NACA M17 AIRFOIL (m17-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M17 AIRFOIL (m17-il) Reynolds number: 200,000 Max Cl/Cd: 74.21 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m17-il-200000-n5.txt Download as CSV file: xf-m17-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4614 0.10770 0.10453 0.0076 1.0000 0.0189
-8.750 -0.4569 0.10392 0.10078 0.0057 1.0000 0.0188
-8.500 -0.4549 0.09981 0.09671 0.0019 1.0000 0.0193
-7.250 -0.4270 0.08273 0.07949 -0.0088 0.8536 0.0227
-7.000 -0.4197 0.07793 0.07459 -0.0125 0.8357 0.0218
-6.750 -0.4091 0.07354 0.07009 -0.0157 0.8205 0.0217
-6.500 -0.3954 0.07036 0.06679 -0.0177 0.8063 0.0228
-6.250 -0.3792 0.06669 0.06299 -0.0204 0.7934 0.0247
-6.000 -0.3619 0.06235 0.05846 -0.0232 0.7819 0.0252
-5.750 -0.3434 0.05793 0.05386 -0.0255 0.7714 0.0253
-5.500 -0.3223 0.05334 0.04903 -0.0277 0.7614 0.0260
-5.250 -0.3004 0.04878 0.04422 -0.0297 0.7517 0.0292
-5.000 -0.2819 0.04764 0.04300 -0.0295 0.7415 0.0312
-4.750 -0.2591 0.04361 0.03869 -0.0304 0.7327 0.0311
-4.500 -0.2358 0.04037 0.03520 -0.0309 0.7238 0.0318
-4.000 -0.1870 0.03442 0.02864 -0.0312 0.7073 0.0348
-3.750 -0.1621 0.03114 0.02497 -0.0309 0.7000 0.0350
-3.500 -0.1362 0.02825 0.02169 -0.0306 0.6921 0.0363
-3.250 -0.1103 0.02569 0.01868 -0.0300 0.6852 0.0371
-3.000 -0.0841 0.02339 0.01599 -0.0296 0.6776 0.0372
-2.750 -0.0577 0.02161 0.01379 -0.0292 0.6710 0.0375
-2.500 -0.0304 0.02020 0.01205 -0.0289 0.6634 0.0379
-2.250 -0.0031 0.01921 0.01076 -0.0286 0.6570 0.0384
-2.000 0.0246 0.01834 0.00964 -0.0285 0.6498 0.0394
-1.750 0.0520 0.01729 0.00834 -0.0283 0.6433 0.0397
-1.500 0.0798 0.01650 0.00739 -0.0282 0.6368 0.0396
-1.250 0.1074 0.01577 0.00653 -0.0280 0.6304 0.0398
-1.000 0.1348 0.01511 0.00575 -0.0279 0.6245 0.0401
-0.750 0.1622 0.01453 0.00511 -0.0277 0.6179 0.0407
-0.500 0.1891 0.01408 0.00460 -0.0275 0.6124 0.0415
-0.250 0.2164 0.01371 0.00422 -0.0274 0.6061 0.0426
0.000 0.2435 0.01343 0.00392 -0.0272 0.6002 0.0440
0.250 0.2707 0.01322 0.00367 -0.0270 0.5945 0.0455
0.500 0.2981 0.01305 0.00348 -0.0270 0.5884 0.0473
0.750 0.3254 0.01294 0.00332 -0.0268 0.5833 0.0505
1.000 0.3529 0.01281 0.00321 -0.0268 0.5773 0.0554
1.250 0.3803 0.01271 0.00312 -0.0267 0.5715 0.0617
1.500 0.4074 0.01259 0.00306 -0.0266 0.5666 0.0878
1.750 0.4340 0.01233 0.00310 -0.0265 0.5607 0.1887
2.250 0.5396 0.01078 0.00338 -0.0373 0.5483 1.0000
2.500 0.5658 0.01087 0.00344 -0.0370 0.5428 1.0000
2.750 0.5917 0.01098 0.00347 -0.0366 0.5380 1.0000
3.000 0.6179 0.01108 0.00358 -0.0364 0.5319 1.0000
3.250 0.6439 0.01119 0.00367 -0.0360 0.5265 1.0000
3.500 0.6699 0.01131 0.00379 -0.0357 0.5213 1.0000
3.750 0.6959 0.01143 0.00393 -0.0354 0.5152 1.0000
4.000 0.7217 0.01156 0.00403 -0.0351 0.5091 1.0000
