GIII BL332 AIRFOIL (giiij-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GIII BL332 AIRFOIL (giiij-il) Reynolds number: 50,000 Max Cl/Cd: 32.35 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-giiij-il-50000-n5.txt Download as CSV file: xf-giiij-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GIII BL332 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 -0.5608 0.11614 0.10891 0.0029 1.0000 0.1267
-9.500 -0.5693 0.11316 0.10602 -0.0008 1.0000 0.1328
-9.250 -0.5789 0.10938 0.10235 -0.0050 1.0000 0.1351
-9.000 -0.5582 0.10499 0.09790 -0.0020 1.0000 0.1394
-8.750 -0.5581 0.10147 0.09443 -0.0038 1.0000 0.1452
-8.500 -0.5730 0.09751 0.09061 -0.0090 1.0000 0.1508
-8.250 -0.5580 0.09369 0.08678 -0.0072 1.0000 0.1551
-7.750 -0.4947 0.06728 0.06048 -0.0281 1.0000 0.0733
-7.500 -0.5635 0.07358 0.06627 -0.0253 1.0000 0.0730
-7.250 -0.5559 0.06932 0.06194 -0.0256 1.0000 0.0691
-7.000 -0.5547 0.06404 0.05608 -0.0275 1.0000 0.0610
-6.750 -0.5442 0.06025 0.05215 -0.0272 1.0000 0.0604
-6.500 -0.5337 0.05670 0.04837 -0.0268 1.0000 0.0604
-6.250 -0.5221 0.05336 0.04472 -0.0262 1.0000 0.0605
-6.000 -0.5090 0.05014 0.04122 -0.0253 1.0000 0.0604
-5.750 -0.4947 0.04703 0.03786 -0.0243 1.0000 0.0599
-5.500 -0.4793 0.04409 0.03464 -0.0232 1.0000 0.0593
-5.250 -0.4630 0.04136 0.03158 -0.0219 1.0000 0.0588
-5.000 -0.4455 0.03883 0.02867 -0.0205 1.0000 0.0586
-4.750 -0.4268 0.03656 0.02602 -0.0191 1.0000 0.0591
-4.500 -0.4068 0.03469 0.02369 -0.0177 1.0000 0.0607
-4.250 -0.3865 0.03265 0.02136 -0.0164 1.0000 0.0624
-4.000 -0.3652 0.03072 0.01924 -0.0154 1.0000 0.0635
-3.750 -0.3429 0.02906 0.01739 -0.0143 1.0000 0.0645
-3.500 -0.3203 0.02764 0.01581 -0.0132 1.0000 0.0667
-3.250 -0.2974 0.02647 0.01450 -0.0122 1.0000 0.0707
-3.000 -0.2734 0.02540 0.01322 -0.0110 1.0000 0.0737
-2.750 -0.2491 0.02434 0.01206 -0.0099 1.0000 0.0757
-2.500 -0.2261 0.02332 0.01104 -0.0090 1.0000 0.0798
-2.250 -0.2038 0.02263 0.01026 -0.0080 1.0000 0.0864
-2.000 -0.1834 0.02196 0.00949 -0.0067 1.0000 0.0913
-1.750 -0.1632 0.02132 0.00883 -0.0056 1.0000 0.0971
-1.500 -0.1423 0.02079 0.00823 -0.0047 1.0000 0.1068
-1.000 -0.0393 0.01700 0.00773 -0.0106 1.0000 1.0000
-0.750 -0.0322 0.01707 0.00757 -0.0077 1.0000 1.0000
-0.500 -0.0222 0.01717 0.00747 -0.0053 1.0000 1.0000
-0.250 0.0005 0.01735 0.00745 -0.0053 0.9960 1.0000
0.000 0.0426 0.01761 0.00749 -0.0089 0.9850 1.0000
