Predicting composite laminated part wripping

Abstract

The invention relates to predicting composite laminate part corrugation using a surrogate model. A composite laminated part, such as a spar with reduced ply wrinkling, is designed by generating a proxy model of the part and performing geodesic strain analysis of the proxy model. The result of the analysis is used to modify the design of the part to reduce strain in areas of the part that may cause ply wrinkling.

Description

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS

[0002] This application claims the benefit of U.S. Provisional Patent Application No. 63 / 590,095, filed on October 13, 2023, which is hereby incorporated by reference in its entirety. Technical Field

[0003] The present disclosure relates generally to the fabrication of composite laminate parts, and more particularly to a method of predicting ply wrinkling during part formation. The present disclosure also relates to a method of designing and fabricating the part based on wrinkling prediction using a surrogate model. Background Art

[0004] The production of large composite laminate parts is sometimes done by forming a desired part shape from a stack of composite plies. The tools used to form the part often contain global and local curvatures, ply bevels and flange angles, and various other features that can cause local strains that lead to undesirable ply wrinkling as the material is stacked to form the shape. One approach to solving the ply wrinkling problem requires that the part be completely designed, reviewed, manufactured, and tested to determine whether specific geometric features of the part will cause ply wrinkling that exceeds the allowed tolerances. In another approach, a full factorial design of experiments is implemented to determine how geometric features interact to cause ply wrinkling, however this approach is costly because it requires a large number of design / build / analysis cycles. In yet another approach, general design rules for the part are written based on perceived sensitivity that assumes the worst-case strain situation, often resulting in the part being overdesigned and therefore more expensive and / or heavier than desired.

[0005] It would therefore be desirable to provide a more efficient method of designing composite laminate parts that identifies part design features that adversely affect part manufacturability early in the design process, before a fully integrated design is completed. Summary of the invention

[0006] The present disclosure relates generally to designing and manufacturing composite laminate parts, and more particularly to a part design method that reduces or eliminates ply wrinkling during part formation. The method allows part features to be adjusted early in the design process, before a fully integrated model is completed. During this early design stage, the part configuration is evaluated to predict ply wrinkling. Ply wrinkling of the part configuration is predicted by performing a geodesic strain analysis of a parametric model of the part to identify part features that cause stress concentrations, which can lead to ply wrinkling. The method generates a parameterization of the part shape through design features, allowing a large number of representative spar shapes to be generated, and the geodesic strains of each spar shape can be measured.

[0007] The parameterization is a mapping from part features to representative part geometry, which allows simulation DOEs (design of experiments) to be run on part shapes. The geodesic strain can then be fitted to a proxy model that associates part features with geodesic strain. The output of this fitting problem is a model that allows designers to evaluate the geodesic strain of any set of spar features without the need to build a CAD model. Therefore, the proxy model is used to predict geodesic strains for multi-spar designs. By allowing early evaluation of design features in this way, the design can be easily adjusted outside the CAD (computer aided design) environment to reduce or avoid stress concentrations, thereby eliminating the need and errors for long design iterations or iterative manufacturing trials. In one instance, the method performs geodesic strain analysis based on a parametric model of a part using a tabular design feature geometry. In another instance, the method calculates geodesic strains based on CAD geometry using a machine learning model.

[0008] According to one aspect, a method for designing a composite laminate part with reduced ply wrinkling is provided. The method includes generating a plurality of geometric models of the part, and performing a geodesic strain analysis of each geometric model. The strain analysis results in generating strain data for each geometric model. The method further includes generating a proxy model that associates the geometric model with the strain data. The proxy model is used to design the part with minimized strain.

[0009] According to yet another aspect, a method for designing a composite laminate spar having reduced ply wrinkling is provided. The method comprises generating a plurality of parametric models of the spar and performing a geodesic strain analysis of each parametric model. The strain analysis results in a set of strain data for each parametric model. A proxy model is generated by fitting the strain data to the parametric model. The proxy model is used to iteratively modify the design of the spar until the strain on the spar is substantially reduced.

[0010] According to yet a further aspect, there is provided a method for generating a proxy model for designing a composite laminate part having reduced ply wrinkling. The method comprises generating a plurality of geometric models of the part, and performing a geodesic strain analysis of each geometric model. The strain analysis generates strain data for each geometric model. Each geometric model is fitted to the strain data, resulting in a proxy model that can be used to iteratively modify a spar design until the strain on the spar is minimized.

