Davis AIRFOIL (davis-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: Davis AIRFOIL (davis-il) Reynolds number: 100,000 Max Cl/Cd: 54.99 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-davis-il-100000-n5.txt Download as CSV file: xf-davis-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: Davis 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4008 0.08966 0.08494 -0.0466 1.0000 0.0266
-9.250 -0.4107 0.08594 0.08130 -0.0466 1.0000 0.0264
-9.000 -0.4258 0.08167 0.07712 -0.0466 1.0000 0.0265
-8.750 -0.4402 0.07852 0.07407 -0.0458 1.0000 0.0262
-8.500 -0.4619 0.07497 0.07063 -0.0448 1.0000 0.0261
-8.250 -0.4865 0.07146 0.06726 -0.0439 1.0000 0.0259
-8.000 -0.5050 0.06513 0.06094 -0.0469 0.9991 0.0255
-7.750 -0.5025 0.05325 0.04862 -0.0597 0.9888 0.0251
-7.500 -0.4941 0.04555 0.04027 -0.0657 0.9801 0.0254
-7.250 -0.4793 0.04011 0.03413 -0.0684 0.9727 0.0258
-7.000 -0.4573 0.03608 0.02937 -0.0702 0.9674 0.0264
-6.750 -0.4372 0.03276 0.02553 -0.0707 0.9609 0.0274
-6.500 -0.4106 0.03075 0.02335 -0.0720 0.9560 0.0298
-6.250 -0.3839 0.02896 0.02123 -0.0726 0.9497 0.0324
-6.000 -0.3540 0.02702 0.01890 -0.0734 0.9448 0.0344
-5.500 -0.2948 0.02401 0.01549 -0.0749 0.9332 0.0418
-5.250 -0.2669 0.02288 0.01414 -0.0750 0.9256 0.0455
-5.000 -0.2326 0.02175 0.01290 -0.0766 0.9212 0.0524
-4.750 -0.2072 0.02098 0.01198 -0.0761 0.9125 0.0591
-4.500 -0.1727 0.02011 0.01107 -0.0776 0.9087 0.0720
-4.250 -0.1495 0.01942 0.01042 -0.0768 0.9005 0.0883
-4.000 -0.1171 0.01877 0.00989 -0.0779 0.8960 0.1321
-3.750 -0.0867 0.01836 0.00949 -0.0785 0.8906 0.1668
-3.500 -0.0593 0.01809 0.00919 -0.0785 0.8836 0.1922
-3.250 -0.0244 0.01779 0.00887 -0.0799 0.8797 0.2237
-3.000 0.0004 0.01760 0.00869 -0.0794 0.8718 0.2541
-2.750 0.0320 0.01725 0.00839 -0.0801 0.8666 0.2862
-2.500 0.0612 0.01695 0.00810 -0.0804 0.8605 0.3133
-2.250 0.0890 0.01668 0.00788 -0.0803 0.8534 0.3428
-2.000 0.1232 0.01633 0.00759 -0.0814 0.8494 0.3814
-1.750 0.1467 0.01616 0.00748 -0.0806 0.8405 0.4195
-1.500 0.1804 0.01578 0.00720 -0.0816 0.8354 0.4706
-1.250 0.2058 0.01548 0.00705 -0.0810 0.8264 0.5266
-1.000 0.2370 0.01486 0.00680 -0.0811 0.8201 0.6271
-0.750 0.2901 0.01427 0.00683 -0.0850 0.8129 0.8779
-0.500 0.3701 0.01401 0.00644 -0.0954 0.8068 0.9907
-0.250 0.4028 0.01398 0.00627 -0.0964 0.7964 1.0000
0.000 0.4322 0.01391 0.00605 -0.0965 0.7865 1.0000
0.250 0.4608 0.01387 0.00587 -0.0964 0.7760 1.0000
0.500 0.4863 0.01388 0.00578 -0.0958 0.7640 1.0000
0.750 0.5128 0.01389 0.00567 -0.0954 0.7515 1.0000
1.000 0.5395 0.01393 0.00560 -0.0949 0.7395 1.0000
1.250 0.5667 0.01397 0.00555 -0.0947 0.7283 1.0000
1.500 0.5923 0.01406 0.00556 -0.0940 0.7165 1.0000
1.750 0.6156 0.01419 0.00563 -0.0931 0.7039 1.0000
2.000 0.6391 0.01431 0.00571 -0.0921 0.6909 1.0000
2.250 0.6625 0.01444 0.00580 -0.0911 0.6776 1.0000
2.500 0.6854 0.01458 0.00590 -0.0900 0.6637 1.0000
2.750 0.7079 0.01471 0.00602 -0.0889 0.6490 1.0000
3.000 0.7303 0.01486 0.00614 -0.0877 0.6340 1.0000
3.250 0.7525 0.01500 0.00627 -0.0865 0.6185 1.0000
