DSMA-523B AIRFOIL (dsma523b-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: DSMA-523B AIRFOIL (dsma523b-il) Reynolds number: 100,000 Max Cl/Cd: 23.34 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-dsma523b-il-100000.txt Download as CSV file: xf-dsma523b-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: DSMA-523B 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 -0.5589 0.10034 0.09466 -0.0156 1.0000 0.1893
-8.750 -0.7755 0.05484 0.04758 -0.0538 1.0000 0.0969
-8.500 -0.7587 0.05064 0.04321 -0.0547 1.0000 0.0960
-8.250 -0.7396 0.04641 0.03871 -0.0562 1.0000 0.0953
-8.000 -0.7170 0.04239 0.03436 -0.0580 1.0000 0.0944
-7.750 -0.6909 0.03865 0.03022 -0.0600 1.0000 0.0935
-7.500 -0.6626 0.03555 0.02669 -0.0617 1.0000 0.0936
-7.250 -0.6329 0.03320 0.02391 -0.0632 1.0000 0.0953
-7.000 -0.6042 0.03112 0.02154 -0.0642 1.0000 0.0978
-6.750 -0.5785 0.02962 0.02003 -0.0643 1.0000 0.1005
-6.500 -0.5516 0.02827 0.01859 -0.0642 1.0000 0.1031
-6.250 -0.5237 0.02713 0.01728 -0.0643 1.0000 0.1070
-6.000 -0.4973 0.02588 0.01601 -0.0642 1.0000 0.1115
-5.750 -0.4708 0.02497 0.01517 -0.0639 1.0000 0.1165
-5.500 -0.4429 0.02420 0.01430 -0.0638 1.0000 0.1223
-5.250 -0.4166 0.02327 0.01356 -0.0636 1.0000 0.1298
-5.000 -0.3884 0.02268 0.01295 -0.0635 1.0000 0.1386
-4.750 -0.3606 0.02193 0.01244 -0.0637 1.0000 0.1501
-4.500 -0.3313 0.02130 0.01199 -0.0640 1.0000 0.1658
-4.250 -0.2997 0.02062 0.01160 -0.0649 1.0000 0.1979
-4.000 -0.2590 0.01909 0.01225 -0.0682 1.0000 0.4803
-3.750 -0.2712 0.02082 0.01441 -0.0559 1.0000 0.5581
-3.500 -0.2370 0.02181 0.01518 -0.0570 1.0000 0.6318
-3.250 -0.2223 0.02282 0.01613 -0.0528 1.0000 0.6538
-3.000 -0.2057 0.02360 0.01686 -0.0493 1.0000 0.6720
-2.750 -0.1899 0.02427 0.01749 -0.0456 1.0000 0.6890
-2.500 -0.1694 0.02476 0.01792 -0.0434 1.0000 0.7052
-2.250 -0.1628 0.02510 0.01829 -0.0374 1.0000 0.7148
-2.000 -0.1469 0.02538 0.01855 -0.0341 1.0000 0.7294
-1.750 -0.1285 0.02562 0.01878 -0.0316 1.0000 0.7450
-1.500 -0.1153 0.02574 0.01891 -0.0277 1.0000 0.7587
-1.250 -0.1048 0.02572 0.01891 -0.0232 1.0000 0.7722
-1.000 -0.0914 0.02569 0.01891 -0.0196 1.0000 0.7878
-0.750 -0.0797 0.02561 0.01885 -0.0155 1.0000 0.8052
-0.500 -0.0683 0.02544 0.01871 -0.0116 1.0000 0.8228
-0.250 -0.0556 0.02521 0.01852 -0.0081 1.0000 0.8396
0.000 -0.0364 0.02505 0.01838 -0.0064 1.0000 0.8531
0.250 -0.0159 0.02489 0.01825 -0.0051 1.0000 0.8639
0.500 0.0021 0.02461 0.01802 -0.0034 1.0000 0.8732
0.750 0.0261 0.02455 0.01801 -0.0032 1.0000 0.8822
1.000 0.0488 0.02448 0.01801 -0.0027 1.0000 0.8910
1.250 0.0690 0.02434 0.01793 -0.0018 1.0000 0.8990
1.500 0.0930 0.02436 0.01805 -0.0018 1.0000 0.9060
1.750 0.1185 0.02450 0.01829 -0.0022 1.0000 0.9114
2.000 0.2150 0.02384 0.01778 -0.0147 0.9639 0.9158
2.250 0.2846 0.02204 0.01617 -0.0211 0.9268 0.9176
2.500 0.3336 0.02016 0.01448 -0.0227 0.8653 0.9185
2.750 0.4398 0.02161 0.01223 -0.0335 0.1837 0.9199
3.000 0.4681 0.02243 0.01286 -0.0339 0.1624 0.9213
3.250 0.4979 0.02335 0.01359 -0.0347 0.1485 0.9223
3.500 0.5300 0.02403 0.01427 -0.0357 0.1381 0.9227
3.750 0.5631 0.02515 0.01520 -0.0371 0.1305 0.9235
4.000 0.5961 0.02611 0.01624 -0.0381 0.1248 0.9248
4.250 0.6285 0.02700 0.01719 -0.0392 0.1184 0.9258
4.500 0.6630 0.02841 0.01854 -0.0408 0.1141 0.9261
4.750 0.6971 0.03036 0.02061 -0.0423 0.1109 0.9264
5.000 0.7273 0.03150 0.02207 -0.0427 0.1073 0.9273
5.250 0.7573 0.03309 0.02392 -0.0432 0.1046 0.9283
5.500 0.7866 0.03508 0.02622 -0.0436 0.1031 0.9289
5.750 0.8146 0.03713 0.02853 -0.0439 0.1014 0.9290
6.000 0.8429 0.03917 0.03063 -0.0447 0.0987 0.9292
6.250 0.8674 0.04225 0.03393 -0.0449 0.0973 0.9294
6.500 0.8792 0.04650 0.03931 -0.0418 0.1026 0.9295
6.750 0.8858 0.05365 0.04739 -0.0390 0.1184 0.9297
7.000 0.8444 0.07323 0.06877 -0.0382 0.2134 0.9295
7.250 0.8298 0.07781 0.07354 -0.0379 0.2022 0.9296
7.500 0.8369 0.08094 0.07675 -0.0378 0.1939 0.9302
7.750 0.8989 0.08800 0.08332 -0.0391 0.1879 0.9315
8.000 0.8057 0.09050 0.08646 -0.0388 0.1817 0.9311
8.250 0.7955 0.09534 0.09137 -0.0414 0.1749 0.9312
8.500 0.9086 0.09994 0.09553 -0.0389 0.1675 0.9323
8.750 0.8208 0.10388 0.09987 -0.0409 0.1659 0.9315
9.000 0.7709 0.11317 0.10918 -0.0548 0.1581 0.9311
|
Polar data table (+)
Polar graphs
<< Back to DSMA-523B AIRFOIL (dsma523b-il)