[0011] The features, functions, and advantages of various examples of the present disclosure may be achieved independently or may also be combined in yet other examples in which further details may be apparent with reference to the following description and drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0012] Figure 1is an illustration of a perspective view of an aircraft with the wing box segments shown in dashed lines.

[0013] Figure 2 yes Figure 1 Illustration of a perspective view of a wing box section is shown with portions of the upper wing skin cut away to reveal several wing spars.

[0014] Figure 3 is an illustration of a perspective view of a simple wing spar.

[0015] Figure 4 yes Figure 3 Specified as " Figure 4 ” area, showing the individual plies of the laminate.

[0016] Figure 5 is an illustration of a portion of a composite laminate spar having a ramp component.

[0017] Figure 6 is a functional block diagram of one embodiment of a method for designing a composite laminate part with reduced ply wrinkling.

[0018] Figure 7 is a graphic illustration of a table of spar data.

[0019] Figures 8 to 10 is an illustration of a graphical flow chart showing how to generate geodesic strain data based on a geodesic analysis of a parametric model.

[0020] Fig.11 is an illustration of a diagrammatic perspective diagrammatically representing the concentration of compression and tension in a composite laminate spar.

[0021] Fig.12 is an illustration of a flow chart of another method for designing a composite laminate part with reduced ply wrinkling using CAD geometry in accordance with another embodiment of the method.

[0022] Fig.13 A diagram that combines diagrams and flow charts showing how to use gage maps to design composite laminate parts and calculate strains in a wing spar.

[0023] Fig.14 is an illustration of a functional block diagram of a computer-based system for implementing the disclosed method.

[0024] Fig.15 is an illustration of a flow chart of a method for designing a composite laminate part having reduced ply wrinkling.

[0025] Fig.16 is an illustration of a flow chart of a method for designing a composite laminate spar with reduced ply wrinkling.

[0026] Fig.17 is an illustration of a flow chart of a method for generating a proxy model for designing a composite laminate part with reduced ply wrinkling.

[0027] Fig.18 It is an illustration of a flow chart of aircraft production and maintenance methodology.

[0028] Fig.19 is an illustration of a block diagram of an aircraft. DETAILED DESCRIPTION

[0029] The disclosed examples described below relate to a method for predicting wrinkling in a composite laminate part, where wrinkling is due to geometric part design features that cause local strains to be generated when the part is formed. The terms "strain", "local strain" and "excessive strain" as used herein refer to strains generated during the formation of the charge or stack of composite plies, which result in undesirable ply wrinkling levels or other unacceptable non-fit. By predicting wrinkling before creating a fully integrated part design, the geometric features of the part can be changed in a way that reduces or eliminates these strains during formation. The method includes establishing a proxy model that operates on part shape geometry or geometric features in order to quickly evaluate the geodesic strains of the part. The evaluation of geodesic strains allows ply wrinkling predictions to be made based on a tabular part definition file or a detailed CAD model itself, which typically forms the basis of a detailed CAD model.

[0030] In one mode of the method, a parametric model is created using tabular data that establishes representative part geometry used in geodesic analysis of the fitted surrogate model. This enables geodesic strain analysis to be performed before a detailed CAD model is generated, allowing part manufacturability analysis to be performed early in the part design process. In another mode of the method, the geodesic analysis results are calculated based on the full CAD geometry using a machine learning (ML) model such as DiffusionNet.

[0031] As will be discussed later, the disclosed method can be used to design any kind of part that is produced by forming a charge (stack) of composite plies 44 of a material such as CFRP (carbon fiber reinforced plastic). Figure 1 and Figure 2, various composite laminate parts used in the aircraft 20 include a fuselage 22, wings 24, and an empennage 26. Each wing 24 includes a wing box section 28. The wing box section 28 includes a grid of structural members including a plurality of composite laminate spars 30 extending from a root 32 to a tip 34 of the wing 24, as well as stringers (not shown) and cross beams or ribs 38. These structural members transfer loads applied to the wing skin 36 during flight. Although not shown in the drawings, the empennage 26 may also have a grid of structural members including composite laminate spars, similar to the wing box section 28. In the following description, the spars 30 are used merely as an example of how to use the method of the present disclosure to design a composite laminate part that is susceptible to strains that may cause ply wrinkling, however the method may be used to design various other types of composite laminate parts used in an aircraft, such as ground beams, stringers, and frames, as well as the composite laminate spars forming part of the empennage 26. Furthermore, it should be noted herein that although an aircraft application is described, the disclosed method may be used to design composite laminate parts for a wide range of applications other than aircraft.

