FX 60-177 AIRFOIL (fx60177-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: FX 60-177 AIRFOIL (fx60177-il) Reynolds number: 50,000 Max Cl/Cd: 24.18 at α=9° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx60177-il-50000.txt Download as CSV file: xf-fx60177-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: FX 60-177 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.4819 0.11925 0.11262 -0.0450 1.0001 0.1496
-11.000 -0.5127 0.10945 0.10294 -0.0490 1.0001 0.1325
-10.750 -0.6546 0.09177 0.08522 -0.0587 1.0001 0.1070
-10.500 -0.6714 0.08786 0.08131 -0.0570 1.0001 0.1060
-10.250 -0.6950 0.08436 0.07781 -0.0546 1.0001 0.1049
-10.000 -0.7235 0.08136 0.07478 -0.0513 1.0001 0.1037
-9.750 -0.7545 0.07851 0.07188 -0.0476 1.0001 0.1022
-9.500 -0.7830 0.07504 0.06826 -0.0445 1.0001 0.1005
-9.250 -0.8074 0.07134 0.06432 -0.0414 1.0001 0.0990
-9.000 -0.8238 0.06765 0.06034 -0.0387 1.0001 0.0984
-8.750 -0.8322 0.06414 0.05654 -0.0363 1.0001 0.0987
-8.500 -0.8348 0.06070 0.05278 -0.0342 1.0001 0.0997
-8.250 -0.8328 0.05731 0.04900 -0.0324 1.0001 0.1012
-8.000 -0.8269 0.05408 0.04532 -0.0308 1.0001 0.1038
-7.750 -0.8172 0.05107 0.04166 -0.0294 1.0001 0.1070
-7.500 -0.8005 0.04788 0.03833 -0.0282 1.0001 0.1109
-7.250 -0.7826 0.04572 0.03602 -0.0269 1.0001 0.1175
-7.000 -0.7635 0.04343 0.03339 -0.0255 1.0001 0.1243
-6.750 -0.7430 0.04168 0.03165 -0.0238 1.0001 0.1327
-6.500 -0.7228 0.04014 0.03014 -0.0218 1.0001 0.1446
-6.250 -0.7022 0.03889 0.02899 -0.0193 1.0001 0.1598
-6.000 -0.6832 0.03777 0.02791 -0.0167 1.0001 0.1837
-5.750 -0.6697 0.03622 0.02692 -0.0139 1.0001 0.2204
-5.500 -0.6637 0.03307 0.02517 -0.0120 1.0001 0.3269
-5.250 -0.6687 0.03600 0.03002 0.0019 1.0001 0.5701
-5.000 -0.6562 0.04072 0.03449 0.0116 1.0001 0.6227
-4.750 -0.6392 0.04454 0.03804 0.0194 1.0001 0.6545
-4.500 -0.6153 0.04842 0.04165 0.0265 1.0001 0.6856
-4.250 -0.5857 0.05203 0.04499 0.0323 1.0001 0.7233
-4.000 -0.5232 0.05514 0.04767 0.0326 1.0001 0.7594
-3.750 -0.4993 0.05551 0.04781 0.0341 1.0001 0.7879
-3.500 -0.4112 0.05624 0.04808 0.0260 1.0001 0.8225
-3.250 -0.3798 0.05586 0.04748 0.0248 1.0001 0.8485
-3.000 -0.3330 0.05518 0.04654 0.0207 1.0001 0.8717
-2.750 -0.3087 0.05450 0.04570 0.0197 1.0001 0.8920
-2.500 -0.2818 0.05377 0.04480 0.0179 1.0001 0.9106
-2.250 -0.2575 0.05305 0.04393 0.0163 1.0001 0.9247
-2.000 -0.2358 0.05244 0.04319 0.0150 1.0001 0.9375
-1.750 -0.2100 0.05191 0.04252 0.0128 1.0001 0.9499
-1.500 -0.1788 0.05147 0.04195 0.0096 1.0001 0.9619
-1.250 -0.1468 0.05114 0.04151 0.0060 1.0001 0.9725
-1.000 -0.1152 0.05095 0.04121 0.0024 1.0001 0.9817
-0.750 -0.0840 0.05092 0.04108 -0.0012 1.0001 0.9899
-0.500 -0.0516 0.05107 0.04115 -0.0052 1.0001 0.9976
-0.250 -0.0400 0.05120 0.04124 -0.0052 1.0001 0.9999
