ROTTER, Daniel M. (17054 26th Avenue NortheastLFP, WA, 98155, US)
WASHBURN, Todd J. (24005 242nd Way Southeast, Maple Valley, WA, 98038-0208, US)
GEORGE, Panagiotis E (17818 Driftwood Drive East, Lake Tapps, WA, 98391, US)
WILLDEN, Kurtis S (20233 SE 290th PL, Kent, WA, 98042, US)
HUNTER, Ethan W (844 Homestead Avenue, Haverton, PA, 19083, US)
MCMANUS, Nika (19-9 Valley Rd, Drexel Hill, PA, 19026, US)
ROBINS, Brian G. (16028 Southeast 148th Street, Renton, WA, 98059, US)
ROTTER, Daniel M. (17054 26th Avenue NortheastLFP, WA, 98155, US)
WASHBURN, Todd J. (24005 242nd Way Southeast, Maple Valley, WA, 98038-0208, US)
GEORGE, Panagiotis E (17818 Driftwood Drive East, Lake Tapps, WA, 98391, US)
WILLDEN, Kurtis S (20233 SE 290th PL, Kent, WA, 98042, US)
HUNTER, Ethan W (844 Homestead Avenue, Haverton, PA, 19083, US)
MCMANUS, Nika (19-9 Valley Rd, Drexel Hill, PA, 19026, US)
| CLAIMS What is claimed is: 1. A composite mandrel, comprising: a generally elongated mandrel body comprising: a resilient mandrel core; and an elastomeric mandrel outer layer disposed outside said mandrel core. 2. The composite mandrel of claim 1 wherein said mandrel core comprises foam. 3. The composite mandrel of claim 1 wherein said mandrel outer layer comprises an elastomeric material. 4. The composite mandrel of claim 1 wherein said mandrel body has a generally triangular cross-section. 5. The composite mandrel of claim 4 wherein said mandrel core comprises a core base, a pair of core sides extending from said core base and a core apex extending between said pair of core sides. 6. The composite mandrel of claim 5 wherein said core apex of said mandrel core is rounded. 7. The composite mandrel of claim 5 wherein said mandrel outer layer comprises a mandrel base disposed adjacent to said core base of said mandrel core; a pair of mandrel sides disposed adjacent to said core sides, respectively, of said mandrel core; and a mandrel apex disposed adjacent to said core apex of said mandrel core. 8. The composite mandrel of claim 7 wherein said mandrel apex of said mandrel outer layer is rounded. 9. A method for fabricating a contoured stiffened composite panel for an aircraft structure, comprising: providing a tooling surface; placing a base composite layer on said tooling surface; placing at least one stiffening element having a stiffening element cavity on said base composite layer; providing at least one composite mandrel having a resilient mandrel core and an elastomeric mandrel outer layer disposed outside said resilient mandrel core; inserting said at least one composite mandrel in said stiffening element cavity of said at least one stiffening element; curing said base composite layer and said at least one stiffening element; and removing said at least one composite mandrel from said stiffening element cavity of said at least one stiffening element. 10. The method of claim 9 wherein said providing a tooling surface comprises providing a tooling surface having a generally concave contour. 11. The method of claim 9 wherein said curing said base composite layer and said at least one stiffening element comprises providing vacuum bagging, placing said base composite layer and said at least one stiffening element in said vacuum bagging and subjecting said vacuum bagging to autoclave conditions. 12. The method of claim 9 wherein said providing at least one composite mandrel comprises providing at least one composite mandrel having a generally triangular cross-section. 13. The method of claim 9 wherein said providing at least one composite mandrel comprises providing at least one composite mandrel having a generally trapezoidal cross-section. 14. The method of claim 9 wherein said providing at least one composite mandrel having a resilient mandrel core comprises providing at least one composite mandrel having a foam mandrel core. 15. The method of claim 9 wherein said providing at least one composite mandrel having an elastomeric mandrel outer layer comprises providing at least one composite mandrel having an elastic rubber mandrel outer layer. |
TECHNICAL FIELD The disclosure relates to mandrels for forming cavities in composite materials. More particularly, the disclosure relates to a composite mandrel which is suitable for autoclave curing applications in the formation of cavities in composite materials.
