AG10 (ag10-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: AG10 (ag10-il) Reynolds number: 500,000 Max Cl/Cd: 67.47 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ag10-il-500000-n5.txt Download as CSV file: xf-ag10-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: AG10
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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.6490 0.12693 0.12459 0.0414 1.0000 0.0093
-9.750 -0.6444 0.12327 0.12094 0.0400 1.0000 0.0094
-9.500 -0.6446 0.11734 0.11502 0.0369 1.0000 0.0086
-9.250 -0.6393 0.11415 0.11183 0.0361 1.0000 0.0085
-9.000 -0.6353 0.11049 0.10818 0.0346 1.0000 0.0084
-8.750 -0.6316 0.10667 0.10438 0.0329 1.0000 0.0083
-8.500 -0.6284 0.10265 0.10037 0.0309 1.0000 0.0083
-7.500 -0.6145 0.08357 0.08136 0.0163 1.0000 0.0087
-7.250 -0.6035 0.07929 0.07707 0.0118 1.0000 0.0085
-7.000 -0.5898 0.07441 0.07217 0.0065 1.0000 0.0084
-6.000 -0.5061 0.01863 0.01387 -0.0308 1.0000 0.0085
-5.750 -0.4793 0.01652 0.01128 -0.0310 1.0000 0.0087
-5.500 -0.4524 0.01502 0.00953 -0.0311 1.0000 0.0091
-5.250 -0.4252 0.01439 0.00881 -0.0311 1.0000 0.0096
-5.000 -0.3977 0.01374 0.00804 -0.0311 1.0000 0.0102
-4.750 -0.3701 0.01288 0.00700 -0.0310 1.0000 0.0109
-4.500 -0.3424 0.01206 0.00600 -0.0309 1.0000 0.0113
-4.250 -0.3147 0.01128 0.00513 -0.0308 1.0000 0.0120
-4.000 -0.2870 0.01079 0.00460 -0.0307 1.0000 0.0128
-3.750 -0.2592 0.01035 0.00410 -0.0306 1.0000 0.0139
-3.500 -0.2313 0.00994 0.00362 -0.0305 1.0000 0.0149
-3.250 -0.2036 0.00956 0.00324 -0.0304 1.0000 0.0167
-3.000 -0.1757 0.00928 0.00294 -0.0303 1.0000 0.0188
-2.750 -0.1479 0.00892 0.00258 -0.0302 1.0000 0.0211
-2.500 -0.1201 0.00869 0.00234 -0.0301 1.0000 0.0241
-2.250 -0.0924 0.00841 0.00209 -0.0299 1.0000 0.0284
-2.000 -0.0646 0.00819 0.00188 -0.0298 1.0000 0.0333
-1.750 -0.0370 0.00799 0.00172 -0.0296 1.0000 0.0404
-1.500 -0.0053 0.00780 0.00158 -0.0304 0.9864 0.0513
-1.250 0.0304 0.00762 0.00146 -0.0320 0.9531 0.0679
-1.000 0.0594 0.00754 0.00138 -0.0318 0.9041 0.0865
-0.750 0.0830 0.00758 0.00130 -0.0304 0.8444 0.1069
-0.500 0.1077 0.00766 0.00123 -0.0294 0.7795 0.1348
-0.250 0.1338 0.00774 0.00118 -0.0289 0.7114 0.1770
0.000 0.1606 0.00778 0.00115 -0.0287 0.6426 0.2429
0.250 0.1878 0.00772 0.00115 -0.0287 0.5751 0.3525
0.500 0.2085 0.00609 0.00112 -0.0271 0.5219 1.0000
0.750 0.2361 0.00633 0.00112 -0.0270 0.4720 1.0000
1.000 0.2639 0.00656 0.00114 -0.0269 0.4309 1.0000
1.250 0.2918 0.00677 0.00118 -0.0268 0.3948 1.0000
1.500 0.3197 0.00697 0.00123 -0.0268 0.3643 1.0000
