in order to prevent or reduce the likelihood of skin buckling, thereby increasing component weight. Accordingly, what is needed are improved composite structures and methods of forming composite structures which address the above-noted concern.
It should be understood that any or all of the features or embodiments described herein can be used or combined in any combination with each and every other feature or embodiment described herein unless expressly noted otherwise.
According to an aspect of the present disclosure, a composite structure includes a first composite skin and a second composite skin defining a longitudinal cavity therebetween. The first composite skin and the second composite skin further define at least one edge where the first composite skin contacts the second composite skin. The composite structure further includes at least one core disposed within the longitudinal cavity. The core includes a first surface and a second surface which define a core edge where the first surface contacts the second surface. The core is positioned with the core edge adjacent the at least one edge with the first surface contacting the first composite skin and the second surface contacting the second composite skin.
In any of the aspects or embodiments described above and herein, the at least one core may include an interior portion including a honeycomb structure including a plurality of cavities defined by a plurality of side walls extending between the first surface and the second surface and an exterior portion surrounding the honeycomb structure portion and defining the first surface and the second surface.
In any of the aspects or embodiments described above and herein, the at least one core may include a foam material.
In any of the aspects or embodiments described above and herein, the composite structure may include a plurality of spars located in the longitudinal cavity and laterally spaced from one another. The plurality of spars may extend between and connect the first composite skin and the second composite skin.
In any of the aspects or embodiments described above and herein, the at least one core may be disposed in a sub-cavity defined between the at least one edge and an adjacent spar of the plurality of spars.
In any of the aspects or embodiments described above and herein, the core may include a third surface extending between the first surface and the second surface. The third surface may contact the adjacent spar.
In any of the aspects or embodiments described above and herein, the first composite skin and the second composite skin may extend between a first longitudinal end and a second longitudinal end opposite the first longitudinal end.
In any of the aspects or embodiments described above and herein, the at least one core may extend a portion of a distance from the first longitudinal end to the second longitudinal end.
In any of the aspects or embodiments described above and herein, the at least one core may extend substantially an entire distance from the first longitudinal end to the second longitudinal end.
In any of the aspects or embodiments described above and herein, the at least one core may be tapered such that one or both of a width and a height of the at least one core changes in a direction from a first longitudinal side of the at least one core to a second longitudinal side of the at least one core opposite the first longitudinal side.
In any of the aspects or embodiments described above and herein, the first composite skin and the second composite skin may form a unitary composite skin.
According to another aspect of the present disclosure, a method for forming a composite structure includes positioning a first composite skin and a second composite skin so that the first composite skin and the second composite skin define a longitudinal cavity therebetween and at least one edge where the first composite skin contacts the second composite skin, curing the first composite skin and the second composite skin, and inserting a core into the longitudinal cavity so that the core is positioned with a first surface of the core contacting the first composite skin, a second surface of the core contacting the second composite skin, and a core edge of the core adjacent the at least one edge. The core edge is defined where the first surface contacts the second surface.
In any of the aspects or embodiments described above and herein, the step of positioning the first composite skin and the second composite skin may include positioning a plurality of spars so that the plurality of spars are located in the longitudinal cavity and laterally spaced from one another with the plurality of spars extending between and connecting the first composite skin and the second composite skin. The step of curing the first composite skin and the second composite skin may include curing the plurality of spars.
In any of the aspects or embodiments described above and herein, the core may be disposed in a sub-cavity defined between the at least one edge and an adjacent spar of the plurality of spars.
In any of the aspects or embodiments described above and herein, the core may include a third surface extending between the first surface and the second surface and the step of inserting the core into the longitudinal cavity may include positioning the third surface in contact with the adjacent spar.
In any of the aspects or embodiments described above and herein, the method may further include inserting at least one mandrel into the longitudinal cavity, prior to the step of curing the first composite skin and the second composite skin.
In any of the aspects or embodiments described above and herein, the step of inserting the core into the longitudinal cavity may be performed subsequent to curing the first composite skin and the second composite skin.
In any of the aspects or embodiments described above and herein, the method may further include applying an adhesive to at least the first surface and the second surface of the core prior to the step of inserting the core into the longitudinal cavity.
