Electronic control method and composite structure manufactured

Abstract

The invention relates to an electronic control method and a manufactured composite structure. The electronic control method comprises the following steps: picking up a first ceramic matrix composite layer sheet clamped between a first bottom backing film and a first top backing film; and stripping the first bottom backing film from the first ceramic matrix composite ply. The electronic control method further includes placing the first ceramic matrix composite ply on the tool surface, wherein the first top backing film faces away from the tool surface; and positioning a vacuum film against the first ceramic matrix composite ply on the tool surface to provide a vacuum seal against the first ceramic matrix composite ply. The electronic control method further comprises the steps that vacuumizing is conducted to pull the vacuum film to abut against the first ceramic-based composite layer piece, and therefore the first ceramic-based composite layer piece is formed into the shape of the tool surface; and releasing the vacuum. The electronic control method also includes peeling the first top backing film from the first ceramic matrix composite ply after releasing the vacuum, thereby providing a ceramic matrix composite structure having a desired shape.

Description

Technical Field

[0001] This application relates to composite structures, and more particularly to ceramic matrix composite structures and methods of making the same. Background Art

[0002] Ceramic matrix composites have different viscosities and textures than polymer matrix composites, which require different processing methods. The ceramic fibers of ceramic matrix composites are more brittle and harder than the carbon fibers of polymer matrix composites. The more brittle and harder fibers, as well as the different organic sizing components of the ceramic fibers, require different processing methods during the manufacture of ceramic matrix composite structures.

[0003] Typical ceramic matrix composite structures are manufactured using a hand lay-up process. A disadvantage of manufacturing ceramic matrix composite structures using a hand lay-up process is the variability in the quality and consistency of the ceramic matrix composite structures. Therefore, manual inspection and rework are typically required. Another disadvantage is that the hand lay-up process is time-consuming and requires skilled technicians. The overall result is an increase in cycle time and an increase in the labor cost of manufacturing ceramic matrix composite structures.

[0004] Despite the progress that has been made, those skilled in the art continue to conduct research and development work in the field of manufacturing ceramic matrix composite structures. Summary of the Invention

[0005] In one aspect, an electronically controlled method for manufacturing a non-polymer structure in a desired shape is provided. The electronically controlled method includes placing a non-polymer material ply on a forming tool.

[0006] In another aspect, an electronically controlled method for manufacturing a ceramic matrix composite structure in a desired shape is provided. The electronically controlled method includes: picking up a first ceramic matrix composite ply sandwiched between a first bottom backing film and a first top backing film; and peeling the first bottom backing film from the first ceramic matrix composite ply. The electronically controlled method further includes: placing the first ceramic matrix composite ply on a tool surface, wherein the first top backing film faces away from the tool surface; and positioning a vacuum film against the first ceramic matrix composite ply on the tool surface to provide a vacuum seal against the first ceramic matrix composite ply. The electronically controlled method further includes: evacuating to pull the vacuum film against the first ceramic matrix composite ply, thereby shaping the first ceramic matrix composite ply into the shape of the tool surface; and releasing the vacuum. The electronically controlled method further includes: after releasing the vacuum, peeling the first top backing film from the first ceramic matrix composite ply, thereby providing a ceramic matrix composite structure having a desired shape.

[0007] In yet another aspect, the manufactured composite structure includes at least one layer of non-polymer material. During the operational use of the manufactured composite structure, each ply is capable of withstanding temperatures up to 2400 degrees Fahrenheit.

[0008] Other aspects will become apparent from the following detailed description, the drawings, and the appended claims. BRIEF DESCRIPTION OF THE DRAWINGS

[0009] Figure 1 is a schematic block diagram of an apparatus for manufacturing a ceramic matrix composite structure according to an embodiment.

[0010] Figure 2A is a front view of a first non-polymer laminate that is processed to provide a ceramic matrix composite structure.

[0011] Figure 2B is with Figure 2A a front view of a second non-polymer laminate that is processed from the first non-polymer laminate to provide a ceramic matrix composite structure.

[0012] Figures 3A to 3J is a front view showing certain components of a Figure 1 manufacturing apparatus at different positions during the manufacture of a ceramic matrix composite structure.

[0013] Figure 4 is according to Figures 3A to 3J a magnified front view of an exemplary ceramic matrix composite structure manufactured.

