PMB Aerospace

Aerospace Structural Analysis Software


*** NEW!!!! *** Feb 2017

Web form for buckling analysis of stiffened plates.Click here.

*** NEW!!!! ***

Analysis of Buckling and Strength of perforated plates (with reinforcement).Click here.

*** NEW!!!! ***

Web access to 48 software!! Run online your jobs! No installation needed.Click here.

Nonuniform, nonhomogeneus composite plate analysis (with longitudinals). BETA version released. Learn more here.

Non linear version under development: Postbuckling analysis of composite shells

48 software

Composite and Metallic stiffened shell analysis

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The following are just relevant examples of how 48 software can be used in structural analysis and design.

They have been selected from those needs that the engineer has everyday in achieving efficient and economic designs with minimum effort. We expect to boost that ability.

If you want to try it by yourself access to the software online right now! The User Manual can be downloaded here.

Want to know more about 48 current capabilities? Write our engineering team to know how to apply the software to your designs.

Fuselage composite stiffened shell

This example is from an actual R&D test carried out during Posicoss development program. Initial buckling is calculated here using 48. We get excellent agreement with test figures. Initial buckling is of local type as can be seen in the modeshape #1 where stiffeners remain in place while skin buckles. Postbuckling analysis can be carried out to determine ultimate strength of the shell. We are working on a nonlinear version of the 48 software that will be able to obtain it directly.

Posicoss Panel P14 - composite stiffened shell

Initial buckling of a composite cilindrical shell (Length 660 x Width 419 x Radius 947 mm) stiffened with 4 stringers and loaded in axial compression. Skin layup is (90,+45,-45,0,0,-45,+45,90).

All other data can be checked in the input file below or in Ref. 1. Lower eigenvalue and mode shape are obtained for applied loads.

Calculated initial buckling load is 38.6 kN, and test buckling load reported is 38.9 kN.

The following picture shows the modeshape #1 corresponding to the lowest eigenvalue. The applied load is an axial flow of 178 N/mm applied on the curved edges.

Example 5 Mode Shape (created with SciLab)

Ref. 1: R. Zimmermann, H. Klein, A. Kling, "Buckling and postbuckling of stringer stiffened fibre composite curved panels - Tests and computations", Composite Structures, Volume 73, Issue 2, May 2006, Pages 150-161, ISSN 0263-8223, 10.1016/j.compstruct.2005.11.050. Read it at ScienceDirect

Here follow the input and output file that you can modify for your own use:
Example 5 Input File
Example 5 Output File

Use it!

Access to the software online right now! The User Manual can be downloaded here.

You can also try the unlimited version for a free period of 2 weeks.

Wing stiffened skin cover

This is the buckling analysis of a typical skin panel between two consecutive ribs and between front and rear spar.

General Aviation Aircraft metallic wing cover

x-axis is placed spanwise. Panel is modeled as flat due to slight curvature, there are 5 equal longitudinals (stringers) along x-axis and in this case the skin is aluminum of uniform thickness (1.2 mm).

Complicated in plane loads can be considered in these analyses. The axial flow Nx has been represented in the next figure on 3d colors and in contours for you to visualize the load distribution. There are as well Nx and Nxy flow, but are not shown here (they are an order of magnitude smaller than Nx).

In this case the internal loads were obtained using a General FEM (coarse) but any other method of obtaining the static prebuckling load distribution is applicable

NX flow

The resulting modeshape for lowest eigenvalue shows local buckling in the skin panel between rear spar and stringer #1. It is plotted in next figure.

mode 1 skin panel PL

In the following figure you can see the front view and identify that there is slight buckling of stringer #1 (yST1 = 164.9 mm, closer to rear spar) and despite the skin buckles clearly we have to conclude that panel would buckle globally and hence there is slight postbuckling capability. This can be fixed easily reinforcing the stringer (increase inertia).

The software provides the stringer deformation relative to maximum to assess stringer buckling vs. skin buckling.

mode 1 skin panel PL - front view

This analysis has been performed with an evolution of 48 software. It allows non-uniform skin panels (different thicknesses) . BETA version is available for evaluation on demand.

More examples

There are a lot of ways in which our software can support your designs.

Here you can find some of the more common applications, but the list is not exhaustive at all.

Let us now if you need engineering support to know how 48 can be applied to your design.

Example 1a (Niu, Airframe Stress Analysis and Sizing, p. 191)

Deflection and stresses under uniformly distributed lateral load of 10 psi on flat metallic plate.

Example 1a Input File

Example 1a Output File

Example 1b (Niu, Airframe Stress Analysis and Sizing, p. 191)

Deflection and stresses under point lateral load of 375 lbs. at the center of a flat metallic plate.

Example 1b Input File

Example 1b Output File

Example 2 (Wing Rib bay example)

Buckling of flate metallic plate under in-plane loads (Nx, Ny, Nxy).

Lower eigenvalue and mode shape are obtained for applied loads.

Example 2 Input File

Example 2 Output File

Example 2 Mode Shape 
                (created with SciLab)

Example 3 (Empennage Leading Edge Panel)

Static strength of a flat sandwich panel under in-plane loads.

Stresses and strains are obtained, reserve factor are calculated (maximum strain criterion).

Local sandwich failure modes are calculated (dimpling, crimping and wrinkling) and reserve factors obtained.

Example 3 Input File

Example 3 Output File

Buckling failure is also checked running a stability solution.

Example 3.1 Input File

Example 3.1 Output File

Example 4 (Wing Trailing Edge panel)

Deflection under lateral load (uniform distribution) of a sandwich plate.

Supported in 3 edges and free on the other, typical condition of a wing trailing edge panel. Aerodinamic pressure (0.00081 MPa = 0.1175 psi) and maximum deflection obtained (0.224 mm = 0.009 in.) are representative of this type of structure under cruise condition for a civil transport aircraft.

Example 4 Input File

Example 4 Output File

Example 4 Deflection (created with SciLab)

More examples will be published soon. Keep visiting this section, please!