0.000 -0.0385 0.05128 0.04130 -0.0033 1.0001 0.9999
0.250 -0.0372 0.05139 0.04140 -0.0014 1.0001 0.9999
0.500 -0.0359 0.05155 0.04155 0.0004 1.0001 0.9999
0.750 -0.0348 0.05176 0.04174 0.0023 1.0001 0.9999
1.000 -0.0338 0.05200 0.04198 0.0041 1.0001 0.9999
1.250 0.0361 0.05330 0.04325 -0.0071 0.9756 0.9999
1.500 0.0865 0.05408 0.04400 -0.0139 0.9509 0.9999
1.750 0.1290 0.05477 0.04466 -0.0189 0.9309 0.9999
2.000 0.1577 0.05510 0.04498 -0.0213 0.9106 0.9999
2.250 0.1866 0.05540 0.04529 -0.0234 0.8909 0.9999
2.500 0.2189 0.05574 0.04562 -0.0259 0.8727 0.9999
2.750 0.2517 0.05604 0.04593 -0.0282 0.8555 0.9999
3.000 0.2721 0.05627 0.04618 -0.0285 0.8382 0.9999
3.250 0.2885 0.05648 0.04642 -0.0281 0.8204 0.9999
3.500 0.3080 0.05665 0.04661 -0.0280 0.8026 0.9999
3.750 0.3304 0.05673 0.04672 -0.0281 0.7849 0.9999
4.000 0.3540 0.05674 0.04676 -0.0282 0.7676 0.9999
4.250 0.3751 0.05680 0.04686 -0.0280 0.7512 0.9999
4.500 0.3931 0.05692 0.04703 -0.0272 0.7353 0.9999
4.750 0.4093 0.05707 0.04722 -0.0262 0.7196 0.9999
5.000 0.4249 0.05720 0.04740 -0.0250 0.7040 0.9999
5.250 0.4429 0.05714 0.04739 -0.0239 0.6879 0.9999
5.500 0.4612 0.05698 0.04728 -0.0226 0.6718 0.9999
5.750 0.4773 0.05687 0.04723 -0.0210 0.6558 0.9999
6.000 0.4922 0.05678 0.04719 -0.0192 0.6402 0.9999
6.250 0.5040 0.05686 0.04732 -0.0172 0.6249 0.9999
6.500 0.5129 0.05713 0.04764 -0.0149 0.6100 0.9999
6.750 0.5222 0.05743 0.04801 -0.0127 0.5952 0.9999
7.000 0.5364 0.05765 0.04829 -0.0111 0.5809 0.9999
7.250 0.6047 0.05454 0.04533 -0.0130 0.5717 0.9999
7.500 0.6235 0.05442 0.04529 -0.0116 0.5566 0.9999
7.750 0.6469 0.05410 0.04505 -0.0107 0.5414 0.9999
8.000 0.6695 0.05407 0.04513 -0.0099 0.5262 0.9999
8.250 0.6945 0.05404 0.04521 -0.0094 0.5112 0.9999
9.000 0.9934 0.04108 0.03251 -0.0254 0.4583 0.9999
9.250 1.0096 0.04186 0.03337 -0.0241 0.4417 0.9999
9.500 1.0295 0.04274 0.03431 -0.0234 0.4253 0.9999
9.750 1.0537 0.04367 0.03530 -0.0233 0.4091 0.9999
10.000 1.0791 0.04469 0.03638 -0.0235 0.3933 0.9999
10.250 1.0511 0.04777 0.03971 -0.0184 0.3840 0.9999
10.500 1.0550 0.04997 0.04204 -0.0168 0.3720 0.9999
10.750 0.7682 0.08192 0.07409 -0.0122 0.3756 0.9999
11.000 0.7129 0.09374 0.08589 -0.0158 0.3674 0.9999
11.250 0.7133 0.09825 0.09048 -0.0173 0.3585 0.9999
11.500 0.6878 0.10624 0.09850 -0.0203 0.3540 0.9999
11.750 0.7181 0.10760 0.10001 -0.0208 0.3451 0.9999
12.000 0.6901 0.11606 0.10850 -0.0244 0.3434 0.9999
12.250 0.6751 0.12273 0.11522 -0.0275 0.3416 0.9999
12.500 0.6691 0.12874 0.12131 -0.0305 0.3421 0.9999
12.750 0.6701 0.13441 0.12706 -0.0334 0.3439 0.9999
13.000 0.6852 0.13983 0.13261 -0.0361 0.3462 0.9999
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Polar data table (+)
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