[0014] Figure 5 is an overall flow chart depicting an exemplary method for manufacturing a ceramic matrix composite structure according to an embodiment.

[0015] Figure 6 is a flow chart depicting an exemplary electronic control method for manufacturing a ceramic matrix composite structure according to an embodiment.

[0016] Figure 7 is a flow chart depicting an exemplary electronic control method for manufacturing a ceramic matrix composite structure according to another embodiment.

[0017] Figure 8 is a flow chart of an aircraft manufacturing and maintenance method.

[0018] Figure 9 is a block diagram of an aircraft. DETAILED DESCRIPTION

[0019] This application relates to ceramic matrix composite structures and methods of manufacturing the same. The specific configurations of the ceramic matrix composite structures, the methods of manufacturing the same, and the industries implementing the structures and methods can vary. It should be understood that the following disclosure provides various embodiments or examples for implementing different features of the various embodiments. Specific examples of components and arrangements are described to simplify the disclosure. These are merely examples and are not intended to be limiting.

[0020] As an example, the following disclosure describes a ceramic matrix composite structure and a method for manufacturing at least a part of an aircraft, such as an aircraft exhaust structure. The ceramic matrix composite structure and its manufacturing method can be implemented by an original equipment manufacturer (OEM) in accordance with commercial, military, and space regulations. It is contemplated that the disclosed ceramic matrix composite structure and its manufacturing method can be implemented in many other ceramic matrix composite manufacturing industries.

[0021] Reference Figure 1 , a schematic block diagram of an apparatus 100 for manufacturing a ceramic matrix composite structure according to an embodiment is shown. The apparatus 100 includes a tool 110 having a tool surface 112. The tool surface 112 has an optional step 114 near the perimeter of the tool 110. The apparatus 100 also includes a plurality of mechanisms, which include a stripping mechanism 120, a pick-up mechanism 130, and a vacuum forming mechanism 140.

[0022] The pick-up mechanism 130 is a gripper end effector that is configured to pick up a sheet of material (e.g., a ply) and place it on the tool surface 112 of the tool 110. The pick-up mechanism 130 can include, for example, an electrostatic gripper or a vacuum gripper. The vacuum forming mechanism 140 includes a vacuum membrane 142. The structures and operations of the stripping mechanism, the pick-up mechanism, and the vacuum forming mechanism are known and conventional and will not be described further.

[0023] Reference Figure 2A , a front view of a first non-polymer ply 210 that has been processed to provide a ceramic matrix composite structure is shown. The first non-polymer ply 210 includes a first ceramic matrix composite ply 212 sandwiched between a first top backing film 211 and a first bottom backing film 213. The first ceramic matrix composite ply 212 has a first fiber reinforcement 214 that is oriented in a first direction as shown by arrow A in Figure 2A . The first fiber reinforcement 214 includes ceramic fibers, and the matrix material is a ceramic matrix material. Alternatively, the first ceramic matrix composite ply 212 can include a ceramic matrix composite ply having a fabric pre-impregnated with a matrix material, such as C f / Si or SiC f / SiC.

[0024] The first ceramic matrix composite ply 212 is a non-polymeric material and has a viscosity between about 3000 poise and 7000 poise. The tackiness of the first ceramic matrix composite ply 212 can vary according to the amount of water contained in the first ceramic matrix composite ply 212. Alternatively, the tackiness of the first ceramic matrix composite ply 212 can vary according to the amount of solvent (e.g., non-aqueous) contained in the first ceramic matrix composite ply 212. Other aqueous and non-aqueous compounds are possible. For a given volume of the first ceramic matrix composite ply 212, the weight of the first ceramic matrix composite ply 212 is less than the weight of an equivalent volume of a metallic material (e.g., steel).

