OAF128 AIRFOIL (oaf128-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: OAF128 AIRFOIL (oaf128-il) Reynolds number: 100,000 Max Cl/Cd: 41.44 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-oaf128-il-100000.txt Download as CSV file: xf-oaf128-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: OAF128 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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.8660 0.07826 0.07293 -0.0052 1.0000 0.0784
-10.750 -0.8977 0.06900 0.06355 -0.0129 1.0000 0.0775
-10.500 -0.9295 0.06145 0.05572 -0.0183 1.0000 0.0766
-10.250 -0.9473 0.05474 0.04856 -0.0214 1.0000 0.0764
-10.000 -0.9510 0.04911 0.04240 -0.0233 1.0000 0.0767
-9.750 -0.9445 0.04439 0.03708 -0.0244 1.0000 0.0775
-9.500 -0.9313 0.04043 0.03246 -0.0252 1.0000 0.0787
-9.250 -0.9093 0.03810 0.03016 -0.0254 1.0000 0.0811
-9.000 -0.8860 0.03628 0.02824 -0.0256 1.0000 0.0843
-8.750 -0.8634 0.03381 0.02526 -0.0260 1.0000 0.0878
-8.500 -0.8386 0.03173 0.02320 -0.0261 1.0000 0.0912
-8.250 -0.8128 0.03017 0.02150 -0.0263 1.0000 0.0964
-8.000 -0.7869 0.02847 0.01976 -0.0265 1.0000 0.1025
-7.750 -0.7597 0.02703 0.01803 -0.0267 1.0000 0.1108
-7.500 -0.7331 0.02562 0.01679 -0.0269 1.0000 0.1209
-7.250 -0.7057 0.02430 0.01549 -0.0272 1.0000 0.1351
-7.000 -0.6781 0.02317 0.01445 -0.0276 1.0000 0.1528
-6.750 -0.6500 0.02247 0.01387 -0.0280 1.0000 0.1721
-6.500 -0.6212 0.02181 0.01318 -0.0284 1.0000 0.1925
-6.250 -0.5922 0.02114 0.01246 -0.0289 1.0000 0.2120
-6.000 -0.5633 0.02054 0.01192 -0.0293 1.0000 0.2303
-5.750 -0.5343 0.02009 0.01158 -0.0296 1.0000 0.2484
-5.500 -0.5050 0.01972 0.01131 -0.0299 1.0000 0.2671
-5.250 -0.4755 0.01934 0.01102 -0.0303 1.0000 0.2858
-5.000 -0.4458 0.01894 0.01067 -0.0306 1.0000 0.3043
-4.750 -0.4159 0.01854 0.01034 -0.0309 1.0000 0.3229
-4.500 -0.3858 0.01816 0.01001 -0.0313 1.0000 0.3417
-4.250 -0.3557 0.01781 0.00973 -0.0318 1.0000 0.3606
-4.000 -0.3254 0.01747 0.00946 -0.0322 1.0000 0.3794
-3.750 -0.2950 0.01716 0.00923 -0.0328 1.0000 0.3984
-3.500 -0.2645 0.01685 0.00902 -0.0334 1.0000 0.4175
-3.250 -0.2343 0.01658 0.00885 -0.0342 1.0000 0.4367
-3.000 -0.2056 0.01636 0.00875 -0.0350 1.0000 0.4556
-2.750 -0.1819 0.01633 0.00885 -0.0355 0.9966 0.4731
-2.500 -0.1268 0.01616 0.00867 -0.0409 0.9708 0.4999
-2.250 -0.0826 0.01609 0.00870 -0.0435 0.9389 0.5225
-2.000 -0.0544 0.01622 0.00883 -0.0427 0.9039 0.5421
-1.750 -0.0322 0.01633 0.00891 -0.0405 0.8731 0.5600
-1.500 -0.0090 0.01636 0.00891 -0.0388 0.8442 0.5777
-1.250 0.0138 0.01637 0.00890 -0.0369 0.8199 0.5951
-1.000 0.0377 0.01634 0.00886 -0.0353 0.7973 0.6126
-0.750 0.0630 0.01630 0.00883 -0.0341 0.7753 0.6309
-0.500 0.0884 0.01628 0.00880 -0.0329 0.7547 0.6511
-0.250 0.1141 0.01625 0.00876 -0.0318 0.7347 0.6728
