NACA M15 AIRFOIL (m15-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M15 AIRFOIL (m15-il) Reynolds number: 1,000,000 Max Cl/Cd: 124.75 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m15-il-1000000.txt Download as CSV file: xf-m15-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M15 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 -0.3987 0.08713 0.08555 -0.0324 0.9780 0.0243
-9.000 -0.3973 0.08336 0.08171 -0.0346 0.9346 0.0249
-8.000 -0.5009 0.03326 0.03005 -0.0575 0.8334 0.0273
-7.750 -0.4883 0.03046 0.02700 -0.0566 0.8182 0.0276
-7.500 -0.4761 0.02668 0.02283 -0.0550 0.8038 0.0264
-7.250 -0.4636 0.02309 0.01873 -0.0534 0.7895 0.0264
-7.000 -0.4434 0.02133 0.01662 -0.0525 0.7749 0.0266
-6.500 -0.4014 0.01767 0.01227 -0.0507 0.7474 0.0274
-6.250 -0.3770 0.01653 0.01097 -0.0503 0.7355 0.0277
-6.000 -0.3515 0.01572 0.01002 -0.0500 0.7246 0.0280
-5.750 -0.3254 0.01505 0.00922 -0.0497 0.7140 0.0282
-5.500 -0.2988 0.01442 0.00848 -0.0495 0.7048 0.0285
-5.250 -0.2721 0.01385 0.00780 -0.0493 0.6965 0.0287
-5.000 -0.2450 0.01331 0.00717 -0.0492 0.6897 0.0290
-4.750 -0.2178 0.01282 0.00661 -0.0491 0.6830 0.0293
-4.500 -0.1905 0.01237 0.00608 -0.0489 0.6770 0.0297
-4.250 -0.1630 0.01194 0.00559 -0.0488 0.6711 0.0301
-4.000 -0.1354 0.01161 0.00519 -0.0487 0.6656 0.0306
-3.750 -0.1076 0.01133 0.00486 -0.0487 0.6608 0.0311
-3.500 -0.0798 0.01102 0.00451 -0.0486 0.6559 0.0315
-3.250 -0.0520 0.01076 0.00421 -0.0486 0.6507 0.0318
-3.000 -0.0242 0.01055 0.00394 -0.0485 0.6458 0.0320
-2.750 0.0024 0.00996 0.00333 -0.0483 0.6418 0.0327
-2.500 0.0299 0.00965 0.00300 -0.0482 0.6373 0.0333
-2.250 0.0576 0.00945 0.00277 -0.0481 0.6326 0.0341
-2.000 0.0857 0.00929 0.00259 -0.0481 0.6281 0.0349
-1.750 0.1139 0.00912 0.00242 -0.0482 0.6236 0.0358
-1.500 0.1422 0.00899 0.00228 -0.0482 0.6191 0.0369
-1.250 0.1704 0.00893 0.00217 -0.0482 0.6142 0.0378
-1.000 0.1986 0.00872 0.00197 -0.0483 0.6099 0.0400
-0.750 0.2269 0.00860 0.00186 -0.0483 0.6048 0.0423
-0.500 0.2552 0.00853 0.00176 -0.0484 0.5998 0.0454
-0.250 0.2833 0.00837 0.00168 -0.0484 0.5950 0.0624
0.000 0.3111 0.00815 0.00164 -0.0485 0.5895 0.1111
0.250 0.3389 0.00802 0.00161 -0.0485 0.5836 0.1526
0.500 0.3664 0.00779 0.00161 -0.0486 0.5781 0.2319
0.750 0.3866 0.00658 0.00156 -0.0476 0.5720 0.6331
1.000 0.4017 0.00582 0.00168 -0.0445 0.5657 0.9159
1.250 0.4285 0.00586 0.00177 -0.0439 0.5585 0.9516
1.500 0.4573 0.00598 0.00185 -0.0438 0.5503 0.9707
1.750 0.4947 0.00612 0.00196 -0.0457 0.5409 0.9809
2.000 0.5431 0.00630 0.00207 -0.0502 0.5290 0.9854
2.250 0.5883 0.00647 0.00216 -0.0540 0.5156 0.9890
2.500 0.6265 0.00662 0.00223 -0.0563 0.5012 0.9925
2.750 0.6872 0.00685 0.00236 -0.0636 0.4836 0.9976
3.000 0.7281 0.00700 0.00242 -0.0666 0.4677 0.9999
3.250 0.7551 0.00712 0.00247 -0.0666 0.4543 1.0000
3.500 0.7808 0.00723 0.00253 -0.0664 0.4430 1.0000
3.750 0.8065 0.00733 0.00258 -0.0661 0.4334 1.0000
4.000 0.8319 0.00745 0.00266 -0.0657 0.4243 1.0000
