FAR 23.572(a)(2) Fail Safe 1

[Lescombes]

Hello everyone,

In consultation with our NAA, i'm currently undertaking an analysis to see whether a primary structural joint on an aircraft can be shown as fail-safe under FAR 23.572(a)(2). I'm using FAA AC 23-13A as guidance, specifically para 3-10(b):

You may also use the analytical approach when conservative failures are assumed. However, the failure must be detected long before the critical crack length is approached, and the margins of safety resulting from the analysis must be considerably more than the fail-safe residual static strength level

Might someone be able to share how the word "considerably" is typically interpreted when relating to the resulting MoS? Say, for argument's sake, the linear ultimate margin is +0.1 against the ultimate static case of 75% limit load with a factor of 1.15. Would that typically be "considerable" in users' experiences? Or should the MoS be more +1.0 to be considerable? Or is there no answer besides whatever your [insert appropriate signatory/authority] interprets it as?

Many thanks,
Greg

 

[SWComposites]

IMO a +0.1 MS is not "considerable". Something more like +0.5 or more is probably expected. Probably also depends on the particuler analysis method used, the pedigree of the input material properties, and the confidence in the applied loads.

 

[rb1957]

it looks to me like an "each way bet"; ie giving you a direction then giving you more constraints rather than a simple original direction (like carry ultimate load with the failed structure, which is equivalent to MS of 0.50 under limit load).

there is a lot of history here. I think the Brits started fail-safe structures and in their mind the failures were detectable. The US picked up on fail-safe structures but lost sight of the detectability of failure. The usual approach (FAR25) to residual strength is limit load capability, with some factoring. I'd argue that an MS of 0.50 is super conservative (like I said ultimate load capability with failed loadpaths) and maybe "considerably" mean MS = 0.33 (although the difference is small !). Detecting the failure is probably "daunting". And remember FAR23 is supposed to be a "lower" standard than FAR25. Personally I think you have much more difficulty showing "failure detected long before the critical crack length is approached" ...
"long" ? "approached" ?? this means you're doing crack growth of the failed structure ... why not go for 25.571 in lieu of 23.57x ? Of course the issue could be the operators having a mandatory inspection program.

Technically you need to show that the critical crack length in the secondary loadpath under failed structure loads (something that can be calculated from failed structure loads) won't be reached within 2 inspection intervals (looking for the original failed loadpath). The analysis behind this is "daunting".

another day in paradise, or is paradise one day closer ?

 

[SAITAETGrad]

Caution -

Lessons Learned - Dan Air 707-300

When an aircraft has been certificated against failsafe criteria, those portions of the structure considered significant in the failsafe design should be identified in the approved inspection schedule and their relative importance defined.

The applicant must develop and implement inspections or other procedures to prevent structural failures due to foreseeable causes of strength degradation, which could result in serious or fatal injuries, or extended periods of operation with reduced safety margins. Each of the inspections or other procedures developed under this section must be included in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness required by §23.1529.

Fail safe features are great but without inspections problematic/difficult.

Sizing for damage tolerant seems like less work. Instead of 75% of limit load plus "fudge_k(?)", simply 100% of limit load. See 23.573(a)(3).

Put another way, what RB said.

 

[rb1957]

sort of what i hinted at ...

but I don't think the earlier FAR23 envisions the same mandatory structural inspections as FAR25. And I supposed that maybe the FAR23 operators would baulk at these requirements. A, in my mind dangerous, way to resolve this is "slow crack growth, non-inspectable" ... show crack growth from flaws exceeds 2* the service life so you don't need to inspect; I believe many military programs use this approach. As far as I know civil inspection programs require an inspection at 1/2 (or 2/3) life, followed by other inspections as required.

A practical issue is that it is, or it can be, very onerous to develop inspection procedures for cracks that you are quite certain aren't there, that you may employ only once in the life of the plane, particularly to detect buried cracks. Thus limit load capability with failed structure sounds a pretty safe and reasonable approach; possibly limit load with an inspection program, and ultimate load without ?

another day in paradise, or is paradise one day closer ?

 

[Lescombes]

Hello everyone,

Thanks for your inputs so far. And a nod to SAITAETGrad - the Dan Air 707 accident report is fascinating reading (actually something I only read a couple of months ago myself).

