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Understanding crack growth analysis
2

Understanding crack growth analysis

Understanding crack growth analysis

(OP)
Hello,

I was looking into learning the RAPID software developed by FAA. I started by viewing the tutorial and the example provided with the software on Antenna Installation. Everything was fine and understandable until I got to the DTA part.

My DTA knowledge is not good but I have some knowledge of crack growth... I am wondering how come the software finds it's limit at a crack growth of 63.78 inches if the plate is only 7 inches. I understand how the limit loads are calculated but I am having a hard time accepting that the threshold will be set at a value equal to a crack length of 63.78 inches which is way larger than the doubler itself (7 inches).

Please clarify this! Also, if you have any elementary DTA books or papers you suggest especially if using RAPID let me know!

Thanks
Craig

Here is the extract from the tutorial I am speaking of.

DAMAGE TOLERANCE ANALYSIS RESULTS OF THE REPAIR
-----------------------------------------------


THE FOLLOWING 3 SETS OF OUTPUT ARE FOR THE LONGITUDINAL CRACKS OF SIDE 2:

     BULGING IS NOT CONSIDERED IN THE ANALYSIS FOR THE LONGITUDINAL CRACK.


     *****   CENTER FASTENER   *****


     FASTENER LOAD =   12.90390 POUNDS (BASED ON 1000.0 PSI FAR FIELD REFERENCE STRESS)


     (A) CRACK GROWTH LIFE:

         EDGE-TO-TIP       CRACK GROWTH
         CRACK LENGTH          LIFE    
           (INCHES)         (FLIGHTS)  
         ------------      ------------
            0.05000                0
            0.08795            13975
            0.12590            24985
            0.16385            34554
            0.20180            43076
            0.23975            50726
        .
        . (etc...)
        .
            0.65720            99539
            0.69515           101629
            0.73310           103283
            0.77105           104483
            0.80900           105293
 
          TIP-TO-TIP       CRACK GROWTH
         CRACK LENGTH          LIFE    
           (INCHES)         (FLIGHTS)  
         ------------      ------------
            1.19100           105293
            1.63063           105293
            1.68955           105731
            1.74846           106156
            1.80738           106572
            1.86630           106978
        .
        . (etc...)
        .
           37.85822           119826
           41.02494           119950
           44.50834           120066
           48.34007           120174
           52.55498           120275
           57.19138           120369
           62.29142           120457
           63.78702           120479


     (B) RESIDUAL STRENGTH:

            CRACK            RESIDUAL  
            LENGTH           STRENGTH  
           (INCHES)           (KSI)    
         ------------      ------------
            0.00000          44.00000
           24.09188          28.59985
           25.85386          27.53323
           27.82014          26.35822
        .
        . (etc...)
        .
           48.34007          19.60196
           52.55498          18.74188
           57.19138          17.90910
           62.29142          17.10355
           63.78702          16.89600
 
 

RE: Understanding crack growth analysis

RAPIDC (or RAPID) does not know panel size limits.  Also be aware the frames and stringers you put in are just for show; they do not reduce the pr/t stress (i.e., Flugge stresses) or arrest the crack (as they would in reality).  

Usually a crack growing towards a free edge would accelerate (finite width effect), but where on an actual aircraft (pressure vessel, that is) do you have an edge hanging out in space?  Skin panels have splices, and you have to look at your output to find your true final crack length and those geometric constraints.  In your example the difference between 7" and 63" is less than 10,000 cycles - 110,000 vs. 120,000 or so.  Tests show that cracks slow down significantly as they approach splices (or frames, stringers, or crack stoppers), so finite width is not a factor.  This assumes, of course, the splice itself is of conventional design and it does not have multiple site damage on its own.

I also sense some confusion on the panel size - or does your doubler take up the entire bay (both 7").

Most of all be aware that despite all the automation the results still need interpretation.  The software is free but nothing else is.

As far as references, I assume you have the Analysis Methods Document.  There are literally dozens of other publications out there; for a start I might suggest DOT/FAA/CT-93/69.I and DOT/FAA/CT-93/69.II, these should get you oriented.

RE: Understanding crack growth analysis

How to use rapidC software for canard Configuration e.g. Piaggio Avanti II. what about Zone I/II.

RE: Understanding crack growth analysis

Go back to basic statics and find the corresponding bending moment for a three lifting surface aircraft.  Look in the Analysis Methods Document for how it was derived for the (2) surface case.

Quick and dirty- you could enter L=S (and D same as radius), betcha on an aircraft that small the penalty would not be horrendous.

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