WACO COOTIE AIRFOIL (cootie-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: WACO COOTIE AIRFOIL (cootie-il) Reynolds number: 50,000 Max Cl/Cd: 39.75 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-cootie-il-50000-n5.txt Download as CSV file: xf-cootie-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: WACO COOTIE 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3145 0.11368 0.10731 -0.0332 1.0000 0.0866
-8.500 -0.3279 0.11331 0.10709 -0.0325 1.0000 0.0868
-8.250 -0.3385 0.11267 0.10656 -0.0328 1.0000 0.0870
-8.000 -0.3199 0.10488 0.09877 -0.0275 1.0000 0.0895
-7.750 -0.3239 0.10271 0.09667 -0.0255 1.0000 0.0908
-7.500 -0.3295 0.10082 0.09487 -0.0239 1.0000 0.0920
-7.250 -0.3352 0.09898 0.09311 -0.0227 1.0000 0.0934
-7.000 -0.3401 0.09712 0.09132 -0.0219 1.0000 0.0950
-6.750 -0.3438 0.09520 0.08947 -0.0218 1.0000 0.0967
-6.500 -0.3149 0.09225 0.08641 -0.0352 0.9912 0.1008
-6.000 -0.2738 0.08281 0.07691 -0.0394 0.9787 0.1078
-5.750 -0.2389 0.07906 0.07299 -0.0495 0.9698 0.1139
-5.500 -0.2006 0.07467 0.06841 -0.0592 0.9622 0.1157
-5.250 -0.1840 0.07052 0.06430 -0.0586 0.9550 0.1195
-4.750 -0.1198 0.06316 0.05664 -0.0692 0.9398 0.1349
-4.500 -0.0564 0.05644 0.04922 -0.0818 0.9341 0.0859
-4.250 -0.0277 0.05300 0.04566 -0.0845 0.9258 0.0831
-4.000 0.0178 0.04933 0.04161 -0.0907 0.9208 0.0840
-3.750 0.0502 0.04637 0.03838 -0.0935 0.9118 0.0818
-3.500 0.0951 0.04300 0.03457 -0.0985 0.9068 0.0791
-3.250 0.1309 0.04040 0.03155 -0.1013 0.8985 0.0776
-3.000 0.1734 0.03787 0.02853 -0.1049 0.8929 0.0771
-2.750 0.2068 0.03632 0.02666 -0.1066 0.8846 0.0802
-2.500 0.2459 0.03462 0.02449 -0.1090 0.8781 0.0834
-2.250 0.2789 0.03317 0.02264 -0.1101 0.8696 0.0841
-2.000 0.3167 0.03171 0.02076 -0.1118 0.8631 0.0852
-1.750 0.3475 0.03071 0.01935 -0.1122 0.8538 0.0868
-1.500 0.3836 0.02959 0.01807 -0.1136 0.8475 0.0896
-1.250 0.4107 0.02899 0.01730 -0.1133 0.8372 0.0935
-1.000 0.4478 0.02821 0.01618 -0.1144 0.8309 0.1025
-0.750 0.4751 0.02782 0.01565 -0.1141 0.8203 0.1152
-0.500 0.5040 0.02718 0.01496 -0.1140 0.8107 0.1309
-0.250 0.5367 0.02641 0.01417 -0.1145 0.8028 0.1605
0.000 0.5618 0.02598 0.01392 -0.1142 0.7912 0.2060
0.250 0.5895 0.02552 0.01364 -0.1140 0.7806 0.2484
0.750 0.6427 0.02370 0.01318 -0.1125 0.7600 1.0000
1.000 0.6684 0.02389 0.01313 -0.1118 0.7486 1.0000
1.250 0.6966 0.02396 0.01300 -0.1114 0.7387 1.0000
1.500 0.7243 0.02404 0.01291 -0.1110 0.7286 1.0000
1.750 0.7488 0.02429 0.01304 -0.1103 0.7169 1.0000
