NACA M22 AIRFOIL (m22-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M22 AIRFOIL (m22-il) Reynolds number: 1,000,000 Max Cl/Cd: 91.02 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m22-il-1000000-n5.txt Download as CSV file: xf-m22-il-1000000-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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4473 0.10991 0.10729 0.0350 0.5882 0.0039
-7.750 -0.4472 0.10439 0.10179 0.0325 0.5849 0.0029
-7.500 -0.4430 0.10106 0.09845 0.0303 0.5806 0.0029
-7.250 -0.4353 0.09740 0.09477 0.0274 0.5768 0.0029
-7.000 -0.4244 0.09335 0.09071 0.0240 0.5727 0.0029
-6.750 -0.4120 0.08913 0.08647 0.0204 0.5683 0.0030
-6.500 -0.3967 0.08581 0.08311 0.0172 0.5639 0.0031
-6.250 -0.3800 0.08254 0.07981 0.0142 0.5598 0.0032
-6.000 -0.3619 0.07923 0.07646 0.0112 0.5554 0.0034
-5.750 -0.3424 0.07554 0.07272 0.0079 0.5511 0.0036
-5.500 -0.3214 0.07160 0.06872 0.0045 0.5473 0.0038
-5.250 -0.2987 0.06733 0.06439 0.0010 0.5433 0.0042
-5.000 -0.2737 0.06217 0.05914 -0.0028 0.5395 0.0045
-4.750 -0.2470 0.05700 0.05387 -0.0065 0.5360 0.0049
-4.500 -0.2217 0.05435 0.05113 -0.0085 0.5322 0.0051
-4.250 -0.1954 0.05143 0.04813 -0.0105 0.5278 0.0055
-4.000 -0.1669 0.04734 0.04390 -0.0129 0.5239 0.0063
-3.750 -0.1374 0.04289 0.03930 -0.0150 0.5204 0.0074
-3.500 -0.1107 0.04117 0.03750 -0.0160 0.5166 0.0079
-3.250 -0.0825 0.03854 0.03475 -0.0171 0.5125 0.0088
-3.000 -0.0513 0.03373 0.02970 -0.0183 0.5090 0.0107
-2.750 -0.0245 0.03292 0.02881 -0.0189 0.5050 0.0112
-2.500 0.0031 0.03148 0.02728 -0.0194 0.5011 0.0121
-2.250 0.0367 0.02550 0.02091 -0.0196 0.4980 0.0157
-2.000 0.0637 0.02513 0.02047 -0.0199 0.4936 0.0161
-1.750 0.0910 0.02444 0.01971 -0.0203 0.4901 0.0165
-1.500 0.1187 0.02351 0.01869 -0.0206 0.4862 0.0172
-1.250 0.1469 0.02243 0.01749 -0.0208 0.4820 0.0184
-1.000 0.1759 0.02077 0.01565 -0.0208 0.4781 0.0195
-0.750 0.2052 0.01886 0.01354 -0.0207 0.4751 0.0201
-0.500 0.2344 0.01678 0.01122 -0.0204 0.4717 0.0209
-0.250 0.2635 0.01453 0.00867 -0.0200 0.4679 0.0219
0.000 0.2918 0.01325 0.00712 -0.0200 0.4639 0.0237
0.250 0.3198 0.01266 0.00646 -0.0202 0.4604 0.0247
0.500 0.3480 0.01242 0.00617 -0.0205 0.4567 0.0257
0.750 0.3763 0.01208 0.00577 -0.0207 0.4526 0.0266
1.000 0.4047 0.01170 0.00530 -0.0209 0.4485 0.0276
1.250 0.4331 0.01123 0.00475 -0.0210 0.4448 0.0281
1.500 0.4615 0.01077 0.00420 -0.0211 0.4409 0.0281
1.750 0.4898 0.01046 0.00383 -0.0213 0.4370 0.0285
2.000 0.5181 0.01017 0.00349 -0.0215 0.4334 0.0289
2.250 0.5462 0.00982 0.00310 -0.0216 0.4296 0.0283
2.500 0.5741 0.00952 0.00277 -0.0217 0.4252 0.0277
2.750 0.6020 0.00931 0.00251 -0.0218 0.4209 0.0271
3.000 0.6299 0.00912 0.00231 -0.0219 0.4175 0.0266
3.250 0.6579 0.00898 0.00217 -0.0220 0.4137 0.0262
