DORNIER A-5 AIRFOIL (doa5-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: DORNIER A-5 AIRFOIL (doa5-il) Reynolds number: 50,000 Max Cl/Cd: 29.86 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-doa5-il-50000.txt Download as CSV file: xf-doa5-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: DORNIER A-5 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 -0.3604 0.11399 0.10685 -0.0109 1.0000 0.4551
-8.500 -0.6304 0.08701 0.08038 -0.0319 1.0000 0.2261
-8.250 -0.7240 0.08294 0.07632 -0.0286 1.0000 0.2281
-8.000 -0.7335 0.07496 0.06818 -0.0294 1.0000 0.1984
-7.750 -0.7579 0.06970 0.06259 -0.0283 1.0000 0.1847
-7.500 -0.7762 0.06480 0.05712 -0.0271 1.0000 0.1721
-7.250 -0.7750 0.06061 0.05262 -0.0260 1.0000 0.1653
-7.000 -0.7739 0.05716 0.04835 -0.0248 1.0000 0.1574
-6.750 -0.7618 0.05405 0.04500 -0.0238 1.0000 0.1560
-6.500 -0.7486 0.05130 0.04191 -0.0229 1.0000 0.1553
-6.250 -0.7335 0.04880 0.03904 -0.0219 1.0000 0.1551
-6.000 -0.7161 0.04649 0.03641 -0.0210 1.0000 0.1548
-5.750 -0.6972 0.04438 0.03399 -0.0201 1.0000 0.1545
-5.500 -0.6771 0.04253 0.03184 -0.0191 1.0000 0.1548
-5.250 -0.6568 0.04104 0.02998 -0.0181 1.0000 0.1572
-5.000 -0.6364 0.03947 0.02821 -0.0171 1.0000 0.1604
-4.750 -0.6158 0.03804 0.02681 -0.0159 1.0000 0.1642
-4.500 -0.5945 0.03695 0.02566 -0.0145 1.0000 0.1684
-4.250 -0.5733 0.03613 0.02471 -0.0129 1.0000 0.1739
-4.000 -0.5537 0.03525 0.02401 -0.0108 1.0000 0.1819
-3.750 -0.5345 0.03472 0.02345 -0.0087 1.0000 0.1916
-3.500 -0.5185 0.03400 0.02295 -0.0061 1.0000 0.2028
-3.250 -0.5037 0.03315 0.02227 -0.0039 1.0000 0.2204
-3.000 -0.4892 0.03198 0.02136 -0.0023 1.0000 0.2471
-2.750 -0.4834 0.02929 0.02084 0.0000 1.0000 0.4144
-2.500 -0.4950 0.03337 0.02558 0.0151 1.0000 0.6769
-2.250 -0.5025 0.03692 0.02904 0.0285 1.0000 0.7386
-2.000 -0.1219 0.04725 0.03795 -0.0073 1.0000 0.9796
-1.750 -0.1071 0.04698 0.03759 -0.0072 1.0000 0.9831
-1.500 -0.1071 0.04673 0.03729 -0.0044 1.0000 0.9822
-1.250 -0.1037 0.04647 0.03699 -0.0023 1.0000 0.9822
-1.000 -0.0947 0.04623 0.03669 -0.0012 1.0000 0.9834
-0.750 -0.0873 0.04602 0.03643 0.0001 1.0000 0.9843
-0.500 -0.0776 0.04585 0.03622 0.0010 1.0000 0.9859
-0.250 -0.0658 0.04576 0.03609 0.0015 1.0000 0.9883
0.000 -0.0544 0.04573 0.03602 0.0020 1.0000 0.9907
0.250 -0.0442 0.04575 0.03601 0.0027 1.0000 0.9927
0.500 -0.0286 0.04583 0.03605 0.0023 1.0000 0.9957
0.750 -0.0089 0.04614 0.03633 0.0011 1.0000 0.9998
1.000 -0.0056 0.04613 0.03632 0.0030 1.0000 1.0000
