DF 101 AIRFOIL (df101-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: DF 101 AIRFOIL (df101-il) Reynolds number: 50,000 Max Cl/Cd: 34.8 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-df101-il-50000-n5.txt Download as CSV file: xf-df101-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: DF 101 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4791 0.08926 0.08219 -0.0417 1.0000 0.0441
-9.750 -0.4863 0.08363 0.07663 -0.0445 1.0000 0.0438
-9.500 -0.4969 0.07758 0.07063 -0.0480 1.0000 0.0435
-9.250 -0.5111 0.07174 0.06483 -0.0520 1.0000 0.0431
-9.000 -0.5329 0.06687 0.06002 -0.0539 1.0000 0.0427
-8.750 -0.5519 0.06272 0.05582 -0.0540 1.0000 0.0425
-8.500 -0.5673 0.05883 0.05180 -0.0533 1.0000 0.0423
-8.250 -0.5800 0.05502 0.04780 -0.0519 1.0000 0.0422
-8.000 -0.5886 0.05137 0.04386 -0.0501 1.0000 0.0424
-7.750 -0.5928 0.04790 0.04004 -0.0480 1.0000 0.0427
-7.500 -0.5924 0.04470 0.03643 -0.0458 1.0000 0.0431
-7.250 -0.5882 0.04175 0.03296 -0.0435 1.0000 0.0439
-7.000 -0.5801 0.03941 0.03050 -0.0416 1.0000 0.0458
-6.750 -0.5695 0.03761 0.02855 -0.0397 1.0000 0.0485
-6.500 -0.5574 0.03563 0.02627 -0.0377 1.0000 0.0515
-6.250 -0.5430 0.03358 0.02381 -0.0358 1.0000 0.0540
-6.000 -0.5282 0.03178 0.02185 -0.0339 1.0000 0.0572
-5.750 -0.5131 0.03045 0.02043 -0.0323 1.0000 0.0628
-5.500 -0.4963 0.02901 0.01871 -0.0305 1.0000 0.0684
-5.250 -0.4802 0.02774 0.01742 -0.0288 1.0000 0.0752
-5.000 -0.4632 0.02657 0.01621 -0.0273 1.0000 0.0865
-4.750 -0.4454 0.02541 0.01503 -0.0259 1.0000 0.1022
-4.500 -0.4269 0.02432 0.01413 -0.0247 1.0000 0.1356
-4.250 -0.4080 0.02364 0.01361 -0.0237 1.0000 0.1907
-4.000 -0.3899 0.02337 0.01342 -0.0226 1.0000 0.2410
-3.750 -0.3649 0.02328 0.01332 -0.0229 0.9963 0.2913
-3.500 -0.3305 0.02312 0.01312 -0.0247 0.9883 0.3340
-3.250 -0.2944 0.02288 0.01277 -0.0266 0.9808 0.3673
-3.000 -0.2589 0.02262 0.01242 -0.0284 0.9729 0.3979
-2.750 -0.2243 0.02235 0.01214 -0.0300 0.9648 0.4303
-2.500 -0.1882 0.02209 0.01193 -0.0319 0.9572 0.4676
-2.250 -0.1570 0.02180 0.01174 -0.0327 0.9481 0.5113
-2.000 -0.1220 0.02150 0.01164 -0.0340 0.9406 0.5718
-1.750 -0.0935 0.02117 0.01161 -0.0337 0.9312 0.6501
-1.500 -0.0540 0.02089 0.01172 -0.0348 0.9249 0.7731
-1.250 0.0233 0.02092 0.01175 -0.0435 0.9217 0.9424
-1.000 0.0872 0.02101 0.01154 -0.0511 0.9163 1.0000
-0.750 0.1202 0.02112 0.01143 -0.0526 0.9054 1.0000
-0.500 0.1620 0.02120 0.01129 -0.0556 0.8974 1.0000
-0.250 0.1904 0.02132 0.01125 -0.0560 0.8851 1.0000
0.000 0.2194 0.02143 0.01124 -0.0564 0.8731 1.0000
0.250 0.2499 0.02153 0.01122 -0.0570 0.8618 1.0000
0.500 0.2846 0.02156 0.01115 -0.0582 0.8518 1.0000
0.750 0.3137 0.02163 0.01114 -0.0584 0.8396 1.0000
1.000 0.3406 0.02173 0.01117 -0.0581 0.8265 1.0000
1.250 0.3681 0.02179 0.01118 -0.0578 0.8134 1.0000
1.500 0.3959 0.02182 0.01117 -0.0575 0.8000 1.0000
1.750 0.4239 0.02182 0.01115 -0.0570 0.7862 1.0000
