NACA M25 AIRFOIL (m25-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M25 AIRFOIL (m25-il) Reynolds number: 500,000 Max Cl/Cd: 94.51 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m25-il-500000-n5.txt Download as CSV file: xf-m25-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M25 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 -0.3242 0.10983 0.10648 0.0194 0.5697 0.0076
-7.750 -0.3161 0.10683 0.10349 0.0175 0.5663 0.0076
-7.500 -0.3084 0.10388 0.10053 0.0154 0.5628 0.0077
-7.250 -0.2989 0.10087 0.09751 0.0129 0.5595 0.0077
-7.000 -0.2863 0.09763 0.09424 0.0100 0.5565 0.0077
-6.750 -0.2723 0.09432 0.09092 0.0070 0.5533 0.0077
-6.500 -0.2569 0.09095 0.08752 0.0039 0.5501 0.0077
-6.250 -0.2402 0.08754 0.08408 0.0007 0.5470 0.0077
-6.000 -0.2224 0.08410 0.08059 -0.0025 0.5439 0.0077
-5.750 -0.2035 0.08060 0.07705 -0.0056 0.5410 0.0077
-5.500 -0.1849 0.07707 0.07350 -0.0084 0.5377 0.0076
-5.250 -0.1670 0.07397 0.07036 -0.0107 0.5344 0.0073
-5.000 -0.1441 0.07061 0.06694 -0.0140 0.5314 0.0071
-4.500 -0.0881 0.06491 0.06111 -0.0212 0.5254 0.0108
-4.250 -0.0683 0.06168 0.05782 -0.0231 0.5221 0.0102
-4.000 -0.0412 0.05821 0.05428 -0.0262 0.5190 0.0088
-3.750 -0.0117 0.05460 0.05056 -0.0293 0.5161 0.0084
-3.500 0.0172 0.05161 0.04746 -0.0319 0.5133 0.0086
-3.250 0.0466 0.04880 0.04457 -0.0342 0.5103 0.0090
-3.000 0.0777 0.04578 0.04144 -0.0365 0.5071 0.0103
-2.750 0.1115 0.04211 0.03762 -0.0390 0.5042 0.0111
-2.500 0.1381 0.04057 0.03598 -0.0401 0.5011 0.0116
-2.250 0.1672 0.03860 0.03389 -0.0415 0.4981 0.0123
-2.000 0.1988 0.03625 0.03142 -0.0428 0.4951 0.0143
-1.750 0.2327 0.03335 0.02833 -0.0441 0.4921 0.0155
-1.500 0.2589 0.03225 0.02716 -0.0448 0.4889 0.0163
-1.250 0.2876 0.03089 0.02567 -0.0456 0.4858 0.0178
-1.000 0.3191 0.02907 0.02370 -0.0462 0.4829 0.0195
-0.750 0.3545 0.02685 0.02124 -0.0467 0.4800 0.0215
-0.500 0.3783 0.02605 0.02042 -0.0473 0.4767 0.0228
-0.250 0.4107 0.02596 0.02021 -0.0476 0.4736 0.0281
0.250 0.4699 0.02307 0.01700 -0.0481 0.4678 0.0284
0.500 0.4991 0.02178 0.01558 -0.0482 0.4645 0.0284
0.750 0.5285 0.02077 0.01443 -0.0483 0.4613 0.0286
1.000 0.5572 0.01956 0.01307 -0.0485 0.4583 0.0286
1.250 0.5846 0.01781 0.01113 -0.0488 0.4556 0.0271
1.500 0.6138 0.01663 0.00980 -0.0487 0.4527 0.0267
1.750 0.6431 0.01553 0.00853 -0.0487 0.4495 0.0273
2.000 0.6720 0.01485 0.00770 -0.0487 0.4460 0.0283
2.500 0.7284 0.01340 0.00597 -0.0490 0.4399 0.0312
2.750 0.7564 0.01314 0.00571 -0.0493 0.4369 0.0328
3.000 0.7849 0.01247 0.00486 -0.0492 0.4337 0.0301
3.250 0.8130 0.01232 0.00466 -0.0494 0.4303 0.0294
3.500 0.8408 0.01225 0.00455 -0.0496 0.4268 0.0290
3.750 0.8684 0.01197 0.00427 -0.0498 0.4236 0.0285
4.000 0.8956 0.01171 0.00400 -0.0498 0.4202 0.0283
4.250 0.9227 0.01155 0.00384 -0.0499 0.4168 0.0282
4.500 0.9498 0.01147 0.00376 -0.0500 0.4136 0.0282
