WO2014168701A1

WO2014168701A1 - Multi-component composite structures

Info

Publication number: WO2014168701A1
Application number: PCT/US2014/021594
Authority: WO
Inventors: Bruno Boursier
Original assignee: Hexcel Corp
Current assignee: Hexcel Corp
Priority date: 2013-04-12
Filing date: 2014-03-07
Publication date: 2014-10-16
Anticipated expiration: 2015-10-12
Also published as: CA2907937A1; EP2983899A1; ES2875791T3; KR20150140379A; CN105189093B; US10239289B2; CN105189093A; JP6337090B2; US20150375481A1; JP2016522101A; EP2983899B1

Abstract

A composite assembly that can be cared to form a multi-component composite structure which does not have micro cracks along the boundaries between the various components. The composite assembly includes a structural component and a moldable component wherein the coefficients of thermal expansion of the structural component and the moldable component at the interface or boundary between the two components are such that micro cracks do not form along the interface when the composite assembly is cured to form the multi -component composite structure.

Description

BACKGROUND OF THE. IN VENTIO

1. Field of the Invention

|§Θ01] The present invention relates generally to iiJiilti-c-oiupoaeiit or hybrid composite structures that are made by molding imcuxed composite assemblies which arc composed of a structural component that is embedded within a moldable component The combination, of a structural component with a moldable component allows one to take advantage of the added strength provided by the siruc ral component while still being able to form composite structures that have relatively complex shapes. More particularl _f the present invention is directed to eliminating the micro cracks that tend to form along the interfaces between the structural component and the moldable component during molding of he imcured composite ass*

2, Description of Related Art.

|08O2J Composite materials typically include fibers and a resin matrix as the two principal components.. Composite materials typically have a rather high strength to weight ratio. As a result, composite materials are being used in. demanding mvironments, such as in the field of aerospace where the high strength and relatively lightweight of composite parts are of particular importance.

100031 A discontinuous fiber composite (DFC) material has been developed that can be accurately molded and machined to form a wide variety of relatively complex, structures- This composite material is composed of randomly oriented segments of unidirectional tape that have been impregnated with thermosetting resin. This type of quasi-isotropic fiber material has been used to make molds and a variety of aerospace components. The material is available from Hexcel Corporation (Dublin, CA) under the trade name HexMC®. Examples of the types of parts that have been made using HexMC® are described in US Patent Nos. 7,510.390; 7,960,674 and published US Patent Application US2032-0040169-A 1, the contents of which are hereby incorporated by reference. 108 41 The fibers used in many load-bearing composite structures or elements are unidirectional and continuous. Such tuiidireetioiial (UD) fibers are particularly useful when the load-bearing structure is relatively long with respect to the width and thickness of the structure. Wing spars, struts_., links, frames, mieroosials, beams, -skins, panels, jet engine blades and vanes are examples of various aircraft structures that can be relatively long and whieh are designed to cany significant loads,

10 )051 UD fibers are generally provided as a tape or layer of parallel continuous fiber that may or may not be impregnated with thermosetting resin. The tape or layer of UD tibers has a width and a thickness with the fibers extending unidireetionally in the length direction. The UD fiber layer can generally be shaped into curved structures provided that the tape is bent in the thickness direction. It is much more difficult to form curved structures in which the UD fiber layer is bent in the direction of the width of the UD layer. Procedures have been developed to allow a UD fiber layer to be bent in the width direction. Such, procedures involve twisting the UD fibers prior to bending the UD layer in the width direction. Such procedures are described in published US Patent Applications US2010-0173143- A I and US2010-0173152- A I, the contents of which, is hereby incorporated by reference. These bending procedures allow one to form UD fiber layers into strong structural parts that have some curvature in the ^"thickness and/or width directions. However, it remains difficult to form complex machinable structures using only UD fiber layers.

10 061 DFC material is entirely suitable for use in those situations where the desired composite structure has a relatively complex shape and/or requires post-curing machining. However, there are many situations where it is desirable to reinforce one or more sections of the DFC structure with continuous UD fibers. Such niulti -component or hybrid composite structures are composed of DFC material, as the moldable component, and continuous UD fibers as the structural component. The UD fibers are embedded within the DFC material to provide structural reinforcement in those areas of the structure that require the extra strength which is provided by continuous UD fibers.

100071 DFC/UD hybrid composite structures are generally made by first forming an uncured composite assembly that includes continuous UD fibers as the structural, component of the assembly and DFC material as the moldable component. This assembly is cured in a mold under high pressure at an elevated temperature to produce a multi-component composite structure. The structural component can be made up of one or more UD structural elements that are placed strategically within the structure to provide the desired degree of reinforcement for the rnoldabte component.

O08^'| DFC materia] and continuous UD fiber layers tend to expand at different rates as the materials are heated and cured during the molding process. The rate at which these materials expand during molding is expressed as the coefficient of thermal expansion (CTE). The micro cracking that may occur along the interfaces or boundaries between the various components is a ^•major concern when molding hybrid composite assemblies to form multi-component composite structures. Micro cracking becomes more of an issue as the difference in CTE between the various components increases. The difference in CTE between DFC material and UD fiber layers is sufficiently large that .micro cracking can become a problem when these two components are combined for molding into nndti -component composite structures.

1 0 1 It would be desirable to provide methods for making multi-component structures from DFC materials and UD fibers where micro cracking along the interfaces ^'betwee the two materials is avoided during high temperature molding. Elimination of micro cracking is especially an issue in those situations where multiple UD structural elements are combined with DFC material and molded to form the hybrid structure.

