WACO COOTIE AIRFOIL (cootie-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: WACO COOTIE AIRFOIL (cootie-il) Reynolds number: 1,000,000 Max Cl/Cd: 136.88 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-cootie-il-1000000.txt Download as CSV file: xf-cootie-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.2597 0.08848 0.08678 -0.0431 0.9267 0.0144
-8.500 -0.2522 0.08603 0.08429 -0.0442 0.9137 0.0147
-8.250 -0.2446 0.08336 0.08158 -0.0457 0.9031 0.0151
-8.000 -0.2370 0.08052 0.07870 -0.0479 0.8942 0.0158
-7.750 -0.2204 0.07510 0.07323 -0.0574 0.8850 0.0173
-7.500 -0.2031 0.07048 0.06857 -0.0634 0.8783 0.0174
-7.250 -0.1844 0.06569 0.06373 -0.0690 0.8717 0.0174
-6.250 -0.0714 0.03532 0.03320 -0.0813 0.8325 0.0220
-6.000 -0.0482 0.03028 0.02806 -0.0862 0.8272 0.0221
-5.750 -0.0282 0.02130 0.01884 -0.0932 0.8230 0.0227
-5.500 -0.0059 0.01969 0.01718 -0.0942 0.8167 0.0232
-5.250 0.0172 0.01840 0.01583 -0.0949 0.8101 0.0239
-5.000 0.0423 0.01654 0.01387 -0.0964 0.8036 0.0253
-4.750 0.0751 0.01384 0.01093 -0.0985 0.7970 0.0282
-3.750 0.1857 0.01268 0.00777 -0.1078 0.7804 0.0292
-3.500 0.2141 0.01196 0.00687 -0.1078 0.7684 0.0289
-3.250 0.2422 0.01085 0.00554 -0.1080 0.7566 0.0293
-3.000 0.2704 0.01009 0.00463 -0.1080 0.7439 0.0297
-2.750 0.2986 0.00958 0.00400 -0.1081 0.7286 0.0301
-2.500 0.3267 0.00920 0.00350 -0.1081 0.7112 0.0305
-2.250 0.3547 0.00891 0.00310 -0.1081 0.6935 0.0308
-2.000 0.3826 0.00869 0.00277 -0.1081 0.6765 0.0310
-1.750 0.4107 0.00850 0.00248 -0.1081 0.6616 0.0312
-1.500 0.4389 0.00834 0.00224 -0.1081 0.6489 0.0315
-1.250 0.4671 0.00822 0.00206 -0.1082 0.6386 0.0323
-1.000 0.4954 0.00812 0.00190 -0.1083 0.6294 0.0330
-0.750 0.5239 0.00802 0.00175 -0.1084 0.6208 0.0337
-0.500 0.5522 0.00796 0.00165 -0.1085 0.6128 0.0347
-0.250 0.5806 0.00792 0.00157 -0.1087 0.6046 0.0356
0.000 0.6090 0.00790 0.00152 -0.1088 0.5973 0.0362
0.250 0.6375 0.00789 0.00147 -0.1089 0.5900 0.0379
0.500 0.6658 0.00789 0.00145 -0.1090 0.5832 0.0416
0.750 0.6942 0.00782 0.00151 -0.1092 0.5758 0.0776
1.000 0.7223 0.00788 0.00158 -0.1093 0.5684 0.0926
1.250 0.7506 0.00791 0.00162 -0.1095 0.5601 0.0991
1.500 0.7787 0.00796 0.00168 -0.1096 0.5520 0.1056
1.750 0.8066 0.00803 0.00172 -0.1097 0.5418 0.1090
2.000 0.8346 0.00806 0.00175 -0.1098 0.5296 0.1162
2.250 0.8624 0.00813 0.00181 -0.1099 0.5167 0.1227
2.750 0.9119 0.00668 0.00213 -0.1093 0.4877 1.0000
3.000 0.9390 0.00686 0.00222 -0.1093 0.4700 1.0000
3.250 0.9659 0.00707 0.00232 -0.1093 0.4515 1.0000
3.500 0.9928 0.00729 0.00245 -0.1093 0.4312 1.0000
3.750 1.0189 0.00758 0.00261 -0.1092 0.4069 1.0000
