AA SEZC method 4.1.6.2 measuring composite laminate strength

Abbott Aerospace – Analysis Method

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Taken from: Analysis and Design of Composite and Metallic Flight Vehicle Structures

Measuring Composite Laminate Strength

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The measured strength of composite laminate and how it is compared with failurecriteria is a bone of contention. A wide variety of opinion exists as to the best andmost accurate way to analyze composite structures.

The aim of any analysis is to identify the load at which the structure fails, the location of failure and the mode of failure.

There are two different approaches to fiber laminate stress analysis, one is to consider each ply within the laminate and develop a failure index (common methods are Tsai-Wu or Hill) for each ply within the laminate, this is called the lamina stress approach. The failure of a single ply causing a sudden internal load redistribution that will cause the laminate to fail








This is a stress based approach where the load applied to a particular laminate has to be distributed onto each ply across the laminate thickness taking the direction of load in relation to the ply orientations into account. This produces a plot of stresses on each ply through the thickness that looks like this:

-0.200

-0.150

-0.100

-0.050

0.000

0.050

0.100

0.150

0.200

-15000.0 -10000.0 -5000.0 0.0 5000.0 10000.0 15000.0 20000.0 25000.0

Thic

kness (

in)

Stress (psi)

Ply Total Stress

The following spreadsheet calculates the individual stress in each ply of a laminate and margins of safety based on a range of failure indices.

AA-SM-101-008 Lamina Stress Analysis




http://www.abbottaerospace.com/wpdm-package/aa-sm-101-008





A margin of safety can then be generated for each ply; the full Tsai-Wu failure criteriaformulation is as follows:

While this appears complex, multiple Tsai-Wu checks on each ply in the laminate, peranalysis location can easily be managed through spreadsheets or finite element postprocessors.

The ‘ply–by-ply’ stress failure approach is not commonly used in aircraft structure sizing and analysis. This is because it is difficult to correlate the actual failure of damaged composite laminates to an individual ply stress failure criterion.








The second analysis approach which is more commonly used for airframe primary structure is the laminate strain approach. In this method, the peak laminate strain is determined by analysis and compared with a strain allowable for a specific laminate configuration determined by test.

Classical laminate theory, the same theory that is used to distribute the loads andmoments into each ply for the ply-by-ply stress analysis approach, assumes that thestrain in the laminate maintains a smooth, linear distribution. There is a necessity forthe laminate to operate as a coherent cross-section when reacting load. A stepchange in strain at a point across the thickness implies an inter-laminar failure.








Most companies developing composite structure for primary airframe applications usethe strain approach. This is because testing can be done to develop laminate strainallowables that take account of manufacturing flaws and in-service damage. This isnecessary to show compliance with FAR 23.573 and 25.573.

In primary structure, the level of stress raiser that is used for composite analysis is dictated by the aircraft certification regulations and has been accepted as impact damage up to the 'BVID' (Barely Visible Damage) level. The certifiable strength of laminate for the primary structure is usually derived using 'compression after impact' testing.

The strength of any structure in a non-pristine condition is referred to as ‘Residual Strength’. Composite laminate structure is brittle compared to metallic structure (no plasticity) and therefore damage and flaws have to be accounted for in the static analysis in ways that are not considered for metal structure.








Residual Strength (expressed as failure strain) for Increasing Damage Level (NASA-CR-3767, 1984)




http://www.abbottaerospace.com/nasa-cr-3767-durability-and-damage-tolerance-of-large-composite-aircraft-primary-structure





This form of expressing strength with respect to damage level is commonly used and is a feature of modern FAA guidance material:

Schematic Diagram of Design Load Levels versus Categories of Damage Severity (AC20-107B, 2009)




http://www.abbottaerospace.com/wpdm-package/faa-ac-20-107b-composite-aircraft-structure

AA SEZC method 4.1.6.2 measuring composite laminate strength

Engineering

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