Abstract: A computer is configured to enable a rapid, consistent, ply-by-ply, quantitative analytical assessment of various Finite Element Method (FEM) material models based on metrics defined for impact damage. Additionally, the computer is configured to provide a method for determining the accuracy of such FEM material model(s) by comparing the output of those models to non-destructive evaluation (NDE) test data.

Abstract: A system comprises a model generation system and an analyzer. The model generation system is configured to modify a model of a laminated composite structure to meet a user defined objective and to meet at least one of user defined performance constraints or user defined manufacturing constraints by changing at least one ply characteristic of at least one ply of the model while maintaining ply boundary geometry constraints for each ply of the model. The ply boundary geometry constraints of the model generation system include a number of defined ply boundary geometries each described by a respective mathematical function. The analyzer is configured to return objective values or constraint function values for at least one ply of the model.

Abstract: A computer is configured to enable a rapid, consistent, ply-by-ply, quantitative analytical assessment of various Finite Element Method (FEM) material models based on metrics defined for impact damage. Additionally, the computer is configured to provide a method for determining the accuracy of such FEM material model(s) by comparing the output of those models to non-destructive evaluation (NDE) test data.

Abstract: A composite structure, such as a laminated composite panel, for example, comprises one or more layers or “plies” embedded in a matrix material or otherwise fixed together in an arrangement, commonly referred to as a “stack up.” Each material in the structure has a corresponding material failure model (MFM) defining the physical characteristics of that material. A ballistic threshold velocity computing device obtains the MFMs for each material in the composite structure, generates a predicted ballistic velocity threshold velocity for each MFM, and then generates a parametric model to compute a composite ballistic velocity threshold velocity for the composite structure.

Abstract: A system for fabrication of an aerospace structure incorporates a mold having a surface and at least one unidirectional SFL head adapted to lay down a plurality of collimated tows in a predetermined laminated pattern on the mold surface to produce a fuselage skin. At least one cross plied laminate SFL head is adapted to lay down a cross plied laminate base interface on the fuselage skin to establish a lattice rib shape for each of a plurality of lattice ribs. The cross plied laminate SFL head has a band placement head steerable to avoid structural design features and to maintain spacing from adjacent steered lattice ribs. The unidirectional SFL head is further adapted to lay down a plurality of collimated tows on the base interface of each of the plurality of lattice ribs for a first plurality of unidirectional tow plies in each lattice rib.

Abstract: A composite structure, such as a laminated composite panel, for example, comprises one or more layers or “plies” embedded in a matrix material or otherwise fixed together in an arrangement, commonly referred to as a “stack up.” Each material in the structure has a corresponding material failure model (MFM) defining the physical characteristics of that material. A ballistic threshold velocity computing device obtains the MFMs for each material in the composite structure, generates a predicted ballistic velocity threshold velocity for each MFM, and then generates a parametric model to compute a composite ballistic velocity threshold velocity for the composite structure.

Abstract: A method of analyzing the durability of a structure. Load-controlled testing is performed on samples of a composite material of the structure to relate critical strain invariants of the material to cyclic rates of strain invariant accumulation and frequencies associated with the cyclic rates. The material is characterized based on effective properties of the material, including the cyclic rates of strain invariant accumulation. Laminate properties and a geometrical definition of the structure are used to obtain a parametric model. Material characterizations are used to determine model element frequency responses to applied load conditions. Each element's frequency responses and critical strain invariants are used to determine whether damage is indicated at the element. Progression of damage is tracked and accounted for in the model.

Abstract: Composite skin-stringer structures which reduce or eliminate the risk of delamination at the skin-stringer interface. This can be accomplished by arranging ply directions (i.e., the angles of the fiber paths of the ply) in a layup in a way such that for the dominant loading, the skin and stringer will each deform in a way that reduces relative opening (fracture Mode I) and/or sliding (fracture Mode II) and/or scissoring (fracture Mode III) at the skin-stringer interface. This is possible when coupling between specific deformations modes is purposefully activated instead of being suppressed. The ply directions in the stringer are adjusted so that the stringer deforms in a controlled fashion to suppress or “close” cracks that are about to form—before the undesirable modes of failure form—as load is applied.

Abstract: A design process that uses lamination parameter inversion to generate a set of baseline layups having desired stiffness properties. Then the underdetermined Newton's method can be applied to explore solution manifolds describing alternative designs having similar if not identical stiffness properties. The manifold of solutions can be methodically examined to find those with desirable properties. Desirable properties include those that have been traditionally captured by design rules or those that improve manufacturability. Combining lamination parameters as design variables with lamination parameter inversion provides an efficient optimization process for non-traditional laminates.

Abstract: A method of generating an optimized design model for a composite laminate may include computing a normalized set of lamination parameters and laminate stiffness matrices of an initial laminate design, and determining, using an optimizer operating on a finite element model, optimum values for the lamination parameters and the laminate thickness. The method may further include adjusting the optimum value of the laminate thickness, and performing an inversion process extracting multiple solutions from the lamination parameters, each solution including a unique set of individual fiber angles for each ply and representing an optimized design model of the composite laminate.