0.250 0.0837 0.01785 0.00758 -0.0122 0.9732 1.0000
0.500 0.1251 0.01808 0.00770 -0.0155 0.9612 1.0000
0.750 0.1678 0.01830 0.00785 -0.0189 0.9491 1.0000
1.000 0.2091 0.01849 0.00801 -0.0220 0.9354 1.0000
1.250 0.2506 0.01865 0.00817 -0.0249 0.9212 1.0000
1.500 0.2898 0.01878 0.00833 -0.0272 0.9061 1.0000
1.750 0.3274 0.01888 0.00847 -0.0290 0.8902 1.0000
2.000 0.3622 0.01896 0.00863 -0.0301 0.8730 1.0000
2.250 0.3922 0.01905 0.00878 -0.0302 0.8531 1.0000
2.500 0.4229 0.01909 0.00889 -0.0302 0.8329 1.0000
2.750 0.4521 0.01910 0.00900 -0.0297 0.8117 1.0000
3.000 0.4780 0.01914 0.00912 -0.0286 0.7881 1.0000
3.250 0.5036 0.01914 0.00920 -0.0273 0.7629 1.0000
3.500 0.5285 0.01912 0.00924 -0.0257 0.7350 1.0000
3.750 0.5518 0.01912 0.00931 -0.0238 0.7025 1.0000
4.000 0.5745 0.01908 0.00927 -0.0216 0.6641 1.0000
4.250 0.5953 0.01913 0.00926 -0.0191 0.6156 1.0000
4.500 0.6154 0.01928 0.00926 -0.0166 0.5595 1.0000
4.750 0.6350 0.01963 0.00942 -0.0143 0.4984 1.0000
5.000 0.6537 0.02021 0.00975 -0.0123 0.4299 1.0000
5.250 0.6709 0.02103 0.01022 -0.0103 0.3510 1.0000
5.500 0.6870 0.02218 0.01092 -0.0086 0.2721 1.0000
5.750 0.7035 0.02355 0.01192 -0.0072 0.2108 1.0000
6.000 0.7212 0.02496 0.01305 -0.0061 0.1736 1.0000
6.250 0.7405 0.02627 0.01425 -0.0050 0.1476 1.0000
6.500 0.7603 0.02761 0.01550 -0.0040 0.1308 1.0000
6.750 0.7807 0.02892 0.01679 -0.0030 0.1166 1.0000
7.000 0.8033 0.03030 0.01835 -0.0021 0.1063 1.0000
7.250 0.8254 0.03177 0.01990 -0.0013 0.0971 1.0000
7.500 0.8471 0.03331 0.02144 -0.0006 0.0898 1.0000
7.750 0.8708 0.03521 0.02361 0.0002 0.0843 1.0000
8.000 0.8917 0.03694 0.02553 0.0009 0.0783 1.0000
8.250 0.9122 0.03906 0.02777 0.0016 0.0741 1.0000
8.500 0.9313 0.04169 0.03089 0.0026 0.0711 1.0000
8.750 0.9477 0.04425 0.03382 0.0036 0.0679 1.0000
9.000 0.9631 0.04649 0.03624 0.0045 0.0647 1.0000
9.250 0.9757 0.04933 0.03930 0.0054 0.0622 1.0000
9.500 0.9806 0.05300 0.04355 0.0069 0.0610 1.0000
9.750 0.9809 0.05697 0.04802 0.0084 0.0601 1.0000
10.000 0.9760 0.06115 0.05263 0.0097 0.0595 1.0000
10.250 0.9653 0.06548 0.05731 0.0108 0.0591 1.0000
10.500 0.9482 0.06985 0.06195 0.0117 0.0589 1.0000
10.750 0.9263 0.07472 0.06701 0.0114 0.0590 1.0000
11.000 0.9018 0.08069 0.07313 0.0089 0.0594 1.0000
11.250 0.8764 0.08808 0.08062 0.0043 0.0600 1.0000
11.500 0.8530 0.09678 0.08933 -0.0018 0.0606 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GIII BL332 AIRFOIL (giiij-il)