FX 63-147 AIRFOIL (fx63147-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: FX 63-147 AIRFOIL (fx63147-il) Reynolds number: 100,000 Max Cl/Cd: 48.6 at α=8.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx63147-il-100000-n5.txt Download as CSV file: xf-fx63147-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: FX 63-147 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 -0.3677 0.10546 0.10057 -0.0485 1.0001 0.0283
-10.250 -0.3632 0.10313 0.09825 -0.0476 1.0001 0.0268
-9.750 -0.3925 0.08746 0.08273 -0.0552 1.0001 0.0225
-9.500 -0.3974 0.08409 0.07942 -0.0556 1.0001 0.0223
-9.250 -0.4085 0.07994 0.07533 -0.0568 1.0001 0.0220
-9.000 -0.4255 0.07598 0.07140 -0.0573 1.0001 0.0218
-8.750 -0.4470 0.07272 0.06817 -0.0565 1.0001 0.0216
-8.500 -0.4726 0.07029 0.06576 -0.0539 1.0001 0.0215
-8.000 -0.4819 0.06117 0.05631 -0.0602 0.9844 0.0207
-7.750 -0.4751 0.05572 0.05053 -0.0636 0.9731 0.0201
-7.500 -0.4680 0.05050 0.04488 -0.0652 0.9613 0.0193
-7.000 -0.4429 0.04130 0.03443 -0.0658 0.9384 0.0179
-6.750 -0.4186 0.03820 0.03096 -0.0670 0.9307 0.0178
-6.500 -0.3950 0.03546 0.02784 -0.0674 0.9215 0.0177
-6.250 -0.3692 0.03300 0.02499 -0.0677 0.9128 0.0176
-6.000 -0.3389 0.03070 0.02234 -0.0685 0.9060 0.0176
-5.750 -0.3110 0.02882 0.02017 -0.0686 0.8975 0.0177
-5.500 -0.2790 0.02709 0.01819 -0.0693 0.8909 0.0179
-5.250 -0.2505 0.02563 0.01660 -0.0693 0.8821 0.0181
-5.000 -0.2184 0.02422 0.01511 -0.0702 0.8750 0.0186
-4.750 -0.1904 0.02308 0.01392 -0.0704 0.8649 0.0192
-4.500 -0.1546 0.02192 0.01264 -0.0722 0.8583 0.0202
-4.250 -0.1256 0.02104 0.01161 -0.0726 0.8468 0.0215
-4.000 -0.0921 0.02018 0.01066 -0.0740 0.8370 0.0244
-3.750 -0.0537 0.01939 0.00979 -0.0762 0.8283 0.0304
-3.500 -0.0222 0.01857 0.00895 -0.0771 0.8166 0.0472
-3.250 0.0052 0.01678 0.00820 -0.0783 0.8057 0.2545
-3.000 0.0252 0.01549 0.00844 -0.0767 0.7956 0.5827
-2.750 0.0441 0.01619 0.00937 -0.0723 0.7836 0.6861
-2.500 0.0676 0.01720 0.01028 -0.0693 0.7719 0.7432
-2.250 0.0930 0.01815 0.01105 -0.0666 0.7612 0.7700
-2.000 0.1219 0.01870 0.01139 -0.0654 0.7510 0.7872
-1.750 0.1483 0.01882 0.01133 -0.0649 0.7393 0.7964
-1.500 0.1765 0.01888 0.01118 -0.0648 0.7286 0.8041
-1.250 0.2067 0.01891 0.01100 -0.0652 0.7190 0.8110
-1.000 0.2312 0.01899 0.01094 -0.0643 0.7078 0.8180
-0.750 0.2569 0.01901 0.01080 -0.0640 0.6976 0.8254
-0.500 0.2856 0.01904 0.01066 -0.0640 0.6886 0.8303
-0.250 0.3080 0.01905 0.01057 -0.0634 0.6783 0.8385
0.000 0.3344 0.01909 0.01049 -0.0628 0.6699 0.8428
0.250 0.3585 0.01913 0.01042 -0.0621 0.6609 0.8493
0.500 0.3827 0.01916 0.01035 -0.0617 0.6527 0.8555
0.750 0.4079 0.01917 0.01028 -0.0611 0.6446 0.8599
1.000 0.4312 0.01921 0.01024 -0.0605 0.6366 0.8662
1.250 0.4558 0.01923 0.01019 -0.0600 0.6290 0.8711
1.500 0.4792 0.01926 0.01017 -0.0592 0.6214 0.8759
