NACA M17 AIRFOIL (m17-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M17 AIRFOIL (m17-il) Reynolds number: 500,000 Max Cl/Cd: 102.4 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m17-il-500000.txt Download as CSV file: xf-m17-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4742 0.10505 0.10307 0.0087 1.0000 0.0198
-8.500 -0.4724 0.10127 0.09932 0.0047 1.0000 0.0199
-8.250 -0.4693 0.09726 0.09534 0.0008 1.0000 0.0200
-8.000 -0.4645 0.09313 0.09123 -0.0034 1.0000 0.0200
-7.750 -0.4572 0.08902 0.08699 -0.0067 0.9041 0.0201
-7.500 -0.4506 0.08504 0.08285 -0.0091 0.8713 0.0201
-7.250 -0.4514 0.07992 0.07765 -0.0087 0.8510 0.0206
-7.000 -0.4415 0.07709 0.07474 -0.0088 0.8323 0.0209
-6.750 -0.4295 0.07407 0.07161 -0.0101 0.8161 0.0213
-6.500 -0.4156 0.07086 0.06831 -0.0121 0.8016 0.0220
-6.250 -0.3995 0.06754 0.06488 -0.0144 0.7884 0.0232
-6.000 -0.3678 0.06317 0.06027 -0.0206 0.7767 0.0258
-5.750 -0.3460 0.05913 0.05605 -0.0228 0.7656 0.0260
-5.500 -0.3249 0.05487 0.05161 -0.0244 0.7550 0.0261
-5.250 -0.3158 0.04822 0.04480 -0.0256 0.7462 0.0269
-5.000 -0.2976 0.04625 0.04273 -0.0257 0.7358 0.0276
-4.750 -0.2767 0.04432 0.04071 -0.0260 0.7257 0.0288
-4.500 -0.2533 0.04173 0.03795 -0.0265 0.7171 0.0307
-4.250 -0.2175 0.03939 0.03526 -0.0269 0.7085 0.0342
-4.000 -0.1999 0.03246 0.02791 -0.0269 0.7019 0.0352
-3.750 -0.1785 0.03081 0.02618 -0.0270 0.6933 0.0363
-3.500 -0.1548 0.02957 0.02484 -0.0270 0.6854 0.0377
-3.250 -0.1296 0.02788 0.02297 -0.0269 0.6778 0.0399
-3.000 -0.0973 0.02756 0.02228 -0.0261 0.6703 0.0449
-2.750 -0.0770 0.02276 0.01713 -0.0258 0.6640 0.0472
-2.500 -0.0514 0.02186 0.01617 -0.0259 0.6571 0.0489
-2.250 -0.0247 0.02093 0.01511 -0.0258 0.6501 0.0523
-2.000 0.0065 0.02203 0.01589 -0.0251 0.6434 0.0574
-1.750 0.0294 0.01833 0.01206 -0.0253 0.6370 0.0617
-1.500 0.0598 0.01556 0.00858 -0.0239 0.6316 0.0458
-1.250 0.0877 0.01410 0.00703 -0.0237 0.6254 0.0434
-1.000 0.1157 0.01299 0.00573 -0.0235 0.6194 0.0422
-0.750 0.1436 0.01222 0.00485 -0.0233 0.6135 0.0419
-0.500 0.1715 0.01165 0.00421 -0.0231 0.6071 0.0422
-0.250 0.1989 0.01126 0.00372 -0.0229 0.6017 0.0426
0.000 0.2267 0.01088 0.00334 -0.0228 0.5956 0.0431
0.250 0.2542 0.01060 0.00302 -0.0227 0.5898 0.0437
0.500 0.2819 0.01042 0.00281 -0.0226 0.5841 0.0446
0.750 0.3095 0.01017 0.00255 -0.0226 0.5781 0.0456
1.000 0.3369 0.00996 0.00228 -0.0224 0.5730 0.0476
1.250 0.3649 0.00981 0.00216 -0.0224 0.5671 0.0513
1.500 0.3929 0.00973 0.00207 -0.0224 0.5613 0.0570
1.750 0.4200 0.00950 0.00208 -0.0223 0.5562 0.1467
2.000 0.4556 0.00752 0.00232 -0.0241 0.5503 0.9836
2.250 0.5307 0.00777 0.00244 -0.0347 0.5433 1.0000
2.500 0.5573 0.00780 0.00246 -0.0344 0.5376 1.0000
