Area Out / Capabilities

[born2fly]

Hello,
I have an area out question.
On my (hypothetical) airplane control surface, some corrosion was discovered and trimmed out.
To determine the area loss/capability, is it the thickness of the material times the length (or width) of the trimmed surface?
So if the material was .063 inches thick, it would be Ftu(.063 x L), with Ftu used to determine a more conservative allowable.

Thank you in advance for answer my question.

 

[WKTaylor]

born2fly... Your question is meaningless without real-world details.

Regards, Wil Taylor

Trust - But Verify!

We believe to be true what we prefer to be true.

For those who believe, no proof is required; for those who cannot believe, no proof is possible.

Unfortunately, in science what You 'believe' is irrelevant – "Orion"

 

[proppastie]

Here is a page from AC 43-13-1b. It is free on line FAA site google search. Go to this section and there will be more than you ever wanted to know about aircraft repair. See attached file.

Don A&P PE

 

[WKTaylor]

Your example [AC43.13-1b for fastener row area-out] does not blend-well with Your original statement...

"... surface, some corrosion was discovered and trimmed out.
To determine the area loss/capability, is it the thickness of the material times the length (or width) of the trimmed surface?
So if the material was .063 inches thick, it would be Ftu(.063 x L), with Ftu used to determine a more conservative allowable."

Also for a [simplified] longitidunal load example, Your equations should probably read Ftu x 0.063 x LT [sheet width] = load-max

Load-after localized corrosion blend-out ~ = Ftu x [(0.063 x LT sheet)-(blend-depth x LT width of blend)]

Note: this equation probably should have a Kt associated with it... to account for localized thinning at an arbitrary location across the psan of the skin.

Absolutely conservative would be to factor the blend depth across the sheet width.

CAUTION: cross-load cases would cause this problem to "blow-up".

I suspect You need to take a liaison engineering course, so You can see all the factors associated with damage analysis.

Regards, Wil Taylor

Trust - But Verify!

We believe to be true what we prefer to be true.

For those who believe, no proof is required; for those who cannot believe, no proof is possible.

Unfortunately, in science what You 'believe' is irrelevant – "Orion"

 

[rb1957]

personally i'm am not a fan of assuming that the material is working at Ftu. sure it is conservative but i've seen it lead to ridiculous repairs. SRMs tend to give allowances for corrosion clean out. I wouldn't be surprised if they excluded control surfaces. I wouldn't like to go beyond 10% thickness.

Quando Omni Flunkus Moritati

 

[proppastie]

Your example [AC43.13-1b for fastener row area-out] does not blend-well with Your original statement...


I am not the OP however it seemed like he was doing a repair of a light aircraft, and reading 43.13 may get him were he needs to go faster than trying to learn stress analysis. I can do the former but not the latter.

 

[WKTaylor]

proppsatie... this is a aero engineering forum for degree'd/working engineers... not for mechanics who want advice interpreting figures from Maintenance Manuals for structural repair intent.

Take it from me... a 36-year practicing liaison/design structural/mechanicl enginurd... there is a LOT more about the direction of this question than meets the eye... and only formal liaison training could get him on a path where we can even begin to have a sensible discussion.

I am pretty close to red-flagging You and born2fly as non-engineers looking for engineering advice.

Regards, Wil Taylor

Trust - But Verify!

We believe to be true what we prefer to be true.

For those who believe, no proof is required; for those who cannot believe, no proof is possible.

Unfortunately, in science what You 'believe' is irrelevant – "Orion"