[0032] Now also refer to Figure 3 and Figure 4 , the spar 30 may have any type of shape, and in the example shown, the cross-section is C-shaped. The C-shape is formed by a pair of flanges 40 that transition into a web 42 at radius corners R1, R2. The flanges 40 have desired widths W1, W2, while the web 42 has a desired width W3. The spar 30 is shown to have a constant cross-section along its length, however, in many applications the spar cross-section may not be constant. In fact, the spar 30 may have various features and changes along its length, such as pads, profiles, steps, cuts, and the like, many of which may appear simultaneously (e.g., in the same area) or overlap each other. For example, Figure 5 A portion of a C-shaped spar 30 is shown having bevel segments 45 along its length. The spar 30 may be produced by any of several forming techniques, such as compression forming or corrugation forming, in which a flat stack of composite plies 44 having different fiber orientations is formed into a desired shape by a mold (not shown). As will be discussed below, these features and variations in spar geometry may cause areas of the spar 30 to be placed in tension or compression during forming, creating strains severe enough to cause unacceptable levels of ply wrinkling or other non-conformity.

[0033] Follow now Figure 6 and Figure 7, which can be used to explain a first mode of implementing the method, which allows for rapid assessment of geodesic strains in a spar 30 that may cause ply wrinkling or other lamination mismatches during the forming process. The ability to predict ply wrinkling allows the designer to change spar design features early in the design process in order to achieve a final design that minimizes localized excessive strains that may cause ply wrinkling. The method includes two stages, the first stage is shown at 48 and 50, and the second stage is shown at 52. In the first step 48, a table 54 of data 56 is collected, which represents the geometry of the spar 30 (or other part). The data 56 is arranged in rows 58 and columns 60. The columns 60 each represent a set of refined geometric features or geometries of the spar 30. The spar features used to populate the table 54 include those features that are most likely to produce strain concentrations during the laminate forming process that may cause wrinkling. These features may generally include, but are not limited to, ramp rate, ramp height, local ply drop, flange angle, flange length, gage, flange tension state, radius design, global and local curvature, and ply arrangement (number, thickness, and orientation of plies).

[0034] The data in row 58 represents the value of each feature or geometry represented by column 60 all along the length of the spar 30. Table 54 is used to create a parametric model of the spar 30, which is then used to associate the spar data 56 with the geodesic strain analysis of each spar design. Parametric modeling is a computer-aided design technique that involves specifying parameters that define the geometry or features of a part, and establishing relationships between the parameters. When one or more parameters are changed, the parametric model of the part can be automatically modified or rebuilt. This modeling technique is implemented using parametric modeling software, which incorporates related parameters in a manner that allows the designer to change the model of the part by changing one or more parameters. Therefore, in this application, a parametric model is essentially a computer-implemented algorithm that parameterizes the spar design represented by table 54, allowing instantaneous changes in the spar design. The parameterization process can be implemented using commercially available computer software to define the size and shape of the spar design, which can be visualized in 3D if desired. The 3D spar model changes instantly when the value of the selected feature (spar data 56) is changed.

[0035] like Figure 6 As shown, a proxy model 50 is developed that includes a mathematical representation of a function that relates the spar data in the table 54 to the geodesic strain in the spar 30 defined by the table 54. As will be discussed in more detail below, to build the proxy model, geodesic strain data is developed for each of a plurality of tables 54 that represent different spar designs. In effect, the use of the proxy model 50 is a way to output geodesic strain information all along the length of the spar 30 using parameterized tables 54 of data 56 that define different designs of the spar 30.

[0036] exist Figure 6 In the second stage 52 of the method shown, the designer uses the proxy model 50 to make instantaneous changes to the design using an iterative technique, wherein the values ​​of the data in the table 54 are changed in order to arrive at design parameters that minimize the geodesic strains in the spar 30 during the forming process. The use of the proxy model 50 allows the designer to avoid long design cycles that would otherwise require intensive calculations to evaluate the geodesic strains in the spar 30. In effect, the method uses easy-to-setup geodesic strain measurements that allow compression and tension hot spots 65 ( Fig.11 ). Rapidly evaluating the strain concentrations of a proxy model of the spar 30 caused by specific geodesic features of the spar 30 allows the spar design to be adjusted before arriving at a fully integrated design.

