CONSOLIDATING THERMOPLASTIC MATERIAL WITH INDUCTION HEATING COIL AND HEATING PLATE(S)

A repair method is provided during which a thermoplastic patch is arranged on a thermoplastic aircraft component. The thermoplastic patch and the thermoplastic aircraft component each include a thermoplastic matrix and a plurality of fibers embedded within the thermoplastic matrix. A metal plate is arranged against a stack. The stack includes the thermoplastic patch and the thermoplastic aircraft component. The thermoplastic patch and the thermoplastic aircraft component are consolidated together. The consolidating includes concurrently induction heating the thermoplastic patch, the thermoplastic aircraft component and the metal plate using an induction heating coil.

BACKGROUND

1. Technical Field

This disclosure relates generally to joining methods and, more particularly, to consolidating thermoplastic material together.

2. Background Information

Various systems and methods are known in the art for joining thermoplastic materials together. While these known joining systems and methods have various benefits, there is still room in the art for improvement.

SUMMARY OF THE DISCLOSURE

According to an aspect of the present disclosure, a repair method is provided during which a thermoplastic patch is arranged on a thermoplastic aircraft component. The thermoplastic patch and the thermoplastic aircraft component each include a thermoplastic matrix and a plurality of fibers embedded within the thermoplastic matrix. A metal plate is arranged against a stack. The stack includes the thermoplastic patch and the thermoplastic aircraft component. The thermoplastic patch and the thermoplastic aircraft component are consolidated together. The consolidating includes concurrently induction heating the thermoplastic patch, the thermoplastic aircraft component and the metal plate using an induction heating coil.

According to another aspect of the present disclosure, a method is provided during which a first thermoplastic body is arranged on a second thermoplastic body. The first thermoplastic body and the second thermoplastic body each include a thermoplastic matrix and a plurality of fibers embedded within the thermoplastic matrix. A stack is arranged between a first magnetic plate and a second magnetic plate. The stack includes the first thermoplastic body and the second thermoplastic body. The stack is pressed between the first magnetic plate and the second magnetic plate. At least the first thermoplastic body and the second thermoplastic body are consolidated together to form a thermoplastic component. The consolidating includes induction heating the first thermoplastic body, the second thermoplastic body, the first magnetic plate and the second magnetic plate using an induction heating coil.

According to still another aspect of the present disclosure, another method is provided during which a first thermoplastic body is arranged on a second thermoplastic body. The first thermoplastic body and the second thermoplastic body each include a thermoplastic matrix and a plurality of fibers embedded within the thermoplastic matrix. A magnetic plate is arranged against a stack which includes the first thermoplastic body and the second thermoplastic body. The first thermoplastic body is pressed against the second thermoplastic body at a pressure less than two atmospheric pressure. The first thermoplastic body, the second thermoplastic body and the magnetic plate are induction heated to consolidate at least the first thermoplastic body and the second thermoplastic body together to form a thermoplastic component for an aircraft.

The first magnetic plate may have a first uninterrupted surface engaging the stack. The second magnetic plate may also or alternatively have a second uninterrupted surface engaging the stack.

The method may also include configuring the stack, the first magnetic plate and the second magnetic plate within a vacuum bag. The stack may be pressed between the first magnetic plate and the second magnetic plate using the vacuum bag.

The induction heating may include exciting at least some of the fibers in the thermoplastic patch, at least some of the fibers in the thermoplastic aircraft component and at least some material of the metal plate with eddy currents generated by the induction heating coil.

The metal plate may be induction heated to provide a heated metal plate. The consolidating may also include conduction heating the stack with heat energy from the heated metal plate.

The metal plate may be arranged against the thermoplastic patch.

The method may also include arranging a second metal plate against the thermoplastic aircraft component. The stack may be located between the metal plate and the second metal plate. The consolidating may include concurrently induction heating the thermoplastic patch, the thermoplastic aircraft component, the metal plate and the second metal plate using the induction heating coil.

The method may also include pressing the stack between the metal plate and the second metal plate during the consolidating.

The method may also include pressing the thermoplastic patch between the metal plate and the thermoplastic aircraft component during the consolidating.

The metal plate may be arranged against the thermoplastic aircraft component.

