GOE 622 AIRFOIL (goe622-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: GOE 622 AIRFOIL (goe622-il) Reynolds number: 100,000 Max Cl/Cd: 52.85 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe622-il-100000-n5.txt Download as CSV file: xf-goe622-il-100000-n5.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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4874 0.09042 0.08557 -0.0182 1.0000 0.0229
-8.250 -0.4880 0.08613 0.08134 -0.0206 1.0000 0.0227
-8.000 -0.4901 0.08177 0.07705 -0.0232 1.0000 0.0225
-7.750 -0.4905 0.07686 0.07221 -0.0273 1.0000 0.0223
-7.500 -0.4885 0.07135 0.06671 -0.0321 1.0000 0.0222
-7.250 -0.4853 0.06540 0.06074 -0.0368 1.0000 0.0222
-7.000 -0.4809 0.05910 0.05435 -0.0408 1.0000 0.0222
-6.750 -0.4755 0.05262 0.04770 -0.0438 1.0000 0.0225
-6.500 -0.4690 0.04600 0.04077 -0.0455 1.0000 0.0229
-6.250 -0.4603 0.03970 0.03397 -0.0462 1.0000 0.0234
-6.000 -0.4480 0.03436 0.02798 -0.0460 1.0000 0.0240
-5.750 -0.4319 0.03013 0.02306 -0.0453 1.0000 0.0245
-5.500 -0.4127 0.02699 0.01926 -0.0442 1.0000 0.0253
-5.250 -0.3932 0.02480 0.01685 -0.0435 1.0000 0.0272
-5.000 -0.3721 0.02362 0.01545 -0.0426 1.0000 0.0307
-4.750 -0.3493 0.02184 0.01327 -0.0416 1.0000 0.0339
-4.500 -0.3269 0.02041 0.01169 -0.0408 1.0000 0.0379
-4.250 -0.3038 0.01955 0.01059 -0.0400 1.0000 0.0453
-4.000 -0.2807 0.01852 0.00954 -0.0394 1.0000 0.0536
-3.750 -0.2573 0.01791 0.00887 -0.0388 1.0000 0.0657
-3.500 -0.2330 0.01753 0.00843 -0.0384 0.9996 0.0806
-3.250 -0.1961 0.01687 0.00764 -0.0404 0.9948 0.0971
-3.000 -0.1602 0.01633 0.00707 -0.0423 0.9889 0.1179
-2.750 -0.1245 0.01586 0.00657 -0.0442 0.9829 0.1403
-2.500 -0.0877 0.01541 0.00616 -0.0463 0.9771 0.1649
-2.250 -0.0535 0.01499 0.00581 -0.0479 0.9697 0.1972
-2.000 -0.0178 0.01447 0.00551 -0.0499 0.9632 0.2540
-1.750 0.0163 0.01371 0.00528 -0.0515 0.9562 0.3908
-1.500 0.0463 0.01267 0.00520 -0.0518 0.9499 0.6487
-1.250 0.0864 0.01207 0.00516 -0.0527 0.9445 0.9341
-1.000 0.1329 0.01199 0.00487 -0.0566 0.9378 1.0000
-0.750 0.1672 0.01198 0.00471 -0.0579 0.9275 1.0000
-0.500 0.2008 0.01198 0.00457 -0.0590 0.9168 1.0000
-0.250 0.2346 0.01198 0.00444 -0.0601 0.9062 1.0000
0.000 0.2684 0.01197 0.00433 -0.0612 0.8955 1.0000
0.250 0.2997 0.01199 0.00426 -0.0617 0.8831 1.0000
0.500 0.3294 0.01202 0.00422 -0.0619 0.8694 1.0000
0.750 0.3584 0.01205 0.00419 -0.0618 0.8551 1.0000
1.000 0.3861 0.01209 0.00419 -0.0615 0.8392 1.0000
1.250 0.4126 0.01213 0.00418 -0.0609 0.8195 1.0000