According to another aspect of the present disclosure, a composite structure includes a first composite skin and a second composite skin mounted to the first composite skin. The first composite skin and the second composite skin define a longitudinal cavity therebetween. The first composite skin and the second composite skin further define a first longitudinal edge where the first composite skin contacts the second composite skin at a first lateral side and a second longitudinal edge where the first composite skin contacts the second composite skin at a second lateral side opposite the first lateral side. The composite structure further includes a first core and a second core disposed within the longitudinal cavity. The first core is positioned adjacent the first longitudinal edge and contacts the first composite skin and the second composite skin and the second core is positioned adjacent the second longitudinal edge and contacts the first composite skin and the second composite skin.
In any of the aspects or embodiments described above and herein, each of the first core and the second core may include a first surface and a second surface which define a core edge where the first surface contacts the second surface and the first surface is in contact with the first composite skin and the second surface is in contact with the second composite skin.
The present disclosure, and all its aspects, embodiments and advantages associated therewith will become more readily apparent in view of the detailed description provided below, including the accompanying drawings.
In accordance with various aspects of the present disclosure, apparatuses, systems, and methods are described in connection with a component of, for example, an aircraft. In some embodiments, the component may be a composite structure such as, but not limited to, an aircraft control structure, an airfoil, or a wing of an aircraft. In some embodiments, a composite structure of the present disclosure may for all or a portion of a stabilizer or a stabilator of an aircraft. However, it should be understood that the composite structures of the present disclosure are not limited to utilization in an aircraft or for aerospace applications and may alternatively be used for other applications.
Referring to
The configuration of the composite structure 20 is discussed above to assist in the description of the present disclosure. It should be understood, however, that composite structures may have a variety of different shapes, forms, and configurations and the present disclosure is not limited to the particular exemplary configuration of the composite structure 20 described above. As used herein, the terms “longitudinal,” “lateral,” and “vertical” may be used to refer to the respective x-axis, y-axis, and z-axis as shown, for example, in
In some embodiments, the composite structure 20 may include a plurality of spars 40 located in the cavity 26 and laterally spaced from one another within the cavity 26. Each spar of the plurality of spars 40 extends between and connects the first composite skin 22 and the second composite skin 24 in order to provide structural support for the composite structure 20. As shown in
The composite structure 20 may include at least one opening 44 between the cavity 26 and an exterior of the composite structure 20. For example, the first composite skin 22 and the second composite skin 24 may define the opening 44 therebetween at one or both of the first longitudinal end 36, as shown in
Referring to
In some embodiments, the at least one core 46 may include an interior portion 62 and an exterior portion 64. The interior portion 62 may include a plurality of cells 66 defined by a corresponding plurality of walls 68 of the interior portion 62 which extend, for example, between the first surface 56 and the second surface 58. In various embodiments, each cell of the plurality of cells 66 may be configured to form a “honeycomb” structure defined by, for example, six adjacent walls of the plurality of walls 68 (see, e.g.,
Referring to
Referring to
Referring to
Step 1002 includes positioning the at least one composite skin and/or the plurality of spars 40 relative to one another in preparation for forming the composite structure 20, as described above. For example, step 1002 may include positioning the first composite skin 22, the second composite skin 24, and the plurality of spars 40 so that the second composite skin 24 is spaced from the first composite skin 22 and the first composite skin 22 and the second composite skin 24 define the cavity 26 therebetween, and so that the plurality of spars 40 may be located in the cavity 26 and laterally spaced from one another with the plurality of spars 40 extending between and connecting the first composite skin 22 and the second composite skin 24.
In some embodiments, the method 1000 may optionally include inserting at least one mandrel 80 into the cavity 26 and/or one or more of the sub-cavities 42 defined by the plurality of spars 40, as provided in step 1004 and shown in
Step 1006 includes curing the composite skins 22, 24 and the plurality of spars 40. In some embodiments, the composite skins 22, 24 and the plurality of spars 40 may be co-cured (e.g., cured simultaneously) to form the composite structure 20. Curing the composite skins 22, 24 and the plurality of spars 40 may include heating the assembled composite skins 22, 24 and the plurality of spars 40 to an elevated temperature and holding the composite skins 22, 24 and the plurality of spars 40 at the elevated temperature for a sufficient time to cure the composite skins 22, 24 and the plurality of spars 40. Various temperatures, pressure, and curing times may be used, depending on the materials selected for the composite skins 22, 24 and the plurality of spars 40. The composite skins 22, 24 and the plurality of spars 40 may be cured, for example, in an oven or autoclave. The present disclosure is not limited to any particular curing temperatures, pressures, curing times, or equipment. In the cured state, the composite skins 22, 24 and the plurality of spars 40 form the composite structure 20.
