NACA M21 AIRFOIL (m21-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M21 AIRFOIL (m21-il) Reynolds number: 1,000,000 Max Cl/Cd: 128.44 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m21-il-1000000-n5.txt Download as CSV file: xf-m21-il-1000000-n5.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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3764 0.08367 0.08075 -0.0244 0.6094 0.0180
-8.750 -0.3733 0.08074 0.07781 -0.0270 0.6047 0.0182
-7.750 -0.4699 0.01926 0.01418 -0.0481 0.6002 0.0249
-7.500 -0.4467 0.01756 0.01221 -0.0477 0.5959 0.0251
-7.250 -0.4214 0.01653 0.01099 -0.0474 0.5911 0.0254
-7.000 -0.3952 0.01576 0.01004 -0.0472 0.5859 0.0256
-6.750 -0.3685 0.01513 0.00925 -0.0471 0.5812 0.0257
-6.500 -0.3413 0.01459 0.00859 -0.0469 0.5764 0.0258
-6.250 -0.3145 0.01393 0.00780 -0.0468 0.5717 0.0261
-5.750 -0.2597 0.01296 0.00664 -0.0466 0.5626 0.0270
-5.500 -0.2317 0.01266 0.00628 -0.0466 0.5577 0.0273
-5.250 -0.2036 0.01237 0.00591 -0.0466 0.5531 0.0277
-5.000 -0.1755 0.01208 0.00556 -0.0466 0.5490 0.0280
-4.750 -0.1472 0.01180 0.00522 -0.0466 0.5445 0.0283
-4.500 -0.1189 0.01155 0.00490 -0.0467 0.5397 0.0286
-4.250 -0.0907 0.01131 0.00459 -0.0467 0.5353 0.0290
-4.000 -0.0623 0.01108 0.00431 -0.0467 0.5315 0.0294
-3.750 -0.0339 0.01086 0.00404 -0.0468 0.5270 0.0298
-3.500 -0.0056 0.01067 0.00380 -0.0468 0.5224 0.0301
-3.250 0.0228 0.01051 0.00358 -0.0469 0.5183 0.0305
-3.000 0.0513 0.01035 0.00338 -0.0469 0.5145 0.0307
-2.750 0.0795 0.01013 0.00313 -0.0469 0.5099 0.0316
-2.500 0.1078 0.00999 0.00296 -0.0470 0.5055 0.0324
-2.250 0.1362 0.00988 0.00283 -0.0471 0.5017 0.0332
-2.000 0.1647 0.00977 0.00271 -0.0472 0.4978 0.0341
-1.750 0.1932 0.00968 0.00259 -0.0473 0.4933 0.0351
-1.500 0.2216 0.00960 0.00248 -0.0473 0.4890 0.0359
-1.250 0.2499 0.00951 0.00237 -0.0474 0.4854 0.0370
-1.000 0.2783 0.00942 0.00229 -0.0475 0.4814 0.0388
-0.750 0.3068 0.00936 0.00223 -0.0476 0.4770 0.0405
-0.500 0.3352 0.00933 0.00217 -0.0478 0.4727 0.0421
-0.250 0.3636 0.00928 0.00212 -0.0479 0.4693 0.0442
0.000 0.3922 0.00924 0.00210 -0.0480 0.4655 0.0467
0.250 0.4207 0.00922 0.00207 -0.0482 0.4610 0.0490
0.500 0.4490 0.00922 0.00205 -0.0484 0.4565 0.0510
0.750 0.4775 0.00921 0.00206 -0.0485 0.4532 0.0540
1.000 0.5061 0.00920 0.00206 -0.0487 0.4494 0.0564
1.250 0.5346 0.00921 0.00206 -0.0489 0.4450 0.0582
1.500 0.5628 0.00923 0.00206 -0.0491 0.4404 0.0598
1.750 0.5912 0.00922 0.00208 -0.0493 0.4369 0.0627
2.000 0.6197 0.00924 0.00211 -0.0495 0.4332 0.0657
2.250 0.6481 0.00927 0.00214 -0.0497 0.4289 0.0681
2.500 0.6763 0.00932 0.00218 -0.0499 0.4244 0.0694
2.750 0.7045 0.00932 0.00219 -0.0501 0.4207 0.0723
3.000 0.7327 0.00934 0.00223 -0.0503 0.4165 0.0750
3.250 0.7609 0.00938 0.00228 -0.0505 0.4124 0.0775
3.500 0.7888 0.00945 0.00234 -0.0507 0.4083 0.0799
3.750 0.8171 0.00948 0.00239 -0.0509 0.4048 0.0819
4.000 0.8451 0.00952 0.00245 -0.0511 0.4002 0.0857
