DAE-11 AIRFOIL (dae11-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: DAE-11 AIRFOIL (dae11-il) Reynolds number: 500,000 Max Cl/Cd: 135.81 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-dae11-il-500000.txt Download as CSV file: xf-dae11-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: DAE-11 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.0484 0.08616 0.08260 -0.0786 0.7081 0.0218
-8.250 -0.0452 0.08299 0.07945 -0.0808 0.7064 0.0218
-8.000 -0.0449 0.07910 0.07562 -0.0830 0.7048 0.0220
-7.750 -0.0324 0.07700 0.07353 -0.0824 0.7031 0.0224
-7.500 -0.0221 0.07490 0.07144 -0.0828 0.7012 0.0227
-7.250 -0.0157 0.07250 0.06906 -0.0837 0.6993 0.0229
-7.000 -0.0108 0.07019 0.06677 -0.0844 0.6974 0.0234
-6.750 -0.0078 0.06784 0.06444 -0.0856 0.6957 0.0239
-6.500 -0.0001 0.06486 0.06144 -0.0885 0.6939 0.0245
-6.250 0.0104 0.05659 0.05311 -0.1029 0.6920 0.0263
-6.000 0.0205 0.05252 0.04903 -0.1056 0.6902 0.0268
-5.750 0.0371 0.05040 0.04691 -0.1070 0.6885 0.0271
-5.500 0.0559 0.04789 0.04437 -0.1092 0.6866 0.0277
-5.250 0.0768 0.04460 0.04101 -0.1126 0.6846 0.0284
-5.000 0.1034 0.03416 0.02991 -0.1218 0.6828 0.0321
-4.750 0.1258 0.03247 0.02823 -0.1225 0.6809 0.0326
-4.500 0.1500 0.03101 0.02670 -0.1233 0.6792 0.0334
-4.250 0.1793 0.02887 0.02409 -0.1245 0.6776 0.0376
-4.000 0.2041 0.02022 0.01429 -0.1257 0.6762 0.0266
-3.750 0.2315 0.01869 0.01257 -0.1259 0.6744 0.0261
-3.500 0.2590 0.01685 0.01050 -0.1261 0.6724 0.0256
-3.250 0.2872 0.01566 0.00910 -0.1262 0.6703 0.0256
-3.000 0.3156 0.01486 0.00813 -0.1263 0.6683 0.0260
-2.750 0.3439 0.01406 0.00720 -0.1263 0.6663 0.0267
-2.500 0.3719 0.01336 0.00648 -0.1264 0.6644 0.0277
-2.250 0.4004 0.01299 0.00606 -0.1265 0.6626 0.0291
-2.000 0.4291 0.01277 0.00578 -0.1266 0.6609 0.0312
-1.750 0.4572 0.01235 0.00534 -0.1267 0.6589 0.0333
-1.500 0.4856 0.01208 0.00511 -0.1269 0.6567 0.0360
-1.250 0.5141 0.01178 0.00480 -0.1270 0.6543 0.0399
-1.000 0.5428 0.01159 0.00462 -0.1272 0.6520 0.0465
-0.750 0.5716 0.01131 0.00437 -0.1274 0.6498 0.0611
-0.500 0.6002 0.01095 0.00419 -0.1276 0.6478 0.1230
-0.250 0.6283 0.01053 0.00414 -0.1280 0.6458 0.2591
0.000 0.6560 0.01023 0.00422 -0.1282 0.6437 0.4081
0.250 0.6826 0.00988 0.00431 -0.1282 0.6410 0.5583
0.500 0.7052 0.00930 0.00439 -0.1271 0.6383 0.7769
0.750 0.7410 0.00893 0.00429 -0.1285 0.6357 1.0000
1.000 0.7698 0.00895 0.00423 -0.1287 0.6333 1.0000
1.250 0.7988 0.00899 0.00417 -0.1289 0.6310 1.0000
1.500 0.8275 0.00910 0.00419 -0.1291 0.6287 1.0000
1.750 0.8555 0.00915 0.00424 -0.1292 0.6258 1.0000
2.000 0.8838 0.00920 0.00428 -0.1293 0.6227 1.0000
2.250 0.9123 0.00922 0.00426 -0.1295 0.6197 1.0000
2.500 0.9410 0.00924 0.00423 -0.1297 0.6170 1.0000
2.750 0.9697 0.00930 0.00421 -0.1299 0.6144 1.0000
3.000 0.9975 0.00938 0.00430 -0.1300 0.6111 1.0000
3.250 1.0253 0.00942 0.00435 -0.1301 0.6074 1.0000
3.500 1.0536 0.00944 0.00436 -0.1303 0.6040 1.0000
