GOE 626 AIRFOIL (goe626-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 626 AIRFOIL (goe626-il) Reynolds number: 100,000 Max Cl/Cd: 48.93 at α=8.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe626-il-100000.txt Download as CSV file: xf-goe626-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 626 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.6088 0.06848 0.06314 -0.0663 1.0000 0.0906
-10.500 -0.6668 0.06262 0.05729 -0.0661 1.0000 0.0897
-10.250 -0.7296 0.05479 0.04932 -0.0704 1.0000 0.0883
-10.000 -0.7571 0.04736 0.04146 -0.0773 1.0000 0.0882
-9.750 -0.7593 0.04320 0.03691 -0.0798 1.0000 0.0891
-9.500 -0.7528 0.03988 0.03313 -0.0816 1.0000 0.0906
-9.250 -0.7390 0.03844 0.03180 -0.0804 1.0000 0.0925
-9.000 -0.7243 0.03719 0.03059 -0.0796 1.0000 0.0947
-8.750 -0.7090 0.03555 0.02882 -0.0796 1.0000 0.0972
-8.500 -0.6922 0.03374 0.02674 -0.0799 1.0000 0.1001
-8.250 -0.6761 0.03265 0.02575 -0.0790 1.0000 0.1031
-8.000 -0.6593 0.03174 0.02492 -0.0783 1.0000 0.1073
-7.750 -0.6246 0.03059 0.02380 -0.0811 0.9953 0.1155
-7.500 -0.5812 0.02940 0.02271 -0.0858 0.9872 0.1331
-7.250 -0.5349 0.02914 0.02276 -0.0904 0.9793 0.1652
-7.000 -0.4900 0.03026 0.02395 -0.0933 0.9689 0.1905
-6.750 -0.4484 0.03333 0.02735 -0.0932 0.9573 0.2006
-6.250 -0.3576 0.03778 0.03188 -0.0957 0.9374 0.2222
-6.000 -0.3222 0.04134 0.03566 -0.0928 0.9250 0.2253
-5.750 -0.2767 0.04104 0.03517 -0.0966 0.9172 0.2382
-5.500 -0.2399 0.04104 0.03504 -0.0986 0.9061 0.2506
-5.250 -0.2048 0.04360 0.03777 -0.0960 0.8963 0.2551
-5.000 -0.1637 0.04298 0.03699 -0.0993 0.8877 0.2703
-4.750 -0.1163 0.04573 0.03993 -0.0969 0.8840 0.2753
-4.250 -0.0416 0.04196 0.03575 -0.1049 0.8663 0.3033
-4.000 -0.0223 0.04185 0.03570 -0.1025 0.8529 0.3089
-3.750 0.0308 0.03940 0.03294 -0.1095 0.8490 0.3247
-3.500 0.0508 0.03990 0.03359 -0.1058 0.8372 0.3290
-3.250 0.1019 0.03790 0.03135 -0.1117 0.8322 0.3444
-3.000 0.1477 0.03786 0.03142 -0.1120 0.8290 0.3551
-2.750 0.1708 0.03649 0.02989 -0.1131 0.8153 0.3648
-2.500 0.2173 0.03547 0.02884 -0.1153 0.8109 0.3734
-2.250 0.2468 0.03379 0.02689 -0.1181 0.7977 0.3789
-2.000 0.2932 0.03215 0.02513 -0.1215 0.7921 0.3844
-1.750 0.3157 0.03161 0.02459 -0.1204 0.7790 0.3888
-1.500 0.3637 0.03006 0.02282 -0.1245 0.7725 0.3934
-1.250 0.3904 0.02916 0.02174 -0.1253 0.7589 0.3968
-1.000 0.4286 0.02803 0.02034 -0.1279 0.7490 0.4010
-0.750 0.4595 0.02725 0.01953 -0.1282 0.7378 0.4037
-0.500 0.4871 0.02675 0.01896 -0.1281 0.7256 0.4067
-0.250 0.5264 0.02600 0.01803 -0.1300 0.7166 0.4112
0.000 0.5514 0.02564 0.01755 -0.1298 0.7033 0.4145
0.250 0.5844 0.02515 0.01685 -0.1309 0.6923 0.4171
0.500 0.6177 0.02468 0.01618 -0.1320 0.6815 0.4203
0.750 0.6421 0.02452 0.01603 -0.1312 0.6697 0.4235
1.000 0.6773 0.02424 0.01560 -0.1323 0.6607 0.4272
1.250 0.6998 0.02428 0.01561 -0.1315 0.6487 0.4299
1.500 0.7338 0.02413 0.01527 -0.1325 0.6398 0.4331
1.750 0.7574 0.02425 0.01532 -0.1319 0.6285 0.4367
2.000 0.7894 0.02424 0.01516 -0.1326 0.6201 0.4405
2.250 0.8108 0.02443 0.01542 -0.1315 0.6101 0.4439
2.500 0.8446 0.02448 0.01533 -0.1324 0.6028 0.4478
2.750 0.8633 0.02479 0.01572 -0.1310 0.5925 0.4510
3.000 0.8957 0.02483 0.01558 -0.1316 0.5843 0.4553
