GOE 548 AIRFOIL (goe548-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 548 AIRFOIL (goe548-il) Reynolds number: 50,000 Max Cl/Cd: 32.77 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe548-il-50000-n5.txt Download as CSV file: xf-goe548-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 548 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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.4533 0.10298 0.09599 -0.0477 1.0000 0.0488
-10.250 -0.4614 0.09826 0.09137 -0.0490 1.0000 0.0487
-9.750 -0.5222 0.08030 0.07356 -0.0602 1.0000 0.0432
-9.500 -0.5420 0.07645 0.06976 -0.0603 1.0000 0.0430
-9.250 -0.5643 0.07351 0.06685 -0.0585 1.0000 0.0428
-9.000 -0.5875 0.07079 0.06412 -0.0556 1.0000 0.0426
-8.750 -0.6074 0.06766 0.06093 -0.0529 1.0000 0.0425
-8.500 -0.6240 0.06452 0.05766 -0.0502 1.0000 0.0425
-8.250 -0.6393 0.06119 0.05411 -0.0473 1.0000 0.0426
-8.000 -0.6506 0.05786 0.05048 -0.0443 1.0000 0.0430
-7.750 -0.6580 0.05454 0.04683 -0.0415 1.0000 0.0437
-7.500 -0.6568 0.05161 0.04382 -0.0393 1.0000 0.0453
-7.250 -0.6526 0.04925 0.04132 -0.0372 1.0000 0.0473
-7.000 -0.6481 0.04657 0.03833 -0.0349 1.0000 0.0493
-6.750 -0.6414 0.04365 0.03498 -0.0325 1.0000 0.0508
-6.500 -0.6316 0.04088 0.03172 -0.0303 1.0000 0.0529
-6.250 -0.6194 0.03863 0.02878 -0.0280 1.0000 0.0566
-6.000 -0.6055 0.03645 0.02660 -0.0266 1.0000 0.0610
-5.750 -0.5887 0.03459 0.02446 -0.0249 1.0000 0.0653
-5.500 -0.5715 0.03296 0.02251 -0.0232 1.0000 0.0722
-5.250 -0.5545 0.03173 0.02118 -0.0217 1.0000 0.0796
-5.000 -0.5371 0.03053 0.01979 -0.0201 1.0000 0.0876
-4.750 -0.5192 0.02957 0.01861 -0.0186 1.0000 0.0977
-4.500 -0.4959 0.02850 0.01749 -0.0181 0.9979 0.1060
-4.250 -0.4653 0.02761 0.01650 -0.0192 0.9931 0.1212
-4.000 -0.4344 0.02675 0.01564 -0.0204 0.9879 0.1429
-3.750 -0.4038 0.02577 0.01486 -0.0218 0.9830 0.1851
-3.500 -0.3761 0.02476 0.01437 -0.0227 0.9776 0.2952
-3.250 -0.3505 0.02421 0.01420 -0.0229 0.9714 0.4071
-3.000 -0.3226 0.02395 0.01423 -0.0230 0.9658 0.5011
-2.750 -0.3008 0.02368 0.01430 -0.0215 0.9588 0.6041
-2.500 -0.2743 0.02361 0.01459 -0.0201 0.9534 0.7088
-2.250 -0.2447 0.02363 0.01472 -0.0195 0.9472 0.7900
-2.000 -0.1978 0.02384 0.01479 -0.0227 0.9426 0.8519
-1.750 -0.1336 0.02419 0.01486 -0.0298 0.9399 0.9074
-1.500 -0.0619 0.02455 0.01488 -0.0387 0.9376 0.9533
-1.250 0.0046 0.02483 0.01486 -0.0473 0.9342 0.9974
-1.000 0.0347 0.02485 0.01464 -0.0490 0.9261 1.0000
-0.750 0.0570 0.02488 0.01447 -0.0490 0.9161 1.0000
-0.500 0.0921 0.02499 0.01438 -0.0512 0.9096 1.0000
-0.250 0.1120 0.02509 0.01433 -0.0505 0.8990 1.0000
0.000 0.1375 0.02524 0.01433 -0.0508 0.8901 1.0000
0.250 0.1692 0.02537 0.01435 -0.0520 0.8827 1.0000
0.500 0.1897 0.02555 0.01442 -0.0512 0.8724 1.0000
0.750 0.2172 0.02571 0.01450 -0.0516 0.8640 1.0000
1.000 0.2451 0.02587 0.01459 -0.0520 0.8556 1.0000
1.250 0.2656 0.02608 0.01475 -0.0510 0.8452 1.0000