3.500 0.6859 0.00890 0.00207 -0.0222 0.4095 0.0259
3.750 0.7141 0.00887 0.00203 -0.0225 0.4053 0.0256
4.000 0.7423 0.00885 0.00201 -0.0227 0.3962 0.0254
4.250 0.7704 0.00891 0.00202 -0.0230 0.3829 0.0254
4.500 0.7987 0.00896 0.00206 -0.0234 0.3745 0.0254
4.750 0.8268 0.00913 0.00215 -0.0238 0.3531 0.0256
5.000 0.8547 0.00939 0.00229 -0.0242 0.3255 0.0264
5.250 0.8822 0.00998 0.00261 -0.0249 0.2710 0.0295
5.500 0.9053 0.01286 0.00436 -0.0268 0.0212 0.0331
5.750 0.9325 0.01308 0.00466 -0.0271 0.0148 0.0786
6.000 0.9587 0.01299 0.00510 -0.0274 0.0120 0.4432
6.250 0.9808 0.01205 0.00543 -0.0265 0.0107 0.9861
6.500 1.0476 0.01278 0.00623 -0.0361 0.0075 1.0000
6.750 1.0720 0.01311 0.00657 -0.0359 0.0069 1.0000
7.000 1.0962 0.01349 0.00699 -0.0357 0.0061 1.0000
7.250 1.1202 0.01391 0.00742 -0.0355 0.0055 1.0000
7.500 1.1431 0.01455 0.00813 -0.0353 0.0048 1.0000
7.750 1.1662 0.01505 0.00868 -0.0351 0.0045 1.0000
8.000 1.1889 0.01558 0.00924 -0.0348 0.0040 1.0000
8.250 1.2107 0.01622 0.00995 -0.0345 0.0037 1.0000
8.500 1.2319 0.01687 0.01065 -0.0341 0.0035 1.0000
8.750 1.2523 0.01757 0.01139 -0.0337 0.0033 1.0000
9.000 1.2702 0.01853 0.01244 -0.0332 0.0030 1.0000
9.250 1.2828 0.02001 0.01403 -0.0324 0.0028 1.0000
9.500 1.2967 0.02119 0.01532 -0.0317 0.0028 1.0000
9.750 1.3064 0.02292 0.01716 -0.0315 0.0027 1.0000
10.000 1.3070 0.02520 0.01956 -0.0310 0.0026 1.0000
10.250 1.3076 0.02770 0.02216 -0.0305 0.0025 1.0000
10.500 1.3080 0.03033 0.02490 -0.0301 0.0024 1.0000
10.750 1.3094 0.03284 0.02749 -0.0297 0.0023 1.0000
11.000 1.3113 0.03527 0.03002 -0.0292 0.0022 1.0000
11.250 1.3149 0.03754 0.03236 -0.0287 0.0021 1.0000
11.500 1.3177 0.03984 0.03474 -0.0282 0.0020 1.0000
11.750 1.3238 0.04189 0.03685 -0.0280 0.0020 1.0000
12.000 1.3277 0.04412 0.03917 -0.0277 0.0019 1.0000
12.250 1.3306 0.04644 0.04157 -0.0272 0.0018 1.0000
12.500 1.3327 0.04883 0.04404 -0.0267 0.0018 1.0000
12.750 1.3352 0.05120 0.04649 -0.0263 0.0018 1.0000
13.000 1.3369 0.05366 0.04904 -0.0258 0.0017 1.0000
13.250 1.3377 0.05622 0.05170 -0.0252 0.0017 1.0000
13.500 1.3371 0.05894 0.05455 -0.0243 0.0016 1.0000
13.750 1.3315 0.06201 0.05783 -0.0217 0.0016 1.0000
14.000 1.3314 0.06487 0.06087 -0.0202 0.0015 1.0000
14.250 1.3249 0.06907 0.06537 -0.0165 0.0014 1.0000
14.500 1.3098 0.07511 0.07171 -0.0149 0.0013 1.0000
14.750 1.2914 0.08146 0.07829 -0.0155 0.0013 1.0000
15.000 1.2737 0.08770 0.08472 -0.0169 0.0012 1.0000
15.250 1.2559 0.09423 0.09143 -0.0188 0.0012 1.0000
15.500 1.2383 0.10094 0.09830 -0.0213 0.0012 1.0000
15.750 1.2199 0.10807 0.10558 -0.0242 0.0012 1.0000
16.000 1.2023 0.11540 0.11306 -0.0276 0.0012 1.0000
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Polar data table (+)
Polar graphs
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