Stress Determination by Crack Inspection

[macoovacany]

Hello all,
Consider a crack in the middle of large sheet of metal, about the same dimensions as, say, the sidewalls of the engine bay mount for the Cessna 210. The crack was probably initiated by a tool scratch, falling screw driver, whatever. There is now a crack that runs at approximately 45 degrees to the vertical about 1-1/2" long.

From this level of information, can we determine the direction of principle stresses? Does the mode of the crack matter? Would the direction of the crack depend on the Shear/Axial stresses? Would this mode change with a different material, i.e. different ratio of shear strength to axial strength. Would the direction of the crack change with a different material?

The reason I ask is that by removing material 'in front of' the crack (a series of decreasing circles, the axis along the line of principal stress), the stress concentration at the crack tip can be reduced dramatically.

Not that that is what we approved for the particular problem above (stop drill crack, doubler 1 gauge thicker than parent material, etc: standard stuff). I'm asking just in case a problem comes along that may require a solution where the standard stuff may not be approriate. In summary, got crack, want to know stress direction.

Timbo

 

[ubrales]

The original stress was in the direction 90 degrees to the crack - the crack has relieved the stress somewhat and may have changed the direction of the new stress. Fatigue may be a factor. In any case, a detailed analysis needs to be conducted since a crack is the first sign of impending failure.

Since it is an aircraft application, can the part be replaced with a beefed-up part, not just an original part?

Whatever repair is considered, vibrations caused by the repair needs to be considered.

One way of stress analysis involves the etching of a grid of small circles and subjecting the part to its normal duty cycle and later observing the distortion pattern of the circles. This technique is sometimes used in deep draw and forming operations, and will work quite well for any kind of stress analysis particularly when failure has occured.

 

[jetmaker]

macoovacany,

I agree with ubrales in that the maximum principal stress tends to be 90 deg. to the crack orientation. However, substructure and changes in thickness can cause a crack to grow in the direction of least resistance (ie. skin padup, stinger/frame, bonded doubler, etc). Therefore, you need to evaluate the structure around to verify that it is not influencing crack propagation.

You mention a repair of stop-drilling the crack, is this an approved method by the aircraft manufacturer for primary structure? There is debate as to whether a stop-drilled crack can actually relieve the stress intensity at the crack tip. Where I work, this repair would have been conducted using a dime/dollar repair, thereby eliminating the crack altogether.

Thanks for the input.

jetmaker

 

[macoovacany]

dime/dollar repair?

Timbo

 

[jetmaker]

Timbo,

Yeah, a dime/dollar repair is 2 concentric circles, one smaller than the other, used for repairing holes in sheet metal. The crack would be cut out to a round hole configuration, then the "dime" skin repair cut to match. The "dollar" patch would be installed on the inside, connecting it to the basic skin structure. The dime would then be riveted to the dollar to ensure aerodynamic smoothness and some hole fill.

Another issue about your repair is that you have now covered the crack with a patch, potentially allowing the crack to grow without detection until it emerges from beneath the patch. Forgive me is this is not the proper configuration you have. I'm picturing a firewall repair where the patch is in the engine compartment, where easy inspection would occur.

Regards,

jetmaker

 

[sabedi]

What kind of substantiation would be needed to get this repair approved?

 

[aviat]

What jetmaker is referring is called a circular patch plate or plug patch and is shown in AC 43.13-1B on page 4-32. It is a standard repair and is in all SRM’s. Its advantage is that it will take stress equally from all directions, unlike other repairs. But it cannot be used too close to reinforcement.

Stop drilling the crack and installing a patch on either outside or inside is usually an approved repair on low stress skins on this type of A/C. However, if I read your question correctly the area concerned is a stressed skin panel. In that case you should have used a plug patch as per 4-58f of 43-13 or your SRM.

As for the direction of loads from the crack I don’t think it is very reliable. It is an engine mount and will have torque, tension, shear, and compression – the whole ball of wax.

As for substantiation for approval you just need to list the chapter and verse of the approved data used for repair (43-13 or SRM etc.). This is just routine maintenance; it is not even a major repair.

 

[Sparweb]

aviat,

I'm pretty sure that the C210 doesn't have an SRM. It does sound like it's within the limits of AC 43.13-1B...

BUT

It also sounds like the problem is on the firewall. Some of the 210's are pressurized. Is the engine firewall also the forward pressure bulkhead? If so, then the firewall is a membrane under pressure, and AC 43.13-1B DOESN'T APPLY EITHER.

Macoovacancy now has to develop a repair scheme of his own, and justify it vis-a-vis the applied pressure loads.


"Simplicate, and add more lightness" - Bill Stout
Steven Fahey, CET

 

[aviat]

Cessna SRMs for a lot of their singles are included in a single book which includes maintenance, wiring diagrams and SRM, and is called a service manual. For many it covers a complete series. For instance, I have one book that covers all 100 series 1962 and prior. The SRM is 24 pages long.

 

[Sparweb]

You got me. How far off the mark am I on the other stuff I posted?

Humbly,

"Simplicate, and add more lightness" - Bill Stout
Steven Fahey, CET