EPPLER STE 871-514 AIRFOIL (ste87151-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: EPPLER STE 871-514 AIRFOIL (ste87151-il) Reynolds number: 100,000 Max Cl/Cd: 40.6 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ste87151-il-100000.txt Download as CSV file: xf-ste87151-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER STE 871-514 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.6165 0.09824 0.09391 -0.0303 1.0020 0.1110
-8.000 -0.6509 0.09483 0.09048 -0.0293 1.0020 0.1115
-7.750 -0.6929 0.09119 0.08662 -0.0289 1.0020 0.1123
-7.500 -0.6617 0.08768 0.08336 -0.0261 1.0020 0.1166
-7.250 -0.6047 0.08309 0.07910 -0.0242 1.0020 0.1286
-7.000 -0.6250 0.07925 0.07519 -0.0229 1.0020 0.1321
-6.750 -0.6839 0.07695 0.07242 -0.0229 1.0020 0.1302
-6.500 -0.6439 0.07126 0.06709 -0.0202 1.0020 0.1414
-6.000 -0.6798 0.06667 0.06182 -0.0202 1.0020 0.1548
-5.750 -0.6740 0.06347 0.05849 -0.0196 1.0020 0.1681
-5.500 -0.6654 0.06058 0.05549 -0.0188 1.0020 0.1837
-5.250 -0.6547 0.05787 0.05274 -0.0178 1.0020 0.2015
-5.000 -0.6429 0.05534 0.05022 -0.0166 1.0020 0.2219
-4.250 -0.5273 0.03692 0.02807 -0.0186 1.0020 0.0616
-4.000 -0.5011 0.03486 0.02548 -0.0178 1.0020 0.0592
-3.750 -0.4775 0.03354 0.02402 -0.0172 1.0020 0.0613
-3.500 -0.4528 0.03226 0.02247 -0.0165 1.0020 0.0621
-3.250 -0.4281 0.03106 0.02111 -0.0156 1.0020 0.0621
-3.000 -0.4038 0.03011 0.02004 -0.0147 1.0020 0.0631
-2.750 -0.3798 0.02925 0.01918 -0.0139 1.0020 0.0650
-2.500 -0.3559 0.02864 0.01861 -0.0134 1.0020 0.0700
-2.250 -0.3314 0.02805 0.01814 -0.0133 1.0020 0.0779
-2.000 -0.3048 0.02748 0.01769 -0.0136 1.0020 0.0885
-1.750 -0.2753 0.02684 0.01736 -0.0145 1.0020 0.1279
-1.500 -0.1317 0.03227 0.02547 -0.0166 1.0020 0.9980
-1.250 -0.1231 0.03218 0.02526 -0.0144 1.0020 0.9980
-1.000 -0.1146 0.03213 0.02510 -0.0122 1.0020 0.9980
-0.750 -0.1061 0.03210 0.02499 -0.0099 1.0020 0.9980
-0.500 -0.0978 0.03211 0.02491 -0.0077 1.0020 0.9980
-0.250 -0.0897 0.03213 0.02486 -0.0055 1.0020 0.9980
0.000 -0.0817 0.03218 0.02483 -0.0032 1.0020 0.9980
0.250 -0.0738 0.03225 0.02484 -0.0010 1.0020 0.9980
0.500 -0.0661 0.03233 0.02487 0.0012 1.0020 0.9980
0.750 -0.0585 0.03243 0.02492 0.0034 1.0020 0.9980
1.000 -0.0510 0.03254 0.02499 0.0056 1.0020 0.9980
1.250 -0.0436 0.03266 0.02508 0.0078 1.0020 0.9980
1.500 -0.0197 0.03333 0.02570 0.0066 0.9979 0.9980
1.750 0.0326 0.03487 0.02718 0.0002 0.9834 0.9980
2.000 0.1483 0.03655 0.02874 -0.0166 0.9346 0.9980
2.250 0.1885 0.03694 0.02911 -0.0196 0.9171 0.9980
2.500 0.2249 0.03731 0.02948 -0.0219 0.9036 0.9980
