NACA M23 AIRFOIL (m23-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M23 AIRFOIL (m23-il) Reynolds number: 1,000,000 Max Cl/Cd: 128.57 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m23-il-1000000.txt Download as CSV file: xf-m23-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M23 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4421 0.12130 0.11881 0.0315 0.6457 0.0095
-9.000 -0.4367 0.11783 0.11535 0.0298 0.6413 0.0095
-8.750 -0.4331 0.11360 0.11111 0.0292 0.6368 0.0097
-8.500 -0.4270 0.11054 0.10801 0.0280 0.6320 0.0098
-8.250 -0.4207 0.10756 0.10504 0.0267 0.6276 0.0099
-8.000 -0.4144 0.10459 0.10206 0.0251 0.6228 0.0101
-7.750 -0.4084 0.10157 0.09902 0.0233 0.6183 0.0103
-7.500 -0.4030 0.09849 0.09594 0.0212 0.6142 0.0106
-7.250 -0.3977 0.09543 0.09287 0.0189 0.6100 0.0111
-7.000 -0.3852 0.09153 0.08894 0.0148 0.6059 0.0123
-6.500 -0.3549 0.08362 0.08095 0.0060 0.5982 0.0125
-6.250 -0.2818 0.06532 0.06276 -0.0011 0.5840 0.0128
-6.000 -0.2706 0.06213 0.05955 -0.0022 0.5799 0.0130
-5.750 -0.2574 0.05897 0.05634 -0.0038 0.5758 0.0134
-5.500 -0.2429 0.05551 0.05283 -0.0057 0.5718 0.0139
-5.250 -0.2268 0.05177 0.04906 -0.0078 0.5683 0.0149
-4.750 -0.2137 0.05933 0.05620 -0.0114 0.5695 0.0160
-4.500 -0.1889 0.05575 0.05252 -0.0132 0.5656 0.0160
-3.750 -0.1263 0.04688 0.04345 -0.0159 0.5537 0.0172
-3.500 -0.1006 0.04464 0.04112 -0.0169 0.5494 0.0184
-3.250 -0.0632 0.04204 0.03830 -0.0183 0.5456 0.0202
-3.000 -0.0354 0.03924 0.03535 -0.0188 0.5419 0.0203
-2.750 -0.0103 0.03417 0.03008 -0.0190 0.5388 0.0207
-2.500 0.0130 0.03245 0.02829 -0.0193 0.5345 0.0211
-2.250 0.0380 0.03105 0.02681 -0.0196 0.5303 0.0216
-2.000 0.0644 0.02958 0.02523 -0.0199 0.5266 0.0223
-1.750 0.0921 0.02795 0.02350 -0.0200 0.5229 0.0236
-1.500 0.1261 0.02660 0.02196 -0.0196 0.5190 0.0259
-1.250 0.1546 0.02452 0.01968 -0.0193 0.5151 0.0260
-1.000 0.1826 0.02244 0.01742 -0.0190 0.5116 0.0260
-0.750 0.2073 0.01862 0.01330 -0.0185 0.5085 0.0272
-0.500 0.2346 0.01790 0.01251 -0.0187 0.5045 0.0278
-0.250 0.2623 0.01714 0.01164 -0.0189 0.5003 0.0287
0.000 0.2906 0.01615 0.01052 -0.0188 0.4967 0.0304
0.250 0.3207 0.01608 0.01035 -0.0187 0.4929 0.0335
0.500 0.3482 0.01345 0.00733 -0.0183 0.4895 0.0350
0.750 0.3763 0.01281 0.00661 -0.0185 0.4855 0.0363
1.000 0.4046 0.01237 0.00614 -0.0187 0.4818 0.0376
1.250 0.4332 0.01194 0.00565 -0.0188 0.4780 0.0393
1.500 0.4619 0.01166 0.00530 -0.0190 0.4742 0.0418
1.750 0.4902 0.01114 0.00467 -0.0192 0.4701 0.0477
2.000 0.5187 0.01082 0.00437 -0.0194 0.4667 0.0502
2.250 0.5470 0.00961 0.00297 -0.0186 0.4631 0.0320
2.500 0.5750 0.00942 0.00275 -0.0187 0.4592 0.0327
2.750 0.6031 0.00934 0.00263 -0.0188 0.4551 0.0333
3.000 0.6312 0.00920 0.00251 -0.0190 0.4519 0.0335
3.250 0.6593 0.00909 0.00241 -0.0191 0.4479 0.0337
3.500 0.6871 0.00897 0.00226 -0.0192 0.4439 0.0345
3.750 0.7152 0.00893 0.00219 -0.0194 0.4400 0.0353
4.000 0.7436 0.00887 0.00217 -0.0197 0.4366 0.0363
