REPUBLIC S-3 AIRFOIL (s3-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: REPUBLIC S-3 AIRFOIL (s3-il) Reynolds number: 500,000 Max Cl/Cd: 73.74 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-s3-il-500000-n5.txt Download as CSV file: xf-s3-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: REPUBLIC S-3 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.6756 0.05989 0.05765 -0.0298 1.0000 0.0056
-10.000 -0.6977 0.05519 0.05284 -0.0316 1.0000 0.0055
-9.750 -0.7910 0.03809 0.03497 -0.0281 1.0000 0.0054
-9.500 -0.8061 0.03122 0.02743 -0.0253 1.0000 0.0055
-9.250 -0.7985 0.02802 0.02382 -0.0236 1.0000 0.0056
-9.000 -0.7892 0.02505 0.02047 -0.0218 1.0000 0.0058
-8.750 -0.7727 0.02336 0.01858 -0.0206 1.0000 0.0059
-8.500 -0.7533 0.02220 0.01726 -0.0196 1.0000 0.0061
-8.250 -0.7334 0.02107 0.01598 -0.0186 1.0000 0.0063
-8.000 -0.7129 0.02001 0.01478 -0.0176 1.0000 0.0064
-7.750 -0.6909 0.01933 0.01402 -0.0168 1.0000 0.0068
-7.500 -0.6695 0.01851 0.01309 -0.0158 1.0000 0.0072
-7.250 -0.6385 0.01742 0.01184 -0.0168 0.9854 0.0076
-7.000 -0.6041 0.01639 0.01064 -0.0185 0.9669 0.0080
-6.500 -0.5410 0.01436 0.00831 -0.0205 0.9210 0.0088
-6.250 -0.5159 0.01382 0.00767 -0.0199 0.8980 0.0092
-5.750 -0.4679 0.01298 0.00659 -0.0182 0.8537 0.0105
-5.500 -0.4438 0.01263 0.00611 -0.0174 0.8305 0.0114
-5.250 -0.4196 0.01226 0.00560 -0.0166 0.8077 0.0121
-5.000 -0.3959 0.01179 0.00499 -0.0157 0.7869 0.0136
-4.500 -0.3448 0.01124 0.00424 -0.0146 0.7575 0.0166
-4.250 -0.3189 0.01096 0.00389 -0.0141 0.7455 0.0198
-4.000 -0.2926 0.01074 0.00362 -0.0137 0.7339 0.0241
-3.750 -0.2662 0.01053 0.00336 -0.0134 0.7228 0.0291
-3.500 -0.2394 0.01039 0.00316 -0.0131 0.7126 0.0345
-3.250 -0.2126 0.01018 0.00295 -0.0128 0.7022 0.0417
-3.000 -0.1857 0.01004 0.00274 -0.0125 0.6905 0.0480
-2.750 -0.1589 0.00988 0.00256 -0.0122 0.6772 0.0574
-2.500 -0.1321 0.00970 0.00237 -0.0120 0.6628 0.0727
-2.250 -0.1057 0.00944 0.00218 -0.0117 0.6489 0.1066
-2.000 -0.0833 0.00852 0.00186 -0.0111 0.6356 0.2878
-1.750 -0.0594 0.00795 0.00169 -0.0106 0.6184 0.4150
-1.500 -0.0340 0.00771 0.00158 -0.0101 0.5941 0.4843
-1.250 -0.0088 0.00756 0.00151 -0.0095 0.5614 0.5419
-1.000 0.0163 0.00757 0.00145 -0.0090 0.5135 0.5835
-0.750 0.0414 0.00765 0.00144 -0.0084 0.4626 0.6216
-0.500 0.0666 0.00770 0.00147 -0.0078 0.4287 0.6629
-0.250 0.0921 0.00773 0.00152 -0.0072 0.4089 0.7012
0.000 0.1190 0.00779 0.00154 -0.0069 0.3938 0.7222
0.250 0.1458 0.00783 0.00158 -0.0066 0.3795 0.7434
0.500 0.1725 0.00788 0.00162 -0.0063 0.3646 0.7638
0.750 0.1995 0.00795 0.00165 -0.0060 0.3481 0.7788
1.000 0.2267 0.00804 0.00168 -0.0058 0.3300 0.7906
1.250 0.2537 0.00814 0.00173 -0.0056 0.3114 0.8028
1.500 0.2805 0.00825 0.00178 -0.0053 0.2950 0.8154
1.750 0.3073 0.00835 0.00185 -0.0051 0.2821 0.8275
