Stress Crack Detection systems

[Higgler]

I'm writing a proposal for a requirement to detect stress cracks in the surfaces of aircraft. What size cracks normally occur that need to be detected. I proposed to detect cracks as small as 0.005", and maybe 10 times smaller, this may seem ridiculously small, hence my writing to all of you experienced aviation wizards. Any hint on what size and shape cracks normally occur that need detecting would be appreciated. Including what size is not of concern and what size is very dangerous. My proposal is to detect cracks in a metal surface only. Another question is regarding non metal surface cracks, are they of great concern (I expect that with composite structures the answer is yes).
Thanks,
KCH

 

[FastMouse]

Hi Higgler,

In my experience at an aircraft engineering organisation, the smallest crack that we would typically ask inspectors to find would be a 1mm x 1 mm corner flaw in aluminium alloy that was typically 4 mm to 20 mm thick. These would usually be cracks emanating from bolt holes and an often-encountered problem was that the hypothetical crack would be situated underneath the bolt head, making detection difficult. We would usually assume that the crack shape for a 1 mm x 1 mm crack was quarter-circular.

We had no minimum crack size below which cracks were of no concern. In addition to structural design, my group was also responsible for analysis of in-service findings, and all detected cracks were a concern, regardless of their size. Under the right conditions, inspectors were able to find cracks significantly smaller than 1 mm x 1 mm, and we sweated over them just as much as larger findings.

As to what size was “very dangerous”, we calculated the critical crack size under limit-load conditions, ie - if the structure experienced a limit load condition, the crack would fast-fracture, leading to structural failure. I don’t know if you have an aviation background, so limit load is, in a nutshell, the maximum expected load that the airframe would ever see in its life, usually due to some abnormal operating condition, such as a 2.5 g manoeuvre, large gust, hard landing, or whatever. We were dealing with commercial aircraft - military aircraft have a much different loading environment.

In some cases we were operating in plane strain conditions, in other cases we were in the plane stress regime, so critical crack sizes ranged from, typically, 10 mm from the bore of the hole to 150 mm tip-to-tip. I emphasise that these are just typical critical crack sizes, not an absolute envelope. If the critical crack size was not a through-crack, our analyses would predict the crack growth rates along the bore of the hole, and growth rates along the surface of the material. These crack rates would be different, so you could end up with, say a 15 mm (bore) x 10 mm (surface) crack from an assumed 1.27 mm x 1.27 mm initial flaw, and we would assume that the profile of the crack tip was elliptical between the bore crack and the surface crack. Hope that makes sense. Too difficult to draw a diagram in ASCII!

I can’t comment on non-metallic structure. While we did analyse plastics from time to time, our core competence was analysis of aluminium.

Your proposal for detecting cracks of 0.005” size, and maybe even and order of magnitude smaller is interesting. These detectable crack sizes are surprisingly small for aircraft inspection and I am sure that bringing this capability to in-service aircraft would provoke some discussion in the industry. Some people would love it. Others would hate it.

I am aware of cases in other organisations where, to keep the aircraft in operation rather than stuck in a hangar for repairs, they would not want to do anything about a crack less than, say, 7mm long, so they would deliberately calibrate their inspection methods such that cracks less than 7 mm long would not be found, despite the fact that the inspection technique could easily find much smaller cracks with appropriate calibration.

Without going into the details, there are situations where this is a sound engineering philosphy, but at other times, it is analogous to simply “looking the other way” and those employing such an approach will not want a technique that can find 0.005” cracks or smaller.

There is also the issue of discriminating between true cracks and simple scratching of the material as a result of bolt installation, etc, but I am sure that your discussions with others have highlighted this and more. I can see an immediate application to better understanding of the benefits of manufacturing techniques such as cold expansion of holes.

I hope that this is useful.

Good luck! I would love to be able to ask our technical publications people to write up a technique that can find a 0.005” crack.

