GIII BL207 AIRFOIL (giiih-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: GIII BL207 AIRFOIL (giiih-il) Reynolds number: 50,000 Max Cl/Cd: 33.4 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-giiih-il-50000.txt Download as CSV file: xf-giiih-il-50000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5613 0.10557 0.09886 0.0068 1.0000 0.2445
-8.500 -0.5614 0.10230 0.09565 0.0066 1.0000 0.2606
-8.250 -0.5667 0.09934 0.09278 0.0060 1.0000 0.2769
-8.000 -0.5529 0.09514 0.08859 0.0079 1.0000 0.2981
-7.750 -0.5617 0.09262 0.08618 0.0079 1.0000 0.3198
-7.500 -0.5558 0.08934 0.08294 0.0101 1.0000 0.3482
-7.250 -0.5344 0.08562 0.07916 0.0141 1.0000 0.3878
-7.000 -0.5318 0.08334 0.07695 0.0179 1.0000 0.4329
-6.500 -0.4636 0.07764 0.07112 0.0310 1.0000 0.6019
-5.000 -0.4840 0.04363 0.03525 -0.0204 1.0000 0.1769
-4.750 -0.4626 0.03997 0.03112 -0.0195 1.0000 0.1584
-4.500 -0.4403 0.03683 0.02739 -0.0182 1.0000 0.1440
-4.250 -0.4173 0.03428 0.02412 -0.0166 1.0000 0.1351
-4.000 -0.3948 0.03171 0.02121 -0.0153 1.0000 0.1327
-3.750 -0.3719 0.02992 0.01887 -0.0138 1.0000 0.1364
-3.500 -0.3486 0.02767 0.01659 -0.0128 1.0000 0.1424
-3.250 -0.3226 0.02591 0.01450 -0.0117 1.0000 0.1467
-3.000 -0.2967 0.02435 0.01287 -0.0107 1.0000 0.1604
-2.750 -0.0952 0.01786 0.00900 -0.0307 1.0000 1.0000
-2.500 -0.0796 0.01764 0.00841 -0.0292 1.0000 1.0000
-2.250 -0.0669 0.01746 0.00799 -0.0272 1.0000 1.0000
-2.000 -0.0576 0.01732 0.00768 -0.0247 1.0000 1.0000
-1.750 -0.0518 0.01723 0.00746 -0.0217 1.0000 1.0000
-1.500 -0.0486 0.01718 0.00730 -0.0182 1.0000 1.0000
-1.250 -0.0470 0.01717 0.00715 -0.0144 1.0000 1.0000
-1.000 -0.0448 0.01720 0.00705 -0.0108 1.0000 1.0000
-0.750 -0.0406 0.01725 0.00697 -0.0075 1.0000 1.0000
-0.500 -0.0334 0.01735 0.00693 -0.0047 1.0000 1.0000
-0.250 -0.0227 0.01749 0.00693 -0.0026 1.0000 1.0000
0.000 -0.0094 0.01767 0.00698 -0.0010 1.0000 1.0000
0.250 0.0058 0.01789 0.00707 0.0003 1.0000 1.0000
0.500 0.0223 0.01815 0.00723 0.0013 1.0000 1.0000
0.750 0.0397 0.01845 0.00745 0.0021 1.0000 1.0000
1.000 0.0578 0.01879 0.00772 0.0027 1.0000 1.0000
1.250 0.0763 0.01917 0.00806 0.0032 1.0000 1.0000
1.500 0.0950 0.01960 0.00845 0.0036 1.0000 1.0000
1.750 0.1140 0.02007 0.00890 0.0038 1.0000 1.0000
2.000 0.1329 0.02058 0.00942 0.0040 1.0000 1.0000
2.250 0.1519 0.02115 0.01001 0.0041 1.0000 1.0000
2.500 0.1707 0.02178 0.01068 0.0041 1.0000 1.0000
2.750 0.1892 0.02247 0.01144 0.0040 1.0000 1.0000
3.000 0.2127 0.02331 0.01237 0.0027 0.9975 1.0000
3.250 0.2765 0.02471 0.01400 -0.0061 0.9754 1.0000
3.500 0.3385 0.02596 0.01556 -0.0141 0.9511 1.0000
3.750 0.3992 0.02697 0.01692 -0.0212 0.9243 1.0000
4.000 0.4557 0.02769 0.01803 -0.0268 0.8942 1.0000
4.250 0.5195 0.02795 0.01885 -0.0324 0.8599 1.0000
4.500 0.6044 0.02551 0.01714 -0.0351 0.8020 1.0000
4.750 0.6453 0.02286 0.01493 -0.0295 0.7420 1.0000
5.000 0.6654 0.02108 0.01334 -0.0223 0.6714 1.0000
5.250 0.6706 0.02008 0.01163 -0.0122 0.4700 1.0000
5.500 0.6681 0.02371 0.01301 -0.0063 0.2440 1.0000
5.750 0.6880 0.02607 0.01481 -0.0046 0.1866 1.0000
6.000 0.7153 0.02835 0.01677 -0.0037 0.1572 1.0000
6.250 0.7440 0.03064 0.01927 -0.0030 0.1398 1.0000
6.500 0.7725 0.03345 0.02209 -0.0025 0.1294 1.0000
6.750 0.7958 0.03582 0.02482 -0.0014 0.1192 1.0000
7.000 0.8192 0.03926 0.02857 -0.0005 0.1155 1.0000
7.250 0.8391 0.04279 0.03269 0.0010 0.1148 1.0000
7.500 0.8546 0.04652 0.03702 0.0025 0.1140 1.0000
7.750 0.8654 0.05038 0.04148 0.0041 0.1127 1.0000
8.000 0.8734 0.05465 0.04630 0.0055 0.1128 1.0000
8.250 0.8829 0.05955 0.05148 0.0064 0.1154 1.0000
8.500 0.8713 0.06528 0.05799 0.0072 0.1236 1.0000
8.750 0.8771 0.07100 0.06390 0.0073 0.1314 1.0000
9.000 0.8631 0.07809 0.07132 0.0060 0.1447 1.0000
9.250 0.8142 0.08521 0.07855 0.0015 0.1523 1.0000
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Polar data table (+)
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