GOE 404 AIRFOIL (goe404-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 404 AIRFOIL (goe404-il) Reynolds number: 1,000,000 Max Cl/Cd: 121.56 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe404-il-1000000.txt Download as CSV file: xf-goe404-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 404 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.500 -0.3548 0.14871 0.14688 -0.0130 1.0000 0.0250
-13.250 -0.3664 0.14075 0.13891 -0.0184 1.0000 0.0263
-12.000 -0.7735 0.03241 0.02955 -0.0918 0.9938 0.0273
-11.750 -0.7511 0.02975 0.02671 -0.0944 0.9901 0.0275
-11.500 -0.7266 0.02778 0.02459 -0.0962 0.9857 0.0277
-11.250 -0.7004 0.02603 0.02269 -0.0979 0.9820 0.0280
-11.000 -0.6781 0.02444 0.02094 -0.0984 0.9755 0.0281
-10.750 -0.6548 0.02300 0.01934 -0.0987 0.9693 0.0283
-10.500 -0.6336 0.02178 0.01797 -0.0983 0.9614 0.0285
-10.250 -0.6122 0.02069 0.01674 -0.0976 0.9527 0.0286
-10.000 -0.5913 0.01973 0.01564 -0.0966 0.9426 0.0288
-9.750 -0.5694 0.01885 0.01462 -0.0957 0.9329 0.0290
-9.500 -0.5467 0.01805 0.01368 -0.0949 0.9227 0.0292
-9.250 -0.5231 0.01732 0.01282 -0.0942 0.9124 0.0293
-9.000 -0.4992 0.01667 0.01204 -0.0934 0.9007 0.0295
-8.750 -0.4743 0.01608 0.01133 -0.0928 0.8883 0.0297
-8.500 -0.4490 0.01556 0.01069 -0.0922 0.8758 0.0298
-8.250 -0.4234 0.01512 0.01012 -0.0916 0.8616 0.0300
-8.000 -0.3997 0.01412 0.00896 -0.0909 0.8446 0.0303
-7.750 -0.3749 0.01346 0.00818 -0.0902 0.8243 0.0306
-7.500 -0.3492 0.01302 0.00763 -0.0897 0.7995 0.0310
-7.250 -0.3233 0.01269 0.00717 -0.0891 0.7740 0.0313
-7.000 -0.2970 0.01241 0.00675 -0.0886 0.7517 0.0316
-6.750 -0.2702 0.01215 0.00639 -0.0882 0.7338 0.0321
-6.500 -0.2430 0.01189 0.00605 -0.0878 0.7197 0.0325
-6.250 -0.2157 0.01164 0.00570 -0.0875 0.7078 0.0329
-6.000 -0.1882 0.01139 0.00536 -0.0871 0.6965 0.0333
-5.750 -0.1605 0.01116 0.00506 -0.0869 0.6865 0.0336
-5.500 -0.1327 0.01096 0.00479 -0.0866 0.6763 0.0339
-5.250 -0.1055 0.01055 0.00432 -0.0862 0.6678 0.0346
-5.000 -0.0778 0.01028 0.00402 -0.0860 0.6591 0.0352
-4.750 -0.0499 0.01009 0.00378 -0.0857 0.6510 0.0359
-4.500 -0.0217 0.00990 0.00356 -0.0855 0.6426 0.0367
-4.250 0.0063 0.00977 0.00336 -0.0852 0.6335 0.0375
-4.000 0.0347 0.00960 0.00316 -0.0851 0.6252 0.0383
-3.750 0.0624 0.00940 0.00293 -0.0848 0.6159 0.0399
-3.500 0.0909 0.00925 0.00277 -0.0846 0.6078 0.0415
-3.250 0.1189 0.00911 0.00260 -0.0844 0.5982 0.0440
-3.000 0.1472 0.00897 0.00246 -0.0842 0.5891 0.0483
-2.750 0.1752 0.00888 0.00239 -0.0840 0.5783 0.0572
-2.500 0.2038 0.00887 0.00238 -0.0838 0.5671 0.0654
-2.250 0.2322 0.00891 0.00240 -0.0837 0.5541 0.0706
-2.000 0.2604 0.00895 0.00238 -0.0835 0.5389 0.0740
-1.750 0.2886 0.00900 0.00240 -0.0833 0.5222 0.0770
-1.500 0.3168 0.00908 0.00241 -0.0831 0.5047 0.0795
-1.250 0.3447 0.00914 0.00239 -0.0829 0.4871 0.0816
-1.000 0.3724 0.00917 0.00237 -0.0826 0.4709 0.0843
-0.750 0.4003 0.00925 0.00241 -0.0824 0.4568 0.0867
-0.500 0.4282 0.00933 0.00242 -0.0822 0.4434 0.0886
0.000 0.4837 0.00938 0.00238 -0.0818 0.4211 0.0932
0.250 0.5113 0.00945 0.00242 -0.0816 0.4119 0.0957
0.500 0.5397 0.00947 0.00243 -0.0815 0.4050 0.0983
0.750 0.5676 0.00956 0.00248 -0.0813 0.3973 0.1002
1.000 0.5954 0.00953 0.00243 -0.0812 0.3915 0.1033
1.250 0.6236 0.00952 0.00243 -0.0810 0.3861 0.1060
1.500 0.6513 0.00957 0.00246 -0.0808 0.3795 0.1087
1.750 0.6793 0.00962 0.00249 -0.0807 0.3740 0.1113
2.000 0.7075 0.00963 0.00250 -0.0806 0.3694 0.1144
2.250 0.7354 0.00965 0.00253 -0.0804 0.3647 0.1202
