NACA M22 AIRFOIL (m22-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M22 AIRFOIL (m22-il) Reynolds number: 100,000 Max Cl/Cd: 54.57 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m22-il-100000.txt Download as CSV file: xf-m22-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M22 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3773 0.11477 0.11112 0.0081 0.9146 0.0341
-8.250 -0.3736 0.11336 0.10968 0.0062 0.8968 0.0347
-8.000 -0.3699 0.11235 0.10865 0.0031 0.8815 0.0351
-7.750 -0.3648 0.11163 0.10788 -0.0008 0.8687 0.0353
-7.500 -0.3543 0.11067 0.10685 -0.0055 0.8573 0.0355
-7.250 -0.3392 0.10916 0.10526 -0.0103 0.8465 0.0356
-7.000 -0.3301 0.10377 0.09985 -0.0109 0.8378 0.0360
-6.750 -0.3280 0.09616 0.09226 -0.0061 0.8300 0.0368
-6.500 -0.3181 0.09194 0.08802 -0.0063 0.8211 0.0378
-6.250 -0.3080 0.08874 0.08475 -0.0070 0.8136 0.0389
-6.000 -0.2929 0.08551 0.08149 -0.0094 0.8047 0.0403
-5.750 -0.2778 0.08260 0.07851 -0.0115 0.7973 0.0424
-5.500 -0.2573 0.08000 0.07581 -0.0149 0.7892 0.0454
-5.250 -0.2136 0.08154 0.07703 -0.0241 0.7810 0.0477
-5.000 -0.1958 0.07730 0.07271 -0.0258 0.7743 0.0484
-4.750 -0.1911 0.07120 0.06668 -0.0240 0.7678 0.0497
-4.500 -0.1759 0.06762 0.06302 -0.0244 0.7608 0.0518
-4.250 -0.1546 0.06482 0.06011 -0.0259 0.7543 0.0550
-4.000 -0.1154 0.06430 0.05929 -0.0304 0.7469 0.0604
-3.750 -0.0931 0.06081 0.05567 -0.0318 0.7413 0.0624
-3.500 -0.0787 0.05686 0.05174 -0.0318 0.7341 0.0659
-3.250 -0.0546 0.05473 0.04940 -0.0324 0.7290 0.0713
-3.000 -0.0151 0.05356 0.04793 -0.0359 0.7211 0.0759
-2.750 -0.0007 0.04991 0.04423 -0.0351 0.7160 0.0791
-2.500 0.0419 0.05062 0.04448 -0.0377 0.7086 0.0888
-2.250 0.0533 0.04582 0.03980 -0.0371 0.7032 0.0930
-2.000 0.0874 0.04522 0.03885 -0.0383 0.6972 0.1033
-1.750 0.1074 0.04218 0.03582 -0.0385 0.6907 0.1088
-1.500 0.1347 0.04083 0.03415 -0.0383 0.6863 0.1188
-1.250 0.1660 0.03981 0.03294 -0.0397 0.6786 0.1314
-1.000 0.1907 0.03825 0.03119 -0.0393 0.6737 0.1452
-0.750 0.2153 0.03633 0.02923 -0.0400 0.6672 0.1616
-0.500 0.2414 0.03508 0.02782 -0.0400 0.6613 0.1814
-0.250 0.2650 0.03359 0.02613 -0.0392 0.6575 0.2044
0.000 0.2922 0.03249 0.02500 -0.0405 0.6494 0.2441
0.250 0.3139 0.03085 0.02328 -0.0396 0.6450 0.2937
0.500 0.3389 0.02949 0.02188 -0.0400 0.6386 0.3427
0.750 0.3652 0.02822 0.02054 -0.0398 0.6329 0.3753
1.000 0.3925 0.02733 0.01942 -0.0389 0.6292 0.3950
1.250 0.4266 0.02748 0.01945 -0.0405 0.6212 0.3830
1.500 0.4761 0.02883 0.01988 -0.0399 0.6169 0.1553
1.750 0.5072 0.02817 0.01898 -0.0399 0.6113 0.1235
2.000 0.5369 0.02777 0.01840 -0.0400 0.6050 0.1116
2.250 0.5647 0.02705 0.01746 -0.0389 0.6014 0.1051
2.500 0.5940 0.02737 0.01773 -0.0400 0.5941 0.1024
2.750 0.6241 0.02694 0.01725 -0.0401 0.5893 0.1027
