Patent Application
20250058502
An example embodiment relates generally to an apparatus, system, and method for manufacturing of composite components, and more particularly, to an apparatus, system, and method for creating complex-geometry composite parts with the use of dissolvable internal support structures.
Composite structures are often formed from composite plies that are stacked upon one another and bonded together to form the composite structure. Composite structures are used in a wide variety of industries as they are generally lightweight, yet very strong. This combination of strong yet lightweight is highly desirable for a wide range of applications, ranging from wind turbine blades, to automotive structures, to aerospace vehicle components among many others. Historically, the complexity of composite structures was limited based on manufacturing techniques. Composite plies were laid over a mold with a resin either applied or embedded within the composite plies. The plies were then vacuum formed to the mold, and the mold was heated, often in an autoclave. This process required that all surfaces of the mold have draft angles, and complex curvatures and angles were often not feasible. Some composite parts are impossible to make using conventional mold-based manufacturing processes.
To form complex structures often required multiple composite parts to be joined together. This increases a weight of the composite structure and can introduce weak points in the structure through events such as delamination. Thus, such composite part formation is not ideal. Embodiments described herein are able to produce composite parts with complex shapes, without the drawbacks of prior processes.
A method, system, and apparatus are provided in accordance with an example embodiment in order to manufacture composite components, and more particularly, to an apparatus, system, and method for creating complex-geometry composite parts with the use of dissolvable internal support structures. Embodiments provided herein include method for forming a composite part including: forming a core from a dissolvable material and defining a channel formed therein; layering composite material layers over the core and into the channel; inserting a spar core into the channel having the composite material layers therein; layering additional composite material layers over the core and spar core disposed within the channel; curing the composite material layers and the additional composite material layers; and dissolving the core to leave cured composite layers of the composite part.
According to some embodiments the spar core is formed of the dissolvable material, where the method further includes dissolving the dissolvable material of the spar core. Forming the core from the dissolvable material of an example embodiment includes using three-dimensional (3D) printing to form the core. According to some embodiments the dissolvable material includes polyvinyl alcohol. The core of some embodiments defines a convoluted shape, and the channel defines a convoluted shape, and the spar defines a convoluted shape congruent with the channel.
The method of some embodiments further includes inserting the core with the composite material layers and the additional composite material layers into a vacuum bag as a bagged component; sealing the bagged component as a sealed, bagged component; and drawing a vacuum on the bag about the bagged component, where curing the composite material layers and the additional composite material layers includes curing the sealed, bagged component. According to certain embodiments the spar core is formed with bulkheads disposed therein defining chamber cores between the bulkheads. The chamber cores of some embodiments are connected to one another by a channel core. According to certain embodiments the spar core is formed of a dissolvable material, where the method further includes dissolving the dissolvable material of the spar core and the chamber cores dissolve to form orifices between chambers of a spar.
Embodiments provided herein include a system for forming a composite part including: a core formed from a dissolvable material and defining a channel formed therein; composite material layers layered over the core and into the channel; a spar core received within the channel having the additional composite material layers therein; additional composite layers are layered over the core and the spar core; where the composite material layers and the additional composite layers are cured, and where the core is dissolved to leave cured composite material layers of the composite part.
According to some embodiments the spar core is formed of the dissolvable material, where the spar core is dissolved to leave a hollow spar of cured additional composite layers. According to certain embodiments the core formed from the dissolvable material is formed using 3D printing. The dissolvable material of some embodiments includes polyvinyl alcohol. The core of some embodiments defines a convoluted shape and wherein the channel defines a convoluted shape, and wherein the spar defines a convoluted shape congruent with the channel.
The system of some embodiments further includes: a vacuum bag, where the core with the composite material layers and the additional composite material layers is inserted into the vacuum bag and sealed as a sealed, bagged component; where the composite material layers and the additional composite layers are cured as the sealed, bagged component. The spar core of an example embodiment includes bulkheads disposed therein defining chamber cores between the bulkheads. The chamber cores of some embodiments are connected to one another by a channel core. According to certain embodiments the spar core is formed of the dissolvable material, where the spar core is dissolved to leave a hollow spar of cured additional composite layers and where the chamber cores dissolve to form orifices between chambers of a spar.
Embodiments provided herein include a method including: forming a core from a dissolvable material; layering composite material layers over the core; curing the composite material layers; and dissolving the core to leave cured composite layers of the composite part.