In some embodiments, for example, where at least one mandrel 80 has been used to support the composite structure 20, the method 1000 may include removing the at least one mandrel 80 from the sub-cavities 42 of the composite structure 20 once the composite structure 20 has sufficiently cooled and solidified, as provided in step 1008.
In some embodiments, the steps 1002, 1004, 1006, and 1008 of method 1000 may be performed during application a composite molding process. Various types of molding techniques may be used to construct composite components of an aircraft. For example, a resin pressure molding (RPM) technique or a Same Qualified Resin Transfer Molding (SQRTM) technique may combine pre-preg processing and liquid molding to produce composite components targeted to aerospace applications. As part of these techniques, pre-preg plies may be arranged within a mold, the mold may be closed, and then a resin may be injected into the mold. The resin maintains hydrostatic pressure within the mold. The present disclosure, however, is not limited to any particular composite formation technique or process for forming the composite structure 20.
In some embodiments, the method 1000 may optionally include applying an adhesive to the at least one core 46, as provided in step 1010, prior to insertion of the at least one core 46 into the composite structure 20. As shown in
Step 1012 includes inserting the at least one core 46 into the cavity 26 of the composite structure 20 and positioning the at least one core 46 within the composite structure 20 as described above. For example, step 1012 may include inserting the at least one core 46 into the cavity 26 so that the at least one core 46 is mounted to or otherwise in contact with one or more of the first composite skin 22, the second composite skin 24, and an adjacent spar of the plurality of spars 40. Insertion of the at least one core 46 into the composite structure 20 may be performed subsequent to curing the composite skins 22, 24 and the plurality of spars 40. In some embodiments, such as with embodiments of the composite structure 20 which have one or more tapered sub-cavities 42, as described above, the at least one core 46 may be inserted into a respective sub-cavity 42 in the taper direction (e.g., a direction extending from the opening 44 toward an opposing end of the respective sub-cavity 42) in which the cross-sectional area of the respective sub-cavity 42 decreases, until the at least one core 46 is tightly fitted within the respective sub-cavity 42 and in contact with one or more of the first composite skin 22, the second composite skin 24, and the adjacent spar of the plurality of spars 40.
In some embodiments, the method 1000 may optionally include curing the adhesive applied to the at least one core 46, as provided in step 1014, subsequent to insertion of the at least one core 46 into the composite structure 20. Similar to the curing process used for the composite skins 22, 24, curing the adhesive may include heating the composite structure 20 to an elevated temperature and holding the composite structure 20 at the elevated temperature for a sufficient time to cure the adhesive. Various temperatures, pressure, and curing times may be used, depending on the particular adhesive selected. In some embodiments, the adhesive may not require the use of a curing process.
It is noted that various connections are set forth between elements in the preceding description and in the drawings. It is noted that these connections are general and, unless specified otherwise, may be direct or indirect and that this specification is not intended to be limiting in this respect. A coupling between two or more entities may refer to a direct connection or an indirect connection. An indirect connection may incorporate one or more intervening entities. It is further noted that various method or process steps for embodiments of the present disclosure are described in the following description and drawings. The description may present the method and/or process steps as a particular sequence. However, to the extent that the method or process does not rely on the particular order of steps set forth herein, the method or process should not be limited to the particular sequence of steps described. As one of ordinary skill in the art would appreciate, other sequences of steps may be possible. Therefore, the particular order of the steps set forth in the description should not be construed as a limitation.
Furthermore, no element, component, or method step in the present disclosure is intended to be dedicated to the public regardless of whether the element, component, or method step is explicitly recited in the claims. No claim element herein is to be construed under the provisions of 35 U.S.C. 112(f) unless the element is expressly recited using the phrase “means for.” As used herein, the terms “comprises”, “comprising”, or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.
While various aspects of the present disclosure have been disclosed, it will be apparent to those of ordinary skill in the art that many more embodiments and implementations are possible within the scope of the present disclosure. For example, the present disclosure as described herein includes several aspects and embodiments that include particular features. Although these particular features may be described individually, it is within the scope of the present disclosure that some or all of these features may be combined with any one of the aspects and remain within the scope of the present disclosure. References to “various embodiments,” “one embodiment,” “an embodiment,” “an example embodiment,” etc., indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Further, when a particular feature, structure, or characteristic is described in connection with an embodiment, it is submitted that it is within the knowledge of one skilled in the art to effect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described. Accordingly, the present disclosure is not to be restricted except in light of the attached claims and their equivalents.