The method may also include pressing the thermoplastic aircraft component between the metal plate and the thermoplastic patch during the consolidating.

The method may also include pressing the thermoplastic patch against the thermoplastic aircraft component at a pressure less than two atmospheric pressure.

A surface of the metal plate contacting the stack may be covered with a protective material.

The aircraft component may be configured with an aircraft during the consolidating.

The method may also include supporting the stack on a support structure during the consolidating with a protective layer between the stack and the support structure.

The fibers may include a plurality of carbon fibers.

The fibers in at least one layer of the thermoplastic patch may be unidirectional.

In addition or alternatively, the fibers in at least one layer of the thermoplastic aircraft component may be unidirectional, at least along the thermoplastic patch.

The thermoplastic patch may include a plurality of unconsolidated patch layers prior to the consolidating.

DETAILED DESCRIPTION

The present disclosure includes methods and systems for manufacturing a thermoplastic component20(e.g., a thermoplastic composite component) of an aircraft, where an exemplary section of the aircraft component20is shown inFIG.1. The aircraft may be an airplane, a helicopter, a drone (e.g., an unmanned aerial vehicle (UAV)) or any other manned or unmanned aerial vehicle or system.

The term “manufacturing” may describe a process for original manufacturing the aircraft component20; e.g., creating a brand new aircraft component. The term “manufacturing” may also or alternatively describe a process for remanufacturing or otherwise repairing the aircraft component20; e.g., restoring one or more features of a previously formed aircraft component to brand new condition, similar to brand new condition, better than brand new condition, etc. The aircraft component20, for example, may be repaired to fix one or more defects (e.g., cracks, wear and/or other damage) imparted during previous use of the aircraft component. The aircraft component20may also or alternatively be repaired to fix one or more defects imparted during the initial formation of the aircraft component. For ease of description, however, the aircraft component20may be described below as a repaired aircraft component, and the methods may be described as repair methods.

Referring toFIG.2, the aircraft component20may be configured as or otherwise included as part of a nacelle22of a propulsion system24for the aircraft. The aircraft component20, for example, may be (or may be part of) a component of a nacelle inlet structure26; e.g., a nacelle inlet lip (e.g., a nose lip), a nacelle outer barrel, a nacelle inner barrel, etc. In another example, the aircraft component20may be (or may be part of) another component of the propulsion system nacelle22such as a cowl28(e.g., a fan cowl), or the like. Referring toFIG.3, the aircraft component20may alternatively be configured as or otherwise included as part of an airframe30of the aircraft. The aircraft component20, for example, may be (or may be part of) an aircraft wing32, an aircraft fuselage skin34, an aircraft stabilizer36, or the like. The aircraft component20may still alternatively be configured as or otherwise included as part of a structure within the aircraft airframe30; e.g., within a cabin of the aircraft. The present disclosure, however, is not limited to manufacturing the foregoing exemplary aircraft components. Moreover, it is contemplated the methods of the present disclosure may be utilized for manufacturing (e.g., repairing or creating) non-aircraft components. However, for ease of description, the thermoplastic component is generally described below as the aircraft component20.

Referring again toFIG.1, the aircraft component20includes a plurality of thermoplastic bodies consolidated (e.g., induction welded) together to form the aircraft component20. Where the aircraft component20is a repaired aircraft component, the first thermoplastic body may be a thermoplastic component base38and the second thermoplastic body may be a thermoplastic component patch40. The component base38may be a portion or an entirety of a damaged aircraft component which is to be repaired (prior to the consolidation). The component patch40may be configured to partially or completely repair damage to the damaged aircraft component (prior to the consolidation). Alternatively, the thermoplastic bodies38and40may be discrete layers of thermoplastic material which are consolidated together to form a consolidated multi-layer patch, which patch may then be bonded to a damaged aircraft component in a subsequent operation. On the other hand, where the aircraft component20is a brand new aircraft component, each of the thermoplastic bodies38and40may be a discrete layer of thermoplastic material (prior to the consolidation), the first thermoplastic body may be a preform of the aircraft component20to which additional thermoplastic material (e.g., at least or only the second thermoplastic body40) is consolidated with, etc. However, for ease of description, the first thermoplastic body and the second thermoplastic body may be generally described below as the component base38and the component patch40.

