NACA M21 AIRFOIL (m21-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M21 AIRFOIL (m21-il) Reynolds number: 500,000 Max Cl/Cd: 107.26 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m21-il-500000.txt Download as CSV file: xf-m21-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M21 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.1989 0.09551 0.09242 -0.0247 0.6925 0.0305
-9.000 -0.1929 0.09261 0.08953 -0.0253 0.6867 0.0310
-8.750 -0.2939 0.09822 0.09510 -0.0184 0.7024 0.0303
-8.500 -0.2855 0.09561 0.09245 -0.0193 0.6958 0.0307
-8.250 -0.2782 0.09289 0.08970 -0.0206 0.6898 0.0312
-8.000 -0.2718 0.08994 0.08675 -0.0225 0.6835 0.0319
-7.750 -0.2670 0.08674 0.08351 -0.0250 0.6778 0.0328
-7.500 -0.2695 0.08189 0.07860 -0.0354 0.6735 0.0346
-7.250 -0.2586 0.07717 0.07380 -0.0410 0.6684 0.0347
-7.000 -0.2513 0.07333 0.06994 -0.0403 0.6631 0.0351
-6.750 -0.2381 0.07091 0.06747 -0.0400 0.6579 0.0354
-6.500 -0.2235 0.06850 0.06505 -0.0406 0.6522 0.0358
-6.250 -0.2079 0.06598 0.06246 -0.0419 0.6468 0.0365
-6.000 -0.1912 0.06324 0.05962 -0.0436 0.6420 0.0373
-5.750 -0.1616 0.05846 0.05458 -0.0502 0.6377 0.0404
-5.500 -0.1470 0.05366 0.04966 -0.0514 0.6332 0.0408
-5.250 -0.1313 0.05134 0.04729 -0.0513 0.6283 0.0412
-5.000 -0.1126 0.04949 0.04539 -0.0513 0.6234 0.0416
-4.750 -0.0921 0.04756 0.04341 -0.0516 0.6182 0.0423
-4.500 -0.0704 0.04547 0.04122 -0.0521 0.6135 0.0433
-4.250 -0.0469 0.04316 0.03874 -0.0527 0.6091 0.0451
-4.000 -0.0184 0.03785 0.03299 -0.0536 0.6054 0.0476
-3.750 0.0022 0.03603 0.03114 -0.0535 0.6004 0.0481
-3.500 0.0245 0.03474 0.02978 -0.0535 0.5958 0.0487
-3.250 0.0480 0.03345 0.02840 -0.0535 0.5914 0.0497
-3.000 0.0730 0.03182 0.02668 -0.0534 0.5867 0.0512
-2.750 0.1057 0.02997 0.02432 -0.0526 0.5824 0.0550
-2.500 0.1215 0.02282 0.01669 -0.0509 0.5792 0.0464
-2.250 0.1455 0.01981 0.01327 -0.0500 0.5754 0.0462
-2.000 0.1708 0.01670 0.00967 -0.0491 0.5713 0.0466
-1.750 0.1980 0.01508 0.00772 -0.0489 0.5668 0.0477
-1.500 0.2257 0.01455 0.00710 -0.0489 0.5624 0.0489
-1.250 0.2540 0.01445 0.00698 -0.0490 0.5580 0.0503
-1.000 0.2825 0.01417 0.00666 -0.0491 0.5534 0.0521
-0.750 0.3109 0.01369 0.00604 -0.0491 0.5489 0.0540
-0.500 0.3388 0.01314 0.00534 -0.0491 0.5447 0.0560
-0.250 0.3672 0.01293 0.00518 -0.0492 0.5404 0.0580
0.000 0.3956 0.01281 0.00505 -0.0493 0.5359 0.0605
0.250 0.4239 0.01275 0.00492 -0.0494 0.5316 0.0635
0.500 0.4515 0.01242 0.00454 -0.0494 0.5273 0.0667
0.750 0.4798 0.01230 0.00448 -0.0495 0.5230 0.0697
1.000 0.5080 0.01222 0.00438 -0.0496 0.5187 0.0731
1.250 0.5357 0.01211 0.00421 -0.0496 0.5145 0.0764
1.500 0.5630 0.01193 0.00405 -0.0496 0.5103 0.0802
1.750 0.5909 0.01184 0.00400 -0.0497 0.5058 0.0839
2.000 0.6190 0.01182 0.00397 -0.0498 0.5016 0.0873
2.250 0.6455 0.01164 0.00375 -0.0496 0.4976 0.0910
