Patent Description:
In one process for producing elongated composite laminate parts having contoured geometries, unidirectional prepreg plies are laid up, either by hand or using automated layup equipment, to form a flat laminate stack. In an initial forming operation, a pair of matched dies punch form the flat stack into a straight part having a desired cross-sectional shape. In a secondary forming operation, the part is formed onto a contoured forming tool which imparts a desired contour to the part along its length. As a result of these two forming operations, the part has contours along two axes.

Traditional laminate layups typically use a combination of <NUM>°, <NUM>° and <NUM>° plies. When using these traditional laminates in the process described above to produce contoured composite laminate parts having high aspect ratios, such as stringers and spars used in aerospace vehicles, ply wrinkling may occur because some of the reinforcing fibers in the plies having <NUM>° orientations. The fibers in the <NUM>° plies strain in response to being loaded in compression along the entire length of the part during the secondary forming operation. The problem of ply wrinkling can be more pronounced where the part has a large number of plies and/or contains joggles or aggressive ply ramps along its length. Ply wrinkling is undesirable. In some applications, parts can be reworked to reduce or eliminate ply wrinkling, however the rework adds to labor costs and may reduce production rate. One solution to the wrinkling problem involves cutting the <NUM>° plies into segments however this may decrease load carrying ability. The reduced load carrying ability can be compensated by adding additional plies to the part, however this approach to the problem increases material costs and part weight.

Document <CIT>, according to its abstract, states a composite laminate having a primary axis of loading and comprises a plurality resin plies each reinforced with unidirectional fibers. The laminate includes cross-plies with fiber orientations optimized to resist bending and torsional loads along the primary axis of loading.

Document <CIT>, according to its abstract, states a method of designing a composite laminate, the laminate comprising a plurality of zones, each zone comprising a plurality of plies of composite material, each ply in each zone having a respective ply orientation angle. A global stacking sequence is determined for the laminate, the global stacking sequence comprising a sequence of stacking sequence elements. A local laminate thickness is determined for each zone. A local stacking sequence is then determined for each zone by extracting a subsequence of stacking sequence elements from the global stacking sequence. The global stacking sequence and the local laminate thicknesses are determined together in an optimization process in which multiple sub-ply selection variables are assigned to each stacking sequence element, each sub-ply selection variable representing the density or sub-ply thickness for a respective candidate ply orientation angle. Optimal values are determined for the sub-ply selection variables and the local laminate thicknesses. A single one of the sub-ply selection variables is assigned to each stacking sequence element thereby forcing a discrete choice of global ply orientation angle for each stacking sequence element.

Document <CIT>, according to its abstract, states a curved composite element being laid up over a tool having first and second curved tool surfaces possessing differing radii of curvatures. A plurality of composite fiber ply Segments are arranged in substantially side-by-side relationship into a group. The ply segments are formed as a group onto the first curved tool surface, and the group is then formed from the first curved tool surface onto the second curved tool surface.

Document <CIT>, according to its abstract, states an aerospace vehicle including a plurality of composite stiffeners. Each stiffener of the plurality has a stack of plies of reinforcing fibers. At least some of the plies in the stack have reinforcing fibers oriented at ±a with respect to an axis of primary loading, where a is between <NUM> and <NUM> degrees. At least some of the plies in the stack have reinforcing fibers oriented at ±ß with respect to the axis of primary loading, where ß is between <NUM> and <NUM> degrees.

According to claim <NUM>, there is provided a method of making a contoured composite laminate part having a high aspect ratio, a major axis of loading and a plurality of zones along its length respectively having desired stiffnesses, comprising selecting a set of fiber angles (+θ<NUM>, +θ<NUM>, ±θ<NUM>) for plies of unidirectional reinforcing fibers, wherein none of the fiber angles is <NUM>° relative to the major axis of loading, wherein the set of fiber angles includes three fiber orientations at off-angles (+θ<NUM>, +θ<NUM>, ±θ<NUM>) relative to the major axis of loading of θ<NUM>, θ<NUM>, θ<NUM>, where <NUM> < θ<NUM> < θ<NUM> < θ<NUM> ≤ <NUM>°, wherein θ<NUM> - θ<NUM> ≤ <NUM>°, and θ<NUM> - θ<NUM> ≤ <NUM>°; determining, for each of the fiber angles (+θ<NUM>, +θ<NUM>, ±θ<NUM>) , a number of plies in each of the zones required to provide a desired set of in-plane laminate properties in the zone); determining a shape and stacking sequence of the plies, wherein some of the plies are continuous and other of the plies are discontinuous along the major axis of loading, and wherein the off-angles (+θ<NUM>, ±θ<NUM>, ±θ<NUM>) include balanced pairs of + and - angles; laying up the plies into a flat stack using the stacking sequence; and forming the flat stack into the shape of the contoured composite laminate part.

