Patent Application
20080277531
The field of the present disclosure relates to composite panel systems and methods, and more specifically, to asymmetric composite panels formed using a hybrid process of automated and non-automated fabrication activities.
Due to the favorable strength and weight characteristics of composite materials, the use of composites in various industries continues to expand. In aircraft manufacturing, the increasing use of composite materials and composite structural assemblies is leading to significant reductions in aircraft weight. These weight savings translate to significant improvements in fuel economy, and substantial reduction of operating costs and atmospheric emissions. For example, due in large measure to the extensive use of composites, it has been estimated that Boeing's 787 “Dreamliner” aircraft will consume an estimated 20% less fuel than comparable contemporary aircraft.
The feasibility of using composite materials to form a structure depends on many factors, including the size and complexity of the structure and the loads the structure will experience. In the context of aircraft manufacturing, the wing skin panels present formidable challenges for the use of composites. The wing skin panels must be capable of carrying very high loads. Current methods use stringers attached to the skin panels to provide stiffness, but since extra wing depth increases aerodynamic drag, the size of the stringers must be kept to a minimum, particularly in the outermost portions of the wings. In addition, composite manufacturing processes that involve extensive hand-layup activities may result in undesirably high costs and slow production rates. Composite panel systems that meet the strength and size requirements imposed by aircraft wing skin panels, and that may be manufactured in an economical manner, would therefore have considerable utility.
Hybrid composite panel systems and methods in accordance with the teachings of the present disclosure may advantageously meet the strength and size requirements imposed by aircraft wing skin panels, and may result in reduced aircraft weight, reduced operating costs, improved fuel economy, and reduced emissions.
In one embodiment, an assembly includes a primary section, a matrix member engaged with the primary section, and a secondary section engaged with the matrix member opposite the primary section. The primary section includes a plurality of first composite layers reinforced with a first reinforcing material, and the secondary section includes a plurality of second composite layers reinforced with a second reinforcing material. The primary and secondary sections are configured to bear an operating load at least partially transversely to the first and second composite layers, and are asymetrically configured such that the primary section bears a majority of the applied operating load.
In another embodiment, a vehicle includes at least one propulsion unit, and a structural assembly coupled to the at least one propulsion unit and configured to support a payload. The structural assembly includes at least one composite panel that includes a primary section, a matrix member engaged with the primary section, and a secondary section engaged with the matrix member opposite the primary section. As noted above, the primary section includes a plurality of first composite layers reinforced with a first reinforcing material, and the secondary section includes a plurality of second composite layers reinforced with a second reinforcing material. The primary and secondary sections are configured to bear an operating load at least partially transversely to the first and second composite layers, and are asymetrically configured such that the primary section bears a majority of the applied operating load.
In a further embodiment, a method of forming a composite structure includes forming a primary section including a plurality of first composite layers reinforced with a first reinforcing material; engaging a matrix member with the primary section; and forming a secondary section including a plurality of second composite layers reinforced with a second reinforcing material. The secondary section is engaged with the matrix member opposite from the primary section, wherein the primary and secondary sections are configured to bear an operating load at least partially transversely to the first and second composite layers, and the primary and secondary sections being asymetrically configured such that the primary section bears a majority of the applied operating load.
The features, functions, and advantages that have been above or will be discussed below 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.
Embodiments of systems and methods in accordance with the teachings of the present disclosure are described in detail below with reference to the following drawings.
The present disclosure teaches hybrid composite panel systems and methods. Many specific details of certain embodiments of the invention are set forth in the following description and in
In general, embodiments of hybrid composite panel systems and methods in accordance with the teachings of the present disclosure include relatively thick, load-carrying outer plies, a honeycomb core, and one or more inner fabric plies. The outer plies may include high-strength, high modulus, toughened epoxy uni-directional composite tape that is applied using one or more automated machines. The bulk of the load-carrying material resides in these outer tape plies. The honeycomb core may be positioned on the outer, load-carrying plies and then covered with a limited number of inner fabric plies that can be laid down by hand. Thus, hybrid composite panel systems and methods in accordance with the present disclosure combine the stiffer, higher strength and more durable uni-directional composite tape layers that are formed using automated processes with the less expensive, lower strength inner fabric plies that may be laid down by hand to provide a carbon composite material system having desirable stiffness, strength, weight, durability and manufacturability characteristics.