 

[Higgler]

FastMouse,
Thankyou for the response. After reading it, additional questions arise.
1) Are surface cracks less important than through cracks (or I think you called them bore hole cracks). I would think that to be true. What depth of surface crack needs detection? I suppose you'd want to actually measure it to see if you call it a scratch or a crack.
2) Seems the 7 mm minimum size to keep aircraft flying is a simple thing to detect. Simple RF systems could find a crack that size. If we propose the cheapest system, should we designate 7 mm as the minimal crack size to be detectable? Or is 5 mm or 3 mm a better proposal size?
3) Does a solution require "inspection from outside the craft only", or is it feasible to propose placing hardware/(detectors or transmitters) inside the aircraft to propagate energy through the craft?
4) What are your thoughts on how the system should function? One piece of hardware used manually, place at surface of the aircraft, push a button to test and get a green light or red light for an answer? Or an automated machine? etc.
Thanks
KCH.

 

[FastMouse]

Hello Higgler.

Interesting Questions! Here are my thoughts...

Question 1)

From an in-service point of view, you can’t really say that a surface crack is more important than a bore crack. They are both equally important as the point is that you have an aircraft with a crack in it, and you have to do something about it. Leaving a crack in situ is rarely an option.

On the more analytical side of things, there may be cases where a crack might run along the bore of the hole relatively quickly to become a through crack, but not extend away radially from the hole very much, so the surface crack length remains small. I can envision this happening with certain residual stress fields existing around the bolt hole, but it’s a debateable point from the perspective of supporting in-service aircraft. That’s just another way of saying that we need a test program!

The answer to the required detectable depth of surface crack is sort of the same. Whatever we can find is important. Discriminating between small cracks and other surface defects is also important.

Question 2)

7 mm is indeed a simple thing to detect. My point was that some people may not wish to have a system that can find very small cracks, because once they find the crack, then they have to do something about it. This might involve grounding the aircraft, which causes major pressure to be focussed on whoever makes that decision. So some people find it easier to deliberately fail to detect existing cracks and not have to ground the aircraft than to find the crack and make that difficult decision. I’m glad to say that these people have no place in any reputable organisation.

If you are looking for a minimum size to detect, I would say go for the smallest crack that you can detect with 95% reliability. Current methods can reliably detect a 1 mm crack, or less, under the right conditions, so to be industrially interesting, maybe a 0.2 mm crack detection capability would be a good objective. The 0.005” (and less!) that you talked about in your original posting sounded great.

Question 3)

If the solution is from the exterior of the aircraft only, then that is an immense plus. If hardware must be placed on both sides of the structure, as with an X-ray process, then the technique becomes less attractive, but still potentially useful.

Fuselages obviously have easy(ish) access to both sides of the structure when all the cabin panels, insulation, etc have been removed, Wings are a more difficult proposition as draining and venting the wings is a big job. We try to avoid it. Of course, if an aircraft is undergoing a D-check, (or 8-C check, depending on the nomenclature that you use), you can expect the fuselage to be emptied of furnishings and the wings to be drained with the access hole covers removed. Getting hardware in place should not be such an issue in that case.

Question 4)

I see two modes of operation.

First, a manual device that can be used to inspect specific holes or areas. Service bulletins often specify very localised inspections, and scheduled maintenance programs often used localised NDT, so there’s a possible application there. Just like rotoprobes, or ultrasonics, etc. Also, there is a need to be able to do quick ad-hoc inspections.

Second, an automated device is also industrially interesting. Applying NDT to huge swathes of structure brings up all sorts of human factors issues that can make the inspections less successful than they might otherwise be. I’ve been a little involved in automated inspection of aircraft wings. It worked well. My experience is that automated techniques are more easily applicable to manufacturing situations and less easily applicable to in-service aircraft. That’s not to say it can’t be done, ‘cause we did it. It was just difficult, and I’ve heard others say the same. I think that other industries make greater used of automated, in-service inspections that the aerospace industries. I’m thinking railway rails, nuke reactors… I’m sure that there are other good examples.

Fast Mouse