NACA M19 AIRFOIL (m19-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M19 AIRFOIL (m19-il) Reynolds number: 200,000 Max Cl/Cd: 75.45 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m19-il-200000-n5.txt Download as CSV file: xf-m19-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.2883 0.11450 0.11118 -0.0017 0.7809 0.0158
-9.250 -0.2868 0.11209 0.10875 -0.0031 0.7727 0.0159
-9.000 -0.2843 0.10929 0.10592 -0.0043 0.7650 0.0160
-8.750 -0.2815 0.10634 0.10295 -0.0056 0.7573 0.0160
-8.500 -0.2777 0.10307 0.09966 -0.0067 0.7500 0.0160
-8.250 -0.2737 0.09967 0.09624 -0.0079 0.7427 0.0160
-8.000 -0.2694 0.09623 0.09278 -0.0091 0.7358 0.0160
-7.750 -0.2649 0.09276 0.08929 -0.0105 0.7286 0.0160
-7.500 -0.2606 0.08925 0.08577 -0.0121 0.7223 0.0160
-7.250 -0.2536 0.08229 0.07880 -0.0103 0.7161 0.0164
-6.750 -0.2436 0.07465 0.07109 -0.0112 0.7034 0.0170
-6.500 -0.2364 0.07105 0.06745 -0.0126 0.6975 0.0173
-6.250 -0.2276 0.06741 0.06379 -0.0145 0.6914 0.0177
-6.000 -0.2605 0.08074 0.07683 -0.0211 0.7046 0.0162
-5.750 -0.2385 0.07754 0.07352 -0.0241 0.6979 0.0160
-5.500 -0.2365 0.07228 0.06829 -0.0220 0.6918 0.0169
-5.250 -0.2209 0.07013 0.06608 -0.0226 0.6846 0.0208
-5.000 -0.1869 0.06763 0.06343 -0.0283 0.6784 0.0238
-4.750 -0.1610 0.06407 0.05976 -0.0318 0.6720 0.0239
-4.500 -0.1370 0.06032 0.05588 -0.0344 0.6665 0.0238
-4.250 -0.1197 0.05591 0.05141 -0.0359 0.6602 0.0212
-4.000 -0.0917 0.05234 0.04767 -0.0387 0.6546 0.0208
-3.750 -0.0509 0.04976 0.04483 -0.0428 0.6487 0.0241
-3.500 -0.0268 0.04606 0.04099 -0.0442 0.6427 0.0236
-3.000 0.0375 0.04220 0.03670 -0.0478 0.6309 0.0343
-2.750 0.0694 0.03964 0.03386 -0.0489 0.6255 0.0346
-2.500 0.0973 0.03594 0.02991 -0.0499 0.6201 0.0351
-2.250 0.1146 0.03354 0.02752 -0.0503 0.6140 0.0383
-2.000 0.1536 0.03334 0.02686 -0.0508 0.6087 0.0462
-1.750 0.1771 0.02962 0.02301 -0.0516 0.6027 0.0478
-1.500 0.2017 0.02802 0.02131 -0.0519 0.5970 0.0501
-1.250 0.2347 0.02713 0.02002 -0.0521 0.5914 0.0604
-1.000 0.2592 0.02533 0.01818 -0.0526 0.5852 0.0625
-0.750 0.2913 0.02279 0.01518 -0.0522 0.5806 0.0440
-0.500 0.3200 0.02113 0.01329 -0.0523 0.5745 0.0413
-0.250 0.3493 0.01954 0.01137 -0.0521 0.5690 0.0398
0.000 0.3782 0.01827 0.00981 -0.0520 0.5638 0.0398
0.250 0.4068 0.01730 0.00861 -0.0520 0.5576 0.0405
0.500 0.4348 0.01664 0.00774 -0.0519 0.5523 0.0421
0.750 0.4631 0.01618 0.00713 -0.0519 0.5463 0.0446
1.000 0.4910 0.01554 0.00634 -0.0519 0.5406 0.0439
1.250 0.5188 0.01502 0.00569 -0.0518 0.5355 0.0431
1.500 0.5465 0.01457 0.00518 -0.0518 0.5292 0.0424