This disclosure also generally relates to mandrels used to form cavities in composite structures, and deals more particularly with a mandrel that is suitable for autoclave curing applications.
BACKGROUND
When composite materials are molded into shapes with cavities, such as hat stringers, for example, there may be a need for some type of tooling that can apply pressure from the cavity outward during the curing step and can be extracted from the cavity after curing. The existing tooling used for this purpose may include without limitation inflatable rubber mandrels; solid mandrels such as metal, rubber or composite mandrels; or dissolvable mandrels. However, the inflatable rubber mandrels may be prone to leaking, which may lead to widespread porosity in the resulting composite laminate. The solid rubber mandrel may result in a cavity with a distorted cross-sectional shape or exert an uneven pressure on the composite laminate and may be too heavy for fabrication of large parts. The solid metal or composite mandrels may not have sufficient flexibility to be removed from parts having any degree of curvature or complexity. The dissolvable mandrels may be expensive to make and difficult to remove from large parts. Existing mandrel designs may not accommodate the dimensional changes of the composite part which occurs during application of heat to the surrounding tooling and part materials at the curing step. This can cause undesirable part material movement resulting in such distortions as waviness, wrinkling and/or bridging in the composite material.
Therefore, a mandrel is needed which is suitable for curing applications in the formation of cavities in composite materials and overcomes some or all of the limitations of conventional composite mandrels.
When composite materials are molded into structures having cavities, such as hat stringers, for example and without limitation, there may be a need for tooling that applies outward pressure from within the cavity during curing, and which can be removed from the cavity after curing. Existing tooling used for this purpose may include, without limitation, inflatable rubber mandrels; solid mandrels such as metal, rubber or composite mandrels; or dissolvable mandrels. Solid rubber or composite mandrel are sometimes preferred for certain applications, however when subjected to heat and pressure, this type of mandrel may produce a cavity having a distorted cross-sectional shape and/or may exert uneven pressure on the composite laminate. Accordingly, there is a need for a removable mandrel that reduces or eliminates distortion of the cavity shape, and exerts substantially uniform pressure over the composite laminate during curing at elevated temperatures.
SUMMARY The disclosure is generally directed to a composite mandrel. An illustrative embodiment of the composite mandrel includes a generally elongated mandrel body comprising a resilient mandrel core and an elastomeric mandrel outer layer disposed outside the mandrel core. The mandrel may combine the desired characteristics of foam and rubber to produce a manufacturing aid for airplane stringers or other similar open cavity parts made from fiber/resin composite materials. The manufacturing aid which is embodied in the composite mandrel may be less costly, more durable and less prone to failures than current inflatable bladder technologies.
The disclosed embodiments provide a removable mandrel used to cure a composite structure layup which may reduce or eliminate distortion of the structure caused by uneven application of pressure by the mandrel during curing. The mandrel includes one or more internal open spaces therein designed to allow the mandrel to thermally expand inwardly, while the exterior of the mandrel expands substantially uniformly to exert an even pressure over the layup. The even application of pressure to the layup may reduce or eliminate distortion of the cured structure. According to one disclosed embodiment, a mandrel comprises an elastic body. At least one internal open space passes through the body, which is configured to allow substantially uniform thermal expansion of the body. The elastic body may comprise an elastomer, and the internal open space may be symmetric about the longitudinal axis of the body.
According to another embodiment, a mandrel is provided for curing a composite part layup comprising a generally solid body that expands when heated to apply pressure to the composite part layup. The body includes at least one internal open space therein positioned and configured to promote substantially uniform support for the part during layup and cure. The internal open space in the body may be elongate, and in one application may possess a generally daisy shaped cross section. In another variation, a plurality of the internal open spaces pass longitudinally through the body and are collectively configured to allow substantially uniform thermal expansion of the body.