Referring toFIG.4, each of the thermoplastic bodies38and40may include a thermoplastic matrix42and fiber-reinforcement44embedded within the thermoplastic matrix42. The thermoplastic matrix42may include a semi-crystalline thermoplastic material and/or an amorphous thermoplastic material. The fiber-reinforcement44may include a plurality of fibers46such as, but not limited to, carbon fibers (e.g., fibers of carbon fiber material). The fiber-reinforcement44and its fibers46may be arranged in one or more layers48within the respective thermoplastic body38,40and its thermoplastic matrix42. The fibers46in each layer48of the fiber-reinforcement44may be arranged in a unidirectional pattern. Alternatively, the fibers46in one or more of the layers48of the fiber-reinforcement44may be arranged in a multi-directional pattern; e.g., woven together in a weave. The fibers46in each layer48of the fiber-reinforcement44include continuous fibers and/or chopped fibers. The present disclosure, however, is not limited to the foregoing exemplary fiber types and/or fiber arrangements. Moreover, while each thermoplastic body38,40is described above as including the fiber-reinforcement44embedded within the thermoplastic matrix42, it is contemplated one of the thermoplastic bodies38and40(e.g., the component patch40) may alternatively be configured without any the fiber-reinforcement44embedded within the thermoplastic matrix42.

FIG.5is a flow diagram of a method500for manufacturing a thermoplastic component; e.g., a thermoplastic composite component. For ease of description, this manufacturing method500is described below as a repair method and with respect to the repaired aircraft component20described above. The present disclosure, however, is not limited thereto. The manufacturing method500, for example, may alternatively be performed to create a brand new aircraft component. Moreover, the manufacturing method500may be performed to repair or create components with various other structural configurations than that described above.

In step502, the component base38(the first thermoplastic body) is provided. The damaged component to be repaired, for example, may be removed from the aircraft and/or otherwise received and prepared for patching. Alternatively, the damaged component to be repaired may be prepared for patching while still installed with or otherwise onboard the aircraft.

In step504, the component patch40(the second thermoplastic body) is provided. A piece of thermoplastic stock material, for example, may be cutout to form the component patch40. The thermoplastic stock material may be a (e.g., laminated) sheet of thermoplastic composite material, prepreg material, etc. The component patch40, of course, may alternatively be laminated, molded, pressed, injection molded, stamped and/or otherwise formed.

In step506, the component patch40is arranged with the component base38for consolidation; e.g., induction welding together. For example, the component patch40ofFIG.6is disposed on the component base38. The component base38ofFIG.6has a base thickness that extends vertically between and to a first (e.g., lower) surface50of the component base38and a second (e.g., upper) surface52of the component base38. The component patch40ofFIG.6has a patch thickness that extends vertically between and to a first (e.g., lower) surface54of the component patch40and a second (e.g., upper) surface56of the component patch40. The patch first surface54ofFIG.6is abutted against and contacts the base second surface52. The patch first surface54, for example, may lay against (e.g., rest on, be disposed in full contact with, be disposed flat against, etc.) the base second surface52.

While the surfaces52and54are shown with straight-line sectional geometries in the plane ofFIG.6, it is contemplated the surfaces52and54may alternatively have non-straight-line (e.g., curved, compound, etc.) sectional geometries in the plane ofFIG.6. Moreover, the surfaces52and54may also or alternatively have straight-line or non-straight-line sectional geometries in a plane perpendicular to the plane ofFIG.6. For example, the surfaces52and54may be flat, planar surfaces, two-dimensional (2D) curved or otherwise non-flat surfaces, or three-dimensional (3D) curved or otherwise non-flat surfaces.

