NASA SC(2)-0612 AIRFOIL (sc20612-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA SC(2)-0612 AIRFOIL (sc20612-il) Reynolds number: 200,000 Max Cl/Cd: 40.53 at α=2.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-sc20612-il-200000.txt Download as CSV file: xf-sc20612-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0612 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.000 -0.5838 0.12573 0.12150 -0.0307 1.0000 0.0785
-12.750 -0.7536 0.07520 0.07064 -0.0595 1.0000 0.0505
-12.500 -0.7726 0.06947 0.06483 -0.0617 1.0000 0.0501
-12.250 -0.7969 0.06411 0.05936 -0.0632 1.0000 0.0497
-12.000 -0.8227 0.05958 0.05470 -0.0632 1.0000 0.0494
-11.750 -0.8497 0.05610 0.05110 -0.0612 1.0000 0.0491
-11.500 -0.8739 0.05273 0.04756 -0.0591 1.0000 0.0489
-11.250 -0.8904 0.04891 0.04348 -0.0578 1.0000 0.0488
-11.000 -0.8993 0.04518 0.03943 -0.0564 1.0000 0.0488
-10.750 -0.9010 0.04170 0.03558 -0.0553 1.0000 0.0492
-10.500 -0.8960 0.03871 0.03214 -0.0544 1.0000 0.0501
-10.250 -0.8856 0.03604 0.02903 -0.0540 1.0000 0.0510
-10.000 -0.8698 0.03372 0.02666 -0.0533 1.0000 0.0521
-9.750 -0.8515 0.03236 0.02525 -0.0527 1.0000 0.0533
-9.500 -0.8324 0.03090 0.02367 -0.0521 1.0000 0.0545
-9.250 -0.8124 0.02937 0.02193 -0.0517 1.0000 0.0561
-9.000 -0.7912 0.02798 0.02027 -0.0513 1.0000 0.0579
-8.750 -0.7700 0.02639 0.01851 -0.0509 1.0000 0.0600
-8.500 -0.7484 0.02547 0.01765 -0.0504 1.0000 0.0622
-8.250 -0.7257 0.02462 0.01672 -0.0500 1.0000 0.0647
-8.000 -0.7021 0.02376 0.01569 -0.0496 1.0000 0.0673
-7.750 -0.6796 0.02255 0.01448 -0.0491 1.0000 0.0704
-7.500 -0.6559 0.02194 0.01391 -0.0489 1.0000 0.0741
-7.250 -0.6311 0.02138 0.01323 -0.0488 1.0000 0.0780
-7.000 -0.6068 0.02038 0.01230 -0.0486 1.0000 0.0822
-6.750 -0.5812 0.01990 0.01183 -0.0487 1.0000 0.0873
-6.500 -0.5550 0.01928 0.01116 -0.0488 1.0000 0.0924
-6.250 -0.5284 0.01869 0.01067 -0.0492 1.0000 0.0980
-6.000 -0.5014 0.01838 0.01028 -0.0495 1.0000 0.1038
-5.750 -0.4727 0.01767 0.00971 -0.0504 1.0000 0.1114
-5.500 -0.4439 0.01725 0.00930 -0.0512 1.0000 0.1196
-5.250 -0.4133 0.01674 0.00889 -0.0525 1.0000 0.1308
-5.000 -0.3806 0.01620 0.00846 -0.0543 1.0000 0.1471
-4.750 -0.3417 0.01535 0.00793 -0.0578 1.0000 0.1873
-4.500 -0.2744 0.01315 0.00760 -0.0685 1.0000 0.5253
-4.250 -0.2517 0.01357 0.00817 -0.0672 1.0000 0.5945
-4.000 -0.2282 0.01400 0.00858 -0.0662 1.0000 0.6232
-3.750 -0.2073 0.01450 0.00908 -0.0646 1.0000 0.6405
-3.500 -0.1857 0.01501 0.00956 -0.0633 1.0000 0.6554
-3.250 -0.1690 0.01560 0.01019 -0.0607 1.0000 0.6653
-3.000 -0.1489 0.01610 0.01068 -0.0590 1.0000 0.6752
-2.750 -0.1304 0.01663 0.01122 -0.0570 1.0000 0.6839
-2.500 -0.1027 0.01717 0.01176 -0.0566 0.9973 0.6932
-2.250 -0.0715 0.01791 0.01252 -0.0565 0.9937 0.7048