[0025] Reference Figure 2B , shows a front view of a second non-polymeric material ply 220 of a ceramic matrix composite structure treated with Figure 2A the first non-polymeric ply 210 to provide a ceramic matrix composite structure. The second non-polymeric material ply 220 includes a second ceramic matrix composite ply 222 sandwiched between a second top backing film 221 and a second bottom backing film 223. The second ceramic matrix composite ply 222 has a second fiber reinforcement 224 oriented in a second direction indicated by arrow B in Figure 2B . The second direction B of the second fiber reinforcement 224 is transverse (e.g., perpendicular) to the first direction A of the first fiber reinforcement 214. It is contemplated that the second direction B of the second fiber reinforcement 224 is not transverse (e.g., parallel) to the first direction A of the first fiber reinforcement 214. The second fiber reinforcement 224 includes ceramic fibers and the matrix material is a ceramic matrix material. Alternatively, the second ceramic matrix composite ply 222 can include a ceramic matrix composite ply having a fabric pre-impregnated with a matrix material, such as f / C f Si or SiC

[0026] The second ceramic matrix composite ply 222 is a non-polymeric material and has a viscosity between about 3000 poise and 7000 poise. The tackiness of the second ceramic matrix composite ply 222 can vary according to the amount of water contained in the second ceramic matrix composite ply 222. Alternatively, the tackiness of the second ceramic matrix composite ply 222 can vary according to the amount of solvent (e.g., non-aqueous) contained in the second ceramic matrix composite ply 222. Other aqueous and non-aqueous compounds are possible. For a given volume of the second ceramic matrix composite ply 222, the weight of the second ceramic matrix composite ply 222 is less than the weight of an equivalent volume of a metallic material (e.g., steel).

[0027] Reference Figures 3A to 3J , a front view shows certain components of a Figure 1 manufacturing apparatus 100 at different positions during the manufacture of a ceramic matrix composite structure. As Figure 3AAs shown, the pick-up mechanism 130 picks up the first non-polymer laminate 210 that is missing the first bottom backing film 213 ( Figure 2A ), and the first bottom backing film 213 has been removed by the stripping mechanism 120 ( Figure 1 ). Alternatively, the first bottom backing film 213 can be removed by a stripping mechanism (not shown) located at a remote stripping station. Thus, Figure 3A the pick-up mechanism 130 in

[0028] As Figure 3B shown, the pick-up mechanism 130 lowers the first ceramic matrix composite laminate 212 and the first top backing film 211 onto the tool 110. As Figure 3C shown, the pick-up mechanism 130 is then lifted away from the tool 110, leaving the first ceramic matrix composite laminate 212 and the first top backing film 211 on the tool 110. Then, as Figure 3D shown, the vacuum film 142 ( Figure 1 ) is positioned above the tool 110, referred to as "bagging" with the vacuum film 142, and the vacuum forming mechanism 140 applies a vacuum to compact the first ceramic matrix composite laminate 212 and the first top backing film 211 into the shape of the tool 110.

[0029] After the first ceramic matrix composite laminate 212 and the first top backing film 211 are formed into the shape of the tool 110, the vacuum and the vacuum film 142 are removed, and then the first top backing film 211 is removed, leaving only the first ceramic matrix composite laminate 212 on the tool 110, as Figure 3E shown. Thus, Figure 3E the first ceramic matrix composite laminate 212 shown in

[0030] Then, as Figure 3F shown, the pick-up mechanism 130 picks up the second non-polymer material laminate 220 ( Figure 2B ) minus the second bottom backing film 223 that has been removed by the stripping mechanism 120 ( Figure 1 ). Thus, Figure 3F the pick-up mechanism 130 in

[0031] As Figure 3G shown, the pick-up mechanism 130 lowers the second ceramic matrix composite laminate 222 and the second top backing film 221 onto Figure 3E the formed first ceramic matrix composite laminate 212 of Figure 3H shown. As Figure 3IAs shown, a vacuum membrane 142 ( Figure 1 ) is positioned above the tool 110, and a vacuum forming mechanism 140 applies a vacuum to compact the second ceramic matrix composite ply 222 and the second top backing film 221 onto the Figure 3E formed first ceramic matrix composite ply 212.

[0032] After the second ceramic matrix composite ply 222 and the second top backing film 221 are formed into the Figure 3E formed first ceramic matrix composite ply 212, the vacuum and the vacuum membrane 142 are removed, and then the second top backing film 211 is removed, leaving only the second ceramic matrix composite ply 222, as Figure 3J shown. Thus, Figure 3J the second ceramic matrix composite ply 222 shown is formed into the Figure 3E shape of the formed first ceramic matrix composite ply 212 and the shape of the tool 110.