3.000 0.6511 0.02632 0.01656 -0.0391 0.5862 0.1063
3.250 0.6776 0.02705 0.01741 -0.0404 0.5775 0.1179
3.500 0.7019 0.02685 0.01716 -0.0393 0.5739 0.1301
3.750 0.7266 0.02740 0.01779 -0.0397 0.5677 0.1514
4.000 0.7968 0.02619 0.01789 -0.0484 0.5613 1.0000
4.250 0.8204 0.02618 0.01774 -0.0469 0.5583 1.0000
4.500 0.8445 0.02784 0.01957 -0.0486 0.5498 1.0000
4.750 0.8683 0.02809 0.01978 -0.0477 0.5459 1.0000
5.000 0.8915 0.02887 0.02061 -0.0475 0.5408 1.0000
5.250 0.9137 0.03013 0.02199 -0.0481 0.5338 1.0000
5.500 0.9377 0.03027 0.02214 -0.0469 0.5308 1.0000
5.750 0.9565 0.03252 0.02465 -0.0483 0.5225 1.0000
6.000 0.9856 0.02955 0.02152 -0.0438 0.5163 1.0000
6.250 1.0121 0.02622 0.01798 -0.0392 0.5014 1.0000
6.500 1.0369 0.02501 0.01679 -0.0372 0.4881 1.0000
6.750 1.0620 0.02329 0.01505 -0.0347 0.4716 1.0000
7.000 1.0866 0.02135 0.01309 -0.0324 0.4411 1.0000
7.250 1.1111 0.02036 0.01218 -0.0311 0.4033 1.0000
7.500 1.1311 0.02073 0.01218 -0.0304 0.2998 1.0000
7.750 1.1210 0.02612 0.01589 -0.0304 0.0733 1.0000
8.000 1.1285 0.02841 0.01818 -0.0294 0.0571 1.0000
8.250 1.1352 0.03049 0.02040 -0.0284 0.0511 1.0000
8.500 1.1390 0.03272 0.02288 -0.0276 0.0486 1.0000
8.750 1.1358 0.03536 0.02568 -0.0267 0.0473 1.0000
9.000 1.1306 0.03827 0.02874 -0.0259 0.0463 1.0000
9.250 1.1254 0.04139 0.03199 -0.0253 0.0455 1.0000
9.500 1.1205 0.04454 0.03525 -0.0247 0.0448 1.0000
9.750 1.1174 0.04739 0.03818 -0.0238 0.0443 1.0000
10.000 1.1194 0.04937 0.04018 -0.0219 0.0437 1.0000
10.250 1.1414 0.04881 0.03941 -0.0170 0.0430 1.0000
10.500 1.1698 0.04888 0.03946 -0.0136 0.0415 1.0000
10.750 1.2224 0.04892 0.03955 -0.0097 0.0410 1.0000
11.000 1.3098 0.05180 0.04254 -0.0096 0.0440 1.0000
11.250 1.3235 0.05417 0.04554 -0.0074 0.0481 1.0000
11.500 1.3721 0.05915 0.05080 -0.0073 0.0559 1.0000
12.000 1.3901 0.06685 0.06001 -0.0023 0.0819 1.0000
12.250 1.3773 0.07100 0.06485 0.0002 0.0951 1.0000
12.500 1.2471 0.06428 0.05781 0.0087 0.0691 1.0000
12.750 1.2186 0.06818 0.06199 0.0099 0.0699 1.0000
13.000 1.1910 0.07271 0.06678 0.0104 0.0707 1.0000
13.250 1.1631 0.07770 0.07199 0.0104 0.0712 1.0000
13.500 1.1356 0.08287 0.07736 0.0099 0.0713 1.0000
13.750 1.1074 0.08823 0.08289 0.0089 0.0710 1.0000
14.000 1.0791 0.09362 0.08844 0.0076 0.0705 1.0000
14.250 1.0506 0.09900 0.09396 0.0060 0.0699 1.0000
14.500 1.0225 0.10433 0.09940 0.0041 0.0693 1.0000
14.750 0.9924 0.10929 0.10448 0.0022 0.0687 1.0000
15.000 0.9632 0.11448 0.10977 -0.0002 0.0682 1.0000
15.250 0.9345 0.12035 0.11572 -0.0034 0.0678 1.0000
15.500 0.9052 0.12696 0.12241 -0.0072 0.0675 1.0000
15.750 0.8741 0.13456 0.13007 -0.0119 0.0673 1.0000
16.000 0.8366 0.14440 0.13994 -0.0184 0.0673 1.0000
16.250 0.7847 0.16078 0.15617 -0.0285 0.0753 1.0000
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