4.250 0.8574 0.00753 0.00272 -0.0654 0.4177 1.0000
4.500 0.8822 0.00765 0.00281 -0.0649 0.4112 1.0000
4.750 0.9071 0.00773 0.00290 -0.0644 0.4058 1.0000
5.000 0.9317 0.00784 0.00300 -0.0639 0.3997 1.0000
5.250 0.9557 0.00798 0.00312 -0.0633 0.3932 1.0000
5.500 0.9805 0.00806 0.00321 -0.0628 0.3883 1.0000
5.750 1.0046 0.00818 0.00333 -0.0621 0.3828 1.0000
6.000 1.0283 0.00833 0.00347 -0.0615 0.3768 1.0000
6.250 1.0521 0.00846 0.00358 -0.0608 0.3663 1.0000
6.500 1.0753 0.00862 0.00371 -0.0600 0.3530 1.0000
6.750 1.0983 0.00881 0.00386 -0.0592 0.3408 1.0000
7.000 1.1209 0.00901 0.00403 -0.0584 0.3283 1.0000
7.250 1.1432 0.00924 0.00421 -0.0575 0.3146 1.0000
7.500 1.1650 0.00951 0.00442 -0.0565 0.2972 1.0000
7.750 1.1857 0.00985 0.00467 -0.0553 0.2756 1.0000
8.000 1.2052 0.01028 0.00498 -0.0540 0.2513 1.0000
8.250 1.2231 0.01081 0.00536 -0.0525 0.2230 1.0000
8.500 1.2398 0.01141 0.00582 -0.0507 0.1938 1.0000
8.750 1.2562 0.01204 0.00630 -0.0490 0.1682 1.0000
9.000 1.2728 0.01267 0.00680 -0.0473 0.1448 1.0000
9.250 1.2873 0.01344 0.00740 -0.0454 0.1159 1.0000
9.500 1.2959 0.01456 0.00826 -0.0426 0.0753 1.0000
9.750 1.3064 0.01552 0.00907 -0.0402 0.0526 1.0000
10.000 1.3185 0.01630 0.00978 -0.0380 0.0406 1.0000
10.250 1.3291 0.01699 0.01045 -0.0354 0.0332 1.0000
10.500 1.3390 0.01773 0.01119 -0.0329 0.0282 1.0000
10.750 1.3506 0.01848 0.01194 -0.0308 0.0251 1.0000
11.000 1.3624 0.01929 0.01277 -0.0290 0.0228 1.0000
11.250 1.3745 0.02014 0.01365 -0.0274 0.0212 1.0000
11.500 1.3830 0.02129 0.01483 -0.0255 0.0196 1.0000
11.750 1.3959 0.02221 0.01581 -0.0243 0.0188 1.0000
12.000 1.4073 0.02327 0.01693 -0.0231 0.0180 1.0000
12.250 1.4173 0.02451 0.01821 -0.0220 0.0173 1.0000
12.500 1.4242 0.02607 0.01981 -0.0208 0.0165 1.0000
12.750 1.4289 0.02788 0.02170 -0.0196 0.0159 1.0000
13.000 1.4389 0.02930 0.02320 -0.0189 0.0155 1.0000
13.250 1.4473 0.03090 0.02485 -0.0182 0.0150 1.0000
13.500 1.4543 0.03267 0.02669 -0.0175 0.0145 1.0000
13.750 1.4601 0.03461 0.02870 -0.0170 0.0142 1.0000
14.000 1.4633 0.03683 0.03098 -0.0164 0.0138 1.0000
14.250 1.4631 0.03945 0.03367 -0.0158 0.0134 1.0000
14.500 1.4578 0.04265 0.03696 -0.0153 0.0131 1.0000
14.750 1.4542 0.04572 0.04013 -0.0149 0.0129 1.0000
15.000 1.4564 0.04826 0.04275 -0.0148 0.0127 1.0000
15.250 1.4574 0.05099 0.04557 -0.0147 0.0125 1.0000
15.500 1.4571 0.05390 0.04856 -0.0147 0.0123 1.0000
15.750 1.4555 0.05706 0.05180 -0.0148 0.0121 1.0000
16.000 1.4534 0.06036 0.05518 -0.0151 0.0119 1.0000
16.250 1.4509 0.06375 0.05866 -0.0154 0.0117 1.0000
16.500 1.4475 0.06734 0.06233 -0.0159 0.0115 1.0000
16.750 1.4433 0.07109 0.06617 -0.0165 0.0114 1.0000
17.000 1.4391 0.07495 0.07010 -0.0172 0.0112 1.0000
17.250 1.4339 0.07895 0.07418 -0.0180 0.0110 1.0000
17.500 1.4273 0.08319 0.07851 -0.0189 0.0109 1.0000
17.750 1.4187 0.08778 0.08318 -0.0200 0.0107 1.0000
18.000 1.4082 0.09270 0.08818 -0.0212 0.0106 1.0000
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