[0033] Figure 3J The result in is a ceramic matrix composite structure 400 that includes the formed second ceramic matrix composite ply 222 and the formed first ceramic matrix composite ply 212 on the tool 110. During placement of the first ceramic matrix composite ply 212 and the second ceramic matrix composite ply 222 onto the tool 110, the first fiber reinforcement 214 ( Figure 2A ) of the first ceramic matrix composite ply 212 and the second fiber reinforcement 224 ( Figure 2B ) of the second ceramic matrix composite ply 222 are oriented relative to each other such that the first fiber reinforcement 214 and the second fiber reinforcement 224 reinforce each other to provide an Figure 3J improved ceramic matrix composite structure 400 having a desired shape.

[0034] The improved ceramic matrix composite structure 400 is shown enlarged in Figure 4 . The formed first ceramic matrix composite ply 212 has an optional flange 215, and the formed second ceramic matrix composite ply 222 has an optional flange 225. The optional flanges 215, 225 depend on the shape of the tool surface 112 of the tool 110 and whether an optional step 114 ( Figure 1 ) is provided near the perimeter of the tool 110. The optional flanges 215, 225 provide an attachment interface for mounting the ceramic matrix composite structure 400.

[0035] As an example, an aircraft component or a part of an aircraft can include the ceramic matrix composite structure 400 that includes the optional flanges 215, 225. Aircraft include, for example, missiles, launch vehicles, high-speed aircraft, and rockets. Aircraft components include, for example, engine exhaust structures. Other types of aircraft and other aircraft components or systems are possible.

[0036] Although the exemplary ceramic matrix composite structure 400 described above includes two laminae (i.e., the first ceramic matrix composite lamina 212 and the second ceramic matrix composite lamina 222), it is contemplated that the ceramic matrix composite structure includes three or more laminae. It is also contemplated that the ceramic matrix composite structure includes only one lamina.

[0037] Moreover, although the above description describes removing the first bottom backing film 213 before placing the first ceramic matrix composite lamina 212 on the tool 110, it is contemplated that the first bottom backing film 213 is removed after the ceramic matrix composite structure 400 having the desired shape has been formed. Figure 3J of the ceramic matrix composite structure 400.

[0038] Reference Figure 5 , the overall flow chart 500 depicts an example method for manufacturing a ceramic matrix composite structure according to an embodiment. In block 502, a ceramic matrix composite lamina is picked up before peeling the bottom backing film, as shown in block 504. Then, in block 506, the ceramic matrix composite lamina is positioned on the tool before positioning the vacuum film onto the tool, as shown in block 508.

[0039] Then a vacuum is applied, as shown in block 510, to compact the ceramic matrix composite lamina to the tool. The process proceeds to block 512, where the vacuum is removed / released before peeling the top backing film, as shown in block 514. The process proceeds to block 516, where an in-situ inspection is provided to verify the successful placement, compaction, and removal of the backing films of the ceramic matrix composite lamina.

[0040] Then in block 518, it is determined whether another ceramic matrix composite lamina is to be added to manufacture the ceramic matrix composite structure. If the determination in block 518 is affirmative (i.e., another ceramic matrix composite lamina is to be added), the process returns to block 502 to process the next ceramic matrix composite lamina. However, if the determination in block 518 is negative (i.e., there are no additional ceramic matrix composite laminae), the process proceeds to block 520, where the ceramic matrix composite structure is provided. The ceramic matrix composite structure includes at least one ceramic matrix composite lamina plus any ceramic matrix composite laminae added in block 518. Then the process ends.

[0041] Reference Figure 6 , the flow chart depicts an example electronic control method 600 for manufacturing a ceramic matrix composite structure according to an embodiment. In block 602, a non-polymeric material lamina is placed on the forming tool. Then the process ends.

[0042] Reference Figure 7, Flowchart 700 depicts an example electronic control method for manufacturing a ceramic matrix composite structure according to another embodiment. In block 702, pick up a first ceramic matrix composite ply sandwiched between a first bottom backing film and a first top backing film. Then in block 704, peel the first bottom backing film from the first ceramic matrix composite ply. The process proceeds to block 706, where the first ceramic matrix composite ply is placed on a tool surface with the first top backing film facing away from the tool surface. The process proceeds to block 708.