SUMMARY OF THE INVENTION eOSiOj In accordance with the present invention. It was discovered that uncured composite assemblies, which contain a moldable component that i reinforced with a structural component, can he molded to form rmilti-coroponent structures that do not have micro cracks along the .interfaces between the two components. The invention is based in part, on the discovery that the size, shape and CTE of the structural, component can be controlled such, that micro crack formation during the molding process does not occur along the interlaces between the components.

gOQOll] The present invention is directed to composite assemblies that can be cured to form inulti-coinponent composite structures. The composite assembly includes a structural component which is composed of one or more structural elements. Each structural element includes uncured thermosetting resin and unidirectional continuous fibers. The composite assembly also includes a moldable component which is composed of a moldable body that includes an uncured thermosetting resin and discontinuous fibers. As a feature of the present invention, the coefficient of thermal expansion of ttie structural component and the moldable component, at fee interface between the two components, is such feat micro cracks do not form along the interface when, said composite assembly is c¾red to form, the m«Iti~componsat composite structure.

[00012] As another feature of me present invention, the CTE of the structural component is made to more closely match, the CTE of the moldable component by including multidirectional continuous fibers ia. the structural component Multi-directional continuous fibers tend to have a CTE tha more closely matches the CTE of the moldable component. The multidirectional, continuous fibers caa be distributed throughout the structural component to provide a structural component wife a relatively uniform CTE. The multi-directional continuous libers may also ^'be concentrated aear the interface with the moldable component to act as a CTE "buffer' between fee UD fibers and fee DFC molding material

[00013] The .invention is particularly useful for making multi-component structures where the structural component, is composed of multiple structural elements embedded within fee moldable component. The use of multiple structural elements increases the complexity and ^•number of the interfaces between the structural, component and moldable component The present invention is designed to avoid the formation of micro cracks that typically form in such complex, multi-interface hybrid structures.

000.1 The above described and many other features and attendant advantages of fee present .invention will become better understood by reference to the following detailed description when taken in conjunction with the accompanying drawings,

BRIEF DESCRIPTIO OF THE DRAWINGS

[000151 FIG. 1 is a perspective view of an exemplary uiicured composite assembly after it has been placed into a mold (not shown) and formed into its final shape prior to being cured under elevated temperature and high pressure to form an aircraft window frame which is a multi- component composite structure in accordance with the present invention,

[00016] FIG. 2 is a perspective view of an exemplary composite assembly after it has been placed into a mold (not shown) and formed into its final shape prior to being cured under elevated temperature and high pressure to form m aircraft access opening cover which is a miilti-eoniponent composite structure in accordance with the esent invention. The uncured access opening cover is shown with attachment holes which are typically not present in the uaeured composite assembly. The attachment holes are typically -machined into the access opening after molding is completed.

[0Θ 1.7] FIG. 3 is a sectional view representation of the exemplary composite assembly shown in FIG, L

[00018] PIG. 4 is a sectional view representation of an exemplary structural, element in. accordance with the present invention where layers of continuous UD fibers are alternated with Layers of multi-directional continuous fibers to provide a structural, component that contains ^'UD fibers_., but which also has a CTE that more closely matches the DFC material that makes up the moidab!c component,

01 019] FIG. 5 is a sectional view representation of an alternate exemplary composite assembly that is located within a mold (no shown) prior to being cured to form an alternate aircraft window frame.

[00020] FIG. 6 is a sectional view representation of the exemplary composite assembly shown in FIG, 2.

[00021] FIG. 7 is a sectional view of a preform prior to the preform being placed in a mold and formed into its final shape for molding to form the exemplary composite assembly shown in

FIGS. 1 and 3.

[08022] F I G, 8 is a sectional view o a preform prior to the preform being placed in a mold and formed into its final shape for molding to form the exemplary composite assembly shown in PIG. 5.

DETAILED DESCRIPTION OF THE INVENTION

[00023] The present invention -may he used in a wide variety of situations where it is desirable to combine unidirectional fibers with a raoldable composite materia to form multi- component or hybrid composite structures. Such hybrid composite structures are useful in situations where the combination of strength provided by the unidirectional fibers and nioidability/niachinabi!ity provided by the composite molding compound is required. The invention is applicable to any situation where a relatively strong structure is inquired that has a complex shape. 88024] The invention is particularly applicable to hybrid aircraft structures which require the use of molding compounds in combination with unidirectional fiber reinforcements to meet both strength and dimensional tolerances. Exemplary aircraft structures include window frames, access opening covers, outlet guide vanes for jet engines, thrust reverse* cascades, various engine airfoils, access doors, brackets, fittings, gussets, clips/cleats, Intercostals, pans,, flanges and stiffeners for aerospace structures

10 )1⁾25] An exemplary uneured composite assembly is shown at 10 in FIGS. 1 and 3. The composite assembly is shown as it looks once it is placed within a suitable moid (not shown) and formed into its final, shape prior to being molded at an elevated temperature under a relatively high pressure to form a multi -component composite structure. As shown in FIG. 3, the composite assembly Ϊ0 includes a structural component 11 that is made up of a structural element that is composed of a UD fiber body 12 and multi-directional continuous fiber body 14. The composite assembly 10 further includes a moldable component 16. The moldable component 16 has a face 18. The structural component 1 1 h s a face that is made up of the face 20 of the UD fiber body 12 and the face 22 of the mulii -directional fiber body 14, The meeting of the moldable component face 18 with the two faces 20 and 22 of the st ctaral component forms the interface 24 between the moldable and structural components.

|β8826] The uneured composite assembly 18 is formed from a preform i which the molding compound is located on top of the structural componen An exemplary preform is shown at 10a in FIG. 7. In the preform 1.0a, the amount of molding compound 16a that is located on the structural component 11a is sufficient to ensure that the molding compound 16a will flow within the mold, as .represented by arrow 15, to fill the mold cavity 16P (shown in phantom) as the preform 10a is being formed into its final shape as shown in FIGS. 1 and 3. There can be substantial -movement of the molding compound when the preform is placed within the mold. However, the structural elements 12a and 14a are located in the preform such that they do not move to any significant degree when the preform 10a is placed within the mold and formed into its final shape prior to curing. This allows one to accurately place the structural elements within the part while at the same time including the molding compound, which is allowed to move as needed in. the mold to form the desired final, shape.

