Modular mandrel for monolithic composite fuselage

A method of manufacturing a self-supporting, monolithic fuselage body, including engaging peripheral mandrel sections around at least one central mandrel section, placing uncured composite material on the mold surface, curing the composite material on the mold surface, and sliding the central mandrel section(s) out of engagement with the peripheral mandrel sections and disengaging the peripheral mandrel sections from the cured composite material without collapsing the mandrel sections. The peripheral mandrel sections each include a shape-retaining core of a thermally insulating material and an outer layer on an outer surface of the shape-retaining core. The outer layer has a coefficient of thermal expansion within the range of variation of that of the coefficient of thermal expansion of the composite material. A mandrel for layup and cure of a predetermined composite material in the manufacture of a monolithic fuselage is also discussed.

FIELD OF THE INVENTION

The application generally relates to the manufacture of large composite structures and, more particularly, to the manufacture of monolithic structures.

BACKGROUND OF THE INVENTION

When manufacturing composite aircraft fuselages, multiple primary and secondary structural parts are typically molded, cured and trimmed separately, and then assembled together, using numerous tools for each of these operations. Each tool and operation usually increases the costs and time of the manufacturing process.

Some large composite airframe sections are manufactured using collapsible mandrels or collapsible tooling in order to be able to extract the mandrel or tooling from within the completed structure after curing. However, collapsible tooling can be relatively complex and/or more prone to failure or damage than solid tooling.

SUMMARY OF THE INVENTION

In one aspect, there is provided a method of manufacturing a self-supporting, monolithic fuselage body, the method comprising: engaging a plurality of peripheral mandrel sections around a central mandrel section, the peripheral mandrel sections each including a shape-retaining core of a thermally insulating material and an outer layer on an outer surface of the shape-retaining core, the outer layer of the peripheral mandrel sections cooperating to define a mold surface; placing uncured composite material on the mold surface to form a skin of the monolithic fuselage; curing the composite material on the mold surface by heating the composite material in a pressurized atmosphere, a coefficient of thermal expansion of the composite material varying within a predetermined range during the cure, the outer layer of the peripheral mandrel sections having a coefficient of thermal expansion within the predetermined range; after the composite material is cured, sliding the central mandrel section out of engagement with the peripheral mandrel sections without collapsing the central mandrel section; disengaging the peripheral mandrel sections from the cured composite material without collapsing the peripheral mandrel sections, the cured composite material forming the fuselage body.

In another aspect, there is provided a mandrel for layup and cure of a predetermined composite material in the manufacture of a monolithic fuselage, the mandrel comprising: a central mandrel section; a plurality of peripheral mandrel sections cooperating to surround the central mandrel section, the peripheral mandrel sections together defining a mold surface for receiving the predetermined composite material; wherein the central mandrel section is slidingly engaged to the peripheral mandrel sections and is configured to be slidable out of the fuselage after the curing of the predetermined composite material; wherein each of the peripheral mandrel sections includes a shape-retaining core of a thermally insulating material and an outer layer on an outer surface of the shape-retaining core and defining the mold surface, the outer layer being made of a material having a coefficient of thermal expansion within a range defined by a variation of a coefficient of thermal expansion of the predetermined composite material during the cure.

DETAILED DESCRIPTION

FIG. 1shows a rotorcraft100according to one example embodiment. Rotorcraft100features a rotor system110, blades120, a fuselage130, a landing gear140, and an empennage150. Rotor system110may rotate blades120. Rotor system110may include a control system for selectively controlling the pitch of each blade120in order to selectively control direction, thrust, and lift of rotorcraft100. Fuselage130represents the body of rotorcraft100and may be coupled to rotor system110such that rotor system110and blades120may move fuselage130through the air. Landing gear140supports rotorcraft100when rotorcraft100is landing and/or when rotorcraft100is at rest on the ground. Empennage150represents the tail section of the aircraft and features components of a rotor system110and blades120′. Blades120′ may provide thrust in the same direction as the rotation of blades120so as to counter the torque effect created by rotor system110and blades120. It should also be appreciated that teachings regarding rotorcraft100may apply to aircraft and vehicles other than rotorcraft, such as airplanes and unmanned aircraft, to name a few examples.

