NACA 63-412 AIRFOIL (n63412-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 63-412 AIRFOIL (n63412-il) Reynolds number: 50,000 Max Cl/Cd: 35.99 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63412-il-50000-n5.txt Download as CSV file: xf-n63412-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-412 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.4759 0.10169 0.09432 -0.0425 1.0000 0.0575
-10.500 -0.4749 0.09720 0.08987 -0.0440 1.0000 0.0571
-10.250 -0.4768 0.09223 0.08495 -0.0462 1.0000 0.0567
-10.000 -0.4815 0.08675 0.07952 -0.0491 1.0000 0.0561
-9.750 -0.4911 0.08100 0.07383 -0.0525 1.0000 0.0556
-9.500 -0.5061 0.07551 0.06838 -0.0559 1.0000 0.0551
-9.250 -0.5259 0.07084 0.06374 -0.0581 1.0000 0.0546
-9.000 -0.5460 0.06689 0.05978 -0.0588 1.0000 0.0542
-8.750 -0.5620 0.06310 0.05590 -0.0587 1.0000 0.0539
-8.500 -0.5747 0.05950 0.05216 -0.0580 1.0000 0.0538
-8.250 -0.5831 0.05611 0.04857 -0.0568 1.0000 0.0538
-8.000 -0.5881 0.05292 0.04513 -0.0554 1.0000 0.0543
-7.750 -0.5903 0.04991 0.04178 -0.0538 1.0000 0.0552
-7.500 -0.5891 0.04703 0.03848 -0.0521 1.0000 0.0564
-7.250 -0.5841 0.04432 0.03525 -0.0505 1.0000 0.0575
-7.000 -0.5745 0.04174 0.03244 -0.0491 1.0000 0.0586
-6.750 -0.5624 0.03952 0.03001 -0.0478 1.0000 0.0596
-6.500 -0.5485 0.03754 0.02781 -0.0465 1.0000 0.0610
-6.250 -0.5334 0.03591 0.02598 -0.0453 1.0000 0.0636
-6.000 -0.5169 0.03434 0.02410 -0.0441 1.0000 0.0670
-5.750 -0.4988 0.03284 0.02222 -0.0428 1.0000 0.0695
-5.500 -0.4673 0.03111 0.02040 -0.0439 0.9948 0.0726
-5.250 -0.4338 0.02988 0.01905 -0.0454 0.9881 0.0790
-5.000 -0.4009 0.02870 0.01773 -0.0465 0.9815 0.0855
-4.750 -0.3671 0.02764 0.01657 -0.0480 0.9749 0.0939
-4.500 -0.3361 0.02661 0.01552 -0.0493 0.9669 0.1064
-4.250 -0.3029 0.02552 0.01439 -0.0511 0.9598 0.1257
-3.750 -0.2443 0.02196 0.01275 -0.0550 0.9451 0.4112
-3.500 -0.2174 0.02206 0.01317 -0.0540 0.9366 0.5528
-3.250 -0.1933 0.02254 0.01373 -0.0520 0.9280 0.6249
-3.000 -0.1707 0.02326 0.01446 -0.0490 0.9199 0.6833
-2.750 -0.1506 0.02372 0.01484 -0.0460 0.9111 0.7197
-2.500 -0.1219 0.02392 0.01488 -0.0452 0.9040 0.7447
-2.250 -0.0966 0.02395 0.01473 -0.0445 0.8955 0.7618
-2.000 -0.0643 0.02393 0.01452 -0.0452 0.8889 0.7773
-1.750 -0.0386 0.02390 0.01434 -0.0449 0.8802 0.7916
-1.500 -0.0070 0.02383 0.01411 -0.0456 0.8734 0.8054
-1.250 0.0176 0.02381 0.01397 -0.0452 0.8647 0.8192
-1.000 0.0489 0.02374 0.01377 -0.0458 0.8580 0.8326
-0.750 0.0728 0.02373 0.01368 -0.0453 0.8492 0.8460
-0.500 0.1040 0.02366 0.01352 -0.0459 0.8425 0.8594
-0.250 0.1285 0.02366 0.01346 -0.0455 0.8338 0.8738
0.000 0.1608 0.02358 0.01332 -0.0464 0.8271 0.8882
0.250 0.1881 0.02360 0.01331 -0.0466 0.8187 0.9038
0.500 0.2238 0.02355 0.01322 -0.0482 0.8120 0.9190
0.750 0.2591 0.02359 0.01325 -0.0501 0.8044 0.9352
