NACA M17 AIRFOIL (m17-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M17 AIRFOIL (m17-il) Reynolds number: 1,000,000 Max Cl/Cd: 122.29 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m17-il-1000000.txt Download as CSV file: xf-m17-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M17 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5070 0.11128 0.10985 0.0167 1.0000 0.0121
-9.250 -0.5003 0.10837 0.10695 0.0160 1.0000 0.0123
-9.000 -0.4938 0.10540 0.10399 0.0146 0.9643 0.0125
-8.750 -0.4934 0.10314 0.10149 0.0151 0.8758 0.0127
-8.500 -0.4891 0.10022 0.09845 0.0139 0.8463 0.0129
-8.250 -0.4848 0.09708 0.09522 0.0123 0.8249 0.0133
-8.000 -0.4807 0.09379 0.09187 0.0103 0.8074 0.0138
-7.500 -0.4716 0.08414 0.08211 -0.0016 0.7783 0.0154
-7.250 -0.4610 0.07938 0.07726 -0.0056 0.7659 0.0155
-7.000 -0.4488 0.07451 0.07231 -0.0093 0.7545 0.0155
-6.750 -0.4349 0.06960 0.06731 -0.0127 0.7438 0.0156
-6.500 -0.4269 0.06307 0.06066 -0.0160 0.7343 0.0159
-6.250 -0.4099 0.06150 0.05903 -0.0164 0.7225 0.0163
-6.000 -0.3911 0.05927 0.05673 -0.0176 0.7119 0.0168
-5.750 -0.3708 0.05654 0.05390 -0.0192 0.7023 0.0178
-5.000 -0.3038 0.03796 0.03460 -0.0250 0.6794 0.0207
-4.750 -0.2815 0.03732 0.03391 -0.0252 0.6707 0.0212
-4.500 -0.2581 0.03644 0.03295 -0.0255 0.6624 0.0220
-3.750 -0.1862 0.02217 0.01754 -0.0237 0.6438 0.0270
-3.500 -0.1609 0.02152 0.01685 -0.0239 0.6366 0.0278
-3.250 -0.1350 0.02085 0.01608 -0.0239 0.6301 0.0285
-3.000 -0.1087 0.01982 0.01492 -0.0238 0.6237 0.0296
-2.750 -0.0819 0.01860 0.01350 -0.0235 0.6173 0.0315
-2.500 -0.0515 0.01959 0.01443 -0.0234 0.6109 0.0342
-2.250 -0.0247 0.01822 0.01282 -0.0230 0.6052 0.0343
-2.000 -0.0008 0.01362 0.00772 -0.0222 0.6005 0.0321
-1.750 0.0269 0.01227 0.00616 -0.0219 0.5948 0.0311
-1.500 0.0548 0.01135 0.00508 -0.0218 0.5889 0.0306
-1.250 0.0827 0.01072 0.00435 -0.0217 0.5835 0.0306
-1.000 0.1107 0.01027 0.00384 -0.0216 0.5779 0.0311
-0.750 0.1385 0.00983 0.00332 -0.0215 0.5724 0.0315
-0.500 0.1664 0.00947 0.00294 -0.0214 0.5670 0.0321
-0.250 0.1942 0.00920 0.00263 -0.0214 0.5612 0.0327
0.000 0.2221 0.00900 0.00239 -0.0213 0.5560 0.0332
0.250 0.2501 0.00879 0.00217 -0.0213 0.5506 0.0336
0.500 0.2781 0.00864 0.00198 -0.0213 0.5451 0.0338
0.750 0.3061 0.00850 0.00182 -0.0214 0.5397 0.0341
1.000 0.3343 0.00837 0.00169 -0.0214 0.5342 0.0345
1.250 0.3622 0.00822 0.00147 -0.0214 0.5288 0.0363
1.500 0.3905 0.00811 0.00137 -0.0215 0.5236 0.0390
1.750 0.4188 0.00806 0.00132 -0.0216 0.5180 0.0425
2.000 0.4470 0.00803 0.00130 -0.0217 0.5127 0.0535
2.250 0.4745 0.00774 0.00137 -0.0217 0.5069 0.1917
2.500 0.4884 0.00585 0.00150 -0.0189 0.5010 0.9658
2.750 0.5381 0.00600 0.00163 -0.0238 0.4926 0.9899
3.000 0.6013 0.00622 0.00178 -0.0318 0.4824 0.9984
3.250 0.6384 0.00630 0.00183 -0.0339 0.4740 1.0000
3.500 0.6650 0.00636 0.00187 -0.0338 0.4685 1.0000
3.750 0.6916 0.00640 0.00192 -0.0336 0.4629 1.0000
4.000 0.7179 0.00648 0.00197 -0.0333 0.4568 1.0000