A method embodiment provides for fabrication of a composite structure. A mandrel is formed and the thermal expansion characteristics of the mandrel are determined. The shape of an internal open space in the mandrel is selected which will result in substantially uniform thermal expansion of the mandrel during curing of the composite structure layup. An internal open space in the mandrel having the desired shape is formed, and a composite structure layup is placed over the mandrel, following which the layup is cured. Selecting the shape of the internal open space in the mandrel may be performed using finite element analysis. In one variation, a plurality of internal open spaces may be formed in the mandrel, which are collectively configured to allow substantially uniform thermal expansion of the mandrel body.
1. A mandrel, comprising: an elastic body; and at least one internal open space passing through the body and configured to allow substantially uniform thermal expansion of the body.
2. The mandrel of claim 1 wherein the elastic body.
3. The mandrel of claim 1 wherein the internal open space passes longitudinally through the body.
4. The mandrel of claim 1, further comprising a plurality of internal open spaces passing through the elastic body and collectively configured to allow substantially uniform thermal expansion of the body.
5. The mandrel of claim 1 wherein the elastic body is an elastomer.
6. The mandrel of claim 1 wherein the elastic body is elongate and the open space passes longitudinally through the body. 7. The mandrel of claim 6 wherein the open space is substantially symmetric about the longitudinal axis of the elastic body.
8. The mandrel of claim 1 wherein said elastic body has a generally trapezoidal shape.
9. A mandrel for curing a composite part layup, comprising: a body formed of a material that thermally expands when heated to apply pressure to the composite part layup, the body including at least one internal open space therein shaped to promote substantially uniform support for the part during layup and cure.
10. The mandrel of claim 9, wherein: the body is elongate, and the internal open space passes longitudinally through the body.
11. The mandrel of claim 10 wherein the internal open space is disposed substantially symmetrically about the longitudinal axis of the body.
12. The mandrel of claim 9 wherein the body is formed of an elastomeric material. 13. The mandrel of claim 9 wherein the internal open space has a generally daisy-shaped cross section.
14. The mandrel of claim 9, further comprising a plurality of internal open spaces passing through the body and collectively configured to allow substantially uniform thermal expansion of the body. 15. The mandrel of claim 14, wherein: the body is elongate and has a longitudinal axis, and the plurality of internal open spaces are arranged generally symmetrically about the longitudinal axis of the body.
16. A method of fabricating a composite part, comprising: forming a mandrel; determining the thermal expansion characteristics of the mandrel; forming at least one internal open space in the mandrel that will result in substantially uniform thermal expansion of the mandrel during curing of the composite part; placing a composite part layup over the mandrel; and curing the composite part. 17. The method of claim 16, further comprising: selecting the shape of the opening using finite element analysis.
18. The method of claim 16, wherein forming the at least one internal open space and forming the mandrel are performed substantially simultaneously. 19. The method of claim 16, wherein forming the at least one internal open space includes forming an opening passing longitudinally through the mandrel. 20. The method of claim 16, further comprising: forming a plurality of internal open spaces in the mandrel collectively configured to allow substantially uniform thermal expansion of the body. 21. The method of claim 16, wherein forming the at least one internal open space includes selecting the cross sectional shape and the position of the open space within the mandrel.
22. The method of claim 16, wherein curing the composite part layup is performed in an autoclave.
23. A composite part cured by the method pf claim 16. 24. A method of fabricating a composite part, comprising: selecting an elastic mandrel material; determining the thermal expansion characteristics of the selected elastic mandrel material; selecting the shape of the mandrel; forming the selected mandrel material into the selected mandrel shape; performing a finite element analysis to determine the size, number and cross sectional shape of internal open spaces within in the mandrel that will result in substantially uniform thermal expansion of the mandrel; forming internal open spaces within the mandrel based on the results of the finite element analysis; placing a composite part layup over the mandrel; and curing the composite part in an autoclave using the mandrel. 25. A mandrel for curing a composite part in an autoclave, comprising: an elongate, generally solid mandrel body formed of an elastomeric material; and, a plurality of internal open spaces within the mandrel body and passing longitudinally through the mandrel body, the open spaces being arranged substantially symmetrically around the longitudinal axis of the mandrel body and each having a substantially daisy shaped cross section.