[0037] Follow now Figures 8 to 10 , which generally illustrates the process of generating the data required to generate the proxy model 50. Geodesic strain data is generated for a plurality of different spar data tables 54, which are represented in the figure as T1, T2, T3, respectively. Although only three spar tables are shown, any number of spar tables can be used, each of which is sampled and used as input to generate strain data g1, g2, g3. Using tables T1, T2, T3 ... T n A parametric spar model 62, 62a, 62b is generated, each of which represents a specific spar design 64, 64a, 64b. Next, a geodesic strain analysis 66, 66a, 66b is performed on each spar design 64, 64a, 64b to generate geodesic strain data g1, g2, g3, ... g n The above process generates multiple data pairs (T1,g1), (T2,g2), (T3,g3)…(T n ,g n ). The surrogate model 50 is a mathematical object or function that fits this data; thus, the surrogate model 50 is the function F(T i) =g i The proxy model is a way of encoding the relationship between the tabular information that defines the spars and the geodesic strains on those spars. The geodesic strain data g1, g2, g3, ... g n In 3D, this can be visualized as strain “hot spots” 65 on the 3D spar model, as shown in Fig.11 As shown. Fig.11, these hot spots 65 are shown as areas on the spar 30 that are in tension 69 or compression 67. Strain can be inferred from the hot spots 65 because they can generate stresses that cause strain during the forming process. Therefore, ply wrinkling caused by excessive strain can be inferred from the hot spots 65. In the example shown, the hot spots 65 are represented by a gray scale, which can be associated with a scale 101 representing the degree of tension or compression, however, if desired, a color scale can be used in the case of a color-coded representation of the spar model.

[0038] Follow now Fig.12 , which may be used to illustrate a second alternative mode of implementing the disclosed method, which is to design a composite laminate part such as a spar 30 based on a complete rather than a refined set of part geometry features. While similar in other respects to the first mode of the method described above, this method does not rely on a fixed set of refined spar geometry features, thereby allowing it to be more flexible, but relies on a complete spar model. In this example, a machine learning (ML) application such as the DuffusionNet learning model is used to establish a proxy model 72 that can directly operate on the spar geometry. The proxy model 72 is a mathematical representation of a function that relates the CAD model of the spar 30 to the geodesic strain data of the CAD model. Using the ML application, a geodesic strain analysis 66 presents a complete design based on the CAD geometry, thereby accelerating the geodesic strain analysis calculations. Similar to the previously described example, a list of geometric features that fully define the spar 30 is generated, and a parametric spar model is generated based on these geometric features. DuffusionNet is then used to generate a proxy model 72 of the part, which is used to optimize the manufacturability of the part. Similar to the previously described examples, this example of the method allows for manufacturability analysis to be performed prior to generating a final, fully integrated CAD model, thereby accelerating the process of designing composite laminate parts optimized for manufacturability.

[0039] Use of the above method provides the designer with an understanding of how certain design features contribute to ply wrinkling, including specific areas of the part that may experience wrinkling during the laminate forming process. Based on the results of the geodesic strain analysis 66, the designer may change certain geometric features of the part to create a basic, prototype part geometry, which may then be used as a basis for designing a detailed, final part design without the use of CAD tools. In other words, a final or near-final part design may be completed relatively simply and quickly outside of a CAD environment. In effect, the method allows the geometric features of the model that affect ply wrinkling to be adjusted and selected before the actual CAD model is generated.

[0040] The disclosed method can be used with any type of computer-based system such as Fig.14The system 102 includes a computer 88, one or more input / output devices 90, and one or more displays 92 for displaying data and 3D models. The system 102 also includes application software programs 94, data storage in the form of memory 96, system programs 98, and CAD files 100. The geodesic strain analysis 66 can be performed using the system 102 running any type of commercially available software package. Depending on the results of the geodesic strain analysis 66, the designer can change any value in the table 54, automatically causing the parametric spar model 62 to be changed accordingly. The designer can repeat the above process, each time generating a revised spar design with a predictable result (strain), until a final design is achieved that eliminates unacceptable levels of strain. Therefore, the method allows for the rapid evaluation of strain concentrations in a spar configuration before a fully integrated design is completed, allowing the design to be easily changed "offline" outside the CAD environment, thereby avoiding long design iteration cycles.