4.250 0.7720 0.00886 0.00216 -0.0199 0.4325 0.0394
4.500 0.8004 0.00890 0.00219 -0.0202 0.4284 0.0440
4.750 0.8282 0.00883 0.00234 -0.0205 0.4244 0.1610
5.000 0.8471 0.00728 0.00258 -0.0187 0.4185 0.9803
5.250 0.9197 0.00764 0.00290 -0.0291 0.4087 0.9956
5.500 0.9805 0.00787 0.00309 -0.0368 0.3983 0.9997
5.750 1.0096 0.00795 0.00318 -0.0374 0.3895 1.0000
6.000 1.0359 0.00809 0.00327 -0.0374 0.3780 1.0000
6.250 1.0620 0.00826 0.00339 -0.0374 0.3639 1.0000
6.500 1.0880 0.00855 0.00359 -0.0375 0.3402 1.0000
6.750 1.1127 0.00949 0.00412 -0.0380 0.2670 1.0000
7.000 1.1266 0.01362 0.00687 -0.0393 0.0193 1.0000
7.250 1.1495 0.01408 0.00735 -0.0389 0.0158 1.0000
7.500 1.1721 0.01452 0.00785 -0.0385 0.0146 1.0000
7.750 1.1941 0.01498 0.00835 -0.0379 0.0138 1.0000
8.000 1.2152 0.01552 0.00894 -0.0373 0.0129 1.0000
8.250 1.2350 0.01614 0.00960 -0.0366 0.0120 1.0000
8.500 1.2513 0.01710 0.01065 -0.0356 0.0110 1.0000
8.750 1.2644 0.01821 0.01188 -0.0341 0.0104 1.0000
9.000 1.2800 0.01891 0.01263 -0.0329 0.0102 1.0000
9.250 1.2924 0.01979 0.01357 -0.0314 0.0099 1.0000
9.500 1.3010 0.02087 0.01472 -0.0296 0.0096 1.0000
9.750 1.2994 0.02250 0.01643 -0.0275 0.0093 1.0000
10.000 1.2976 0.02437 0.01838 -0.0255 0.0091 1.0000
10.250 1.3003 0.02644 0.02052 -0.0248 0.0089 1.0000
10.500 1.3035 0.02865 0.02281 -0.0243 0.0086 1.0000
10.750 1.3068 0.03091 0.02513 -0.0239 0.0084 1.0000
11.000 1.3089 0.03333 0.02760 -0.0236 0.0081 1.0000
11.250 1.3074 0.03612 0.03045 -0.0232 0.0079 1.0000
11.500 1.3005 0.03947 0.03387 -0.0227 0.0076 1.0000
11.750 1.2873 0.04336 0.03785 -0.0219 0.0074 1.0000
12.000 1.2771 0.04677 0.04133 -0.0208 0.0073 1.0000
12.250 1.2827 0.04887 0.04349 -0.0206 0.0072 1.0000
12.500 1.2873 0.05109 0.04578 -0.0203 0.0071 1.0000
12.750 1.2911 0.05340 0.04815 -0.0200 0.0069 1.0000
13.000 1.2951 0.05567 0.05049 -0.0196 0.0068 1.0000
13.250 1.2992 0.05791 0.05279 -0.0192 0.0067 1.0000
13.500 1.3036 0.06008 0.05502 -0.0187 0.0066 1.0000
13.750 1.3088 0.06215 0.05715 -0.0181 0.0064 1.0000
14.000 1.3145 0.06417 0.05923 -0.0176 0.0063 1.0000
14.250 1.3200 0.06621 0.06133 -0.0170 0.0061 1.0000
14.500 1.3254 0.06830 0.06349 -0.0165 0.0060 1.0000
14.750 1.3306 0.07046 0.06572 -0.0161 0.0058 1.0000
15.000 1.3356 0.07269 0.06803 -0.0158 0.0057 1.0000
15.250 1.3399 0.07510 0.07049 -0.0158 0.0056 1.0000
15.500 1.3431 0.07765 0.07311 -0.0157 0.0055 1.0000
15.750 1.3459 0.08030 0.07583 -0.0157 0.0054 1.0000
16.000 1.3475 0.08320 0.07880 -0.0161 0.0052 1.0000
16.250 1.3487 0.08605 0.08172 -0.0162 0.0051 1.0000
16.500 1.3469 0.08849 0.08431 -0.0136 0.0049 1.0000
16.750 1.3361 0.09310 0.08915 -0.0124 0.0049 1.0000
17.000 1.3290 0.09797 0.09415 -0.0154 0.0048 1.0000
17.250 1.3209 0.10274 0.09907 -0.0167 0.0048 1.0000
17.500 1.3114 0.10786 0.10435 -0.0184 0.0048 1.0000
17.750 1.2999 0.11340 0.11005 -0.0202 0.0048 1.0000
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