FAUVEL 14% AIRFOIL (fauvel-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FAUVEL 14% AIRFOIL (fauvel-il) Reynolds number: 100,000 Max Cl/Cd: 22.74 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fauvel-il-100000.txt Download as CSV file: xf-fauvel-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: FAUVEL 14% 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4982 0.09846 0.09456 0.0027 1.0000 0.1499
-9.000 -0.4699 0.09417 0.09003 0.0020 0.8474 0.1548
-8.500 -0.5324 0.08603 0.08171 -0.0072 0.8122 0.1662
-8.250 -0.4723 0.08301 0.07862 -0.0034 0.7914 0.1723
-8.000 -0.4848 0.07943 0.07499 -0.0053 0.7805 0.1789
-7.750 -0.5411 0.07567 0.07114 -0.0061 0.7748 0.1846
-7.500 -0.4994 0.07287 0.06838 -0.0046 0.7601 0.1904
-7.250 -0.5428 0.07017 0.06546 -0.0045 0.7542 0.2012
-7.000 -0.5084 0.06692 0.06230 -0.0033 0.7415 0.2063
-6.750 -0.5267 0.06447 0.05965 -0.0021 0.7339 0.2194
-6.500 -0.5626 0.04574 0.03886 -0.0018 0.7371 0.1044
-6.250 -0.5476 0.04259 0.03548 -0.0003 0.7279 0.1022
-6.000 -0.5329 0.03923 0.03169 0.0015 0.7178 0.0982
-5.750 -0.5187 0.03703 0.02844 0.0049 0.7098 0.0926
-5.500 -0.4962 0.03466 0.02585 0.0059 0.6989 0.0920
-5.250 -0.4739 0.03301 0.02386 0.0072 0.6897 0.0921
-5.000 -0.4505 0.03144 0.02200 0.0083 0.6792 0.0933
-4.750 -0.4249 0.02983 0.02026 0.0088 0.6699 0.0949
-4.500 -0.3981 0.02860 0.01889 0.0094 0.6598 0.0961
-4.250 -0.3704 0.02757 0.01775 0.0099 0.6499 0.0977
-4.000 -0.3428 0.02666 0.01671 0.0105 0.6407 0.1000
-3.750 -0.3148 0.02601 0.01596 0.0108 0.6308 0.1038
-3.500 -0.2870 0.02512 0.01502 0.0113 0.6227 0.1078
-3.250 -0.2593 0.02434 0.01438 0.0115 0.6128 0.1127
-3.000 -0.2337 0.02377 0.01373 0.0125 0.6054 0.1191
-2.750 -0.2097 0.02306 0.01322 0.0133 0.5959 0.1291
-2.500 -0.1879 0.02241 0.01267 0.0149 0.5893 0.1490
-2.250 -0.1727 0.02110 0.01235 0.0171 0.5816 0.2774
-2.000 -0.1446 0.01956 0.01312 0.0201 0.5743 0.8026
-1.750 -0.0883 0.02110 0.01451 0.0184 0.5641 0.8723
-1.500 0.0872 0.02400 0.01685 -0.0021 0.5483 0.9568
-1.250 0.1631 0.02366 0.01619 -0.0104 0.5394 0.9806
-1.000 0.2133 0.02339 0.01583 -0.0152 0.5302 0.9918
-0.750 0.2572 0.02292 0.01518 -0.0185 0.5238 1.0000
-0.500 0.2785 0.02301 0.01519 -0.0178 0.5183 1.0000
-0.250 0.3003 0.02313 0.01529 -0.0174 0.5113 1.0000
0.000 0.3210 0.02307 0.01511 -0.0164 0.5067 1.0000
0.250 0.3423 0.02323 0.01521 -0.0157 0.5015 1.0000
0.500 0.3643 0.02354 0.01555 -0.0154 0.4950 1.0000
0.750 0.3852 0.02357 0.01548 -0.0144 0.4906 1.0000
1.000 0.4057 0.02353 0.01530 -0.0132 0.4873 1.0000
1.250 0.4278 0.02417 0.01608 -0.0133 0.4796 1.0000
1.500 0.4488 0.02431 0.01616 -0.0124 0.4749 1.0000
1.750 0.4696 0.02437 0.01612 -0.0113 0.4715 1.0000
2.000 0.4906 0.02496 0.01677 -0.0109 0.4665 1.0000
2.250 0.5116 0.02561 0.01749 -0.0106 0.4603 1.0000
2.500 0.5323 0.02570 0.01752 -0.0095 0.4563 1.0000
2.750 0.5530 0.02571 0.01742 -0.0081 0.4532 1.0000
3.000 0.5724 0.02704 0.01894 -0.0084 0.4469 1.0000
3.250 0.5918 0.02784 0.01981 -0.0079 0.4417 1.0000
3.500 0.6120 0.02795 0.01985 -0.0065 0.4382 1.0000
3.750 0.6327 0.02791 0.01971 -0.0049 0.4354 1.0000
4.000 0.6468 0.03018 0.02226 -0.0054 0.4280 1.0000
4.250 0.6637 0.03120 0.02334 -0.0046 0.4232 1.0000
4.500 0.6838 0.03122 0.02330 -0.0029 0.4200 1.0000
4.750 0.7053 0.03102 0.02299 -0.0011 0.4175 1.0000
5.000 0.7059 0.03517 0.02752 -0.0020 0.4083 1.0000
5.250 0.7218 0.03583 0.02820 -0.0005 0.4040 1.0000
5.500 0.7449 0.03527 0.02756 0.0016 0.4014 1.0000
5.750 0.7705 0.03449 0.02663 0.0038 0.3992 1.0000
6.000 0.7464 0.04082 0.03339 0.0033 0.3867 1.0000
6.250 0.7798 0.03873 0.03117 0.0060 0.3841 1.0000
6.500 0.8159 0.03651 0.02877 0.0084 0.3819 1.0000
6.750 0.7658 0.04500 0.03766 0.0085 0.3683 1.0000
7.000 0.8112 0.04180 0.03436 0.0112 0.3665 1.0000
7.250 0.8534 0.03926 0.03168 0.0132 0.3647 1.0000
7.750 0.7637 0.05352 0.04627 0.0148 0.3440 1.0000
8.000 0.8399 0.04712 0.03982 0.0181 0.3442 1.0000
8.250 0.6962 0.06384 0.05650 0.0166 0.3274 1.0000
8.500 0.8005 0.05536 0.04813 0.0201 0.3280 1.0000
8.750 0.8824 0.04847 0.04122 0.0234 0.3279 1.0000
9.000 0.7618 0.06305 0.05576 0.0219 0.3122 1.0000
9.250 0.8389 0.05633 0.04912 0.0257 0.3121 1.0000
9.500 0.9657 0.04574 0.03851 0.0284 0.3125 1.0000
10.000 0.6424 0.08753 0.08018 0.0164 0.2703 1.0000
10.250 0.6748 0.08627 0.07896 0.0185 0.2661 1.0000
10.500 0.7246 0.08250 0.07522 0.0219 0.2641 1.0000
10.750 0.6636 0.09397 0.08668 0.0169 0.2517 1.0000
11.000 0.7109 0.09038 0.08312 0.0203 0.2491 1.0000
11.250 0.6445 0.10354 0.09628 0.0142 0.2396 1.0000
11.500 0.6813 0.10137 0.09414 0.0167 0.2351 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FAUVEL 14% AIRFOIL (fauvel-il)