4.000 1.0448 0.00789 0.00279 -0.1090 0.3807 1.0000
4.250 1.0708 0.00817 0.00298 -0.1089 0.3604 1.0000
4.500 1.0970 0.00843 0.00317 -0.1088 0.3443 1.0000
4.750 1.1230 0.00870 0.00336 -0.1087 0.3276 1.0000
5.000 1.1486 0.00901 0.00358 -0.1085 0.3086 1.0000
5.250 1.1742 0.00930 0.00381 -0.1084 0.2920 1.0000
5.500 1.1996 0.00961 0.00404 -0.1082 0.2732 1.0000
5.750 1.2235 0.01006 0.00433 -0.1078 0.2419 1.0000
6.000 1.2407 0.01119 0.00499 -0.1064 0.1646 1.0000
6.250 1.2546 0.01264 0.00594 -0.1046 0.0817 1.0000
6.500 1.2754 0.01334 0.00649 -0.1036 0.0547 1.0000
6.750 1.2923 0.01440 0.00730 -0.1021 0.0174 1.0000
7.000 1.3144 0.01490 0.00783 -0.1013 0.0149 1.0000
7.250 1.3366 0.01537 0.00836 -0.1006 0.0137 1.0000
7.500 1.3584 0.01584 0.00887 -0.0998 0.0129 1.0000
7.750 1.3792 0.01637 0.00945 -0.0989 0.0121 1.0000
8.000 1.3985 0.01699 0.01012 -0.0977 0.0114 1.0000
8.250 1.4144 0.01786 0.01107 -0.0960 0.0107 1.0000
8.500 1.4286 0.01877 0.01208 -0.0940 0.0102 1.0000
8.750 1.4459 0.01936 0.01272 -0.0925 0.0099 1.0000
9.000 1.4611 0.01998 0.01340 -0.0907 0.0096 1.0000
9.250 1.4731 0.02069 0.01415 -0.0884 0.0092 1.0000
9.500 1.4843 0.02148 0.01500 -0.0861 0.0088 1.0000
9.750 1.4941 0.02242 0.01601 -0.0837 0.0085 1.0000
10.000 1.5029 0.02347 0.01713 -0.0814 0.0082 1.0000
10.250 1.5090 0.02478 0.01851 -0.0790 0.0079 1.0000
10.500 1.5099 0.02655 0.02036 -0.0762 0.0077 1.0000
10.750 1.5018 0.02918 0.02311 -0.0728 0.0075 1.0000
11.000 1.4992 0.03153 0.02556 -0.0704 0.0073 1.0000
11.250 1.5081 0.03296 0.02707 -0.0691 0.0072 1.0000
11.500 1.5139 0.03472 0.02891 -0.0678 0.0071 1.0000
11.750 1.5185 0.03666 0.03094 -0.0664 0.0070 1.0000
12.000 1.5217 0.03878 0.03315 -0.0652 0.0069 1.0000
12.250 1.5233 0.04112 0.03558 -0.0640 0.0068 1.0000
12.500 1.5247 0.04356 0.03811 -0.0629 0.0066 1.0000
12.750 1.5252 0.04614 0.04078 -0.0619 0.0065 1.0000
13.000 1.5256 0.04882 0.04355 -0.0612 0.0064 1.0000
13.250 1.5259 0.05159 0.04641 -0.0606 0.0063 1.0000
13.500 1.5264 0.05442 0.04933 -0.0602 0.0061 1.0000
13.750 1.5269 0.05735 0.05234 -0.0600 0.0060 1.0000
14.000 1.5269 0.06043 0.05552 -0.0599 0.0059 1.0000
14.250 1.5263 0.06365 0.05883 -0.0600 0.0057 1.0000
14.500 1.5251 0.06700 0.06225 -0.0602 0.0056 1.0000
14.750 1.5224 0.07057 0.06590 -0.0605 0.0055 1.0000
15.000 1.5184 0.07433 0.06975 -0.0607 0.0054 1.0000
15.250 1.5130 0.07827 0.07377 -0.0609 0.0053 1.0000
15.500 1.5060 0.08231 0.07790 -0.0608 0.0053 1.0000
15.750 1.4953 0.08651 0.08223 -0.0599 0.0051 1.0000
|
Polar data table (+)
Polar graphs
<< Back to WACO COOTIE AIRFOIL (cootie-il)