would anybody like to comment on this can of worms

[berkshire]

CLARIFY CRACKS POLICY TO COVER ALL OLDER AIRCRAFT, AOPA SAYS
AOPA is asking the FAA to revise a proposed advisory circular (AC) to
clarify that it can be applied to all older general aviation aircraft.
The new AC would set guidelines for allowing aircraft to continue flying
with known structural cracks, something most aircraft develop as they
age. The AC would publicize a long-existing FAA policy that deems an
aircraft airworthy if it can still withstand the ultimate design load.
"The FAA left out the majority of older GA aircraft from this guidance
document," said Luis Gutierrez, AOPA director of regulatory and
certification policy. "It only applies to Part 23-certificated aircraft.
But most aircraft flying today were certificated under the old CAR 3
standards. It's important that the policy be applied uniformly and
predictably to all aircraft in order to keep them flying safely and
affordably." See AOPA Online
(

http://www.aopa.org/whatsnew/newsitems/2005/050705cracks.html

).
What do you think about this?
Brian Evans.

 

[rb1957]

i understand that it is FAA policy to allow flight with cracks so long as the residual strength is greater than ultimate load. this means that the structure can still withstand ultimate load, just like any plane leaving the factory, so it is compliant with the reg's. I understand that the C-130s (as an example) sometimes fly with known cracks which presumably are closely monitored and presumably comply with this rule.

whilst i have never tried to get something like this approved, i'd think ...
1) the approval would be for a very limited period, eg return to repair depot;
2) proving this would be pretty difficult, not much stress data available for these 'planes, potential of plasticity effects, load redistribution to other load paths;
3) the approval would require substantiation of a repeat inspection, ie demonstrate crack growth for an inspection interval is still not critical. From this you'd develop a not-to-exceed crack length such that crack growth anticipated during the next inspection interval would produce an unsafe condition (a crack with less than ultimate load residual strength).

the distinction between this (flying with known cracks) and conventional DTA is that conventional DTA (with limit load as the residual strength) is trying to anticipate cracking and develops a maintenance schedule to ensure safe flight (given the possibility of cracking).

seeing the analytical difficulties with this, and the rather questionable analytical assumptions (what flight spectrum would be considered "reasonably" conservative), i would have thought that this approval would only be for emergency, rather than routine, cases.

 

[TheTsquare]

Hey Berkshire,

I gotta Agree with rb1957 on the crack issue. I'm new to this arena as I am still in shool studying for my A&P license (1 month left). I am also a private pilot and I'm looking into getting my own plane (ehough about that) The deal is that I sympathize with you in that thet repair of a crack (one would have to assume that the crack is on the skins) is quite an ordeal especially if its a corporate jet of worse a private jet and I understand that everything in the small aircraft arena is expensive but unless you cut it out and patch it (not quite as expensive as replacing the skin) then you're just YES ...OPENING UPA CAN OF WORMS!!

Even if you do get approval there has to be a real 'pain in the neck' 'AD' (airworthyness directive) that goes along with it and the research involved to start an AD was pretty well laid out by rb1957 in the last post. \

which woud eventually measn that the plane would be grounded when the crack grew to a sufficient problem as to facilitate repair.

 

[berkshire]

The T square,
Before you comment on this, you need to go back and read the proposed changes to the advisory circular (AC23-13).
As it is written, it does not cover older aircraft for which there is no manufacturing support, except for a brief mention in passing.
Very often the original data used for certification is simply not available. This reqires new test data to evaluate a repair or modification.
As the advisory circular is written, the data it contains, is not applicable to the older aircraft. Which is what the AOPA is trying to change
Brian Evans.

 

[rb1957]

is there an NPRM for 23-13 ?

i quickly looked over 23-13 this morning. granted there isn't an explicit analysis path for airplanes without structural data (ie internal stresses). but how does this extend to "flight with known cracks" ... which (hopefully) should be much more difficult to substantiate.

i think that each of these older planes will have analytical problems (and solutions) of their own. i can't see what the FAA could write to cover these cases, other than their standard "or other rational analysis".

 

[berkshire]

Yes,Rb 1957.
AC23-13A.If you click on the AOPA article I cited, then click on Proposed advisory circular, it will take you to the draft copy of AC23-13A.

You are absolutely right about the fact that older planes have problems of their own. The other complicating factor is that not all of the older machines are made of metal. They are composite aircraft. An example is the mainspar cracking problem on Champion aircraft ( wooden Spar) The language of the AC deals primarily with aircraft certified under part 23 of the FARs. The AOPA would like to apply this to CAM 3/CAR3 standards.
As far as the known cracks issue is concerned I feel this can only appy to secondary structure. The AC itself is pretty specific about cracks in critical structure, it does not allow them.
As rb1957 says, if the strength is reduced below safe levels you are not going to fly.
B.E.

 

[Sparweb]

I just can't picture this working. For most aircraft, new or old, the design loads are not available. Phone Raytheon and ask politely for the loads analysis and structural test reports for the Beech 90's. Sit back and listen to the laughter. It's proprietary.

There are probably only a handful of DER's who can both reproduce credible loads data and then create damage-tolerant repair designs on light aircraft (or non-repair, in the case of "flight with known cracks"). The expense would be huge, and if done piecemeal, one repair at a time, one fleet at a time, would represent a lot of duplicated effort.

The original AC 23-13 (1993 vintage) is short and sweet and likely relates only to new aircraft going into production. This revised AC 23-13A is stepping into new ground on the "flight with known cracks" issue. It will take a lot of time for some engineer to massage the data out of the AC's load spectrum charts and do the whole aerodynamic analysis that has to go along with it. By then, the owner could have disassembled the plane, trucked it to the nearest MRO, and put it back together - the way things are done these days anyway!


Steven Fahey, CET

 

[rb1957]

ok, now i've looked at the link !

i'd relax about the "flight-with-known-cracks" issue, para 6.3 spells out the hoops you have to jump through ... if anyone can !! ... so really it's a non-starter.

i'd also relax about the AC ... it is meant as one method for demonstrating compliance.

what i would say about the AC, i think like steve above, is why have they made it soooo difficult ? my experience with commuter airplanes has lead me to think that for the fuselage pressure stress is always the most significant fatigue stress and reasonably sufficient for the forward fuselage; for the aft fuselage i include a factor, something like 1.5, to account for the effect of gust and manoeuvre cycles (ie a GAG cycle of 1.5*hoop stress). the unpressurised fuselage ought to be able to survive with the hoop stress applied (way conservative). wings and H.stabs are less clear, obviously being more affected by gusts and manoeuvres; however, i'd suggested that their durability should be something like the fuselage's and that a GAG stress would be some factor of the hoop stress (2?).

the AC approach, being very rigorous and complete, looks as though it is for new design certification. but surely any flight assumption for this class of airplane is doomed to failure (how many angels can danced on the head of a pin?).

i'd propose a more pragmatic approach ...
show a good fatigue life of hoop stress, factored if you want,
worry (a lot) about concentrated load paths (lift struts) and load transfer/redistribution sites,
a nominal fatigue stress (GAG) cycle for the wings ... 1/3 limit stress ?, = hoop stress (factored) ...
i think this type of guidence would have been useful.

as berkshire says above, many materials in these airplanes would not conform well to detailed F/DT analysis, so why "guild the lily" ?