1.750 0.5031 0.01930 0.01014 -0.0588 0.6138 0.8810
2.000 0.5263 0.01932 0.01013 -0.0581 0.6061 0.8850
2.250 0.5502 0.01933 0.01010 -0.0575 0.5985 0.8887
2.500 0.5734 0.01938 0.01012 -0.0569 0.5912 0.8931
2.750 0.5963 0.01945 0.01018 -0.0564 0.5838 0.8977
3.000 0.6202 0.01947 0.01018 -0.0557 0.5772 0.9013
3.250 0.6419 0.01953 0.01026 -0.0548 0.5698 0.9057
3.500 0.6678 0.01962 0.01030 -0.0548 0.5641 0.9099
3.750 0.6877 0.01969 0.01046 -0.0536 0.5567 0.9136
4.000 0.7129 0.01975 0.01050 -0.0533 0.5506 0.9169
4.250 0.7341 0.01986 0.01066 -0.0524 0.5436 0.9208
4.500 0.7578 0.01998 0.01080 -0.0520 0.5368 0.9243
4.750 0.7803 0.02006 0.01094 -0.0513 0.5301 0.9275
5.000 0.8012 0.02015 0.01109 -0.0503 0.5226 0.9312
5.250 0.8251 0.02027 0.01122 -0.0499 0.5161 0.9346
5.500 0.8449 0.02044 0.01151 -0.0489 0.5086 0.9385
5.750 0.8708 0.02053 0.01161 -0.0488 0.5029 0.9414
6.000 0.8887 0.02070 0.01194 -0.0474 0.4952 0.9453
6.250 0.9131 0.02086 0.01214 -0.0471 0.4891 0.9487
6.500 0.9336 0.02109 0.01249 -0.0462 0.4823 0.9523
6.750 0.9561 0.02123 0.01272 -0.0456 0.4750 0.9553
7.000 0.9770 0.02142 0.01303 -0.0448 0.4676 0.9589
7.250 0.9979 0.02159 0.01329 -0.0440 0.4593 0.9627
7.500 1.0167 0.02177 0.01360 -0.0428 0.4502 0.9665
7.750 1.0369 0.02184 0.01374 -0.0418 0.4396 0.9701
8.000 1.0530 0.02196 0.01394 -0.0402 0.4273 0.9745
8.250 1.0669 0.02206 0.01414 -0.0383 0.4130 0.9796
8.750 1.0974 0.02258 0.01483 -0.0354 0.3826 0.9921
9.000 1.1103 0.02285 0.01520 -0.0337 0.3680 0.9998
9.250 1.1256 0.02339 0.01583 -0.0326 0.3529 0.9999
9.500 1.1405 0.02405 0.01655 -0.0317 0.3370 0.9999
9.750 1.1547 0.02480 0.01740 -0.0307 0.3206 0.9999
10.000 1.1683 0.02568 0.01839 -0.0299 0.3037 0.9999
10.250 1.1811 0.02668 0.01948 -0.0291 0.2868 0.9999
10.500 1.1922 0.02786 0.02075 -0.0284 0.2681 0.9999
10.750 1.2003 0.02929 0.02222 -0.0275 0.2473 0.9999
11.000 1.2056 0.03101 0.02396 -0.0267 0.2262 0.9999
11.500 1.2058 0.03575 0.02865 -0.0255 0.1777 0.9999
11.750 1.2037 0.03859 0.03149 -0.0253 0.1559 0.9999
12.000 1.1998 0.04175 0.03465 -0.0254 0.1368 0.9999
12.250 1.1950 0.04518 0.03809 -0.0257 0.1207 0.9999
12.750 1.1822 0.05282 0.04575 -0.0270 0.0926 0.9999
13.000 1.1773 0.05664 0.04964 -0.0278 0.0823 0.9999
13.250 1.1724 0.06055 0.05364 -0.0287 0.0729 0.9999
13.500 1.1661 0.06470 0.05785 -0.0298 0.0653 0.9999
13.750 1.1596 0.06898 0.06220 -0.0309 0.0602 0.9999
14.000 1.1515 0.07357 0.06682 -0.0323 0.0563 0.9999
14.250 1.1476 0.07768 0.07108 -0.0335 0.0525 0.9999
14.500 1.1444 0.08172 0.07527 -0.0348 0.0488 0.9999
14.750 1.1398 0.08608 0.07972 -0.0363 0.0460 0.9999
15.000 1.1364 0.09021 0.08394 -0.0376 0.0436 0.9999
15.250 1.1368 0.09393 0.08781 -0.0389 0.0410 0.9999
|
Polar data table (+)
Polar graphs
<< Back to FX 63-147 AIRFOIL (fx63147-il)