[0041] Reference now Fig.13 As previously mentioned, using the above process, a preliminary version of the part geometry 76 of the spar 30 is designed that can be refined by modifying the specification diagram 78, which specifies the number of plies used in various components of the spar 30. In some cases, it is also necessary as part of this design process to consider and, if necessary, modify the interface part geometry that can cause local excessive strain. An example of an interface part geometry that affects strain is the slope of the skin ply inclination at the interface between the skin and the spar. The slope of the incline directly affects the specification diagram 78, which in turn affects the geodesic strain value. The specification diagram 78 can be modified by the designer to change the part interface geometry to reduce local excessive strain. For each set of values ​​in the specification diagram 78, the strain in the proxy spar model can be calculated to determine whether there are strain hot spots in the model. If such hot spots exist, the values ​​in the specification diagram 78 can be modified to reduce or eliminate the hot spots. The process of modifying the specification diagram 78 and then calculating the laminate strain 80 is iterated 84 until the calculated laminate strain 80 is reduced to a point indicating that ply wrinkling is minimized or eliminated during the forming process. This process results in a final, detailed spar design82.

[0042] Figures 15 to 17 The above-described method is generally illustrated and summarized. Fig.15Steps of a method for designing a composite laminate part that reduces possible ply wrinkling are shown and can be performed before a CAD model is completed. The method begins at 104, where multiple geometric models of a part are generated. As previously discussed, these geometric models can be represented as parametric models based on tabular data that define the geometric features of the part. At 106, a geodesic strain analysis is performed on each geometric model. This strain analysis results in the generation of strain data for each geometric model. At 108, a proxy model is generated that associates the geometric model with the strain data. Then at 110, the proxy model can be used to design a part with minimal excessive local strain.

[0043] Fig.16 Steps of a method for designing a composite laminate part with reduced ply wrinkling are shown. At 112, a plurality of parametric models of a spar are generated. Then, at 114, a geodesic strain analysis is performed on each parametric model, resulting in the generation of a set of strain data for each parametric model. At 116, a proxy model is generated by fitting the strain data to the parametric model. Then, at 118, the proxy model is used to iteratively modify the design of the spar until the strain on the spar is substantially reduced.

[0044] Fig.17 The steps of a method for generating a proxy model for designing a composite laminate part with reduced ply wrinkling are generally shown. At 120, a plurality of parametric models of a part are generated. At 122, a geodesic strain analysis is performed on each parametric model, resulting in the generation of strain data for each parametric model. Finally, at 124, each parametric model is fitted to the strain data, forming a proxy model that can be used to iteratively modify the spar design until excessive strain is minimized.

[0045] Examples of the present disclosure may find use in a variety of potential applications, particularly in the transportation industry, including, for example, aerospace, marine, automotive applications, and other applications using shaped composite laminate parts. Fig.18 and Fig.19 , examples of the present disclosure can be used as Fig.18 Aircraft manufacturing and service method 126 and the like are shown Fig.19 128 is shown in the context of an aircraft 128. Aircraft applications of examples of the present disclosure may include various composite laminate parts, such as wing spar forming parts of a fuselage 128 of an aircraft. During pre-production, the method 126 may include specification and design 130 of the aircraft 128 and material procurement 132. During production, component and assembly manufacturing 134 and system integration 136 of the aircraft occur. Thereafter, the aircraft 128 may be certified and delivered 138 for entry into service 140. While in service with a customer, the aircraft 128 is scheduled for routine maintenance and overhaul 142, which may also include modification, reconfiguration, refurbishment, etc.

[0046] Each process of method 126 may be performed or executed by a system integrator, a third party, and / or an operator (e.g., a customer). For purposes of this description, a system integrator may include, but is not limited to, any number of aircraft manufacturers and major system subcontractors; a third party may include, but is not limited to, any number of vendors, subcontractors, and suppliers; and an operator may be an airline, a leasing company, the military, a service organization, and the like.

[0047] like Fig.19 As shown, an aircraft 128 produced by method 126 may include an airframe 144 and an interior 148 having a plurality of high-level systems 146. Examples of high-level systems 146 include one or more propulsion systems 150, electrical systems 152, hydraulic systems 154, and environmental systems 156. Any number of other systems may be included. Although an aerospace example is shown, the principles of the present disclosure may be applied to other industries, such as the marine and automotive industries.

[0048] The systems and methods included herein may be used in any one or more stages of the aircraft manufacturing and service method 126. For example, components or assemblies corresponding to component and assembly manufacturing 134 may be fabricated or manufactured in a manner similar to components or assemblies produced while in service with the aircraft 128. Furthermore, one or more equipment instances, method instances, or a combination thereof may be used during component and assembly manufacturing 134 and system integration 136, for example, by significantly speeding up assembly of the aircraft 128 or reducing the cost of the aircraft 128. Similarly, one or more equipment instances, method instances, or a combination thereof may be used while the aircraft 128 is in service, such as, but not limited to, maintenance and service 142.