[0043] In block 708, position a vacuum film against the first ceramic matrix composite ply on the tool surface to provide a vacuum seal against the first ceramic matrix composite ply. Then in block 710, evacuate the air to pull the vacuum film against the first ceramic matrix composite ply, thereby shaping the first ceramic matrix composite ply into the shape of the tool surface.

[0044] Before proceeding to block 714, release the vacuum in block 712. In block 714, after releasing the vacuum, peel the first top backing film from the first ceramic matrix composite ply, thereby providing a ceramic matrix composite structure having a desired shape. Then the process ends.

[0045] By providing the above-described ceramic matrix composite structure (e.g., Figure 4 the ceramic matrix composite structure 400 shown in

[0046] and its manufacturing method, many advantages are achieved. One advantage is that laying the ceramic matrix composite ply onto the tool is a fully automated process. The placement and compaction of the ply are automated, and in-situ inspection of quality measurements is provided. Quality measurements that can be inspected in-situ include, but are not limited to, ply position, fiber orientation, uncompacted areas, rework path determination, and large defects of different types and sizes.

[0047] Another advantage is that since the placement and compaction of the ceramic matrix composite ply onto the tool are automated, both the first-pass quality and the final product consistency are improved. The result is reduced rework, reduced touch labor, reduced cycle time, and thus reduced overall manufacturing cost. Figure 4The weight of the ceramic matrix composite structure 400 is less than the weight of the same structure made of non-ceramic materials (such as metals). In addition, the ability of the ceramic matrix material to withstand high temperatures during the operational use of the structure is much higher than that of non-ceramic materials to withstand the same high temperatures. The ceramic matrix material can withstand temperatures up to 2400 degrees Fahrenheit. The high-temperature capability of the ceramic matrix material allows structures made of this material (such as thermal shield aircraft components or aircraft exhaust structures) to be exposed to constant high temperatures (e.g., exceeding 1500 degrees Fahrenheit, which is the limit for most metals) during the operational use of the structure. Thus, the ceramic matrix composite structures manufactured in accordance with the present disclosure not only have the desired weight advantages but also have the desired thermal properties in applications where weight and thermal characteristics are considered important.

[0048] Examples of the present disclosure can be described in the context of an aircraft manufacturing and maintenance method 1100 as shown in Figure 8 and an aircraft 1102 as shown in Figure 9 During pre-production, the aircraft manufacturing and maintenance method 1100 can include the specification and design 1104 of the aircraft 1102 and the procurement of materials 1106. During production, component / subassembly manufacturing 1108 and system integration 1110 of the aircraft 1102 are performed. Thereafter, the aircraft 1102 can undergo certification and delivery 1112 for entry into service 1114. When used by a customer, the aircraft 1102 is scheduled for routine maintenance and repair 1116, which can also include modifications, reconfigurations, refurbishments, etc.

[0049] Each process of the aircraft manufacturing and maintenance method 1100 can 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 specification, a system integrator can include, but is not limited to, any number of aircraft manufacturers and prime system subcontractors; a third party can include, but is not limited to, any number of vendors, subcontractors, and suppliers; and an operator can be an airline, a leasing company, a military entity, a service organization, etc.

[0050] As Figure 9 shown, the aircraft 1102 produced by the aircraft manufacturing and maintenance method 1100 can include a fuselage 1118 having a plurality of systems 1120 and an interior 1122. Examples of the plurality of systems 1120 can include one or more of a propulsion system 1124, an electrical system 1126, a hydraulic system 1128, and an environmental system 1130. Any number of other systems can be included.

[0051] The disclosed devices and methods can be employed during any one or more stages of an aircraft manufacturing and maintenance method 1100. As an example, the disclosed device method can be used to assemble components or sub-components corresponding to component / sub-component manufacturing 1108, system integration 1110, and / or maintenance and repair 1116. As another example, the fuselage 1118 can be constructed using the disclosed devices and methods. Additionally, one or more device examples, method examples, or combinations thereof can be utilized during component / sub-component manufacturing 1108 and / or system integration 1110, for example, by significantly accelerating the assembly of the aircraft 1102 (such as the fuselage 1118 and / or the interior 1122) or reducing the cost of the aircraft 1102 (such as the fuselage 1118 and / or the interior 1122). Similarly, one or more of the system examples, method examples, or combinations thereof can be utilized when the aircraft 1102 is in use, such as, but not limited to, maintenance and repair 1116.