9. A composite mandrel for fabricating an aircraft part, comprising: a generally elongated mandrel body having a generally trapezoidal cross-section and comprising: a resilient mandrel core; and an elastomeric mandrel outer layer disposed outside said mandrel core.
10. The composite mandrel of claim 9 wherein said mandrel core comprises foam. 11. The composite mandrel of claim 9 wherein said mandrel outer layer comprises an elastomeric material.
12. The composite mandrel of claim 9 wherein said mandrel core comprises a core base, a pair of core sides extending from said core base and a generally planar core top extending between said pair of core sides. 13. The composite mandrel of claim 12 wherein said mandrel outer layer comprises a mandrel base disposed adjacent to said core base of said mandrel core; a pair of mandrel sides disposed adjacent to said core sides, respectively, of said mandrel core; and a mandrel top surface disposed adjacent to said core top of said mandrel core.
21. A composite mandrel for fabricating an aircraft part, comprising: a generally elongated mandrel body having a generally trapezoidal cross-section and comprising: a resilient foam mandrel core having a core base, a pair of core sides extending from said core base and a generally planar core top extending between said pair of core sides; and an elastic rubber mandrel outer layer disposed outside said mandrel core and having a mandrel base disposed adjacent to said core base of said mandrel core; a pair of mandrel sides disposed adjacent to said core sides, respectively, of said mandrel core; and a mandrel top surface disposed adjacent to said core top of said mandrel core. 22. A method for fabricating a contoured stiffened composite panel for an aircraft structure, comprising: providing a tooling surface having a generally concave contour; placing a base composite layer on said tooling surface; placing at least one stiffening element having a stiffening element cavity on said base composite layer; providing at least one composite mandrel having a generally triangular or trapezoidal cross-section and including a resilient foam mandrel core and an elastic rubber mandrel outer layer disposed outside said resilient mandrel core; inserting said at least one composite mandrel in said stiffening element cavity of said at least one stiffening element; curing said base composite layer and said at least one stiffening element by providing vacuum bagging, placing said base composite layer and said at least one stiffening element in said vacuum bagging and subjecting said vacuum bagging to autoclave conditions; and removing said at least one composite mandrel from said stiffening element cavity of said at least one stiffening element. BRIEF DESCRIPTION OF THE ILLUSTRATIONS FIG. 1 is a top view of an illustrative embodiment of a mandrel.
FIG. 2 is a cross-sectional view, taken along section lines 2-2 in FIG. 1, of the mandrel.
FIG. 3 is a cross-sectional view of an alternative illustrative embodiment of the mandrel.
FIG. 4 is an exploded top view of a composite assembly, more particularly illustrating insertion of multiple mandrels into respective stiffening elements in the composite assembly preparatory to curing of the composite assembly.
FIG. 5 is a cross-sectional view, taken along section lines 5-5 in FIG. 4, of the composite assembly.
FIG. 6 is a top view of the composite assembly, with the mandrels inserted in the respective stiffening elements of the assembly.
FIG. 7 is a top view of the composite assembly, contained in vacuum bagging preparatory to curing of the assembly.
FIG. 8 is an exploded top view of the composite assembly, more particularly illustrating removal of the mandrels from the respective stiffening elements in the composite assembly after curing of the composite assembly.
FIG. 9 is a flow diagram illustrating a method for fabricating a contoured stiffened composite panel. FIG. 10 is a cross sectional view of an alternate form of the mandrel having an internal open space for controlling thermal expansion, shown installed within a composite layup on a tool.