0.000 0.1391 0.01620 0.00875 -0.0304 0.7153 0.6949
0.250 0.1639 0.01617 0.00874 -0.0290 0.6966 0.7200
0.500 0.1887 0.01615 0.00874 -0.0276 0.6793 0.7499
0.750 0.2116 0.01613 0.00875 -0.0256 0.6640 0.7827
1.000 0.2337 0.01606 0.00876 -0.0235 0.6482 0.8203
1.250 0.2527 0.01590 0.00870 -0.0207 0.6332 0.8648
1.500 0.2713 0.01556 0.00843 -0.0177 0.6192 0.9303
1.750 0.3122 0.01550 0.00825 -0.0206 0.6028 1.0000
2.000 0.3468 0.01580 0.00831 -0.0223 0.5883 1.0000
2.250 0.3796 0.01608 0.00848 -0.0235 0.5724 1.0000
2.500 0.4113 0.01639 0.00869 -0.0243 0.5567 1.0000
2.750 0.4422 0.01672 0.00893 -0.0248 0.5418 1.0000
3.000 0.4724 0.01705 0.00914 -0.0251 0.5275 1.0000
3.250 0.5020 0.01737 0.00929 -0.0252 0.5136 1.0000
3.500 0.5322 0.01767 0.00957 -0.0255 0.4973 1.0000
3.750 0.5618 0.01793 0.00978 -0.0257 0.4803 1.0000
4.000 0.5911 0.01816 0.00993 -0.0257 0.4626 1.0000
4.250 0.6202 0.01839 0.01009 -0.0257 0.4451 1.0000
4.500 0.6493 0.01864 0.01028 -0.0258 0.4280 1.0000
4.750 0.6783 0.01892 0.01050 -0.0258 0.4113 1.0000
5.000 0.7072 0.01920 0.01074 -0.0258 0.3947 1.0000
5.250 0.7359 0.01948 0.01097 -0.0258 0.3778 1.0000
5.500 0.7645 0.01976 0.01121 -0.0258 0.3604 1.0000
5.750 0.7928 0.02005 0.01147 -0.0258 0.3421 1.0000
6.000 0.8209 0.02036 0.01173 -0.0258 0.3226 1.0000
6.250 0.8487 0.02071 0.01199 -0.0257 0.3019 1.0000
6.500 0.8759 0.02119 0.01236 -0.0256 0.2799 1.0000
6.750 0.9029 0.02179 0.01293 -0.0256 0.2551 1.0000
7.000 0.9292 0.02255 0.01350 -0.0255 0.2337 1.0000
7.250 0.9552 0.02355 0.01434 -0.0254 0.2151 1.0000
7.500 0.9811 0.02458 0.01527 -0.0253 0.1995 1.0000
7.750 1.0071 0.02554 0.01627 -0.0252 0.1863 1.0000
8.000 1.0328 0.02664 0.01745 -0.0251 0.1754 1.0000
8.250 1.0583 0.02781 0.01853 -0.0249 0.1670 1.0000
8.500 1.0836 0.02889 0.01969 -0.0247 0.1588 1.0000
8.750 1.1085 0.03025 0.02103 -0.0245 0.1524 1.0000
9.000 1.1328 0.03153 0.02246 -0.0242 0.1460 1.0000
9.250 1.1581 0.03303 0.02375 -0.0241 0.1408 1.0000
9.500 1.1798 0.03462 0.02575 -0.0237 0.1357 1.0000
9.750 1.2028 0.03607 0.02728 -0.0233 0.1311 1.0000
10.000 1.2262 0.03801 0.02913 -0.0232 0.1272 1.0000
10.250 1.2431 0.04012 0.03171 -0.0225 0.1234 1.0000
10.500 1.2619 0.04207 0.03386 -0.0220 0.1198 1.0000
10.750 1.2838 0.04381 0.03555 -0.0217 0.1165 1.0000
11.000 1.2962 0.04679 0.03883 -0.0211 0.1139 1.0000
11.250 1.3000 0.05008 0.04264 -0.0201 0.1116 1.0000
11.500 1.3025 0.05336 0.04629 -0.0191 0.1094 1.0000
11.750 1.3103 0.05601 0.04911 -0.0184 0.1071 1.0000
12.000 1.3293 0.05798 0.05101 -0.0180 0.1045 1.0000
12.250 1.3280 0.06220 0.05541 -0.0173 0.1031 1.0000
12.500 1.3017 0.06700 0.06063 -0.0161 0.1028 1.0000
12.750 1.2676 0.07228 0.06620 -0.0157 0.1028 1.0000
13.000 1.2296 0.07950 0.07367 -0.0184 0.1031 1.0000
|
Polar data table (+)
Polar graphs
<< Back to OAF128 AIRFOIL (oaf128-il)