According to some embodiments, forming the core from the dissolvable material includes using 3D printing to form the core. The dissolvable material of an example embodiment includes polyvinyl alcohol. The core of an example embodiment includes a channel formed therein, where layering composite material layers includes layering composite material layers within the channel of the core. The method of an example embodiment further comprises: inserting a spar core formed from a dissolvable material into the channel; and layering additional composite material layers over the spar and the composite material layers. The method of some embodiments includes curing the additional composite material layers; and dissolving the spar core from the cured composite part. According to certain embodiments, the composite material layers and the additional composite material layers are cured at the same time.
Embodiments provided herein include a system for forming a composite part including: a core formed from a dissolvable material; and composite material layers layered over the core, where the composite material layers are cured, and where the core is dissolved to leave cured composite layers of the composite part. The core formed from the dissolvable material is, in some embodiments, formed using 3D printing. The dissolvable material of some embodiments is polyvinyl alcohol.
According to some embodiments, the core includes a channel, where the composite material layers layered over the core include composite material layers layered within the channel of the core. The system of some embodiments further includes a spar core formed from a dissolvable material inserted into the channel, and additional composite material layers layered over the spar core and the composite material layers. The additional composite material layers are, in some embodiments, cured and the spar core is dissolved from the cured composite part. The composite material layers and the additional composite material layers are, in some embodiments, cured at a same time.
Having described certain examples of the present disclosure in general terms above, reference will now be made to the accompanying drawings, which are not necessarily drawn to scale and wherein:
Some examples of the present disclosure will now be described more fully hereinafter with reference to the accompanying drawings, in which some, but not all examples of the present disclosure are shown. Indeed, the present disclosure may be embodied in many different forms and should not be construed as limited to the examples set forth herein; rather, these examples are provided so that this disclosure will satisfy applicable legal requirements. Like numbers refer to like elements throughout. As used herein, the terms “data,” “content,” “information,” and similar terms may be used interchangeably to refer to data capable of being transmitted, received, and/or stored in accordance with examples of the present disclosure. Thus, use of any such terms should not be taken to limit the spirit and scope of the present disclosure.
An apparatus, system, and method are provided in order to manufacture composite components, and more particularly, to an apparatus, system, and method for creating complex-geometry composite parts with dissolvable internal support structures. Embodiments described herein enable the creation of complex geometry parts and composites that would otherwise require core material to be permanently trapped inside a final composite structure if the structure were to be made of a unitary piece. Embodiments create high-accuracy, ultra-low weight composite parts without requiring full composite molds.
Embodiments described herein provide a manufacturing method that uses dissolvable material to act as a sacrificial internal structure to support the composite structure while in its uncured sate. In the uncured state, composite structures are flexible and, in most scenarios, cannot support their own weight without deflection, much less hold their shape. Only after the curing process do composite parts retain their final material properties or are then sufficiently rigid to be cured in an autoclave to achieve their final material properties and become rigid. Many composite parts that include cavities and complex shapes require core materials to support the composite before the composite structure is cured. In current manufacturing methods, the core material is typically non-removable and becomes a permanent addition to the final part. In some instances, this core material is not useful and provides little structural strength to the final part. This is because the majority of strength derived from these structural additions made by these methods is from an I-beam effect, created by two separate plates. As these plates separate, they add rigidity to the composite structure without requiring additional weight.
In some instances, the core material is not always unnecessary or even unwanted. There are some instances when large surface areas need to have two plates or surfaces separated, such as in composite structures that employ a honeycomb core material sandwiched between composite sheets, such as carbon fiber. The large surfaces allow the effects of buckling to become prominent, negating the I-beam effect. In such scenarios, the core materials are a welcome addition as they can do a great deal to prevent this buckling. Although in many applications, such core material is undesirable.
The fabrication of composite parts having complex geometries is challenging, particularly closed-shape parts requiring a core.
For cross-sectional scales such as the illustrated embodiment, the core material offers very little structural support and negates the use of the stringer as a conduit for wires, pipes, sensors, and other helpful equipment. The stringer core 110 of an example embodiment is formed of a dissolvable material, such as being printed in PVA (polyvinyl alcohol) filament. After the composite layers are laid up over the core 100 including the stringer core 110, the stringer core can be dissolved out of the final part along with the core 100, reducing a weight of the final part and opening up the channel to be used for other components as described above. This process further strengthens the composite part, even after the stringer core has been dissolved away. The core can be formed with holes or voids, or the holes can be drilled after curing of the part to encourage faster dissolving of the core once the composite part is cured.