FIG. 11 is an enlarged view of a portion of the mandrel shown in FIG. 10, illustrating the substantially uniform thermal expansion of the mandrel during curing.
FIG. 12 is a top view of the mandrel shown in FIGS, 10 and 11;
FIG. 13 is a cross sectional view similar to FIG. 10, but showing another form of the mandrel having multiple internal open spaces for controlling thermal expansion.
FIG. 14 is a flow diagram illustrating a method of producing and using the mandrel shown in FIGS. 10-13.
FIG. 15 is a flow diagram of an aircraft production and service methodology.
FIG. 16 is a block diagram of an aircraft.
DETAILED DESCRIPTION
Referring initially to FIGS. 1 and 2, an illustrative embodiment of the mandrel is generally indicated by reference numeral 1. The mandrel 1 may be used to fill a cavity (not shown) in an airplane stringer or other open-cavity part (not shown) made from fiber/resin composite materials to prevent collapse of the cavity during curing of the composite materials. The mandrel 1 may be less costly, more durable and more effective and reliable than current inflatable bladder mandrel technologies. The mandrel 1 includes a generally elongated mandrel body 7 having a mandrel core 2 which is a resilient material and a mandrel outer layer 10 which is disposed outside the mandrel core 2, as shown in FIG. 2, and is an elastomeric material. In some embodiments, the mandrel core 2 is foam or other such material which incorporates open space and/or air pockets to prevent bulk modulus behavior during thermal expansion and the mandrel outer layer 10 may be an elastomeric material such as elastic rubber, for example and without limitation. The mandrel core 2 and the mandrel outer layer 10 may be generally coextensive with the mandrel body 7.
The mandrel core 2 and the mandrel outer layer 10 may have any cross-sectional shape depending on the particular use requirements of the mandrel 1. In some applications, for example, each of multiple mandrels 1 may be suitably configured to fill respective stiffening elements (such as stringers) 27 during the curing and/or cocuring of a composite panel assembly 24, as shown in FIGS. 4-8 and will be hereinafter described. As shown in FIG. 2, in some embodiments of the mandrel 1, the mandrel body 7 may have a generally triangular cross- sectional shape. Accordingly, the mandrel core 2 has a generally flat or planar core base 3 with lateral core edges 6. Core sides 4 angle from the respective core edges 6. A core apex 5, which may be rounded, extends between the core sides 4. The shape of the mandrel outer layer 10 may generally correspond to that of the mandrel core 2, defining a mandrel base 11 which extends adjacent to the core base 3; a pair of mandrel sides 12 which extend adjacent to the respective core sides 4; a mandrel apex 13 which may be rounded and is disposed adjacent to the core apex 5; and mandrel edges 14 which correspond positionally to the respective core edges 6 of the mandrel core 2.
