High Speed Aircraft

Global-Local-Detail analysis

The figure represents the HyperSizer Global-Local-Detail process that very accurately determines ply-by-ply stresses throughout a panel’s cross section. Starting with a vehicle FEM, an arbitrary location on the HyperSizer generated transparent graphic (a), is identified with surface skins as being Tee shaped stiffened panels (b), which are separated by an unstiffened web. Both are modeled in the global loads FEM with a single plane of elements. Note the mesh refinement does not have to align with the stiffener spacing and the user can construct the mesh with as few elements as appropriate to get overall running loads in the skins (b). The image depicts a 6 x 4 element mesh per panel bay, but for this specific model, only one element was needed to span the substructure, full depth webs.

Each panel bay can be modeled with a single finite element because for any general, uniformly applied edge forces or moments including out-of-plane surface pressure, HyperSizer can compute the resulting local panel deformation as portrayed at (c). This includes thermoelastic deformations caused by in-plane and out-of-plane temperature gradients.

HyperSizer automatically couples to FEA codes, and is FEA code independent. However, NASTRAN is the FEA solver most used by our customers and HyperSizer is certified to support MSC, NX, and NEi versions. In essence, the NASTRAN file format is a defacto standard. After HyperSizer has optimized the design of the vehicle, each user identified component will have its generalized temperature dependent stiffness terms updated in the FEM. HyperSizer accomplishes this by regenerating NASTRAN PSHELL and MAT2 data types for shell elements and PBAR and MAT1 for beams. With the new material and design data, another FEA is submitted for the next round on internal loads that capture changed load paths. At this point HyperSizer reads the new computed element forces from the FEA output file. In this manner, the optimizer can evaluate any stiffened panel cross sectional shape without having to remesh the model. Trades between honeycomb sandwich, blade stiffened, and/or hat stiffened panels are lightning fast.

There is no limit to the number of FEM elements, grids, or load cases, permitting HyperSizer to rapidly handle large FEMs. HyperSizer has a linear relationship between run times and model size, not exponential which can become detrimental when going from demonstration to full production FEMs.

HyperSizer can analyze and optimize all structural components of entire airframes to thousands of load cases. Statistical post processing of the FEA computed element forces provide appropriate design-to loads. These loads are used for panel buckling and beamcolumn type failure analyses and are further resolved into individual panel segment forces (d) for other instability analyses such as local buckling and crippling, and then even further for concentrated stresses/strains.

Specializing in composite analyses and optimization, HyperSizer’s progressive Global-Local-Detail process of computing stresses and strains allows hundreds of different failure analyses to be included. Material strength failure predictions for the laminates include the panel span segments (e,left image) and the bonded joint between skin and flange of a stiffened panel (e,right image).

Interlaminar shear and peel stress variation is computed in the adhesive for linear and five different non-linear material methods. The Z axis stress variation is also computed throughout the laminate depth, and also for each individual ply as required for the last ply of a stepped joint, (e,right). The number of integration points and characteristic distance for failure prediction can be selected by user.

In addition to material strength based on damage initiation, damage tolerance residual strength of strain energy release rates (SERR) are computed using a rapid, non-FEA, virtual crack closure technique (VCCT). These values are compared to critical energy release rates GIc and GIIc to predict delamination propagation for a crack between laminate plies and/or a crack between the skin and bonded flange.