Abstract: A system for fabrication of an aerospace structure incorporates a mold having a surface and at least one unidirectional SFL head adapted to lay down a plurality of collimated tows in a predetermined laminated pattern on the mold surface to produce a fuselage skin. At least one cross plied laminate SFL head is adapted to lay down a cross plied laminate base interface on the fuselage skin to establish a lattice rib shape for each of a plurality of lattice ribs. The cross plied laminate SFL head has a band placement head steerable to avoid structural design features and to maintain spacing from adjacent steered lattice ribs. The unidirectional SFL head is further adapted to lay down a plurality of collimated tows on the base interface of each of the plurality of lattice ribs for a first plurality of unidirectional tow plies in each lattice rib.

Abstract: A method of predicting the strength characteristics of a composite laminate may include loading a structural model of a composite laminate formed of a material system. The method may additionally include comparing strain invariants from loading the composite laminate to critical strain invariant values of the material system. The method may also include identifying as a first significant event (FSE) a strain invariant of the matrix and/or the fibers reaching a critical strain invariant value.

Abstract: A system comprises a model generation system and an analyzer. The model generation system is configured to modify a model of a laminated composite structure to meet a user defined objective and to meet at least one of user defined performance constraints or user defined manufacturing constraints by changing at least one ply characteristic of at least one ply of the model while maintaining ply boundary geometry constraints for each ply of the model. The ply boundary geometry constraints of the model generation system include a number of defined ply boundary geometries each described by a respective mathematical function. The analyzer is configured to return objective values or constraint function values for at least one ply of the model.

Abstract: A system for fabrication of an aerospace structure incorporates a mold having a surface and at least one unidirectional SFL head adapted to lay down a plurality of collimated tows in a predetermined laminated pattern on the mold surface to produce a fuselage skin. At least one cross plied laminate SFL head is adapted to lay down a cross plied laminate base interface on the fuselage skin to establish a lattice rib shape for each of a plurality of lattice ribs. The cross plied laminate SFL head has a band placement head steerable to avoid structural design features and to maintain spacing from adjacent steered lattice ribs. The unidirectional SFL head is further adapted to lay down a plurality of collimated tows on the base interface of each of the plurality of lattice ribs for a first plurality of unidirectional tow plies in each lattice rib.

Abstract: A system and method for creating an optimized composite laminate structure containing a plurality of plies. The system has a processor and a memory, including an application interface. The application interface, when executed by the processor, is configured to operably: receive an input file having one or more of a maximum number of plies, design variables, material properties, and design constraints; determine an initial layup sequence defining parameters of a fiber orientation angle for each ply, and a total percentage of plies at a given fiber orientation angle; iteratively adjust the parameters, until an optimum set of parameters is obtained that achieves one or more predetermined margins of safety, and that achieves optimization of the composite laminate structure; and generate an output file for creating a layup, according to the parameters. The system further has a layup system for creating the optimized composite laminate structure.

Abstract: Methods and systems for reducing finite element simulation time for acoustic response analysis are disclosed. In one embodiment, a method includes analytically creating a finite element model, the finite element model including a plurality of subdivisions. A plurality of cross-correlations between respective pairings of the subdivisions is then specified. A portion of the cross-correlations are then eliminated to provide a reduced set of cross-correlations between respective pairings of the subdivisions. The elimination includes determining a spatial distance value between at least two subdivisions, and discarding at least one of the cross-correlations for which the spatial distance value is greater than a specified threshold value. The finite element simulation is then performed using the reduced set of cross-correlations.

Abstract: A design process that uses lamination parameter inversion to generate a set of baseline layups having desired stiffness properties. Then the underdetermined Newton's method can be applied to explore solution manifolds describing alternative designs having similar if not identical stiffness properties. The manifold of solutions can be methodically examined to find those with desirable properties. Desirable properties include those that have been traditionally captured by design rules or those that improve manufacturability. Combining lamination parameters as design variables with lamination parameter inversion provides an efficient optimization process for non-traditional laminates.

Abstract: A method of predicting the strength characteristics of a composite laminate may include loading a structural model of a composite laminate formed of a material system. The method may additionally include comparing strain invariants from loading the composite laminate to critical strain invariant values of the material system. The method may also include identifying as a first significant event (FSE) a strain invariant of the matrix and/or the fibers reaching a critical strain invariant value.

Abstract: Composite skin-stringer structures which reduce or eliminate the risk of delamination at the skin-stringer interface. This can be accomplished by arranging ply directions (i.e., the angles of the fiber paths of the ply) in a layup in a way such that for the dominant loading, the skin and stringer will each deform in a way that reduces relative opening (fracture Mode I) and/or sliding (fracture Mode II) and/or scissoring (fracture Mode III) at the skin-stringer interface. This is possible when coupling between specific deformations modes is purposefully activated instead of being suppressed. The ply directions in the stringer are adjusted so that the stringer deforms in a controlled fashion to suppress or “close” cracks that are about to form—before the undesirable modes of failure form—as load is applied.

Abstract: A system and method for creating an optimized composite laminate structure containing a plurality of plies. The system has a processor and a memory, including an application interface. The application interface, when executed by the processor, is configured to operably: receive an input file having one or more of a maximum number of plies, design variables, material properties, and design constraints; determine an initial layup sequence defining parameters of a fiber orientation angle for each ply, and a total percentage of plies at a given fiber orientation angle; iteratively adjust the parameters, until an optimum set of parameters is obtained that achieves one or more predetermined margins of safety, and that achieves optimization of the composite laminate structure; and generate an output file for creating a layup, according to the parameters. The system further has a layup system for creating the optimized composite laminate structure.