A composite structure in an aircraft is stressed to limit or ult as the case calls for. Defects are introduced during manufacture and may be inflicted on the aircraft in service. The materials test programme will establish acceptable levels of damage viz-a-viz the far field allowable strain. Hole allowables will be established by a bypass tool possibly.
However on a recent project, a tool was developed to calculate progressive destruction of a laminate to its last ply failure. The method involved applying a load to break the first fiber and then the matrix which ever came first, and so on sequentially.
The applied load regime was discussed among the people specifying the tool. An expert in composite residual strength (actually the composites programme chief) on a previous programme, questioned what this tool was for. His position was that we comply with FAR 25, static limit, ult, DT and that is it. We don't look to see what the remaining strength is after we have passed the specified FAR load so why develop a residual damage tool that could cater for up to 60 plies and give the load associated with the last ply failure.
Am I missing something? Perhaps a single load path subject to some kind of degradation or ply reduction by abrasion, one could establish if we are still meeting the load bearing requirement. Perhaps if the structure is overbuilt, then we could allow say scrapes of say 25% of thickness if justified by a residual damage tolerance tool to last ply fail.9.
The only one I can recall is the case of a landing gear door bracket lug in composite that had this performed for inflicted ply damage. And I may be wrong about this latter exercise.
Could this be really to establish MRB guidelines?
Slightly Baffled by need for this tool.
Can anybody give a real life example?
Para g of this EASA ref gives something or suggests something.
www.easa.europa.eu
However on a recent project, a tool was developed to calculate progressive destruction of a laminate to its last ply failure. The method involved applying a load to break the first fiber and then the matrix which ever came first, and so on sequentially.
The applied load regime was discussed among the people specifying the tool. An expert in composite residual strength (actually the composites programme chief) on a previous programme, questioned what this tool was for. His position was that we comply with FAR 25, static limit, ult, DT and that is it. We don't look to see what the remaining strength is after we have passed the specified FAR load so why develop a residual damage tool that could cater for up to 60 plies and give the load associated with the last ply failure.
Am I missing something? Perhaps a single load path subject to some kind of degradation or ply reduction by abrasion, one could establish if we are still meeting the load bearing requirement. Perhaps if the structure is overbuilt, then we could allow say scrapes of say 25% of thickness if justified by a residual damage tolerance tool to last ply fail.9.
The only one I can recall is the case of a landing gear door bracket lug in composite that had this performed for inflicted ply damage. And I may be wrong about this latter exercise.
Could this be really to establish MRB guidelines?
Slightly Baffled by need for this tool.
Can anybody give a real life example?
Para g of this EASA ref gives something or suggests something.
Easy Access Rules for Acceptable Means of Compliance for Airworthiness of Products, Parts and Appliances (AMC-20) - Initial issue & Amendments 1 – 22 | EASA
EASA | European Union Aviation Safety Agency: The European Union Authority for aviation safety
www.easa.europa.eu
