RG 15A-1.8/11.0 AIRFOIL (rg15a111-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: RG 15A-1.8/11.0 AIRFOIL (rg15a111-il) Reynolds number: 50,000 Max Cl/Cd: 35.08 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rg15a111-il-50000-n5.txt Download as CSV file: xf-rg15a111-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RG 15A-1.8/11.0 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4980 0.10791 0.10048 -0.0331 1.0000 0.0500
-10.750 -0.4993 0.10318 0.09580 -0.0347 1.0000 0.0497
-10.500 -0.5022 0.09825 0.09094 -0.0366 1.0000 0.0495
-10.250 -0.5071 0.09294 0.08569 -0.0389 1.0000 0.0493
-10.000 -0.5148 0.08697 0.07979 -0.0420 1.0000 0.0490
-9.750 -0.5258 0.08064 0.07352 -0.0457 1.0000 0.0486
-9.500 -0.5419 0.07484 0.06775 -0.0490 1.0000 0.0482
-9.250 -0.5627 0.07004 0.06298 -0.0508 1.0000 0.0479
-9.000 -0.5833 0.06598 0.05889 -0.0506 1.0000 0.0476
-8.750 -0.6002 0.06182 0.05461 -0.0502 1.0000 0.0474
-8.500 -0.6128 0.05776 0.05036 -0.0492 1.0000 0.0475
-8.250 -0.6216 0.05375 0.04604 -0.0479 1.0000 0.0480
-8.000 -0.6261 0.04985 0.04171 -0.0463 1.0000 0.0491
-7.750 -0.6252 0.04629 0.03770 -0.0446 1.0000 0.0504
-7.500 -0.6152 0.04406 0.03541 -0.0433 1.0000 0.0525
-7.250 -0.6047 0.04160 0.03271 -0.0417 1.0000 0.0544
-7.000 -0.5925 0.03890 0.02962 -0.0402 1.0000 0.0561
-6.750 -0.5780 0.03635 0.02667 -0.0386 1.0000 0.0582
-6.500 -0.5613 0.03413 0.02391 -0.0370 1.0000 0.0618
-6.250 -0.5448 0.03258 0.02240 -0.0357 1.0000 0.0663
-6.000 -0.5263 0.03096 0.02055 -0.0343 1.0000 0.0715
-5.750 -0.5068 0.02930 0.01868 -0.0327 1.0000 0.0764
-5.500 -0.4884 0.02808 0.01739 -0.0312 1.0000 0.0846
-5.250 -0.4701 0.02694 0.01623 -0.0298 1.0000 0.0961
-5.000 -0.4516 0.02586 0.01511 -0.0284 1.0000 0.1114
-4.750 -0.4334 0.02491 0.01420 -0.0271 1.0000 0.1341
-4.500 -0.4151 0.02401 0.01339 -0.0259 1.0000 0.1627
-4.250 -0.3963 0.02311 0.01260 -0.0248 1.0000 0.1948
-4.000 -0.3770 0.02226 0.01190 -0.0239 1.0000 0.2367
-3.750 -0.3584 0.02147 0.01133 -0.0229 1.0000 0.2907
-3.500 -0.3404 0.02070 0.01092 -0.0217 1.0000 0.3621
-3.250 -0.3237 0.02006 0.01075 -0.0200 1.0000 0.4536
-3.000 -0.3099 0.01975 0.01094 -0.0169 1.0000 0.5536
-2.750 -0.2963 0.01976 0.01111 -0.0137 1.0000 0.6395
-2.500 -0.2725 0.01997 0.01133 -0.0124 0.9945 0.7096
-2.250 -0.2422 0.02023 0.01153 -0.0121 0.9847 0.7672
-2.000 -0.2124 0.02043 0.01166 -0.0116 0.9747 0.8144
-1.750 -0.1797 0.02059 0.01170 -0.0118 0.9651 0.8571
-1.500 -0.1366 0.02078 0.01174 -0.0140 0.9572 0.8974
-1.250 -0.0843 0.02091 0.01168 -0.0184 0.9493 0.9314
-1.000 -0.0181 0.02107 0.01162 -0.0259 0.9439 0.9597
-0.750 0.0456 0.02110 0.01145 -0.0333 0.9364 0.9829
-0.500 0.1067 0.02109 0.01127 -0.0405 0.9292 1.0000
-0.250 0.1392 0.02101 0.01109 -0.0423 0.9155 1.0000
0.000 0.1715 0.02094 0.01093 -0.0439 0.9018 1.0000
0.250 0.2042 0.02087 0.01079 -0.0454 0.8882 1.0000
0.500 0.2372 0.02080 0.01065 -0.0468 0.8747 1.0000
0.750 0.2686 0.02072 0.01053 -0.0478 0.8608 1.0000
