EPPLER 426 AIRFOIL (e426-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 426 AIRFOIL (e426-il) Reynolds number: 500,000 Max Cl/Cd: 74.88 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e426-il-500000.txt Download as CSV file: xf-e426-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 426 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-15.750 -1.0615 0.08206 0.07826 -0.0161 1.0000 0.0172
-15.500 -1.0779 0.07552 0.07161 -0.0199 1.0000 0.0174
-15.250 -1.0918 0.06957 0.06556 -0.0233 1.0000 0.0173
-15.000 -1.1062 0.06351 0.05938 -0.0270 1.0000 0.0174
-14.750 -1.1200 0.05764 0.05339 -0.0306 1.0000 0.0175
-14.500 -1.1315 0.05214 0.04776 -0.0341 1.0000 0.0176
-14.250 -1.1412 0.04713 0.04262 -0.0373 1.0000 0.0176
-14.000 -1.1489 0.04257 0.03792 -0.0402 1.0000 0.0178
-13.750 -1.1549 0.03863 0.03384 -0.0425 1.0000 0.0179
-13.500 -1.1579 0.03543 0.03050 -0.0440 1.0000 0.0181
-13.250 -1.1597 0.03289 0.02782 -0.0445 1.0000 0.0182
-13.000 -1.1707 0.03027 0.02510 -0.0433 1.0000 0.0185
-12.750 -1.1760 0.02849 0.02324 -0.0407 1.0000 0.0188
-12.500 -1.1698 0.02699 0.02167 -0.0392 1.0000 0.0191
-12.250 -1.1587 0.02574 0.02034 -0.0379 1.0000 0.0196
-12.000 -1.1447 0.02463 0.01914 -0.0368 1.0000 0.0202
-11.750 -1.1285 0.02363 0.01804 -0.0357 1.0000 0.0210
-11.500 -1.1101 0.02276 0.01706 -0.0348 1.0000 0.0217
-11.250 -1.0967 0.02131 0.01555 -0.0335 1.0000 0.0227
-11.000 -1.0776 0.02041 0.01460 -0.0327 1.0000 0.0239
-10.750 -1.0572 0.01960 0.01370 -0.0318 1.0000 0.0252
-10.500 -1.0383 0.01855 0.01258 -0.0308 1.0000 0.0269
-10.250 -1.0174 0.01772 0.01170 -0.0300 1.0000 0.0290
-10.000 -0.9967 0.01681 0.01073 -0.0292 1.0000 0.0320
-9.250 -0.9314 0.01437 0.00836 -0.0268 1.0000 0.0539
-9.000 -0.9067 0.01400 0.00806 -0.0262 1.0000 0.0667
-8.750 -0.8806 0.01387 0.00786 -0.0258 1.0000 0.0740
-8.500 -0.8562 0.01350 0.00753 -0.0252 1.0000 0.0799
-8.250 -0.8303 0.01336 0.00730 -0.0246 1.0000 0.0840
-8.000 -0.8061 0.01297 0.00693 -0.0240 1.0000 0.0884
-7.750 -0.7810 0.01273 0.00669 -0.0234 1.0000 0.0924
-7.500 -0.7554 0.01257 0.00646 -0.0228 1.0000 0.0957
-7.250 -0.7310 0.01221 0.00611 -0.0221 1.0000 0.0994
-7.000 -0.7059 0.01200 0.00592 -0.0215 1.0000 0.1036
-6.750 -0.6803 0.01185 0.00573 -0.0208 1.0000 0.1071
-6.500 -0.6553 0.01161 0.00545 -0.0202 1.0000 0.1099
-6.250 -0.6308 0.01130 0.00519 -0.0194 1.0000 0.1139
-6.000 -0.6056 0.01111 0.00500 -0.0187 1.0000 0.1173
-5.750 -0.5804 0.01096 0.00481 -0.0180 1.0000 0.1206
-5.500 -0.5560 0.01066 0.00453 -0.0172 1.0000 0.1247
-5.250 -0.5315 0.01042 0.00433 -0.0165 1.0000 0.1289
-5.000 -0.5068 0.01024 0.00414 -0.0157 1.0000 0.1329
-4.750 -0.4825 0.01004 0.00393 -0.0148 1.0000 0.1372
-4.500 -0.4590 0.00977 0.00373 -0.0139 1.0000 0.1435
-4.250 -0.4316 0.00967 0.00362 -0.0137 0.9991 0.1497
-4.000 -0.3905 0.00941 0.00343 -0.0165 0.9941 0.1583
-3.750 -0.3498 0.00928 0.00329 -0.0190 0.9870 0.1656
-3.500 -0.3076 0.00901 0.00308 -0.0218 0.9756 0.1738
-3.250 -0.2660 0.00883 0.00290 -0.0244 0.9550 0.1811
-3.000 -0.2222 0.00858 0.00267 -0.0274 0.9261 0.1893
-2.750 -0.1862 0.00849 0.00249 -0.0286 0.8888 0.1970
-2.500 -0.1594 0.00844 0.00235 -0.0280 0.8514 0.2045
-2.250 -0.1333 0.00840 0.00223 -0.0272 0.8186 0.2133
