E10(08%) (e228-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: E10(08%) (e228-il) Reynolds number: 100,000 Max Cl/Cd: 38.67 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e228-il-100000-n5.txt Download as CSV file: xf-e228-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: E10(08%)
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.7545 0.07959 0.07479 -0.0102 1.0000 0.0205
-11.000 -0.7940 0.06788 0.06285 -0.0200 1.0000 0.0196
-10.750 -0.8193 0.06146 0.05621 -0.0241 1.0000 0.0194
-10.500 -0.8391 0.05695 0.05152 -0.0251 1.0000 0.0193
-10.250 -0.8553 0.05333 0.04763 -0.0240 1.0000 0.0193
-10.000 -0.8661 0.05019 0.04424 -0.0221 1.0000 0.0195
-9.750 -0.8700 0.04704 0.04081 -0.0205 1.0000 0.0198
-9.500 -0.8690 0.04409 0.03756 -0.0189 1.0000 0.0203
-9.250 -0.8626 0.04168 0.03488 -0.0175 1.0000 0.0214
-9.000 -0.8532 0.03939 0.03228 -0.0161 1.0000 0.0229
-8.750 -0.8419 0.03691 0.02939 -0.0146 1.0000 0.0244
-8.500 -0.8276 0.03441 0.02647 -0.0132 1.0000 0.0255
-8.250 -0.8104 0.03213 0.02385 -0.0119 1.0000 0.0263
-8.000 -0.7940 0.02952 0.02108 -0.0107 1.0000 0.0278
-7.750 -0.7755 0.02817 0.01965 -0.0098 1.0000 0.0304
-7.500 -0.7554 0.02682 0.01812 -0.0088 1.0000 0.0335
-7.250 -0.7350 0.02532 0.01638 -0.0076 1.0000 0.0354
-7.000 -0.7160 0.02386 0.01480 -0.0062 1.0000 0.0374
-6.750 -0.6979 0.02257 0.01347 -0.0049 1.0000 0.0405
-6.500 -0.6780 0.02159 0.01238 -0.0037 1.0000 0.0455
-6.250 -0.6590 0.02053 0.01130 -0.0024 1.0000 0.0532
-6.000 -0.6400 0.01949 0.01024 -0.0010 1.0000 0.0637
-5.750 -0.6211 0.01848 0.00929 0.0004 1.0000 0.0845
-5.500 -0.6019 0.01762 0.00852 0.0016 1.0000 0.1152
-5.250 -0.5830 0.01685 0.00796 0.0029 1.0000 0.1577
-5.000 -0.5653 0.01610 0.00752 0.0043 1.0000 0.2133
-4.750 -0.5474 0.01551 0.00713 0.0059 1.0000 0.2640
-4.500 -0.5293 0.01502 0.00677 0.0075 1.0000 0.3039
-4.250 -0.5113 0.01459 0.00646 0.0091 1.0000 0.3433
-4.000 -0.4937 0.01421 0.00619 0.0109 1.0000 0.3839
-3.750 -0.4761 0.01387 0.00593 0.0127 1.0000 0.4236
-3.500 -0.4587 0.01357 0.00574 0.0145 1.0000 0.4640
-3.250 -0.4280 0.01329 0.00559 0.0137 0.9929 0.5117
-3.000 -0.3917 0.01305 0.00547 0.0119 0.9836 0.5616
-2.750 -0.3559 0.01284 0.00535 0.0104 0.9742 0.6090
-2.500 -0.3183 0.01269 0.00530 0.0086 0.9657 0.6555
-2.250 -0.2806 0.01257 0.00526 0.0069 0.9569 0.6993
-2.000 -0.2447 0.01249 0.00525 0.0057 0.9466 0.7390
-1.750 -0.2076 0.01245 0.00523 0.0044 0.9365 0.7755
-1.500 -0.1688 0.01245 0.00524 0.0028 0.9267 0.8077
-1.250 -0.1311 0.01247 0.00525 0.0014 0.9152 0.8355
-1.000 -0.0936 0.01250 0.00526 0.0000 0.9027 0.8596
-0.750 -0.0555 0.01254 0.00527 -0.0015 0.8895 0.8803
-0.500 -0.0192 0.01258 0.00526 -0.0027 0.8752 0.8996
-0.250 0.0189 0.01262 0.00526 -0.0042 0.8603 0.9153
