GOE 622 AIRFOIL (goe622-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 622 AIRFOIL (goe622-il) Reynolds number: 200,000 Max Cl/Cd: 69.82 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe622-il-200000.txt Download as CSV file: xf-goe622-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 622 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3870 0.10179 0.09839 -0.0158 1.0000 0.0465
-9.500 -0.3864 0.09802 0.09464 -0.0168 1.0000 0.0481
-9.250 -0.5084 0.10317 0.09967 -0.0116 1.0000 0.0425
-9.000 -0.4971 0.10070 0.09717 -0.0100 1.0000 0.0446
-8.750 -0.4927 0.09733 0.09382 -0.0112 1.0000 0.0463
-8.500 -0.4902 0.09364 0.09016 -0.0131 1.0000 0.0478
-8.250 -0.4894 0.08972 0.08629 -0.0158 1.0000 0.0496
-8.000 -0.4936 0.08535 0.08198 -0.0211 1.0000 0.0515
-7.750 -0.4940 0.07860 0.07521 -0.0350 1.0000 0.0527
-7.500 -0.4895 0.07322 0.06972 -0.0404 1.0000 0.0530
-7.250 -0.4897 0.06719 0.06377 -0.0397 1.0000 0.0541
-7.000 -0.4807 0.06464 0.06125 -0.0382 1.0000 0.0553
-6.750 -0.4715 0.06167 0.05827 -0.0381 1.0000 0.0568
-6.500 -0.4625 0.05802 0.05456 -0.0392 1.0000 0.0591
-6.250 -0.4525 0.05333 0.04914 -0.0439 1.0000 0.0655
-6.000 -0.4470 0.04719 0.04319 -0.0433 1.0000 0.0672
-5.750 -0.4409 0.03413 0.02933 -0.0428 1.0000 0.0379
-5.500 -0.4283 0.02890 0.02353 -0.0415 1.0000 0.0362
-5.250 -0.4111 0.02586 0.02001 -0.0401 1.0000 0.0378
-5.000 -0.3915 0.02281 0.01642 -0.0388 1.0000 0.0383
-4.750 -0.3696 0.02053 0.01367 -0.0377 1.0000 0.0393
-4.500 -0.3465 0.01917 0.01194 -0.0366 1.0000 0.0408
-4.250 -0.3229 0.01729 0.00994 -0.0362 1.0000 0.0457
-4.000 -0.2987 0.01638 0.00886 -0.0355 1.0000 0.0505
-3.750 -0.2732 0.01516 0.00754 -0.0350 1.0000 0.0565
-3.500 -0.2488 0.01479 0.00708 -0.0345 1.0000 0.0658
-3.250 -0.2234 0.01425 0.00656 -0.0342 1.0000 0.0773
-3.000 -0.1985 0.01396 0.00624 -0.0339 1.0000 0.0918
-2.750 -0.1635 0.01340 0.00579 -0.0358 0.9974 0.1164
-2.500 -0.1227 0.01285 0.00532 -0.0387 0.9934 0.1493
-2.250 -0.0831 0.01242 0.00500 -0.0415 0.9879 0.1800
-2.000 -0.0411 0.01199 0.00475 -0.0448 0.9835 0.2191
-1.750 -0.0048 0.01110 0.00455 -0.0472 0.9770 0.3637
-1.500 0.0211 0.00942 0.00470 -0.0460 0.9734 0.8408
-1.250 0.0778 0.00927 0.00451 -0.0514 0.9710 1.0000
-1.000 0.1180 0.00923 0.00433 -0.0540 0.9625 1.0000
-0.750 0.1649 0.00915 0.00412 -0.0579 0.9574 1.0000
-0.500 0.2030 0.00907 0.00395 -0.0600 0.9476 1.0000
-0.250 0.2423 0.00896 0.00378 -0.0622 0.9385 1.0000
0.000 0.2809 0.00884 0.00358 -0.0641 0.9291 1.0000
0.250 0.3132 0.00873 0.00342 -0.0645 0.9152 1.0000
0.500 0.3402 0.00864 0.00326 -0.0636 0.8950 1.0000
0.750 0.3658 0.00857 0.00310 -0.0624 0.8738 1.0000
1.000 0.3900 0.00857 0.00303 -0.0611 0.8519 1.0000
1.250 0.4147 0.00859 0.00296 -0.0599 0.8310 1.0000
1.500 0.4394 0.00862 0.00293 -0.0588 0.8085 1.0000