In step508, one or more conduction heating plates58A and58B (generally referred to as “58”) are arranged with a stack60of the thermoplastic bodies38and40. Each heating plate58may be constructed from a magnetic material such as stainless steel (SS). The first (e.g., lower) heating plate58A has a plate thickness that extends vertically between and to a first (e.g., lower) surface62of the first heating plate58A and a second (e.g., upper) surface64of the first heating plate58A. The first plate second surface64is abutted against and engages (e.g., contacts) the stack60at a first (e.g., lower) side of the stack60. The first heating plate58A, for example, is abutted against the component base38. The first plate second surface64, for example, may lay against (e.g., rest on, be disposed in full contact with, be disposed flat against, etc.) the base first surface50. Similarly, the second (e.g., upper) heating plate58B has a plate thickness that extends vertically between and to a first (e.g., lower) surface66of the second heating plate58B and a second (e.g., upper) surface68of the second heating plate58B. The second plate first surface66is abutted against and engages (e.g., contacts) the stack60at a second (e.g., upper) side of the stack60. The second heating plate58B, for example, is abutted against the component patch40(and the component base38in some embodiments; e.g., seeFIG.7). The second plate first surface66, for example, may lay against (e.g., rest on, be disposed in full contact with, be disposed flat against, etc.) the patch second surface56(and the base second surface52in some embodiments; e.g., seeFIG.7). With this arrangement, the stack60is located (e.g., sandwiched) vertically between the first heating plate58A and the second heating plate58B.

While the surfaces50and64,56and66are shown with straight-line sectional geometries in the plane ofFIG.6, it is contemplated the surfaces50and64,56and66may alternatively have non-straight-line (e.g., curved, compound, etc.) sectional geometries in the plane ofFIG.6. Moreover, the surfaces50and64,56and66may also or alternatively have straight-line or non-straight-line sectional geometries in a plane perpendicular to the plane ofFIG.6. For example, the surfaces50and64,56and66may be flat, planar surfaces, two-dimensional (2D) curved or otherwise non-flat surfaces, or three-dimensional (3D) curved or otherwise non-flat surfaces. However, each surface50,56,64,66may be an uninterrupted surface. Herein, the term “uninterrupted” may describe a surface without any discrete apertures (e.g., channels, dimples, notches, grooves, etc.) and/or without any discrete protrusions (e.g., ribs, bumps, etc.) The present disclosure, however, is not limited to such an uninterrupted surface arrangement.

In some embodiments, one or more of the heating plates58may each be coated with a protective coating. One or more of the heating plate surfaces64and/or66ofFIG.6, for example, may be coated with the protective coating. This protective coating may be selected to facilitate release of the thermoplastic bodies38and40from the heating plates58following consolidation. For example, the protective coating may be constructed from Frekote® material manufactured by Henkel Adhesive Technologies of Düsseldorf, Germany.

In step510, the stack60is compressed. The stack60and the heating plates58ofFIG.6, for example, may be configured within a vacuum bag70. This vacuum bag70may be formed by a first (e.g., lower) sheet72and a second (e.g., upper) sheet74. Each vacuum bag sheet72,74may be constructed from a material which does not (e.g., permanently) bond with the heating plates58during the manufacturing method500. For example, the material may be a polyimide such as Kapton® polyimide material manufactured by DowDuPont Company (formerly DuPont de Nemours, Inc. of Delaware, United States). The second sheet74ofFIG.6is attached to the first sheet72at (e.g., on, adjacent or proximate) an outer periphery of the second sheet74to form an internal volume76(e.g., a vacuum bag cavity) between the first sheet72and the second sheet74. The stack60and the heating plates58are disposed within this internal volume76, and a vacuum is applied to the vacuum bag70to compress the stack60vertically between the heating plates58. More particularly, the component base38is pressed between the first heating plate58A and the component patch40, and the component patch40is pressed between the second heating plate58B and the component base38. Of course, various other techniques are known in the art for applying pressure to (e.g., compressing) a stack of thermoplastic material for consolidation, and the present disclosure is not limited to any particular ones thereof.

The vacuum bagged (or otherwise compressed) arrangement of the thermoplastic bodies38and40and the heating plates58may be supported on a rigid support structure78; e.g., a processing table. The vacuum bagged arrangement ofFIG.6, for example, is disposed on a protective layer80(e.g., a silicon sheet), which protective layer80is disposed on the support structure78. Of course, it is contemplated the vacuum bagged arrangement may alternatively be disposed directly on the support structure78without the protective layer80in other embodiments.

