GIII BL207 AIRFOIL (giiih-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GIII BL207 AIRFOIL (giiih-il) Reynolds number: 50,000 Max Cl/Cd: 32.7 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-giiih-il-50000-n5.txt Download as CSV file: xf-giiih-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GIII BL207 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4463 0.10610 0.09953 -0.0106 1.0000 0.1234
-9.500 -0.4642 0.10310 0.09663 -0.0140 1.0000 0.1276
-9.250 -0.4616 0.09847 0.09204 -0.0143 1.0000 0.1298
-9.000 -0.4550 0.09417 0.08775 -0.0141 1.0000 0.1320
-8.750 -0.4556 0.09002 0.08364 -0.0150 1.0000 0.1335
-8.500 -0.4599 0.08569 0.07936 -0.0166 1.0000 0.1336
-8.000 -0.5791 0.08101 0.07412 -0.0254 1.0000 0.0576
-7.500 -0.5688 0.07235 0.06534 -0.0262 1.0000 0.0557
-7.250 -0.5638 0.06806 0.06096 -0.0267 1.0000 0.0541
-7.000 -0.5580 0.06375 0.05649 -0.0272 1.0000 0.0526
-6.750 -0.5508 0.05948 0.05200 -0.0273 1.0000 0.0512
-6.500 -0.5419 0.05533 0.04757 -0.0270 1.0000 0.0501
-6.250 -0.5312 0.05137 0.04327 -0.0263 1.0000 0.0493
-6.000 -0.5189 0.04767 0.03919 -0.0253 1.0000 0.0489
-5.750 -0.5049 0.04461 0.03579 -0.0241 1.0000 0.0503
-5.500 -0.4897 0.04184 0.03264 -0.0228 1.0000 0.0530
-5.250 -0.4732 0.03905 0.02937 -0.0213 1.0000 0.0549
-5.000 -0.4551 0.03633 0.02619 -0.0197 1.0000 0.0555
-4.750 -0.4353 0.03382 0.02321 -0.0182 1.0000 0.0563
-4.500 -0.4138 0.03160 0.02046 -0.0167 1.0000 0.0575
-4.250 -0.3927 0.02950 0.01815 -0.0157 1.0000 0.0606
-4.000 -0.3714 0.02806 0.01657 -0.0146 1.0000 0.0666
-3.750 -0.3471 0.02651 0.01469 -0.0134 1.0000 0.0708
-3.500 -0.3223 0.02498 0.01307 -0.0124 1.0000 0.0750
-3.250 -0.2993 0.02405 0.01196 -0.0113 1.0000 0.0854
-3.000 -0.2767 0.02294 0.01081 -0.0103 1.0000 0.0935
-2.750 -0.2558 0.02203 0.00976 -0.0090 1.0000 0.1021
-2.500 -0.2373 0.02125 0.00901 -0.0077 1.0000 0.1196
-2.250 -0.2183 0.02046 0.00822 -0.0064 1.0000 0.1423
-2.000 -0.0587 0.01733 0.00771 -0.0247 1.0000 1.0000
-1.750 -0.0536 0.01724 0.00750 -0.0215 1.0000 1.0000
-1.500 -0.0510 0.01720 0.00734 -0.0179 1.0000 1.0000
-1.250 -0.0495 0.01720 0.00721 -0.0142 1.0000 1.0000
-1.000 -0.0470 0.01723 0.00709 -0.0106 1.0000 1.0000
-0.750 -0.0423 0.01729 0.00701 -0.0074 1.0000 1.0000
-0.500 -0.0343 0.01739 0.00696 -0.0048 1.0000 1.0000
-0.250 -0.0056 0.01759 0.00699 -0.0060 0.9939 1.0000
0.000 0.0316 0.01783 0.00708 -0.0089 0.9847 1.0000
0.250 0.0693 0.01808 0.00720 -0.0117 0.9757 1.0000
0.500 0.1056 0.01833 0.00737 -0.0142 0.9658 1.0000
0.750 0.1415 0.01857 0.00757 -0.0166 0.9554 1.0000
1.000 0.1787 0.01882 0.00780 -0.0191 0.9449 1.0000
1.250 0.2192 0.01905 0.00804 -0.0221 0.9340 1.0000
1.500 0.2623 0.01924 0.00828 -0.0255 0.9227 1.0000
1.750 0.3012 0.01940 0.00853 -0.0278 0.9098 1.0000
2.000 0.3365 0.01955 0.00877 -0.0294 0.8962 1.0000
2.250 0.3701 0.01970 0.00903 -0.0305 0.8823 1.0000
2.500 0.4025 0.01984 0.00933 -0.0313 0.8680 1.0000
2.750 0.4334 0.01995 0.00960 -0.0316 0.8529 1.0000
3.000 0.4604 0.02010 0.00990 -0.0312 0.8357 1.0000
3.250 0.4876 0.02021 0.01018 -0.0306 0.8179 1.0000
3.500 0.5155 0.02027 0.01048 -0.0299 0.7997 1.0000
3.750 0.5419 0.02030 0.01073 -0.0289 0.7795 1.0000
4.000 0.5671 0.02027 0.01091 -0.0274 0.7557 1.0000
4.250 0.5887 0.01999 0.01079 -0.0244 0.7177 1.0000
4.500 0.6078 0.01952 0.01039 -0.0205 0.6579 1.0000
4.750 0.6261 0.01936 0.01013 -0.0169 0.5799 1.0000
5.000 0.6429 0.01966 0.01002 -0.0134 0.4652 1.0000
5.250 0.6542 0.02096 0.01043 -0.0103 0.3182 1.0000
5.500 0.6644 0.02297 0.01157 -0.0081 0.1843 1.0000
5.750 0.6787 0.02488 0.01297 -0.0066 0.1285 1.0000
6.000 0.6958 0.02651 0.01452 -0.0052 0.1060 1.0000
6.250 0.7150 0.02808 0.01612 -0.0039 0.0927 1.0000
6.500 0.7353 0.02958 0.01765 -0.0029 0.0800 1.0000
6.750 0.7592 0.03117 0.01948 -0.0018 0.0722 1.0000
7.000 0.7840 0.03326 0.02156 -0.0011 0.0670 1.0000
7.250 0.8081 0.03512 0.02382 -0.0003 0.0602 1.0000
7.500 0.8299 0.03736 0.02618 0.0003 0.0554 1.0000
7.750 0.8522 0.04030 0.02956 0.0013 0.0533 1.0000
8.000 0.8706 0.04347 0.03331 0.0026 0.0513 1.0000
8.250 0.8845 0.04651 0.03687 0.0039 0.0482 1.0000
8.500 0.8972 0.04916 0.03977 0.0050 0.0449 1.0000
8.750 0.9072 0.05251 0.04336 0.0061 0.0432 1.0000
9.000 0.9129 0.05672 0.04789 0.0073 0.0424 1.0000
9.250 0.9128 0.06084 0.05242 0.0088 0.0421 1.0000
9.500 0.9085 0.06502 0.05700 0.0101 0.0420 1.0000
9.750 0.9003 0.06924 0.06154 0.0112 0.0419 1.0000
10.000 0.8879 0.07340 0.06594 0.0120 0.0419 1.0000
10.250 0.8722 0.07756 0.07027 0.0124 0.0420 1.0000
10.500 0.8562 0.08220 0.07503 0.0114 0.0421 1.0000
10.750 0.8408 0.08740 0.08033 0.0092 0.0423 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GIII BL207 AIRFOIL (giiih-il)