The present disclosure provides for a mandrel and method of manufacturing the body of the fuselage130in a monolithic manner. In a particular embodiment, the monolithic fuselage body includes at least 80% of the sum of parts forming the finished fuselage130of the rotorcraft100. Accordingly, 80% of the parts that make up the fuselage130can be molded, cured and bonded simultaneously without the need for separate molds and major assembly tools, such as to allow for a reduction the number of operations, tools, and/or labor to produce the finished fuselage130of the rotorcraft100.

Referring toFIG. 2-4, a mandrel200configured for forming the monolithic (i.e. integrally formed as a single piece, monocoque) fuselage body according to a particular embodiment is shown. As can be best seen inFIG. 2, the mandrel200is a modular male mandrel having a plurality of dismountable sections, including a central mandrel section202and a plurality of peripheral mandrel sections204,206,208. The peripheral mandrel sections include side mandrel sections204, a bottom mandrel section206, and a top mandrel section208, which cooperate to surround the central mandrel section202and together define a male mold surface210for receiving the uncured composite material used to form the fuselage body. When engaged to one another, the outer surfaces of the adjacent mandrel sections204,206,208are aligned with each other so as to define the mold surface210in a continuous manner.

In the embodiment shown, the mold surface210is shaped to correspond to the shape of side walls160, a top wall170, a bottom wall180, and a rear wall190of the fuselage130(seeFIG. 1). The rear of the central mandrel section202also defines part of the mold surface210, to form part of the real wall. Alternatively, the central mandrel section202could be configured so as not to contact the composite material, for example, by being completely surrounded by the peripheral mandrel sections204,206,208.

Each of the mandrel sections202,204,206,208is solid and not collapsible; it has a permanent shape, i.e., its shape remains the same throughout the layup, cure and unmolding process. The central mandrel section202is slidingly engaged to the peripheral mandrel sections204,206,208and is configured to be slidable out of the fuselage130after the composite material is cured, so that the mandrel sections202,204,206,208can be disengaged from the cured material. In the embodiment shown and as can be best seen inFIG. 4, mating surfaces of the central mandrel section202and of the peripheral mandrel sections204,206,208include parallel sets of complementary tongues212and grooves214defining the sliding engagement between the mandrel sections202,204,206,208. It is understood that any other mating configuration or mating element(s) allowing the relative sliding movement can alternatively be used.

In the embodiment shown, the mandrel sections202,204,206,208are configured so that the front surface216(seeFIG. 2) of the central mandrel section202remains unobstructed by the composite material after cure, so as to be able to slide the central mandrel section202out of the cured fuselage130through the opening configured to receive the front windshield. In an embodiment where the central mandrel section202contacts the composite material such as shown, the relative sliding movement between the mandrel sections202,204,206,208is defined along a direction at least substantially perpendicular (i.e., substantially perpendicular or perpendicular) to the part of the mold surface210defined by the central mandrel section202, so as to avoid interference from and/or damage to the cured composite material upon sliding of the central mandrel section202out of the cured body of the fuselage130.

As can be best seen inFIG. 2, additional mandrel sections218may be provided in detachable engagement with one or more of the peripheral mandrel sections204,206,208(i.e., not directly connected to the central mandrel section202) to define the mold surface210, for example, at corner junctions between walls. Some or all of the peripheral mandrel sections202,204,206,208can be made of two or more detachably interconnected parts, for example, to facilitate unmolding operations. In one example embodiment, the mandrel200includes 33 separate solid sections detachably connected to one another to form the final shape of the mandrel200.

In the embodiment shown, the mandrel200includes a shaft220and a drive tray222drivingly connected to each other and extending from opposed sides of the mandrel200, used to support and rotate the mandrel200when the composite material is placed on the mold surface210. The mandrel200can be rotated through rotation of the drive tray222. Other configurations are also possible.

Referring toFIG. 4, the central and peripheral mandrel sections202,204,206,208(and, in the embodiment shown, the additional mandrel sections218) include a shape-retaining core230and an outer layer232on an outer surface of the shape-retaining core230, in at least the portion(s) of the mandrel section defining the mold surface210. In the embodiment shown, each mandrel section202,204,206,208,218has the outer layer232completely surrounding the core230, so that the entire outer surface of the mandrel section202,204,206,208,218is defined by the outer layer232.