The process of an example embodiment is shown in
The dissolvable material can, in some embodiments, be recaptured, processed, and re-used. While the illustrated embodiment of
In an example embodiment in which the wrapped core 220 is first cured before the spar core 210 is received and subsequently wrapped, the cured, wrapped core would require sanding and surface preparation to be able to accept the adhesion of the next, outer layer of composites to bond to it mechanically. This mechanical bond is typically weaker than the chemical bond the layers can experience when cured together as one part. This is significant as a poor connection between layers can lead to delamination and failure of the part. There is also considerable time savings involved when requiring only one curing cycle. A typical composite epoxy requires around 24 hours to cure before it can be removed from a mold or sanded and modified. In the scenario where the core is made from a mold ahead of receiving the spar core, the process requires two complete curing cycles, one for each layer. This contrasts with a single curing cycle needed if the core is made from a dissolvable material. Forming the component monolithically with a single cure cycle requires no secondary cure cycle and thus does not necessitate any mechanical bonds. This increases the strength of parts molded according to example embodiments described herein, while lowing manufacturing time and increasing a lifespan of the part. Further, the chances of separation or delamination of support structures such as ribs from a main component of the skin or shell is less likely when they are formed monolithically instead of in separate cures as separate parts.
For the parts described above to be formed using conventional molding techniques without a permanent core, the parts would require separate fabrication with subsequent joining of the parts. For example, a shell molded in multiple pieces with the spar molded separately, and then joined. In such an embodiment, the shell would be very delicate and easily structurally compromised. The addition of the spar would require adhesive, and the strength of the assembly would rely on the spar fitting sufficiently well within the shell with the adhesive there between. This introduces potential failure points at the adhesive and the interface between the spar and shell. Further, this two-part assembly would generally be heavier than if the part were formed as described by the processes of the present disclosure.
While composite parts have a wide array of applications, composite parts are particularly desirable in the aerospace industry. Weight reduction in aircraft is imperative to improving performance and fuel economy. Further, airline parts have unique requirements for strength, flexibility, and performance that is less critical in many terrestrial applications. Many aircraft share remarkable similarities in their internal structures. Aircraft frames generally use a combination of stringers, spars, and other structural elements to form their internal frames. The purpose of these different components is to add rigidity and distribute loads experienced by the airframe. Employing embodiments described herein of dissolvable core composite parts, many of these stringers and potentially even main wing spars can be replaced with more efficient composite versions of these parts.
Flight efficiency is heavily dependent upon properties of a vehicle that takes flight. Vehicles designed to fly can include aircraft and spacecraft, which may be manned or unmanned and can be of any size; however, all vehicles designed for flight will herein be referred to generally as aircraft, despite some such vehicles designed to operate in space in the absence of air. Large aircraft can employ large propulsion systems that carry heavier payloads such as passengers and/or cargo for aircraft and payloads such as satellites or passengers into orbit for spacecraft.
One factor in scaling propulsion systems is a power-to-weight ratio of the propulsion systems. Small UAVs may use various types of propulsion systems; however, most cost-effective propulsion systems include electric propulsion, which often includes a motor-driven rotor. Such propulsion systems require a power source, such as a battery to power the motor, and batteries are often relatively heavy and dense compared to other components of UAVs. As such, weight reduction in small UAVs may rely on reducing the weight of any component in an effort to improve the flight range for the UAV. Every gram of weight can impact the range and efficiency of small UAVs.
One specific type of UAV that faces great challenges is micro-aerial vehicles or MAVs. This market is highly focused on weight savings as every opportunity to reduce a weight of the craft by as little as a gram is of significant value to adding to battery life and, therefore, flight time/range. MAVs can be very compact, such as 100 millimeters long or less in some embodiments. To help save weight, composite materials are often employed. However, conventional manufacturing techniques limit the amount of weight reduction of complex parts, particularly when the parts are very small and potentially very delicate.
In the ever growing aircraft fields of UAVs and vertical take-off and landing vehicles (VTOLs), specifically MAVs, there is a major need for weight reduction in parts without compromising structural integrity. While battery power density is increasing substantially, batteries still represent a substantial load for battery-powered vehicles. Thus, with batteries being a limiting factor, it is important to find other areas where weight savings can be achieved. Embodiments provided herein can reduce the weight of structural components while forming smooth, strong, and slender parts that reduce weight and improve efficiency. Embodiments can be employed, for example, to form long, slender wings for a plane-like electric drone, while enabling the wings to carry batteries thereby reducing the components needed to be housed in a fuselage. A smaller fuselage reduces drag and material cost, while increasing range. For MAVs, parts can be manufactured faster as they can be made monolithically and can take on shapes otherwise not possible for improved strength-to-weight ratio.
