NACA M11 AIRFOIL (m11-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M11 AIRFOIL (m11-il) Reynolds number: 100,000 Max Cl/Cd: 48.6 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m11-il-100000-n5.txt Download as CSV file: xf-m11-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M11 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.4104 0.10997 0.10507 -0.0105 1.0000 0.0578
-10.250 -0.4101 0.10620 0.10132 -0.0118 1.0000 0.0593
-9.750 -0.5053 0.10556 0.10038 -0.0108 1.0000 0.0556
-9.500 -0.5028 0.10196 0.09681 -0.0120 1.0000 0.0568
-9.250 -0.5022 0.09813 0.09302 -0.0137 1.0000 0.0573
-8.750 -0.5228 0.08471 0.07970 -0.0238 1.0000 0.0384
-8.500 -0.5281 0.08040 0.07543 -0.0258 1.0000 0.0380
-8.250 -0.5336 0.07618 0.07122 -0.0272 1.0000 0.0378
-8.000 -0.5384 0.07125 0.06624 -0.0294 1.0000 0.0380
-7.750 -0.5418 0.06615 0.06104 -0.0309 1.0000 0.0384
-7.500 -0.5432 0.06116 0.05587 -0.0315 1.0000 0.0388
-7.250 -0.5427 0.05644 0.05090 -0.0310 1.0000 0.0392
-7.000 -0.5397 0.05220 0.04643 -0.0299 1.0000 0.0392
-6.750 -0.5358 0.04810 0.04201 -0.0282 1.0000 0.0393
-6.500 -0.5325 0.04403 0.03770 -0.0263 1.0000 0.0402
-6.250 -0.5202 0.04256 0.03629 -0.0250 1.0000 0.0422
-6.000 -0.5090 0.04040 0.03396 -0.0232 1.0000 0.0438
-5.750 -0.4992 0.03737 0.03060 -0.0209 1.0000 0.0444
-5.500 -0.4884 0.03453 0.02739 -0.0185 1.0000 0.0447
-5.250 -0.4759 0.03219 0.02470 -0.0161 1.0000 0.0457
-5.000 -0.4621 0.03029 0.02242 -0.0138 1.0000 0.0479
-4.750 -0.4475 0.02838 0.02012 -0.0115 1.0000 0.0490
-4.500 -0.4316 0.02662 0.01801 -0.0093 1.0000 0.0493
-4.250 -0.4148 0.02511 0.01618 -0.0073 1.0000 0.0498
-4.000 -0.3881 0.02366 0.01440 -0.0072 0.9968 0.0504
-3.750 -0.3521 0.02234 0.01275 -0.0090 0.9896 0.0523
-3.500 -0.3164 0.02101 0.01135 -0.0109 0.9826 0.0544
-3.250 -0.2796 0.01991 0.01013 -0.0128 0.9757 0.0552
-3.000 -0.2447 0.01897 0.00912 -0.0143 0.9673 0.0562
-2.750 -0.2089 0.01815 0.00824 -0.0160 0.9594 0.0575
-2.500 -0.1731 0.01743 0.00747 -0.0176 0.9513 0.0591
-2.250 -0.1395 0.01683 0.00680 -0.0187 0.9417 0.0615
-2.000 -0.1041 0.01632 0.00621 -0.0202 0.9331 0.0642
-1.750 -0.0717 0.01581 0.00572 -0.0211 0.9223 0.0695
-1.500 -0.0401 0.01541 0.00534 -0.0218 0.9108 0.0809
-1.250 -0.0085 0.01487 0.00497 -0.0224 0.8994 0.1134
-1.000 0.1042 0.01218 0.00520 -0.0390 0.9112 0.9618
-0.750 0.1746 0.01217 0.00498 -0.0477 0.9065 1.0000
-0.500 0.2069 0.01210 0.00476 -0.0485 0.8881 1.0000
-0.250 0.2366 0.01205 0.00458 -0.0487 0.8691 1.0000
0.000 0.2646 0.01201 0.00442 -0.0486 0.8505 1.0000
0.250 0.2907 0.01200 0.00430 -0.0480 0.8321 1.0000
0.500 0.3157 0.01203 0.00422 -0.0472 0.8133 1.0000
0.750 0.3405 0.01206 0.00417 -0.0464 0.7953 1.0000
