NACA M22 AIRFOIL (m22-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M22 AIRFOIL (m22-il) Reynolds number: 200,000 Max Cl/Cd: 73.91 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m22-il-200000-n5.txt Download as CSV file: xf-m22-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4274 0.11838 0.11470 0.0274 0.6962 0.0151
-8.000 -0.4221 0.11582 0.11214 0.0245 0.6909 0.0152
-7.750 -0.4167 0.11331 0.10961 0.0213 0.6853 0.0152
-7.500 -0.4064 0.11030 0.10656 0.0176 0.6804 0.0153
-7.250 -0.3942 0.10683 0.10307 0.0144 0.6751 0.0153
-7.000 -0.3878 0.10042 0.09667 0.0156 0.6702 0.0155
-6.750 -0.3796 0.09584 0.09204 0.0157 0.6656 0.0158
-6.500 -0.3685 0.09210 0.08828 0.0141 0.6602 0.0161
-6.250 -0.3555 0.08862 0.08476 0.0119 0.6549 0.0165
-5.750 -0.3238 0.08185 0.07789 0.0067 0.6449 0.0175
-5.500 -0.3054 0.07849 0.07445 0.0038 0.6399 0.0182
-5.250 -0.2851 0.07519 0.07105 0.0009 0.6356 0.0195
-5.000 -0.2571 0.07234 0.06811 -0.0032 0.6299 0.0217
-4.750 -0.2254 0.06969 0.06531 -0.0077 0.6250 0.0223
-4.500 -0.1988 0.06635 0.06183 -0.0107 0.6207 0.0224
-4.250 -0.1740 0.06267 0.05805 -0.0130 0.6155 0.0224
-4.000 -0.1508 0.05883 0.05409 -0.0147 0.6108 0.0223
-3.750 -0.1330 0.05464 0.04981 -0.0156 0.6067 0.0215
-3.500 -0.0867 0.05295 0.04782 -0.0194 0.6014 0.0228
-3.250 -0.0696 0.04802 0.04279 -0.0202 0.5970 0.0237
-3.000 -0.0506 0.04552 0.04019 -0.0205 0.5928 0.0253
-2.750 -0.0125 0.04460 0.03905 -0.0223 0.5872 0.0321
-2.500 0.0219 0.04254 0.03671 -0.0233 0.5826 0.0326
-2.250 0.0506 0.03990 0.03381 -0.0238 0.5787 0.0327
-2.000 0.0736 0.03566 0.02943 -0.0246 0.5741 0.0338
-1.750 0.0944 0.03392 0.02765 -0.0250 0.5690 0.0373
-1.500 0.1334 0.03372 0.02702 -0.0251 0.5649 0.0443
-1.250 0.1605 0.03056 0.02363 -0.0255 0.5604 0.0453
-1.000 0.1840 0.02849 0.02151 -0.0260 0.5556 0.0471
-0.750 0.2177 0.02910 0.02174 -0.0258 0.5512 0.0575
-0.500 0.2430 0.02613 0.01864 -0.0263 0.5470 0.0590
0.000 0.3015 0.02234 0.01436 -0.0260 0.5381 0.0423
0.250 0.3302 0.02094 0.01269 -0.0260 0.5343 0.0411
0.500 0.3595 0.01970 0.01123 -0.0260 0.5295 0.0409
0.750 0.3885 0.01872 0.01003 -0.0259 0.5251 0.0417
1.000 0.4174 0.01842 0.00946 -0.0258 0.5212 0.0444
1.250 0.4459 0.01733 0.00825 -0.0260 0.5166 0.0432
1.500 0.4744 0.01659 0.00738 -0.0261 0.5120 0.0422
1.750 0.5027 0.01601 0.00666 -0.0261 0.5081 0.0414
2.000 0.5308 0.01552 0.00608 -0.0261 0.5040 0.0406
2.250 0.5588 0.01511 0.00563 -0.0262 0.4992 0.0400
2.500 0.5862 0.01477 0.00524 -0.0261 0.4949 0.0395
2.750 0.6134 0.01452 0.00494 -0.0261 0.4913 0.0392
3.000 0.6408 0.01431 0.00478 -0.0261 0.4866 0.0391
3.250 0.6680 0.01416 0.00465 -0.0261 0.4821 0.0392