1.250 -0.0033 0.04610 0.03629 0.0050 1.0000 1.0000
1.500 -0.0014 0.04609 0.03628 0.0070 1.0000 1.0000
1.750 0.0001 0.04608 0.03628 0.0090 1.0000 1.0000
2.000 0.0010 0.04608 0.03630 0.0111 1.0000 1.0000
2.250 0.0013 0.04610 0.03634 0.0131 1.0000 1.0000
2.500 0.0903 0.04811 0.03837 -0.0011 0.9652 1.0000
2.750 0.1397 0.04894 0.03921 -0.0074 0.9385 1.0000
3.000 0.1808 0.04956 0.03986 -0.0117 0.9151 1.0000
3.250 0.2105 0.04988 0.04022 -0.0139 0.8926 1.0000
3.500 0.2492 0.05034 0.04072 -0.0173 0.8718 1.0000
3.750 0.2689 0.05049 0.04091 -0.0176 0.8498 1.0000
4.000 0.2910 0.05076 0.04123 -0.0182 0.8293 1.0000
4.250 0.3179 0.05100 0.04152 -0.0192 0.8091 1.0000
4.500 0.3464 0.05115 0.04173 -0.0202 0.7894 1.0000
4.750 0.3780 0.05108 0.04172 -0.0212 0.7701 1.0000
5.000 0.3851 0.05109 0.04176 -0.0188 0.7493 1.0000
5.250 0.4029 0.05059 0.04131 -0.0173 0.7264 1.0000
5.500 0.4446 0.04932 0.04010 -0.0180 0.7050 1.0000
5.750 0.4837 0.04800 0.03886 -0.0182 0.6862 1.0000
6.000 0.5048 0.04750 0.03843 -0.0169 0.6693 1.0000
6.250 0.5297 0.04695 0.03795 -0.0159 0.6523 1.0000
6.500 0.5616 0.04608 0.03718 -0.0155 0.6350 1.0000
6.750 0.6034 0.04445 0.03568 -0.0156 0.6175 1.0000
7.000 0.6604 0.04137 0.03276 -0.0161 0.6000 1.0000
7.250 0.7217 0.03768 0.02923 -0.0165 0.5757 1.0000
7.500 0.8508 0.03134 0.02286 -0.0240 0.5283 1.0000
7.750 0.8787 0.03099 0.02231 -0.0227 0.4829 1.0000
8.000 0.9204 0.03122 0.02216 -0.0238 0.4321 1.0000
8.250 0.9710 0.03252 0.02299 -0.0273 0.3862 1.0000
8.500 1.0071 0.03406 0.02423 -0.0290 0.3536 1.0000
8.750 1.0228 0.03549 0.02571 -0.0277 0.3324 1.0000
9.000 1.0433 0.03701 0.02723 -0.0271 0.3132 1.0000
9.250 1.0670 0.03862 0.02883 -0.0271 0.2966 1.0000
9.500 1.0940 0.04022 0.03041 -0.0276 0.2816 1.0000
9.750 1.1022 0.04194 0.03239 -0.0254 0.2710 1.0000
10.000 1.1201 0.04385 0.03441 -0.0248 0.2606 1.0000
10.250 1.1538 0.04558 0.03601 -0.0263 0.2490 1.0000
10.500 1.1449 0.04787 0.03875 -0.0221 0.2433 1.0000
10.750 1.1723 0.04975 0.04064 -0.0227 0.2342 1.0000
11.000 1.1644 0.05237 0.04357 -0.0190 0.2296 1.0000
11.250 1.1503 0.05504 0.04654 -0.0147 0.2256 1.0000
11.500 1.1893 0.05714 0.04853 -0.0169 0.2164 1.0000
11.750 1.1601 0.06037 0.05211 -0.0114 0.2150 1.0000
12.000 1.1271 0.06425 0.05629 -0.0066 0.2140 1.0000
12.250 1.0879 0.06903 0.06134 -0.0028 0.2139 1.0000
12.500 1.0419 0.07517 0.06769 -0.0004 0.2147 1.0000
12.750 0.9898 0.08328 0.07595 -0.0001 0.2162 1.0000
|
Polar data table (+)
Polar graphs
<< Back to DORNIER A-5 AIRFOIL (doa5-il)