The core230is made of a material sufficiently rigid to withstand autoclave pressure (e.g., 85 psi, 90 psi) while allowing the mandrel section to retain its shape. The material of the core230is a thermally insulating material, i.e., a material having a thermal conductivity sufficiently low so as to inhibit conductive heat transfer across the core230. Accordingly, the core230does not define a substantial heat sink (defines no heat sink or heat sink sufficiently low so as not to have a substantial impact on the temperature of the composite material during cure). For example, in a particular embodiment, the material of the core230has a thermal conductivity having a value corresponding to one or more of the following: less than 1 W/m° K; less than 0.6 W/m° K; approximately 0.25 W/m° K. In a particular embodiment, the core230is made of thermally insulating carbon foam such as a thermally insulating grade of Cfoam®, for example, a grade of Cfoam® having a thermal conductivity of about 0.25 W/m° K. The core230forms the bulk of the central and peripheral mandrel sections202,204,206,208, so that the central and peripheral mandrel sections202,204,206,208are not significantly heated during the cure cycle of the composite material.

Each of the central and peripheral mandrel sections202,204,206,208,218has a rigidity sufficient to withstand autoclave pressure (e.g. 85 psi, 90 psi) at the temperatures of the cure cycle of the composite material.

The outer layer232is made of a material having a coefficient of thermal expansion within a range defined by the variation of a coefficient of thermal expansion of the composite material during the cure cycle. Accordingly, heating of the mandrel200during the cure cycle is focused in the outer layer232and not throughout the entire mandrel200. For example, in a particular embodiment, the composite material being cured includes carbon fibers in an epoxy resin and has a coefficient of thermal expansion varying from 1.10×10−6to 1.70×10−6when heated from 70° F. to 356° F., while the material of the outer layer232is a nickel-iron alloy such as invar36with a coefficient of thermal expansion of 1.17×10−6at 300° F., i.e., within the range of the variation of coefficient of thermal expansion of the carbon/epoxy composite. Other materials and values are also possible.

In a particular embodiment, the outer layer232is made of the same material as the composite material being cured, but of a different grade suitable for use in tooling (e.g. with greater heat resistance). For example, in a particular embodiment where the mandrel200is configured for use with a composite material including carbon fibers in an epoxy resin, the outer layer232is made of a tooling composite material including carbon fibers in an epoxy resin. Accordingly, the coefficient of thermal expansion for the outer layer232and for the material being cured may be identical or substantially identical.

It is understood that the particular geometry shown and described for the mandrel200is configured for a particular rotorcraft100and that the geometry can be adapted to suit any fuselage configuration. For example, more than one central mandrel section202can be provided. The number of mandrel sections may vary in accordance with the geometry and/or size of the fuselage body being manufactured. Other configurations are, of course, possible.

In use and in accordance with a particular embodiment, the monolithic body of the fuselage130is manufactured by engaging the peripheral mandrel sections204,206,208around the central mandrel section(s)202, and placing a plurality of layers of uncured composite material on the mold surface210defined by the mandrel sections202,204,206,208to form at least the skin of the monolithic fuselage body. The layup of composite material can be done using any suitable method or combination of methods, including, but not limited to, manual or automated layup of prepreg layers, and automated fiber placement (AFP).

In a particular embodiment, the layers of uncured composite material are disposed to also form additional structural elements (e.g., inner structural primary parts such as stiffeners, lift frame, roof beam, floor beam, bulkhead, secondary parts, etc.) in contact with the skin, which may be assembled prior to being placed on the mandrel200. In addition, or alternatively, cured structural elements (e.g., inner structural primary parts, secondary parts) may be disposed in contact with the uncured skin. It is understood that the mold surface210of the mandrel200is shaped to receive and support the additional structural elements when provided, and to adequately position the elements with respect to the skin during layup. An example of a structural primary part234is shown inFIG. 2, with the mold surface210including grooves236complementary to the structural primary part234to receive the part therein before the composite material of the skin is placed in contact with the part.

