RAE 103 AIRFOIL (rae103-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 103 AIRFOIL (rae103-il) Reynolds number: 1,000,000 Max Cl/Cd: 62.23 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae103-il-1000000.txt Download as CSV file: xf-rae103-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 103 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.750 -0.7714 0.09112 0.08954 -0.0121 1.0000 0.0076
-12.500 -0.7666 0.08745 0.08588 -0.0144 1.0000 0.0077
-12.250 -0.9275 0.04775 0.04541 -0.0384 1.0000 0.0065
-11.750 -0.9505 0.04101 0.03832 -0.0381 1.0000 0.0067
-11.500 -0.9657 0.03792 0.03503 -0.0359 1.0000 0.0067
-11.000 -0.9968 0.03256 0.02917 -0.0273 1.0000 0.0069
-10.750 -1.0044 0.02949 0.02578 -0.0236 1.0000 0.0069
-10.500 -1.0095 0.02610 0.02204 -0.0199 1.0000 0.0071
-10.250 -1.0020 0.02412 0.01985 -0.0174 1.0000 0.0073
-10.000 -0.9886 0.02287 0.01848 -0.0156 1.0000 0.0075
-9.750 -0.9751 0.02145 0.01689 -0.0137 1.0000 0.0076
-9.500 -0.9583 0.02053 0.01588 -0.0121 1.0000 0.0078
-9.250 -0.9416 0.01953 0.01475 -0.0105 1.0000 0.0080
-9.000 -0.9237 0.01871 0.01383 -0.0091 1.0000 0.0082
-8.750 -0.9060 0.01781 0.01283 -0.0075 1.0000 0.0085
-8.500 -0.8878 0.01698 0.01190 -0.0059 1.0000 0.0088
-8.250 -0.8686 0.01633 0.01118 -0.0045 1.0000 0.0092
-8.000 -0.8489 0.01580 0.01058 -0.0032 1.0000 0.0096
-7.750 -0.8285 0.01537 0.01010 -0.0019 1.0000 0.0099
-7.500 -0.8141 0.01422 0.00882 0.0005 1.0000 0.0103
-7.250 -0.8006 0.01316 0.00765 0.0030 1.0000 0.0110
-7.000 -0.7824 0.01269 0.00717 0.0048 1.0000 0.0117
-6.750 -0.7642 0.01229 0.00673 0.0065 1.0000 0.0123
-6.500 -0.7305 0.01181 0.00621 0.0049 0.9981 0.0134
-6.250 -0.6964 0.01144 0.00581 0.0033 0.9956 0.0142
-6.000 -0.6653 0.01068 0.00495 0.0022 0.9923 0.0160
-5.750 -0.6326 0.01030 0.00458 0.0009 0.9885 0.0182
-5.500 -0.5981 0.01003 0.00430 -0.0008 0.9853 0.0201
-5.250 -0.5655 0.00951 0.00374 -0.0020 0.9812 0.0242
-5.000 -0.5330 0.00922 0.00345 -0.0032 0.9754 0.0281
-4.750 -0.4980 0.00883 0.00309 -0.0050 0.9707 0.0369
-4.500 -0.4660 0.00849 0.00280 -0.0061 0.9617 0.0498
-4.250 -0.4335 0.00812 0.00252 -0.0073 0.9516 0.0749
-4.000 -0.4046 0.00770 0.00225 -0.0078 0.9371 0.1184
-3.750 -0.3794 0.00726 0.00201 -0.0075 0.9194 0.1760
-3.500 -0.3569 0.00675 0.00176 -0.0066 0.9008 0.2585
-3.250 -0.3361 0.00617 0.00152 -0.0055 0.8827 0.3623
-3.000 -0.3162 0.00556 0.00129 -0.0041 0.8656 0.4812
-2.750 -0.2931 0.00528 0.00121 -0.0032 0.8506 0.5579
-2.500 -0.2678 0.00519 0.00115 -0.0026 0.8373 0.5927
-2.250 -0.2417 0.00514 0.00110 -0.0022 0.8250 0.6146
-2.000 -0.2153 0.00511 0.00106 -0.0019 0.8134 0.6329
-1.750 -0.1886 0.00508 0.00103 -0.0016 0.8024 0.6490
-1.500 -0.1618 0.00506 0.00100 -0.0013 0.7921 0.6634
-1.250 -0.1351 0.00506 0.00097 -0.0010 0.7823 0.6765