As shown in FIG. 3, in some embodiments of the mandrel Ia, the mandrel body 7a may have a generally trapezoidal shape. Accordingly, the mandrel core 2a has a generally flat or planar core base 3; a pair of core sides 4 which angle from the core base 3; and a generally flat or planar mandrel core top 8 which extends between the core sides 4. The mandrel outer layer 10a defines a mandrel base 11 which extends adjacent to the core base 3; a pair of mandrel sides 12 which extend adjacent to the respective core sides 4; a generally flat or planar mandrel top surface 16 which is disposed adjacent to the mandrel core top 8; and mandrel edges 14 which correspond to the respective core edges 6 of the mandrel core 2a. Referring next to FIGS. 4-8, in typical application, multiple mandrels 1 are inserted in respective stiffening elements 27 provided in a stiffening layer 26 of a composite panel assembly 24 during curing of the composite panel assembly 24. The composite panel assembly 24 will ultimately form an airplane stringer (not shown); however, it will be appreciated by those skilled in the art that the mandrels 1 can be adapted to fill cavities in any other type of open-cavity or closed-cavity composite material part made from fiber/resin composite materials during curing of the composite material part. The mandrels 1 can be adapted to fill cavities having a constant cross-sectional shape or a cross-sectional shape which varies along the length of the composite material, such as cavities which taper or curve along the length of the cavity, for example and without limitation. As illustrated in FIG. 5, in an embodiment of fabrication of the composite panel assembly
24, a base composite layer 25 may initially be placed on a tooling surface 20 of OML tooling or IML tooling, for example and without limitation. The tooling surface 20 may have a generally concave contour, as shown. Alternatively, the tooling surface 20 may have a generally planar or convex contour, depending on the particular application. The stiffening layer 26 may be placed on the base composite layer 25. The stiffening elements 27 may be shaped in the stiffening layer 26 and extend along the longitudinal axis of the tooling surface 20 in generally parallel relationship with respect to each other, as shown in FIG. 4, and in generally perpendicular relationship with respect to the concave contour of the tooling surface 20. Alternatively, the stiffening elements 27 may be separate or discrete units. As further shown in FIG. 5, each stiffening element 27 has a stiffening element cavity 28. In some embodiments, the stiffening elements 27 may be oriented in orientations other than along the longitudinal axis of the tooling surface 20 and may converge or diverge, for example and without limitation.
As shown in FIGS. 4 and 6, multiple mandrels 1 may be inserted into the stiffening element cavities 28 of the respective stiffening elements 27. The elastomeric mandrel outer layer 10 of each mandrel 1 allows for a proper fit of the mandrel 1 into the stiffening element cavity 28 of each stiffening element 27 and conforms to pad-ups and ramps. As shown in FIG. 7, the composite panel assembly 24 may then be enclosed in vacuum bagging 30 and cured by autoclaving. During the curing process, the mandrels 1 maintain the shape and prevent collapse of the respective stiffening elements 27 as the composite material of the base composite layer 25 and the stiffening layer 26 hardens.
After curing, the composite panel assembly 24 is removed from the vacuum bagging 30. The mandrels 1 may be removed from the stiffening element cavities 28 of the respective stiffening elements 27, as shown in FIG. 8. During removal, the elastomeric mandrel outer layer 10 of each mandrel 1 may easily be deformed; this reduces the effort required for removal. The cured composite panel assembly 24 may then be processed to complete fabrication of the airplane assembly (not shown) or other composite part, according to the knowledge of those skilled in the art.
It will be appreciated by those skilled in the art that the resilient mandrel core 2 of the mandrel 1 enhances the structural and compressive characteristics of the mandrel 1 relative to the designs of conventional mandrels. This structural and compressive support may be necessary to maintain the shape of the stringer or other composite part during automated composite fiber placement as well as autoclave curing. Since the outer mandrel layer 10 may be a constant thickness, it may expand uniformly during curing, thus avoiding the problems associated with uneven expansion of a solid rubber material. The cross-sectional area and type of foam used for the mandrel core 2 may be engineered to impart compression compliance under autoclave pressure, thus offsetting the combined thermal expansion behavior of the foam and rubber.