00827] The molding compound 16 can be preformed prior to placement in the mold so that it closely matches the shape of the final, composite structure. However, it is difficult to accurately form a preform that matches the shape shown hi FIGS. 1 and 3 where an elliptical wall extends perpendicularly from a base plate. Accordingly, it is preferred that the molding compound 16a simply be located on the structural elements 12a and 14a, as shown in FIG, 7, with the moid being used to form the molding compound 16a and structural elements 12a and 14a into tlie desired final shape of the composite assembly 10, Once in die mold, the principal difference between the composite assembly 1.0 and the resulting aircraft window frame is that the thermosetting resin present in die iracured assembly 10 roast be completely cured,

100028] In accordance with the present invention, the coefficient of thermal expansion (CTE) of the moldable component^" at the interface 24 and the CTE of the structural component at the interface 24 are such that micro cracks do not form along the interface when the composite assembly 10 is cured molded. The CTE's of t e two components at the interface 24 must be sufficiently close to each other so that micro cracks do not form. Micro cracks typically form when the CTE of two adjoining materials differ, at least in one direction, creating a local strain from expansion or contraction that exceeds the materials ability to resist cracking. The CTE of the UD fiber body 12 and multidirectional fiber body 14, as well as the size, shape and relative orientation of the two bodies must be taken into consideration when designing a composite assembly that can be molded without micro cracking.

00029] in general, larger structural components should have a CTE that more closely matche the CTE of the moldable component. Likewise, inter aces that are larger and/or more complex require that the structural component have a CTE at the interface that more closely matches the CTE of the moldable component. For any given combination of structural and moldable components, the difference in CTE's that, can be tolerated without molding-induced micro cracking is determinable by routine experimentation.

00030] The molding compounds and structural elements that are used in accordance with the present invention are composed of fibers and resin. Fibers, such as carbon fibers, have a CTE (pails per millionfC) that is close to zero. Accordingly, the majority of the CTE of a composite material Is due to expansion and contraction of the resin matrix, in bulk molding compounds, where the fibers are discontinuous and randomly oriented, the CTE tends to be uniform in all directions. For sheets of molding compound where the fibers are quasi- Isotropically oriented, the CTE is uniform in all planar directions. The planar CTE of a typical sheet of quasi-isotropie carbon fiber/epoxy molding compound is on the order of 2 - 4 ppn /'C. The CTE in a direction perpendicular to the plane of the sheet of molding compound is controlled more by the resin matrix and tends to be higher than the planar CTE's, CTE's in the perpendicular direction are typically on the order of 20 to 40 ppmf C for a sheet of quasi - isoiropie molding compound.

00e31j The CTE for structural, elements is highly dependent upon the direction of the fibers. For example, the CTE of UD prepreg n ^'the direction parallel to the fibers (X direction) is due maiiily to fee fiber,. The matrix resin contributes very little to the CTE. As a result, fee CTE of UD prepreg in the X direction is close to zero. The CTE of an exemplary carbon fiber/epoxy UD prepreg in the X directi on is 0.01. ppm ^'C, The CTE of the same UD prepreg in the direction perpendicular to die UD fibers (Y and Z directions) is 30 - 40 ppmTC, which is due mainly to fee CTE of fee resin matrix.

00 2] Structural components made from woven fibers also exhibit CTE's that vary due to the orientation of fee fibers. However, the differences between X, Y and Z direction CTE's is not as great as in UD prepreg, since all of the fibers are not parallel to each other. The CTE in the X and Y direction of woven fiber prepreg tends to he controlled more by the matrix resi as compared to UD prepreg in the X direction. As a result, the CTE's in the X and Y direction of woven prepreg generally lie somewhere between the CTE of a similar UD prepreg in the X direction and the CTE of the UD prepreg in the Y/Z direction. The CTE of woven fiber prepreg in the Z direction is similar to the CTE of a similar UD prepreg in the Z direction, since the matrix resin in both types of prepreg contributes to the Z-direetion CT.E in a similar manner. 00033] The CTE of a molding component or structural component is measured using conventional procedures after complete curing of the individual components. The components contract as they coo! and expand as they are heated. Micro cracks can form during either contraction or expansion of the components. The CTE of the components, as measured after complete coring of the component, is considered to also be a. measure of the thermal contraction that occurs during cooling of the component from molding/curing temperature. In practice, the coring temperature of the component is most likely the highest temperature to which the component is exposed during the life of the component The contraction that occurs during fee initial cool down of fee molded part can be a source of considerable strain between the various components which does result in micro cracking when the strain exceeds the ability of the components to resist cracking.

