NACA M19 AIRFOIL (m19-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M19 AIRFOIL (m19-il) Reynolds number: 1,000,000 Max Cl/Cd: 117.89 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m19-il-1000000.txt Download as CSV file: xf-m19-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4163 0.11120 0.10895 0.0179 0.7178 0.0076
-8.250 -0.4133 0.10678 0.10452 0.0166 0.7105 0.0077
-8.000 -0.4070 0.10360 0.10131 0.0155 0.7024 0.0078
-7.750 -0.4002 0.10063 0.09833 0.0140 0.6947 0.0079
-7.500 -0.3931 0.09772 0.09539 0.0121 0.6871 0.0081
-7.250 -0.3852 0.09477 0.09242 0.0098 0.6801 0.0082
-7.000 -0.3732 0.09155 0.08916 0.0067 0.6731 0.0084
-6.750 -0.3596 0.08824 0.08582 0.0035 0.6667 0.0088
-6.500 -0.3444 0.08484 0.08238 0.0000 0.6601 0.0092
-6.250 -0.3272 0.08125 0.07875 -0.0036 0.6543 0.0100
-6.000 -0.3018 0.07729 0.07474 -0.0094 0.6484 0.0107
-5.500 -0.2547 0.06925 0.06658 -0.0184 0.6373 0.0108
-5.250 -0.2394 0.06443 0.06171 -0.0208 0.6320 0.0112
-5.000 -0.2185 0.06181 0.05903 -0.0228 0.6264 0.0115
-4.750 -0.1949 0.05918 0.05633 -0.0253 0.6205 0.0122
-4.500 -0.1679 0.05632 0.05338 -0.0280 0.6151 0.0140
-4.250 -0.1300 0.05294 0.04988 -0.0325 0.6100 0.0147
-4.000 -0.0996 0.04921 0.04603 -0.0353 0.6048 0.0148
-3.750 -0.0700 0.04556 0.04224 -0.0375 0.5999 0.0149
-3.500 -0.0481 0.04105 0.03764 -0.0392 0.5951 0.0155
-3.250 -0.0241 0.03942 0.03594 -0.0400 0.5894 0.0161
-3.000 0.0029 0.03776 0.03418 -0.0410 0.5843 0.0179
-2.750 0.0414 0.03573 0.03198 -0.0423 0.5791 0.0199
-2.000 0.1266 0.02582 0.02152 -0.0445 0.5645 0.0213
-1.750 0.1525 0.02480 0.02041 -0.0449 0.5591 0.0221
-1.500 0.1801 0.02359 0.01909 -0.0452 0.5539 0.0235
-1.250 0.2107 0.02254 0.01791 -0.0452 0.5482 0.0267
-1.000 0.2408 0.02151 0.01672 -0.0450 0.5429 0.0275
-0.750 0.2700 0.01975 0.01478 -0.0449 0.5379 0.0276
-0.500 0.2988 0.01796 0.01276 -0.0448 0.5328 0.0277
-0.250 0.3268 0.01435 0.00870 -0.0447 0.5283 0.0288
0.000 0.3545 0.01333 0.00756 -0.0448 0.5227 0.0298
0.250 0.3823 0.01281 0.00696 -0.0450 0.5168 0.0307
0.500 0.4104 0.01229 0.00636 -0.0451 0.5112 0.0319
0.750 0.4386 0.01170 0.00565 -0.0451 0.5056 0.0333
1.000 0.4669 0.01121 0.00505 -0.0452 0.5002 0.0351
1.250 0.4955 0.01178 0.00561 -0.0453 0.4944 0.0381
1.500 0.5235 0.00993 0.00354 -0.0451 0.4889 0.0356
1.750 0.5516 0.00952 0.00307 -0.0451 0.4829 0.0361
2.000 0.5794 0.00918 0.00268 -0.0451 0.4759 0.0358
2.250 0.6071 0.00891 0.00239 -0.0451 0.4699 0.0353
2.500 0.6349 0.00872 0.00218 -0.0451 0.4641 0.0351