Referring next to FIG. 9 of the drawings, a flow diagram 32 which illustrates an illustrative method for fabricating a contoured stiffened composite panel is shown. At 34, a tooling surface, such as the tooling surface 20 which was heretofore described with respect to FIG. 5, for example and without limitation, is provided. The tooling surface may have a concave, planar, convex or alternative contour. At 36, a base composite layer is laminated on the tooling surface. In step 38, open-section stiffening elements are positioned on the base composite layer. At step 40, mandrels are provided. Each mandrel includes a resilient mandrel core and an elastomeric mandrel outer layer disposed outside the resilient mandrel core. At step 42, mandrels are inserted in the respective stiffening elements. At 44, the composite panel and stiffening elements are sealed in vacuum bagging. In step 46, the composite panel and the stiffening elements are cured. An autoclave may be used during curing. Finally at step 48, the mandrels are removed from the stiffening elements. Attention is now directed to FIGS. 10-12 which illustrate another embodiment of the mandrel 50 used to apply pressure to a composite structure 49 layup, such as for example and without limitation, a hat-shaped stiffer 52 joined to a composite skin 54. In this embodiment, the mandrel 50 is used to form hat stiffeners 52 that are made with cure tooling forming the surface of the hat and the skin side subject to autoclave pressure, in contrast to the previously described embodiments in which the cure tooling is used against the skin and the hat surface is subject to the autoclave pressure. The mandrel 50 defines and maintains the shape of the stiffener 52 during autoclave curing, with substantially even pressure and without substantial distortion. The composite structure layup 49 may be supported on a tool 60 within an autoclave (not shown) used to cocure the stiffener 52 and skin 54. The mandrel 50 includes a generally solid body 50a that may extend the length of the stiffener 52. As used herein, "generally solid body" refers to a body that is substantially solid but may contain one or more open spaces 62 therein, as will be discussed below in more detail. The body 50a may be formed of an elastic material such as, without limitation, an elastomer or a rubber that is relatively soft, resilient and possesses a relatively low CTE (coefficient of thermal expansion) which may be less than the CTE of the composite structure layup 49. The resilient, elastic nature of the mandrel 50 allows it to conform to slight variations in the shape of the layup 49, while permitting the mandrel 50 to flex or bend slightly as it is being removed from the layup
49 after curing. In the illustrated embodiment, the mandrel body 50a has a substantially trapezoidal cross sectional shape, generally matching that of the stiffener 52. However the mandrel body 50a may possess any of a variety of cross sectional shapes depending on the application and the particular shape of the composite structure layup 49. Although not shown in the drawings, one or more of the open spaces 62 may be filled with material such as a foam having a low CTE. The mandrel body 50a includes at least one internal open space 62 which, in the illustrated embodiment, passes longitudinally through the body 50a, substantially parallel to the longitudinal axis 64 (FIG. 12) of the body 50a. The location and shape of the internal open space 62 is configured to control the expansion of the body 50a, in a manner that results in the mandrel
50 exerting substantially uniform cure pressure against the stiffener 52 and the skin 54. The open space 62 has a substantially daisy-shaped cross section which, in the illustrated example, is generally symmetrically disposed around the longitudinal axis 64. The longitudinal axis 64 passes the centroid of the cross section of the body 50b in the illustrated embodiment. In other embodiments, the cross sectional shape of the open space 62 may not be symmetric about the longitudinal axis 64. Generally, however, the cross sectional shape of the open space 62 will be symmetric about the mid-plane of the stiffener 52, defined as a plane that is perpendicular to the skin 54. The daisy-shaped open space 62 comprises a plurality of circumferentially spaced lobes 66, and is merely illustrative of a wide variety of shapes that may be possible for achieving substantially even thermal expansion of the outer surfaces 50b of the mandrel body 50a.
As previously noted, the particular shape chosen for the open space 62 will depend upon the geometry of the composite structure layup 49, including the shape of the cavity filled by the mandrel 50. The shape and placement of the open space 62 may be selected and optimized using finite element analysis which may indicate the amount of expansion of the mandrel body 50a at various temperatures during cure, and the corresponding pressures applied to the composite structure layup 49. The finite element analysis may be carried out using any of several commercially available software packages. Finite element analysis may be used as a predictive numerical tool to model and analyze the laminate cure process of the composite structure and to optimize the internal open space 62 in the mandrel 50 to specified tolerances.
During the thermal cycling used to effect cure of the composite structure layup 49, substantially uniform expansion of the outer surfaces 50b of the mandrel body 50a is achieved as a result of the mandrel body 50a expanding into the open space 62. The essentially concurrent outward expansion of the mandrel surfaces 50b and the inward expansion of the mandrel body 50a into the open space 62 is indicated by the arrows 68.