1.750 0.4648 0.00868 0.00290 -0.0579 0.7879 1.0000
2.000 0.4902 0.00876 0.00293 -0.0571 0.7659 1.0000
2.250 0.5156 0.00885 0.00293 -0.0563 0.7435 1.0000
2.500 0.5411 0.00897 0.00298 -0.0556 0.7198 1.0000
2.750 0.5668 0.00911 0.00307 -0.0549 0.6981 1.0000
3.000 0.5922 0.00928 0.00316 -0.0542 0.6749 1.0000
3.250 0.6176 0.00948 0.00328 -0.0535 0.6513 1.0000
3.500 0.6431 0.00970 0.00343 -0.0528 0.6289 1.0000
3.750 0.6682 0.00994 0.00362 -0.0521 0.6042 1.0000
4.000 0.6932 0.01018 0.00381 -0.0514 0.5775 1.0000
4.250 0.7181 0.01044 0.00402 -0.0507 0.5504 1.0000
4.500 0.7430 0.01069 0.00425 -0.0500 0.5225 1.0000
4.750 0.7673 0.01099 0.00451 -0.0492 0.4886 1.0000
5.000 0.7902 0.01134 0.00472 -0.0482 0.4324 1.0000
5.250 0.8110 0.01197 0.00499 -0.0470 0.3532 1.0000
5.500 0.8315 0.01277 0.00544 -0.0460 0.2868 1.0000
5.750 0.8533 0.01350 0.00595 -0.0451 0.2375 1.0000
6.000 0.8762 0.01413 0.00649 -0.0445 0.2005 1.0000
6.250 0.8989 0.01479 0.00701 -0.0438 0.1622 1.0000
6.500 0.9201 0.01567 0.00765 -0.0430 0.1183 1.0000
6.750 0.9365 0.01737 0.00885 -0.0414 0.0474 1.0000
7.000 0.9555 0.01869 0.01020 -0.0399 0.0372 1.0000
7.250 0.9766 0.01965 0.01131 -0.0387 0.0332 1.0000
7.500 0.9953 0.02086 0.01255 -0.0374 0.0294 1.0000
7.750 1.0124 0.02240 0.01418 -0.0357 0.0276 1.0000
8.000 1.0314 0.02381 0.01571 -0.0343 0.0266 1.0000
8.250 1.0507 0.02544 0.01746 -0.0329 0.0257 1.0000
8.500 1.0707 0.02732 0.01947 -0.0316 0.0251 1.0000
8.750 1.0910 0.02949 0.02186 -0.0303 0.0248 1.0000
9.000 1.1103 0.03160 0.02421 -0.0290 0.0240 1.0000
9.250 1.1276 0.03345 0.02625 -0.0278 0.0228 1.0000
9.500 1.1431 0.03590 0.02894 -0.0264 0.0222 1.0000
9.750 1.1550 0.03920 0.03267 -0.0244 0.0225 1.0000
10.000 1.1580 0.04372 0.03780 -0.0215 0.0235 1.0000
10.250 1.1516 0.04881 0.04348 -0.0183 0.0248 1.0000
10.500 1.1415 0.05351 0.04857 -0.0154 0.0258 1.0000
10.750 1.1267 0.05767 0.05301 -0.0124 0.0266 1.0000
11.000 1.1085 0.06184 0.05739 -0.0100 0.0271 1.0000
11.250 1.0892 0.06663 0.06236 -0.0090 0.0276 1.0000
11.500 0.9786 0.05968 0.05570 0.0004 0.0263 1.0000
11.750 0.9529 0.06537 0.06156 -0.0006 0.0264 1.0000
12.000 0.9260 0.07182 0.06817 -0.0027 0.0265 1.0000
12.250 0.8984 0.07894 0.07544 -0.0058 0.0265 1.0000
12.500 0.8704 0.08673 0.08335 -0.0097 0.0266 1.0000
12.750 0.8402 0.09545 0.09218 -0.0146 0.0267 1.0000
13.000 0.8085 0.10517 0.10198 -0.0201 0.0270 1.0000
13.250 0.7784 0.11467 0.11148 -0.0249 0.0276 1.0000
13.500 0.7154 0.13176 0.12857 -0.0314 0.0374 1.0000
15.250 0.8921 0.19338 0.18985 -0.0779 0.0460 1.0000
15.500 0.8960 0.19774 0.19420 -0.0806 0.0440 1.0000
15.750 0.6848 0.17357 0.17035 -0.0444 0.0489 1.0000
16.000 0.6816 0.17691 0.17369 -0.0465 0.0487 1.0000
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Polar data table (+)
Polar graphs
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