1.500 0.4396 0.01213 0.00411 -0.0602 0.7982 1.0000
1.750 0.4651 0.01218 0.00412 -0.0594 0.7738 1.0000
2.000 0.4914 0.01224 0.00413 -0.0587 0.7520 1.0000
2.250 0.5173 0.01233 0.00417 -0.0579 0.7293 1.0000
2.500 0.5431 0.01244 0.00424 -0.0572 0.7053 1.0000
2.750 0.5689 0.01257 0.00430 -0.0564 0.6809 1.0000
3.000 0.5944 0.01274 0.00439 -0.0556 0.6555 1.0000
3.250 0.6197 0.01295 0.00457 -0.0548 0.6300 1.0000
3.500 0.6449 0.01319 0.00476 -0.0541 0.6055 1.0000
3.750 0.6701 0.01345 0.00498 -0.0533 0.5811 1.0000
4.000 0.6950 0.01372 0.00524 -0.0526 0.5560 1.0000
4.250 0.7201 0.01400 0.00558 -0.0520 0.5322 1.0000
4.500 0.7449 0.01431 0.00590 -0.0513 0.5065 1.0000
4.750 0.7692 0.01463 0.00625 -0.0505 0.4756 1.0000
5.000 0.7928 0.01500 0.00660 -0.0496 0.4358 1.0000
5.250 0.8155 0.01547 0.00701 -0.0486 0.3872 1.0000
5.500 0.8355 0.01622 0.00744 -0.0473 0.3234 1.0000
5.750 0.8544 0.01720 0.00804 -0.0461 0.2551 1.0000
6.000 0.8752 0.01804 0.00870 -0.0452 0.2091 1.0000
6.250 0.8968 0.01883 0.00939 -0.0444 0.1683 1.0000
6.500 0.9164 0.01987 0.01023 -0.0434 0.1242 1.0000
6.750 0.9359 0.02100 0.01116 -0.0424 0.0674 1.0000
7.000 0.9529 0.02256 0.01247 -0.0409 0.0368 1.0000
7.250 0.9715 0.02386 0.01378 -0.0396 0.0276 1.0000
7.500 0.9902 0.02512 0.01517 -0.0383 0.0238 1.0000
7.750 1.0067 0.02661 0.01682 -0.0367 0.0218 1.0000
8.000 1.0212 0.02830 0.01868 -0.0349 0.0207 1.0000
8.250 1.0370 0.02981 0.02043 -0.0332 0.0196 1.0000
8.500 1.0522 0.03144 0.02228 -0.0316 0.0182 1.0000
8.750 1.0670 0.03327 0.02441 -0.0299 0.0172 1.0000
9.000 1.0815 0.03536 0.02673 -0.0282 0.0167 1.0000
9.250 1.0958 0.03768 0.02931 -0.0265 0.0163 1.0000
9.500 1.1088 0.04025 0.03217 -0.0248 0.0159 1.0000
9.750 1.1190 0.04308 0.03534 -0.0230 0.0156 1.0000
10.000 1.1255 0.04607 0.03868 -0.0210 0.0154 1.0000
10.250 1.1274 0.04911 0.04206 -0.0188 0.0151 1.0000
10.500 1.1233 0.05205 0.04531 -0.0163 0.0148 1.0000
10.750 1.1156 0.05516 0.04871 -0.0140 0.0145 1.0000
11.000 1.1051 0.05860 0.05242 -0.0124 0.0143 1.0000
11.250 1.0921 0.06250 0.05659 -0.0118 0.0142 1.0000
11.500 1.0771 0.06701 0.06135 -0.0123 0.0141 1.0000
11.750 1.0600 0.07236 0.06695 -0.0140 0.0142 1.0000
12.000 1.0410 0.07868 0.07350 -0.0172 0.0143 1.0000
12.250 1.0191 0.08647 0.08152 -0.0220 0.0147 1.0000
12.500 0.9933 0.09624 0.09149 -0.0286 0.0152 1.0000
12.750 0.9633 0.10879 0.10420 -0.0372 0.0160 1.0000
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Polar data table (+)
Polar graphs
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