NACA M15 AIRFOIL (m15-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M15 AIRFOIL (m15-il) Reynolds number: 50,000 Max Cl/Cd: 29.99 at α=8.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m15-il-50000-n5.txt Download as CSV file: xf-m15-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M15 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3541 0.11169 0.10460 -0.0281 1.0000 0.1331
-9.250 -0.3618 0.10939 0.10240 -0.0309 1.0000 0.1380
-9.000 -0.3817 0.10755 0.10072 -0.0354 1.0000 0.1396
-8.750 -0.3595 0.10254 0.09572 -0.0332 1.0000 0.1424
-8.500 -0.3483 0.09930 0.09250 -0.0326 1.0000 0.1463
-8.250 -0.3490 0.09648 0.08977 -0.0334 1.0000 0.1496
-8.000 -0.3639 0.09430 0.08775 -0.0345 1.0000 0.1534
-7.750 -0.4062 0.09389 0.08755 -0.0356 1.0000 0.1559
-7.250 -0.4323 0.08277 0.07635 -0.0358 1.0000 0.0914
-6.750 -0.4047 0.07029 0.06331 -0.0469 0.9771 0.0785
-6.500 -0.3813 0.06632 0.05925 -0.0496 0.9678 0.0775
-6.250 -0.3571 0.06194 0.05462 -0.0534 0.9588 0.0771
-6.000 -0.3361 0.05786 0.05018 -0.0561 0.9477 0.0775
-5.750 -0.3130 0.05405 0.04600 -0.0585 0.9380 0.0777
-5.500 -0.2860 0.05050 0.04211 -0.0608 0.9299 0.0774
-5.250 -0.2634 0.04751 0.03878 -0.0617 0.9199 0.0770
-5.000 -0.2345 0.04443 0.03528 -0.0634 0.9126 0.0767
-4.750 -0.2119 0.04200 0.03243 -0.0635 0.9024 0.0767
-4.500 -0.1811 0.03949 0.02944 -0.0648 0.8958 0.0773
-4.250 -0.1577 0.03767 0.02712 -0.0644 0.8855 0.0788
-4.000 -0.1259 0.03597 0.02529 -0.0657 0.8793 0.0812
-3.750 -0.1023 0.03472 0.02382 -0.0653 0.8689 0.0828
-3.500 -0.0700 0.03317 0.02194 -0.0660 0.8621 0.0842
-3.250 -0.0435 0.03199 0.02050 -0.0656 0.8521 0.0858
-3.000 -0.0121 0.03078 0.01899 -0.0659 0.8440 0.0881
-2.750 0.0178 0.02983 0.01774 -0.0659 0.8347 0.0917
-2.500 0.0473 0.02900 0.01689 -0.0660 0.8258 0.0967
-2.250 0.0798 0.02825 0.01599 -0.0664 0.8175 0.1031
-2.000 0.1111 0.02756 0.01513 -0.0666 0.8087 0.1091
-1.750 0.1408 0.02691 0.01445 -0.0666 0.8008 0.1193
-1.500 0.1659 0.02646 0.01400 -0.0660 0.7916 0.1351
-1.250 0.1929 0.02581 0.01344 -0.0656 0.7841 0.1622
-1.000 0.3102 0.02220 0.01262 -0.0812 0.7812 1.0000
-0.750 0.3313 0.02242 0.01253 -0.0799 0.7706 1.0000
-0.500 0.3545 0.02257 0.01240 -0.0790 0.7617 1.0000
-0.250 0.3774 0.02272 0.01231 -0.0779 0.7525 1.0000
0.000 0.3990 0.02295 0.01235 -0.0768 0.7428 1.0000
0.250 0.4232 0.02305 0.01225 -0.0759 0.7348 1.0000
0.500 0.4436 0.02336 0.01242 -0.0747 0.7245 1.0000
0.750 0.4690 0.02342 0.01229 -0.0738 0.7172 1.0000
1.000 0.4884 0.02380 0.01259 -0.0725 0.7063 1.0000
1.250 0.5118 0.02399 0.01265 -0.0715 0.6981 1.0000
1.500 0.5334 0.02427 0.01284 -0.0704 0.6885 1.0000
1.750 0.5548 0.02459 0.01307 -0.0693 0.6792 1.0000
2.000 0.5784 0.02479 0.01318 -0.0683 0.6709 1.0000