1.000 0.3650 0.01212 0.00415 -0.0455 0.7780 1.0000
1.250 0.3892 0.01219 0.00415 -0.0446 0.7613 1.0000
1.750 0.4367 0.01239 0.00425 -0.0425 0.7283 1.0000
2.000 0.4602 0.01250 0.00433 -0.0415 0.7117 1.0000
2.250 0.4837 0.01263 0.00444 -0.0405 0.6953 1.0000
2.500 0.5071 0.01277 0.00457 -0.0394 0.6788 1.0000
2.750 0.5306 0.01292 0.00471 -0.0383 0.6621 1.0000
3.000 0.5539 0.01307 0.00487 -0.0372 0.6452 1.0000
3.250 0.5772 0.01324 0.00504 -0.0361 0.6278 1.0000
3.500 0.6000 0.01341 0.00523 -0.0349 0.6071 1.0000
3.750 0.6224 0.01359 0.00538 -0.0336 0.5843 1.0000
4.000 0.6444 0.01379 0.00556 -0.0322 0.5579 1.0000
4.250 0.6657 0.01402 0.00574 -0.0307 0.5266 1.0000
4.500 0.6865 0.01429 0.00594 -0.0291 0.4905 1.0000
4.750 0.7070 0.01461 0.00616 -0.0275 0.4516 1.0000
5.000 0.7275 0.01497 0.00646 -0.0261 0.4111 1.0000
5.250 0.7470 0.01544 0.00681 -0.0245 0.3642 1.0000
5.500 0.7655 0.01604 0.00722 -0.0228 0.3092 1.0000
5.750 0.7827 0.01683 0.00775 -0.0211 0.2439 1.0000
6.000 0.7982 0.01787 0.00844 -0.0193 0.1763 1.0000
6.250 0.8135 0.01901 0.00932 -0.0175 0.1290 1.0000
6.500 0.8296 0.02007 0.01024 -0.0157 0.1037 1.0000
6.750 0.8462 0.02104 0.01120 -0.0139 0.0875 1.0000
7.000 0.8625 0.02202 0.01218 -0.0122 0.0741 1.0000
7.250 0.8777 0.02309 0.01322 -0.0104 0.0638 1.0000
7.500 0.8940 0.02406 0.01433 -0.0086 0.0557 1.0000
7.750 0.9082 0.02527 0.01561 -0.0066 0.0488 1.0000
8.000 0.9234 0.02642 0.01685 -0.0047 0.0435 1.0000
8.250 0.9362 0.02793 0.01837 -0.0026 0.0396 1.0000
8.500 0.9525 0.02915 0.01978 -0.0009 0.0354 1.0000
8.750 0.9671 0.03047 0.02116 0.0008 0.0326 1.0000
9.000 0.9810 0.03239 0.02311 0.0025 0.0308 1.0000
9.250 0.9972 0.03422 0.02524 0.0041 0.0289 1.0000
9.500 1.0108 0.03591 0.02723 0.0059 0.0266 1.0000
9.750 1.0224 0.03756 0.02905 0.0076 0.0249 1.0000
10.000 1.0326 0.03954 0.03119 0.0094 0.0239 1.0000
10.250 1.0404 0.04199 0.03380 0.0113 0.0232 1.0000
10.500 1.0440 0.04493 0.03702 0.0135 0.0227 1.0000
10.750 1.0419 0.04774 0.04021 0.0162 0.0223 1.0000
11.000 1.0333 0.05058 0.04338 0.0194 0.0220 1.0000
11.250 1.0223 0.05358 0.04668 0.0219 0.0218 1.0000
11.500 1.0077 0.05702 0.05041 0.0235 0.0214 1.0000
11.750 0.9917 0.06096 0.05461 0.0239 0.0213 1.0000
12.000 0.9735 0.06550 0.05940 0.0231 0.0211 1.0000
12.250 0.9547 0.07075 0.06485 0.0212 0.0213 1.0000
12.500 0.9349 0.07675 0.07103 0.0182 0.0215 1.0000
12.750 0.9115 0.08414 0.07858 0.0136 0.0216 1.0000
13.000 0.8910 0.09188 0.08643 0.0086 0.0219 1.0000
13.250 0.8705 0.10044 0.09506 0.0033 0.0223 1.0000
13.500 0.8470 0.11083 0.10550 -0.0031 0.0226 1.0000
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Polar data table (+)
Polar graphs
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