DSMA-523A AIRFOIL (dsma523a-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: DSMA-523A AIRFOIL (dsma523a-il) Reynolds number: 50,000 Max Cl/Cd: 18.65 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-dsma523a-il-50000-n5.txt Download as CSV file: xf-dsma523a-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: DSMA-523A 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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.7266 0.08055 0.07168 -0.0475 1.0000 0.0718
-10.250 -0.7524 0.07558 0.06663 -0.0480 1.0000 0.0715
-10.000 -0.7788 0.07150 0.06246 -0.0472 1.0000 0.0713
-9.750 -0.8004 0.06727 0.05808 -0.0469 1.0000 0.0712
-9.500 -0.8152 0.06299 0.05356 -0.0470 1.0000 0.0712
-9.250 -0.8238 0.05891 0.04914 -0.0469 1.0000 0.0713
-9.000 -0.8266 0.05495 0.04476 -0.0468 1.0000 0.0718
-8.750 -0.8142 0.05275 0.04255 -0.0461 1.0000 0.0730
-8.500 -0.8021 0.05055 0.04024 -0.0455 1.0000 0.0745
-8.250 -0.7902 0.04798 0.03741 -0.0452 1.0000 0.0762
-8.000 -0.7765 0.04520 0.03426 -0.0450 1.0000 0.0779
-7.750 -0.7600 0.04251 0.03113 -0.0447 1.0000 0.0793
-7.500 -0.7409 0.04061 0.02914 -0.0440 1.0000 0.0806
-7.250 -0.7212 0.03910 0.02757 -0.0433 1.0000 0.0829
-7.000 -0.7007 0.03754 0.02582 -0.0426 1.0000 0.0860
-6.750 -0.6796 0.03601 0.02409 -0.0417 1.0000 0.0889
-6.500 -0.6588 0.03485 0.02295 -0.0406 1.0000 0.0915
-6.250 -0.6375 0.03374 0.02176 -0.0392 1.0000 0.0949
-6.000 -0.6161 0.03275 0.02062 -0.0377 1.0000 0.0993
-5.750 -0.5958 0.03193 0.01989 -0.0361 1.0000 0.1037
-5.500 -0.5751 0.03120 0.01907 -0.0341 1.0000 0.1087
-5.250 -0.5550 0.03047 0.01838 -0.0323 1.0000 0.1143
-5.000 -0.5347 0.02977 0.01769 -0.0308 1.0000 0.1213
-4.750 -0.5144 0.02904 0.01702 -0.0294 1.0000 0.1292
-4.500 -0.4934 0.02833 0.01635 -0.0283 1.0000 0.1398
-4.250 -0.4720 0.02748 0.01568 -0.0276 1.0000 0.1540
-4.000 -0.4493 0.02655 0.01493 -0.0274 1.0000 0.1756
-3.750 -0.4242 0.02535 0.01415 -0.0281 1.0000 0.2181
-3.500 -0.3966 0.02379 0.01342 -0.0300 1.0000 0.3176
-3.250 -0.3886 0.02389 0.01454 -0.0242 1.0000 0.4383
-3.000 -0.3919 0.02573 0.01683 -0.0129 1.0000 0.5135
-2.500 -0.3320 0.02686 0.01761 -0.0144 1.0000 0.6513
-2.000 -0.2890 0.02795 0.01848 -0.0108 1.0000 0.7012
-1.750 -0.2720 0.02828 0.01876 -0.0078 1.0000 0.7193
-1.500 -0.2568 0.02854 0.01899 -0.0043 1.0000 0.7375
-1.250 -0.2441 0.02871 0.01915 -0.0002 1.0000 0.7563
-1.000 -0.2328 0.02876 0.01920 0.0043 1.0000 0.7756