3.500 0.6952 0.01407 0.00453 -0.0261 0.4782 0.0395
3.750 0.7223 0.01402 0.00450 -0.0261 0.4737 0.0402
4.000 0.7497 0.01402 0.00452 -0.0262 0.4692 0.0414
4.250 0.7772 0.01406 0.00453 -0.0263 0.4653 0.0436
4.500 0.8047 0.01411 0.00463 -0.0264 0.4613 0.0532
5.000 0.8964 0.01302 0.00527 -0.0349 0.4496 1.0000
5.250 0.9224 0.01306 0.00528 -0.0347 0.4340 1.0000
5.500 0.9484 0.01317 0.00545 -0.0346 0.4216 1.0000
5.750 0.9744 0.01330 0.00559 -0.0345 0.4034 1.0000
6.000 1.0000 0.01353 0.00576 -0.0345 0.3769 1.0000
6.250 1.0239 0.01432 0.00616 -0.0348 0.2987 1.0000
6.500 1.0334 0.01888 0.00919 -0.0364 0.0259 1.0000
6.750 1.0555 0.01964 0.01006 -0.0361 0.0204 1.0000
7.000 1.0767 0.02045 0.01104 -0.0357 0.0174 1.0000
7.250 1.0963 0.02143 0.01221 -0.0352 0.0158 1.0000
7.500 1.1131 0.02270 0.01368 -0.0346 0.0146 1.0000
7.750 1.1224 0.02469 0.01591 -0.0338 0.0131 1.0000
8.000 1.1297 0.02661 0.01801 -0.0330 0.0121 1.0000
8.250 1.1349 0.02869 0.02022 -0.0326 0.0117 1.0000
8.500 1.1327 0.03102 0.02266 -0.0313 0.0115 1.0000
8.750 1.1318 0.03347 0.02519 -0.0303 0.0112 1.0000
9.000 1.1322 0.03592 0.02772 -0.0294 0.0109 1.0000
9.250 1.1335 0.03828 0.03014 -0.0284 0.0107 1.0000
9.500 1.1364 0.04046 0.03238 -0.0272 0.0105 1.0000
9.750 1.1415 0.04238 0.03435 -0.0257 0.0103 1.0000
10.000 1.1491 0.04408 0.03610 -0.0243 0.0098 1.0000
10.250 1.1575 0.04574 0.03780 -0.0231 0.0090 1.0000
10.500 1.1660 0.04740 0.03952 -0.0219 0.0084 1.0000
10.750 1.1782 0.04878 0.04089 -0.0198 0.0079 1.0000
11.000 1.2031 0.04976 0.04187 -0.0162 0.0077 1.0000
11.250 1.2283 0.05165 0.04388 -0.0137 0.0075 1.0000
11.500 1.2407 0.05398 0.04641 -0.0124 0.0075 1.0000
11.750 1.2511 0.05666 0.04930 -0.0112 0.0074 1.0000
12.000 1.2595 0.05974 0.05260 -0.0101 0.0074 1.0000
12.500 1.2704 0.06659 0.05987 -0.0084 0.0076 1.0000
12.750 1.2662 0.06974 0.06322 -0.0081 0.0076 1.0000
13.000 1.2602 0.07301 0.06670 -0.0081 0.0077 1.0000
13.250 1.2528 0.07649 0.07040 -0.0083 0.0078 1.0000
13.500 1.2436 0.08046 0.07458 -0.0087 0.0079 1.0000
13.750 1.2330 0.08482 0.07914 -0.0095 0.0079 1.0000
14.000 1.2210 0.08949 0.08401 -0.0105 0.0079 1.0000
14.250 1.2081 0.09440 0.08916 -0.0118 0.0078 1.0000
14.500 1.1944 0.09949 0.09442 -0.0135 0.0078 1.0000
14.750 1.1799 0.10495 0.10005 -0.0155 0.0078 1.0000
15.000 1.1658 0.11061 0.10588 -0.0180 0.0078 1.0000
15.250 1.1518 0.11662 0.11206 -0.0209 0.0078 1.0000
15.500 1.1382 0.12282 0.11844 -0.0243 0.0079 1.0000
15.750 1.1245 0.12950 0.12528 -0.0281 0.0079 1.0000
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Polar data table (+)
Polar graphs
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