Allowable Damage in Composite 1

[Fatstress]

Dear All,

Can anybody suggest me what would be the allowable damage accepted in aircraft composite (CFRP) structural component in production and in-service ? Or is there a thread about this already or a reference book that you can suggest to read? Thanks...

 

[SWComposites]

The allowable damage depends on many things: damage type, material, layup, structural configuration, damage assumed at design ultimate load, margin of safety in the part, etc. There is no single answer and it gets very complex. What is the specific basis for your question?

 

[Fatstress]

Hello SWComposite,

I am new in this subject, but I would like try to explain what I am looking for.
Well, as you know we have allowable damage limit for scratches and dents for aircraft metallic materials. Beyond this limit you have to contact the OEM for further instruction. The question, what is the most common in service damages in composite of aircraft structural components (damage due to impact, temperature, moisture, etc...?) and how to establish this allowable damage limit. I think, forinstance, Boeing has introduced what they called BVID, or Barely Visual Impact Damage to categorized the detectable damage size that the operator should consider. Do we simply simulate an impact damage on lots of specimens and test them and make a curve to show whether there will be a certain damage size that would be critical ? Or can we actually perform an analysis to show this to minimize the test program ?

FATSTRESS

 

[SWComposites]

Well, where to start -

There are a number of damage types for composite structures that need ADLs in the Structural Repair Manuals (SRMs):
scratches/nicks/gouge
dents
delaminations
cracks
holes/punctures
edge damage
lightning damage

Each of the above damage types need test data and/or validated analysis methods to determine the allowable damage sizes. Attempting to predict the strength reduction due to most damage types is difficult, at best, and still requires lots of data to validate the analysis method.

The ADL sizes are directly linked to the allowables used to size the structure and the margins at ultimate load. The ADL sizes are determined as the damage size that has a MS = 0 at ultimate load. If a part is designed to unnotched/undamaged allowables and the MS = 0, then the ADL sizes will be 0. On the other hand if the part is designed to notched/impact damaged allowables and the MS's are larger than 0, then the ADL sizes can be larger than the hole sizes or impact damage sizes corresponding to the allowables for ultimate load.

Barely visible impact damage refers to the minimum damage size that is considerd for ultimate strength sizing. ADLs for impact damage may be equal to or larger than BVID depending on the allowables and margins.

For more info on BVID, damage, ADLs, etc, see the following:

FAA Advisory Circular 20-107A (download at

http://www.faa.gov/certification/aircraft/index.cfm)

Mil-Handbook-17, Volume 3, Rev F, chapter on Damage Tolerance (purchase at

http://www.astm.org)

What specific composite structure and application are you involved with?

 

[Fatstress]

Hello SWComposites,

Thanks for the tips. Then except delamination, the ADLs for composite are the same as for metallic materials. And there is no easy way, we really have to work alot with the test department to determine the ADLs. One other question if I may, is it possible/common to repair a puncture in a composite skin panel (fuselage and moving surfaces) with a bolted aluminum (or Titanium) doubler as for metallic structure ? Or it is always preferable a bonded doubler repair ?

FATSTRESS

 

[SWComposites]

The damage types are more or less the same between composite and metal structure; the ADL sizes are NOT the same. Yes you need a lot of data to establish ADLs, unless the part in question is secondary/tertiary structure with high strength margins, in which case conservative ADLs can be set with engineering judgement.

Bolted repairs are used on composite structure. They are not generally prefered on honeycomb sandwich panels due to problems with sealing the repair to keop water out of the core (these repairs are typicaly only temporary repairs).

Bolted repairs to composites require a lot more design/analysis work to develop a structurally adequate repair compared to those for metal structure. You CANNOT JUST SLAP ON A DOUBLER TO A COMPOSITE PRIMARY STURCTURE PART (especialy a fuselage panel). The fastener load share adn joint strength in the composite is tricky to determine, and with a metal repair fatigue and damage tolerance issues have to be considered and addressed. Even design of a bonded repair to a highly loaded composite panel requires a significant amount of analysis and data.

Do you work for an aircraft OEM, part manufacturer or airline?

 

[ARDTL]

BVID is a genearic term that refers to a common characteristic of impact damage to CC or glass fibre composite structure. From a small impact point the structure will delaminate outwards in cone formation, hence the "barely visable" tag.

Depending upon the size of the damage and where it is will determine the repair design solution, I have seen damage to primary structure (wing spar)scarfed out, replaced with a bonded plug of the same material and backed up with a titanium doubler. Components and lighly loaded structure can be recovered with simple scarf repairs. i.e. panels u/c doors etc. Most of which should be in the aircraft SRM.


NDI is recommended for all such impact damage. This type of damage does not occur in metal to metal composites of the honycomb type.

 

[Fatstress]

ARDTL,

Do you have a limitation how deep you can scarf a component made of CC ? And do you need to implement a dedicated inspection program or can it be just fire and forget?

 

[SWComposites]

Fatstress - do you work for an OEM or an airline or ? What type of components and materials are you working with?

If an airline, then the allowable repair depth and size is given in the aircraft SRM; for repairs larger than those limits, you need to contact the manufacturer.

If an OEM, repair sizes and depths have to be validated by test data. For thin skin sandwich panels, repair depths can usually be the full thickness of the facesheet, and peridoic inspections are usually not required. On the other hand for thick solid laminate structure, repair sizes and depths may be more limited, and inspections may be required, depending on the results of the validating tests, etc.

 

[ARDTL]

Fastress,
I would reccommend with any bonded repair to aircraft structure, that you employ an inspection regime until there is a level of confidence the repair is going to stay put.

Any composite repair is process driven and therefore, only as good as the technician that carried out the repair. As there is no NDI test that will determine the integrity of a bond line, it is impossible to say with total conviction the repair is good for the life of the aircraft or component. Current NDI techniques will only give an indication of voids and disbonds in a composite.

 

[tatsco]

REF AMERICAN AIRLINES FLT 587 - LOOKING AT THE PHOTOS OF THE FIN MOUNTING LUGS (6) IT APPEARS TO BE "VERY THIN EDGE MARGINS" - OR IT WOULD BE ON METAL STRUCTURE? HOW DO COMPOSITE GUYS DESCRIBE THIS AND WHAT WOULD A LIMIT BE?

JIM - TATSCO

 

[40818]

tatsco, its bad form to jump onto somebody elses post, you should create your own post.
In anycase, what picture are you refering to (and create a new post to put in in).
Cheers.

 

[tatsco]

MEA CULPA! Now, if a Professional person would point out the correct road for me I'll leave. (I had searched using AA587, COMPOSITE LUG FAILURES, COMPOSITE SHEAR and i didn't find anything.)

Jim

 

[FEMdude]

tatsco, correct road is

http://www.airdisaster.com

forum(therr is a lot of pics and conversation about AA587)

 

[magom]

SWComposites said:
"For thin skin sandwich panels, repair depths can usually be the full thickness of the facesheet, and peridoic inspections are usually not required."

Do you mean that only the thin thickness can be riveted using a doubler? do you have any available document or reference that shows this statement?! Thank you in advance!

Regards,

 

[SWComposites]

I was referring to bonded scarf repairs of facesheets. "Mechanically fastened" doublers are NOT recommended for repairs to sandwich panels (they cause lots of damage and provide a path for moisture ingression into the core). And "rivets" should almost NEVER be used in composite structures (the only exceptions are the trailing edges of some control surfaces where double sided flush fasteners are needed).

What is the issue behind your question?