Understanding crack growth analysis 2

[Craig1983]

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

 

[der8110]

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.

 

[msm1981]

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

 

[der8110]

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.