Referring now to FIG. 13, in contrast to the single internal open space 62 shown in FIGS. 10 and 11, it may be useful to employ a plurality of longitudinal internal open spaces 62a within the mandrel body 50a to control thermal expansion of the mandrel body 50a. In the illustrated example, three longitudinal internal open spaces 62a are arranged in a generally triangular pattern, roughly evenly spaced from the longitudinal axis 64 of the mandrel 50. Each of the longitudinal open spaces 62a possesses a daisy-shaped cross section, however as previously indicated, a variety of other shapes may be possible. Attention is now directed to FIG. 14 which broadly illustrates the steps of a method of fabricating a composite structure using the mandrel 50 shown in FIGS. 10-13. Beginning at step 70, the mandrel 50 is formed by any of a variety of fabrication techniques, such as molding an elastomeric material. Next at 72, the expansion characteristics of the mandrel 50 are determined based on its geometry, dimensions and material. At 74, the number, shape and position of the internal open spaces 62 are selected which may be accomplished, at least in part, by performing a finite element analysis, as indicated at step 75. At 76, the selected open spaces 62 may be formed in the mandrel 50 by any of a variety of the fabrication techniques. The openings 62 may be formed in the mandrel 50 at the same time the mandrel body 50a is formed, in step 70, as by molding. At 78, a composite structure layup 49 is placed over the mandrel 50, or alternatively, the mandrel 50 is placed inside an existing layup 49. The assembly of the mandrel 50 and layup 49 is then vacuum bagged at 79 in order to form and consolidate of the layup 49. The vacuum bagged assembly of the layup 49 and the mandrel 50 is placed in an autoclave (not shown) where it is cured at step 80. Following curing, the mandrel 50 is removed from the cured layup 49 at 81. The elasticity of the mandrel 50 allows it to deform slightly to facilitate its removal, as it is either pushed or pulled from the layup 49 during the removal.
Referring next to FIGS. 15 and 16, embodiments of the disclosure may be used in the context of an aircraft manufacturing and service method 82 as shown in FIG. 15 and an aircraft 84 as shown in FIG. 16. During pre-production, exemplary method 82 may include specification and design 86 of the aircraft 84 and material procurement 88. During production, component and subassembly manufacturing 90 and system integration 92 of the aircraft 84 takes place. Thereafter, the aircraft 44 may go through certification and delivery 94 in order to be placed in service 96. While in service by a customer, the aircraft 84 may be scheduled for routine maintenance and service 98 (which may also include modification, reconfiguration, refurbishment, and so on). Each of the processes of method 82 may 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 description, a system integrator may include without limitation any number of aircraft manufacturers and major-system subcontractors; a third party may include without limitation any number of vendors, subcontractors, and suppliers; and an operator may be an airline, leasing company, military entity, service organization, and so on.
As shown in FIG. 16, the aircraft 84 produced by exemplary method 82 may include an airframe 100 with a plurality of systems 102 and an interior 104. Examples of high-level systems 102 include one or more of a propulsion system 106, an electrical system 108, a hydraulic system 110, and an environmental system 112. Any number of other systems may be included. Although an aerospace example is shown, the principles of the invention may be applied to other industries, such as the automotive industry.
The apparatus embodied herein may be employed during any one or more of the stages of the production and service method 82. For example, components or subassemblies corresponding to production process 90 may be fabricated or manufactured in a manner similar to components or subassemblies produced while the aircraft 84 is in service. Also, one or more apparatus embodiments may be utilized during the production stages 90 and 92, for example, by substantially expediting assembly of or reducing the cost of an aircraft 84. Similarly, one or more apparatus embodiments may be utilized while the aircraft 84 is in service, for example and without limitation, to maintenance and service 98.
Although the embodiments of this disclosure have been described with respect to certain exemplary embodiments, it is to be understood that the specific embodiments are for purposes of illustration and not limitation, as other variations will occur to those of skill in the art.
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