NACA M22 AIRFOIL (m22-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M22 AIRFOIL (m22-il) Reynolds number: 500,000 Max Cl/Cd: 81.88 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m22-il-500000-n5.txt Download as CSV file: xf-m22-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4380 0.11050 0.10749 0.0321 0.6297 0.0071
-7.750 -0.4321 0.10710 0.10408 0.0303 0.6251 0.0071
-7.250 -0.4214 0.10033 0.09730 0.0268 0.6161 0.0070
-7.000 -0.4126 0.09716 0.09410 0.0248 0.6111 0.0064
-6.750 -0.4012 0.09351 0.09041 0.0217 0.6067 0.0061
-6.500 -0.3879 0.08971 0.08660 0.0185 0.6021 0.0060
-6.250 -0.3729 0.08589 0.08274 0.0152 0.5975 0.0060
-5.750 -0.3361 0.07751 0.07426 0.0077 0.5891 0.0069
-5.500 -0.3145 0.07330 0.06999 0.0039 0.5845 0.0071
-5.250 -0.2939 0.06948 0.06609 0.0009 0.5800 0.0070
-5.000 -0.2708 0.06546 0.06200 -0.0022 0.5758 0.0070
-4.750 -0.2454 0.06108 0.05750 -0.0055 0.5715 0.0071
-4.500 -0.2211 0.05793 0.05426 -0.0079 0.5668 0.0074
-4.250 -0.1960 0.05539 0.05163 -0.0099 0.5624 0.0078
-4.000 -0.1695 0.05250 0.04865 -0.0119 0.5577 0.0085
-3.750 -0.1396 0.04838 0.04437 -0.0144 0.5536 0.0099
-3.250 -0.0826 0.04209 0.03780 -0.0175 0.5453 0.0111
-3.000 -0.0554 0.04027 0.03586 -0.0185 0.5406 0.0121
-2.750 -0.0238 0.03681 0.03218 -0.0197 0.5367 0.0143
-2.500 0.0084 0.03298 0.02809 -0.0206 0.5333 0.0158
-2.250 0.0333 0.03250 0.02758 -0.0210 0.5284 0.0167
-2.000 0.0609 0.03121 0.02617 -0.0216 0.5239 0.0183
-1.750 0.0906 0.02907 0.02381 -0.0219 0.5201 0.0195
-1.500 0.1227 0.02672 0.02123 -0.0220 0.5163 0.0217
-1.250 0.1516 0.02475 0.01906 -0.0222 0.5120 0.0242
-1.000 0.1781 0.02391 0.01812 -0.0226 0.5078 0.0254
-0.750 0.2064 0.02268 0.01675 -0.0228 0.5041 0.0266
-0.500 0.2354 0.02124 0.01514 -0.0228 0.5001 0.0274
-0.250 0.2645 0.02000 0.01371 -0.0229 0.4958 0.0289
0.000 0.2939 0.01819 0.01163 -0.0227 0.4922 0.0276
0.250 0.3233 0.01654 0.00975 -0.0225 0.4887 0.0276
0.500 0.3522 0.01531 0.00829 -0.0224 0.4848 0.0289
0.750 0.3809 0.01443 0.00722 -0.0224 0.4807 0.0300
1.000 0.4095 0.01379 0.00642 -0.0225 0.4768 0.0306
1.250 0.4379 0.01326 0.00579 -0.0226 0.4728 0.0310
1.500 0.4664 0.01301 0.00546 -0.0228 0.4688 0.0319
1.750 0.4947 0.01253 0.00489 -0.0229 0.4648 0.0319
2.000 0.5228 0.01200 0.00428 -0.0230 0.4610 0.0313
2.250 0.5509 0.01160 0.00384 -0.0231 0.4567 0.0309
2.500 0.5787 0.01129 0.00349 -0.0231 0.4526 0.0306
2.750 0.6064 0.01106 0.00322 -0.0232 0.4490 0.0302
3.000 0.6341 0.01085 0.00302 -0.0233 0.4452 0.0300
3.250 0.6618 0.01069 0.00287 -0.0234 0.4409 0.0298
