Fatigue Critical Baseline Structure List for Cessna 560 1

[asistructures]

Does anyone have the fatigue critical baseline structures list for the Cessna 560. I am not sure one exists. This list will identify principle structural elements of the airplane that would require damage tolerance evaluation when modified or repaired

 

[rb1957]

if this is for Part 26, Cessna Have to produce the data.

if this is for an STC you're doing (and you're looking for a definition of primary structure), that is probably more generic (wing, fuselage, empennage, ...).

note that the TC shows that the Citation is not a damage tolerant a/c (they're at rev 25.17)

 

[GBor]

I've never seen one from Cessna, but I haven't worked on a 560. Boeing pretty much placed every structural member including clips, floor beams, and any type of stiffener on their commercial products. Boeing also charges an extreme fee for the list.

Hopefully Cessna didn't do the same thing, but it is safe to assume they did.

 

[asistructures]

I know the lists can be expensive, my next step will be direct contact with Cessna. I was looking for the list in support of a skin repair near the L1 door, lower aft corner

 

[GBor]

Skin repair...I'm sure skin, frames, and stringers make the list. PSE was the baseline for FCBS, then they added anything else to the list.

 

[rb1957]

ok, you're repairing the fuselage. it's almost certainly primary structure (to answer your original question). i would propose that you complete a fatgiue analysis of your repair, Michael Niu's book has a nice method for establishing the stress concentration at a loaded hole., as the a/c's basis of certification doesn't include damage tolerance. you Could do a DTA and implement your own inspection program (i sometimes have to do this for Kingairs, even though they are not DT ... whatever to keep the cert people happy)

 

[der8110]

Part 26 would not apply to the 560 by any stretch of the definitions. SRM is next to useless for what you seek. Based on your names of the involved elements involved and AC25.571-1C, you do have fatigue critical structure. Your choice whether to comply with fatigue (cert basis) or DT.

 

[rb1957]

the Part 26 was only in response to the OP's terminology (FCBS).

i'd suggest you consider carefully before committing to doing DTA. if the plane doesn't have a mandatory structural inspection manual, then it almost certainly won't have defined inspection procedures. in my experience using an OEM's inspection technique beats the heck out of creating (and proving) your own. mind you your plane is probably a "hangar queen" (no slight implied) and the design life goal is short (compared to a workin' plane) so you may be able to get away with a "non-inspectable" result (the USAF permits you not to inspect a location if the threshold interal > 2*service life, normally the FAA doesn't).

 

[der8110]

Damage tolerance does not always result in a mandatory inspection. Following that logic, by doing a fatigue analysis, then a mandatory retirement time would have to be imposed (both are Airworthiness Limitations). Damage tolerance is just another analytical method, what you do with it depends entirely on where the numbers fall.

Since it's on an airplane it is going to have some kind of inspection, and the 560 indeed does have a nice phase inspection program. Cabin door frame is every 1200 hours or 36 months. The trick is to do a design that remains inspectable, and have an analysis that shows these existing inspections remain appropriate.

Something else that may sway your approach: the 560 has Special Conditions (25-ANM-31) which would apply if your repair involves the pressure vessel boundary. DT makes compliance much easier.

 

[rb1957]

agreed, what can get approved varies from place to place. my local cert people prefer to cut-off the threshold at 1/2 life so someone looks at the structure just in case, the USAF allows (allowed?) structure not to be inspected.

i'd say the difference with fatigue analysis is that if the factored safe life exceeds the design life then no limitation is imposed, always.

do the special conditions in ICAs ? i'd think this'd be a little odd (since it isn't a DT (admt 45) airplane).

 

[GBor]

Threshold inspection is defined by the least of: N-detectable, 1/2 N-critical, 3/4 Design Service Goal, or 1/4 N-fatigue, where 'N' is a GAG cycle. In short, if you perform a DTA, you can't end up with inspections exceeding the aircraft DSG or the fatigue life.

Subsequent inspections are then defined by (N-critical - N-detectable)/3 so that you get at least 2 inspections between the time you are supposed to be able to detect the crack and the time when it fails catastrophically.

 

[rb1957]

that's interesting (the 3/4 DSG particularly) ... is that FAA practice ? (TC, in my experience, applies 1/2 DSG)

 

[GBor]

That's what has been used for the dozen or so reports I've turned in over the last couple of years, and the ACO seems good with it. I don't believe that has been the inspection driver in any of the cases I've dealt with. If it were 1/2 DSG it may have come in to play, but I don't think so. Detectable and Critical seem to drive more often. Fatigue has driven once.

I have to admit that I'm not a DER (yet), so DER8110 may have better knowledge and different guidelines. My information comes from a close relationship with a DTA DER, and I believe she follows the FAA pretty closely. She attends the various DTA seminars hosted by the FAA and argues with them about things with which she disagrees until she finally submits to their requirements

 

[der8110]

Your knowledge is pretty thorough GBor, especially the part about knuckling under to whatever the ACO wants. The problem being some ACO's are different, and have requirements that are not based on consistent guidelines and sometimes even contradicts the FAA's own information. I say "consistent" because as it started off, USAF and FAA DT terminology and methodology was remarkably close. And for "contradicts", one example is AC25.571-1C says threshold inspection can be set by EITHER (emphasis added) fatigue with an appropriate scatter factor (yours being 4) or slow crack growth with initial manufacturing damage. Not "lesser of" though there is additional conservatism there. Like you, I have only seen fatigue trump Ndet or Ncr/2 in one instance. Another variable I see is what you have for repeat inspections, the 3 in the denominator seems to be an EASA thing that has been randomly adopted. Sometimes it's not even 3, it's a fatigue scatter factor. And using fatigue scatter factors in DT is like nails on a chalkboard to me (along with stress concentration factors on top of stress intensity solutions). Again conservative, but not very indicative of an understanding of the definitions or their origins. Design Service Goal is defined in AC25.571-1 but is not a regulatory requirement. Mainly a customer expectation or manufacturer warranty, if you will. And good luck getting the DSG for a 560. With DT, is there really a DSG?- conceptually fly it forever (as long as you follow the program!).

rb: I may have been too brief about Special Conditions. They are for "high altitude aircraft" of which the 560 is one. Nowadays the requirements are part of regulation, amendment 25-87. Read the preamble for some background. Basically whenever "high altitude" applies there needs to be a leakage analysis through cracks propagating over four (and number can vary) inspection intervals. So what better way to know how long/wide your crack is than by DT (LEFM). Fatigue can't get you that. So depending on the analytical results, there may or may not be anything to go into the ICA. Rarely see leakage driving any maintanance, even on a small cabin like a 560. There are going to be standard hull checks anyway; that is when cracks are typically found.