According to claim <NUM>, there is provided a composite laminate stiffener contoured along a major axis of loading, comprising a plurality of laminated plies of unidirectional reinforcing fibers held in a plastic matrix, wherein all of the plies have fiber orientations at off-angles (+θ<NUM>, +θ<NUM>, ±θ<NUM>) relative to the major axis of loading, wherein the plies have three fiber orientations relative to the major axis of loading of θ<NUM>, θ<NUM>, θ<NUM>, and <NUM> < θ<NUM> < θ<NUM> < θ<NUM> ≤ <NUM>°, wherein θ<NUM> - θ<NUM> ≤ <NUM>°, and θ<NUM> - θ<NUM> ≤ <NUM>°, wherein the composite laminate stiffener is devoid of <NUM>° plies, wherein some of the plies are continuous and other of the plies are discontinuous along the major axis of loading, and wherein the off-angles (+θ<NUM>, +θ<NUM>, ±θ<NUM>) include balanced pairs of + and - angles.

One of the advantages of the fabrication method is the elimination of plies having <NUM>° fiber orientations which have a tendency to wrinkle when formed to a contoured geometry.

Another advantage of the fabrication method is that the plies having fiber orientations primarily intended to provide strength and stiffness along the longitudinal axis of the part are shorter in length, compared to <NUM>° fibers which extend entire length of the laminate, and permit increased axial strain before being subject to buckling. The reduction in the length of these fibers reduces the amount of friction between the plies, allowing transverse slip to take place between them, which in turn reduces the compression of those fibers having the highest tendency to buckle.

A further advantage of the fabrication method is that composite laminate parts with contoured geometries and high aspect ratios can be produced in which the plies having fiber orientations providing strength and stiffness along the longitudinal axis of the part are oriented such that they transition from a compressive state to a neutral and then a tensile state during forming, permitting these plies to relax rather than buckle.

Still another advantage of the embodiments is that a highly contoured composite laminate part, such as a contoured stiffener, can be produced that provides adequate stiffness along a major axis of loading without the need for plies having a <NUM>° orientation and without increasing the weight of the part.

The features, functions, and advantages can be achieved independently in various embodiments of the present disclosure or may be combined in yet other embodiments in which further details can be seen with reference to the following description and drawings.

The illustrative embodiments, however, as well as a preferred mode of use, further objectives and advantages thereof, will best be understood by reference to the following detailed description of an illustrative embodiment of the present disclosure when read in conjunction with the accompanying drawings, wherein:.

Referring first to <FIG>, a composite laminate part <NUM> is contoured along its length L and has a radius of curvature R. In the illustrated example, the part <NUM> is a stringer <NUM>, also referred to herein as a stiffener <NUM>, used to transmit loads in a structure such as the airframe of an aircraft, however principles of the disclosed embodiments may be used in the fabrication of a wide range of other types of contoured composite parts, especially structural stiffeners, having various cross sectional shapes. As used herein "contour" and "contoured" are each used in its broadest sense, and includes but is not limited to curvatures in any portion, or throughout the length of the part <NUM>. "Contour" and "contoured" also include curvatures or other geometric features having either a constant or a changing radius of curvature, as well as local changes in geometry such as, without limitation, joggles. The stiffener <NUM> has a hat section <NUM> defined by a cap <NUM> and a pair of webs <NUM>. The webs <NUM> connect the cap <NUM> with a pair of flanges <NUM> that extend outwardly. The stiffener <NUM> has a length L that is significantly greater than its width W, and thus has a high aspect ratio.

The stiffener <NUM> has a major axis of loading <NUM>, which in the illustrated example, is aligned with the X axis in the coordinate system shown at <NUM>. The stiffener <NUM> thus possesses double contour. The first contour is along the length of the stiffener <NUM> in the XZ plane, and the second contour defined by the hat section <NUM> is in the YZ plane. While a hat stringer <NUM> is illustrated, principles of the disclosed embodiments are also applicable to other types of stiffeners, including but not limited to stiffeners having other cross-sectional shapes such as a Z-shape, a C-shape, a rounded hat shape, or a blade (an I-shape), etc. Principles of the disclosed embodiments are likewise applicable to other types of composite laminate structural members such as spars and floor beams that are contoured in one or more planes and/or have cross-sectional shapes that vary along the length of the member.

