GOE 164 (MVA MK.10) AIRFOIL (goe164-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 164 (MVA MK.10) AIRFOIL (goe164-il) Reynolds number: 200,000 Max Cl/Cd: 77.35 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe164-il-200000-n5.txt Download as CSV file: xf-goe164-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 164 (MVA MK.10) 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.2685 0.11373 0.11036 -0.0269 1.0000 0.0215
-8.750 -0.2673 0.11136 0.10803 -0.0265 1.0000 0.0218
-8.500 -0.2669 0.10905 0.10577 -0.0260 1.0000 0.0222
-8.250 -0.2622 0.10630 0.10306 -0.0267 0.9992 0.0226
-8.000 -0.2447 0.10231 0.09906 -0.0313 0.9960 0.0232
-7.750 -0.2285 0.09840 0.09516 -0.0355 0.9923 0.0235
-7.500 -0.2141 0.09482 0.09160 -0.0439 0.9855 0.0260
-7.250 -0.1954 0.09058 0.08738 -0.0500 0.9786 0.0262
-7.000 -0.1752 0.08621 0.08301 -0.0554 0.9716 0.0262
-6.500 -0.1434 0.07608 0.07287 -0.0611 0.9591 0.0175
-6.250 -0.1190 0.07146 0.06824 -0.0675 0.9521 0.0170
-5.750 -0.0741 0.06187 0.05861 -0.0796 0.9283 0.0163
-5.500 -0.0477 0.05630 0.05300 -0.0870 0.9139 0.0162
-5.250 -0.0121 0.04916 0.04577 -0.0975 0.9001 0.0163
-5.000 0.0880 0.02266 0.01764 -0.1372 0.8945 0.0177
-4.750 0.1284 0.01955 0.01382 -0.1410 0.8844 0.0185
-4.500 0.1686 0.01739 0.01127 -0.1441 0.8740 0.0198
-4.250 0.2119 0.01635 0.01003 -0.1475 0.8635 0.0222
-4.000 0.2559 0.01530 0.00867 -0.1507 0.8521 0.0256
-3.750 0.2994 0.01419 0.00734 -0.1539 0.8386 0.0288
-3.500 0.3397 0.01364 0.00658 -0.1563 0.8221 0.0347
-3.250 0.3755 0.01301 0.00573 -0.1579 0.8057 0.0408
-3.000 0.4077 0.01271 0.00520 -0.1585 0.7893 0.0484
-2.750 0.4379 0.01231 0.00461 -0.1589 0.7723 0.0594
-2.500 0.4666 0.01195 0.00421 -0.1590 0.7552 0.0812
-2.250 0.4942 0.01195 0.00424 -0.1587 0.7382 0.1340
-2.000 0.5214 0.01211 0.00423 -0.1582 0.7226 0.1514
-1.750 0.5483 0.01224 0.00420 -0.1578 0.7075 0.1627
-1.500 0.5750 0.01242 0.00419 -0.1573 0.6925 0.1741
-1.250 0.6014 0.01250 0.00420 -0.1568 0.6774 0.1808
-1.000 0.6278 0.01261 0.00413 -0.1563 0.6616 0.1886
-0.750 0.6539 0.01266 0.00413 -0.1557 0.6455 0.1955
-0.500 0.6801 0.01269 0.00404 -0.1552 0.6296 0.2002
-0.250 0.7062 0.01272 0.00393 -0.1547 0.6135 0.2036
0.000 0.7320 0.01272 0.00388 -0.1541 0.5977 0.2071
0.250 0.7576 0.01277 0.00385 -0.1536 0.5808 0.2111
0.500 0.7830 0.01286 0.00382 -0.1529 0.5621 0.2151
0.750 0.8080 0.01297 0.00381 -0.1522 0.5418 0.2189
1.000 0.8327 0.01309 0.00385 -0.1515 0.5214 0.2229
1.250 0.8575 0.01323 0.00390 -0.1508 0.5005 0.2277
1.750 0.9066 0.01357 0.00407 -0.1495 0.4573 0.2398