IMO a +0.1 MS is not "considerable". Something more like +0.5 or more is probably expected.

I acknowledge it's hard to pick an MS out of the air without knowing the analysis method (FEM with a complete part failure) pedigree of the input material properties (just MMPDS) or confidence in loads (FAR 23 Loads equivalent quality) so I appreciate your thoughts.

Some more background, which I hastily omitted: the aircraft was certified as safe-life, and the structural parts in the joint already have published safe-life limits. It's a long story, but we've found some cracks starting well below the published life for the affected part. So, we've thought to ourselves, if the structure can actually be shown to be fail-safe then perhaps we don't need to worry (they say, hopefully...) The part(s) are reasonably inspectable, not buried deep in the aircraft.

Technically you need to show that the critical crack length in the secondary loadpath under failed structure loads (something that can be calculated from failed structure loads) won't be reached within 2 inspection intervals (looking for the original failed loadpath). The analysis behind this is "daunting".

rb1957 - I understand that this may be the case for FAR 25, but I don't understand that as a requirement from FAR 23? I do concur, however, with the 'daunting' "failure must be detected long before critical crack length" statement.

 

[verymadmac]

It shouldn't be forgotten that design loads for part 23 are quite a bit more conservative than the comparative part 25.

How well do you know the history of this aircraft, Part 23 tend to subject to a wider range of possible operational load spectrum than Part 25. It can be quite simple for an aircraft to be operated in a manner that results in accelerated life consumption. I encountered a classic example, aircraft was flying what was described as cable patrols, aircraft went out to ensure recreational and commercial boats were staying out side submarine cable exclusion zone, they omitted to include as that when they found boats they drove then off the exclusion by aggressive low level maneuvers. Military use of turbine part 23 aircraft for pilot training is also another classic for ending outside the expected spectrum.

 

[WKTaylor]

Some Fud-4-thot, RE military evolution of Safe-Life, Fail-Safe, Damage Tolerance

A few reference I’ve found over the years... that refined the concepts...

MIL-A-8860 AIRPLANE STRENGTH AND RIGIDITY, GENERAL SPECIFICATION FOR
3.1.10 Fall-safe. The complete airframe shall be designed such that
failure of a single structural element will neither cause catastrophic failure
of the airplane nor prevent safe continuance of its flight to a planned
destination or to an aircraft carrier. Redundancy, alternate load paths and
systems, and other fail-safe principles are required to achieve this
capability. For this fall-safe requirement, the airframe is defined as
including all of the structural elements of major systems and all of the
structural connecting and supporting elements of powerplant installations,
whereby the failure of these elements can cause uncontrollable motions of the
airplane within the speed limits for its structural design, prevent the
airplane from achieving speeds sufficiently low to effect a safe landing, and
reduce the ultimate factor of safety for flight design conditions from 1.5 to
a value less than 1.0.

AGARD-AG-292 HELICOPTER FATIGUE DESIGN GUIDE

AVRADCOM SAFE-LIFE AND DAMAGE-TOLERANT DESIGN APPROACHES FOR HELICOPTER STRUCTURES

USAAMRDL-TR-75-12 FAIL-SAFE/SAFE-LIFE INTERFACE CRITERIA

USARADCOM Report SAFE-LIFE AND DAMAGE-TOLERANT DESIGN APPROACHES FOR HELICOPTER STRUCTURES

DESIGN OF MODERN AIRCRAFT STRUCTURE AND THE ROLE OF NDI

http://www.ndt.net/article/ecndt98/aero/001/001.htm

UNIVERSITY OF MONTANA... DESIGN AND ANALYSIS OF AIRCRAFT STRUCTURES ~ DESIGN PHILOSOPHY

http://www.montana.edu/dcairns/documents/introduction_to_aerospace-web/04-DesignPhilosophy.pdf

UNIVERSITY OF WASHINGTON... FATIGUE AND DAMAGE TOLERANCE REQUIREMENTS OF CIVIL AVIATION

https://www.aa.washington.edu/files/mae_colloquium/mae_winter2014/Safarian_2-3-2014.pdf

Regards, Wil Taylor

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