NACA 63-412 AIRFOIL (n63412-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63-412 AIRFOIL (n63412-il) Reynolds number: 200,000 Max Cl/Cd: 77.24 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63412-il-200000.txt Download as CSV file: xf-n63412-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-412 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4467 0.08769 0.08412 -0.0429 1.0000 0.0652
-9.500 -0.4538 0.08272 0.07921 -0.0460 1.0000 0.0676
-9.250 -0.4725 0.07531 0.07188 -0.0526 1.0000 0.0690
-9.000 -0.5037 0.06876 0.06535 -0.0580 1.0000 0.0693
-8.750 -0.5326 0.06517 0.06176 -0.0584 1.0000 0.0695
-8.500 -0.5687 0.06296 0.05939 -0.0564 1.0000 0.0705
-8.250 -0.6022 0.06274 0.05895 -0.0516 1.0000 0.0712
-7.750 -0.5964 0.05274 0.04895 -0.0524 0.9962 0.0742
-7.500 -0.5786 0.03930 0.03390 -0.0582 0.9873 0.0481
-7.250 -0.5474 0.03412 0.02852 -0.0611 0.9837 0.0453
-7.000 -0.5176 0.03044 0.02435 -0.0632 0.9774 0.0452
-6.750 -0.4826 0.02699 0.02043 -0.0656 0.9729 0.0447
-6.500 -0.4443 0.02451 0.01754 -0.0681 0.9696 0.0454
-6.250 -0.4104 0.02355 0.01623 -0.0694 0.9625 0.0474
-6.000 -0.3745 0.02078 0.01331 -0.0715 0.9591 0.0495
-5.750 -0.3342 0.01934 0.01184 -0.0742 0.9565 0.0520
-5.500 -0.3011 0.01843 0.01086 -0.0754 0.9500 0.0557
-5.250 -0.2642 0.01746 0.00979 -0.0772 0.9452 0.0595
-5.000 -0.2272 0.01613 0.00855 -0.0794 0.9416 0.0650
-4.750 -0.1974 0.01552 0.00788 -0.0800 0.9335 0.0720
-4.500 -0.1635 0.01458 0.00698 -0.0815 0.9279 0.0841
-4.250 -0.1343 0.01367 0.00621 -0.0823 0.9202 0.1121
-4.000 -0.1094 0.01142 0.00546 -0.0836 0.9128 0.4221
-3.750 -0.0823 0.01120 0.00551 -0.0833 0.9049 0.5244
-3.500 -0.0533 0.01116 0.00548 -0.0831 0.8976 0.5694
-3.250 -0.0263 0.01119 0.00551 -0.0826 0.8894 0.6004
-3.000 0.0014 0.01123 0.00551 -0.0821 0.8820 0.6263
-2.750 0.0270 0.01137 0.00567 -0.0811 0.8737 0.6505
-2.500 0.0536 0.01152 0.00578 -0.0803 0.8663 0.6752
-2.250 0.0784 0.01166 0.00593 -0.0791 0.8578 0.6933
-2.000 0.1047 0.01174 0.00598 -0.0782 0.8507 0.7093
-1.750 0.1309 0.01178 0.00598 -0.0777 0.8421 0.7217
-1.500 0.1578 0.01177 0.00592 -0.0771 0.8350 0.7319
-1.250 0.1836 0.01179 0.00592 -0.0765 0.8264 0.7415
-1.000 0.2115 0.01177 0.00582 -0.0763 0.8195 0.7520
-0.750 0.2370 0.01179 0.00585 -0.0756 0.8108 0.7606
-0.500 0.2644 0.01177 0.00577 -0.0752 0.8040 0.7700
-0.250 0.2907 0.01180 0.00580 -0.0748 0.7953 0.7799
0.000 0.3173 0.01178 0.00575 -0.0743 0.7886 0.7882
0.250 0.3437 0.01182 0.00579 -0.0740 0.7799 0.7984
0.500 0.3701 0.01181 0.00576 -0.0733 0.7732 0.8073
0.750 0.3955 0.01185 0.00584 -0.0728 0.7644 0.8168
1.000 0.4232 0.01186 0.00580 -0.0725 0.7578 0.8275
1.250 0.4470 0.01190 0.00591 -0.0715 0.7491 0.8365