2.750 0.5838 0.00785 0.00248 -0.0342 0.5318 1.0000
3.000 0.6102 0.00792 0.00251 -0.0340 0.5254 1.0000
3.250 0.6365 0.00796 0.00254 -0.0337 0.5176 1.0000
3.500 0.6627 0.00803 0.00257 -0.0334 0.5100 1.0000
3.750 0.6889 0.00809 0.00263 -0.0331 0.5022 1.0000
4.000 0.7151 0.00817 0.00270 -0.0329 0.4957 1.0000
4.250 0.7413 0.00825 0.00279 -0.0326 0.4895 1.0000
4.500 0.7674 0.00837 0.00290 -0.0323 0.4840 1.0000
4.750 0.7936 0.00844 0.00302 -0.0320 0.4774 1.0000
5.000 0.8194 0.00854 0.00311 -0.0317 0.4679 1.0000
5.250 0.8452 0.00864 0.00319 -0.0314 0.4559 1.0000
5.500 0.8710 0.00873 0.00332 -0.0311 0.4427 1.0000
5.750 0.8965 0.00887 0.00344 -0.0308 0.4263 1.0000
6.000 0.9221 0.00901 0.00358 -0.0305 0.4067 1.0000
6.250 0.9472 0.00925 0.00377 -0.0301 0.3825 1.0000
6.500 0.9712 0.00966 0.00403 -0.0297 0.3358 1.0000
6.750 0.9891 0.01116 0.00484 -0.0291 0.2098 1.0000
7.000 0.9981 0.01392 0.00654 -0.0276 0.0387 1.0000
7.250 1.0189 0.01472 0.00733 -0.0267 0.0277 1.0000
7.500 1.0400 0.01542 0.00811 -0.0258 0.0243 1.0000
7.750 1.0613 0.01603 0.00879 -0.0250 0.0220 1.0000
8.000 1.0805 0.01687 0.00968 -0.0240 0.0201 1.0000
8.250 1.0945 0.01826 0.01119 -0.0223 0.0187 1.0000
8.500 1.1133 0.01903 0.01204 -0.0213 0.0181 1.0000
8.750 1.1300 0.01996 0.01305 -0.0200 0.0174 1.0000
9.000 1.1459 0.02092 0.01410 -0.0187 0.0165 1.0000
9.250 1.1611 0.02188 0.01511 -0.0174 0.0156 1.0000
9.500 1.1735 0.02301 0.01627 -0.0159 0.0149 1.0000
9.750 1.1805 0.02453 0.01786 -0.0137 0.0144 1.0000
10.000 1.1804 0.02649 0.01988 -0.0106 0.0140 1.0000
10.250 1.1860 0.02853 0.02200 -0.0083 0.0137 1.0000
10.500 1.1977 0.02977 0.02334 -0.0070 0.0134 1.0000
10.750 1.2094 0.03120 0.02489 -0.0058 0.0132 1.0000
11.000 1.2213 0.03280 0.02660 -0.0044 0.0129 1.0000
11.250 1.2333 0.03456 0.02849 -0.0031 0.0127 1.0000
11.500 1.2443 0.03641 0.03048 -0.0018 0.0124 1.0000
11.750 1.2535 0.03831 0.03252 -0.0007 0.0121 1.0000
12.000 1.2609 0.04025 0.03459 0.0002 0.0117 1.0000
12.250 1.2671 0.04233 0.03680 0.0011 0.0114 1.0000
12.500 1.2714 0.04486 0.03952 0.0021 0.0112 1.0000
12.750 1.2723 0.04786 0.04271 0.0030 0.0112 1.0000
13.000 1.2682 0.05153 0.04664 0.0039 0.0113 1.0000
13.250 1.2580 0.05618 0.05160 0.0046 0.0116 1.0000
13.500 1.2430 0.06128 0.05698 0.0046 0.0119 1.0000
13.750 1.2261 0.06659 0.06253 0.0041 0.0122 1.0000
14.000 1.2082 0.07211 0.06826 0.0029 0.0125 1.0000
14.250 1.1894 0.07788 0.07422 0.0013 0.0127 1.0000
14.500 1.1692 0.08413 0.08064 -0.0008 0.0130 1.0000
14.750 1.0229 0.08170 0.07833 0.0099 0.0124 1.0000
15.000 1.0002 0.08791 0.08468 0.0080 0.0125 1.0000
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Polar data table (+)
Polar graphs
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