I want to analyze my airfoil, here it is.

[e-sailpilot86]

I kinda understand polars, and I'm in a rush so here's the polar img(s,), next is the attachment. I did it in X-foil, and I'm clueless...:p whoops, can't post it till formatted right, L/D is about 50.

[e-sailpilot86]

:rolleyes: That wasn't a very informative post, was it? Well, that's what happens when you're in a rush to go somewhere but just can't get off E-zone! :D Here's the polar, and I really don't know whether anything was tuned right in the program. In other words, this polar could be completely far from even being an estimate. I've modified the colors in the file so it doesn't have a black background. I wanted to print it out.

BTW, I'm not sure what combination of airfoils this is, all I know is it was a blend of two from way back when. I'll post the coordinates on the next post since I cannot attach the file. So, what means what? I don't know what a realistic L/D is for the airfoil alone is. I know the L/D for a whole plane is about 20-25 (guestimate). None of what I did was scientific, it's a bit like coming up with an airfoil from what your shoe looks like.

[e-sailpilot86]

pm001
1.000000 0.000000
0.993606 0.002548
0.982702 0.006981
0.970311 0.011203
0.956398 0.015132
0.941068 0.018978
0.925133 0.022666
0.908905 0.026183
0.892399 0.029575
0.875726 0.032854
0.858976 0.036020
0.842167 0.039086
0.825301 0.042060
0.808388 0.044948
0.791443 0.047754
0.774473 0.050481
0.757480 0.053131
0.740467 0.055707
0.723438 0.058209
0.706395 0.060639
0.689340 0.062997
0.672275 0.065282
0.655202 0.067495
0.638122 0.069634
0.621038 0.071698
0.603951 0.073687
0.586861 0.075598
0.569769 0.077429
0.552677 0.079178
0.535591 0.080841
0.518517 0.082414
0.501464 0.083889
0.484434 0.085256
0.467431 0.086510
0.450451 0.087643
0.433485 0.088648
0.416525 0.089517
0.399586 0.090242
0.382712 0.090813
0.365957 0.091208
0.349332 0.091394
0.332795 0.091343
0.316280 0.091043
0.299749 0.090498
0.283208 0.089711
0.266683 0.088683
0.250191 0.087413
0.233730 0.085897
0.217309 0.084133
0.200963 0.082120
0.184750 0.079851
0.168704 0.077302
0.152821 0.074448
0.137105 0.071268
0.121603 0.067751
0.106404 0.063882
0.091606 0.059644
0.077333 0.055020
0.063774 0.050020
0.051234 0.044662
0.040072 0.039061
0.030358 0.033345
0.021996 0.027598
0.015543 0.022432
0.011000 0.018229
0.007667 0.014623
0.005059 0.011284
0.003037 0.008173
0.001596 0.005492
0.000630 0.003272
0.000079 0.001346
-0.000090 -0.000490
0.000262 -0.002353
0.001239 -0.004217
0.002813 -0.006023
0.004964 -0.007659
0.007838 -0.009107
0.011681 -0.010448
0.016762 -0.011941
0.023429 -0.013633
0.031919 -0.015330
0.042822 -0.016923
0.056463 -0.018419
0.071589 -0.019656
0.087227 -0.020547
0.103386 -0.021139
0.119891 -0.021469
0.136481 -0.021556
0.152989 -0.021412
0.169397 -0.021009
0.185829 -0.020323
0.202443 -0.019386
0.219297 -0.018251
0.236349 -0.016968
0.253532 -0.015582
0.270799 -0.014136
0.288114 -0.012649
0.305459 -0.011133
0.322824 -0.009600
0.340205 -0.008058
0.357590 -0.006518
0.374966 -0.004989
0.392322 -0.003477
0.409658 -0.001988
0.426979 -0.000526
0.444292 0.000905
0.461599 0.002303
0.478904 0.003667
0.496210 0.004997
0.513523 0.006292
0.530849 0.007555
0.548190 0.008789
0.565537 0.009995
0.582872 0.011172
0.600181 0.012311
0.617454 0.013402
0.634688 0.014437
0.651883 0.015404
0.669044 0.016293
0.686176 0.017094
0.703283 0.017800
0.720366 0.018404
0.737424 0.018897
0.754454 0.019273
0.771449 0.019521
0.788406 0.019634
0.805321 0.019601
0.822193 0.019412
0.839018 0.019055
0.855778 0.018519
0.872448 0.017789
0.889003 0.016844
0.905418 0.015661
0.921638 0.014216
0.937541 0.012480
0.952913 0.010404
0.967486 0.007945
0.981065 0.005062
0.993195 0.001950
1.000000 0.000000

Copy and paste this into wordpad, and remove the extesion. X-foil won't recognize anything with an extension like "pm001.txt". It needs to be "pm001" in order to work. Thanks in advance!

[e-sailpilot86]

What I think I know is that this airfoil will have a sharp laminar flow separation= bad stall. The Reynolds are set to 100000, says so in the scientific notation. I think the other line low on the graph represents the pitching moment. This wouldn't be a very good airfoil on the wingtips, it might be okay if thinned and put on a slope plane. So, is the L/D too high to be real or is this a breakthrough? :confused: :cool:

I think the text in the upper right is most helpful.

[Ollie]

The graph shows the pressure distribution along the chord and the lines above and below the airfoil profile show the boundary layer thickness.

Polars are plots of the coefficient of lift versus the coefficient of drag for a range of angles of attack. Also, the coefficient of lift and the pitching moment coefficient vs angle of attack.

An L/D of a little more than 50, at a reynolds number of 100,000, puts this airfoil up there with the best for thermal soaring. The "hook" in the trailing edge acts like an airfoil with the flaps down a little. The difference in pressure near the trailing edge results in a thick, turbulent wake. Removing the "hook" will thin the wake and reduce drag a bit. This would benefit a slope soarer in all but the lightest of conditions. The bump in the upper surface pressure distribution is indicitive if a laminar seperation bubble. Modifying the contour to straighten the upper surface pressure distribution would also reduce drag a bit and possibly improve stall characteristics.

[e-sailpilot86]

Thanks again Ollie! Now I've got to figure out how to alter the airfoil shape a bit to see if it helps any bit. BTW, I put the parameters in a second time, I didn't have things set up perfectly. It had a L/D of 60. Now to know if that's even going to be the case I need a windtunnel... ;)