3.500 0.6895 0.01060 0.00276 -0.0235 0.4365 0.0296
3.750 0.7174 0.01053 0.00271 -0.0237 0.4326 0.0296
4.000 0.7453 0.01048 0.00270 -0.0239 0.4286 0.0296
4.250 0.7733 0.01047 0.00271 -0.0241 0.4245 0.0299
4.500 0.8013 0.01051 0.00273 -0.0244 0.4158 0.0303
4.750 0.8293 0.01058 0.00277 -0.0247 0.4012 0.0312
5.000 0.8573 0.01069 0.00285 -0.0250 0.3911 0.0328
5.250 0.8851 0.01081 0.00298 -0.0253 0.3724 0.0479
5.750 0.9678 0.01370 0.00593 -0.0365 0.0252 1.0000
6.250 1.0171 0.01461 0.00692 -0.0362 0.0125 1.0000
6.500 1.0414 0.01503 0.00743 -0.0360 0.0115 1.0000
6.750 1.0652 0.01554 0.00803 -0.0358 0.0102 1.0000
7.000 1.0884 0.01613 0.00866 -0.0356 0.0087 1.0000
7.250 1.1088 0.01728 0.00995 -0.0354 0.0075 1.0000
7.500 1.1301 0.01804 0.01078 -0.0350 0.0072 1.0000
7.750 1.1495 0.01902 0.01188 -0.0346 0.0067 1.0000
8.000 1.1667 0.02015 0.01311 -0.0340 0.0063 1.0000
8.250 1.1834 0.02116 0.01419 -0.0334 0.0057 1.0000
8.500 1.1997 0.02212 0.01519 -0.0328 0.0052 1.0000
8.750 1.2085 0.02382 0.01697 -0.0321 0.0049 1.0000
9.000 1.2065 0.02648 0.01974 -0.0316 0.0047 1.0000
9.250 1.2081 0.02862 0.02198 -0.0307 0.0046 1.0000
9.500 1.2105 0.03091 0.02437 -0.0300 0.0045 1.0000
9.750 1.2122 0.03329 0.02687 -0.0292 0.0043 1.0000
10.000 1.2141 0.03562 0.02929 -0.0283 0.0042 1.0000
10.250 1.2169 0.03781 0.03156 -0.0271 0.0040 1.0000
10.500 1.2211 0.03980 0.03363 -0.0256 0.0039 1.0000
10.750 1.2275 0.04156 0.03546 -0.0238 0.0038 1.0000
11.000 1.2363 0.04312 0.03711 -0.0217 0.0037 1.0000
11.250 1.2477 0.04461 0.03868 -0.0195 0.0036 1.0000
11.500 1.2587 0.04628 0.04047 -0.0178 0.0035 1.0000
11.750 1.2658 0.04823 0.04250 -0.0172 0.0033 1.0000
12.000 1.2714 0.05034 0.04471 -0.0169 0.0032 1.0000
12.250 1.2749 0.05266 0.04712 -0.0168 0.0031 1.0000
12.500 1.2771 0.05521 0.04978 -0.0164 0.0030 1.0000
12.750 1.2801 0.05815 0.05293 -0.0150 0.0028 1.0000
13.000 1.2826 0.06134 0.05638 -0.0138 0.0027 1.0000
13.250 1.2811 0.06503 0.06030 -0.0129 0.0027 1.0000
13.500 1.2768 0.06916 0.06465 -0.0124 0.0026 1.0000
13.750 1.2685 0.07361 0.06930 -0.0125 0.0026 1.0000
14.000 1.2576 0.07842 0.07431 -0.0131 0.0025 1.0000
14.250 1.2450 0.08348 0.07956 -0.0140 0.0025 1.0000
14.500 1.2315 0.08890 0.08516 -0.0155 0.0025 1.0000
14.750 1.2172 0.09472 0.09116 -0.0173 0.0025 1.0000
15.000 1.2026 0.10079 0.09740 -0.0196 0.0025 1.0000
15.250 1.1875 0.10724 0.10400 -0.0223 0.0025 1.0000
15.500 1.1724 0.11398 0.11089 -0.0254 0.0025 1.0000
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