2.750 0.2579 0.03754 0.02972 -0.0236 0.8911 0.9980
3.000 0.2820 0.03754 0.02974 -0.0235 0.8769 0.9980
3.250 0.3059 0.03755 0.02977 -0.0235 0.8633 0.9980
3.500 0.3336 0.03756 0.02982 -0.0239 0.8509 0.9980
3.750 0.3787 0.03760 0.02991 -0.0272 0.8425 0.9980
4.000 0.3997 0.03730 0.02968 -0.0263 0.8284 0.9980
4.250 0.4238 0.03696 0.02940 -0.0258 0.8150 0.9980
4.500 0.4519 0.03651 0.02903 -0.0259 0.8022 0.9980
4.750 0.4876 0.03583 0.02842 -0.0269 0.7906 0.9980
5.000 0.5389 0.03460 0.02734 -0.0298 0.7809 0.9980
5.250 0.5705 0.03337 0.02621 -0.0295 0.7663 0.9980
5.500 0.6054 0.03182 0.02478 -0.0294 0.7512 0.9980
5.750 0.6404 0.03011 0.02320 -0.0290 0.7357 0.9980
6.000 0.6741 0.02836 0.02162 -0.0283 0.7198 0.9980
6.250 0.7054 0.02661 0.02001 -0.0271 0.7029 0.9980
6.500 0.7388 0.02481 0.01835 -0.0262 0.6822 0.9980
6.750 0.7722 0.02324 0.01689 -0.0254 0.6500 0.9980
7.000 0.8238 0.02122 0.01474 -0.0268 0.5825 0.9980
7.250 0.8530 0.02101 0.01370 -0.0255 0.4408 0.9980
7.500 0.8493 0.02242 0.01431 -0.0207 0.3271 0.9980
7.750 0.8490 0.02413 0.01531 -0.0170 0.2489 0.9980
8.000 0.8597 0.02556 0.01637 -0.0150 0.2050 0.9980
8.250 0.8774 0.02684 0.01736 -0.0141 0.1805 0.9980
8.500 0.9012 0.02797 0.01839 -0.0140 0.1632 0.9980
8.750 0.9302 0.02912 0.01943 -0.0147 0.1503 0.9980
9.000 0.9654 0.03036 0.02054 -0.0164 0.1399 0.9980
9.250 1.0013 0.03160 0.02179 -0.0182 0.1315 0.9980
9.500 1.0401 0.03309 0.02339 -0.0203 0.1245 0.9980
9.750 1.0898 0.03504 0.02527 -0.0245 0.1183 0.9980
10.000 1.1192 0.03668 0.02723 -0.0253 0.1138 0.9980
10.250 1.1699 0.03922 0.02971 -0.0300 0.1090 0.9980
10.500 1.1857 0.04106 0.03205 -0.0285 0.1066 0.9980
10.750 1.2059 0.04314 0.03446 -0.0281 0.1037 0.9980
11.000 1.2412 0.04570 0.03710 -0.0303 0.1007 0.9980
11.250 1.2540 0.04849 0.04027 -0.0289 0.0989 0.9980
11.500 1.2531 0.05094 0.04322 -0.0254 0.0976 0.9980
11.750 1.2521 0.05368 0.04638 -0.0222 0.0964 0.9980
12.000 1.2491 0.05663 0.04971 -0.0190 0.0954 0.9980
12.250 1.2443 0.05955 0.05294 -0.0158 0.0944 0.9980
12.500 1.2776 0.06269 0.05601 -0.0181 0.0912 0.9980
12.750 1.2572 0.06585 0.05955 -0.0133 0.0911 0.9980
13.000 1.2366 0.06934 0.06336 -0.0091 0.0910 0.9980
13.250 1.2137 0.07306 0.06739 -0.0054 0.0911 0.9980
13.500 1.1898 0.07713 0.07172 -0.0023 0.0911 0.9980
13.750 1.1664 0.08157 0.07639 0.0001 0.0913 0.9980
14.000 1.1442 0.08637 0.08139 0.0016 0.0915 0.9980
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Polar data table (+)
Polar graphs
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