DAE-21 AIRFOIL (dae21-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: DAE-21 AIRFOIL (dae21-il) Reynolds number: 200,000 Max Cl/Cd: 88.71 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-dae21-il-200000-n5.txt Download as CSV file: xf-dae21-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: DAE-21 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.0411 0.09218 0.08760 -0.0757 0.7611 0.0285
-8.250 -0.0419 0.08958 0.08504 -0.0788 0.7579 0.0291
-8.000 -0.0408 0.08671 0.08221 -0.0814 0.7547 0.0292
-7.750 -0.0244 0.08394 0.07942 -0.0804 0.7524 0.0296
-7.500 -0.0147 0.08149 0.07696 -0.0809 0.7499 0.0300
-7.250 -0.0067 0.07908 0.07452 -0.0818 0.7476 0.0304
-7.000 0.0004 0.07665 0.07214 -0.0828 0.7446 0.0309
-6.750 0.0058 0.07428 0.06981 -0.0839 0.7413 0.0313
-6.500 0.0103 0.07196 0.06751 -0.0852 0.7380 0.0318
-6.250 0.0184 0.06912 0.06468 -0.0879 0.7351 0.0326
-6.000 0.0285 0.06526 0.06079 -0.0942 0.7323 0.0337
-5.750 0.0441 0.06104 0.05650 -0.1006 0.7300 0.0339
-5.500 0.0581 0.05463 0.05005 -0.1057 0.7267 0.0249
-5.250 0.0754 0.05270 0.04812 -0.1068 0.7236 0.0243
-5.000 0.0967 0.04880 0.04414 -0.1112 0.7205 0.0234
-4.750 0.1228 0.04305 0.03820 -0.1178 0.7178 0.0224
-4.500 0.1502 0.03857 0.03347 -0.1223 0.7156 0.0233
-4.250 0.1792 0.03319 0.02767 -0.1265 0.7136 0.0236
-4.000 0.2080 0.02843 0.02244 -0.1294 0.7105 0.0230
-3.750 0.2368 0.02472 0.01815 -0.1311 0.7074 0.0230
-3.500 0.2653 0.02233 0.01528 -0.1319 0.7045 0.0233
-3.250 0.2939 0.02058 0.01308 -0.1322 0.7018 0.0239
-3.000 0.3227 0.01925 0.01128 -0.1324 0.6994 0.0251
-2.750 0.3505 0.01869 0.01066 -0.1325 0.6973 0.0263
-2.500 0.3784 0.01800 0.00984 -0.1326 0.6941 0.0275
-2.250 0.4064 0.01728 0.00897 -0.1326 0.6906 0.0284
-2.000 0.4344 0.01666 0.00823 -0.1326 0.6874 0.0293
-1.750 0.4623 0.01618 0.00774 -0.1326 0.6847 0.0307
-1.500 0.4906 0.01582 0.00730 -0.1326 0.6822 0.0333
-1.250 0.5186 0.01544 0.00690 -0.1326 0.6797 0.0355
-1.000 0.5462 0.01519 0.00668 -0.1327 0.6757 0.0379
-0.750 0.5741 0.01494 0.00642 -0.1327 0.6721 0.0414
-0.500 0.6023 0.01472 0.00617 -0.1327 0.6689 0.0478
-0.250 0.6308 0.01447 0.00591 -0.1328 0.6663 0.0578
0.000 0.6590 0.01422 0.00573 -0.1329 0.6635 0.0860
0.250 0.6860 0.01393 0.00579 -0.1330 0.6591 0.1803
0.500 0.7127 0.01351 0.00587 -0.1332 0.6553 0.3493
0.750 0.7382 0.01298 0.00592 -0.1328 0.6520 0.5623
1.000 0.7676 0.01211 0.00577 -0.1325 0.6493 1.0000
1.250 0.7950 0.01225 0.00583 -0.1325 0.6449 1.0000
1.500 0.8225 0.01237 0.00587 -0.1324 0.6404 1.0000
1.750 0.8504 0.01244 0.00584 -0.1324 0.6366 1.0000
2.000 0.8789 0.01247 0.00576 -0.1324 0.6335 1.0000
2.250 0.9056 0.01264 0.00590 -0.1323 0.6285 1.0000
2.500 0.9328 0.01276 0.00598 -0.1322 0.6237 1.0000
2.750 0.9606 0.01281 0.00596 -0.1321 0.6196 1.0000
3.000 0.9881 0.01291 0.00600 -0.1321 0.6153 1.0000
3.250 1.0143 0.01308 0.00617 -0.1319 0.6095 1.0000
3.500 1.0416 0.01316 0.00620 -0.1318 0.6046 1.0000
3.750 1.0688 0.01325 0.00625 -0.1317 0.5997 1.0000
4.000 1.0945 0.01343 0.00646 -0.1314 0.5932 1.0000
