RAE 103 AIRFOIL (rae103-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 103 AIRFOIL (rae103-il) Reynolds number: 500,000 Max Cl/Cd: 54.84 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae103-il-500000.txt Download as CSV file: xf-rae103-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.5764 0.07874 0.07663 -0.0214 1.0000 0.0238
-10.500 -0.6726 0.05434 0.05198 -0.0371 1.0000 0.0219
-9.750 -0.8439 0.04431 0.04093 -0.0283 1.0000 0.0139
-9.500 -0.8559 0.04018 0.03650 -0.0249 1.0000 0.0137
-9.250 -0.8626 0.03611 0.03209 -0.0215 1.0000 0.0135
-9.000 -0.8623 0.03252 0.02813 -0.0184 1.0000 0.0135
-8.750 -0.8546 0.02986 0.02514 -0.0158 1.0000 0.0137
-8.500 -0.8411 0.02831 0.02333 -0.0137 1.0000 0.0141
-8.250 -0.8258 0.02729 0.02208 -0.0118 1.0000 0.0144
-8.000 -0.8145 0.02417 0.01861 -0.0094 1.0000 0.0145
-7.750 -0.8021 0.02063 0.01471 -0.0070 1.0000 0.0150
-7.500 -0.7852 0.01893 0.01289 -0.0053 1.0000 0.0156
-7.250 -0.7666 0.01799 0.01190 -0.0038 1.0000 0.0163
-7.000 -0.7476 0.01721 0.01105 -0.0023 1.0000 0.0172
-6.750 -0.7291 0.01639 0.01016 -0.0006 1.0000 0.0181
-6.500 -0.7110 0.01564 0.00934 0.0012 1.0000 0.0191
-6.250 -0.6910 0.01533 0.00897 0.0026 1.0000 0.0202
-6.000 -0.6794 0.01400 0.00755 0.0055 1.0000 0.0220
-5.750 -0.6642 0.01338 0.00691 0.0078 1.0000 0.0237
-5.500 -0.6449 0.01293 0.00643 0.0092 0.9995 0.0258
-5.250 -0.6080 0.01249 0.00593 0.0070 0.9961 0.0285
-5.000 -0.5745 0.01162 0.00504 0.0054 0.9917 0.0350
-4.750 -0.5395 0.01107 0.00444 0.0035 0.9870 0.0424
-4.500 -0.5029 0.01052 0.00397 0.0013 0.9834 0.0592
-4.250 -0.4711 0.00988 0.00352 0.0001 0.9765 0.1003
-4.000 -0.4373 0.00902 0.00309 -0.0019 0.9716 0.1936
-3.750 -0.4089 0.00799 0.00267 -0.0028 0.9633 0.3382
-3.500 -0.3807 0.00693 0.00235 -0.0036 0.9557 0.5163
-3.250 -0.3521 0.00664 0.00226 -0.0038 0.9443 0.5885
-3.000 -0.3220 0.00651 0.00216 -0.0041 0.9328 0.6251
-2.750 -0.2930 0.00641 0.00208 -0.0042 0.9204 0.6497
-2.500 -0.2652 0.00636 0.00200 -0.0041 0.9070 0.6696
-2.250 -0.2383 0.00632 0.00195 -0.0037 0.8935 0.6864
-2.000 -0.2119 0.00630 0.00189 -0.0032 0.8806 0.7020
-1.750 -0.1855 0.00630 0.00185 -0.0027 0.8685 0.7164
-1.500 -0.1593 0.00629 0.00182 -0.0023 0.8565 0.7298
-1.250 -0.1330 0.00629 0.00179 -0.0018 0.8450 0.7423
-1.000 -0.1066 0.00631 0.00178 -0.0014 0.8343 0.7546
-0.750 -0.0800 0.00630 0.00177 -0.0010 0.8246 0.7653
-0.500 -0.0534 0.00629 0.00176 -0.0006 0.8145 0.7755
-0.250 -0.0267 0.00630 0.00175 -0.0003 0.8048 0.7858
0.000 0.0000 0.00631 0.00175 0.0000 0.7957 0.7957
0.250 0.0268 0.00630 0.00175 0.0003 0.7858 0.8048
0.500 0.0534 0.00629 0.00176 0.0006 0.7755 0.8145
0.750 0.0800 0.00630 0.00177 0.0010 0.7653 0.8247
1.000 0.1066 0.00631 0.00178 0.0014 0.7546 0.8343
1.250 0.1330 0.00629 0.00180 0.0018 0.7424 0.8450
1.500 0.1593 0.00629 0.00182 0.0023 0.7298 0.8566
1.750 0.1855 0.00630 0.00185 0.0027 0.7163 0.8685
2.000 0.2119 0.00630 0.00189 0.0032 0.7020 0.8806
2.250 0.2384 0.00632 0.00194 0.0037 0.6866 0.8935
2.500 0.2652 0.00636 0.00200 0.0041 0.6696 0.9070
2.750 0.2930 0.00641 0.00208 0.0042 0.6500 0.9204
3.000 0.3220 0.00651 0.00216 0.0041 0.6255 0.9328
3.250 0.3520 0.00664 0.00226 0.0038 0.5883 0.9443
3.500 0.3806 0.00694 0.00235 0.0036 0.5147 0.9557
3.750 0.4089 0.00799 0.00267 0.0028 0.3381 0.9633
4.000 0.4374 0.00901 0.00309 0.0019 0.1965 0.9716
4.250 0.4711 0.00987 0.00352 -0.0001 0.1005 0.9765
4.500 0.5029 0.01051 0.00396 -0.0013 0.0590 0.9834
4.750 0.5395 0.01106 0.00444 -0.0035 0.0426 0.9870
5.000 0.5745 0.01162 0.00503 -0.0053 0.0349 0.9917
5.250 0.6079 0.01250 0.00594 -0.0070 0.0285 0.9961
5.500 0.6449 0.01292 0.00642 -0.0092 0.0258 0.9995
5.750 0.6640 0.01340 0.00692 -0.0077 0.0237 1.0000
6.000 0.6793 0.01401 0.00755 -0.0055 0.0220 1.0000
6.250 0.6910 0.01532 0.00896 -0.0026 0.0202 1.0000
6.500 0.7104 0.01574 0.00944 -0.0010 0.0192 1.0000
6.750 0.7290 0.01641 0.01018 0.0007 0.0181 1.0000
7.000 0.7476 0.01719 0.01103 0.0023 0.0172 1.0000
7.250 0.7665 0.01802 0.01193 0.0038 0.0164 1.0000
7.500 0.7851 0.01899 0.01296 0.0054 0.0157 1.0000
7.750 0.8015 0.02084 0.01493 0.0071 0.0149 1.0000
8.000 0.8162 0.02358 0.01797 0.0091 0.0146 1.0000
8.250 0.8257 0.02723 0.02201 0.0118 0.0144 1.0000
8.500 0.8416 0.02817 0.02318 0.0136 0.0141 1.0000
8.750 0.8550 0.02978 0.02505 0.0157 0.0137 1.0000
9.000 0.8638 0.03224 0.02782 0.0182 0.0133 1.0000
9.250 0.8632 0.03602 0.03199 0.0214 0.0134 1.0000
9.500 0.8573 0.04001 0.03634 0.0248 0.0136 1.0000
9.750 0.8428 0.04449 0.04113 0.0284 0.0139 1.0000
10.000 0.8494 0.04675 0.04342 0.0300 0.0148 1.0000
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Polar data table (+)
Polar graphs
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