DSMA-523B AIRFOIL (dsma523b-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: DSMA-523B AIRFOIL (dsma523b-il) Reynolds number: 50,000 Max Cl/Cd: 20.86 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-dsma523b-il-50000.txt Download as CSV file: xf-dsma523b-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5156 0.11381 0.10577 0.0005 1.0000 0.3697
-8.500 -0.4790 0.10854 0.10044 0.0012 1.0000 0.3788
-8.250 -0.6937 0.07479 0.06680 -0.0363 1.0000 0.1778
-8.000 -0.6869 0.06967 0.06161 -0.0373 1.0000 0.1718
-7.750 -0.6969 0.05735 0.04810 -0.0511 1.0000 0.1572
-7.500 -0.6775 0.05275 0.04326 -0.0528 1.0000 0.1558
-7.250 -0.6533 0.04832 0.03836 -0.0558 1.0000 0.1559
-7.000 -0.6260 0.04429 0.03392 -0.0587 1.0000 0.1582
-6.750 -0.6014 0.04142 0.03097 -0.0594 1.0000 0.1615
-6.500 -0.5730 0.03870 0.02800 -0.0607 1.0000 0.1643
-6.250 -0.5436 0.03645 0.02545 -0.0619 1.0000 0.1702
-6.000 -0.5113 0.03431 0.02285 -0.0635 1.0000 0.1769
-5.750 -0.4873 0.03251 0.02116 -0.0627 1.0000 0.1838
-5.500 -0.4588 0.03113 0.01955 -0.0628 1.0000 0.1949
-5.250 -0.4372 0.02981 0.01841 -0.0610 1.0000 0.2055
-5.000 -0.4138 0.02872 0.01732 -0.0594 1.0000 0.2203
-4.750 -0.3917 0.02775 0.01647 -0.0575 1.0000 0.2400
-4.500 -0.3703 0.02665 0.01574 -0.0556 1.0000 0.2691
-4.250 -0.3427 0.02455 0.01478 -0.0557 1.0000 0.3696
-4.000 -0.3592 0.02575 0.01759 -0.0408 1.0000 0.5060
-3.750 -0.3520 0.02910 0.02081 -0.0313 1.0000 0.6321
-3.500 -0.3446 0.03089 0.02245 -0.0237 1.0000 0.6763
-3.250 -0.3393 0.03191 0.02339 -0.0160 1.0000 0.7109
-3.000 -0.3335 0.03242 0.02381 -0.0090 1.0000 0.7450
-2.750 -0.3300 0.03237 0.02369 -0.0015 1.0000 0.7742
-2.500 -0.3226 0.03197 0.02321 0.0043 1.0000 0.8036
-2.250 -0.3132 0.03142 0.02256 0.0091 1.0000 0.8336
-2.000 -0.3061 0.03046 0.02154 0.0148 1.0000 0.8626
-1.750 -0.2958 0.02940 0.02041 0.0192 1.0000 0.8932
-1.500 -0.2785 0.02818 0.01909 0.0222 1.0000 0.9234
-1.250 -0.0456 0.02724 0.01764 -0.0146 1.0000 1.0000
-1.000 -0.0364 0.02683 0.01723 -0.0125 1.0000 1.0000
-0.750 -0.0273 0.02645 0.01686 -0.0104 1.0000 1.0000
-0.500 -0.0182 0.02611 0.01654 -0.0082 1.0000 1.0000
-0.250 -0.0093 0.02579 0.01626 -0.0059 1.0000 1.0000
0.000 -0.0006 0.02550 0.01601 -0.0035 1.0000 1.0000
0.250 0.0080 0.02524 0.01579 -0.0012 1.0000 1.0000
0.500 0.0164 0.02499 0.01561 0.0012 1.0000 1.0000
0.750 0.0246 0.02477 0.01546 0.0036 1.0000 1.0000
1.000 0.0329 0.02457 0.01535 0.0060 1.0000 1.0000
1.250 0.0412 0.02441 0.01527 0.0083 1.0000 1.0000
1.500 0.0507 0.02429 0.01526 0.0103 1.0000 1.0000
1.750 0.0635 0.02429 0.01536 0.0116 1.0000 1.0000
2.000 0.0802 0.02441 0.01561 0.0121 1.0000 1.0000
2.250 0.1001 0.02467 0.01602 0.0120 1.0000 1.0000
2.500 0.1223 0.02505 0.01658 0.0114 1.0000 1.0000
2.750 0.1458 0.02557 0.01730 0.0103 1.0000 1.0000
3.000 0.1700 0.02624 0.01822 0.0089 1.0000 1.0000
3.250 0.5409 0.02667 0.01539 -0.0384 0.2318 1.0000
3.500 0.5778 0.02785 0.01651 -0.0400 0.2120 1.0000
3.750 0.6100 0.02924 0.01800 -0.0407 0.2011 1.0000
4.000 0.6348 0.03051 0.01927 -0.0403 0.1914 1.0000
4.250 0.6577 0.03209 0.02100 -0.0395 0.1844 1.0000
4.500 0.6782 0.03362 0.02284 -0.0381 0.1802 1.0000
4.750 0.6977 0.03526 0.02472 -0.0366 0.1761 1.0000
5.000 0.7184 0.03704 0.02655 -0.0357 0.1709 1.0000
5.250 0.7391 0.03974 0.02936 -0.0351 0.1673 1.0000
5.500 0.7587 0.04219 0.03223 -0.0340 0.1666 1.0000
5.750 0.7780 0.04503 0.03548 -0.0330 0.1665 1.0000
6.000 0.7953 0.04799 0.03888 -0.0320 0.1661 1.0000
6.250 0.8113 0.05123 0.04254 -0.0310 0.1657 1.0000
6.500 0.8289 0.05526 0.04685 -0.0307 0.1669 1.0000
6.750 0.8195 0.06046 0.05319 -0.0280 0.1805 1.0000
7.000 0.8417 0.06603 0.05878 -0.0290 0.1872 1.0000
7.250 0.6144 0.09744 0.09181 -0.0644 0.4639 1.0000
7.500 0.6272 0.10104 0.09541 -0.0643 0.4412 1.0000
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Polar data table (+)
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