[0052] Aspects of the disclosed embodiments can be implemented in software, hardware, firmware, or a combination thereof. The various elements of the system (individually or in combination) can be implemented as a computer program product (program of instructions) tangibly embodied in a machine-readable storage device (storage medium) for execution by a processor. The various steps of the embodiments can be performed by a computer processor executing a program tangibly embodied on a computer-readable medium to perform functions by operating on inputs and generating outputs. The computer-readable medium can be, for example, a memory, a transportable medium such as an optical disc or a flash drive, such that a computer program embodying aspects of the disclosed embodiments can be loaded onto a computer.

[0053] The above-described devices and methods are described in the context of an aircraft. However, those of ordinary skill in the art will readily recognize that the disclosed devices and methods are applicable to a variety of applications, and the present disclosure is not limited to aircraft manufacturing applications. For example, the disclosed devices and methods can be implemented in various types of transportation vehicles, including, for example, helicopters, passenger ships, automobiles, marine products (boats, motors, etc.). Non-vehicle applications are also contemplated.

[0054] Moreover, although the above description describes devices and methods for manufacturing ceramic matrix composite structures for aircraft parts in the aviation industry in accordance with military and space regulations, it is contemplated that the devices and methods can be implemented to facilitate the manufacture of ceramic matrix composite structures in any industry in accordance with applicable industry standards. Specific devices and methods can be selected and customized according to a particular application.

[0055] In addition, although various aspects of the disclosed embodiments have been shown and described, modifications may occur to those skilled in the art upon reading the specification. This application includes such modifications and is limited only by the scope of the claims.

[0056] Examples of the present disclosure may be described in accordance with one or more of the following clauses.

[0057] Clause 1. An electronic control method for manufacturing a non-polymer structure in a desired shape, the electronic control method comprising: picking up a first non-polymer material laminate sandwiched between a bottom backing film and a top backing film; peeling the bottom backing film from the first non-polymer material laminate; and placing the first non-polymer material laminate on a tool surface, wherein the top backing film faces away from the tool surface.

[0058] Clause 2. The electronic control method according to Clause 1, wherein placing the non-polymer material laminate on the tool surface comprises: placing a ceramic matrix composite laminate on the tool surface.

[0059] Clause 3. The electronic control method according to Clause 2, wherein placing the non-polymer material laminate on the tool surface comprises: placing a ceramic matrix composite laminate having a fiber reinforcement on the tool surface.

[0060] Clause 4. The electronic control method according to Clause 1, wherein placing the first non-polymer material laminate on the tool surface comprises: placing a fabric pre-impregnated with a base material on the tool surface.

[0061] Clause 5. The electronic control method according to Clause 1, the electronic control method further comprising: applying a vacuum to form the first non-polymer material laminate into the shape of the tool surface, thereby providing a non-polymer structure having a desired shape.

[0062] Clause 6. The electronic control method according to Clause 5, the electronic control method further comprising: positioning a vacuum film against the first non-polymer material laminate on the tool surface before applying the vacuum to provide a vacuum seal for the first non-polymer material laminate.

[0063] Clause 7. The electronic control method according to Clause 1, the electronic control method further comprising: peeling the top backing film from the first non-polymer material laminate that has become the shape of the tool surface.

[0064] Clause 8. The electronic control method according to Clause 1, the electronic control method further comprising: placing a second non-polymer material laminate on the first non-polymer material laminate.

[0065] Clause 9. The electronic control method according to Clause 8, the electronic control method further comprising: applying a vacuum to form the second non-polymer material laminate into the shape of the first non-polymer material laminate and the tool surface, thereby forming a non-polymer structure of multiple non-polymer material laminates having a desired shape.

[0066] Clause 10. The electronic control method according to Clause 9, the electronic control method further comprising: inspecting each non-polymeric material ply and non-polymeric structure to verify ply placement and compaction and removal of the bottom backing film and the top backing film.

[0067] Clause 11. The electronic control method according to Clause 1, wherein, for a given volume of non-polymeric structure, the weight of the non-polymeric structure is less than the weight of a metallic structure of equal volume.