2.250 0.5983 0.02521 0.01356 -0.0670 0.6609 1.0000
2.500 0.6233 0.02535 0.01361 -0.0662 0.6534 1.0000
2.750 0.6417 0.02588 0.01413 -0.0648 0.6430 1.0000
3.000 0.6678 0.02599 0.01416 -0.0641 0.6363 1.0000
3.250 0.6850 0.02662 0.01481 -0.0626 0.6254 1.0000
3.500 0.7099 0.02681 0.01495 -0.0618 0.6183 1.0000
3.750 0.7278 0.02743 0.01562 -0.0604 0.6082 1.0000
4.000 0.7507 0.02777 0.01594 -0.0595 0.6005 1.0000
4.250 0.7701 0.02832 0.01652 -0.0582 0.5913 1.0000
4.500 0.7912 0.02880 0.01703 -0.0572 0.5833 1.0000
4.750 0.8116 0.02930 0.01757 -0.0560 0.5747 1.0000
5.000 0.8310 0.02989 0.01821 -0.0548 0.5663 1.0000
5.250 0.8524 0.03035 0.01872 -0.0538 0.5584 1.0000
5.500 0.8701 0.03106 0.01951 -0.0525 0.5497 1.0000
5.750 0.8926 0.03145 0.01995 -0.0516 0.5422 1.0000
6.000 0.9084 0.03228 0.02088 -0.0501 0.5332 1.0000
6.250 0.9328 0.03253 0.02120 -0.0493 0.5261 1.0000
6.500 0.9461 0.03350 0.02229 -0.0477 0.5165 1.0000
6.750 0.9741 0.03346 0.02231 -0.0471 0.5097 1.0000
7.000 0.9849 0.03453 0.02351 -0.0452 0.4992 1.0000
7.250 1.0108 0.03456 0.02363 -0.0444 0.4914 1.0000
7.500 1.0269 0.03522 0.02441 -0.0428 0.4814 1.0000
7.750 1.0426 0.03594 0.02525 -0.0413 0.4717 1.0000
8.000 1.0707 0.03583 0.02525 -0.0406 0.4639 1.0000
8.250 1.0792 0.03706 0.02663 -0.0385 0.4534 1.0000
8.500 1.1081 0.03695 0.02661 -0.0380 0.4456 1.0000
8.750 1.1174 0.03811 0.02795 -0.0360 0.4353 1.0000
9.000 1.1319 0.03896 0.02895 -0.0344 0.4258 1.0000
9.250 1.1589 0.03896 0.02909 -0.0337 0.4168 1.0000
9.500 1.1598 0.04064 0.03096 -0.0310 0.4059 1.0000
9.750 1.1803 0.04103 0.03149 -0.0298 0.3962 1.0000
10.000 1.1959 0.04166 0.03228 -0.0281 0.3856 1.0000
10.250 1.1889 0.04354 0.03434 -0.0247 0.3740 1.0000
10.500 1.1931 0.04447 0.03538 -0.0221 0.3607 1.0000
10.750 1.1945 0.04545 0.03644 -0.0195 0.3455 1.0000
11.000 1.1918 0.04686 0.03793 -0.0171 0.3295 1.0000
11.250 1.1872 0.04865 0.03980 -0.0151 0.3129 1.0000
11.500 1.1831 0.05069 0.04191 -0.0136 0.2964 1.0000
11.750 1.1805 0.05284 0.04415 -0.0124 0.2802 1.0000
12.000 1.1795 0.05491 0.04627 -0.0113 0.2632 1.0000
12.250 1.1764 0.05736 0.04877 -0.0105 0.2455 1.0000
12.500 1.1691 0.06059 0.05206 -0.0101 0.2275 1.0000
12.750 1.1634 0.06362 0.05508 -0.0097 0.2081 1.0000
13.000 1.1578 0.06661 0.05794 -0.0092 0.1883 1.0000
13.250 1.1478 0.07051 0.06183 -0.0093 0.1696 1.0000
13.500 1.1376 0.07457 0.06581 -0.0095 0.1516 1.0000
13.750 1.1275 0.07881 0.06996 -0.0099 0.1355 1.0000
14.000 1.1179 0.08316 0.07423 -0.0104 0.1213 1.0000
14.250 1.1096 0.08750 0.07851 -0.0111 0.1094 1.0000
14.500 1.1029 0.09167 0.08262 -0.0117 0.0998 1.0000
14.750 1.0976 0.09564 0.08647 -0.0124 0.0923 1.0000
15.000 1.0939 0.09961 0.09049 -0.0131 0.0852 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M15 AIRFOIL (m15-il)