-0.750 -0.2211 0.02868 0.01912 0.0084 1.0000 0.7941
-0.500 -0.2057 0.02851 0.01895 0.0112 1.0000 0.8103
-0.250 -0.1896 0.02823 0.01866 0.0137 1.0000 0.8223
0.000 -0.1705 0.02794 0.01837 0.0152 1.0000 0.8315
0.250 -0.1489 0.02772 0.01815 0.0159 1.0000 0.8394
0.500 -0.1256 0.02754 0.01798 0.0161 1.0000 0.8452
0.750 -0.1032 0.02733 0.01778 0.0165 1.0000 0.8493
1.000 -0.0781 0.02722 0.01772 0.0161 1.0000 0.8527
1.250 -0.0514 0.02720 0.01774 0.0153 1.0000 0.8556
1.500 -0.0235 0.02726 0.01787 0.0142 1.0000 0.8579
1.750 0.0035 0.02732 0.01802 0.0132 1.0000 0.8600
2.000 0.0448 0.02760 0.01842 0.0098 0.9909 0.8616
2.250 0.1040 0.02796 0.01894 0.0033 0.9705 0.8625
2.500 0.1739 0.02731 0.01849 -0.0036 0.9263 0.8629
2.750 0.2208 0.02589 0.01724 -0.0048 0.8543 0.8638
3.000 0.2493 0.02505 0.01656 -0.0036 0.7807 0.8652
3.250 0.3455 0.02523 0.01464 -0.0109 0.3528 0.8644
3.500 0.3658 0.02651 0.01513 -0.0103 0.2337 0.8653
3.750 0.3921 0.02748 0.01570 -0.0107 0.1837 0.8661
4.000 0.4209 0.02830 0.01632 -0.0114 0.1599 0.8670
4.250 0.4514 0.02905 0.01700 -0.0122 0.1439 0.8683
4.500 0.4845 0.02988 0.01781 -0.0134 0.1328 0.8696
4.750 0.5195 0.03074 0.01866 -0.0149 0.1235 0.8708
5.000 0.5552 0.03174 0.01968 -0.0165 0.1157 0.8717
5.250 0.5904 0.03283 0.02087 -0.0180 0.1099 0.8723
5.500 0.6237 0.03408 0.02203 -0.0194 0.1049 0.8729
5.750 0.6560 0.03541 0.02356 -0.0205 0.1000 0.8736
6.000 0.6869 0.03685 0.02523 -0.0213 0.0960 0.8743
6.250 0.7148 0.03833 0.02687 -0.0216 0.0932 0.8753
6.500 0.7409 0.03984 0.02843 -0.0219 0.0902 0.8766
6.750 0.7645 0.04168 0.03057 -0.0216 0.0874 0.8779
7.000 0.7862 0.04369 0.03305 -0.0210 0.0847 0.8792
7.250 0.8066 0.04603 0.03579 -0.0205 0.0832 0.8803
7.500 0.8251 0.04847 0.03860 -0.0199 0.0816 0.8812
7.750 0.8430 0.05077 0.04117 -0.0195 0.0797 0.8820
8.000 0.8626 0.05283 0.04336 -0.0194 0.0777 0.8828
8.250 0.8792 0.05549 0.04612 -0.0192 0.0761 0.8836
8.500 0.8833 0.05903 0.05025 -0.0176 0.0752 0.8844
8.750 0.8856 0.06287 0.05455 -0.0164 0.0748 0.8854
9.000 0.8842 0.06694 0.05902 -0.0153 0.0746 0.8865
9.250 0.8775 0.07133 0.06378 -0.0144 0.0745 0.8876
9.500 0.8654 0.07584 0.06860 -0.0136 0.0746 0.8887
9.750 0.8482 0.08043 0.07343 -0.0130 0.0748 0.8898
10.000 0.8294 0.08588 0.07909 -0.0141 0.0750 0.8908
10.250 0.8112 0.09204 0.08540 -0.0170 0.0753 0.8915
10.500 0.7952 0.09904 0.09252 -0.0216 0.0756 0.8922
|
Polar data table (+)
Polar graphs
<< Back to DSMA-523A AIRFOIL (dsma523a-il)