1.000 0.2993 0.02066 0.01042 -0.0484 0.8466 1.0000
1.250 0.3295 0.02062 0.01034 -0.0489 0.8321 1.0000
1.500 0.3595 0.02058 0.01027 -0.0493 0.8173 1.0000
1.750 0.3895 0.02055 0.01022 -0.0495 0.8021 1.0000
2.000 0.4194 0.02052 0.01020 -0.0496 0.7865 1.0000
2.250 0.4491 0.02050 0.01017 -0.0497 0.7705 1.0000
2.500 0.4790 0.02048 0.01015 -0.0496 0.7540 1.0000
2.750 0.5076 0.02050 0.01018 -0.0493 0.7368 1.0000
3.000 0.5327 0.02060 0.01030 -0.0485 0.7175 1.0000
3.250 0.5598 0.02067 0.01037 -0.0480 0.6984 1.0000
3.500 0.5882 0.02072 0.01045 -0.0475 0.6796 1.0000
3.750 0.6125 0.02089 0.01064 -0.0466 0.6582 1.0000
4.000 0.6388 0.02102 0.01077 -0.0458 0.6372 1.0000
4.250 0.6634 0.02122 0.01098 -0.0448 0.6148 1.0000
4.500 0.6887 0.02141 0.01120 -0.0439 0.5923 1.0000
4.750 0.7120 0.02168 0.01150 -0.0428 0.5682 1.0000
5.000 0.7362 0.02195 0.01175 -0.0417 0.5443 1.0000
5.250 0.7591 0.02229 0.01209 -0.0406 0.5192 1.0000
5.500 0.7812 0.02267 0.01252 -0.0393 0.4934 1.0000
5.750 0.8031 0.02310 0.01294 -0.0381 0.4675 1.0000
6.000 0.8245 0.02356 0.01338 -0.0368 0.4416 1.0000
6.250 0.8450 0.02409 0.01390 -0.0354 0.4151 1.0000
6.500 0.8643 0.02468 0.01450 -0.0339 0.3879 1.0000
6.750 0.8831 0.02533 0.01518 -0.0325 0.3608 1.0000
7.000 0.9012 0.02604 0.01588 -0.0309 0.3342 1.0000
7.250 0.9186 0.02682 0.01663 -0.0294 0.3085 1.0000
7.500 0.9352 0.02766 0.01745 -0.0278 0.2836 1.0000
7.750 0.9514 0.02861 0.01842 -0.0263 0.2592 1.0000
8.000 0.9666 0.02962 0.01945 -0.0246 0.2367 1.0000
8.250 0.9815 0.03073 0.02058 -0.0230 0.2153 1.0000
8.500 0.9948 0.03191 0.02170 -0.0213 0.1960 1.0000
8.750 1.0082 0.03319 0.02306 -0.0197 0.1773 1.0000
9.000 1.0202 0.03453 0.02444 -0.0180 0.1606 1.0000
9.250 1.0318 0.03596 0.02590 -0.0163 0.1461 1.0000
9.500 1.0424 0.03746 0.02750 -0.0146 0.1331 1.0000
9.750 1.0526 0.03905 0.02916 -0.0128 0.1223 1.0000
10.000 1.0623 0.04069 0.03075 -0.0111 0.1133 1.0000
10.250 1.0710 0.04250 0.03281 -0.0093 0.1040 1.0000
10.500 1.0811 0.04448 0.03490 -0.0079 0.0969 1.0000
10.750 1.0901 0.04654 0.03713 -0.0064 0.0905 1.0000
11.000 1.0976 0.04874 0.03943 -0.0050 0.0851 1.0000
11.250 1.1000 0.05119 0.04218 -0.0035 0.0798 1.0000
11.500 1.1077 0.05342 0.04447 -0.0024 0.0759 1.0000
11.750 1.1084 0.05652 0.04785 -0.0013 0.0727 1.0000
12.000 1.1015 0.06000 0.05170 -0.0003 0.0699 1.0000
12.250 1.0945 0.06346 0.05539 0.0001 0.0672 1.0000
12.500 1.0906 0.06663 0.05868 0.0003 0.0647 1.0000
12.750 1.0928 0.06981 0.06185 0.0004 0.0624 1.0000
13.000 1.0741 0.07516 0.06754 -0.0008 0.0619 1.0000
13.250 1.0522 0.08136 0.07402 -0.0032 0.0615 1.0000
13.500 1.0281 0.08850 0.08140 -0.0067 0.0615 1.0000
13.750 1.0018 0.09675 0.08983 -0.0114 0.0618 1.0000
14.000 0.9731 0.10641 0.09964 -0.0173 0.0621 1.0000
14.250 0.9438 0.11736 0.11063 -0.0241 0.0625 1.0000
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Polar data table (+)
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