EPPLER 1210 AIRFOIL (e1210-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 1210 AIRFOIL (e1210-il) Reynolds number: 500,000 Max Cl/Cd: 92.2 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e1210-il-500000.txt Download as CSV file: xf-e1210-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 1210 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
------ -------- --------- --------- -------- -------- --------
-15.000 -0.6285 0.09063 0.08743 -0.0504 1.0000 0.0246
-14.750 -0.6641 0.08115 0.07781 -0.0564 1.0000 0.0246
-14.500 -0.6862 0.07473 0.07133 -0.0604 1.0000 0.0249
-14.250 -0.7076 0.06859 0.06513 -0.0642 1.0000 0.0251
-14.000 -0.7300 0.06253 0.05902 -0.0680 1.0000 0.0252
-13.750 -0.7624 0.05522 0.05162 -0.0728 1.0000 0.0252
-13.500 -0.8022 0.04681 0.04305 -0.0785 1.0000 0.0250
-13.250 -0.8447 0.03688 0.03290 -0.0870 1.0000 0.0249
-13.000 -0.8730 0.03181 0.02762 -0.0898 1.0000 0.0249
-12.750 -0.8771 0.02961 0.02537 -0.0887 1.0000 0.0252
-12.500 -0.8734 0.02796 0.02366 -0.0873 1.0000 0.0256
-12.250 -0.8652 0.02666 0.02229 -0.0858 1.0000 0.0260
-12.000 -0.8557 0.02546 0.02103 -0.0842 1.0000 0.0265
-11.500 -0.7878 0.02261 0.01791 -0.0898 0.9925 0.0285
-11.250 -0.7527 0.02141 0.01668 -0.0923 0.9853 0.0296
-11.000 -0.7177 0.02041 0.01559 -0.0944 0.9755 0.0310
-10.750 -0.6875 0.01934 0.01444 -0.0955 0.9578 0.0325
-10.500 -0.6466 0.01847 0.01351 -0.0984 0.9418 0.0343
-10.250 -0.5958 0.01746 0.01240 -0.1034 0.9230 0.0368
-10.000 -0.5502 0.01682 0.01159 -0.1069 0.8900 0.0394
-9.750 -0.5225 0.01635 0.01096 -0.1070 0.8529 0.0415
-9.500 -0.4977 0.01611 0.01051 -0.1062 0.8234 0.0438
-9.250 -0.4742 0.01572 0.01001 -0.1053 0.7984 0.0458
-9.000 -0.4493 0.01550 0.00964 -0.1045 0.7770 0.0481
-8.750 -0.4248 0.01516 0.00917 -0.1038 0.7576 0.0503
-8.500 -0.3994 0.01495 0.00887 -0.1031 0.7399 0.0525
-8.250 -0.3735 0.01470 0.00846 -0.1024 0.7231 0.0548
-8.000 -0.3478 0.01445 0.00817 -0.1019 0.7077 0.0570
-7.750 -0.3212 0.01430 0.00789 -0.1013 0.6932 0.0596
-7.500 -0.2954 0.01404 0.00756 -0.1007 0.6790 0.0620
-7.250 -0.2683 0.01390 0.00736 -0.1003 0.6657 0.0646
-7.000 -0.2417 0.01369 0.00703 -0.0998 0.6533 0.0672
-6.750 -0.2147 0.01356 0.00686 -0.0993 0.6408 0.0697
-6.500 -0.1871 0.01349 0.00667 -0.0989 0.6291 0.0727
-6.250 -0.1605 0.01330 0.00645 -0.0984 0.6176 0.0756
-6.000 -0.1326 0.01322 0.00632 -0.0981 0.6069 0.0786
-5.500 -0.0781 0.01289 0.00589 -0.0973 0.5856 0.0842
-5.000 -0.0232 0.01259 0.00545 -0.0965 0.5661 0.0901