-2.000 -0.0481 0.01859 0.01326 -0.0545 0.9881 0.7135
-1.750 -0.0129 0.01918 0.01384 -0.0555 0.9840 0.7269
-1.500 0.0055 0.01978 0.01451 -0.0522 0.9777 0.7329
-1.250 0.0420 0.02009 0.01481 -0.0538 0.9729 0.7434
-1.000 0.0767 0.02022 0.01497 -0.0546 0.9679 0.7482
-0.750 0.1077 0.02022 0.01499 -0.0549 0.9603 0.7518
-0.500 0.1515 0.02015 0.01494 -0.0580 0.9564 0.7561
-0.250 0.1952 0.01997 0.01474 -0.0619 0.9481 0.7630
0.000 0.2413 0.01933 0.01416 -0.0640 0.9366 0.7654
0.250 0.2945 0.01856 0.01343 -0.0677 0.9276 0.7677
0.500 0.3385 0.01789 0.01281 -0.0699 0.9165 0.7702
0.750 0.3873 0.01699 0.01195 -0.0729 0.9048 0.7731
1.000 0.4328 0.01627 0.01128 -0.0756 0.8941 0.7768
1.250 0.4755 0.01572 0.01076 -0.0785 0.8786 0.7802
1.500 0.5023 0.01529 0.01038 -0.0774 0.8586 0.7818
1.750 0.5332 0.01484 0.00995 -0.0770 0.8316 0.7836
2.000 0.5615 0.01454 0.00960 -0.0761 0.7854 0.7856
2.250 0.5893 0.01454 0.00920 -0.0750 0.6769 0.7879
2.500 0.6004 0.01592 0.00931 -0.0715 0.4408 0.7909
2.750 0.6154 0.01755 0.00987 -0.0700 0.2445 0.7942
3.000 0.6409 0.01862 0.01040 -0.0704 0.1663 0.7974
3.250 0.6613 0.01919 0.01079 -0.0689 0.1432 0.7993
3.500 0.6847 0.01967 0.01121 -0.0681 0.1293 0.8012
3.750 0.7095 0.02019 0.01169 -0.0676 0.1192 0.8033
4.000 0.7344 0.02092 0.01229 -0.0674 0.1104 0.8056
4.250 0.7631 0.02126 0.01269 -0.0677 0.1036 0.8080
4.500 0.7910 0.02208 0.01339 -0.0682 0.0969 0.8105
4.750 0.8228 0.02258 0.01396 -0.0693 0.0910 0.8133
5.000 0.8464 0.02311 0.01445 -0.0685 0.0860 0.8155
5.250 0.8705 0.02397 0.01535 -0.0678 0.0814 0.8177
5.500 0.8960 0.02449 0.01592 -0.0673 0.0767 0.8202
5.750 0.9228 0.02581 0.01712 -0.0675 0.0723 0.8229
6.000 0.9510 0.02647 0.01797 -0.0676 0.0691 0.8259
6.250 0.9804 0.02726 0.01880 -0.0682 0.0655 0.8285
6.500 1.0088 0.02886 0.02030 -0.0688 0.0624 0.8306
6.750 1.0323 0.02960 0.02130 -0.0678 0.0602 0.8327
7.000 1.0567 0.03071 0.02259 -0.0672 0.0580 0.8350
7.250 1.0811 0.03175 0.02372 -0.0667 0.0561 0.8379
7.500 1.1071 0.03335 0.02528 -0.0670 0.0543 0.8410
7.750 1.1300 0.03518 0.02747 -0.0664 0.0526 0.8439
8.000 1.1514 0.03706 0.02971 -0.0656 0.0510 0.8463
8.250 1.1686 0.03903 0.03200 -0.0639 0.0499 0.8485
8.500 1.1852 0.04110 0.03433 -0.0624 0.0490 0.8510
8.750 1.2017 0.04316 0.03662 -0.0611 0.0483 0.8538
9.000 1.2201 0.04509 0.03865 -0.0603 0.0475 0.8567
9.250 1.2361 0.04782 0.04150 -0.0596 0.0467 0.8599
9.500 1.2382 0.05191 0.04596 -0.0570 0.0462 0.8625
9.750 1.2312 0.05533 0.04990 -0.0529 0.0459 0.8651
10.000 1.2200 0.05938 0.05439 -0.0491 0.0457 0.8678
10.250 1.2063 0.06361 0.05896 -0.0456 0.0458 0.8705
10.500 1.1896 0.06753 0.06312 -0.0420 0.0460 0.8732
10.750 1.1741 0.07187 0.06763 -0.0394 0.0462 0.8760
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