GIII BL0 AIRFOIL (giiia-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GIII BL0 AIRFOIL (giiia-il) Reynolds number: 50,000 Max Cl/Cd: 27.79 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-giiia-il-50000-n5.txt Download as CSV file: xf-giiia-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GIII BL0 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.6260 0.10021 0.09328 -0.0142 1.0000 0.0486
-10.500 -0.6405 0.09142 0.08454 -0.0205 1.0000 0.0475
-10.250 -0.6864 0.07968 0.07269 -0.0291 1.0000 0.0448
-9.250 -0.7559 0.06257 0.05482 -0.0247 1.0000 0.0443
-9.000 -0.7560 0.05957 0.05171 -0.0230 1.0000 0.0454
-8.750 -0.7517 0.05715 0.04920 -0.0213 1.0000 0.0473
-8.500 -0.7510 0.05415 0.04593 -0.0190 1.0000 0.0489
-8.250 -0.7496 0.05086 0.04228 -0.0165 1.0000 0.0504
-8.000 -0.7457 0.04748 0.03845 -0.0139 1.0000 0.0516
-7.750 -0.7382 0.04418 0.03466 -0.0112 1.0000 0.0528
-7.500 -0.7267 0.04111 0.03105 -0.0088 1.0000 0.0543
-7.250 -0.7123 0.03860 0.02817 -0.0069 1.0000 0.0573
-7.000 -0.6958 0.03687 0.02633 -0.0054 1.0000 0.0616
-6.750 -0.6754 0.03470 0.02380 -0.0039 1.0000 0.0655
-6.500 -0.6513 0.03255 0.02133 -0.0029 1.0000 0.0696
-6.250 -0.6309 0.03118 0.01994 -0.0018 1.0000 0.0767
-6.000 -0.6059 0.02961 0.01824 -0.0010 1.0000 0.0847
-5.750 -0.5812 0.02830 0.01676 -0.0001 1.0000 0.0947
-5.500 -0.5597 0.02719 0.01560 0.0011 1.0000 0.1092
-5.250 -0.5392 0.02600 0.01448 0.0024 1.0000 0.1251
-5.000 -0.5216 0.02499 0.01356 0.0041 1.0000 0.1518
-4.750 -0.5063 0.02390 0.01266 0.0061 1.0000 0.1859
-4.500 -0.4926 0.02280 0.01180 0.0083 1.0000 0.2339
-4.250 -0.4814 0.02154 0.01094 0.0108 1.0000 0.3012
-4.000 -0.4771 0.01992 0.01042 0.0150 1.0000 0.4390
-3.750 -0.4506 0.01957 0.01108 0.0180 1.0000 0.7004
-3.500 -0.4101 0.02025 0.01170 0.0185 1.0000 0.7920
-3.250 -0.3711 0.02092 0.01207 0.0183 1.0000 0.8393
-3.000 -0.3306 0.02151 0.01236 0.0174 1.0000 0.8747
-2.750 -0.2753 0.02205 0.01256 0.0133 1.0000 0.9040
-2.500 -0.2082 0.02239 0.01254 0.0066 1.0000 0.9305
-2.250 -0.1457 0.02240 0.01228 0.0001 1.0000 0.9535
-2.000 -0.1021 0.02214 0.01183 -0.0037 1.0000 0.9672
-1.750 -0.0686 0.02184 0.01142 -0.0059 1.0000 0.9768
-1.500 -0.0347 0.02154 0.01103 -0.0082 1.0000 0.9858
-1.250 0.0010 0.02124 0.01065 -0.0109 1.0000 0.9945
-1.000 0.0269 0.02101 0.01039 -0.0119 1.0000 1.0000
-0.750 0.0297 0.02097 0.01034 -0.0087 1.0000 1.0000
-0.500 0.0309 0.02099 0.01036 -0.0051 1.0000 1.0000
-0.250 0.0309 0.02106 0.01042 -0.0015 1.0000 1.0000
0.000 0.0544 0.02114 0.01049 -0.0022 0.9926 1.0000
0.250 0.1042 0.02116 0.01051 -0.0076 0.9765 1.0000
0.500 0.1552 0.02114 0.01051 -0.0130 0.9599 1.0000
0.750 0.2105 0.02097 0.01039 -0.0188 0.9391 1.0000