00034] The CTE of continuous multi-directional fibers tends to more closely match the CTE of .DFC material than UD fiber layers, provided that the resin matrices are the same or similar. The multi-directional fiber body 14 is composed mainly of woven fibers so that the CTE of the ibex body 14 is relatively close to the CTE of the rnoldable component 16 at the interface 24, Accordingly, the risk of micro cracking along the interface between the niulti- directional fiber body 14 nd .rnoldable component 16 is minimal However, if the UD fiber body 12 is composed only of UD fiber layers, then the risk of micro cracking increases due to the increased difference in CTE between the DEC material of the rnoldable component and the body made up entirely of UD fibers.. This is especially a problem in the Z direction relative to the UD fiber body due to the relatively high CTE of the UD fiber body in this direction, which, is driven by the CTE of the resin matrix. Locating the UD fiber body 12 along the surface of the assembly allows for this added strain to occur in. the Z direction without causing micro cracking. The UD fiber body 12 is in the form of an ellipse in the X-Y plane. The CTE in this direction is due mainly to the UD fibers, so that the CTE is lower than in the Z direction. The expansion or contraction (for example during the cure cycle cool down phase) in the X-Y plane is also restrained by the surrounding molding compound.

00035] in accordance with the present invention, the risk of micro cracking is substantially eliminated by insuring that the UD fiber body 12 is located in the assembly and oriented so that the directional CTE's of the UD fiber body 12 are close enough to the CTE of the rnoldable component to avoid micro cracking along the component interfaces. FIG. 4 shows a detailed representation of an exemplary UD fiber body 12 in which layers 26 of continuous UD fibers are alternated with layers 28 of malii-dircctiona! continuous libers. The addition of multidirectional continuous fibers to the UD fiber body 12 changes the overall CTE of the body 12 so that it more closely matches the CTE of the rnoldable component 16. This provides a reduction in the potential, for micro cracking, but also reduces the unidirectional nature of the fiber body. The number of Iayers of UD fibers and mulfi -directional fibers can be varied to provide the desired structural properties while maintaining CTE profiles that are sufficient to avoid micro cracking at the interface 24. It wa found that the use of multi -directional fibers along the interface 24 is preferred because the CTE of the muM-directional fibers more closely matches the CTE of the random discontinuous fibers in the molding compound, hence reducing local strain at their resin interface.

[00036] In FIG, 4, the multi-direction continuous fiber layers 28 are intended to diagrammatieally represent woven fibers layers wherein the fiber orientation relative to the UD fibers alternates between 0/9(1 and +45/~4S, This is for exem lary purposes only. The woven fibers can be in a wide variety of orientations relative to the UD fibers and the various layers may have the same or different orientations. It is not necessary t¾at the layers alternate between UD and woven fiber layers. It is preferred thai one or more layers of woven fibers be located ^•next to the interface 24 as shown in. FIG, 4, The layeifs) of woven, fibers act as a CTE. buffer zone between the relatively low CTE moldable component and relativel high CET UD fiber layers,

108 37] The molti-directional fiber body 14 is located adjacent to the UD fiber body 12 to provide dissipation of CTE-indueed stress during curing of the composite assembly .10, The joint 23 between the two bodies is located at a relative thick section of the assembly 10 with the two bodies extending laterally from, each other into relatively thin sections of assembly 10. In an alternate orientation, the multi-directional fiber body 14 can be located between the UD fiber body 1.2 and the moldable component 16 to act as a CTE buffer, in either case, the multidirectional fiber body is located on the side of UD fiber body, which is the Mgh CTE direction, as opposed to abutting the end of the U D fiber body, which i the low CTE direction.

00038] An alternate embodiment of a composite assembly that is used to form an exemplary window frame is shown at 30 in FIG. 5. In this embodiment, the structural com onent 32 is relatively thin, as compared to the structural component 1 1 shown in FIG. 3. The moldable component of the window frame is also composed of DEC material The moldable component is shown at 34. I this embodiment, up to a few layers of contio ous UD fiber layers are used to form the structural component. This relatively thin sireotaral component can be molded with the DFC material to produce a window frame that is not micro cracked along the UD fiber/DFC material interface.

| K)39] The high CTE direction of the structural component 32 is the Z-direetion, The moldable component is located on only one side of the structural component 32. The difference in CTE between the DFC material and the UD fiber layer(s) in the Z-direciion is chosen such that micro cracking does not occur during cool down of the molded part. The difference in CTE's can he relatively la ge since the UD fiber iayer(s) are located at the surface of the part where thermal contraction/expansion can take place without micro cracking. The CTE in the X ~ Y plane is much less due to the fact that the UD fibers form a continuous loop that does not have an exposed end that forms an interface with the DFC material. This type of closed loop configuration is preferred over using multiple segments of UD fiber layers that form multiple interfaces at the ends of each segment

40- 108 40] If the combination of structural component thickness and CTE difference results hi the formation of micro cracks during molding of the window frame 30, then it is preferred, that ^•mutti~di^'fectio»al fibers be incorporated with the (ID fiber layers in accordance with the embodiment shown in FIG. 4. Alternatively, one or more layers of woven fibers can e placed between, the UD fiber layers 32 and the moklabie component. 34 to provide a CTE buffer between the two components,

10 )041] As shown in FIG. 8, it is preferred that a preform 30a he prepared in which the molding compound 34a is located adjacent to the UD fiber layer(s) 32a, The amount of molding compound 34a that is located on the U D fiber layers 32a is sufficient to ensure that the molding compound 34a will flow within the mold, as represented by arrow 35, to fill the mold cavity 34P (shown in phantom) as the preform 30a is being formed into its final shape as shown in. FIG. 5. There can be substantia! movement of the molding compound 34a when the preform is place within the mold. However, the UD fiber !ayer(s) 32a are located on the preform surface that, is in contact with the mold bottom surface that, is perpendicular to the direction of the press closure and pressure such that they do not move to any significant degree when the preform 30a is placed within the mold and formed into its final, shape prior to molding. This allows one to accurately place the UD fiber layer(s) within the part while at the same lime including the molding compound, which is allowed to move as needed in the mold to form the desired final shape.