3.750 1.0820 0.00947 0.00435 -0.1304 0.6009 1.0000
4.000 1.1100 0.00955 0.00441 -0.1306 0.5975 1.0000
4.250 1.1373 0.00961 0.00451 -0.1306 0.5931 1.0000
4.500 1.1651 0.00964 0.00454 -0.1307 0.5890 1.0000
4.750 1.1932 0.00968 0.00454 -0.1309 0.5854 1.0000
5.000 1.2203 0.00977 0.00466 -0.1309 0.5809 1.0000
5.250 1.2474 0.00983 0.00477 -0.1309 0.5759 1.0000
5.500 1.2749 0.00989 0.00480 -0.1310 0.5714 1.0000
5.750 1.3017 0.00999 0.00492 -0.1310 0.5662 1.0000
6.000 1.3282 0.01007 0.00505 -0.1310 0.5602 1.0000
6.250 1.3551 0.01017 0.00510 -0.1309 0.5548 1.0000
6.500 1.3808 0.01028 0.00530 -0.1308 0.5478 1.0000
6.750 1.4068 0.01041 0.00541 -0.1307 0.5413 1.0000
7.000 1.4321 0.01056 0.00562 -0.1305 0.5337 1.0000
7.250 1.4572 0.01073 0.00577 -0.1302 0.5262 1.0000
7.500 1.4816 0.01092 0.00602 -0.1299 0.5174 1.0000
7.750 1.5054 0.01115 0.00623 -0.1295 0.5087 1.0000
8.000 1.5284 0.01139 0.00650 -0.1289 0.4985 1.0000
8.250 1.5507 0.01167 0.00681 -0.1283 0.4878 1.0000
8.500 1.5715 0.01200 0.00714 -0.1274 0.4763 1.0000
8.750 1.5905 0.01239 0.00751 -0.1263 0.4637 1.0000
9.000 1.6074 0.01285 0.00796 -0.1249 0.4499 1.0000
9.250 1.6214 0.01339 0.00848 -0.1230 0.4349 1.0000
9.500 1.6304 0.01402 0.00908 -0.1203 0.4197 1.0000
9.750 1.6319 0.01482 0.00986 -0.1165 0.4050 1.0000
10.000 1.6318 0.01592 0.01093 -0.1131 0.3897 1.0000
10.250 1.6322 0.01724 0.01222 -0.1102 0.3742 1.0000
10.500 1.6330 0.01872 0.01367 -0.1076 0.3588 1.0000
10.750 1.6336 0.02033 0.01524 -0.1052 0.3431 1.0000
11.000 1.6338 0.02204 0.01692 -0.1030 0.3276 1.0000
11.250 1.6330 0.02388 0.01872 -0.1007 0.3113 1.0000
11.500 1.6318 0.02582 0.02061 -0.0986 0.2953 1.0000
11.750 1.6302 0.02784 0.02259 -0.0965 0.2795 1.0000
12.000 1.6277 0.02998 0.02468 -0.0944 0.2631 1.0000
12.250 1.6259 0.03212 0.02678 -0.0925 0.2479 1.0000
12.500 1.6237 0.03436 0.02898 -0.0907 0.2324 1.0000
12.750 1.6219 0.03670 0.03126 -0.0891 0.2167 1.0000
13.000 1.6205 0.03908 0.03360 -0.0877 0.2020 1.0000
13.250 1.6190 0.04156 0.03603 -0.0863 0.1872 1.0000
13.500 1.6177 0.04408 0.03850 -0.0851 0.1727 1.0000
13.750 1.6163 0.04670 0.04109 -0.0840 0.1589 1.0000
14.000 1.6151 0.04938 0.04372 -0.0830 0.1453 1.0000
14.250 1.6137 0.05213 0.04643 -0.0821 0.1322 1.0000
14.500 1.6125 0.05493 0.04921 -0.0813 0.1202 1.0000
14.750 1.6115 0.05778 0.05203 -0.0806 0.1084 1.0000
15.000 1.6100 0.06077 0.05500 -0.0800 0.0978 1.0000
15.250 1.6082 0.06384 0.05805 -0.0795 0.0877 1.0000
15.500 1.6064 0.06698 0.06118 -0.0791 0.0788 1.0000
15.750 1.6031 0.07041 0.06460 -0.0789 0.0707 1.0000
16.000 1.6025 0.07354 0.06776 -0.0787 0.0635 1.0000
16.250 1.6005 0.07693 0.07116 -0.0786 0.0568 1.0000
16.500 1.5959 0.08073 0.07497 -0.0787 0.0513 1.0000
16.750 1.5959 0.08398 0.07827 -0.0788 0.0466 1.0000
17.000 1.5913 0.08793 0.08226 -0.0792 0.0425 1.0000
17.250 1.5902 0.09143 0.08582 -0.0795 0.0390 1.0000
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