0.000 0.0563 0.01265 0.00524 -0.0056 0.8447 0.9299
0.250 0.0939 0.01268 0.00522 -0.0072 0.8278 0.9430
0.500 0.1309 0.01269 0.00520 -0.0087 0.8094 0.9557
0.750 0.1692 0.01268 0.00514 -0.0106 0.7903 0.9671
1.000 0.2077 0.01265 0.00508 -0.0125 0.7701 0.9779
1.250 0.2461 0.01260 0.00499 -0.0146 0.7480 0.9884
1.500 0.2852 0.01254 0.00490 -0.0169 0.7247 0.9985
1.750 0.3088 0.01254 0.00485 -0.0160 0.7024 1.0000
2.000 0.3299 0.01257 0.00484 -0.0147 0.6787 1.0000
2.250 0.3511 0.01263 0.00485 -0.0134 0.6551 1.0000
2.500 0.3725 0.01272 0.00487 -0.0120 0.6310 1.0000
2.750 0.3942 0.01283 0.00496 -0.0108 0.6055 1.0000
3.000 0.4161 0.01297 0.00505 -0.0095 0.5787 1.0000
3.250 0.4380 0.01314 0.00517 -0.0083 0.5509 1.0000
3.500 0.4600 0.01334 0.00530 -0.0071 0.5219 1.0000
3.750 0.4819 0.01357 0.00547 -0.0059 0.4915 1.0000
4.000 0.5039 0.01383 0.00571 -0.0048 0.4595 1.0000
4.250 0.5258 0.01413 0.00594 -0.0036 0.4264 1.0000
4.750 0.5690 0.01487 0.00652 -0.0014 0.3570 1.0000
5.000 0.5903 0.01531 0.00688 -0.0003 0.3211 1.0000
5.250 0.6113 0.01581 0.00728 0.0008 0.2832 1.0000
5.500 0.6321 0.01636 0.00779 0.0018 0.2459 1.0000
5.750 0.6523 0.01700 0.00831 0.0029 0.2094 1.0000
6.000 0.6724 0.01769 0.00890 0.0040 0.1748 1.0000
6.250 0.6919 0.01845 0.00956 0.0052 0.1450 1.0000
6.500 0.7112 0.01926 0.01031 0.0064 0.1192 1.0000
6.750 0.7300 0.02015 0.01115 0.0076 0.0990 1.0000
7.000 0.7485 0.02108 0.01208 0.0089 0.0814 1.0000
7.250 0.7663 0.02211 0.01318 0.0103 0.0685 1.0000
7.500 0.7837 0.02318 0.01428 0.0117 0.0565 1.0000
7.750 0.8001 0.02437 0.01551 0.0131 0.0474 1.0000
8.000 0.8156 0.02562 0.01675 0.0145 0.0393 1.0000
8.250 0.8322 0.02684 0.01816 0.0160 0.0329 1.0000
8.500 0.8452 0.02840 0.01974 0.0176 0.0284 1.0000
8.750 0.8603 0.02997 0.02155 0.0193 0.0256 1.0000
9.000 0.8740 0.03170 0.02346 0.0209 0.0234 1.0000
9.250 0.8865 0.03349 0.02544 0.0224 0.0220 1.0000
9.500 0.8969 0.03558 0.02765 0.0240 0.0209 1.0000
9.750 0.9049 0.03819 0.03048 0.0257 0.0199 1.0000
10.000 0.9121 0.04051 0.03314 0.0274 0.0190 1.0000
10.250 0.9153 0.04276 0.03571 0.0292 0.0177 1.0000
10.500 0.9135 0.04514 0.03835 0.0313 0.0171 1.0000
10.750 0.9074 0.04762 0.04105 0.0332 0.0165 1.0000
11.000 0.8990 0.05063 0.04429 0.0343 0.0163 1.0000
11.250 0.8877 0.05412 0.04799 0.0343 0.0160 1.0000
11.500 0.8741 0.05832 0.05242 0.0332 0.0159 1.0000
11.750 0.8564 0.06365 0.05797 0.0307 0.0161 1.0000
12.000 0.8367 0.07003 0.06456 0.0267 0.0163 1.0000
12.250 0.8129 0.07810 0.07281 0.0210 0.0164 1.0000
12.500 0.7772 0.09026 0.08515 0.0125 0.0176 1.0000
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Polar data table (+)
Polar graphs
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