GOE 310 (MVA H.42) AIRFOIL (goe310-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 310 (MVA H.42) AIRFOIL (goe310-il) Reynolds number: 500,000 Max Cl/Cd: 99.91 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe310-il-500000.txt Download as CSV file: xf-goe310-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 310 (MVA H.42) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3479 0.09320 0.09103 -0.0348 1.0000 0.0210
-8.500 -0.3577 0.09099 0.08888 -0.0334 1.0000 0.0210
-8.250 -0.3708 0.08919 0.08712 -0.0313 0.9998 0.0210
-8.000 -0.3564 0.08370 0.08163 -0.0391 0.9954 0.0211
-7.750 -0.3466 0.07700 0.07493 -0.0448 0.9905 0.0215
-7.500 -0.3268 0.07298 0.07089 -0.0496 0.9866 0.0218
-7.250 -0.3037 0.06988 0.06776 -0.0537 0.9839 0.0223
-7.000 -0.2842 0.06627 0.06413 -0.0580 0.9785 0.0232
-6.750 -0.2599 0.06164 0.05945 -0.0643 0.9738 0.0244
-6.500 -0.1894 0.02927 0.02696 -0.0831 0.9592 0.0278
-6.250 -0.1641 0.02826 0.02594 -0.0841 0.9552 0.0285
-6.000 -0.1463 0.02573 0.02333 -0.0852 0.9464 0.0295
-5.750 -0.1262 0.02153 0.01898 -0.0878 0.9398 0.0313
-5.500 -0.1148 0.01455 0.01120 -0.0900 0.9300 0.0348
-5.250 -0.1124 0.00934 0.00562 -0.0888 0.9210 0.0371
-5.000 -0.0903 0.00845 0.00468 -0.0884 0.9130 0.0394
-3.750 0.0184 0.01383 0.00822 -0.0857 0.8841 0.0411
-3.500 0.0436 0.01268 0.00677 -0.0845 0.8733 0.0381
-3.250 0.0697 0.01199 0.00592 -0.0837 0.8619 0.0374
-3.000 0.0954 0.01137 0.00520 -0.0830 0.8492 0.0376
-2.750 0.1203 0.01083 0.00457 -0.0821 0.8348 0.0386
-2.500 0.1452 0.01038 0.00403 -0.0812 0.8184 0.0387
-2.250 0.1699 0.01002 0.00358 -0.0803 0.7995 0.0390
-2.000 0.1944 0.00975 0.00319 -0.0793 0.7793 0.0393
-1.750 0.2188 0.00954 0.00286 -0.0784 0.7591 0.0397
-1.500 0.2432 0.00939 0.00260 -0.0775 0.7398 0.0404
-1.250 0.2676 0.00927 0.00237 -0.0766 0.7224 0.0416
-1.000 0.2924 0.00920 0.00218 -0.0757 0.7069 0.0423
-0.750 0.3173 0.00914 0.00203 -0.0750 0.6932 0.0431
-0.500 0.3425 0.00912 0.00191 -0.0743 0.6807 0.0440
-0.250 0.3677 0.00911 0.00182 -0.0736 0.6696 0.0459
0.000 0.3927 0.00910 0.00176 -0.0728 0.6586 0.0519
0.250 0.4158 0.00888 0.00173 -0.0718 0.6464 0.1173
0.500 0.4399 0.00885 0.00180 -0.0710 0.6333 0.1652
0.750 0.4648 0.00890 0.00186 -0.0703 0.6216 0.1860
1.000 0.4898 0.00897 0.00189 -0.0697 0.6107 0.1985
1.250 0.5151 0.00901 0.00194 -0.0691 0.6009 0.2116
1.500 0.5399 0.00905 0.00197 -0.0684 0.5893 0.2236
1.750 0.5640 0.00908 0.00199 -0.0676 0.5755 0.2359
2.000 0.5879 0.00910 0.00202 -0.0667 0.5622 0.2471
2.250 0.6119 0.00913 0.00206 -0.0659 0.5509 0.2637
2.500 0.6359 0.00909 0.00212 -0.0651 0.5400 0.2935
2.750 0.6584 0.00897 0.00224 -0.0641 0.5289 0.3886
3.000 0.7907 0.00818 0.00270 -0.0877 0.4957 0.9959
3.250 0.8329 0.00835 0.00275 -0.0911 0.4667 1.0000
3.500 0.8542 0.00855 0.00283 -0.0898 0.4401 1.0000
3.750 0.8737 0.00885 0.00295 -0.0881 0.4001 1.0000
4.000 0.8902 0.00938 0.00317 -0.0860 0.3390 1.0000
