TSAGI 8% AIRFOIL (tsagi8-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: TSAGI 8% AIRFOIL (tsagi8-il) Reynolds number: 100,000 Max Cl/Cd: 45.29 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-tsagi8-il-100000.txt Download as CSV file: xf-tsagi8-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: TSAGI 8% AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5619 0.09918 0.09438 -0.0066 1.0000 0.0914
-8.750 -0.5779 0.09546 0.09075 -0.0127 1.0000 0.0924
-8.500 -0.5994 0.09231 0.08762 -0.0169 1.0000 0.0928
-8.250 -0.5655 0.08684 0.08219 -0.0113 1.0000 0.0976
-8.000 -0.5688 0.08330 0.07868 -0.0124 1.0000 0.1009
-7.750 -0.4934 0.07001 0.06572 -0.0152 1.0000 0.1133
-7.500 -0.5068 0.06602 0.06176 -0.0164 1.0000 0.1163
-7.250 -0.5357 0.06190 0.05749 -0.0200 1.0000 0.1200
-7.000 -0.5245 0.05668 0.05236 -0.0185 1.0000 0.1237
-6.750 -0.5772 0.06388 0.05906 -0.0175 1.0000 0.1234
-6.500 -0.5800 0.06184 0.05663 -0.0175 1.0000 0.1348
-6.250 -0.5623 0.05710 0.05215 -0.0159 1.0000 0.1394
-6.000 -0.5551 0.05401 0.04897 -0.0148 1.0000 0.1528
-5.750 -0.5457 0.05131 0.04620 -0.0132 1.0000 0.1692
-5.500 -0.5402 0.04872 0.04344 -0.0115 1.0000 0.1919
-5.250 -0.5285 0.04581 0.04055 -0.0095 1.0000 0.2079
-5.000 -0.5014 0.03472 0.02744 -0.0074 1.0000 0.0771
-4.750 -0.4837 0.03135 0.02317 -0.0038 1.0000 0.0659
-4.500 -0.4656 0.02796 0.01952 -0.0019 1.0000 0.0644
-4.250 -0.4455 0.02552 0.01668 0.0001 1.0000 0.0643
-4.000 -0.4252 0.02347 0.01431 0.0019 1.0000 0.0683
-3.750 -0.4033 0.02176 0.01251 0.0032 1.0000 0.0726
-3.500 -0.3798 0.02021 0.01073 0.0048 1.0000 0.0771
-3.250 -0.3585 0.01882 0.00936 0.0062 1.0000 0.0860
-3.000 -0.3385 0.01762 0.00816 0.0078 1.0000 0.0979
-2.750 -0.0974 0.01299 0.00641 -0.0289 1.0000 1.0000
-2.500 -0.0795 0.01278 0.00602 -0.0275 1.0000 1.0000
-2.250 -0.0620 0.01261 0.00570 -0.0261 1.0000 1.0000
-2.000 -0.0451 0.01248 0.00545 -0.0245 1.0000 1.0000
-1.750 -0.0294 0.01239 0.00527 -0.0227 1.0000 1.0000
-1.500 -0.0154 0.01235 0.00514 -0.0206 1.0000 1.0000
-1.250 -0.0042 0.01239 0.00513 -0.0181 1.0000 1.0000
-1.000 0.0012 0.01255 0.00526 -0.0147 1.0000 1.0000
-0.750 -0.0014 0.01287 0.00555 -0.0102 1.0000 1.0000
-0.500 -0.0062 0.01331 0.00594 -0.0057 1.0000 1.0000
-0.250 0.0447 0.01349 0.00605 -0.0114 0.9882 1.0000
0.000 0.1038 0.01357 0.00606 -0.0184 0.9742 1.0000
0.250 0.1626 0.01355 0.00602 -0.0251 0.9606 1.0000
0.500 0.2218 0.01344 0.00592 -0.0317 0.9469 1.0000
0.750 0.2758 0.01328 0.00579 -0.0370 0.9308 1.0000
1.000 0.3248 0.01310 0.00565 -0.0411 0.9130 1.0000
1.250 0.3597 0.01305 0.00565 -0.0423 0.8899 1.0000
1.500 0.3894 0.01305 0.00566 -0.0422 0.8677 1.0000
1.750 0.4139 0.01313 0.00574 -0.0411 0.8454 1.0000
2.000 0.4358 0.01327 0.00589 -0.0395 0.8234 1.0000
2.250 0.4573 0.01342 0.00608 -0.0377 0.8029 1.0000
2.500 0.4782 0.01361 0.00630 -0.0359 0.7817 1.0000
2.750 0.4994 0.01380 0.00651 -0.0342 0.7617 1.0000
3.000 0.5206 0.01401 0.00676 -0.0324 0.7413 1.0000
3.250 0.5418 0.01421 0.00701 -0.0307 0.7207 1.0000
3.500 0.5633 0.01442 0.00732 -0.0289 0.7001 1.0000
3.750 0.5846 0.01460 0.00758 -0.0272 0.6782 1.0000
4.000 0.6042 0.01463 0.00761 -0.0246 0.6491 1.0000
4.250 0.6201 0.01438 0.00727 -0.0209 0.5998 1.0000
4.500 0.6367 0.01429 0.00710 -0.0177 0.5450 1.0000
4.750 0.6522 0.01440 0.00708 -0.0146 0.4655 1.0000
5.000 0.6485 0.01730 0.00789 -0.0093 0.1223 1.0000
5.250 0.6600 0.01930 0.00955 -0.0065 0.0842 1.0000
5.500 0.6746 0.02077 0.01098 -0.0041 0.0707 1.0000
5.750 0.6926 0.02209 0.01237 -0.0019 0.0646 1.0000
6.000 0.7118 0.02399 0.01413 -0.0002 0.0601 1.0000
6.250 0.7341 0.02575 0.01603 0.0012 0.0557 1.0000
6.500 0.7570 0.02758 0.01803 0.0026 0.0527 1.0000
6.750 0.7800 0.02997 0.02073 0.0041 0.0519 1.0000
7.000 0.8008 0.03270 0.02388 0.0059 0.0527 1.0000
7.250 0.8187 0.03589 0.02752 0.0080 0.0546 1.0000
7.500 0.8337 0.03952 0.03153 0.0101 0.0566 1.0000
7.750 0.8453 0.04305 0.03545 0.0123 0.0572 1.0000
8.000 0.8546 0.04702 0.03974 0.0143 0.0579 1.0000
8.250 0.8585 0.05185 0.04534 0.0178 0.0697 1.0000
10.500 0.7012 0.11027 0.10521 0.0018 0.1497 1.0000
10.750 0.7209 0.11310 0.10811 0.0055 0.1372 1.0000
11.000 0.6891 0.11866 0.11353 -0.0021 0.1353 1.0000
|
Polar data table (+)
Polar graphs
<< Back to TSAGI 8% AIRFOIL (tsagi8-il)