00042] As previously mentioned in connection with the embodiment shown in FIGS, 1 and 3_S the molding compound 34 can he preformed prior to placement in the mold so that it closely matches the shape of the final composite structure. However, it als is difficult to accurately form a preform that matches the shape shown in FIG. 5 where an elliptical wall also extends perpendicularly from, a base plate. Accordingly, it is preferred that the molding compound 34a. simply be located on the structural elements 12a and 14a, as shown in FIG. 8, with the mold being used to form the molding compound 34a and UD fiber layer(s) 32a into the desired final shape of the composite assembly 30. Once in the mold, the principal difference between the composite assembly 30 and the resulting aircraft window frame is that the thermosetting resin present in the nncnred assembly .30 must he completely cured.

00043] it should be noted that the UD fiber layers shown in FIGS. 3 and 5 are bent in the width direction in order to form a planar reinforcing hoop within the molded window frame. In other words, the UD fibers form a co-planar layer in the X-Y plane and the UD fiber layer

1- remains within the X-Y plane as it is bent to form the reinforcing hoop. This type of UD fiber hoop configuration is preferably made using the UD pre -twist procedure mentioned in the Description of Related Ait.

|0ββ44] An exemplary composite assembly that is molded to Form a cover for an aircraft access opening is shown, in FIGS- 2 and 6 at 40. The assembly 40 includes a .rnoida^'ble component 42, which is composed of DFC material and a structural component, which is made up of structural elements 44 and 46. Attachment holes 48 ate shown in. the composite assembly. These holes can be Formed in the uncured composite assembly (as shown) or they can be drilled or otherwise machined into the multi-component structure after molding is completed.

β0045] The two structural elements 44 and 46 axe completely surrounded by the molding component 42. Structural element 44 is in the form of a hoop that includes faces 44a, 44b, 44c and 44d, which form a tubular interface with the moldable component 42, This tabular interface has a rectangular cross section. Structural element 46 is also in the form of a hoop that includes faces 46a, 46b, 46c and 46d. The structural clement 46 Forms a second tubular interlace with the moldable component 42. This second tabula interface also has a rectangular cross section. 100046] The structural element 44 includes a combination of UD fiber layers represented at 56 and woven fiber layers represented at 50, 52 and 54. The structural element 46 is shown having a single UD fiber layer 62 sandwiched between woven fiber layers 58 and 60. The number and orientation of UD fiber layers and woven fiber layers in structural elements 44 and 46 are exemplary only. Other orientations with different numbers of layers are possible. For example, the structural elements of the type shown In FIG. 4 may be tised. As previously mentioned, it is preferred that one or more layers of woven fibers he located between the UD fiber Iayers and the molding component in order to act as a CTE bnller between the UD fiber Iayers and the molding component,

100047] The UD fiber layers that make up structural clement 44 are bent in the width direction in the same manner as the UD structural elements shown in FIGS. 3 and 5. The UD fibers form a co~p!anar laye in the X-Y plane and they remain in the X-Y plane as the hoop is formed. As previously mentioned, this type of UD fiber hoop configuration is preferably made using the pre-twist process mentioned in the Description of Related Art, In contrast, the til^") fiber layer in sirac-nral element 46 is bent in the thickness direction. The UD fibers form a layer that, extends in the Z direction (substantially perpendicular to the ^'UD fiber layers of element 44). The UD fiber reinforcing hoop is formed by bending the UD layer hi the X-Y direction. This type of UD fiber hoop configuration is formed without the pre-twisting required for structural element 44. The CTE's of both types of UD fiber layers are similar m the direction perpendicular to the U^'D fibers as well, as hi the direction parallel to the UD fibers.

00048] The use of multiple structural elements within the moldable component, as shown in PIG. 6, presents a relatively complex set of interfaces which increases the chances of micro crack formation. The present invention solves the micro cracking problem associated with ^•multiple complex, interface configurations by reducing the CTE differential, between the structural elements and the molding component and/or by providing integral CTE buffers between the components. The UD fibers in adjacent structural elements 44 and 46 are preferably oriented in the same direction, as shown in FIG. 6, in order to match the directional CTE's of the UD fiber layers of the adjacent elements. Orienting the structural elements so that the UD fibers of adjacent elements are perpendicular to each other should be avoided doe to the significant differences in directional CTE's that would be present, at the interface or area between the two elements.

00049] The oneured resin used in the DEC, UD fiber layers and/or midti -directional fiber layers may be composed of any of the thermosetting or thermoplastic resins that are typically used for structural applications. The resins in the three different fiber materials can be different. However, it is preferred that the resins that are used in the structural, component and moldable component are the same or substantially similar to minimize CTE differences between the components. In addition, the .resin should be chosen such that, the strain to failure properties of the resin are sufficiently high to withstand the thermal strain to which adjacent components are subjected without forming micro cracks or otherwise failing. The fiber orientations and types are chosen, as discussed above, to minimize the strain to which a resin is subjected at any given interface between components,

00050] Preferably, the amount of uncured resin in the moldable component and structural component will be between.25 to 45 weight percent of the overall weight of the component. The uncured resin, may be any of the epoxy resins, bismaleimide resins, poiyimide resins, polyester resins, vinylester resins, cyanate ester resins, phenolic resins or thermoplastic resins that are used in structural composite materials. Exemplary thermoplastic resins include polyphenylene sulfide (PPS), polysulfone (PS), polyetlieretlierketone (PEEK), polyetlierketoneketone (PEf K), polyethersulfone (PES), polyetherimide (PET), polyamide-imide (PAl), Epoxy resins that are toughened with a thermoplastic, such as PES, PEI and/or PAl, are preferred resin matrices. Resins that axe typically present in UD tape of the type osed in the aerospace industry axe preferred. Exemplary feemioplastic toughened resins that are suitable for use as the resin matrix are described in United States Patent Mos. 7,754,322 and 7,968,179 and United States Patent Application No, 12/764,636, the contents of which are hereby Incorporated by reference.