3.500 0.7750 0.01516 0.00639 -0.0853 0.6024 1.0000
3.750 0.7978 0.01532 0.00653 -0.0842 0.5862 1.0000
4.000 0.8207 0.01551 0.00667 -0.0831 0.5698 1.0000
4.250 0.8434 0.01573 0.00683 -0.0820 0.5534 1.0000
4.500 0.8660 0.01600 0.00707 -0.0809 0.5374 1.0000
4.750 0.8883 0.01630 0.00733 -0.0798 0.5215 1.0000
5.000 0.9101 0.01663 0.00765 -0.0786 0.5053 1.0000
5.250 0.9311 0.01697 0.00802 -0.0774 0.4886 1.0000
5.500 0.9518 0.01732 0.00841 -0.0760 0.4716 1.0000
5.750 0.9722 0.01768 0.00881 -0.0747 0.4552 1.0000
6.000 0.9929 0.01806 0.00922 -0.0734 0.4398 1.0000
6.250 1.0137 0.01845 0.00965 -0.0721 0.4260 1.0000
6.500 1.0345 0.01886 0.01013 -0.0709 0.4134 1.0000
6.750 1.0554 0.01928 0.01060 -0.0697 0.4020 1.0000
7.000 1.0757 0.01973 0.01112 -0.0684 0.3901 1.0000
7.250 1.0925 0.02021 0.01158 -0.0665 0.3738 1.0000
7.500 1.1051 0.02076 0.01207 -0.0639 0.3512 1.0000
8.000 1.1298 0.02201 0.01327 -0.0589 0.3116 1.0000
8.250 1.1387 0.02260 0.01392 -0.0558 0.2874 1.0000
8.500 1.1441 0.02327 0.01458 -0.0522 0.2562 1.0000
8.750 1.1472 0.02418 0.01534 -0.0486 0.2120 1.0000
9.000 1.1366 0.02615 0.01662 -0.0437 0.1162 1.0000
9.250 1.1289 0.02851 0.01854 -0.0395 0.0652 1.0000
9.500 1.1284 0.03041 0.02031 -0.0364 0.0430 1.0000
9.750 1.1303 0.03213 0.02206 -0.0336 0.0357 1.0000
10.000 1.1317 0.03390 0.02390 -0.0310 0.0315 1.0000
10.250 1.1322 0.03577 0.02588 -0.0287 0.0287 1.0000
10.500 1.1332 0.03766 0.02794 -0.0266 0.0269 1.0000
10.750 1.1317 0.03984 0.03033 -0.0246 0.0253 1.0000
11.000 1.1289 0.04226 0.03290 -0.0229 0.0243 1.0000
11.250 1.1251 0.04489 0.03568 -0.0215 0.0237 1.0000
11.500 1.1185 0.04794 0.03885 -0.0203 0.0230 1.0000
11.750 1.1121 0.05109 0.04210 -0.0193 0.0225 1.0000
12.000 1.1114 0.05380 0.04498 -0.0185 0.0219 1.0000
12.250 1.1118 0.05648 0.04780 -0.0179 0.0212 1.0000
12.500 1.1128 0.05915 0.05063 -0.0173 0.0203 1.0000
12.750 1.1149 0.06178 0.05338 -0.0167 0.0194 1.0000
13.000 1.1191 0.06424 0.05595 -0.0161 0.0189 1.0000
13.250 1.1247 0.06662 0.05844 -0.0153 0.0183 1.0000
13.500 1.1313 0.06900 0.06095 -0.0146 0.0178 1.0000
13.750 1.1374 0.07156 0.06364 -0.0140 0.0174 1.0000
14.000 1.1427 0.07433 0.06657 -0.0135 0.0170 1.0000
14.250 1.1454 0.07747 0.06986 -0.0134 0.0166 1.0000
14.500 1.1459 0.08092 0.07352 -0.0136 0.0162 1.0000
14.750 1.1443 0.08477 0.07754 -0.0140 0.0158 1.0000
15.000 1.1409 0.08897 0.08193 -0.0147 0.0156 1.0000
15.250 1.1319 0.09427 0.08744 -0.0158 0.0152 1.0000
15.500 1.1219 0.09958 0.09293 -0.0177 0.0150 1.0000
15.750 1.1103 0.10519 0.09878 -0.0202 0.0150 1.0000
16.000 1.0990 0.11105 0.10485 -0.0230 0.0149 1.0000
16.250 1.0863 0.11737 0.11139 -0.0263 0.0149 1.0000
16.500 1.0728 0.12418 0.11839 -0.0301 0.0149 1.0000
16.750 1.0590 0.13139 0.12579 -0.0344 0.0150 1.0000
17.000 1.0445 0.13911 0.13368 -0.0392 0.0150 1.0000
17.250 1.0306 0.14708 0.14179 -0.0441 0.0150 1.0000
17.500 1.0170 0.15553 0.15042 -0.0496 0.0152 1.0000
17.750 0.9957 0.16770 0.16274 -0.0576 0.0158 1.0000
|
Polar data table (+)
Polar graphs
<< Back to Davis AIRFOIL (davis-il)