4.250 0.7443 0.00654 0.00203 -0.0331 0.4504 1.0000
4.500 0.7705 0.00664 0.00211 -0.0329 0.4403 1.0000
4.750 0.7968 0.00672 0.00218 -0.0327 0.4299 1.0000
5.000 0.8229 0.00682 0.00226 -0.0324 0.4188 1.0000
5.250 0.8488 0.00696 0.00236 -0.0322 0.4026 1.0000
5.500 0.8744 0.00715 0.00249 -0.0319 0.3835 1.0000
5.750 0.8998 0.00737 0.00264 -0.0316 0.3579 1.0000
6.000 0.9241 0.00784 0.00289 -0.0313 0.3081 1.0000
6.250 0.9456 0.00888 0.00345 -0.0309 0.2156 1.0000
6.500 0.9659 0.01004 0.00415 -0.0303 0.1301 1.0000
6.750 0.9824 0.01172 0.00525 -0.0293 0.0251 1.0000
7.000 1.0051 0.01224 0.00575 -0.0286 0.0190 1.0000
7.250 1.0286 0.01259 0.00613 -0.0280 0.0175 1.0000
7.500 1.0517 0.01300 0.00659 -0.0274 0.0160 1.0000
7.750 1.0734 0.01360 0.00723 -0.0266 0.0144 1.0000
8.000 1.0945 0.01424 0.00795 -0.0258 0.0135 1.0000
8.250 1.1168 0.01469 0.00843 -0.0251 0.0129 1.0000
8.500 1.1387 0.01517 0.00895 -0.0244 0.0121 1.0000
8.750 1.1600 0.01572 0.00954 -0.0237 0.0114 1.0000
9.000 1.1798 0.01642 0.01029 -0.0228 0.0108 1.0000
9.250 1.1920 0.01789 0.01187 -0.0210 0.0100 1.0000
9.500 1.2094 0.01873 0.01279 -0.0199 0.0098 1.0000
9.750 1.2277 0.01945 0.01357 -0.0189 0.0095 1.0000
10.000 1.2441 0.02030 0.01449 -0.0178 0.0092 1.0000
10.250 1.2586 0.02125 0.01551 -0.0164 0.0089 1.0000
10.500 1.2712 0.02226 0.01659 -0.0149 0.0086 1.0000
10.750 1.2805 0.02325 0.01765 -0.0130 0.0083 1.0000
11.000 1.2897 0.02432 0.01878 -0.0115 0.0080 1.0000
11.250 1.2998 0.02553 0.02003 -0.0105 0.0078 1.0000
11.500 1.3067 0.02717 0.02172 -0.0097 0.0075 1.0000
11.750 1.3075 0.02951 0.02416 -0.0086 0.0073 1.0000
12.000 1.2980 0.03298 0.02776 -0.0063 0.0070 1.0000
12.250 1.3079 0.03456 0.02944 -0.0061 0.0069 1.0000
12.500 1.3156 0.03639 0.03137 -0.0057 0.0068 1.0000
12.750 1.3218 0.03841 0.03349 -0.0052 0.0067 1.0000
13.000 1.3264 0.04061 0.03580 -0.0048 0.0066 1.0000
13.250 1.3298 0.04297 0.03827 -0.0043 0.0065 1.0000
13.500 1.3320 0.04550 0.04092 -0.0040 0.0064 1.0000
13.750 1.3328 0.04821 0.04375 -0.0037 0.0063 1.0000
14.000 1.3323 0.05112 0.04680 -0.0035 0.0062 1.0000
14.250 1.3305 0.05422 0.05004 -0.0035 0.0061 1.0000
14.500 1.3267 0.05763 0.05359 -0.0036 0.0060 1.0000
14.750 1.3216 0.06127 0.05736 -0.0040 0.0060 1.0000
15.000 1.3154 0.06516 0.06139 -0.0046 0.0059 1.0000
15.250 1.3083 0.06938 0.06575 -0.0056 0.0058 1.0000
15.500 1.3001 0.07391 0.07042 -0.0068 0.0058 1.0000
15.750 1.2888 0.07905 0.07572 -0.0084 0.0057 1.0000
16.000 1.2780 0.08436 0.08117 -0.0103 0.0057 1.0000
16.250 1.2654 0.09014 0.08709 -0.0126 0.0056 1.0000
16.500 1.2515 0.09634 0.09344 -0.0152 0.0056 1.0000
16.750 1.2338 0.10349 0.10076 -0.0183 0.0056 1.0000
17.000 1.2145 0.11126 0.10869 -0.0220 0.0056 1.0000
17.250 1.1919 0.12004 0.11764 -0.0264 0.0056 1.0000
17.500 1.1647 0.13042 0.12821 -0.0319 0.0057 1.0000
17.750 1.1312 0.14313 0.14114 -0.0390 0.0058 1.0000
18.000 1.0914 0.15898 0.15719 -0.0481 0.0060 1.0000
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