[ΟΘ051] The moidabie component is preferably composed of randomly oriented segments of unidirectional tape that are impregnated with resin. This type of material is commonl referred to as nasi sotropie chopped ptepreg> Quasi -isotropic chopped prepreg is a form of random discontinuous fiber composite (DFC) that is available commercially from Hexcel Corporation (Dublin, OA) under the trade name BexMC^. As previously mentioned, Hex C*" has beers used for a variety of purposes including aerospace articles and high-strength molds. 0 )052] Quasi-isotropic (Q~l) prepreg is composed of segments or "chips" of unidirectional fiber tape and a resin matrix. Q-I prepreg is typically supplied as a mat made up of randomly oriented chi ps of chopped unidirectional tape prepreg. The size of the chi ps may be varied as well as the type of fibers depending upon the size and shape of the pre form as well as how precisely the pre-form must be machined o meet dimensional tolerances, if any. It is preferred that the chips be 1/3 inch wide, 2 inches long and 0.006 inch thick. The chips include unidirectional fibers that can be carbon, glass, aramid, polyethylene or any of the fibers types that, are commonly used in the aerospace industry. Carbon fibers are preferred. The chips are randomly oriented in the mat and they lay relatively flat This provides the mat with its transverse Isotropic properties.

01 053] The UD tape prepreg that is chopped to form the chips or segments includes a resin -matrix thai can be any of the resins mentioned previously that, are commonly used in aerospace prepregs. Thermosetting epoxy resins that are toughened with thermoplastics are preferred because they tend to be more resistant to fracturing or delamination if machming of the final composite part is required. The resin content of the chips may also be varied, between 25 and 45 weight percent of the total prepreg weight. Chips with resin contents of between 35 and 40 weight percent are preferred. No additional resin is typically added to the prepreg chips when forming the quasi-isolxopic chopped prepreg. The resin present in the initial UD tape prepreg is sufficient to bond the chips together to form the mat.

101 054] The quasi-isotropic (Q-i) chopped prepreg can be made by purchasing or making unidirectional prepreg tape or tow of desired width. The tape or tow is then chopped into chips of desired, length and the chips are laid randomly in layers to form the moidabie component. The randomly placed U^'D prepreg chips are combined with the structural component and pressed together to form the tm ored composite assembly (pre- form). When pressed together, the individual randomly oriented UD prepreg chips inherently bond together due to the presence of the prepreg resin. The preferred method, however, is to purchase HexMC* or equivalent commercially available uas SOiropic chopped prepregs, which are supplied as sheets of material that, are theft used to form the moldable component of the composite assembly

00055] An exemplary preferred qnasi-isotropic chopped prepreg material is HexMC* AS4/S552. This quasi- isotropic chopped prepreg material is supplied as a continuous roll of a mat. that is 46 cm wide and 0.20 cm thick. HexP!y* A$4/8552umdirectiona] fiber prepreg is used to make the chips that are randomly oriented in the quasi- isotropic mat HexPiy^* AS4/8552 prepreg is a carbon fiber (AS4)/epoxy (8552) unidirectional tape that is 0.016 cm thick and has a fiber area! weight of about .145 grams /square meter. The resin content of the tape is 38 weight percent with the resin (8552) being a thernioplastie-tonghened epoxy. The tape is slit to provide 0.85 cm strips and chopped to provide chips that are 5 cm long. The chip density is about 1.52 gram / cubic centimeter. Other exemplary quasi-isotropie chopped prepreg can be made using other HexPly* unidirectional prepreg tape, such as EMC A84/1M7 (epoxy/carbon fiber), IM7/8552 (thermoplastic-toughened epoxy/carbon fiber), 3501~6VT650 (epoxy/carbon fiber) and M7/M21 (thermoplastie-toiighened epoxy/caxhon fiber). HexMC^ AS4/8552 and Μ2Ί ΪΜ7 are preferred quasi-isotropie chopped prepregs for use in forming moldable components in accordance with the present invention,

|β00§6] Other types of DFC molding materials may be used to form the moldable component provided that they meet the necessary strength and machinahihty requirements for the intended structure. Such molding materials typically include randomly oriented chopped fibers which are impregnated with resin. However, in order to ensure that the molding material is sufficiently strong and both moldable and machinable, it is preferred that chopped unidirectional fibers or tape form at least 90 weight percent of the moldable component,

|0I 057| The structural element(s) thai make up the structural component include one or more layers of UD fibers. The same UD fiber prepreg tape that is used to make the DFC molding material can also be used to form the structural elements. The difference being that the structural elements are formed by one or more layers of continuous U^'D fibers, whereas the UD fibers in the moldable component are discontinuou and quasi -isotropieal!y oriented.