As further shown in
The low-strength portion 124 includes a secondary section 132 formed from a plurality of fabric-reinforced composite layers. In some embodiments, the layers of the secondary section 132 are formed using manual or “hand-layup” processes. A second bonding layer 134 is coupled between a stiffener section 136 and the secondary section 132. The stiffener section 136 provides stiffness to the hybrid composite panel 120. In some embodiments, the stiffener section 136 is formed of a lightweight matrix material having a plurality of open-space cells defined by intersecting thin walls of a relatively-rigid material. More specifically, in particular embodiments, the stiffener section 136 is formed of a matrix material (e.g. aluminum, titanium, non metallic resin impregnated material, Al and Ti alloys, other metals or non-metals, etc.) having polygonal or “honeycomb”-shaped cells. The low-strength portion 124 is coupled to the bonding layer 130 of the high-strength portion 122.
It may be appreciated that specific design details of the hybrid composite panel 120 (e.g. dimensions, materials, thermo-mechanical properties, etc.) may be variably adjusted to satisfy a wide variety of requirements and operating conditions. For example, in some embodiments, the primary section 126 is formed from successive layers of a fiber-reinforced, composite tape material having unidirectional fibers that are generally aligned along one axis (e.g. the principal stress direction). In alternate embodiments, however, the reinforcing fibers of the primary section 126 may be multi-directionally oriented.
In particular embodiments, the thick, durable load carrying outer plies of the primary section 126 are toughened epoxy uni-directional tape that is laid on a tool surface by automated machines. The bulk of the load carrying material may reside in these outer tape plies. Automated systems for forming composite structures using successive layers of fiber-reinforced composite tape include those systems disclosed, for example, in U.S. Pat. No. 6,799,619 B2 issued to Holmes et al., and U.S. Pat. No. 6,871,684 B2 issued to Engelbart et al. A honeycomb core may be laid over these plies and then covered with a limited number of inner fabric-reinforced plies that can be laid down by hand. This configuration combines the higher strength and stiffness uni-directional tape that is built using automation with the less expensive lower strength and stiffness inner fabric plies laid down by hand.
The reinforcing fibers may be formed using a variety of materials, including fibers containing metals, alloys, polymers, ceramics, naturally-occurring materials, synthetic materials, or any other suitable materials. A range of thermo-setting and thermo-plastic fiber-reinforced composite tape materials are generally known. For example, suitable fiber-reinforced composite tape materials that may be used in the high-strength portion 122 include those materials commercially available from Specialty Materials, Inc. of Lowell, Mass., and those materials developed by (or on behalf of) the NASA Langley Research Center of Langley, Va., and the NASA Goddard Space Flight Center of Greenbelt, Md., or any other suitable fiber-reinforced composite materials. Similarly, the fabric-reinforced composite materials used in the low-strength portion 124 may include those materials commercially available from Argosy International, Inc. of New York, N.Y., or those materials developed by (or on behalf of) the NASA Glenn Research Center of Cleveland, Ohio, or any other suitable fabric-reinforced composite materials.
Hybrid composite panels in accordance with the teachings of the present disclosure may be fabricated in a variety of ways. For example,
In this embodiment, the process 200 includes providing a suitable forming tool (or mandrel) upon which a hybrid composite panel will be partially or completely formed at 202. For example, in some embodiments, the forming tool may be shaped to form an aircraft component (e.g. a wing skin panel). At 204, the primary section 126 of the high-strength portion 122 is formed on the forming tool using an automated process. The forming of the primary section 126 at 204 may include both application and curing of the successive fiber-reinforced composite layers. Alternately, the forming at 204 may include application of the fiber-reinforced composite layers, and curing of the fiber-reinforced composite layers may occur at another portion of the process 200.
In addition, in some embodiments, the primary section 126 may be formed at 204 using automated systems for application and consolidation (e.g. positioning, compaction, curing, etc.) of fiber-reinforced composite tape materials. The reinforcing fibers within the composite layers of the primary section 126 may be unidirectional (e.g. extending along a longitudinal axis of the wing assembly 110), or alternately, may be multi-directionally oriented. As previously noted, the primary section 126 is configured to carry a majority of the applied loads experienced by the hybrid composite panel during normal operating conditions. At an optional block 205, assuming the primary section 126 has been cured during the forming at 204, the primary section 126 may be non-destructively tested for any desired characteristics (e.g. strength, porosity, flaws, etc.).