The composite material on the mold surface210, and the uncured elements, if present, are cured and bonded by heating the composite material in a pressurized atmosphere, for example, in an autoclave. It is understood that the uncured composite material and mandrel assembly is suitably prepared before the cure cycle, such as by vacuum bagging with suitable breather material and caul plates or pressure pads; such preparation methods are well known in the art and will not be discussed further herein. If cured elements were disposed in contact with the uncured material of the skin before the cure cycle, the cured elements are bonded with the skin by the heat and pressure of the cure cycle. The co-curing and/or co-bonding of the elements and skin results in a monolithic fuselage body once the cure cycle is performed.

In a particular embodiment, the uncured composite material and mandrel assembly are cured in a free vacuum bag containing the entirety of the central and peripheral mandrel sections202,204,206,208,218. The rigidity of the central and peripheral mandrel sections202,204,206,208,218allows for the mandrel200to form a self-supporting structure able to withstand the pressure applied by the autoclave environment in the interior of the vacuum bag. By contrast, tooling not adapted to withstand autoclave pressures typically requires that the vacuum bag(s) be tailored to the periphery of the uncured composite material so as to minimize the portion of the tooling contained with the vacuum bag(s). The mandrel200may thus allow for a simplification of the bagging process.

In a particular embodiment, the thermally insulating material of the mandrel core230allows for heating the composite material without heating the entire mandrel200to the cure temperature, which in a particular embodiment facilitates application of the required cure cycle to the composite material. Since the mandrel200has a relatively large size (for example, corresponding substantially to that of a rotorcraft fuselage130), the length of time require to heat a similar size mandrel without thermally insulating material and accordingly acting as a heat sink may prevent the desired ramp-up in temperature of the desired cure cycle to be applied to the composite material.

In a particular embodiment, the matching of the coefficient of thermal expansion of the mandrel outer layer232and of the composite material allows to reduce, minimize, or avoid thermal stress and distortions of the composite material during the cure cycle.

After the composite material is cured, the central mandrel section202is slid out of engagement with the peripheral mandrel sections204,206,208and out of the fuselage130. For example, central mandrel section202may exit the fuselage130from the front window opening in the embodiment shown, while the central mandrel section202maintains its shape. Once the central mandrel section202is removed, the peripheral mandrel sections204,206,208and additional sections218are free to be disengaged from the cured material and removed from the fuselage130, again while maintaining their shapes. The slidable configuration of the mandrel sections202,204,206,208allows for the mandrel sections to be removable from the cured fuselage body without the need for a collapsing mandrel structure.

In a particular embodiment, the molded body of the fuselage130is self-supporting upon disengagement from the mandrel sections202,204,206,208,218. That is, the fuselage130does not require any additional support structure to maintain its shape once disengaged from the mandrel200. Accordingly, in a particular embodiment, this may allow for the tooling required to perform subsequent finishing operations to be less complex, which may reduce time and/or costs of such finishing operations as compared to a body requiring a support structure.

In a particular embodiment, the dismountable mandrel sections202,204,206,208,218allow for layup of the composite materials, and co-cure and/or co-bond of the skin and inner structural primary parts234. The mandrel200may reduce the number of operations, tools and labor to produce a monolithic fuselage130, as compared to the separate manufacture and subsequent assembly of fuselage sections.

In the embodiment shown, the mandrel200is exposed to the autoclave pressure only along its outer surface, which may reduce the risk of breaks in the vacuum bag(s) and/or leaks during the cure cycle. In an alternative embodiment, some or all of the mandrel sections may be open (i.e., hollow) so as to be exposed to the autoclave pressure on inside surfaces as well.

In a particular embodiment, the configuration of the mandrel200allows for the mandrel200to be scalable to be used to manufacture different sizes of fuselage without significant changes being required.

Although the present mandrel200and method have been discussed in relation to the fuselage130of a rotorcraft100, it is understood that, alternatively, the mandrel200can be configured to manufacture any other suitable type of monolithic composite fuselage. For instance, a modular mandrel, similar to that of the depicted embodiments, may be used to manufacture the tail section empennage150. The mandrel200can also be configured to manufacture any other suitable type of monolithic composite structures, including, but not limited to, in aerospace, boats, automobiles, etc.