In step512, an induction heating coil82(e.g., an induction welding coil) is arranged with the (e.g., compressed) stack60. The induction heating coil82, for example, may be disposed over and slightly spaced from) the vacuum bagged arrangement. Here, the second heating plate58B is positioned vertically between the stack60and the induction heating coil82, and the arrangement of the stack60and the heating plates58is arranged vertically between the support structure78and the induction heating coil82. Referring toFIG.8, the induction heating coil82may include one or more coil segments84, where each coil segment84may be arranged to be substantially parallel with at least some or all of the fibers46(schematically shown) within one or more of the thermoplastic bodies38and40.

In step514, the thermoplastic bodies38and40ofFIG.6are consolidated together while being compressed at a consolidation pressure; e.g., via the vacuum bag70. This consolidation is performed using (e.g., only) the induction heating coil82; e.g., without any other heating devices such as other electric heater(s), an oven, an autoclave, etc. A power source86coupled to the induction heating coil82, for example, may provide a high frequency (e.g., alternating) current to the induction heating coil82. The induction heating coil82may subsequently generate electromagnetic waves which concurrently excite (a) at least some (or all) of the fibers46(seeFIGS.4and8) within each thermoplastic composite body38,40and/or (b) at least some (or all) of the material forming each heating plate58. The fibers46and heating plate material are excited by eddy currents propagated within the respective elements38,40,58A and/or58B. The excitation of the fibers46may elevate a temperature of each thermoplastic body38,40. The excitation of the heating plate material may elevate a temperature of each heating plate58. Each heated heating plate58may subsequently further elevate the temperature of each thermoplastic body38,40by conducting heat energy into the respective thermoplastic body38,40. The thermoplastic bodies38and40are heated to a melting point temperature of its thermoplastic matrix42(seeFIG.4) such that the thermoplastic matrix42of each thermoplastic body38,40melts. A melt layer88(seeFIG.1) may form at an interface between the thermoplastic bodies38and40. This melt layer88may bond the thermoplastic bodies38and40together upon cooling thereof. By induction heating the thermoplastic bodies38and40directly and indirectly heating the thermoplastic bodies38and40through the heating plates58, the thermoplastic bodies38and40may be consolidated together using relatively low power from the power source86as compared to, for example, power required for a typical autoclave, oven or hot stamp press.

Using the foregoing consolidation step514, the stack60may be compressed during the consolidation at a relatively low consolidation pressure. The consolidation pressure, for example, may be equal to or less than two atmospheric pressure (2 atm) during the consolidation. For example, the consolidation pressure may be maintained at about (e.g., +/−0.1) or exactly one atmospheric pressure (1 atm). By contrast, other thermoplastic welding techniques may require pressures greater than three atmospheric pressure (3 atm).

Following the consolidation, the now consolidated thermoplastic bodies38and40are removed from the vacuum bag70. The heating plates58are removed as well to provide the aircraft component20; e.g., the repaired aircraft component.

While the aircraft component20ofFIG.6includes the single component patch40, it is contemplated the aircraft component20may alternatively include multiple component patches40A-E (generally referred to as “40”) as shown inFIGS.9-11; e.g., a multi-layered patch/patch structure. Here, the multiple component patches40may be consolidated with the component base38simultaneously using the consolidation step514(e.g., induction welding) described above. Of course, in other embodiments, it is contemplated each component patch40(or sets of some of the component patches40) may alternatively be discretely consolidated with the component base38where the multi-layered patch/patch structure, for example, is relatively thick.

In some embodiments, referring toFIG.10, the one or more component patches40may be configured to cover a damaged portion90of the component base38. In other embodiments, referring toFIG.11, the one or more component patches40may be configured to at least partially or completely replace the damaged portion of the component base38. For example, some or all of the damaged portion of the component base38may be removed from the component base38via one or more removal techniques; e.g., machining, milling, cutting, grit blasting, etc. This removal may form an aperture92(e.g., a tapered aperture, a plateaued aperture, etc.) in the component base38. The aperture92ofFIG.11projects (e.g., partially) vertically into the component base38from an exterior surface (e.g.,52) of the component base38to an interior surface94of the component base38. Following this removal, the one or more of the component patches40disposed with (e.g., in, along and/or over) the aperture92and subsequently consolidated with the component base38as described above.