1.750 0.3476 0.00715 0.00129 -0.0267 0.3373 1.0000
2.000 0.3755 0.00734 0.00136 -0.0267 0.3118 1.0000
2.250 0.4033 0.00754 0.00144 -0.0266 0.2872 1.0000
2.500 0.4312 0.00773 0.00155 -0.0266 0.2642 1.0000
2.750 0.4590 0.00794 0.00166 -0.0265 0.2410 1.0000
3.000 0.4868 0.00815 0.00178 -0.0264 0.2181 1.0000
3.250 0.5145 0.00839 0.00193 -0.0264 0.1961 1.0000
3.500 0.5421 0.00863 0.00209 -0.0263 0.1741 1.0000
3.750 0.5697 0.00889 0.00227 -0.0263 0.1534 1.0000
4.000 0.5972 0.00916 0.00247 -0.0262 0.1326 1.0000
4.250 0.6247 0.00947 0.00270 -0.0261 0.1131 1.0000
4.500 0.6520 0.00977 0.00294 -0.0260 0.0946 1.0000
4.750 0.6792 0.01013 0.00321 -0.0260 0.0764 1.0000
5.000 0.7064 0.01048 0.00351 -0.0259 0.0605 1.0000
5.250 0.7334 0.01087 0.00385 -0.0258 0.0462 1.0000
5.500 0.7603 0.01129 0.00422 -0.0257 0.0345 1.0000
5.750 0.7872 0.01172 0.00462 -0.0255 0.0265 1.0000
6.000 0.8139 0.01217 0.00507 -0.0254 0.0207 1.0000
6.250 0.8406 0.01262 0.00554 -0.0252 0.0166 1.0000
6.500 0.8671 0.01310 0.00607 -0.0250 0.0140 1.0000
6.750 0.8934 0.01364 0.00665 -0.0248 0.0123 1.0000
7.000 0.9196 0.01421 0.00732 -0.0246 0.0111 1.0000
7.250 0.9456 0.01480 0.00799 -0.0243 0.0102 1.0000
7.500 0.9712 0.01545 0.00872 -0.0241 0.0094 1.0000
7.750 0.9960 0.01632 0.00969 -0.0238 0.0088 1.0000
8.000 1.0213 0.01700 0.01048 -0.0234 0.0083 1.0000
8.250 1.0462 0.01772 0.01131 -0.0231 0.0077 1.0000
8.500 1.0706 0.01853 0.01223 -0.0228 0.0073 1.0000
8.750 1.0944 0.01943 0.01325 -0.0224 0.0070 1.0000
9.000 1.1174 0.02049 0.01443 -0.0220 0.0067 1.0000
9.250 1.1383 0.02202 0.01611 -0.0214 0.0065 1.0000
9.500 1.1602 0.02321 0.01750 -0.0209 0.0063 1.0000
9.750 1.1810 0.02459 0.01908 -0.0203 0.0062 1.0000
10.000 1.2006 0.02615 0.02088 -0.0196 0.0060 1.0000
10.250 1.2187 0.02792 0.02291 -0.0189 0.0059 1.0000
10.500 1.2349 0.02993 0.02519 -0.0182 0.0057 1.0000
10.750 1.2488 0.03221 0.02775 -0.0174 0.0056 1.0000
11.000 1.2596 0.03480 0.03064 -0.0166 0.0055 1.0000
11.250 1.2665 0.03775 0.03390 -0.0158 0.0054 1.0000
11.500 1.2680 0.04114 0.03760 -0.0151 0.0053 1.0000
11.750 1.2622 0.04504 0.04180 -0.0148 0.0053 1.0000
12.000 1.2445 0.04976 0.04677 -0.0156 0.0053 1.0000
12.250 1.2232 0.05984 0.05711 -0.0263 0.0053 1.0000
12.500 1.1929 0.07427 0.07176 -0.0387 0.0054 1.0000
12.750 1.1577 0.08724 0.08485 -0.0468 0.0055 1.0000
13.000 1.1184 0.10060 0.09830 -0.0541 0.0057 1.0000
13.250 1.0723 0.11593 0.11370 -0.0619 0.0058 1.0000
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Polar data table (+)
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