RAF 15 AIRFOIL (raf15-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 15 AIRFOIL (raf15-il) Reynolds number: 100,000 Max Cl/Cd: 43.88 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf15-il-100000.txt Download as CSV file: xf-raf15-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: RAF 15 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4024 0.09131 0.08651 -0.0101 1.0000 0.0998
-8.250 -0.4121 0.08851 0.08377 -0.0119 1.0000 0.1030
-8.000 -0.5014 0.09494 0.08996 -0.0064 1.0000 0.0958
-7.750 -0.5028 0.09210 0.08717 -0.0067 1.0000 0.0993
-7.500 -0.5105 0.08952 0.08465 -0.0081 1.0000 0.1021
-7.250 -0.5266 0.08842 0.08347 -0.0169 1.0000 0.1047
-7.000 -0.5138 0.08229 0.07748 -0.0120 1.0000 0.1066
-6.750 -0.5041 0.07892 0.07414 -0.0096 1.0000 0.1101
-6.500 -0.4993 0.07589 0.07110 -0.0105 1.0000 0.1151
-6.250 -0.4983 0.07244 0.06752 -0.0149 1.0000 0.1199
-6.000 -0.4882 0.06869 0.06386 -0.0119 1.0000 0.1226
-5.750 -0.4786 0.06570 0.06082 -0.0113 1.0000 0.1284
-5.500 -0.4705 0.06217 0.05714 -0.0130 1.0000 0.1349
-5.250 -0.4593 0.05923 0.05425 -0.0107 1.0000 0.1403
-5.000 -0.4483 0.05609 0.05097 -0.0111 1.0000 0.1500
-4.750 -0.4344 0.05374 0.04853 -0.0101 1.0000 0.1595
-4.500 -0.4229 0.05024 0.04495 -0.0091 1.0000 0.1648
-4.250 -0.4039 0.03796 0.03145 -0.0101 1.0000 0.0922
-4.000 -0.3897 0.03162 0.02427 -0.0071 1.0000 0.0763
-3.750 -0.3733 0.02842 0.02069 -0.0048 1.0000 0.0753
-3.500 -0.3546 0.02600 0.01766 -0.0025 1.0000 0.0768
-3.250 -0.3342 0.02359 0.01480 -0.0005 1.0000 0.0779
-3.000 -0.3121 0.02162 0.01262 0.0010 1.0000 0.0802
-2.750 -0.2894 0.02053 0.01139 0.0024 1.0000 0.0860
-2.500 -0.2650 0.01926 0.00983 0.0038 1.0000 0.0918
-2.250 -0.2413 0.01824 0.00884 0.0049 1.0000 0.0989
-2.000 -0.2181 0.01748 0.00809 0.0062 1.0000 0.1116
-1.750 -0.1946 0.01669 0.00733 0.0075 1.0000 0.1298
-1.500 -0.1735 0.01631 0.00719 0.0092 1.0000 0.1789
-1.250 -0.1544 0.01623 0.00713 0.0111 1.0000 0.2359
-1.000 -0.1342 0.01604 0.00701 0.0126 1.0000 0.2671
-0.750 -0.1136 0.01581 0.00687 0.0138 1.0000 0.2993
-0.500 -0.0913 0.01549 0.00668 0.0148 1.0000 0.3312
-0.250 -0.0652 0.01511 0.00644 0.0150 1.0000 0.3619
0.000 0.0989 0.01305 0.00617 -0.0129 1.0000 1.0000
0.250 0.1184 0.01320 0.00618 -0.0112 1.0000 1.0000
0.500 0.1370 0.01339 0.00626 -0.0095 1.0000 1.0000
0.750 0.1547 0.01362 0.00643 -0.0078 1.0000 1.0000
1.000 0.1716 0.01391 0.00668 -0.0061 1.0000 1.0000
1.250 0.2064 0.01423 0.00701 -0.0083 0.9943 1.0000
1.500 0.2836 0.01402 0.00686 -0.0185 0.9672 1.0000
1.750 0.3494 0.01341 0.00635 -0.0256 0.9281 1.0000
2.000 0.4504 0.01245 0.00546 -0.0390 0.8558 1.0000
2.250 0.5518 0.01258 0.00480 -0.0527 0.6610 1.0000
2.500 0.5790 0.01324 0.00499 -0.0521 0.5879 1.0000
2.750 0.6020 0.01382 0.00527 -0.0508 0.5424 1.0000
3.000 0.6243 0.01435 0.00561 -0.0494 0.5080 1.0000
3.250 0.6464 0.01486 0.00597 -0.0480 0.4791 1.0000
3.500 0.6685 0.01535 0.00634 -0.0466 0.4536 1.0000
3.750 0.6905 0.01585 0.00673 -0.0452 0.4301 1.0000
4.000 0.7122 0.01633 0.00721 -0.0438 0.4066 1.0000
4.250 0.7339 0.01687 0.00763 -0.0424 0.3844 1.0000
4.500 0.7540 0.01732 0.00807 -0.0407 0.3592 1.0000
4.750 0.7732 0.01774 0.00843 -0.0388 0.3329 1.0000
5.000 0.7923 0.01816 0.00878 -0.0369 0.3076 1.0000
5.250 0.8111 0.01853 0.00918 -0.0350 0.2821 1.0000
5.500 0.8297 0.01891 0.00958 -0.0331 0.2557 1.0000
5.750 0.8478 0.01934 0.01000 -0.0310 0.2275 1.0000
6.000 0.8648 0.01995 0.01050 -0.0288 0.1952 1.0000
6.250 0.8809 0.02099 0.01132 -0.0265 0.1632 1.0000
6.500 0.8986 0.02211 0.01239 -0.0245 0.1391 1.0000
6.750 0.9183 0.02343 0.01369 -0.0228 0.1256 1.0000
7.000 0.9394 0.02493 0.01524 -0.0214 0.1165 1.0000
7.250 0.9608 0.02646 0.01667 -0.0203 0.1087 1.0000
7.500 0.9802 0.02781 0.01836 -0.0184 0.1020 1.0000
7.750 1.0012 0.02943 0.02004 -0.0172 0.0971 1.0000
8.000 1.0195 0.03159 0.02245 -0.0155 0.0927 1.0000
8.250 1.0356 0.03323 0.02442 -0.0133 0.0877 1.0000
8.500 1.0527 0.03521 0.02651 -0.0117 0.0837 1.0000
8.750 1.0633 0.03798 0.02968 -0.0092 0.0804 1.0000
9.000 1.0695 0.04052 0.03278 -0.0058 0.0771 1.0000
9.250 1.0744 0.04359 0.03628 -0.0027 0.0751 1.0000
9.500 1.0883 0.04560 0.03830 -0.0011 0.0713 1.0000
9.750 1.0888 0.04977 0.04268 0.0014 0.0692 1.0000
10.000 1.0789 0.05328 0.04673 0.0054 0.0683 1.0000
10.250 1.0650 0.05743 0.05130 0.0091 0.0680 1.0000
10.500 1.0499 0.06169 0.05584 0.0121 0.0683 1.0000
10.750 1.0292 0.06593 0.06030 0.0152 0.0685 1.0000
11.000 1.0086 0.07009 0.06458 0.0176 0.0689 1.0000
11.250 0.9893 0.07500 0.06959 0.0185 0.0693 1.0000
11.500 0.8410 0.10882 0.10355 -0.0080 0.0939 1.0000
11.750 0.8125 0.13249 0.12715 -0.0186 0.1496 1.0000
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Polar data table (+)
Polar graphs
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