1.500 0.4734 0.01190 0.00588 -0.0710 0.7424 0.8466
1.750 0.4986 0.01195 0.00599 -0.0704 0.7336 0.8579
2.000 0.5235 0.01194 0.00597 -0.0695 0.7270 0.8680
2.250 0.5467 0.01196 0.00607 -0.0684 0.7179 0.8792
2.500 0.5720 0.01194 0.00604 -0.0676 0.7114 0.8909
2.750 0.5944 0.01194 0.00613 -0.0664 0.7020 0.9039
3.000 0.6182 0.01190 0.00609 -0.0653 0.6948 0.9173
3.250 0.6401 0.01183 0.00610 -0.0639 0.6853 0.9323
3.500 0.6658 0.01175 0.00608 -0.0632 0.6760 0.9485
3.750 0.6989 0.01164 0.00596 -0.0640 0.6659 0.9651
4.000 0.7376 0.01155 0.00592 -0.0662 0.6520 0.9852
4.250 0.7689 0.01153 0.00593 -0.0672 0.6373 1.0000
4.500 0.7973 0.01153 0.00591 -0.0675 0.6211 1.0000
4.750 0.8245 0.01151 0.00591 -0.0675 0.5989 1.0000
5.000 0.8521 0.01156 0.00593 -0.0675 0.5773 1.0000
5.250 0.8794 0.01167 0.00606 -0.0676 0.5556 1.0000
5.500 0.9058 0.01183 0.00618 -0.0674 0.5307 1.0000
5.750 0.9307 0.01205 0.00638 -0.0669 0.4976 1.0000
6.000 0.9534 0.01238 0.00660 -0.0661 0.4487 1.0000
6.250 0.9692 0.01321 0.00698 -0.0642 0.3553 1.0000
6.500 0.9740 0.01510 0.00798 -0.0614 0.2118 1.0000
6.750 0.9777 0.01732 0.00939 -0.0585 0.1050 1.0000
7.000 0.9900 0.01876 0.01063 -0.0565 0.0767 1.0000
7.250 1.0041 0.01996 0.01178 -0.0548 0.0648 1.0000
7.500 1.0193 0.02102 0.01284 -0.0532 0.0575 1.0000
7.750 1.0315 0.02235 0.01420 -0.0511 0.0529 1.0000
8.000 1.0466 0.02339 0.01527 -0.0494 0.0489 1.0000
8.250 1.0583 0.02483 0.01666 -0.0475 0.0455 1.0000
8.500 1.0751 0.02619 0.01809 -0.0461 0.0429 1.0000
8.750 1.0947 0.02746 0.01943 -0.0450 0.0409 1.0000
9.000 1.1161 0.02884 0.02087 -0.0444 0.0391 1.0000
9.250 1.1399 0.03042 0.02247 -0.0441 0.0376 1.0000
9.500 1.1763 0.03365 0.02572 -0.0460 0.0357 1.0000
9.750 1.1955 0.03527 0.02758 -0.0450 0.0347 1.0000
10.000 1.2142 0.03707 0.02962 -0.0440 0.0339 1.0000
10.250 1.2327 0.03948 0.03232 -0.0430 0.0333 1.0000
10.500 1.2475 0.04223 0.03537 -0.0417 0.0330 1.0000
10.750 1.2574 0.04528 0.03876 -0.0400 0.0329 1.0000
11.000 1.2622 0.04866 0.04250 -0.0379 0.0331 1.0000
11.250 1.2600 0.05207 0.04623 -0.0352 0.0334 1.0000
11.500 1.2527 0.05542 0.04986 -0.0322 0.0336 1.0000
11.750 1.2410 0.05883 0.05355 -0.0295 0.0337 1.0000
12.000 1.2268 0.06254 0.05751 -0.0273 0.0339 1.0000
12.250 1.2100 0.06648 0.06169 -0.0257 0.0340 1.0000
12.500 1.1923 0.07096 0.06639 -0.0250 0.0342 1.0000
12.750 1.1730 0.07591 0.07154 -0.0250 0.0345 1.0000
13.000 1.1533 0.08150 0.07731 -0.0258 0.0348 1.0000
13.250 1.1583 0.08612 0.08201 -0.0261 0.0360 1.0000
13.500 0.7877 0.13131 0.12857 -0.0524 0.0573 1.0000
13.750 0.7291 0.15114 0.14834 -0.0648 0.0739 1.0000
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Polar data table (+)
Polar graphs
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