1.500 0.2929 0.02625 0.01488 -0.0512 0.8368 1.0000
1.750 0.3204 0.02639 0.01501 -0.0514 0.8281 1.0000
2.000 0.3403 0.02664 0.01526 -0.0503 0.8175 1.0000
2.250 0.3671 0.02681 0.01543 -0.0503 0.8087 1.0000
2.500 0.3960 0.02693 0.01560 -0.0506 0.8004 1.0000
2.750 0.4170 0.02719 0.01589 -0.0496 0.7898 1.0000
3.000 0.4480 0.02727 0.01603 -0.0502 0.7823 1.0000
3.250 0.4742 0.02743 0.01627 -0.0500 0.7730 1.0000
3.500 0.4971 0.02769 0.01663 -0.0493 0.7629 1.0000
3.750 0.5346 0.02759 0.01666 -0.0507 0.7568 1.0000
4.000 0.5548 0.02794 0.01711 -0.0495 0.7457 1.0000
4.250 0.5818 0.02813 0.01746 -0.0494 0.7364 1.0000
4.500 0.6201 0.02794 0.01746 -0.0507 0.7284 1.0000
4.750 0.6492 0.02782 0.01750 -0.0503 0.7152 1.0000
5.000 0.6880 0.02712 0.01703 -0.0509 0.6994 1.0000
5.250 0.7152 0.02668 0.01675 -0.0495 0.6789 1.0000
5.500 0.7404 0.02639 0.01663 -0.0479 0.6585 1.0000
5.750 0.7684 0.02611 0.01657 -0.0470 0.6399 1.0000
6.000 0.7808 0.02634 0.01698 -0.0439 0.6185 1.0000
6.250 0.8024 0.02613 0.01694 -0.0418 0.5947 1.0000
6.500 0.8217 0.02600 0.01697 -0.0393 0.5680 1.0000
6.750 0.8376 0.02576 0.01673 -0.0360 0.5260 1.0000
7.000 0.8447 0.02578 0.01642 -0.0311 0.4561 1.0000
7.250 0.8505 0.02639 0.01681 -0.0271 0.3895 1.0000
7.500 0.8576 0.02724 0.01747 -0.0237 0.3300 1.0000
7.750 0.8568 0.02862 0.01832 -0.0196 0.2615 1.0000
8.000 0.8564 0.03027 0.01957 -0.0161 0.2107 1.0000
8.250 0.8592 0.03190 0.02100 -0.0133 0.1705 1.0000
8.500 0.8637 0.03353 0.02251 -0.0108 0.1252 1.0000
8.750 0.8627 0.03569 0.02424 -0.0081 0.0837 1.0000
9.000 0.8673 0.03754 0.02608 -0.0059 0.0618 1.0000
9.250 0.8729 0.03935 0.02787 -0.0039 0.0510 1.0000
9.500 0.8778 0.04127 0.02976 -0.0020 0.0458 1.0000
9.750 0.8835 0.04318 0.03177 -0.0001 0.0432 1.0000
10.000 0.8896 0.04513 0.03387 0.0016 0.0410 1.0000
10.250 0.8951 0.04720 0.03613 0.0032 0.0387 1.0000
10.500 0.9001 0.04938 0.03838 0.0046 0.0369 1.0000
10.750 0.9095 0.05141 0.04060 0.0060 0.0349 1.0000
11.000 0.9217 0.05337 0.04281 0.0073 0.0333 1.0000
11.250 0.9369 0.05536 0.04503 0.0086 0.0323 1.0000
11.500 0.9531 0.05759 0.04752 0.0098 0.0314 1.0000
11.750 0.9666 0.06017 0.05036 0.0109 0.0307 1.0000
12.000 0.9750 0.06311 0.05356 0.0119 0.0302 1.0000
12.250 0.9784 0.06632 0.05701 0.0127 0.0297 1.0000
12.500 0.9775 0.06979 0.06071 0.0133 0.0292 1.0000
12.750 0.9733 0.07368 0.06480 0.0137 0.0288 1.0000
13.000 0.9630 0.07820 0.06951 0.0136 0.0283 1.0000
13.250 0.9536 0.08260 0.07415 0.0132 0.0284 1.0000
13.500 0.9383 0.08763 0.07943 0.0120 0.0284 1.0000
13.750 0.9232 0.09293 0.08500 0.0101 0.0286 1.0000
14.000 0.9034 0.09931 0.09163 0.0071 0.0288 1.0000
14.250 0.8770 0.10781 0.10041 0.0022 0.0294 1.0000
14.500 0.8331 0.12278 0.11568 -0.0074 0.0316 1.0000
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Polar data table (+)
Polar graphs
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