As further shown in
The secondary section 132 of the low-strength portion 124 is formed on the stiffener section 136 using a manual application process at 208. More specifically, in some embodiments, the secondary section 132 may be formed by applying successive layers of fabric-reinforced composite materials using manual or “hand-layup” processes. The forming of the secondary section 132 (at 208) may include both application and curing of the successive fabric-reinforced composite layers, or alternately, the curing of the fabric-reinforced composite layers may occur at another portion of the process 200.
At an optional block 210, one or more portions of the hybrid composite panel assembly may be cured and finished. For example, the curing at 210 may include curing (e.g. using an elevated temperature, an elevated pressure, or both) the primary section 126, the secondary section 132, or both. In particular embodiments, the primary section 126 is cured during the forming at 204, while the secondary section 132 is cured at 210 by placing the hybrid composite panel assembly into an autoclave and using a curing process involving the controlled application of elevated temperatures and/or pressures. The finishing at 210 may also include forming the protective outer layer 128 on the primary section 126, or any other desired shaping, machining, or conditioning operations.
It should be appreciated that the exemplary process 200 is one possible embodiment, and that a variety of processes in accordance with the present disclosure may be conceived. For example, in an alternate embodiment, a process for forming a composite panel assembly may include forming a high-strength build up of composite plies, curing the high-strength build up at a first elevated temperature and pressure, and non-destructively testing the high-strength build up for porosity or other characteristics. After testing, the process includes applying a stiffening matrix member to the high-strength build up, forming a low-strength build up of composite plies over the stiffening matrix member, and then curing the assembly at a second temperature and/or pressure less than the first elevated temperature and/or pressure. This alternate process advantageously allows the high-strength build up to be thoroughly inspected (e.g. for porosity) in a manner that may not be practical or possible after the high-strength build up is coupled to the stiffener and low-strength build up.
Embodiments of fabrication processes (e.g. process 200) in accordance with the present disclosure may be used to fabricate a variety of components. For example, in alternate embodiments, hybrid composite panels in accordance with the present disclosure may be used in various portions of an aircraft. More specifically, as shown in
Although the aircraft 100 shown in
Embodiments of hybrid composite panel systems and methods in accordance with the teachings of the present disclosure may provide significant advantages. For example, such hybrid composite panel systems and methods may advantageously meet the strength, weight, and size requirements imposed by demanding operating environments, such as aircraft wing skin panels and other high-load, highly-constrained environments. More specifically, embodiments of hybrid composite panels allow for thin wing development while meeting the high load carrying requirements. Thin wing development increases wing performance, resulting in reduced aircraft operating costs, improved fuel economy, and reduced emissions.
Furthermore, hybrid composite panels in accordance with the present disclosure allow the outer plies (e.g. the high-strength portion 122) to carry the bulk of the wing load. The outer ply manufacturing allows automated machines to do most of the fabrication, reducing labor hours and overall manufacturing costs. Furthermore, uni-directional tape is typically much cheaper than the comparable fabric material of similar strength, providing additional cost reduction. As noted above, in some embodiments, the outer plies may be cured and processed to a higher strength specification by curing prior to the addition of the stiffener section and the inner, fabric-reinforced layers of the secondary section. By adding the stiffener section and inner fabric layers (e.g. the low-strength portion 124) after the outer tape layers are applied, the hybrid composite panel assembly may be processed to a lower manufacturing specification which allows the use of less expensive inner fabric material and limits the number of plies needed. This advantageously reduces the amount of hand fabrication time and reduces labor costs.
It will be appreciated that the method of application of composite plies in a build up or finished product may be determined through inspection. Typically, components created using automated build up processes exhibit greater uniformity than do components formed using manual build up processes. In some embodiments, automated processes may also leave readily discernable characteristics and features (e.g. cyclic or repetitive features) within the build up that can be detected by inspection, and which may be used to ascertain the manner in which the build up was formed.
While specific embodiments of the invention have been illustrated and described herein, as noted above, many changes can be made without departing from the spirit and scope of the invention. Accordingly, the scope of the invention should not be limited by the disclosure of the specific embodiments set forth above. Instead, the invention should be determined entirely by reference to the claims that follow.