-4.750 0.0042 0.01253 0.00533 -0.0961 0.5563 0.0930
-4.500 0.0324 0.01246 0.00519 -0.0957 0.5471 0.0962
-4.250 0.0595 0.01226 0.00494 -0.0953 0.5380 0.0994
-4.000 0.0874 0.01218 0.00485 -0.0950 0.5298 0.1026
-3.750 0.1154 0.01214 0.00473 -0.0947 0.5208 0.1059
-3.500 0.1427 0.01196 0.00451 -0.0943 0.5126 0.1094
-3.250 0.1707 0.01187 0.00442 -0.0940 0.5041 0.1128
-3.000 0.1983 0.01188 0.00433 -0.0936 0.4960 0.1164
-2.750 0.2262 0.01169 0.00415 -0.0933 0.4888 0.1202
-2.500 0.2539 0.01163 0.00406 -0.0930 0.4808 0.1242
-2.250 0.2818 0.01162 0.00400 -0.0926 0.4734 0.1282
-2.000 0.3096 0.01150 0.00385 -0.0923 0.4659 0.1326
-1.750 0.3369 0.01146 0.00379 -0.0920 0.4586 0.1372
-1.500 0.3652 0.01143 0.00374 -0.0917 0.4520 0.1419
-1.250 0.3928 0.01134 0.00364 -0.0914 0.4449 0.1472
-1.000 0.4200 0.01136 0.00361 -0.0910 0.4379 0.1529
-0.750 0.4484 0.01133 0.00357 -0.0908 0.4315 0.1584
-0.500 0.4758 0.01126 0.00351 -0.0905 0.4249 0.1662
-0.250 0.5031 0.01131 0.00351 -0.0901 0.4186 0.1736
0.000 0.5312 0.01122 0.00349 -0.0899 0.4125 0.1838
0.250 0.5587 0.01123 0.00348 -0.0895 0.4061 0.1950
0.500 0.5857 0.01126 0.00351 -0.0892 0.4001 0.2110
0.750 0.6136 0.01118 0.00353 -0.0890 0.3947 0.2353
1.500 0.6943 0.01092 0.00375 -0.0881 0.3780 0.4177
1.750 0.7204 0.01072 0.00387 -0.0877 0.3728 0.5309
2.000 0.7443 0.01048 0.00402 -0.0868 0.3678 0.6777
2.250 0.7621 0.01014 0.00419 -0.0842 0.3633 0.8640
2.500 0.8062 0.01011 0.00425 -0.0871 0.3577 1.0000
2.750 0.8328 0.01027 0.00433 -0.0867 0.3527 1.0000
3.000 0.8585 0.01051 0.00446 -0.0861 0.3478 1.0000
3.250 0.8856 0.01064 0.00457 -0.0858 0.3440 1.0000
3.500 0.9125 0.01078 0.00468 -0.0854 0.3398 1.0000
3.750 0.9387 0.01096 0.00480 -0.0850 0.3354 1.0000
4.000 0.9641 0.01122 0.00497 -0.0844 0.3306 1.0000
4.250 0.9908 0.01137 0.00511 -0.0840 0.3270 1.0000
4.500 1.0174 0.01152 0.00525 -0.0837 0.3232 1.0000
4.750 1.0434 0.01172 0.00540 -0.0832 0.3193 1.0000
5.000 1.0686 0.01197 0.00559 -0.0827 0.3155 1.0000
5.250 1.0939 0.01223 0.00581 -0.0821 0.3117 1.0000
5.500 1.1201 0.01238 0.00598 -0.0817 0.3085 1.0000
5.750 1.1459 0.01257 0.00615 -0.0813 0.3048 1.0000
6.000 1.1708 0.01280 0.00635 -0.0807 0.3011 1.0000
6.250 1.1947 0.01313 0.00661 -0.0800 0.2973 1.0000
6.500 1.2198 0.01336 0.00685 -0.0795 0.2944 1.0000
6.750 1.2450 0.01355 0.00707 -0.0790 0.2915 1.0000
7.000 1.2696 0.01377 0.00729 -0.0784 0.2883 1.0000
7.250 1.2934 0.01403 0.00753 -0.0777 0.2850 1.0000
7.500 1.3159 0.01439 0.00783 -0.0768 0.2814 1.0000