2.000 0.9314 0.01375 0.00418 -0.1488 0.4372 0.2471
2.250 0.9559 0.01392 0.00432 -0.1482 0.4190 0.2547
2.500 0.9804 0.01412 0.00448 -0.1475 0.4024 0.2646
3.000 1.0294 0.01451 0.00489 -0.1463 0.3755 0.3027
3.250 1.0535 0.01471 0.00513 -0.1456 0.3641 0.3452
3.500 1.0776 0.01489 0.00538 -0.1449 0.3544 0.4039
3.750 1.0976 0.01422 0.00558 -0.1432 0.3467 1.0000
4.000 1.1216 0.01454 0.00583 -0.1425 0.3387 1.0000
4.250 1.1462 0.01482 0.00609 -0.1419 0.3313 1.0000
4.500 1.1701 0.01515 0.00637 -0.1412 0.3239 1.0000
4.750 1.1943 0.01544 0.00665 -0.1405 0.3165 1.0000
5.000 1.2178 0.01578 0.00697 -0.1398 0.3087 1.0000
5.250 1.2418 0.01608 0.00728 -0.1391 0.3019 1.0000
5.500 1.2654 0.01639 0.00761 -0.1383 0.2946 1.0000
5.750 1.2888 0.01671 0.00795 -0.1376 0.2877 1.0000
6.000 1.3123 0.01702 0.00831 -0.1368 0.2802 1.0000
6.250 1.3354 0.01735 0.00868 -0.1360 0.2733 1.0000
6.500 1.3584 0.01767 0.00905 -0.1352 0.2652 1.0000
6.750 1.3810 0.01802 0.00945 -0.1343 0.2571 1.0000
7.000 1.4028 0.01839 0.00988 -0.1333 0.2480 1.0000
7.250 1.4248 0.01876 0.01031 -0.1323 0.2380 1.0000
7.500 1.4448 0.01922 0.01077 -0.1311 0.2228 1.0000
7.750 1.4610 0.01992 0.01134 -0.1293 0.1986 1.0000
8.250 1.4902 0.02152 0.01282 -0.1252 0.1701 1.0000
8.500 1.5041 0.02226 0.01358 -0.1230 0.1599 1.0000
8.750 1.5169 0.02309 0.01441 -0.1207 0.1473 1.0000
9.000 1.5325 0.02375 0.01512 -0.1189 0.1269 1.0000
9.250 1.5304 0.02559 0.01658 -0.1148 0.0834 1.0000
9.500 1.5356 0.02696 0.01799 -0.1116 0.0719 1.0000
9.750 1.5377 0.02858 0.01952 -0.1083 0.0345 1.0000
10.000 1.5362 0.03054 0.02141 -0.1046 0.0219 1.0000
10.250 1.5385 0.03226 0.02323 -0.1017 0.0180 1.0000
10.500 1.5431 0.03386 0.02501 -0.0992 0.0161 1.0000
10.750 1.5468 0.03558 0.02690 -0.0969 0.0147 1.0000
11.000 1.5484 0.03758 0.02907 -0.0947 0.0137 1.0000
11.250 1.5475 0.03993 0.03159 -0.0926 0.0130 1.0000
11.500 1.5438 0.04272 0.03456 -0.0908 0.0123 1.0000
11.750 1.5368 0.04605 0.03808 -0.0893 0.0119 1.0000
12.000 1.5285 0.04972 0.04194 -0.0882 0.0115 1.0000
12.250 1.5244 0.05307 0.04546 -0.0875 0.0112 1.0000
12.500 1.5188 0.05673 0.04929 -0.0871 0.0108 1.0000
12.750 1.5119 0.06072 0.05344 -0.0869 0.0104 1.0000
13.000 1.5038 0.06501 0.05790 -0.0870 0.0101 1.0000
13.250 1.4946 0.06960 0.06264 -0.0874 0.0098 1.0000
13.500 1.4850 0.07440 0.06759 -0.0880 0.0095 1.0000
13.750 1.4746 0.07945 0.07279 -0.0888 0.0093 1.0000
14.000 1.4638 0.08465 0.07813 -0.0898 0.0092 1.0000
14.250 1.4529 0.08998 0.08359 -0.0910 0.0090 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 164 (MVA MK.10) AIRFOIL (goe164-il)