GOE 101 AIRFOIL (goe101-il) Xfoil prediction polar at RE=200,000 Ncrit=9
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Airfoil: GOE 101 AIRFOIL (goe101-il) Reynolds number: 200,000 Max Cl/Cd: 77.93 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe101-il-200000.txt Download as CSV file: xf-goe101-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 101 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3745 0.09743 0.09393 -0.0170 1.0000 0.0278
-8.000 -0.3730 0.09521 0.09177 -0.0192 1.0000 0.0282
-7.750 -0.3747 0.09324 0.08987 -0.0215 1.0000 0.0284
-7.500 -0.3694 0.09045 0.08712 -0.0253 1.0000 0.0286
-7.250 -0.3602 0.08738 0.08407 -0.0303 1.0000 0.0287
-7.000 -0.3499 0.08394 0.08063 -0.0334 1.0000 0.0288
-6.750 -0.3520 0.07786 0.07463 -0.0312 1.0000 0.0295
-6.500 -0.3469 0.07454 0.07134 -0.0293 1.0000 0.0302
-6.250 -0.3394 0.07145 0.06827 -0.0296 1.0000 0.0308
-6.000 -0.3308 0.06837 0.06521 -0.0305 1.0000 0.0316
-5.750 -0.3212 0.06524 0.06207 -0.0318 1.0000 0.0325
-5.500 -0.3106 0.06213 0.05895 -0.0333 1.0000 0.0335
-5.250 -0.2991 0.05894 0.05574 -0.0348 1.0000 0.0349
-5.000 -0.2851 0.05575 0.05249 -0.0366 1.0000 0.0366
-4.750 -0.2543 0.05334 0.04973 -0.0411 1.0000 0.0391
-4.500 -0.2372 0.04864 0.04483 -0.0430 0.9990 0.0398
-4.250 -0.2114 0.04381 0.04016 -0.0465 0.9939 0.0426
-4.000 -0.1686 0.04029 0.03643 -0.0518 0.9871 0.0489
-3.750 -0.0775 0.01849 0.01457 -0.0636 0.9641 0.0588
-3.500 -0.0366 0.01497 0.01064 -0.0678 0.9569 0.0683
-3.250 -0.0023 0.01260 0.00800 -0.0702 0.9454 0.0818
-3.000 -0.0007 0.02695 0.02199 -0.0691 0.9579 0.0842
-2.750 0.0370 0.02470 0.01951 -0.0717 0.9474 0.0984
-2.500 0.0745 0.02317 0.01761 -0.0739 0.9353 0.1234
-2.250 0.1069 0.02132 0.01578 -0.0754 0.9201 0.1464
-1.500 0.2118 0.01528 0.00814 -0.0739 0.8617 0.0705
-1.250 0.2399 0.01444 0.00696 -0.0729 0.8373 0.0674
-1.000 0.2667 0.01418 0.00646 -0.0718 0.8113 0.0655
-0.750 0.2924 0.01338 0.00549 -0.0708 0.7855 0.0651
-0.500 0.3179 0.01286 0.00483 -0.0699 0.7598 0.0653
-0.250 0.3430 0.01233 0.00419 -0.0690 0.7362 0.0661
0.000 0.3680 0.01193 0.00374 -0.0682 0.7148 0.0705
0.250 0.3936 0.01175 0.00347 -0.0676 0.6943 0.0732
0.500 0.4195 0.01164 0.00325 -0.0670 0.6755 0.0758
0.750 0.4457 0.01159 0.00305 -0.0664 0.6589 0.0804
1.000 0.4720 0.01146 0.00291 -0.0659 0.6441 0.1029
1.250 0.5065 0.00942 0.00289 -0.0674 0.6305 1.0000
1.500 0.5328 0.00961 0.00292 -0.0669 0.6178 1.0000
1.750 0.5591 0.00981 0.00299 -0.0665 0.6059 1.0000
2.000 0.5854 0.01001 0.00307 -0.0661 0.5948 1.0000
