RAF 38 AIRFOIL (raf38-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 38 AIRFOIL (raf38-il) Reynolds number: 100,000 Max Cl/Cd: 51.64 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf38-il-100000.txt Download as CSV file: xf-raf38-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: RAF 38 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4205 0.10149 0.09627 -0.0315 1.0000 0.1573
-9.750 -0.4088 0.09877 0.09355 -0.0307 1.0000 0.1635
-9.500 -0.4635 0.09540 0.09041 -0.0384 1.0000 0.1693
-9.250 -0.4214 0.09144 0.08636 -0.0334 1.0000 0.1732
-9.000 -0.4137 0.08859 0.08353 -0.0327 1.0000 0.1786
-8.750 -0.4939 0.08495 0.08019 -0.0398 1.0000 0.1850
-8.500 -0.4340 0.08140 0.07652 -0.0346 1.0000 0.1895
-8.250 -0.6093 0.05890 0.05363 -0.0445 1.0000 0.1097
-8.000 -0.6554 0.05222 0.04631 -0.0384 1.0000 0.0966
-7.750 -0.6619 0.04926 0.04327 -0.0348 1.0000 0.0961
-7.500 -0.6680 0.04647 0.04034 -0.0312 1.0000 0.0955
-7.250 -0.6721 0.04364 0.03729 -0.0278 1.0000 0.0948
-7.000 -0.6733 0.04091 0.03424 -0.0246 1.0000 0.0947
-6.750 -0.6721 0.03856 0.03141 -0.0215 1.0000 0.0958
-6.500 -0.6514 0.03528 0.02760 -0.0219 0.9957 0.0979
-6.250 -0.6178 0.03300 0.02521 -0.0242 0.9895 0.1012
-6.000 -0.5819 0.03124 0.02312 -0.0265 0.9833 0.1061
-5.750 -0.5504 0.02952 0.02083 -0.0276 0.9758 0.1098
-5.500 -0.5122 0.02778 0.01901 -0.0302 0.9704 0.1153
-5.250 -0.4806 0.02666 0.01771 -0.0313 0.9626 0.1201
-5.000 -0.4412 0.02549 0.01628 -0.0337 0.9570 0.1263
-4.750 -0.4092 0.02459 0.01541 -0.0348 0.9495 0.1323
-4.500 -0.3706 0.02377 0.01443 -0.0369 0.9432 0.1404
-4.250 -0.3358 0.02300 0.01376 -0.0385 0.9364 0.1486
-4.000 -0.3007 0.02226 0.01306 -0.0400 0.9289 0.1596
-3.750 -0.2573 0.02146 0.01238 -0.0429 0.9239 0.1787
-3.500 -0.2325 0.02073 0.01184 -0.0425 0.9140 0.2059
-3.250 -0.1964 0.01907 0.01107 -0.0443 0.9090 0.3340
-3.000 -0.1805 0.01844 0.01106 -0.0421 0.8985 0.4617
-2.750 -0.1421 0.01776 0.01093 -0.0434 0.8936 0.5950
-2.500 -0.1198 0.01762 0.01113 -0.0415 0.8845 0.6911
-2.250 -0.0764 0.01747 0.01121 -0.0429 0.8795 0.7900
-2.000 -0.0101 0.01760 0.01137 -0.0479 0.8771 0.8680
-1.750 0.0671 0.01802 0.01165 -0.0552 0.8741 0.9144
-1.500 0.1372 0.01823 0.01168 -0.0623 0.8685 0.9400
-1.250 0.2123 0.01807 0.01134 -0.0709 0.8638 0.9572
-1.000 0.2851 0.01773 0.01083 -0.0793 0.8591 0.9736
-0.750 0.3632 0.01727 0.01027 -0.0894 0.8511 0.9962
-0.500 0.3959 0.01703 0.00992 -0.0909 0.8422 1.0000
-0.250 0.4101 0.01712 0.00995 -0.0890 0.8301 1.0000
0.000 0.4248 0.01724 0.01002 -0.0871 0.8185 1.0000
0.250 0.4448 0.01724 0.00993 -0.0858 0.8091 1.0000
0.500 0.4599 0.01741 0.01006 -0.0838 0.7975 1.0000
0.750 0.4757 0.01761 0.01022 -0.0818 0.7863 1.0000
1.000 0.4982 0.01760 0.01013 -0.0805 0.7778 1.0000
1.250 0.5122 0.01790 0.01042 -0.0783 0.7655 1.0000