45- ICMKM5S] The UD fibers used in the structural elements may contain fioro a few hundred filaments to 12,000 or more filaments. UD fibers are typically supplied as a tape made up of eontinuons fibers in. a unidirectional orientation. UD tape is the preferred type of prepreg that is used to form the fibrous structure. Unidirectional tape is available from commercial sources or it .may be fabricated using known prepreg formation processes. The dimensions of the UD tape may be varied, widely depending upon the particular composite part being made. Fo example, the width of the U^'D tape may range from 0.5 inch, to a foot or more. The tape will, typically be from 0.004 to 0.012 inch (0.01 to 0.03 cm) thick and the length of the UD tape (the dimension parallel to the continuous UD fibers) may vary from 0.5 inch (1.3 cm) »p to a few feet (one meter) or more depending upon the size and shape of the structural element^". 011059] A preferred exemplary commercially available unidirectional prepreg thai can be used to make the structural elements is BexPly® 8552, which is available from Hexcel Corporation (Dublin, California). HexPIy€>8552 is available in a variety of unidirectional tape configurations that contain an. amine cured toughened epoxy resin matrix in amounts ranging from 34 to 38 weight percent and carbon or glass UD fibers having from 3,000 to 12,000 filaments, Tbe fibers typically account for 60 volume percent of the U^'D tape. The preferred UD fibers are carbon fibers. Other HexPly* unidirectional prepreg tape may be used in. the structural elements. These UD prepreg tapes include EMC .6/AS4 (epoxy/carbon fiber), 855.2/1M7 (thermoplastic-toughened epoxy/carbon fiber), 3501-6/T650 (epoxy/carbon fiber) and M21/1M7 (thermoplastic-toughened epoxy/carbon fiber). The CTE of this type of UD tape in the direction parallel to the fibers is close to 0 (O.OippmTC) and between 30 and 40 ppmfC in directions perpendicular to the direction o the UD fibers.

100060] The multidirectional fiber layers that are combined with the UD fiber layers to form the structural elements can he non-woven or woven fiber fabric or randomly oriented continuous fibers in tbe fomi of a veil Other types of mulii-directiona! continuous fiber orientations may be used, but it is preferred that the CTE of the melti -directional fiber layer (including .matrix resin) is betwee the CTE's of the DEC molding material and U^'D fiber layer prepreg. It is preferred that the CTE of the multi-directional fiber layer is close to being half way between the CTE of the DEC material, in the molding component and the CTE of the UD fiber layers in the structural component, "close to" means that the CTE of the multi-directional layer is within 20 percent of the halfwa point between ttie CTE's of the DFC material and the OD fiber layer(s),

00061 j in some situations, it is possible to «se one or more layers of DFC molding material in place of woven rnulti -directional fiber layers within a given structural element. However, it is preferred that such use of DFC -molding ^"material layers be limited to the central portion of the structural clement and that the use of DFC molding material in the interior of the structural element be kept low enough to avoid possible micro cracking internally within the structural element. The layers of DFC molding material may be alternated with layers of UD fibers in the same manner as the woven, nmlti-direetional fiber layers, if desired.

100062] The resins used as the resin matrix in the multi-directional fiber layers should be the same as those used in the .DFC molding material and UD fiber layers. The fibers should also be the same. Since the CTE's of the three different types of materials depends on the resin matrix, fiber orientation, fiber type and resin loading of the fibers, it is possible to t ne nine the CTE's of the materials by varying these four parameters. Preferably,, the resin type, fiber type and resin loading for the DFC molding material UD fiber layers and multi-directional fiber layers are the same or similar. In addition, the fibers of adjacent components should be oriented so that the high CTE directions and low CTE directions of the adjacent components, if any, arc matched to minimize differences in directional CTE's at the interfaced) between the components.

00063] Molding of the composite assemblies is carried out. according to known molding procedures of DFC. Tire uncured composite assembly is placed in a mold that is typically composed of two mold halves ami formed into the desired shape. Once formed in the mold, the tmcored composite assembly is heated to the coring temperature of the resin(s) and molded at high pressure to form the muM-cmnponent composite structure. Typical, high-pressure curing temperatures for epoxy resins range from I ?£FC to 225 "C. Preferred curing temperatures range fiom nVfC to 205 "C. .Internal, pressures within the mold are preferably above 500 psi and below 2000 psi at the cure temperatures. Once the uncured composite assembly has been completely cured (typically 5 minutes to 1 hour at curing temperature), the part is removed from the mold and cooled to form the final multi-component composite structure. It is dining this initial post-molding cooling process that micro cracking will most likely occur, if required, the mult? -component composite structure may be machined to form .final surface shapes and provide any precise dimensions that are required. 108 6 ] If desired, the luiciired composite assembly can be "B-sfaged" prior to being placed in the mold in ord r to increase the viscosity of the resin. B-siagiog is a known partial exiling procedure that Invol ves heating the marred composite assembly at ambient pressure to a temperature of 165 "C to I S0^!!C for just enough time to substantially increase the ^"viscosity of the resin. B-staging times on. She order of 5 to I 5 minutes at the B-staging temperature are preferred. The B-staged composite assembly is preferably cooled to room temperature prior to being placed in. the mold for final shaping and curing. In addition, the viscosity of the resin in the unenred composite assembly tends to drop as the assembly is heated to cure temperature and then rapidly increases as the resin cures. It is preferred that the mold not be pressurized until after the resin has reached the minimum viscosity, in practice, the B-staged composite assembly is placed in the mold, which has already been heated to the curing temperature. Presswization of the mold is delayed from a few second to a minute or more in order to allow the resin time to move through the minimum, viscosity phase.

f 08065] Micro cracking occurs when the local tensile strain at the interface between two components exceeds the maximum strain capability of the resin matrix. For a typical epoxy resin, the maximum strain capability is 2,4% before micro cracks form. Accordingly, it is preferred that the resin matrix, fiber orientation, fiber type and resin loading of the fibers combinations, as well as the size, shape and relative orientation of the various components be chosen so that the strain at any given interface does not exceed 2,0% when the cured composite part is cooled down from the earing temperature to room temperature.

|β8866] Examples of practice are as follows;