Referring now to <FIG>, the stiffener <NUM> shown in <FIG> is fabricated by forming a flat stack <NUM> of composite plies <NUM> into a desired cross sectional shape and longitudinal contour. The plies <NUM> each comprising unidirectional fibers <NUM> held in a suitable plastic matrix <NUM>. The fibers <NUM> may be any material suitable for the application including, but not limited to carbon, glass, aramids, ceramic or any combination thereof. The plastic matrix <NUM> may be a thermoset or a thermoplastic, or a hybrid material system that includes both a thermoset and a thermoplastic. In the illustrated example, prepreg plies are laid up to form the stack <NUM>, however principles of the embodiments are also applicable to the layup of a stack of dry fibers which are subsequently infused with the plastic matrix <NUM>.

The ply <NUM> shown in <FIG> is a full, continuous ply, however the flat stack <NUM> may include partial, or discontinuous plies (not shown). The fibers <NUM> in each of the plies <NUM> are oriented at various angles θ relative to the major axis of loading <NUM>, as will be discussed later in more detail. In the illustrated example, the plies <NUM> comprising the flat stack <NUM> are balanced. In a stack <NUM> with balanced pairs of fiber angles, the plies <NUM> are arranged in pairs of equal positive and negative angular orientations. In other examples, however the plies <NUM> may be unbalanced. Further, the flat stack <NUM> may be symmetric or unsymmetric. In a symmetric stack <NUM>, the sequence of the plies on either side of a mid-plane <NUM> of the stack <NUM> are mirror images of each other. As will be discussed below, all of the plies <NUM> are oriented at off-angles θ relative to the major axis of loading <NUM>, thus, none of the fibers <NUM> have <NUM>° fiber orientations.

Attention is now directed to <FIG> which illustrate a die set <NUM> used to stamp form the flat stack <NUM> into a straight stiffener 30a having a desired cross-sectional shape, which in the illustrated example is a hat shape. The die set <NUM> comprises matching male and female dies <NUM>, <NUM> respectively, that are placed in a press (not shown) or other machine which forces the dies <NUM>, <NUM> together. The male die <NUM> includes a punch <NUM> and a pair of die flanges <NUM>. The female die <NUM> includes a die cavity <NUM> having a cross sectional shape that matches that of the punch <NUM>. In preparation for forming operation, the flat stack <NUM> is placed on upper surfaces 60a of the female die <NUM>. Then, as shown in <FIG>, the die set <NUM> is closed causing the punch <NUM> to force a portion of the flat stack <NUM> into the die cavity <NUM>, while the die flanges <NUM> compress other portions of the stack <NUM> against the upper surfaces 60a of the female die <NUM>.

<FIG> illustrate a cure tool <NUM> that is used to form the straight stiffener 30a to a desired contour along its length, and maintain the shape of the fully formed stiffener <NUM> during curing. The cure tool <NUM> is provided with contoured tool surfaces <NUM> that match the shape of the contoured stiffener <NUM> shown in <FIG>. In preparation for contour forming, the straight stiffener 30a is placed on the cure tool <NUM>, and the assembly of the stiffener 30a and cure tool <NUM> is then vacuum bagged (not shown) and placed in an autoclave (not shown). The combination of heat and pressure P applied to the stiffener 30a in the autoclave, form it down onto the contoured tool surfaces <NUM> and cure the stiffener <NUM>.

It should be noted here that while a two-stage process for forming the part <NUM> has been described in the illustrated embodiment, other processes, including a single stage process may be employed in which all contours, both longitudinal and traverse, are formed of a single forming operation. For example, where the plastic matrix is a thermoplastic, the flat stack can be heated to forming temperature and stamped formed to final shape in a consolidation press. Moreover, while thermal curing may be used where the plastic matrix is a thermoset, other curing methods may be employed, depending upon the particular material system being used, including but not limited to curing the formed thermoset part <NUM> at room temperature.

<FIG> illustrates several unidirectional plies 44a-44e of the stiffener <NUM> which comprises a balanced laminate that is devoid of <NUM>° plies. The plies 44a-44e have fiber angles relative to the major axis of loading <NUM>, of ±θ<NUM>, ±θ<NUM> and ±θ<NUM>, where<MAT><MAT> and<MAT>.

