RAE 103 AIRFOIL (rae103-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE 103 AIRFOIL (rae103-il) Reynolds number: 200,000 Max Cl/Cd: 39.19 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae103-il-200000-n5.txt Download as CSV file: xf-rae103-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.7014 0.08641 0.08283 -0.0177 1.0000 0.0134
-11.250 -0.7328 0.07253 0.06887 -0.0276 1.0000 0.0130
-11.000 -0.7602 0.06456 0.06075 -0.0329 1.0000 0.0129
-10.750 -0.7827 0.05906 0.05509 -0.0352 1.0000 0.0128
-10.500 -0.8058 0.05439 0.05022 -0.0353 1.0000 0.0127
-10.250 -0.8266 0.05054 0.04617 -0.0334 1.0000 0.0127
-10.000 -0.8454 0.04708 0.04246 -0.0299 1.0000 0.0127
-9.750 -0.8571 0.04333 0.03837 -0.0268 1.0000 0.0128
-9.500 -0.8625 0.03968 0.03433 -0.0238 1.0000 0.0129
-9.250 -0.8620 0.03634 0.03057 -0.0211 1.0000 0.0131
-9.000 -0.8563 0.03333 0.02713 -0.0186 1.0000 0.0134
-8.750 -0.8458 0.03081 0.02420 -0.0165 1.0000 0.0138
-8.500 -0.8318 0.02870 0.02174 -0.0146 1.0000 0.0142
-8.250 -0.8170 0.02660 0.01938 -0.0130 1.0000 0.0149
-8.000 -0.7999 0.02540 0.01810 -0.0117 1.0000 0.0157
-7.750 -0.7815 0.02454 0.01712 -0.0105 1.0000 0.0168
-7.500 -0.7626 0.02334 0.01576 -0.0091 1.0000 0.0179
-7.250 -0.7436 0.02205 0.01430 -0.0077 1.0000 0.0188
-7.000 -0.7247 0.02091 0.01301 -0.0061 1.0000 0.0197
-6.750 -0.7064 0.01987 0.01186 -0.0045 1.0000 0.0206
-6.500 -0.6902 0.01878 0.01076 -0.0027 1.0000 0.0225
-6.250 -0.6715 0.01815 0.01007 -0.0012 1.0000 0.0246
-6.000 -0.6535 0.01746 0.00931 0.0004 1.0000 0.0268
-5.750 -0.6357 0.01681 0.00856 0.0022 1.0000 0.0285
-5.500 -0.6206 0.01597 0.00774 0.0043 1.0000 0.0316
-5.250 -0.6030 0.01550 0.00723 0.0061 1.0000 0.0361
-5.000 -0.5762 0.01485 0.00655 0.0058 0.9961 0.0421
-4.750 -0.5434 0.01428 0.00592 0.0044 0.9889 0.0519
-4.500 -0.5106 0.01366 0.00538 0.0028 0.9821 0.0711
-4.250 -0.4797 0.01301 0.00489 0.0016 0.9738 0.1064
-4.000 -0.4495 0.01222 0.00441 0.0004 0.9657 0.1775
-3.750 -0.4213 0.01122 0.00395 -0.0005 0.9567 0.2979
-3.500 -0.3982 0.01016 0.00360 -0.0003 0.9453 0.4551
-3.250 -0.3713 0.00971 0.00356 -0.0001 0.9353 0.5702
-3.000 -0.3396 0.00954 0.00346 -0.0008 0.9270 0.6205
-2.750 -0.3100 0.00944 0.00338 -0.0010 0.9160 0.6536
-2.500 -0.2802 0.00936 0.00332 -0.0012 0.9055 0.6804
-2.250 -0.2503 0.00929 0.00326 -0.0014 0.8955 0.7026
-2.000 -0.2212 0.00925 0.00318 -0.0014 0.8852 0.7222
-1.750 -0.1933 0.00921 0.00313 -0.0013 0.8745 0.7386
-1.500 -0.1652 0.00917 0.00307 -0.0011 0.8642 0.7520
-1.250 -0.1370 0.00914 0.00300 -0.0010 0.8545 0.7631