4.750 0.9770 0.01143 0.00373 -0.0502 0.4106 0.0283
5.000 1.0045 0.01140 0.00375 -0.0504 0.4074 0.0286
5.250 1.0318 0.01139 0.00377 -0.0506 0.4036 0.0293
5.500 1.0590 0.01144 0.00381 -0.0508 0.3949 0.0320
5.750 1.0862 0.01158 0.00391 -0.0512 0.3801 0.0347
6.000 1.1133 0.01178 0.00408 -0.0515 0.3651 0.0368
6.250 1.1394 0.01217 0.00437 -0.0519 0.3376 0.0451
6.500 1.1638 0.01294 0.00494 -0.0524 0.2895 0.0759
7.000 1.2095 0.01734 0.00946 -0.0577 0.0167 0.9953
7.250 1.2456 0.01807 0.01028 -0.0607 0.0132 1.0000
7.500 1.2643 0.01889 0.01119 -0.0601 0.0108 1.0000
7.750 1.2806 0.01989 0.01231 -0.0594 0.0097 1.0000
8.000 1.2966 0.02079 0.01329 -0.0586 0.0090 1.0000
8.250 1.3105 0.02182 0.01438 -0.0578 0.0080 1.0000
8.500 1.3209 0.02329 0.01592 -0.0575 0.0073 1.0000
8.750 1.3214 0.02527 0.01800 -0.0563 0.0070 1.0000
9.000 1.3203 0.02770 0.02053 -0.0553 0.0066 1.0000
9.250 1.3139 0.03099 0.02394 -0.0546 0.0063 1.0000
9.500 1.3122 0.03387 0.02692 -0.0542 0.0062 1.0000
9.750 1.3107 0.03675 0.02990 -0.0537 0.0060 1.0000
10.000 1.3071 0.03990 0.03315 -0.0533 0.0059 1.0000
10.250 1.3023 0.04318 0.03653 -0.0528 0.0057 1.0000
10.500 1.2969 0.04655 0.03999 -0.0524 0.0056 1.0000
10.750 1.2924 0.04980 0.04334 -0.0519 0.0054 1.0000
11.000 1.2893 0.05295 0.04656 -0.0516 0.0051 1.0000
11.250 1.2886 0.05592 0.04960 -0.0513 0.0049 1.0000
11.500 1.2895 0.05877 0.05251 -0.0512 0.0046 1.0000
11.750 1.2908 0.06159 0.05538 -0.0511 0.0044 1.0000
12.000 1.2896 0.06472 0.05859 -0.0510 0.0042 1.0000
12.250 1.2869 0.06801 0.06194 -0.0508 0.0041 1.0000
12.500 1.2822 0.07136 0.06534 -0.0503 0.0040 1.0000
12.750 1.2842 0.07391 0.06796 -0.0497 0.0039 1.0000
13.000 1.2879 0.07615 0.07028 -0.0490 0.0038 1.0000
13.250 1.2935 0.07791 0.07210 -0.0478 0.0036 1.0000
13.500 1.3024 0.07892 0.07319 -0.0458 0.0034 1.0000
13.750 1.3162 0.07901 0.07334 -0.0429 0.0032 1.0000
14.000 1.3326 0.07887 0.07330 -0.0396 0.0031 1.0000
14.250 1.3493 0.07911 0.07366 -0.0363 0.0030 1.0000
14.500 1.3572 0.08103 0.07574 -0.0351 0.0029 1.0000
14.750 1.3601 0.08374 0.07855 -0.0350 0.0028 1.0000
15.000 1.3613 0.08672 0.08165 -0.0353 0.0027 1.0000
15.250 1.3606 0.09002 0.08506 -0.0358 0.0026 1.0000
15.500 1.3591 0.09349 0.08865 -0.0364 0.0025 1.0000
15.750 1.3556 0.09728 0.09257 -0.0372 0.0024 1.0000
16.000 1.3495 0.10155 0.09699 -0.0379 0.0024 1.0000
16.250 1.3442 0.10589 0.10154 -0.0377 0.0023 1.0000
16.500 1.3377 0.11062 0.10648 -0.0381 0.0023 1.0000
16.750 1.3286 0.11586 0.11192 -0.0392 0.0023 1.0000
17.000 1.3184 0.12146 0.11772 -0.0407 0.0022 1.0000
17.250 1.3069 0.12743 0.12387 -0.0427 0.0022 1.0000
17.500 1.2943 0.13381 0.13043 -0.0453 0.0022 1.0000
17.750 1.2814 0.14047 0.13726 -0.0484 0.0022 1.0000
18.000 1.2683 0.14736 0.14431 -0.0518 0.0022 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M25 AIRFOIL (m25-il)