Plies 44a-44e are termed "off-angle" plies because the fibers <NUM> in these plies form angles with respect to the major axis of loading <NUM>. +θ<NUM> is within the ranges of approximately +<NUM>° up to approximately +<NUM>°, and -θ<NUM> is within the ranges of approximately -<NUM>° up to approximately - <NUM>°. The fibers <NUM> having orientations of ±θ<NUM> provide the laminate stiffener <NUM> with primary axial or longitudinal stiffness, while the fibers <NUM> having ±θ<NUM> fiber orientations provide the laminate with a lesser amount of axial stiffness, and some degree of transverse stiffness. As used herein, "primary axial stiffness" means that the fibers <NUM> in the ply <NUM> primarily provide the part <NUM> with longitudinal or axial stiffness, rather than with traverse stiffness. In the illustrated example, the plies having a <NUM>° orientation (θ<NUM> = <NUM>°) provide the stiffener <NUM> with transverse stiffness.

Attention is now directed to <FIG> which illustrate the off-angle orientation of one of the fibers <NUM> providing the stiffener <NUM> with primary axial stiffness. The fiber <NUM> may form part of the ply <NUM> shown in <FIG> that has an off-angle fiber orientation of +θ<NUM>. As shown in <FIG>, the fiber <NUM> has a length L' that is less than the length L (<FIG>) of the stiffener <NUM>, and is thus shorter in length than fibers in a <NUM>° ply (not shown) of a conventional laminate which would otherwise extend the entire length L of the stiffener <NUM>. Referring now also to <FIG>, buckling <NUM> of the fiber <NUM> during forming of the straight stiffener 30a to a longitudinal contour is a function of the longitudinal strain εx on the fiber <NUM>, the length L' over which the strain εx is applied and boundary conditions affecting the fiber <NUM>. The tendency of the fiber <NUM> to buckle <NUM> can be reduced by reducing the longitudinal strain εx on the fiber <NUM>. Reducing the length L' of the fiber <NUM>, allowing the plies and thus the fiber <NUM>, to slip <NUM> in plane and placing the fiber <NUM> in shear <NUM> due to in-plane twisting during forming, all contribute to reducing the longitudinal strain εx on the fiber <NUM>, and thus the potential for buckling <NUM>.

As will be discussed below, off-angle plies <NUM> are less likely to wrinkle than <NUM>° plies when the straight stiffener 30a (<FIG>) is formed to the desired longitudinal contour. The use of off-angle plies <NUM> reduces ply wrinkling for several reasons. First, off-angle plies <NUM> reduce the length L over which the individual fibers <NUM> are compressed <NUM> (<FIG>) during forming, and convert a portion of the stretching (εx) into shear deformation <NUM> (<FIG>). Second, the off-angle plies <NUM> are allowed to relax <NUM> to some degree during forming because the fibers <NUM> having orientation angles of ±θ<NUM> that provide the primary axial stiffness transition from a compressive state <NUM> at the caps <NUM> to a neutral state <NUM> at the webs <NUM>, and then to a tensile state <NUM> at the flanges <NUM>. This relaxation <NUM> of a portion of the length L of the fibers <NUM> reduces their tendency to buckle <NUM> during the forming process. Third, because the off-angle fibers <NUM> are shorter in length L (than <NUM>° fibers), some degree of transverse slip <NUM> between the plies <NUM> (<FIG>) takes place during forming, and this ply slippage resulting in a reduction of the compression <NUM> of the fibers <NUM>. Fourth, due to the lower loading on the fibers <NUM> in the off-angle plies <NUM>, the strain εfiber on the off-angle fibers <NUM> is reduced according to εfiber = εx*COS<NUM>(θ), where θ is the angular orientation of fiber <NUM> relative to the major axis of loading <NUM>, and εx is the strain of a ply <NUM> in the longitudinal direction <NUM> (<FIG>).

Using plies <NUM> with selected combinations of off-angle orientations, and preselected ply sequences, a laminate part <NUM> may be produced without the need for <NUM>° plies which provides essentially the same stiffness and performance as an equivalent laminate of comparable weight that relies on <NUM>° plies for axial stiffness. Thus, an existing stiffener design uses <NUM>° plies may be redesigned using off-angle plies <NUM> in order to reduce ply wrinkling without sacrificing laminate stiffness or increasing the weight of the part <NUM>.