2.000 0.3340 0.00846 0.00193 -0.0048 0.2713 0.8393
2.250 0.3608 0.00856 0.00202 -0.0045 0.2616 0.8517
2.500 0.3876 0.00865 0.00211 -0.0042 0.2533 0.8651
2.750 0.4141 0.00876 0.00222 -0.0039 0.2446 0.8800
3.000 0.4411 0.00884 0.00233 -0.0036 0.2358 0.8961
3.250 0.4686 0.00897 0.00245 -0.0035 0.2243 0.9130
3.500 0.4980 0.00913 0.00257 -0.0038 0.2117 0.9302
3.750 0.5303 0.00930 0.00270 -0.0048 0.2006 0.9460
4.000 0.5653 0.00946 0.00286 -0.0064 0.1936 0.9600
4.250 0.6005 0.00965 0.00302 -0.0081 0.1869 0.9727
4.500 0.6351 0.00984 0.00318 -0.0097 0.1812 0.9844
4.750 0.6698 0.01004 0.00337 -0.0114 0.1745 0.9948
5.000 0.6999 0.01026 0.00357 -0.0121 0.1694 1.0000
5.250 0.7246 0.01043 0.00375 -0.0115 0.1668 1.0000
5.500 0.7493 0.01061 0.00395 -0.0110 0.1630 1.0000
5.750 0.7736 0.01087 0.00417 -0.0104 0.1562 1.0000
6.000 0.7988 0.01105 0.00436 -0.0100 0.1499 1.0000
6.250 0.8235 0.01131 0.00458 -0.0096 0.1417 1.0000
6.500 0.8487 0.01152 0.00481 -0.0092 0.1333 1.0000
6.750 0.8731 0.01184 0.00505 -0.0087 0.1164 1.0000
7.000 0.8925 0.01277 0.00565 -0.0078 0.0688 1.0000
7.250 0.9149 0.01335 0.00616 -0.0072 0.0545 1.0000
7.500 0.9374 0.01392 0.00666 -0.0065 0.0407 1.0000
7.750 0.9575 0.01479 0.00742 -0.0056 0.0146 1.0000
8.000 0.9802 0.01533 0.00803 -0.0049 0.0122 1.0000
8.250 1.0027 0.01589 0.00868 -0.0043 0.0104 1.0000
8.500 1.0257 0.01636 0.00922 -0.0037 0.0096 1.0000
8.750 1.0478 0.01690 0.00984 -0.0030 0.0086 1.0000
9.000 1.0690 0.01753 0.01056 -0.0022 0.0080 1.0000
9.250 1.0885 0.01833 0.01149 -0.0013 0.0073 1.0000
9.500 1.1092 0.01895 0.01220 -0.0005 0.0070 1.0000
9.750 1.1286 0.01966 0.01301 0.0004 0.0066 1.0000
10.000 1.1472 0.02041 0.01386 0.0014 0.0062 1.0000
10.250 1.1651 0.02115 0.01467 0.0024 0.0058 1.0000
10.500 1.1807 0.02205 0.01564 0.0037 0.0055 1.0000
10.750 1.1911 0.02327 0.01697 0.0054 0.0052 1.0000
11.000 1.2008 0.02438 0.01818 0.0073 0.0050 1.0000
11.250 1.2071 0.02541 0.01933 0.0098 0.0049 1.0000
11.500 1.2107 0.02661 0.02065 0.0121 0.0048 1.0000
11.750 1.2126 0.02808 0.02223 0.0141 0.0046 1.0000
12.000 1.2106 0.03003 0.02430 0.0157 0.0045 1.0000
12.250 1.2117 0.03198 0.02637 0.0165 0.0044 1.0000
12.500 1.2088 0.03458 0.02909 0.0167 0.0043 1.0000
12.750 1.2019 0.03790 0.03254 0.0163 0.0043 1.0000
13.000 1.1965 0.04136 0.03613 0.0154 0.0041 1.0000
13.250 1.1886 0.04548 0.04038 0.0139 0.0041 1.0000
13.500 1.1821 0.04977 0.04479 0.0119 0.0040 1.0000
13.750 1.1664 0.05579 0.05097 0.0089 0.0041 1.0000
14.000 1.1544 0.06154 0.05685 0.0059 0.0040 1.0000
14.250 1.1405 0.06758 0.06301 0.0029 0.0040 1.0000
14.500 1.1255 0.07379 0.06933 -0.0002 0.0040 1.0000
14.750 1.1121 0.07977 0.07542 -0.0031 0.0039 1.0000
15.000 1.0978 0.08597 0.08172 -0.0061 0.0040 1.0000
|
Polar data table (+)
Polar graphs
<< Back to REPUBLIC S-3 AIRFOIL (s3-il)