1.500 0.5294 0.01814 0.01065 -0.0764 0.7547 1.0000
1.750 0.5519 0.01819 0.01063 -0.0751 0.7459 1.0000
2.000 0.5674 0.01852 0.01097 -0.0730 0.7338 1.0000
2.250 0.5870 0.01874 0.01117 -0.0713 0.7238 1.0000
2.500 0.6090 0.01883 0.01123 -0.0699 0.7142 1.0000
2.750 0.6262 0.01913 0.01154 -0.0680 0.7023 1.0000
3.000 0.6479 0.01919 0.01157 -0.0664 0.6915 1.0000
3.250 0.6721 0.01911 0.01145 -0.0650 0.6811 1.0000
3.500 0.6902 0.01934 0.01171 -0.0631 0.6685 1.0000
3.750 0.7112 0.01951 0.01189 -0.0615 0.6579 1.0000
4.000 0.7367 0.01950 0.01186 -0.0605 0.6488 1.0000
4.250 0.7546 0.01984 0.01226 -0.0587 0.6370 1.0000
4.500 0.7772 0.02000 0.01243 -0.0574 0.6271 1.0000
4.750 0.8023 0.01999 0.01243 -0.0563 0.6170 1.0000
5.000 0.8218 0.02016 0.01266 -0.0545 0.6041 1.0000
5.250 0.8437 0.02015 0.01267 -0.0529 0.5908 1.0000
5.500 0.8667 0.01999 0.01251 -0.0513 0.5764 1.0000
5.750 0.8894 0.01981 0.01234 -0.0497 0.5610 1.0000
6.000 0.9115 0.01968 0.01222 -0.0480 0.5450 1.0000
6.250 0.9328 0.01957 0.01212 -0.0462 0.5281 1.0000
6.500 0.9536 0.01946 0.01201 -0.0443 0.5099 1.0000
6.750 0.9744 0.01941 0.01194 -0.0425 0.4914 1.0000
7.000 0.9935 0.01950 0.01204 -0.0405 0.4724 1.0000
7.250 1.0098 0.01970 0.01229 -0.0382 0.4512 1.0000
7.500 1.0267 0.01988 0.01246 -0.0359 0.4287 1.0000
7.750 1.0396 0.02017 0.01278 -0.0330 0.4021 1.0000
8.000 1.0498 0.02051 0.01311 -0.0297 0.3701 1.0000
8.250 1.0548 0.02100 0.01352 -0.0257 0.3278 1.0000
8.500 1.0509 0.02198 0.01415 -0.0206 0.2703 1.0000
8.750 1.0406 0.02359 0.01528 -0.0150 0.2173 1.0000
9.000 1.0322 0.02518 0.01653 -0.0098 0.1851 1.0000
9.250 1.0307 0.02665 0.01779 -0.0059 0.1638 1.0000
9.500 1.0340 0.02803 0.01904 -0.0028 0.1489 1.0000
9.750 1.0403 0.02943 0.02031 -0.0003 0.1374 1.0000
10.000 1.0495 0.03078 0.02157 0.0017 0.1274 1.0000
10.250 1.0622 0.03215 0.02296 0.0034 0.1183 1.0000
10.500 1.0791 0.03367 0.02437 0.0046 0.1101 1.0000
10.750 1.0961 0.03510 0.02580 0.0057 0.1029 1.0000
11.000 1.1194 0.03691 0.02762 0.0062 0.0959 1.0000
11.250 1.1407 0.03859 0.02934 0.0068 0.0902 1.0000
11.500 1.1682 0.04098 0.03177 0.0065 0.0846 1.0000
11.750 1.1828 0.04295 0.03396 0.0077 0.0809 1.0000
12.000 1.2052 0.04521 0.03625 0.0078 0.0770 1.0000
12.250 1.2218 0.04843 0.03965 0.0083 0.0739 1.0000
12.500 1.2163 0.05049 0.04206 0.0115 0.0724 1.0000
12.750 1.2125 0.05306 0.04494 0.0139 0.0712 1.0000
13.000 1.2066 0.05581 0.04797 0.0162 0.0700 1.0000
13.250 1.1989 0.05878 0.05119 0.0181 0.0690 1.0000
13.500 1.1902 0.06187 0.05451 0.0196 0.0682 1.0000
13.750 1.1950 0.06464 0.05734 0.0203 0.0663 1.0000
14.000 1.1904 0.06812 0.06096 0.0210 0.0652 1.0000
14.250 1.1819 0.07263 0.06561 0.0214 0.0643 1.0000
14.500 1.1615 0.07708 0.07030 0.0217 0.0642 1.0000
14.750 1.1363 0.08169 0.07515 0.0213 0.0644 1.0000
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