Reference is now made to 11A and 11B which respectively show two possible layup sequences <NUM>, <NUM> for a contoured laminate part, wherein the ply orientation angles <NUM> are shown for each of the plies <NUM> in the layup sequence. <FIG> shows the sequencing of a <NUM> ply laminate part using a traditional combination of <NUM>°, ±<NUM>° and <NUM>° plies. <FIG> shows a redesigned sequencing of the same <NUM> ply laminate part having the same laminate thickness which avoids the use of <NUM>° plies in order to reduce ply wrinkling during forming. The layup sequence shown in <FIG> uses a combination of ±<NUM>°, ±<NUM>°, ±<NUM>° and <NUM>° plies, and results in a contoured laminate part that exhibits stiffness equivalent to the laminate part produced using the ply sequence shown in <FIG>, and without increasing part weight.

In some applications, a contoured composite laminate part <NUM> may have different stiffness requirements in different areas of the part. For example, referring now to <FIG>, the contoured composite stiffener <NUM> may have differing stiffness requirements in different zones <NUM> along its length. Different stiffness properties in the different zones <NUM> may be achieved by varying the ply orientations, and/or varying the number of plies of a given orientation in each of the zones <NUM>. For example, referring to <FIG>, the stiffener <NUM> may have a thickness T<NUM> in zone <NUM> that is greater than the thickness T<NUM> in zone <NUM> but less than the thickness T<NUM> zone <NUM>. Ply ramps <NUM> are used to transition between zones have differing thicknesses T.

Referring also now to <FIG> and <FIG>, a laminate part <NUM> having differing thicknesses along its length to provide individual zones of tailored stiffness properties may be achieved by laying up a combination of full plies <NUM>' and partial plies <NUM>" (<FIG>) of selected fiber orientations in a predetermined sequence. <FIG> illustrates the layup sequence for producing differing stiffnesses in each of zone <NUM>-<NUM>. In this example, the laminate part <NUM> has differing ply thicknesses T (<FIG>) in various ones of the zones <NUM>-<NUM> based on whether a full ply <NUM>' or partial ply <NUM>" (<FIG>) stretches (εx) over that zone. The laminate part <NUM> represented by the layup sequence shown in <FIG> and <FIG> includes a combination of full and partial plies having angular orientations of ±<NUM>°, ±<NUM>°, and <NUM>° cross plies. In this example, the ±<NUM>° off-angle plies provide the primary axial stiffness.

<FIG> broadly illustrates the overall steps of one method for producing a contoured composite laminate part <NUM> having reduced wrinkling and exhibiting differing stiffnesses along its length. In this example, the method is used to produce an existing part design that utilizes <NUM>° plies with a new part design that avoids the use of <NUM>° plies. As will be discussed below, the shapes of the plies are selected and optimized only after the ply orientations (fiber angles) and the number of plies per angle are determined for each zone <NUM> having particular stiffness requirements.

Beginning at <NUM>, an existing part <NUM> to be replaced is selected which has part specifications that are required to be met including but not limited to differing stiffness properties along its length. At <NUM>, information is extracted from the existing part design such as, without limitation, the number of plies per orientation in each zone, material properties and zone dimensions. At <NUM>, continuous ply thickness values tij are determined for various ply angle combinations that match the existing part laminate stiffness and thickness. The determination made at <NUM> includes selecting the number of new fiber orientations θ used for the part 30a, which includes limiting the laminate to three fiber orientations θ<NUM>, θ<NUM>, θ<NUM>, between <NUM>° and <NUM>°, wherein <NUM> < θ<NUM> < θ<NUM> < θ<NUM> ≤ <NUM>. Limiting the number of fiber orientations to a relatively small number, such as three fiber orientations θ<NUM>, θ<NUM>, θ<NUM> allows a full design space to be investigated by iterating over all possible combinations of θ<NUM>, θ<NUM>, θ<NUM>. In practice, fiber orientations are limited to integer numbers between <NUM> and <NUM>°.

From classical lamination theory (CLT), the stiffness properties of a laminate may be expressed as a function of a set of interrelated stiffness parameters. Assuming a balanced laminate, there are two equations defining the lamination parameters and one equation for the total laminate thickness. Selecting a set of three fiber angles therefore results in the following three equations per layup zone <NUM>, with the three ply thicknesses as unknowns, where the fiber angles are the same for all zones: <MAT> <MAT> <MAT>.