2.750 0.6627 0.00862 0.00204 -0.0452 0.4585 0.0354
3.000 0.6906 0.00848 0.00190 -0.0453 0.4529 0.0345
3.500 0.7464 0.00844 0.00178 -0.0455 0.4325 0.0337
3.750 0.7744 0.00844 0.00176 -0.0457 0.4228 0.0340
4.000 0.8022 0.00850 0.00177 -0.0458 0.4095 0.0354
4.250 0.8298 0.00863 0.00183 -0.0460 0.3911 0.0382
4.500 0.8573 0.00874 0.00196 -0.0462 0.3701 0.0860
4.750 0.8842 0.00750 0.00233 -0.0468 0.3325 0.9933
5.000 0.9133 0.01038 0.00386 -0.0497 0.0659 1.0000
5.500 0.9630 0.01137 0.00475 -0.0492 0.0149 1.0000
5.750 0.9883 0.01170 0.00512 -0.0489 0.0141 1.0000
6.000 1.0134 0.01210 0.00556 -0.0487 0.0130 1.0000
6.250 1.0380 0.01258 0.00609 -0.0484 0.0117 1.0000
6.500 1.0617 0.01320 0.00678 -0.0480 0.0105 1.0000
6.750 1.0829 0.01426 0.00796 -0.0474 0.0098 1.0000
7.000 1.1021 0.01551 0.00933 -0.0467 0.0093 1.0000
7.250 1.1257 0.01597 0.00983 -0.0463 0.0089 1.0000
7.500 1.1468 0.01675 0.01067 -0.0457 0.0084 1.0000
7.750 1.1651 0.01785 0.01184 -0.0448 0.0080 1.0000
8.000 1.1806 0.01916 0.01324 -0.0435 0.0079 1.0000
8.250 1.1969 0.02028 0.01442 -0.0424 0.0076 1.0000
8.500 1.2134 0.02130 0.01547 -0.0413 0.0072 1.0000
8.750 1.2266 0.02265 0.01687 -0.0398 0.0070 1.0000
9.000 1.2391 0.02413 0.01841 -0.0381 0.0070 1.0000
9.250 1.2527 0.02552 0.01985 -0.0366 0.0069 1.0000
9.500 1.2670 0.02695 0.02134 -0.0350 0.0068 1.0000
9.750 1.2797 0.02812 0.02252 -0.0336 0.0064 1.0000
10.000 1.2978 0.02989 0.02429 -0.0326 0.0061 1.0000
10.250 1.3267 0.03297 0.02745 -0.0325 0.0060 1.0000
10.500 1.3461 0.03647 0.03137 -0.0304 0.0068 1.0000
10.750 1.3590 0.03860 0.03362 -0.0293 0.0066 1.0000
11.000 1.3703 0.04086 0.03599 -0.0281 0.0064 1.0000
11.250 1.3804 0.04331 0.03854 -0.0269 0.0063 1.0000
11.500 1.3851 0.04604 0.04142 -0.0252 0.0062 1.0000
11.750 1.3854 0.04947 0.04501 -0.0233 0.0060 1.0000
12.000 1.3602 0.05933 0.05528 -0.0205 0.0057 1.0000
12.250 1.3464 0.06214 0.05825 -0.0184 0.0057 1.0000
12.500 1.3319 0.06532 0.06159 -0.0171 0.0057 1.0000
12.750 1.3167 0.06883 0.06526 -0.0165 0.0057 1.0000
13.000 1.3010 0.07267 0.06925 -0.0164 0.0057 1.0000
13.250 1.2847 0.07684 0.07356 -0.0168 0.0057 1.0000
13.500 1.2680 0.08134 0.07820 -0.0177 0.0057 1.0000
13.750 1.2507 0.08613 0.08314 -0.0190 0.0057 1.0000
14.000 1.2332 0.09123 0.08837 -0.0207 0.0057 1.0000
14.250 1.2152 0.09666 0.09393 -0.0228 0.0057 1.0000
14.500 1.1977 0.10235 0.09974 -0.0253 0.0057 1.0000
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Polar data table (+)
Polar graphs
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