EXAMPLE 1 00867] An xuieured composite assembly for making an aircraft window frame having the shape and structure shown in FIGS. I and 3 was prepared. The moldab e component 6 was formed from HexMC€⁾AS4/8552, which is a DFC molding material composed of discontinuous iJD A.S4 carbon fibers with an epoxy 8552 resin matrix. The BexMC®AS4/S552 had an area! weight of about 1925 gsm with the resin content being about 38% of the total weight of the molding .material. The multi-directional fiber body .14 was made up of S layers of plain weave AS4 carbon fibers in an epoxy 8552 resin matrix (AS4/8552). The area! weight of each woven

48- fiber layer was aboot 200 gsm wife the resin content being about 40% of the total weight of the multidirectional fiber body 14. The UD body 12 was formed as 16 layer laminate made up of alternating layers of HexCurve® IMA/8552 and A.S4/SSS2 plain weave fabric. BexCurve® IMA/8552 is a carbon fiber UD tape that has been pre-twisted as previously described to allow bending in the width direction. The BexCurve® .IMA/8552 had an area! weight of about 268 gsm with the resin content being about 34% of the total weight of the HexCt?.rve# UD tape. f0f)068] The isolda ie component 16, UD body 12 and multi-directional fabric body .14 where formed into the composited assembly as shown in FIGS. I and 3 for molding into an aircraft window frame. The composite assembly was placed in a suitable mold and cured at 18CTC for 30 minutes at an interna! mold pressure of 1500 psL The cored composite assembly was cooled from the curing temperature to room temperature and then removed from the mold. The resulting multi-component window frame did not have any micro cracks at the boundaries between the moldable component 1.6_» UD body 12 and multi-directional body 1

0006 ] A comparative window frame was prepared in the same manner as the exemplary window frame, except that the UD body 12 was made up only of 16 layers of HexCurve® UD fibers. Micro cracks were observed at the boundary between the 16-layer UD laminate and the moldable component 16.

[00070] Having thus described exemplary embodiments of the present invention, that various other alternatives, adaptations and modifications may be made within the scope of the present invention. Accordingly, the present invention is not limited by the above-described embodiments, but is only limited by the following claims.

Claims

What is claimed is:

1.. A composite assembly which can be core to form a multi-component composite structure, said composite assembly comprising: a structural component which comprises a sfcractural element that comprises an uncured thermosetting resin and unidirectional continuous fibers, said structural component comprising a stmcteal component face; and a rnoldable component which comprises a moldable body comprising an uncured thermosetting .resits and discontinuous fibers, said moldable component comprising a moldable component face wherein said structural component face and said moldable component face share a common interface located within the interior of said composite assembly and wherein the coefficient of thermal expansion of said structural component at said interface and the coefficient of thermal, expansion of said moldable component at said interface are such that micro cracks do not form along said interface when said composite assembly is eared to form said multi-component composite structure.

2. A composite assembly according to claim 1 wherein said structural component comprises multi-directional continuous fibers.

3. A composite assembly according to claim 2 wherein said muM-direciional continuous fibers comprise woven fibers.

4. A composite assembly according to claim 2 wherein said structural component comprises alternating layers of unidirectional, continuous fibers and mufti -directional continuous fibers,

5. A composite assembly according to claim 2 wherein said multi-directional continuous fibers are located between said unidirectional fibers and said interface.

6. A composite assembly according to claim 4 wherein a layer of said multi-directional continuous fibers is located between said unidirectional fibers and said interface.

7. A composite assembly according to claim 1 wherein said structural component is completely surrounded by said moldable component.

8. A composite assembly according to claim 7 wherein at least two structural components are located within said moldable component

9. A composite assembly according to claim 8 wherem a first structural component comprises continuous wiidirectioiial fibers that are oriented in a first planar arrangement and a second structural component comprises unidirectional fibers that are oriented in a second planar arrangement, wherein said first planar arrangement and said second planar arrangement are not coplanar.

10. A multi-component composite structure which is formed by curing a composite assembly according to claim. I .

1 1. A multi-component composite structure according to claim 10 wherein said structural component comprises multi-directional continuous fibers.

12. A multi-component composite structure according to claim 1 1 wherein said structural component comprises alternating layers of unidirectional continuous fibers and multi-directional continuous fibers.

13. A inu -componeni composite structure according to claim 1 wherem said multidirectional continuous fibers are located between said unidirectional fibers and. said interface.

14. A multi-component composite structure according to claim 12 wherein a layer of multi- directtonal continuous fibers is located between said unidirectional fibers and said interface.

13, A multi-component composite structure according to claim 10 wherein said structural component is completely surrounded by said moldable component

16. A multi-component composite structure according to claim 15 wherein at least two structural components are located within said moldable component.

17. A multi-componeni composite structure according to claim 16 wherem a first structural component comprises continuous unidirectional fibers that are oriented in a first planar arrangement and a second structural component comprises unidirectional fibers that are oriented in a second planar arrangement^", wherein said first planar arrangement and said second planar arrangement are not coplanar. 18, A multi-cornponent composite structure according to claim 10 which forms pari of an aircraft,

19, .A. method for making a composite assembly which can be cured to form a .m¾ -component composite structure, said method comprising the steps of; providing a structural component which comprises a structural element that comprises aii uneured thermosetting resin and unidirectional continuous fibers, said structural component comprising a structural component face; and providing a xnoldable component which comprises a moldable body comprising an uncured thermosetting resin and discontinuous fibers, said moldable component comprising a nioldabie component face; and combining said structural component and said moldable component together to form said composite assembly such that said structural component face and said moldable component face share a common interface located within the interior of said composite assembly and wherein the coefficient of thermal expansion of said structural, component at said interface and the coefficient of thermal expansion of said moldable component at said interface are such th at micro cracks do not form along said interface when said composite assembly is cored to form said multi-component composite structure.

20, A method for making a multi-component composite structure comprising the step of curing a composite assembly according to claim 1.