2.500 0.7629 0.00973 0.00259 -0.0802 0.3595 0.1256
2.750 0.7909 0.00973 0.00265 -0.0801 0.3560 0.1397
3.000 0.8187 0.00965 0.00273 -0.0800 0.3523 0.2017
3.250 0.8379 0.00835 0.00295 -0.0788 0.3484 0.8011
3.500 0.8851 0.00822 0.00312 -0.0827 0.3415 1.0000
3.750 0.9118 0.00831 0.00319 -0.0823 0.3378 1.0000
4.000 0.9385 0.00841 0.00327 -0.0819 0.3342 1.0000
4.250 0.9649 0.00854 0.00336 -0.0815 0.3298 1.0000
4.500 0.9907 0.00872 0.00348 -0.0810 0.3240 1.0000
4.750 1.0177 0.00882 0.00357 -0.0807 0.3202 1.0000
5.000 1.0445 0.00894 0.00366 -0.0804 0.3154 1.0000
5.250 1.0706 0.00911 0.00379 -0.0800 0.3099 1.0000
5.500 1.0969 0.00926 0.00392 -0.0797 0.3052 1.0000
5.750 1.1238 0.00938 0.00403 -0.0794 0.3005 1.0000
6.000 1.1499 0.00956 0.00417 -0.0790 0.2943 1.0000
6.250 1.1759 0.00974 0.00432 -0.0787 0.2884 1.0000
6.500 1.2022 0.00989 0.00446 -0.0784 0.2813 1.0000
6.750 1.2275 0.01013 0.00463 -0.0779 0.2726 1.0000
7.000 1.2529 0.01035 0.00481 -0.0775 0.2621 1.0000
7.250 1.2775 0.01062 0.00502 -0.0770 0.2485 1.0000
7.500 1.3000 0.01106 0.00532 -0.0762 0.2275 1.0000
7.750 1.3193 0.01173 0.00578 -0.0749 0.1951 1.0000
8.000 1.3397 0.01229 0.00622 -0.0738 0.1773 1.0000
8.250 1.3615 0.01274 0.00661 -0.0729 0.1678 1.0000
8.500 1.3839 0.01313 0.00696 -0.0721 0.1621 1.0000
8.750 1.4056 0.01354 0.00735 -0.0713 0.1567 1.0000
9.000 1.4291 0.01381 0.00763 -0.0706 0.1543 1.0000
9.250 1.4517 0.01413 0.00796 -0.0699 0.1515 1.0000
9.500 1.4732 0.01450 0.00832 -0.0690 0.1485 1.0000
9.750 1.4936 0.01492 0.00874 -0.0679 0.1451 1.0000
10.000 1.5140 0.01529 0.00912 -0.0669 0.1426 1.0000
10.250 1.5347 0.01557 0.00944 -0.0658 0.1410 1.0000
10.500 1.5541 0.01591 0.00980 -0.0646 0.1390 1.0000
10.750 1.5725 0.01632 0.01021 -0.0632 0.1364 1.0000
11.000 1.5893 0.01681 0.01071 -0.0618 0.1335 1.0000
11.250 1.6046 0.01740 0.01130 -0.0601 0.1303 1.0000
11.500 1.6240 0.01776 0.01171 -0.0591 0.1290 1.0000
11.750 1.6423 0.01820 0.01219 -0.0580 0.1270 1.0000
12.000 1.6593 0.01873 0.01274 -0.0568 0.1244 1.0000
12.250 1.6742 0.01940 0.01341 -0.0554 0.1212 1.0000
12.500 1.6880 0.02016 0.01418 -0.0540 0.1180 1.0000
12.750 1.7049 0.02073 0.01480 -0.0530 0.1154 1.0000
13.000 1.7186 0.02154 0.01561 -0.0517 0.1109 1.0000
13.250 1.7298 0.02254 0.01660 -0.0503 0.1057 1.0000
13.500 1.7390 0.02372 0.01776 -0.0489 0.0982 1.0000
13.750 1.7445 0.02524 0.01924 -0.0473 0.0886 1.0000
14.000 1.7435 0.02734 0.02128 -0.0455 0.0771 1.0000
14.250 1.7417 0.02969 0.02361 -0.0441 0.0698 1.0000
14.500 1.7424 0.03193 0.02588 -0.0430 0.0663 1.0000
14.750 1.7423 0.03439 0.02837 -0.0422 0.0636 1.0000
15.000 1.7453 0.03662 0.03066 -0.0416 0.0623 1.0000
15.250 1.7463 0.03910 0.03321 -0.0411 0.0613 1.0000
15.500 1.7451 0.04186 0.03605 -0.0407 0.0603 1.0000
15.750 1.7424 0.04486 0.03911 -0.0404 0.0594 1.0000
16.000 1.7378 0.04812 0.04244 -0.0402 0.0585 1.0000
16.250 1.7323 0.05156 0.04596 -0.0401 0.0577 1.0000
16.500 1.7242 0.05541 0.04990 -0.0402 0.0569 1.0000
16.750 1.7195 0.05890 0.05348 -0.0404 0.0566 1.0000
17.000 1.7150 0.06245 0.05713 -0.0406 0.0563 1.0000
17.250 1.7092 0.06622 0.06100 -0.0411 0.0559 1.0000
17.500 1.7022 0.07019 0.06506 -0.0416 0.0555 1.0000
17.750 1.6943 0.07434 0.06930 -0.0422 0.0551 1.0000
18.000 1.6855 0.07866 0.07371 -0.0429 0.0547 1.0000
18.250 1.6760 0.08311 0.07825 -0.0438 0.0543 1.0000
18.500 1.6656 0.08769 0.08292 -0.0447 0.0539 1.0000
18.750 1.6547 0.09240 0.08771 -0.0457 0.0534 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 404 AIRFOIL (goe404-il)