4.250 0.7475 0.01166 0.00418 -0.0348 0.4995 1.0000
4.500 0.7731 0.01177 0.00430 -0.0344 0.4892 1.0000
4.750 0.7987 0.01190 0.00442 -0.0340 0.4805 1.0000
5.000 0.8245 0.01204 0.00466 -0.0337 0.4729 1.0000
5.250 0.8501 0.01221 0.00485 -0.0334 0.4663 1.0000
5.500 0.8759 0.01237 0.00511 -0.0331 0.4589 1.0000
5.750 0.9013 0.01255 0.00533 -0.0327 0.4511 1.0000
6.000 0.9265 0.01271 0.00557 -0.0323 0.4374 1.0000
6.250 0.9513 0.01289 0.00579 -0.0318 0.4183 1.0000
6.500 0.9751 0.01314 0.00600 -0.0313 0.3851 1.0000
6.750 0.9972 0.01362 0.00629 -0.0306 0.3337 1.0000
7.000 1.0151 0.01472 0.00695 -0.0298 0.2454 1.0000
7.250 1.0287 0.01644 0.00811 -0.0286 0.1530 1.0000
7.500 1.0355 0.01888 0.00979 -0.0269 0.0413 1.0000
7.750 1.0520 0.02004 0.01092 -0.0257 0.0282 1.0000
8.000 1.0705 0.02089 0.01188 -0.0246 0.0236 1.0000
8.250 1.0870 0.02191 0.01301 -0.0234 0.0208 1.0000
8.500 1.0996 0.02324 0.01448 -0.0218 0.0190 1.0000
8.750 1.1140 0.02428 0.01567 -0.0204 0.0179 1.0000
9.000 1.1264 0.02541 0.01693 -0.0189 0.0166 1.0000
9.250 1.1363 0.02667 0.01829 -0.0172 0.0154 1.0000
9.500 1.1410 0.02810 0.01982 -0.0150 0.0148 1.0000
9.750 1.1437 0.02977 0.02158 -0.0130 0.0143 1.0000
10.000 1.1466 0.03169 0.02359 -0.0116 0.0138 1.0000
10.250 1.1490 0.03385 0.02582 -0.0103 0.0133 1.0000
10.500 1.1508 0.03618 0.02826 -0.0089 0.0129 1.0000
10.750 1.1547 0.03846 0.03061 -0.0074 0.0124 1.0000
11.000 1.1640 0.04015 0.03243 -0.0066 0.0120 1.0000
11.250 1.1727 0.04194 0.03436 -0.0058 0.0114 1.0000
11.500 1.1806 0.04388 0.03642 -0.0049 0.0109 1.0000
11.750 1.1880 0.04595 0.03863 -0.0039 0.0105 1.0000
12.000 1.1949 0.04815 0.04098 -0.0029 0.0102 1.0000
12.250 1.2005 0.05052 0.04350 -0.0021 0.0100 1.0000
12.500 1.2046 0.05309 0.04624 -0.0014 0.0098 1.0000
12.750 1.2070 0.05587 0.04919 -0.0009 0.0096 1.0000
13.000 1.2076 0.05889 0.05243 -0.0007 0.0094 1.0000
13.250 1.2061 0.06218 0.05591 -0.0006 0.0093 1.0000
13.500 1.2025 0.06576 0.05968 -0.0008 0.0092 1.0000
13.750 1.1970 0.06965 0.06376 -0.0013 0.0091 1.0000
14.000 1.1897 0.07382 0.06812 -0.0020 0.0091 1.0000
14.250 1.1810 0.07825 0.07273 -0.0031 0.0090 1.0000
14.500 1.1710 0.08299 0.07765 -0.0045 0.0089 1.0000
14.750 1.1598 0.08813 0.08297 -0.0063 0.0088 1.0000
15.000 1.1477 0.09369 0.08870 -0.0086 0.0088 1.0000
15.250 1.1346 0.09969 0.09487 -0.0113 0.0087 1.0000
15.500 1.1201 0.10624 0.10162 -0.0145 0.0087 1.0000
15.750 1.1050 0.11332 0.10888 -0.0182 0.0087 1.0000
16.000 1.0887 0.12113 0.11688 -0.0227 0.0088 1.0000
16.250 1.0696 0.13024 0.12620 -0.0281 0.0089 1.0000
16.500 1.0435 0.14222 0.13841 -0.0353 0.0091 1.0000
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Polar data table (+)
Polar graphs
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