NACA M19 AIRFOIL (m19-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: NACA M19 AIRFOIL (m19-il) Reynolds number: 100,000 Max Cl/Cd: 57.42 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m19-il-100000-n5.txt Download as CSV file: xf-m19-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M19 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3544 0.11707 0.11309 -0.0052 0.9296 0.0245
-8.250 -0.3474 0.11431 0.11026 -0.0074 0.8893 0.0245
-8.000 -0.3419 0.11170 0.10759 -0.0091 0.8664 0.0245
-7.750 -0.2372 0.08950 0.08537 -0.0145 0.8080 0.0257
-7.500 -0.2320 0.08582 0.08164 -0.0145 0.7988 0.0264
-6.750 -0.2964 0.09172 0.08738 -0.0144 0.8033 0.0257
-6.500 -0.2858 0.08812 0.08369 -0.0153 0.7939 0.0264
-6.250 -0.2725 0.08478 0.08031 -0.0175 0.7845 0.0270
-6.000 -0.2571 0.08149 0.07696 -0.0201 0.7754 0.0279
-5.750 -0.2404 0.07827 0.07365 -0.0228 0.7676 0.0290
-5.500 -0.2207 0.07501 0.07032 -0.0261 0.7586 0.0305
-5.250 -0.1958 0.07212 0.06730 -0.0304 0.7510 0.0330
-5.000 -0.1524 0.07094 0.06584 -0.0391 0.7425 0.0346
-4.750 -0.1325 0.06649 0.06130 -0.0412 0.7356 0.0352
-4.500 -0.1244 0.06181 0.05664 -0.0400 0.7283 0.0366
-4.250 -0.1042 0.05863 0.05334 -0.0413 0.7215 0.0387
-4.000 -0.0775 0.05577 0.05031 -0.0437 0.7140 0.0423
-3.750 -0.0374 0.05341 0.04762 -0.0484 0.7075 0.0472
-3.500 -0.0236 0.04997 0.04418 -0.0480 0.7003 0.0513
-3.250 0.0236 0.04981 0.04346 -0.0519 0.6939 0.0592
-3.000 0.0412 0.04514 0.03880 -0.0524 0.6869 0.0608
-2.750 0.0604 0.04257 0.03610 -0.0524 0.6812 0.0682
-2.500 0.0952 0.04060 0.03380 -0.0543 0.6738 0.0747
-2.250 0.1180 0.03807 0.03114 -0.0545 0.6678 0.0781
-2.000 0.1583 0.03828 0.03075 -0.0557 0.6607 0.0870
-1.750 0.1748 0.03435 0.02694 -0.0559 0.6547 0.0920
-1.500 0.2067 0.03321 0.02541 -0.0565 0.6482 0.1023
-1.250 0.2324 0.03120 0.02326 -0.0567 0.6417 0.1047
-1.000 0.2715 0.02845 0.01975 -0.0563 0.6369 0.0586
-0.750 0.2987 0.02676 0.01790 -0.0566 0.6297 0.0565
-0.500 0.3270 0.02533 0.01617 -0.0565 0.6243 0.0550
-0.250 0.3561 0.02407 0.01462 -0.0566 0.6174 0.0541
0.000 0.3849 0.02313 0.01334 -0.0564 0.6115 0.0570
0.250 0.4138 0.02242 0.01230 -0.0562 0.6052 0.0590
0.500 0.4416 0.02147 0.01118 -0.0562 0.5988 0.0581
0.750 0.4694 0.02070 0.01019 -0.0560 0.5935 0.0572
1.000 0.4974 0.02006 0.00944 -0.0560 0.5865 0.0564
1.250 0.5251 0.01949 0.00871 -0.0558 0.5812 0.0557
1.500 0.5530 0.01904 0.00819 -0.0558 0.5745 0.0552
1.750 0.5808 0.01863 0.00770 -0.0557 0.5686 0.0550
2.000 0.6088 0.01833 0.00731 -0.0556 0.5628 0.0550
