REPUBLIC S-3 AIRFOIL (s3-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: REPUBLIC S-3 AIRFOIL (s3-il) Reynolds number: 1,000,000 Max Cl/Cd: 91.28 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s3-il-1000000.txt Download as CSV file: xf-s3-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.8332 0.02948 0.02617 -0.0215 1.0000 0.0081
-9.250 -0.8397 0.02379 0.01984 -0.0182 1.0000 0.0083
-9.000 -0.8270 0.02129 0.01707 -0.0166 1.0000 0.0085
-8.750 -0.8069 0.02030 0.01599 -0.0156 1.0000 0.0088
-8.500 -0.7853 0.01956 0.01517 -0.0148 1.0000 0.0090
-8.250 -0.7644 0.01855 0.01404 -0.0137 1.0000 0.0092
-8.000 -0.7429 0.01763 0.01302 -0.0128 1.0000 0.0095
-7.750 -0.7213 0.01676 0.01203 -0.0117 1.0000 0.0098
-7.500 -0.6998 0.01589 0.01106 -0.0106 1.0000 0.0101
-7.250 -0.6783 0.01513 0.01021 -0.0094 1.0000 0.0104
-7.000 -0.6572 0.01441 0.00941 -0.0082 1.0000 0.0106
-6.750 -0.6221 0.01387 0.00879 -0.0098 0.9908 0.0109
-6.500 -0.5886 0.01228 0.00707 -0.0114 0.9778 0.0117
-6.250 -0.5502 0.01180 0.00657 -0.0138 0.9597 0.0125
-6.000 -0.5186 0.01147 0.00616 -0.0146 0.9309 0.0133
-5.750 -0.4943 0.01120 0.00577 -0.0136 0.8994 0.0141
-5.500 -0.4706 0.01091 0.00535 -0.0126 0.8706 0.0145
-5.250 -0.4485 0.01030 0.00458 -0.0114 0.8424 0.0156
-5.000 -0.4237 0.01003 0.00421 -0.0106 0.8190 0.0167
-4.750 -0.3980 0.00983 0.00392 -0.0101 0.7980 0.0180
-4.500 -0.3718 0.00969 0.00369 -0.0096 0.7800 0.0191
-4.250 -0.3464 0.00933 0.00325 -0.0090 0.7655 0.0219
-4.000 -0.3196 0.00918 0.00305 -0.0087 0.7529 0.0242
-3.750 -0.2933 0.00893 0.00274 -0.0083 0.7392 0.0280
-3.500 -0.2663 0.00881 0.00258 -0.0080 0.7271 0.0322
-3.250 -0.2396 0.00862 0.00236 -0.0077 0.7171 0.0388
-3.000 -0.2120 0.00849 0.00221 -0.0075 0.7076 0.0442
-2.750 -0.1849 0.00831 0.00204 -0.0072 0.6988 0.0533
-2.500 -0.1577 0.00815 0.00189 -0.0070 0.6888 0.0673
-2.250 -0.1314 0.00782 0.00171 -0.0067 0.6775 0.1159
-2.000 -0.1110 0.00662 0.00136 -0.0058 0.6664 0.3617
-1.750 -0.0866 0.00612 0.00125 -0.0053 0.6546 0.4834
-1.500 -0.0602 0.00592 0.00118 -0.0049 0.6405 0.5383
-1.250 -0.0334 0.00578 0.00113 -0.0046 0.6240 0.5830
-1.000 -0.0065 0.00570 0.00108 -0.0043 0.6025 0.6169
-0.750 0.0200 0.00567 0.00104 -0.0039 0.5729 0.6512
-0.500 0.0462 0.00571 0.00103 -0.0035 0.5323 0.6830
-0.250 0.0720 0.00584 0.00105 -0.0030 0.4808 0.7114
0.000 0.0978 0.00599 0.00110 -0.0025 0.4376 0.7398
0.250 0.1246 0.00609 0.00114 -0.0022 0.4144 0.7615
0.500 0.1520 0.00617 0.00118 -0.0020 0.3970 0.7759
0.750 0.1793 0.00623 0.00122 -0.0018 0.3817 0.7907
1.000 0.2064 0.00629 0.00126 -0.0015 0.3655 0.8069
1.250 0.2334 0.00637 0.00131 -0.0013 0.3472 0.8220
1.500 0.2605 0.00645 0.00135 -0.0010 0.3270 0.8345
1.750 0.2875 0.00656 0.00140 -0.0008 0.3066 0.8461
2.000 0.3143 0.00667 0.00147 -0.0005 0.2897 0.8583
2.250 0.3411 0.00676 0.00154 -0.0002 0.2766 0.8714