[0068] Clause 12. A method of manufacturing an aircraft part, the method comprising the electronic control method according to Clause 1.

[0069] Clause 13. A method of manufacturing an aircraft component having a plurality of flanges, the method comprising the electronic control method according to Clause 1.

[0070] Clause 14. A method of manufacturing a heat shield aircraft component, the method comprising the electronic control method according to Clause 1.

[0071] Clause 15. An electronic control method for manufacturing a ceramic matrix composite structure in a desired shape, the electronic control method comprising: picking up a first ceramic matrix composite ply sandwiched between a first bottom backing film and a first top backing film; peeling the first bottom backing film from the first ceramic matrix composite ply; placing the first ceramic matrix composite ply on a tool surface with the first top backing film facing away from the tool surface; positioning a vacuum film against the first ceramic matrix composite ply on the tool surface to provide a vacuum seal for the first ceramic matrix composite ply; evacuating the air, pulling the vacuum film against the first ceramic matrix composite ply so that the first ceramic matrix composite ply assumes the shape of the tool surface; releasing the vacuum; after releasing the vacuum, peeling the first top backing film from the first ceramic matrix composite ply, thereby providing a ceramic matrix composite structure having the desired shape.

[0072] Clause 16. The electronic control method according to Clause 15, wherein picking up the first ceramic matrix composite ply sandwiched between the bottom backing film and the top backing film comprises: picking up the first ceramic matrix composite ply having a substrate and a fiber reinforcement within the substrate.

[0073] Clause 17. The electronic control method according to Clause 16, wherein the substrate comprises a ceramic matrix material and the fiber reinforcement comprises ceramic fibers.

[0074] Clause 18. The electronic control method according to Clause 15, wherein picking up the first ceramic matrix composite ply sandwiched between the bottom backing film and the top backing film comprises: picking up the first ceramic matrix composite ply having a fabric pre-impregnated with a substrate material.

[0075] Clause 19. The electronic control method according to Clause 15, the electronic control method further comprising: picking up a second ceramic matrix composite laminate sandwiched between a second bottom backing film and a second top backing film.

[0076] Clause 20. The electronic control method according to Clause 19, the electronic control method further comprising: peeling the second bottom backing film from the second ceramic matrix composite laminate; placing the second ceramic matrix composite laminate on a formed first ceramic matrix composite laminate on a tool surface, wherein the second top backing film faces away from the first ceramic matrix composite laminate and the tool surface; positioning a second vacuum film against the second ceramic matrix composite laminate to provide a vacuum seal for the second ceramic matrix composite laminate; evacuating a second vacuum to pull the second vacuum film against the second ceramic matrix composite laminate, thereby shaping the first ceramic matrix composite laminate and the second ceramic matrix composite laminate to the shape of the tool surface; releasing the second vacuum; and after releasing the second vacuum, peeling the second top backing film from the second ceramic matrix composite laminate, thereby providing a ceramic matrix composite structure of a plurality of ceramic matrix composite laminates having the desired shape.

[0077] Clause 21. The electronic control method according to Clause 20, the electronic control method further comprising: during placement of the ceramic matrix composite laminate, orienting the fiber reinforcements of each of the first ceramic matrix composite laminate and the second ceramic matrix composite laminate such that the fiber reinforcements reinforce each other to provide an improved ceramic matrix composite structure having the desired shape.

[0078] Clause 22. A method of manufacturing an aircraft part, the method comprising the electronic control method according to Clause 15.

[0079] Clause 23. A method of manufacturing an aircraft component having a plurality of flanges, the method comprising the electronic control method according to Clause 15.

[0080] Clause 24. A method of manufacturing a heat shield aircraft component, the method comprising the electronic control method according to Clause 15.

[0081] Clause 25. A manufactured composite structure, the manufactured composite structure comprising: at least one non-polymeric material laminate, wherein each non-polymeric material laminate is capable of withstanding temperatures up to 2400 degrees Fahrenheit during the operational use of the manufactured composite structure.

[0082] Clause 26. The manufactured composite structure according to Clause 25, wherein the non-polymeric material laminate comprises a ceramic matrix composite laminate.

