GOE 372 AIRFOIL (goe372-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 372 AIRFOIL (goe372-il) Reynolds number: 100,000 Max Cl/Cd: 60.09 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe372-il-100000-n5.txt Download as CSV file: xf-goe372-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 372 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
------ -------- --------- --------- -------- -------- --------
-7.500 -0.3465 0.09017 0.08553 -0.0290 1.0000 0.0261
-7.250 -0.3515 0.08774 0.08321 -0.0283 1.0000 0.0253
-7.000 -0.3553 0.08497 0.08053 -0.0283 1.0000 0.0247
-6.750 -0.3577 0.08207 0.07773 -0.0289 1.0000 0.0243
-6.500 -0.3599 0.07905 0.07479 -0.0297 1.0000 0.0238
-6.250 -0.3613 0.07599 0.07179 -0.0305 1.0000 0.0236
-6.000 -0.3614 0.07277 0.06862 -0.0316 1.0000 0.0234
-5.750 -0.3533 0.06997 0.06583 -0.0330 0.9988 0.0244
-5.500 -0.3209 0.06517 0.06094 -0.0410 0.9924 0.0272
-5.250 -0.2834 0.05863 0.05426 -0.0511 0.9858 0.0283
-5.000 -0.2435 0.05149 0.04690 -0.0611 0.9788 0.0290
-4.750 -0.1983 0.04208 0.03706 -0.0721 0.9716 0.0313
-4.500 -0.1587 0.03794 0.03258 -0.0778 0.9662 0.0362
-4.250 -0.1182 0.02889 0.02238 -0.0840 0.9585 0.0425
-4.000 -0.0818 0.02832 0.02175 -0.0865 0.9523 0.0489
-3.750 -0.0474 0.02483 0.01738 -0.0887 0.9445 0.0561
-3.500 -0.0106 0.02302 0.01519 -0.0909 0.9401 0.0594
-3.250 0.0201 0.02150 0.01323 -0.0917 0.9328 0.0634
-3.000 0.0556 0.02003 0.01120 -0.0931 0.9280 0.0677
-2.750 0.0864 0.01940 0.01046 -0.0939 0.9208 0.0737
-2.500 0.1200 0.01858 0.00939 -0.0950 0.9148 0.0791
-2.250 0.1516 0.01771 0.00821 -0.0956 0.9080 0.0830
-2.000 0.1835 0.01697 0.00748 -0.0964 0.9012 0.0886
-1.750 0.2146 0.01641 0.00677 -0.0969 0.8939 0.0972
-1.500 0.2461 0.01593 0.00629 -0.0976 0.8868 0.1074
-1.250 0.2759 0.01562 0.00597 -0.0979 0.8784 0.1213
-1.000 0.3085 0.01534 0.00567 -0.0988 0.8715 0.1412
-0.750 0.3369 0.01505 0.00540 -0.0988 0.8620 0.1567
-0.500 0.3670 0.01475 0.00513 -0.0991 0.8538 0.1775
-0.250 0.3978 0.01445 0.00489 -0.0995 0.8452 0.2059
0.000 0.4265 0.01415 0.00476 -0.0996 0.8354 0.2585
0.250 0.4554 0.01348 0.00469 -0.0999 0.8266 0.4539
0.500 0.4983 0.01238 0.00451 -0.1026 0.8182 1.0000
0.750 0.5257 0.01247 0.00449 -0.1023 0.8074 1.0000
1.000 0.5534 0.01255 0.00447 -0.1020 0.7968 1.0000
1.250 0.5816 0.01263 0.00446 -0.1018 0.7866 1.0000
1.500 0.6081 0.01275 0.00453 -0.1014 0.7750 1.0000
1.750 0.6344 0.01288 0.00466 -0.1009 0.7631 1.0000
2.000 0.6608 0.01301 0.00478 -0.1004 0.7514 1.0000
2.250 0.6872 0.01315 0.00492 -0.0999 0.7394 1.0000
2.500 0.7132 0.01328 0.00509 -0.0993 0.7259 1.0000
2.750 0.7388 0.01342 0.00525 -0.0986 0.7112 1.0000
3.000 0.7642 0.01356 0.00543 -0.0978 0.6956 1.0000
3.250 0.7893 0.01370 0.00564 -0.0970 0.6792 1.0000
3.500 0.8104 0.01370 0.00555 -0.0948 0.6353 1.0000
3.750 0.8323 0.01385 0.00560 -0.0931 0.5937 1.0000
4.000 0.8501 0.01429 0.00559 -0.0905 0.5194 1.0000
4.250 0.8684 0.01490 0.00592 -0.0884 0.4467 1.0000
4.500 0.8762 0.01663 0.00653 -0.0850 0.2451 1.0000
4.750 0.8801 0.01951 0.00802 -0.0819 0.0321 1.0000
5.000 0.9003 0.02039 0.00899 -0.0804 0.0261 1.0000
5.250 0.9207 0.02122 0.01001 -0.0789 0.0241 1.0000
5.500 0.9394 0.02225 0.01123 -0.0773 0.0214 1.0000
5.750 0.9545 0.02366 0.01286 -0.0752 0.0192 1.0000
6.000 0.9702 0.02492 0.01429 -0.0731 0.0186 1.0000
6.250 0.9853 0.02623 0.01581 -0.0710 0.0182 1.0000
6.500 1.0002 0.02771 0.01739 -0.0689 0.0179 1.0000
6.750 1.0173 0.02934 0.01908 -0.0671 0.0177 1.0000
7.000 1.0380 0.03105 0.02087 -0.0657 0.0171 1.0000
7.250 1.0606 0.03279 0.02275 -0.0646 0.0160 1.0000
7.500 1.0843 0.03484 0.02497 -0.0637 0.0152 1.0000
7.750 1.1087 0.03721 0.02757 -0.0628 0.0153 1.0000
8.000 1.1312 0.03983 0.03048 -0.0617 0.0154 1.0000
8.250 1.1505 0.04267 0.03366 -0.0602 0.0157 1.0000
8.500 1.1660 0.04566 0.03712 -0.0584 0.0161 1.0000
8.750 1.1775 0.04894 0.04079 -0.0562 0.0165 1.0000
9.000 1.1855 0.05229 0.04451 -0.0539 0.0169 1.0000
9.250 1.1894 0.05597 0.04854 -0.0515 0.0174 1.0000
9.500 1.1898 0.05996 0.05284 -0.0491 0.0179 1.0000
9.750 1.2018 0.06239 0.05557 -0.0469 0.0205 1.0000
10.000 1.1873 0.06551 0.05925 -0.0425 0.0222 1.0000
10.250 1.1716 0.06899 0.06303 -0.0391 0.0233 1.0000
10.500 1.1559 0.07254 0.06682 -0.0367 0.0239 1.0000
10.750 1.1387 0.07651 0.07101 -0.0354 0.0245 1.0000
11.000 1.1213 0.08085 0.07553 -0.0352 0.0248 1.0000
11.250 1.1026 0.08570 0.08056 -0.0361 0.0251 1.0000
11.500 1.0831 0.09116 0.08619 -0.0381 0.0251 1.0000
11.750 1.0629 0.09731 0.09249 -0.0414 0.0249 1.0000
12.000 1.0422 0.10455 0.09986 -0.0460 0.0246 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 372 AIRFOIL (goe372-il)