3.250 0.9157 0.02513 0.01594 -0.1303 0.5744 0.4592
3.500 0.9440 0.02518 0.01597 -0.1301 0.5663 0.4642
3.750 0.9673 0.02549 0.01632 -0.1294 0.5581 0.4695
4.000 0.9919 0.02569 0.01651 -0.1287 0.5496 0.4749
4.250 1.0225 0.02578 0.01652 -0.1290 0.5427 0.4802
4.500 1.0396 0.02614 0.01705 -0.1272 0.5338 0.4853
4.750 1.0698 0.02622 0.01706 -0.1274 0.5271 0.4927
5.000 1.0898 0.02657 0.01752 -0.1260 0.5192 0.4996
5.250 1.1142 0.02666 0.01766 -0.1252 0.5109 0.5088
5.500 1.1432 0.02677 0.01773 -0.1252 0.5038 0.5196
5.750 1.1611 0.02715 0.01827 -0.1235 0.4955 0.5305
6.000 1.1921 0.02717 0.01829 -0.1238 0.4894 0.5453
6.250 1.2113 0.02762 0.01892 -0.1224 0.4823 0.5620
6.500 1.2333 0.02785 0.01933 -0.1213 0.4750 0.5879
6.750 1.2673 0.02748 0.01921 -0.1219 0.4691 0.6650
7.000 1.2766 0.02769 0.01996 -0.1188 0.4605 1.0000
7.250 1.3079 0.02775 0.01979 -0.1190 0.4526 1.0000
7.500 1.3239 0.02830 0.02038 -0.1170 0.4437 1.0000
7.750 1.3513 0.02832 0.02026 -0.1165 0.4346 1.0000
8.000 1.3666 0.02880 0.02077 -0.1144 0.4250 1.0000
8.250 1.3950 0.02875 0.02055 -0.1140 0.4154 1.0000
8.500 1.4051 0.02934 0.02126 -0.1111 0.4055 1.0000
8.750 1.4355 0.02934 0.02107 -0.1111 0.3962 1.0000
9.000 1.4401 0.03003 0.02194 -0.1075 0.3861 1.0000
9.250 1.4696 0.03011 0.02183 -0.1074 0.3766 1.0000
9.500 1.4711 0.03081 0.02271 -0.1033 0.3662 1.0000
9.750 1.4908 0.03111 0.02293 -0.1018 0.3565 1.0000
10.000 1.4974 0.03167 0.02355 -0.0985 0.3462 1.0000
10.250 1.5068 0.03222 0.02412 -0.0956 0.3369 1.0000
10.500 1.5159 0.03270 0.02456 -0.0927 0.3271 1.0000
10.750 1.5179 0.03352 0.02545 -0.0890 0.3179 1.0000
11.000 1.5297 0.03402 0.02589 -0.0867 0.3089 1.0000
11.250 1.5290 0.03512 0.02709 -0.0831 0.3003 1.0000
11.500 1.5389 0.03574 0.02766 -0.0808 0.2918 1.0000
11.750 1.5360 0.03705 0.02908 -0.0774 0.2835 1.0000
12.000 1.5434 0.03793 0.02994 -0.0751 0.2756 1.0000
12.250 1.5439 0.03936 0.03145 -0.0725 0.2683 1.0000
12.500 1.5462 0.04067 0.03281 -0.0702 0.2608 1.0000
12.750 1.5502 0.04203 0.03419 -0.0681 0.2537 1.0000
13.000 1.5467 0.04389 0.03617 -0.0658 0.2463 1.0000
13.250 1.5535 0.04518 0.03739 -0.0642 0.2391 1.0000
13.500 1.5447 0.04766 0.04005 -0.0621 0.2319 1.0000
13.750 1.5519 0.04902 0.04129 -0.0608 0.2240 1.0000
14.000 1.5387 0.05215 0.04463 -0.0591 0.2165 1.0000
14.250 1.5423 0.05395 0.04633 -0.0580 0.2079 1.0000
14.500 1.5299 0.05737 0.04988 -0.0569 0.1995 1.0000
14.750 1.5276 0.06005 0.05251 -0.0561 0.1902 1.0000
15.000 1.5388 0.06152 0.05372 -0.0551 0.1793 1.0000
15.250 1.5195 0.06627 0.05871 -0.0548 0.1717 1.0000
15.500 1.5347 0.06765 0.05987 -0.0540 0.1630 1.0000
15.750 1.5222 0.07198 0.06441 -0.0540 0.1572 1.0000
16.000 1.5528 0.07188 0.06403 -0.0530 0.1506 1.0000
16.250 1.5382 0.07659 0.06900 -0.0532 0.1473 1.0000
16.500 1.5271 0.08106 0.07369 -0.0537 0.1440 1.0000
16.750 1.5290 0.08408 0.07679 -0.0538 0.1408 1.0000
17.000 1.5597 0.08391 0.07647 -0.0528 0.1375 1.0000
17.250 1.5596 0.08746 0.08015 -0.0530 0.1355 1.0000
17.500 1.5319 0.09420 0.08722 -0.0548 0.1343 1.0000
17.750 1.4970 0.10240 0.09576 -0.0577 0.1333 1.0000
18.000 1.4313 0.11604 0.10982 -0.0641 0.1328 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 626 AIRFOIL (goe626-il)