-1.000 -0.1096 0.00912 0.00295 -0.0008 0.8442 0.7732
-0.750 -0.0824 0.00910 0.00290 -0.0006 0.8349 0.7831
-0.250 -0.0274 0.00907 0.00285 -0.0002 0.8169 0.8000
0.000 0.0000 0.00907 0.00285 0.0000 0.8086 0.8086
0.250 0.0275 0.00907 0.00285 0.0002 0.7999 0.8170
0.750 0.0824 0.00910 0.00290 0.0006 0.7831 0.8349
1.000 0.1097 0.00912 0.00295 0.0008 0.7732 0.8442
1.250 0.1370 0.00914 0.00300 0.0010 0.7632 0.8544
1.500 0.1652 0.00917 0.00307 0.0011 0.7519 0.8642
1.750 0.1933 0.00921 0.00313 0.0013 0.7386 0.8745
2.000 0.2213 0.00925 0.00318 0.0014 0.7223 0.8852
2.250 0.2503 0.00929 0.00326 0.0014 0.7027 0.8955
2.500 0.2802 0.00936 0.00332 0.0012 0.6802 0.9055
2.750 0.3100 0.00944 0.00338 0.0010 0.6537 0.9160
3.250 0.3713 0.00971 0.00356 0.0001 0.5700 0.9353
3.500 0.3982 0.01016 0.00360 0.0003 0.4555 0.9453
3.750 0.4212 0.01122 0.00395 0.0006 0.2974 0.9567
4.000 0.4494 0.01223 0.00441 -0.0004 0.1765 0.9657
4.250 0.4797 0.01302 0.00489 -0.0016 0.1059 0.9738
4.500 0.5105 0.01366 0.00538 -0.0028 0.0711 0.9821
4.750 0.5434 0.01428 0.00592 -0.0043 0.0521 0.9889
5.000 0.5761 0.01486 0.00653 -0.0058 0.0421 0.9960
5.250 0.6031 0.01550 0.00723 -0.0061 0.0359 1.0000
5.500 0.6206 0.01597 0.00773 -0.0043 0.0316 1.0000
5.750 0.6357 0.01681 0.00856 -0.0022 0.0285 1.0000
6.000 0.6535 0.01746 0.00931 -0.0004 0.0268 1.0000
6.250 0.6715 0.01816 0.01009 0.0013 0.0247 1.0000
6.500 0.6902 0.01877 0.01075 0.0027 0.0224 1.0000
6.750 0.7065 0.01985 0.01184 0.0045 0.0206 1.0000
7.000 0.7247 0.02091 0.01302 0.0061 0.0197 1.0000
7.250 0.7436 0.02205 0.01430 0.0077 0.0189 1.0000
7.500 0.7626 0.02336 0.01578 0.0091 0.0180 1.0000
7.750 0.7815 0.02448 0.01705 0.0105 0.0167 1.0000
8.000 0.7999 0.02554 0.01826 0.0117 0.0158 1.0000
8.250 0.8170 0.02659 0.01936 0.0130 0.0148 1.0000
8.500 0.8318 0.02874 0.02177 0.0146 0.0143 1.0000
8.750 0.8459 0.03080 0.02419 0.0165 0.0139 1.0000
9.000 0.8564 0.03334 0.02714 0.0186 0.0135 1.0000
9.250 0.8620 0.03636 0.03059 0.0211 0.0131 1.0000
9.500 0.8624 0.03974 0.03439 0.0238 0.0129 1.0000
9.750 0.8561 0.04350 0.03856 0.0269 0.0127 1.0000
10.000 0.8451 0.04716 0.04254 0.0299 0.0127 1.0000
10.250 0.8271 0.05053 0.04615 0.0333 0.0127 1.0000
10.500 0.8052 0.05452 0.05036 0.0353 0.0127 1.0000
10.750 0.7846 0.05889 0.05491 0.0351 0.0128 1.0000
11.000 0.7597 0.06476 0.06095 0.0327 0.0129 1.0000
11.250 0.7355 0.07209 0.06841 0.0279 0.0131 1.0000
11.500 0.7050 0.08527 0.08169 0.0182 0.0134 1.0000
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Polar data table (+)
Polar graphs
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