7.750 1.3394 0.01468 0.00813 -0.0761 0.2785 1.0000
8.000 1.3632 0.01490 0.00839 -0.0754 0.2758 1.0000
8.250 1.3863 0.01515 0.00866 -0.0746 0.2729 1.0000
8.500 1.4085 0.01543 0.00895 -0.0737 0.2701 1.0000
8.750 1.4294 0.01579 0.00927 -0.0727 0.2672 1.0000
9.000 1.4495 0.01627 0.00971 -0.0716 0.2639 1.0000
9.250 1.4709 0.01650 0.01001 -0.0705 0.2617 1.0000
9.500 1.4904 0.01677 0.01032 -0.0692 0.2591 1.0000
9.750 1.5086 0.01707 0.01066 -0.0677 0.2565 1.0000
10.000 1.5257 0.01744 0.01103 -0.0660 0.2541 1.0000
10.250 1.5418 0.01788 0.01146 -0.0643 0.2516 1.0000
10.500 1.5579 0.01847 0.01201 -0.0627 0.2487 1.0000
10.750 1.5746 0.01882 0.01244 -0.0612 0.2468 1.0000
11.000 1.5907 0.01921 0.01290 -0.0596 0.2445 1.0000
11.250 1.6063 0.01967 0.01341 -0.0581 0.2422 1.0000
11.500 1.6210 0.02019 0.01396 -0.0565 0.2400 1.0000
11.750 1.6343 0.02079 0.01457 -0.0549 0.2377 1.0000
12.000 1.6471 0.02154 0.01529 -0.0533 0.2351 1.0000
12.250 1.6607 0.02224 0.01604 -0.0519 0.2329 1.0000
12.500 1.6734 0.02292 0.01681 -0.0505 0.2311 1.0000
12.750 1.6856 0.02368 0.01764 -0.0492 0.2289 1.0000
13.000 1.6970 0.02452 0.01853 -0.0479 0.2266 1.0000
13.250 1.7076 0.02546 0.01951 -0.0467 0.2243 1.0000
13.500 1.7173 0.02652 0.02058 -0.0454 0.2220 1.0000
13.750 1.7281 0.02762 0.02167 -0.0444 0.2193 1.0000
14.000 1.7366 0.02881 0.02297 -0.0434 0.2175 1.0000
14.250 1.7454 0.03004 0.02430 -0.0425 0.2155 1.0000
14.500 1.7534 0.03139 0.02571 -0.0417 0.2132 1.0000
14.750 1.7609 0.03282 0.02719 -0.0410 0.2110 1.0000
15.000 1.7672 0.03437 0.02877 -0.0403 0.2087 1.0000
15.250 1.7749 0.03585 0.03023 -0.0395 0.2061 1.0000
15.500 1.7803 0.03760 0.03208 -0.0390 0.2043 1.0000
15.750 1.7844 0.03951 0.03411 -0.0386 0.2024 1.0000
16.000 1.7882 0.04149 0.03618 -0.0383 0.2002 1.0000
16.250 1.7913 0.04360 0.03836 -0.0381 0.1980 1.0000
16.500 1.7945 0.04571 0.04051 -0.0380 0.1959 1.0000
16.750 1.7984 0.04773 0.04252 -0.0377 0.1933 1.0000
17.000 1.8001 0.05012 0.04500 -0.0377 0.1911 1.0000
17.250 1.7984 0.05303 0.04803 -0.0380 0.1889 1.0000
17.500 1.7977 0.05585 0.05095 -0.0384 0.1866 1.0000
18.000 1.7961 0.06164 0.05683 -0.0393 0.1817 1.0000
18.250 1.7991 0.06401 0.05921 -0.0395 0.1789 1.0000
18.500 1.7909 0.06810 0.06346 -0.0406 0.1766 1.0000
18.750 1.7845 0.07200 0.06747 -0.0417 0.1740 1.0000
19.000 1.7789 0.07584 0.07138 -0.0428 0.1713 1.0000
19.250 1.7775 0.07905 0.07459 -0.0437 0.1685 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 1210 AIRFOIL (e1210-il)