1.750 0.5738 0.01421 0.00474 -0.0517 0.5238 0.0418
2.000 0.6011 0.01391 0.00443 -0.0516 0.5180 0.0413
2.250 0.6282 0.01369 0.00420 -0.0516 0.5124 0.0410
2.500 0.6551 0.01353 0.00402 -0.0515 0.5075 0.0407
2.750 0.6824 0.01341 0.00393 -0.0515 0.5014 0.0407
3.000 0.7095 0.01336 0.00385 -0.0514 0.4959 0.0407
3.250 0.7367 0.01334 0.00383 -0.0514 0.4903 0.0411
3.500 0.7638 0.01335 0.00386 -0.0514 0.4846 0.0417
3.750 0.7909 0.01341 0.00389 -0.0513 0.4800 0.0428
4.000 0.8183 0.01348 0.00401 -0.0514 0.4744 0.0450
4.250 0.8455 0.01355 0.00418 -0.0514 0.4692 0.0591
4.500 0.8871 0.01225 0.00450 -0.0547 0.4624 1.0000
4.750 0.9129 0.01235 0.00454 -0.0545 0.4427 1.0000
5.000 0.9383 0.01252 0.00463 -0.0543 0.4216 1.0000
5.250 0.9635 0.01277 0.00482 -0.0540 0.3940 1.0000
5.500 0.9881 0.01315 0.00508 -0.0538 0.3548 1.0000
5.750 1.0013 0.01557 0.00632 -0.0533 0.1613 1.0000
6.000 1.0149 0.01791 0.00804 -0.0525 0.0231 1.0000
6.250 1.0373 0.01861 0.00890 -0.0520 0.0185 1.0000
6.500 1.0586 0.01943 0.00989 -0.0513 0.0164 1.0000
6.750 1.0775 0.02054 0.01126 -0.0505 0.0149 1.0000
7.000 1.0961 0.02155 0.01242 -0.0496 0.0135 1.0000
7.250 1.1128 0.02270 0.01373 -0.0487 0.0124 1.0000
7.500 1.1256 0.02414 0.01532 -0.0474 0.0119 1.0000
7.750 1.1350 0.02577 0.01711 -0.0459 0.0114 1.0000
8.000 1.1413 0.02752 0.01896 -0.0441 0.0111 1.0000
8.250 1.1435 0.02936 0.02089 -0.0421 0.0109 1.0000
8.500 1.1462 0.03138 0.02298 -0.0405 0.0106 1.0000
8.750 1.1512 0.03334 0.02499 -0.0392 0.0099 1.0000
9.000 1.1563 0.03534 0.02700 -0.0379 0.0092 1.0000
9.250 1.1626 0.03746 0.02905 -0.0355 0.0085 1.0000
9.500 1.1826 0.03920 0.03077 -0.0328 0.0082 1.0000
9.750 1.1995 0.04078 0.03248 -0.0314 0.0081 1.0000
10.000 1.2191 0.04258 0.03448 -0.0299 0.0079 1.0000
10.250 1.2392 0.04473 0.03680 -0.0286 0.0079 1.0000
10.500 1.2582 0.04732 0.03959 -0.0275 0.0079 1.0000
10.750 1.2741 0.05038 0.04288 -0.0263 0.0079 1.0000
11.750 1.2824 0.06331 0.05676 -0.0212 0.0080 1.0000
12.000 1.2733 0.06601 0.05969 -0.0205 0.0079 1.0000
12.250 1.2641 0.06894 0.06287 -0.0202 0.0076 1.0000
12.500 1.2548 0.07253 0.06669 -0.0201 0.0075 1.0000
12.750 1.2445 0.07638 0.07076 -0.0205 0.0073 1.0000
13.000 1.2331 0.08068 0.07528 -0.0212 0.0071 1.0000
13.250 1.2206 0.08534 0.08015 -0.0222 0.0071 1.0000
13.500 1.2068 0.09032 0.08532 -0.0236 0.0070 1.0000
13.750 1.1923 0.09564 0.09083 -0.0253 0.0071 1.0000
14.000 1.1770 0.10127 0.09663 -0.0275 0.0071 1.0000
14.250 1.1616 0.10727 0.10280 -0.0301 0.0071 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M19 AIRFOIL (m19-il)