1.000 0.2649 0.02061 0.01009 -0.0237 0.9101 1.0000
1.250 0.3082 0.02018 0.00968 -0.0259 0.8739 1.0000
1.500 0.3380 0.01989 0.00938 -0.0255 0.8354 1.0000
1.750 0.3621 0.01972 0.00923 -0.0243 0.7994 1.0000
2.000 0.3839 0.01962 0.00913 -0.0228 0.7638 1.0000
2.250 0.4063 0.01952 0.00900 -0.0212 0.7287 1.0000
2.500 0.4286 0.01946 0.00886 -0.0195 0.6927 1.0000
2.750 0.4501 0.01949 0.00880 -0.0177 0.6555 1.0000
3.000 0.4706 0.01963 0.00882 -0.0158 0.6171 1.0000
3.250 0.4904 0.01985 0.00893 -0.0140 0.5774 1.0000
3.500 0.5094 0.02014 0.00910 -0.0121 0.5350 1.0000
3.750 0.5278 0.02051 0.00936 -0.0101 0.4899 1.0000
4.000 0.5456 0.02095 0.00964 -0.0082 0.4438 1.0000
4.250 0.5641 0.02146 0.00999 -0.0066 0.4017 1.0000
4.500 0.5833 0.02203 0.01040 -0.0051 0.3710 1.0000
4.750 0.6034 0.02263 0.01094 -0.0039 0.3485 1.0000
5.000 0.6241 0.02325 0.01153 -0.0028 0.3305 1.0000
5.250 0.6454 0.02386 0.01215 -0.0017 0.3162 1.0000
5.500 0.6672 0.02448 0.01281 -0.0007 0.3046 1.0000
5.750 0.6894 0.02513 0.01350 0.0002 0.2956 1.0000
6.000 0.7123 0.02580 0.01438 0.0010 0.2877 1.0000
6.250 0.7352 0.02652 0.01515 0.0018 0.2811 1.0000
6.500 0.7574 0.02728 0.01614 0.0027 0.2735 1.0000
6.750 0.7801 0.02807 0.01703 0.0036 0.2672 1.0000
7.000 0.8018 0.02898 0.01824 0.0045 0.2609 1.0000
7.250 0.8236 0.02988 0.01936 0.0055 0.2541 1.0000
7.500 0.8409 0.03067 0.02038 0.0071 0.2418 1.0000
7.750 0.8514 0.03105 0.02077 0.0096 0.2200 1.0000
8.000 0.8578 0.03131 0.02106 0.0125 0.1957 1.0000
8.250 0.8662 0.03181 0.02179 0.0152 0.1754 1.0000
8.500 0.8730 0.03234 0.02251 0.0181 0.1524 1.0000
8.750 0.8801 0.03303 0.02350 0.0209 0.1155 1.0000
9.000 0.8842 0.03424 0.02442 0.0237 0.0854 1.0000
9.250 0.8870 0.03612 0.02614 0.0265 0.0723 1.0000
9.500 0.8895 0.03811 0.02814 0.0292 0.0637 1.0000
9.750 0.8912 0.04018 0.03027 0.0318 0.0579 1.0000
10.000 0.8916 0.04223 0.03243 0.0344 0.0536 1.0000
10.250 0.8892 0.04439 0.03465 0.0371 0.0512 1.0000
10.500 0.8880 0.04678 0.03712 0.0394 0.0493 1.0000
10.750 0.8886 0.04926 0.03989 0.0411 0.0472 1.0000
11.000 0.8867 0.05200 0.04286 0.0424 0.0452 1.0000
11.250 0.8828 0.05497 0.04602 0.0433 0.0434 1.0000
11.500 0.8777 0.05818 0.04937 0.0435 0.0418 1.0000
11.750 0.8721 0.06167 0.05295 0.0434 0.0406 1.0000
12.000 0.8665 0.06545 0.05682 0.0429 0.0396 1.0000
12.250 0.8593 0.06975 0.06124 0.0420 0.0389 1.0000
12.500 0.8479 0.07488 0.06660 0.0399 0.0388 1.0000
12.750 0.8346 0.08066 0.07259 0.0371 0.0387 1.0000
13.000 0.8196 0.08708 0.07921 0.0336 0.0387 1.0000
13.250 0.8034 0.09417 0.08646 0.0297 0.0388 1.0000
13.500 0.7863 0.10184 0.09427 0.0254 0.0390 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GIII BL0 AIRFOIL (giiia-il)