The above sets of equations for each zone are independent from the equations for the other zones. Only solutions with positive thickness values for all plies in all of the zones are selected. Only those combinations of three fiber angles that result in the desired laminate properties are selected. At this point in the process, all of these combinations result in the same stiffness, but not all of them can be made, as a practical matter, because the thicknesses typically do not correspond to an integer number. Certain combinations of the fiber angles may be eliminated based on certain composite laminate design rules. Only fiber angle combinations that meet the following constraints are considered:<MAT><MAT> From the above description, it may be appreciated that the process of determining the continuous ply thickness values in step <NUM> comprises selecting, from multiple possible combinations of fiber angles, a set of fiber angles and determining, for each of the zones, the thickness of the laminate within that zone that will provide the desired stiffness properties.

After a continuous thickness solution is obtained at <NUM>, the continuous solution is reduced to a solution with a discrete value or integer number of plies <NUM> that, based on the set of fiber angles selected from the possible combinations of angles, provides the desired stiffness within a zone. At step <NUM>, the ply thickness values T are refined by performing discrete ply thickness integer optimization. The discrete ply thickness integer optimization process is a mixed integer optimization problem with an objective of minimizing the difference between the resulting and optimum lamination parameters. The process performed at step <NUM> comprises calculating the number of plies <NUM> with discrete ply thicknesses for layups in all of the zones <NUM>, thereby ensuring balance and nonzero ply counts. The completion of steps <NUM> and <NUM> results in multiple possible combinations of sets of fiber angles and ply thicknesses that may provide the desired stiffnesses in each zone. These possible combinations are subsequently refined and filtered in order to optimize lamination properties for each of the zones.

Thus, at <NUM>, the results of the ply thickness integer optimization performed at step <NUM> are filtered. Filtering the results at <NUM> determines the integer number of plies that will optimize the desired in-plane laminate properties, and results in multiple possible solutions. This filtering the results of step <NUM>, i.e. the optimization process, involves filtering a number of possible optimized solutions based on an allowed deviation of effective laminate properties from a desired set of laminate properties, and results in multiple candidate fiber angle combinations and ply counts for each of the angle in each of the layup zones <NUM>-<NUM> (<FIG>) of the part 30a. This filtering process results in the selection of a laminate design that best reduces wrinkles, matches given laminate stiffness, and minimizes the number of ply sequences.

Steps <NUM>, <NUM> and <NUM> result in multiple candidate fiber angle combinations and ply counts for each of these angles for each of the layup zones. At <NUM>, layup information is generated, which may include determining the ply shapes and a stacking sequence that conform to a desired set of stacking sequence and manufacturability rules. Stacking sequence rules avoid undesirable laminate modes. The stacking sequence is chosen, at least in part to achieve substantially homogeneous bending stiffness properties in the laminate. When ply spices are required, naturally created splices are preferred which can be achieved by overlapping the ends of medium length plies. The use of natural splices improves layup efficiency by avoiding the need for short plies required to reinforce splices between long plies, while maintaining structural integrity. Also, in determining the ply shapes, the plies should be continuous wherever possible in order to maximize the transfer loads from one zone to another, as well as to optimize layup efficiency. At <NUM>, the flat stack of plies <NUM> is laid up based on the layup information generated at <NUM>. Then, at <NUM>, the flat stack <NUM> is formed, as by stamp forming, into a straight part <NUM> having a desired cross-sectional shape, such a hat or other shape. At <NUM>, the laminate part 30a is then formed to a desired contour along its major axis of loading. Finally, at <NUM> the fully formed laminate part 30a is cured.

Attention is now directed to <FIG> which broadly illustrates the steps of a method of producing a composite laminate part <NUM> of a new, rather than an existing design. The process for producing a newly designed laminate part <NUM> with reduced wrinkling is similar to that previously described with reference to <FIG> but without the need for matching the stiffness of an existing part. Briefly, a determination is made of how many plies of each selected ply orientation are required to satisfy specifications for the new part, followed by an optimization of the ply shapes and stacking sequence.