[0049] Descriptions of different exemplary embodiments are presented for purposes of explanation and illustration, and are not intended to be exhaustive or limiting of the examples of the disclosed form. Many modifications and variations will be apparent to those of ordinary skill in the art. Further, different exemplary embodiments may provide different advantages over other exemplary embodiments. The selected example or examples are chosen and described in order to best explain the principles of the examples, practical applications, and to enable those of ordinary skill in the art to understand the disclosure of various examples with various modifications suitable for the particular use under consideration.

Claims

1. A method of designing a composite laminate part (30) having reduced ply wrinkling, comprising: generating a plurality of geometric models (62) of the part (30); Performing a geodesic strain analysis (66) of each of the geometric models (62), including generating strain data (g) of each of the geometric models (62); generating a proxy model (50, 72) that relates the geometric model (62) to the strain data (g); and The proxy model (50, 72) is used to design a part (30) with minimized strain.

2. The method of claim 1, wherein generating a plurality of geometric models (62) includes producing a parametric model (62) of the part (30) that allows features of the part (30) to be modified.

3. The method of claim 2, wherein generating the parametric model (62) includes creating a plurality of tables (54), each table containing data (56) representing values ​​of geometric features (62) of the part (30).

4. The method of claim 3, wherein generating the table (54) includes selecting geometric features (62) of the part that cause ply wrinkling.

5. The method of claim 1, wherein generating the proxy model (50, 72) comprises fitting the geometric model (62) with the strain data (g).

6. The method of claim 1, wherein using the proxy model (50, 72) includes iteratively (84) modifying a design of the part (30) until the strain is minimized.

7. The method of claim 1, wherein performing the geodesic strain analysis (66) and generating the proxy model (50, 72) is performed using a programmed computer (88).

8. The method of claim 1, further comprising generating a 3D representation of the part showing areas of stress concentration (65).

9. The method of claim 1, wherein generating a plurality of geometric models (62) of a part comprises generating a CAD model (70) of the part (30), and modifying the CAD model (70).

10. A method of designing a composite laminate spar (30) having reduced ply wrinkling, comprising: generating a multi-parameter model (62) of the spar (30); performing a geodesic strain analysis (66) of each of the parameter models (62), including generating a set of strain data (g) for each of the parameter models (62); generating a surrogate model by fitting the strain data to the parametric model; and The design of the spar (30) is iteratively modified (84) using the proxy model (50, 72) until the strain on the spar (30) is substantially reduced.

11. The method of claim 10, wherein generating a plurality of parametric models (62) includes selecting geometric features of the spar that may affect strain and assigning a value (56) to each of the geometric features.

12. The method of claim 11, wherein generating the plurality of parametric models (62) comprises arranging the geometric features and the values ​​(56) for each of the geometric features in a table (54), and converting the table (54) into a parametric model (62) that can be electronically modified.

13. The method of claim 10, further comprising generating a 3D representation representing the design of the spar (30), including displaying areas of stress concentration (65) on the spar (30).

14. The method of claim 10, wherein generating a plurality of parametric models (62) of a spar (30) comprises providing a CAD model (70) of the spar (30) and modifying the CAD model (70).

15. A method of generating a proxy model (50, 72) for designing a composite laminate part (30) having reduced ply wrinkling, comprising: generating a plurality of geometric models (62) of the part (30); performing a geodesic strain analysis (66) of each of said geometric models (52), including generating strain data (g) for each of said geometric models (62); and Each of the geometric models (62) is fitted to the strain data (g).

16. The method of claim 15, wherein generating a plurality of geometric models (62) comprises generating parametric models (62) defining geometric features of the part (30).

17. The method of claim 16, wherein generating the parametric model (62) includes selecting geometric features of the part (30) that may cause stress concentrations (65) in the part (30) and assigning a value (56) to each of the geometric features.

18. The method of claim 17, wherein generating the parametric model (62) comprises arranging the geometric features and the values ​​(56) in a table (54), wherein each of the geometric features is assigned a plurality of values ​​(56).

19. The method of claim 15, wherein generating a plurality of geometric models comprises providing a CAD model (70) of the part (30) and modifying the CAD model (70).

20. The method of claim 15, further comprising generating a 3D representation of each of the geometric models (62) showing stress concentrations (65) resulting from the geodesic strain analysis (66) of the geometric models (62).

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