4.250 0.8729 0.00959 0.00253 -0.0513 0.3956 0.0897
4.500 0.9007 0.00967 0.00261 -0.0515 0.3919 0.0936
4.750 0.9287 0.00971 0.00270 -0.0517 0.3884 0.0998
5.250 0.9813 0.00950 0.00301 -0.0517 0.3797 0.4066
5.500 1.0017 0.00839 0.00327 -0.0503 0.3759 0.9712
5.750 1.0356 0.00854 0.00344 -0.0519 0.3720 0.9827
6.000 1.0780 0.00874 0.00363 -0.0554 0.3675 0.9875
6.250 1.1136 0.00895 0.00381 -0.0575 0.3625 0.9906
6.500 1.1469 0.00912 0.00399 -0.0590 0.3564 0.9936
6.750 1.1810 0.00934 0.00417 -0.0608 0.3488 0.9946
7.000 1.2169 0.00955 0.00436 -0.0630 0.3398 0.9959
7.250 1.2544 0.00982 0.00459 -0.0657 0.3309 0.9978
7.500 1.2908 0.01005 0.00481 -0.0681 0.3212 0.9993
7.750 1.3212 0.01032 0.00505 -0.0692 0.3122 1.0000
8.000 1.3429 0.01064 0.00532 -0.0684 0.3006 1.0000
8.250 1.3637 0.01100 0.00562 -0.0675 0.2859 1.0000
8.500 1.3841 0.01137 0.00594 -0.0665 0.2720 1.0000
8.750 1.4028 0.01183 0.00632 -0.0653 0.2557 1.0000
9.000 1.4182 0.01245 0.00682 -0.0636 0.2314 1.0000
9.250 1.4230 0.01363 0.00773 -0.0603 0.1881 1.0000
9.500 1.4207 0.01497 0.00881 -0.0558 0.1468 1.0000
9.750 1.3937 0.01688 0.01042 -0.0472 0.0957 1.0000
10.000 1.3655 0.01908 0.01244 -0.0398 0.0594 1.0000
10.250 1.3490 0.02190 0.01510 -0.0361 0.0253 1.0000
10.500 1.3545 0.02352 0.01672 -0.0349 0.0189 1.0000
10.750 1.3635 0.02494 0.01818 -0.0340 0.0167 1.0000
11.000 1.3721 0.02647 0.01974 -0.0332 0.0153 1.0000
11.250 1.3804 0.02808 0.02139 -0.0325 0.0141 1.0000
11.500 1.3890 0.02968 0.02304 -0.0319 0.0135 1.0000
11.750 1.3964 0.03143 0.02484 -0.0314 0.0128 1.0000
12.000 1.4022 0.03335 0.02681 -0.0308 0.0121 1.0000
12.250 1.4064 0.03543 0.02894 -0.0302 0.0115 1.0000
12.500 1.4097 0.03765 0.03121 -0.0296 0.0110 1.0000
12.750 1.4145 0.03972 0.03334 -0.0291 0.0107 1.0000
13.000 1.4180 0.04192 0.03560 -0.0286 0.0104 1.0000
13.250 1.4209 0.04423 0.03797 -0.0282 0.0101 1.0000
13.500 1.4233 0.04667 0.04048 -0.0279 0.0098 1.0000
13.750 1.4249 0.04925 0.04311 -0.0276 0.0095 1.0000
14.000 1.4259 0.05193 0.04585 -0.0274 0.0093 1.0000
14.250 1.4263 0.05473 0.04871 -0.0272 0.0090 1.0000
14.500 1.4257 0.05767 0.05171 -0.0271 0.0087 1.0000
14.750 1.4232 0.06089 0.05499 -0.0270 0.0084 1.0000
15.000 1.4232 0.06386 0.05803 -0.0270 0.0083 1.0000
15.250 1.4240 0.06676 0.06100 -0.0271 0.0081 1.0000
15.500 1.4237 0.06981 0.06411 -0.0272 0.0079 1.0000
15.750 1.4228 0.07296 0.06733 -0.0274 0.0077 1.0000
16.000 1.4223 0.07608 0.07052 -0.0276 0.0076 1.0000
16.250 1.4204 0.07945 0.07395 -0.0279 0.0074 1.0000
16.500 1.4188 0.08278 0.07735 -0.0282 0.0072 1.0000
16.750 1.4170 0.08614 0.08078 -0.0286 0.0071 1.0000
17.000 1.4141 0.08971 0.08441 -0.0290 0.0070 1.0000
17.250 1.4122 0.09316 0.08792 -0.0295 0.0068 1.0000
17.500 1.4086 0.09690 0.09173 -0.0301 0.0067 1.0000
17.750 1.4052 0.10059 0.09549 -0.0308 0.0065 1.0000
18.000 1.4007 0.10448 0.09944 -0.0315 0.0064 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M21 AIRFOIL (m21-il)