[0083] Clause 27. The fabricated composite structure according to Clause 25, wherein for at least one non-polymeric material laminate of a given volume, the weight of the at least one non-polymeric material laminate is less than the weight of a metallic material of the same volume.

[0084] Clause 28. The fabricated composite structure according to Clause 25, wherein the viscosity of the at least one non-polymeric material laminate is between approximately 3000 poise and approximately 7000 poise.

[0085] Clause 29. The fabricated composite structure according to Clause 28, wherein the tackiness of the at least one non-polymeric material laminate varies with the change in the water content contained in the at least one non-polymeric material laminate.

[0086] Clause 30. The fabricated composite structure according to Clause 28, wherein the tackiness of the at least one non-polymeric material laminate varies with the change in the solvent content contained in the at least one non-polymeric material laminate.

[0087] Clause 31. The fabricated composite structure according to Clause 25, wherein the at least one non-polymeric material laminate includes a plurality of ceramic matrix composite laminates, and each ceramic matrix composite laminate can withstand a temperature of up to 2400 degrees Fahrenheit during the operational use of the fabricated composite structure.

[0088] Clause 32. The fabricated composite structure according to Clause 31, wherein for each ceramic matrix composite laminate of a given volume, the weight of the ceramic matrix composite laminate is less than the weight of a metallic material of the same volume.

[0089] Clause 33. The fabricated composite structure according to Clause 31, wherein the viscosity of each ceramic matrix composite laminate is between approximately 3000 poise and approximately 7000 poise.

[0090] Priority

[0091] This application claims the priority of U.S. Application No. 63 / 603,867, filed on November 29, 2023, the entire content of which is incorporated herein by reference.

Claims

1. An electronically controlled method for manufacturing a non-polymer structure having a desired shape, the electronically controlled method comprising: picking up a first non-polymer material ply sandwiched between a bottom backing film and a top backing film; peeling the bottom backing film from the first non-polymeric material layer; as well as The first non-polymer material ply is placed on a tool surface with the top backing film facing away from the tool surface.

2. The electronic control method according to claim 1, wherein: The steps of placing a ply of non-polymer material on the tool surface include: A ceramic matrix composite ply is placed on the tool surface.

3. The electronic control method according to claim 2, wherein: The step of placing a ply of non-polymer material on the tool surface comprises: A ceramic matrix composite ply having fiber reinforcement is placed on the tool surface.

4. The electronic control method according to claim 1, wherein: The step of placing a first non-polymer material ply on the tool surface comprises: A fabric pre-impregnated with a substrate material is placed on the tool surface.

5. The electronic control method according to claim 1, further comprising: A vacuum is applied to form the first ply of non-polymer material into the shape of the tool surface, thereby providing a non-polymer structure having a desired shape.

6. The electronic control method according to claim 5, further comprising: Prior to applying a vacuum, a vacuum film is positioned against the first ply of non-polymer material on the tool surface to provide a vacuum seal against the first ply of non-polymer material.

7. The electronic control method according to claim 1, further comprising: The top backing film is peeled from the first non-polymer material ply having the shape of the tool surface.

8. The electronic control method according to claim 1, further comprising: A second ply of non-polymer material is placed over the first ply of non-polymer material.

9. An electronically controlled method for manufacturing a ceramic matrix composite structure having a desired shape, the electronically controlled method comprising: picking up a first ceramic matrix composite ply sandwiched between a first bottom backing film and a first top backing film; peeling the first bottom backing film from the first ceramic matrix composite ply; placing the first ceramic matrix composite ply on a tool surface with the first top backing film facing away from the tool surface; positioning a vacuum membrane against the first ceramic matrix composite ply on the tool surface to provide a vacuum seal against the first ceramic matrix composite ply; drawing a vacuum to draw the vacuum membrane against the first ceramic matrix composite ply to form the first ceramic matrix composite ply into the shape of the tool surface; Release the vacuum; as well as After releasing the vacuum, the first top backing film is peeled off from the first ceramic matrix composite ply, thereby providing the ceramic matrix composite structure having a desired shape.

10. A manufactured composite structure, the manufactured composite structure comprising: At least one ply of non-polymer material, wherein each ply of non-polymer material is capable of withstanding temperatures up to 2400 degrees Fahrenheit during operational use of the fabricated composite structure.

Images