1.000 0.3013 0.02360 0.01325 -0.0532 0.7974 0.9508
1.250 0.3476 0.02362 0.01328 -0.0572 0.7911 0.9662
1.500 0.3902 0.02373 0.01342 -0.0608 0.7828 0.9869
1.750 0.4180 0.02388 0.01356 -0.0616 0.7749 1.0000
2.000 0.4383 0.02413 0.01381 -0.0611 0.7659 1.0000
2.250 0.4599 0.02448 0.01416 -0.0609 0.7568 1.0000
2.500 0.4894 0.02467 0.01437 -0.0616 0.7495 1.0000
2.750 0.5119 0.02512 0.01485 -0.0616 0.7398 1.0000
3.000 0.5443 0.02527 0.01505 -0.0625 0.7332 1.0000
3.250 0.5662 0.02580 0.01564 -0.0623 0.7228 1.0000
3.500 0.6008 0.02586 0.01576 -0.0632 0.7167 1.0000
3.750 0.6216 0.02645 0.01646 -0.0628 0.7054 1.0000
4.000 0.6482 0.02682 0.01693 -0.0629 0.6962 1.0000
4.250 0.6784 0.02699 0.01721 -0.0631 0.6877 1.0000
4.500 0.7013 0.02748 0.01785 -0.0627 0.6762 1.0000
4.750 0.7294 0.02769 0.01820 -0.0626 0.6661 1.0000
5.000 0.7613 0.02762 0.01828 -0.0626 0.6560 1.0000
5.250 0.7860 0.02782 0.01865 -0.0619 0.6423 1.0000
5.500 0.8128 0.02777 0.01881 -0.0611 0.6275 1.0000
5.750 0.8424 0.02738 0.01859 -0.0602 0.6106 1.0000
6.000 0.8654 0.02716 0.01854 -0.0585 0.5884 1.0000
6.250 0.8910 0.02650 0.01802 -0.0565 0.5606 1.0000
6.500 0.9104 0.02608 0.01767 -0.0539 0.5234 1.0000
6.750 0.9270 0.02590 0.01748 -0.0511 0.4754 1.0000
7.000 0.9401 0.02612 0.01757 -0.0483 0.4122 1.0000
7.250 0.9485 0.02693 0.01791 -0.0451 0.3202 1.0000
7.500 0.9472 0.02883 0.01909 -0.0417 0.2324 1.0000
7.750 0.9432 0.03108 0.02083 -0.0386 0.1761 1.0000
8.000 0.9425 0.03336 0.02280 -0.0362 0.1413 1.0000
8.250 0.9449 0.03557 0.02483 -0.0343 0.1206 1.0000
8.500 0.9496 0.03769 0.02684 -0.0326 0.1057 1.0000
8.750 0.9578 0.03965 0.02879 -0.0312 0.0943 1.0000
9.000 0.9689 0.04152 0.03072 -0.0299 0.0862 1.0000
9.250 0.9817 0.04332 0.03250 -0.0288 0.0788 1.0000
9.500 1.0031 0.04488 0.03423 -0.0278 0.0722 1.0000
9.750 1.0260 0.04648 0.03585 -0.0271 0.0667 1.0000
10.000 1.0540 0.04828 0.03785 -0.0266 0.0615 1.0000
10.250 1.0829 0.05037 0.04020 -0.0264 0.0578 1.0000
10.500 1.1069 0.05267 0.04255 -0.0263 0.0549 1.0000
10.750 1.1236 0.05550 0.04568 -0.0257 0.0525 1.0000
11.000 1.1309 0.05848 0.04908 -0.0245 0.0507 1.0000
11.250 1.1352 0.06171 0.05266 -0.0235 0.0494 1.0000
11.500 1.1352 0.06522 0.05650 -0.0225 0.0487 1.0000
11.750 1.1305 0.06897 0.06057 -0.0216 0.0482 1.0000
12.000 1.1213 0.07308 0.06499 -0.0212 0.0478 1.0000
12.250 1.1079 0.07762 0.06981 -0.0213 0.0476 1.0000
12.500 1.0904 0.08277 0.07525 -0.0222 0.0476 1.0000
12.750 1.0688 0.08870 0.08144 -0.0240 0.0477 1.0000
13.000 1.0430 0.09570 0.08869 -0.0272 0.0480 1.0000
13.250 1.0136 0.10409 0.09729 -0.0319 0.0485 1.0000
13.500 0.9815 0.11425 0.10762 -0.0384 0.0493 1.0000
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Polar data table (+)
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