-1.000 -0.1081 0.00504 0.00096 -0.0008 0.7723 0.6887
-0.750 -0.0810 0.00504 0.00094 -0.0006 0.7627 0.7002
-0.250 -0.0270 0.00504 0.00093 -0.0002 0.7428 0.7222
0.000 0.0000 0.00503 0.00093 0.0000 0.7324 0.7324
0.250 0.0271 0.00504 0.00093 0.0002 0.7221 0.7428
0.750 0.0810 0.00504 0.00094 0.0006 0.7003 0.7628
1.000 0.1081 0.00504 0.00096 0.0008 0.6888 0.7724
1.250 0.1351 0.00506 0.00097 0.0010 0.6765 0.7823
1.500 0.1618 0.00506 0.00100 0.0013 0.6634 0.7921
1.750 0.1886 0.00508 0.00103 0.0016 0.6494 0.8024
2.000 0.2153 0.00511 0.00106 0.0019 0.6331 0.8134
2.250 0.2417 0.00514 0.00110 0.0022 0.6147 0.8250
2.500 0.2677 0.00519 0.00115 0.0026 0.5919 0.8373
2.750 0.2931 0.00528 0.00121 0.0032 0.5576 0.8507
3.000 0.3159 0.00558 0.00129 0.0041 0.4762 0.8656
3.250 0.3361 0.00617 0.00152 0.0055 0.3618 0.8827
3.500 0.3571 0.00673 0.00176 0.0066 0.2619 0.9007
3.750 0.3794 0.00726 0.00201 0.0075 0.1759 0.9195
4.000 0.4045 0.00771 0.00226 0.0078 0.1167 0.9372
4.250 0.4335 0.00812 0.00252 0.0073 0.0750 0.9516
4.500 0.4659 0.00850 0.00280 0.0061 0.0497 0.9617
4.750 0.4979 0.00883 0.00309 0.0050 0.0370 0.9707
5.000 0.5329 0.00922 0.00345 0.0032 0.0281 0.9754
5.250 0.5654 0.00951 0.00374 0.0020 0.0242 0.9813
5.500 0.5980 0.01003 0.00430 0.0008 0.0201 0.9853
5.750 0.6325 0.01031 0.00459 -0.0009 0.0182 0.9886
6.000 0.6652 0.01069 0.00496 -0.0022 0.0161 0.9923
6.250 0.6963 0.01145 0.00582 -0.0033 0.0142 0.9956
6.500 0.7303 0.01184 0.00624 -0.0049 0.0133 0.9981
6.750 0.7640 0.01231 0.00675 -0.0065 0.0124 1.0000
7.000 0.7825 0.01268 0.00715 -0.0048 0.0117 1.0000
7.250 0.8008 0.01314 0.00764 -0.0031 0.0112 1.0000
7.500 0.8160 0.01400 0.00858 -0.0008 0.0105 1.0000
7.750 0.8293 0.01527 0.00998 0.0017 0.0099 1.0000
8.000 0.8486 0.01584 0.01062 0.0032 0.0097 1.0000
8.250 0.8689 0.01629 0.01113 0.0045 0.0092 1.0000
8.500 0.8879 0.01696 0.01189 0.0059 0.0088 1.0000
8.750 0.9059 0.01782 0.01284 0.0075 0.0085 1.0000
9.000 0.9243 0.01860 0.01371 0.0090 0.0082 1.0000
9.250 0.9418 0.01948 0.01470 0.0105 0.0080 1.0000
9.500 0.9591 0.02037 0.01570 0.0120 0.0078 1.0000
9.750 0.9755 0.02134 0.01676 0.0136 0.0076 1.0000
10.000 0.9896 0.02264 0.01821 0.0154 0.0075 1.0000
10.250 1.0029 0.02392 0.01961 0.0173 0.0072 1.0000
10.500 1.0075 0.02635 0.02231 0.0201 0.0071 1.0000
10.750 1.0029 0.02966 0.02597 0.0237 0.0069 1.0000
11.000 0.9825 0.03383 0.03057 0.0289 0.0068 1.0000
11.500 0.9608 0.03840 0.03553 0.0362 0.0067 1.0000
11.750 0.9378 0.04233 0.03970 0.0383 0.0067 1.0000
12.000 0.9342 0.04470 0.04220 0.0387 0.0067 1.0000
12.250 0.9113 0.04982 0.04753 0.0376 0.0067 1.0000
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Polar data table (+)
Polar graphs
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