4.250 1.1216 0.01351 0.00649 -0.1313 0.5878 1.0000
4.500 1.1474 0.01368 0.00667 -0.1310 0.5814 1.0000
4.750 1.1733 0.01383 0.00682 -0.1307 0.5747 1.0000
5.000 1.1995 0.01395 0.00692 -0.1305 0.5684 1.0000
5.250 1.2243 0.01415 0.00714 -0.1301 0.5604 1.0000
5.500 1.2501 0.01428 0.00724 -0.1298 0.5536 1.0000
5.750 1.2742 0.01451 0.00752 -0.1293 0.5450 1.0000
6.000 1.2990 0.01469 0.00768 -0.1289 0.5373 1.0000
6.250 1.3224 0.01494 0.00795 -0.1283 0.5281 1.0000
6.500 1.3459 0.01518 0.00821 -0.1277 0.5191 1.0000
6.750 1.3688 0.01543 0.00845 -0.1270 0.5093 1.0000
7.000 1.3902 0.01575 0.00879 -0.1261 0.4984 1.0000
7.250 1.4112 0.01607 0.00911 -0.1252 0.4874 1.0000
7.500 1.4312 0.01643 0.00945 -0.1240 0.4758 1.0000
7.750 1.4495 0.01685 0.00987 -0.1227 0.4632 1.0000
8.000 1.4663 0.01731 0.01033 -0.1212 0.4501 1.0000
8.250 1.4810 0.01784 0.01085 -0.1194 0.4365 1.0000
8.500 1.4919 0.01842 0.01142 -0.1170 0.4228 1.0000
8.750 1.5000 0.01913 0.01211 -0.1143 0.4088 1.0000
9.000 1.5070 0.01998 0.01292 -0.1117 0.3944 1.0000
9.250 1.5136 0.02094 0.01386 -0.1092 0.3798 1.0000
9.500 1.5194 0.02203 0.01492 -0.1069 0.3646 1.0000
9.750 1.5248 0.02323 0.01609 -0.1046 0.3495 1.0000
10.000 1.5296 0.02453 0.01736 -0.1025 0.3342 1.0000
10.250 1.5341 0.02592 0.01872 -0.1004 0.3194 1.0000
10.500 1.5380 0.02741 0.02020 -0.0984 0.3044 1.0000
11.000 1.5440 0.03071 0.02343 -0.0946 0.2745 1.0000
11.250 1.5465 0.03248 0.02518 -0.0929 0.2598 1.0000
11.500 1.5491 0.03431 0.02700 -0.0912 0.2455 1.0000
11.750 1.5513 0.03623 0.02890 -0.0896 0.2316 1.0000
12.000 1.5530 0.03824 0.03090 -0.0881 0.2179 1.0000
12.250 1.5546 0.04034 0.03299 -0.0867 0.2047 1.0000
12.500 1.5563 0.04251 0.03516 -0.0855 0.1922 1.0000
12.750 1.5577 0.04479 0.03744 -0.0843 0.1801 1.0000
13.000 1.5585 0.04721 0.03986 -0.0833 0.1684 1.0000
13.250 1.5587 0.04977 0.04241 -0.0823 0.1569 1.0000
13.500 1.5605 0.05225 0.04490 -0.0815 0.1459 1.0000
13.750 1.5616 0.05486 0.04753 -0.0808 0.1356 1.0000
14.000 1.5621 0.05759 0.05028 -0.0801 0.1263 1.0000
14.250 1.5612 0.06057 0.05326 -0.0796 0.1175 1.0000
14.500 1.5628 0.06332 0.05607 -0.0792 0.1090 1.0000
14.750 1.5630 0.06630 0.05909 -0.0789 0.1018 1.0000
15.000 1.5621 0.06948 0.06230 -0.0787 0.0948 1.0000
15.250 1.5627 0.07253 0.06541 -0.0785 0.0884 1.0000
15.750 1.5616 0.07912 0.07212 -0.0786 0.0774 1.0000
16.000 1.5586 0.08285 0.07590 -0.0788 0.0728 1.0000
16.250 1.5590 0.08615 0.07929 -0.0791 0.0683 1.0000
16.500 1.5564 0.08992 0.08312 -0.0795 0.0643 1.0000
16.750 1.5550 0.09359 0.08689 -0.0801 0.0608 1.0000
17.000 1.5533 0.09735 0.09073 -0.0807 0.0574 1.0000
17.500 1.5485 0.10521 0.09878 -0.0823 0.0517 1.0000
17.750 1.5455 0.10930 0.10297 -0.0833 0.0490 1.0000
18.250 1.5401 0.11752 0.11137 -0.0857 0.0447 1.0000
18.500 1.5375 0.12166 0.11562 -0.0870 0.0427 1.0000
18.750 1.5331 0.12612 0.12015 -0.0886 0.0410 1.0000
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Polar data table (+)
Polar graphs
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