2.000 0.4516 0.02180 0.01110 -0.0565 0.7718 1.0000
2.250 0.4789 0.02175 0.01104 -0.0558 0.7568 1.0000
2.500 0.5059 0.02170 0.01098 -0.0550 0.7411 1.0000
2.750 0.5328 0.02164 0.01094 -0.0541 0.7250 1.0000
3.000 0.5596 0.02159 0.01087 -0.0532 0.7085 1.0000
3.250 0.5867 0.02153 0.01081 -0.0522 0.6917 1.0000
3.500 0.6094 0.02163 0.01095 -0.0508 0.6724 1.0000
3.750 0.6336 0.02169 0.01101 -0.0496 0.6529 1.0000
4.000 0.6594 0.02171 0.01102 -0.0484 0.6335 1.0000
4.250 0.6817 0.02187 0.01121 -0.0470 0.6114 1.0000
4.500 0.7059 0.02198 0.01133 -0.0457 0.5894 1.0000
4.750 0.7275 0.02221 0.01157 -0.0441 0.5651 1.0000
5.000 0.7499 0.02243 0.01178 -0.0427 0.5404 1.0000
5.250 0.7720 0.02270 0.01201 -0.0412 0.5150 1.0000
5.500 0.7924 0.02307 0.01240 -0.0396 0.4880 1.0000
5.750 0.8128 0.02349 0.01279 -0.0380 0.4613 1.0000
6.000 0.8329 0.02397 0.01323 -0.0365 0.4353 1.0000
6.250 0.8532 0.02452 0.01369 -0.0351 0.4112 1.0000
6.500 0.8725 0.02515 0.01435 -0.0336 0.3875 1.0000
6.750 0.8925 0.02583 0.01497 -0.0323 0.3665 1.0000
7.000 0.9121 0.02658 0.01575 -0.0310 0.3464 1.0000
7.250 0.9319 0.02736 0.01653 -0.0297 0.3283 1.0000
7.500 0.9521 0.02819 0.01740 -0.0286 0.3120 1.0000
7.750 0.9720 0.02905 0.01829 -0.0275 0.2965 1.0000
8.000 0.9919 0.02997 0.01927 -0.0263 0.2820 1.0000
8.250 1.0110 0.03091 0.02031 -0.0252 0.2682 1.0000
8.500 1.0297 0.03191 0.02143 -0.0240 0.2550 1.0000
8.750 1.0478 0.03296 0.02264 -0.0227 0.2423 1.0000
9.000 1.0649 0.03404 0.02385 -0.0214 0.2299 1.0000
9.250 1.0813 0.03514 0.02508 -0.0200 0.2179 1.0000
9.500 1.0957 0.03625 0.02628 -0.0184 0.2057 1.0000
9.750 1.1088 0.03734 0.02745 -0.0167 0.1937 1.0000
10.000 1.1170 0.03852 0.02882 -0.0145 0.1811 1.0000
10.250 1.1222 0.03979 0.03027 -0.0121 0.1688 1.0000
10.500 1.1237 0.04109 0.03176 -0.0093 0.1570 1.0000
10.750 1.1231 0.04250 0.03329 -0.0065 0.1456 1.0000
11.000 1.1211 0.04405 0.03491 -0.0040 0.1348 1.0000
11.250 1.1178 0.04577 0.03663 -0.0019 0.1245 1.0000
11.500 1.1132 0.04800 0.03905 -0.0002 0.1139 1.0000
11.750 1.1080 0.05056 0.04175 0.0011 0.1041 1.0000
12.000 1.1030 0.05326 0.04450 0.0020 0.0962 1.0000
12.250 1.0968 0.05631 0.04764 0.0024 0.0887 1.0000
12.500 1.0906 0.05979 0.05125 0.0026 0.0823 1.0000
12.750 1.0842 0.06336 0.05491 0.0023 0.0769 1.0000
13.000 1.0770 0.06728 0.05891 0.0018 0.0724 1.0000
13.250 1.0676 0.07192 0.06380 0.0007 0.0685 1.0000
13.500 1.0600 0.07625 0.06823 -0.0005 0.0654 1.0000
13.750 1.0554 0.08022 0.07216 -0.0014 0.0627 1.0000
14.000 1.0412 0.08640 0.07865 -0.0038 0.0612 1.0000
14.250 1.0237 0.09349 0.08602 -0.0070 0.0602 1.0000
14.500 1.0021 0.10183 0.09460 -0.0113 0.0597 1.0000
14.750 0.9741 0.11225 0.10521 -0.0171 0.0599 1.0000
15.000 0.9400 0.12518 0.11826 -0.0244 0.0605 1.0000
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Polar data table (+)
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