2.000 0.7749 0.02446 0.01312 -0.1097 0.7061 1.0000
2.250 0.8047 0.02442 0.01298 -0.1095 0.6973 1.0000
2.500 0.8294 0.02467 0.01318 -0.1088 0.6854 1.0000
2.750 0.8545 0.02492 0.01339 -0.1082 0.6740 1.0000
3.000 0.8814 0.02508 0.01353 -0.1077 0.6634 1.0000
3.250 0.9105 0.02510 0.01350 -0.1075 0.6537 1.0000
3.500 0.9347 0.02546 0.01387 -0.1068 0.6414 1.0000
3.750 0.9596 0.02579 0.01424 -0.1062 0.6295 1.0000
4.000 0.9857 0.02606 0.01453 -0.1057 0.6181 1.0000
4.250 1.0142 0.02622 0.01468 -0.1054 0.6078 1.0000
4.500 1.0374 0.02671 0.01527 -0.1046 0.5951 1.0000
4.750 1.0605 0.02722 0.01586 -0.1039 0.5826 1.0000
5.000 1.0843 0.02771 0.01644 -0.1032 0.5706 1.0000
5.250 1.1095 0.02812 0.01696 -0.1027 0.5595 1.0000
5.500 1.1344 0.02855 0.01749 -0.1021 0.5482 1.0000
5.750 1.1553 0.02924 0.01834 -0.1012 0.5359 1.0000
6.000 1.1770 0.02989 0.01914 -0.1004 0.5242 1.0000
6.250 1.2005 0.03042 0.01984 -0.0997 0.5133 1.0000
6.500 1.2254 0.03083 0.02038 -0.0990 0.5027 1.0000
6.750 1.2448 0.03150 0.02123 -0.0978 0.4897 1.0000
7.000 1.2644 0.03208 0.02198 -0.0965 0.4763 1.0000
7.250 1.2838 0.03259 0.02266 -0.0951 0.4624 1.0000
7.500 1.3026 0.03308 0.02335 -0.0936 0.4479 1.0000
7.750 1.3204 0.03354 0.02398 -0.0919 0.4325 1.0000
8.000 1.3362 0.03375 0.02423 -0.0897 0.4125 1.0000
8.250 1.3439 0.03444 0.02505 -0.0867 0.3905 1.0000
8.500 1.3522 0.03500 0.02560 -0.0837 0.3675 1.0000
8.750 1.3541 0.03603 0.02670 -0.0803 0.3432 1.0000
9.000 1.3531 0.03718 0.02792 -0.0766 0.3195 1.0000
9.250 1.3545 0.03859 0.02946 -0.0737 0.2992 1.0000
9.500 1.3563 0.04015 0.03116 -0.0711 0.2798 1.0000
9.750 1.3537 0.04219 0.03335 -0.0687 0.2542 1.0000
10.000 1.3466 0.04473 0.03587 -0.0664 0.2173 1.0000
10.250 1.3352 0.04793 0.03879 -0.0644 0.1757 1.0000
10.500 1.3208 0.05187 0.04242 -0.0628 0.1391 1.0000
10.750 1.3072 0.05613 0.04650 -0.0618 0.1104 1.0000
11.000 1.2941 0.06063 0.05091 -0.0612 0.0935 1.0000
11.250 1.2813 0.06534 0.05561 -0.0610 0.0819 1.0000
11.500 1.2679 0.07030 0.06058 -0.0610 0.0730 1.0000
11.750 1.2584 0.07492 0.06535 -0.0611 0.0640 1.0000
12.000 1.2470 0.07988 0.07035 -0.0616 0.0585 1.0000
12.250 1.2394 0.08443 0.07505 -0.0619 0.0529 1.0000
12.500 1.2310 0.08919 0.07988 -0.0626 0.0492 1.0000
12.750 1.2244 0.09373 0.08448 -0.0632 0.0461 1.0000
13.000 1.2201 0.09797 0.08886 -0.0638 0.0431 1.0000
13.250 1.2160 0.10222 0.09325 -0.0645 0.0408 1.0000
|
Polar data table (+)
Polar graphs
<< Back to WACO COOTIE AIRFOIL (cootie-il)