Thus, referring particularly to <FIG>, a new part is selected at <NUM>, and at <NUM> the material, fiber orientations and structural size of the part are chosen. In performing step <NUM>, the zones <NUM> of the part <NUM> are defined, and the number of plies per orientation in each zone is determined. Next, at <NUM>, layup information is generated, which comprises determining the ply shapes and a stacking sequence that conforms to a desired set of stacking sequence and manufacturability rules. Then, at <NUM>, the flat laminate stack is laid up, following which at <NUM>, the flat laminate stack is formed into a straight part having a desired cross-sectional shape, as by stamp forming or other processes previously described. At <NUM>, laminate part is formed to the desired longitudinal contour and is thereafter cured at <NUM>. As previously mentioned, steps <NUM> and <NUM> may be simultaneously performed where the forming is carried out in a single operation.

Embodiments of the disclosure may find use in a variety of potential applications, particularly in the transportation industry, including for example, aerospace, marine, automotive applications and other application where contoured composite laminate structural members may be used. Thus, referring now to <FIG>, embodiments of the disclosure may be used in the context of an aircraft manufacturing and service method <NUM> as shown in <FIG> and an aircraft <NUM> as shown in <FIG>. Aircraft applications of the disclosed embodiments may include, for example, without limitation, spars, stringers, beams and similar structural members that are contoured along a major axis of loading. During pre-production, exemplary method <NUM> may include specification and design <NUM> of the aircraft <NUM> and material procurement <NUM>. During production, component and subassembly manufacturing <NUM> and system integration 128of the aircraft <NUM> takes place. Thereafter, the aircraft <NUM> may go through certification and delivery <NUM> in order to be placed in service <NUM>. While in service by a customer, the aircraft <NUM> is scheduled for routine maintenance and service <NUM>, which may also include modification, reconfiguration, refurbishment, and so on.

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>, the aircraft <NUM> produced by exemplary method <NUM> may include an airframe <NUM> with a plurality of systems <NUM> and an interior <NUM>. The airframe <NUM> may include spars, stringers, beams and similar structural members <NUM> having one or more contours. Examples of high-level systems <NUM> include one or more of a propulsion system <NUM> an electrical system <NUM> a hydraulic system <NUM> and an environmental system <NUM>. Any number of other systems may be included. Although an aerospace example is shown, the principles of the disclosure may be applied to other industries, such as the marine and automotive industries.

Systems and methods embodied herein may be employed during any one or more of the stages of the production and service method <NUM>. For example, components or subassemblies corresponding to production process <NUM> may be fabricated or manufactured in a manner similar to components or subassemblies produced while the aircraft <NUM> is in service. Also, one or more apparatus embodiments, method embodiments, or a combination thereof may be utilized during the production stages <NUM> and <NUM>, for example, by substantially expediting assembly of or reducing the cost of an aircraft <NUM>. Similarly, one or more of apparatus embodiments, method embodiments, or a combination thereof may be utilized while the aircraft <NUM> is in service, for example and without limitation, to maintenance and service <NUM>.

Claim 1:
A method of making a contoured composite laminate part (<NUM>) having a high aspect ratio, a major axis of loading (<NUM>) and a plurality of zones (<NUM>) along its length (L) respectively having desired stiffnesses, comprising:
selecting a set of fiber angles (±θ<NUM>, ±θ<NUM>, ±θ<NUM>) for plies (<NUM>) of unidirectional reinforcing fibers (<NUM>), wherein none of the fiber angles is <NUM>° relative to the major axis of loading (<NUM>), wherein the set of fiber angles includes three fiber orientations at off-angles (±θ<NUM>, ±θ<NUM>, ±θ<NUM>) relative to the major axis of loading of θ<NUM>, θ<NUM>, θ<NUM>, where <NUM> < θ<NUM> < θ<NUM> < θ<NUM> ≤ <NUM>°, wherein θ<NUM> - θ<NUM> ≤ <NUM>°, and θ<NUM> - Θ<NUM> ≤ <NUM>°;
determining, for each of the fiber angles (±θ<NUM>, ±θ<NUM>, ±θ<NUM>), a number of plies (<NUM>) in each of the zones (<NUM>) required to provide a desired set of in-plane laminate properties in the zone (<NUM>);
determining a shape and stacking sequence of the plies (<NUM>), wherein some of the plies (<NUM>) are continuous and other of the plies are discontinuous along the major axis of loading (<NUM>), and wherein the off-angles (±θ<NUM>, ±θ<NUM>, ±θ<NUM>) include balanced pairs of + and - angles;
laying up the plies (<NUM>) into a flat stack (<NUM>) using the stacking sequence; and
forming the flat stack (<NUM>) into the shape of the contoured composite laminate part (<NUM>).