2.250 0.6358 0.01812 0.00706 -0.0555 0.5565 0.0553
2.500 0.6622 0.01798 0.00680 -0.0551 0.5517 0.0560
2.750 0.6890 0.01797 0.00677 -0.0551 0.5448 0.0572
3.000 0.7156 0.01796 0.00668 -0.0548 0.5393 0.0591
3.250 0.7423 0.01800 0.00674 -0.0547 0.5336 0.0643
3.500 0.7691 0.01807 0.00691 -0.0547 0.5277 0.0828
4.000 0.8373 0.01698 0.00739 -0.0578 0.5161 1.0000
4.250 0.8632 0.01721 0.00761 -0.0575 0.5108 1.0000
4.500 0.8891 0.01748 0.00790 -0.0573 0.5056 1.0000
4.750 0.9149 0.01779 0.00830 -0.0572 0.4999 1.0000
5.000 0.9408 0.01805 0.00860 -0.0569 0.4955 1.0000
5.250 0.9658 0.01802 0.00864 -0.0564 0.4812 1.0000
5.500 0.9901 0.01794 0.00860 -0.0557 0.4606 1.0000
5.750 1.0141 0.01790 0.00855 -0.0550 0.4362 1.0000
6.000 1.0376 0.01807 0.00875 -0.0544 0.4046 1.0000
6.250 1.0588 0.01854 0.00906 -0.0536 0.3461 1.0000
6.500 1.0603 0.02195 0.01095 -0.0521 0.1084 1.0000
6.750 1.0686 0.02448 0.01301 -0.0507 0.0310 1.0000
7.000 1.0846 0.02579 0.01448 -0.0496 0.0256 1.0000
7.250 1.0986 0.02721 0.01611 -0.0484 0.0232 1.0000
7.500 1.1119 0.02859 0.01775 -0.0471 0.0223 1.0000
7.750 1.1217 0.03019 0.01959 -0.0456 0.0215 1.0000
8.000 1.1276 0.03200 0.02168 -0.0441 0.0207 1.0000
8.250 1.1278 0.03405 0.02391 -0.0423 0.0197 1.0000
8.500 1.1263 0.03648 0.02651 -0.0411 0.0187 1.0000
8.750 1.1247 0.03915 0.02931 -0.0403 0.0179 1.0000
9.000 1.1236 0.04182 0.03211 -0.0394 0.0174 1.0000
9.250 1.1243 0.04429 0.03466 -0.0382 0.0171 1.0000
9.500 1.1280 0.04636 0.03678 -0.0366 0.0167 1.0000
9.750 1.1370 0.04789 0.03830 -0.0344 0.0164 1.0000
10.000 1.1538 0.04897 0.03937 -0.0316 0.0161 1.0000
10.250 1.1803 0.04993 0.04033 -0.0287 0.0159 1.0000
10.500 1.0096 0.04807 0.03947 -0.0214 0.0168 1.0000
10.750 1.0430 0.04807 0.03941 -0.0178 0.0164 1.0000
11.000 1.0857 0.04891 0.04033 -0.0148 0.0157 1.0000
11.250 1.1325 0.05319 0.04475 -0.0127 0.0142 1.0000
11.500 1.1287 0.05635 0.04816 -0.0115 0.0140 1.0000
11.750 1.1290 0.06010 0.05216 -0.0103 0.0140 1.0000
12.000 1.1267 0.06418 0.05647 -0.0092 0.0141 1.0000
12.250 1.1227 0.06864 0.06116 -0.0082 0.0142 1.0000
12.750 1.2505 0.08243 0.07523 -0.0184 0.0150 1.0000
13.000 1.2400 0.08605 0.07909 -0.0187 0.0152 1.0000
13.250 1.2280 0.09004 0.08333 -0.0195 0.0154 1.0000
13.500 1.2143 0.09451 0.08805 -0.0207 0.0157 1.0000
13.750 1.1992 0.09949 0.09327 -0.0224 0.0160 1.0000
14.000 1.1804 0.10519 0.09923 -0.0250 0.0162 1.0000
14.250 1.1616 0.11134 0.10562 -0.0280 0.0165 1.0000
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Polar data table (+)
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