2.500 0.3679 0.00685 0.00162 0.0001 0.2654 0.8852
2.750 0.3944 0.00695 0.00171 0.0005 0.2534 0.8999
3.000 0.4210 0.00705 0.00180 0.0008 0.2417 0.9162
3.250 0.4483 0.00715 0.00190 0.0010 0.2307 0.9339
3.500 0.4784 0.00726 0.00200 0.0006 0.2196 0.9514
3.750 0.5125 0.00743 0.00211 -0.0008 0.2063 0.9651
4.000 0.5484 0.00761 0.00224 -0.0026 0.1949 0.9747
4.250 0.5830 0.00780 0.00238 -0.0041 0.1864 0.9832
4.500 0.6196 0.00794 0.00251 -0.0061 0.1816 0.9879
4.750 0.6537 0.00813 0.00266 -0.0075 0.1765 0.9938
5.250 0.7199 0.00845 0.00294 -0.0101 0.1651 1.0000
5.500 0.7434 0.00866 0.00312 -0.0093 0.1594 1.0000
5.750 0.7682 0.00877 0.00324 -0.0087 0.1562 1.0000
6.000 0.7930 0.00892 0.00339 -0.0081 0.1519 1.0000
6.250 0.8173 0.00915 0.00358 -0.0075 0.1448 1.0000
6.500 0.8426 0.00927 0.00371 -0.0071 0.1388 1.0000
6.750 0.8672 0.00950 0.00390 -0.0066 0.1298 1.0000
7.000 0.8908 0.00987 0.00413 -0.0059 0.1065 1.0000
7.250 0.9103 0.01073 0.00472 -0.0049 0.0659 1.0000
7.500 0.9329 0.01125 0.00516 -0.0041 0.0518 1.0000
7.750 0.9523 0.01220 0.00587 -0.0031 0.0186 1.0000
8.000 0.9757 0.01265 0.00635 -0.0024 0.0144 1.0000
8.250 0.9995 0.01306 0.00682 -0.0018 0.0131 1.0000
8.500 1.0225 0.01356 0.00740 -0.0012 0.0120 1.0000
8.750 1.0445 0.01418 0.00810 -0.0004 0.0109 1.0000
9.000 1.0677 0.01462 0.00859 0.0002 0.0105 1.0000
9.250 1.0903 0.01512 0.00915 0.0009 0.0099 1.0000
9.500 1.1121 0.01570 0.00979 0.0016 0.0094 1.0000
9.750 1.1328 0.01636 0.01052 0.0025 0.0089 1.0000
10.000 1.1513 0.01720 0.01144 0.0036 0.0084 1.0000
10.250 1.1643 0.01852 0.01290 0.0053 0.0080 1.0000
10.500 1.1808 0.01939 0.01386 0.0066 0.0078 1.0000
10.750 1.1982 0.02012 0.01466 0.0078 0.0076 1.0000
11.000 1.2145 0.02089 0.01550 0.0090 0.0074 1.0000
11.250 1.2283 0.02175 0.01644 0.0106 0.0071 1.0000
11.500 1.2395 0.02269 0.01746 0.0124 0.0069 1.0000
11.750 1.2495 0.02342 0.01824 0.0145 0.0066 1.0000
12.000 1.2535 0.02443 0.01933 0.0172 0.0064 1.0000
12.250 1.2530 0.02586 0.02085 0.0197 0.0063 1.0000
12.500 1.2539 0.02742 0.02250 0.0214 0.0062 1.0000
12.750 1.2536 0.02934 0.02450 0.0225 0.0060 1.0000
13.000 1.2482 0.03202 0.02728 0.0229 0.0060 1.0000
13.250 1.2454 0.03483 0.03019 0.0227 0.0059 1.0000
13.500 1.2283 0.03951 0.03500 0.0214 0.0057 1.0000
13.750 1.2201 0.04360 0.03920 0.0199 0.0057 1.0000
14.000 1.2084 0.04853 0.04426 0.0177 0.0057 1.0000
14.250 1.1927 0.05428 0.05013 0.0149 0.0056 1.0000
14.500 1.1809 0.05966 0.05562 0.0123 0.0056 1.0000
14.750 1.1681 0.06511 0.06116 0.0098 0.0056 1.0000
15.000 1.1570 0.07030 0.06644 0.0074 0.0056 1.0000
15.250 1.1421 0.07583 0.07205 0.0052 0.0054 1.0000
15.500 1.1345 0.08074 0.07704 0.0029 0.0054 1.0000
15.750 1.1282 0.08518 0.08154 0.0010 0.0054 1.0000
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Polar data table (+)
Polar graphs
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