Further benefits of employing composite parts as described herein allows many of these structural elements to function as conduits and enables a more efficient manner of hollowing out of wing space. Many structural elements are used in the structure of a wing, particularly in the main body of a wing. Fuel tanks are often formed within cavities of the wings. Employing embodiments described herein, the internal webbing of these parts could be removed and allow more volume for fuel tanks or batteries, and/or to permit cooling channels for batteries.
Current construction methos often result in straight and flat spars being used. While these components are optimized in their positional placement and general construction, they are still typically limited by reasonable and traditional manufacturing methods. With the implementation of more and more composites in aerospace manufacturing, there is opportunity to create structural elements with shapes that are optimized to the load distributions in the aircraft.
One particular example includes wing spars that may not be perfectly straight, but instead follow the paths of the highest load much more precisely, such as found through finite element analysis of the forces acting on the wing during flight. Embodiments can be employed for webbing material that better matches the curvature of complex wing and fuselage sections. Optimized parts, as found through finite element analysis, include more organic shapes that cannot be readily manufactured using traditional methods, but are possible using composite forming techniques described herein. In this scenario, a core of the shape of the optimized part could be printed, laid up around, and then the internals dissolved. When this type of optimization is applied to wings, similar organic shapes would be formed. These shapes would be extremely difficult to produce with traditional manufacturing methods.
According to some embodiments, the process described herein can be used to create both the wing core and the structural elements such as the spar.
According to example embodiments, the core can be formed without a mold, such as using three-dimensional printing. This can eliminate the need for expensive fixed molds, while still achieving a dimensionally accurate part with all of the structural benefits. Such production can be in small or large scale, and is ideally suited to prototyping with accurate structural, weight, and performance properties relative to later production models. For larger production molds of spars and other components, the dissolvable material may be cast to form the shapes needed such that a three-dimensional printing time latency can be avoided. The channels inside the ribs may be prioritized and dissolved first to increase efficiency of the dissolving of the dissolvable material.
Embodiments can be employed in conjunction with conventional molds while still benefiting from a degree of flexibility and efficiency. A non-dissolvable mold of the wing itself may be used to create the outermost skin of the wing, while dissolvable spars can be used to create spars within the wing frame. For example, as shown in
Embodiments of the present disclosure can create composite parts monolithically without the use of bolts or fasteners, which are vulnerable parts of a structure along with the apertures through which they are fastened. In many conventional parts, rivets, holes, and other connecting assemblies experience loading cycles and fatigue. This can lead to premature failure of a part and can be a significant factor when determining the lifetime of a part.
The absence of core material in larger, thin-walled parts can lead to component buckling and failing and making the application of the dissolvable structure method less versatile. However, for larger structural elements, such as wing spars, pre-formed bulkheads can be set in select locations inside of a wing spar.
As shown in
In some embodiments, certain operations of the operations above may be modified or further amplified. Furthermore, in some embodiments, additional optional operations may be included. Modifications, additions, or amplifications to the operations above may be performed in any order and in any combination.
Many modifications and other examples of the present disclosure set forth herein will come to mind to one skilled in the art to which the present disclosure pertains having the benefit of the teachings presented in the foregoing descriptions and the associated drawings. Therefore, it is to be understood that the present disclosure is not to be limited to the specific examples disclosed and that modifications and other examples are intended to be included within the scope of the appended claims. Moreover, although the foregoing descriptions and the associated drawings describe examples in the context of certain combinations of elements and/or functions, it should be appreciated that different combinations of elements and/or functions may be provided by alternative examples without departing from the scope of the appended claims. In this regard, for example, different combinations of elements and/or functions than those explicitly described above are also contemplated as may be set forth in some of the appended claims. Although specific terms are employed herein, they are used in a generic and descriptive sense only and not for purpose of limitation.
This application claims priority to U.S. Provisional Application No. 63/519,668, filed on Aug. 15, 2023, the contents of which are herein incorporated by reference in their entirety.
| Number | Date | Country | |
|---|---|---|---|
| 63519668 | Aug 2023 | US |