2.500 0.6729 0.01157 0.00373 -0.0496 0.4934 0.0953
2.750 0.7006 0.01151 0.00370 -0.0497 0.4889 0.0991
3.000 0.7281 0.01150 0.00367 -0.0497 0.4845 0.1021
3.250 0.7547 0.01146 0.00362 -0.0496 0.4805 0.1082
3.500 0.7824 0.01143 0.00364 -0.0497 0.4764 0.1134
3.750 0.8099 0.01140 0.00365 -0.0497 0.4719 0.1201
4.000 0.8371 0.01142 0.00368 -0.0498 0.4676 0.1312
4.250 0.8541 0.01039 0.00380 -0.0482 0.4637 0.6872
4.500 0.9864 0.01026 0.00442 -0.0707 0.4562 1.0000
4.750 1.0116 0.01038 0.00450 -0.0704 0.4519 1.0000
5.000 1.0367 0.01053 0.00461 -0.0701 0.4476 1.0000
5.250 1.0620 0.01062 0.00473 -0.0697 0.4432 1.0000
5.500 1.0870 0.01073 0.00485 -0.0694 0.4389 1.0000
5.750 1.1116 0.01091 0.00498 -0.0690 0.4348 1.0000
6.000 1.1363 0.01106 0.00514 -0.0686 0.4306 1.0000
6.250 1.1610 0.01117 0.00530 -0.0682 0.4260 1.0000
6.500 1.1853 0.01133 0.00545 -0.0678 0.4216 1.0000
6.750 1.2091 0.01157 0.00565 -0.0673 0.4173 1.0000
7.000 1.2333 0.01167 0.00583 -0.0668 0.4120 1.0000
7.250 1.2568 0.01182 0.00598 -0.0663 0.4057 1.0000
7.500 1.2798 0.01202 0.00618 -0.0656 0.3995 1.0000
7.750 1.3032 0.01216 0.00638 -0.0651 0.3932 1.0000
8.000 1.3253 0.01242 0.00659 -0.0643 0.3871 1.0000
8.250 1.3482 0.01257 0.00682 -0.0637 0.3804 1.0000
8.500 1.3698 0.01281 0.00705 -0.0628 0.3734 1.0000
8.750 1.3916 0.01302 0.00732 -0.0621 0.3664 1.0000
9.000 1.4119 0.01330 0.00759 -0.0610 0.3577 1.0000
9.250 1.4327 0.01354 0.00788 -0.0601 0.3492 1.0000
9.500 1.4511 0.01390 0.00822 -0.0588 0.3398 1.0000
9.750 1.4694 0.01424 0.00858 -0.0575 0.3278 1.0000
10.000 1.4852 0.01468 0.00900 -0.0558 0.3133 1.0000
10.250 1.4962 0.01531 0.00956 -0.0534 0.2924 1.0000
10.500 1.4978 0.01630 0.01040 -0.0497 0.2607 1.0000
10.750 1.4826 0.01775 0.01164 -0.0434 0.2258 1.0000
11.000 1.4606 0.02025 0.01388 -0.0380 0.1832 1.0000
11.250 1.4370 0.02376 0.01714 -0.0342 0.1401 1.0000
11.500 1.4089 0.02825 0.02138 -0.0313 0.0953 1.0000
11.750 1.3824 0.03300 0.02596 -0.0293 0.0594 1.0000
12.000 1.3625 0.03737 0.03023 -0.0278 0.0374 1.0000
12.250 1.3546 0.04070 0.03358 -0.0269 0.0304 1.0000
12.500 1.3493 0.04382 0.03675 -0.0262 0.0273 1.0000
12.750 1.3481 0.04659 0.03961 -0.0257 0.0257 1.0000
13.000 1.3451 0.04964 0.04273 -0.0252 0.0244 1.0000
13.250 1.3399 0.05303 0.04619 -0.0249 0.0233 1.0000
13.500 1.3348 0.05652 0.04976 -0.0247 0.0224 1.0000
13.750 1.3336 0.05959 0.05292 -0.0246 0.0217 1.0000
14.000 1.3310 0.06289 0.05631 -0.0246 0.0211 1.0000
14.250 1.3273 0.06638 0.05989 -0.0246 0.0205 1.0000
14.500 1.3228 0.07003 0.06361 -0.0248 0.0200 1.0000
14.750 1.3169 0.07389 0.06756 -0.0250 0.0195 1.0000
15.000 1.3097 0.07799 0.07173 -0.0253 0.0191 1.0000
15.250 1.3000 0.08248 0.07631 -0.0257 0.0187 1.0000
15.500 1.2910 0.08691 0.08082 -0.0261 0.0184 1.0000
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