2.250 0.6117 0.01023 0.00316 -0.0657 0.5841 1.0000
2.500 0.6379 0.01041 0.00327 -0.0653 0.5730 1.0000
2.750 0.6641 0.01059 0.00342 -0.0649 0.5615 1.0000
3.000 0.6901 0.01078 0.00355 -0.0645 0.5497 1.0000
3.250 0.7161 0.01097 0.00368 -0.0641 0.5383 1.0000
3.500 0.7423 0.01119 0.00383 -0.0637 0.5288 1.0000
3.750 0.7687 0.01138 0.00409 -0.0634 0.5201 1.0000
4.000 0.7950 0.01162 0.00431 -0.0631 0.5123 1.0000
4.250 0.8212 0.01181 0.00454 -0.0627 0.5030 1.0000
4.500 0.8471 0.01200 0.00477 -0.0623 0.4924 1.0000
4.750 0.8729 0.01220 0.00502 -0.0618 0.4820 1.0000
5.000 0.8989 0.01242 0.00527 -0.0614 0.4727 1.0000
5.250 0.9246 0.01261 0.00556 -0.0610 0.4621 1.0000
5.500 0.9492 0.01273 0.00571 -0.0603 0.4450 1.0000
5.750 0.9732 0.01280 0.00586 -0.0595 0.4210 1.0000
6.000 0.9974 0.01296 0.00605 -0.0588 0.3998 1.0000
6.250 1.0209 0.01310 0.00622 -0.0580 0.3634 1.0000
6.500 1.0434 0.01343 0.00646 -0.0571 0.3087 1.0000
6.750 1.0593 0.01464 0.00710 -0.0556 0.2101 1.0000
7.000 1.0692 0.01688 0.00847 -0.0536 0.0824 1.0000
7.250 1.0840 0.01858 0.00982 -0.0518 0.0399 1.0000
7.500 1.1031 0.01964 0.01104 -0.0504 0.0348 1.0000
7.750 1.1187 0.02104 0.01255 -0.0487 0.0315 1.0000
8.000 1.1325 0.02255 0.01419 -0.0467 0.0293 1.0000
8.250 1.1481 0.02382 0.01559 -0.0451 0.0276 1.0000
8.500 1.1622 0.02534 0.01722 -0.0431 0.0264 1.0000
8.750 1.1767 0.02698 0.01898 -0.0413 0.0256 1.0000
9.000 1.1924 0.02881 0.02091 -0.0396 0.0250 1.0000
9.250 1.2101 0.03086 0.02309 -0.0382 0.0247 1.0000
9.500 1.2287 0.03311 0.02550 -0.0370 0.0243 1.0000
9.750 1.2452 0.03550 0.02800 -0.0359 0.0230 1.0000
10.000 1.2614 0.03883 0.03152 -0.0349 0.0223 1.0000
10.250 1.2751 0.04150 0.03451 -0.0333 0.0225 1.0000
10.500 1.2843 0.04435 0.03781 -0.0312 0.0232 1.0000
10.750 1.2805 0.04909 0.04323 -0.0278 0.0253 1.0000
11.000 1.2709 0.05373 0.04830 -0.0248 0.0270 1.0000
11.250 1.2570 0.05808 0.05293 -0.0219 0.0282 1.0000
11.500 1.1586 0.05356 0.04886 -0.0155 0.0271 1.0000
11.750 1.1370 0.05884 0.05437 -0.0153 0.0276 1.0000
12.000 1.1143 0.06458 0.06031 -0.0159 0.0279 1.0000
12.250 1.0902 0.07085 0.06678 -0.0174 0.0281 1.0000
12.500 1.0659 0.07749 0.07358 -0.0197 0.0282 1.0000
12.750 1.0409 0.08452 0.08078 -0.0228 0.0283 1.0000
13.000 1.0147 0.09190 0.08833 -0.0266 0.0284 1.0000
13.250 0.9874 0.09960 0.09618 -0.0312 0.0283 1.0000
13.500 0.9589 0.10767 0.10441 -0.0364 0.0283 1.0000
13.750 0.9247 0.11612 0.11300 -0.0426 0.0282 1.0000
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Polar data table (+)
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