4.250 0.9046 0.01010 0.00352 -0.0836 0.2694 1.0000
4.500 0.9204 0.01074 0.00387 -0.0814 0.2160 1.0000
4.750 0.9384 0.01122 0.00417 -0.0796 0.1834 1.0000
5.000 0.9574 0.01160 0.00445 -0.0780 0.1600 1.0000
5.250 0.9768 0.01196 0.00474 -0.0764 0.1424 1.0000
5.500 0.9941 0.01245 0.00509 -0.0745 0.1129 1.0000
5.750 1.0072 0.01321 0.00561 -0.0718 0.0681 1.0000
6.000 1.0248 0.01366 0.00605 -0.0699 0.0633 1.0000
6.250 1.0429 0.01405 0.00648 -0.0681 0.0592 1.0000
6.500 1.0614 0.01442 0.00689 -0.0663 0.0573 1.0000
6.750 1.0791 0.01481 0.00735 -0.0645 0.0554 1.0000
7.000 1.0956 0.01527 0.00787 -0.0624 0.0535 1.0000
7.250 1.1102 0.01582 0.00848 -0.0600 0.0513 1.0000
7.500 1.1204 0.01659 0.00932 -0.0569 0.0483 1.0000
7.750 1.1348 0.01710 0.00989 -0.0546 0.0461 1.0000
8.000 1.1534 0.01736 0.01021 -0.0530 0.0430 1.0000
8.250 1.1679 0.01783 0.01071 -0.0507 0.0393 1.0000
8.500 1.1810 0.01824 0.01116 -0.0481 0.0348 1.0000
8.750 1.2003 0.01837 0.01124 -0.0466 0.0289 1.0000
9.000 1.2147 0.01871 0.01159 -0.0442 0.0251 1.0000
9.250 1.2257 0.01925 0.01210 -0.0414 0.0223 1.0000
9.500 1.2336 0.01998 0.01291 -0.0381 0.0207 1.0000
9.750 1.2435 0.02068 0.01367 -0.0354 0.0193 1.0000
10.000 1.2521 0.02149 0.01453 -0.0326 0.0183 1.0000
10.250 1.2563 0.02262 0.01572 -0.0294 0.0172 1.0000
10.500 1.2550 0.02417 0.01738 -0.0258 0.0163 1.0000
10.750 1.2634 0.02518 0.01849 -0.0236 0.0157 1.0000
11.000 1.2687 0.02648 0.01987 -0.0212 0.0152 1.0000
11.250 1.2739 0.02785 0.02133 -0.0190 0.0145 1.0000
11.500 1.2792 0.02928 0.02285 -0.0170 0.0142 1.0000
11.750 1.2821 0.03099 0.02463 -0.0151 0.0136 1.0000
12.000 1.2845 0.03283 0.02656 -0.0133 0.0133 1.0000
12.250 1.2845 0.03499 0.02880 -0.0115 0.0129 1.0000
12.500 1.2810 0.03762 0.03155 -0.0095 0.0125 1.0000
12.750 1.2799 0.04022 0.03428 -0.0076 0.0123 1.0000
13.000 1.2844 0.04225 0.03643 -0.0064 0.0121 1.0000
13.250 1.2883 0.04440 0.03872 -0.0052 0.0120 1.0000
13.500 1.2918 0.04661 0.04107 -0.0042 0.0117 1.0000
13.750 1.2946 0.04892 0.04352 -0.0034 0.0115 1.0000
14.000 1.2953 0.05156 0.04631 -0.0025 0.0113 1.0000
14.250 1.2951 0.05433 0.04922 -0.0020 0.0110 1.0000
14.500 1.2941 0.05720 0.05224 -0.0017 0.0107 1.0000
14.750 1.2910 0.06047 0.05567 -0.0014 0.0106 1.0000
15.000 1.2896 0.06345 0.05875 -0.0017 0.0103 1.0000
15.250 1.2830 0.06738 0.06285 -0.0018 0.0102 1.0000
15.500 1.2776 0.07110 0.06670 -0.0025 0.0100 1.0000
15.750 1.2669 0.07585 0.07164 -0.0031 0.0100 1.0000
16.000 1.2559 0.08077 0.07673 -0.0043 0.0099 1.0000
16.250 1.2470 0.08543 0.08152 -0.0059 0.0098 1.0000
16.500 1.2288 0.09191 0.08823 -0.0079 0.0099 1.0000
16.750 1.2160 0.09766 0.09411 -0.0103 0.0098 1.0000
17.000 1.1782 0.10862 0.10544 -0.0149 0.0102 1.0000
17.250 1.1768 0.11263 0.10947 -0.0171 0.0099 1.0000
17.500 1.1242 0.12839 0.12563 -0.0254 0.0103 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 310 (MVA H.42) AIRFOIL (goe310-il)