-2.000 -0.1071 0.00833 0.00211 -0.0266 0.7899 0.2242
-1.750 -0.0808 0.00825 0.00200 -0.0259 0.7630 0.2397
-1.500 -0.0544 0.00813 0.00190 -0.0254 0.7381 0.2658
-1.250 -0.0286 0.00789 0.00180 -0.0249 0.7137 0.3203
-1.000 -0.0028 0.00764 0.00175 -0.0244 0.6894 0.3962
-0.750 0.0235 0.00750 0.00172 -0.0239 0.6647 0.4537
-0.500 0.0499 0.00742 0.00171 -0.0234 0.6415 0.5021
-0.250 0.0766 0.00734 0.00171 -0.0229 0.6177 0.5442
0.000 0.1033 0.00732 0.00171 -0.0224 0.5936 0.5795
0.250 0.1300 0.00731 0.00173 -0.0219 0.5690 0.6148
0.500 0.1566 0.00732 0.00176 -0.0213 0.5431 0.6490
0.750 0.1832 0.00734 0.00180 -0.0208 0.5160 0.6796
1.000 0.2097 0.00741 0.00184 -0.0202 0.4851 0.7070
1.250 0.2358 0.00751 0.00189 -0.0196 0.4482 0.7345
1.500 0.2614 0.00764 0.00195 -0.0189 0.4042 0.7632
1.750 0.2861 0.00786 0.00204 -0.0181 0.3476 0.7943
2.000 0.3104 0.00808 0.00215 -0.0171 0.3001 0.8277
2.250 0.3345 0.00823 0.00229 -0.0160 0.2744 0.8659
2.500 0.3597 0.00832 0.00243 -0.0150 0.2584 0.9189
2.750 0.4004 0.00851 0.00257 -0.0173 0.2442 0.9710
3.000 0.4455 0.00868 0.00269 -0.0209 0.2329 0.9981
3.250 0.4723 0.00890 0.00284 -0.0206 0.2251 1.0000
3.500 0.4980 0.00904 0.00297 -0.0200 0.2190 1.0000
3.750 0.5236 0.00927 0.00312 -0.0194 0.2128 1.0000
4.000 0.5497 0.00945 0.00329 -0.0189 0.2075 1.0000
4.250 0.5762 0.00961 0.00343 -0.0184 0.2019 1.0000
4.500 0.6023 0.00985 0.00362 -0.0180 0.1963 1.0000
4.750 0.6287 0.01005 0.00382 -0.0175 0.1915 1.0000
5.000 0.6556 0.01020 0.00398 -0.0171 0.1865 1.0000
5.250 0.6819 0.01044 0.00417 -0.0167 0.1807 1.0000
5.500 0.7083 0.01068 0.00441 -0.0163 0.1755 1.0000
5.750 0.7354 0.01083 0.00458 -0.0160 0.1702 1.0000
6.000 0.7614 0.01110 0.00480 -0.0155 0.1638 1.0000
6.250 0.7883 0.01128 0.00501 -0.0152 0.1578 1.0000
6.500 0.8150 0.01147 0.00518 -0.0149 0.1509 1.0000
6.750 0.8412 0.01173 0.00543 -0.0145 0.1446 1.0000
7.000 0.8679 0.01192 0.00563 -0.0141 0.1385 1.0000
7.250 0.8937 0.01223 0.00590 -0.0137 0.1320 1.0000
7.500 0.9204 0.01242 0.00613 -0.0133 0.1261 1.0000
7.750 0.9457 0.01277 0.00642 -0.0129 0.1195 1.0000
8.000 0.9721 0.01299 0.00668 -0.0125 0.1140 1.0000
8.250 0.9970 0.01339 0.00702 -0.0120 0.1075 1.0000
8.500 1.0229 0.01366 0.00734 -0.0116 0.1027 1.0000
8.750 1.0476 0.01405 0.00772 -0.0111 0.0973 1.0000
9.000 1.0725 0.01443 0.00814 -0.0106 0.0927 1.0000
9.250 1.0975 0.01479 0.00851 -0.0101 0.0878 1.0000
9.500 1.1212 0.01529 0.00900 -0.0095 0.0827 1.0000
9.750 1.1462 0.01561 0.00938 -0.0090 0.0783 1.0000
10.000 1.1694 0.01613 0.00988 -0.0084 0.0733 1.0000
10.250 1.1939 0.01649 0.01030 -0.0078 0.0689 1.0000
10.500 1.2168 0.01702 0.01080 -0.0072 0.0639 1.0000
10.750 1.2403 0.01745 0.01129 -0.0066 0.0598 1.0000
11.000 1.2623 0.01803 0.01186 -0.0058 0.0555 1.0000
11.250 1.2849 0.01853 0.01242 -0.0051 0.0522 1.0000
11.500 1.3060 0.01916 0.01305 -0.0043 0.0481 1.0000
11.750 1.3274 0.01972 0.01367 -0.0035 0.0445 1.0000
12.000 1.3468 0.02046 0.01441 -0.0026 0.0404 1.0000
12.250 1.3660 0.02117 0.01515 -0.0016 0.0357 1.0000
12.500 1.3827 0.02208 0.01607 -0.0004 0.0300 1.0000
12.750 1.3972 0.02313 0.01714 0.0010 0.0243 1.0000
13.000 1.4086 0.02433 0.01836 0.0027 0.0191 1.0000
13.250 1.4175 0.02558 0.01965 0.0047 0.0166 1.0000
13.500 1.4225 0.02674 0.02088 0.0072 0.0155 1.0000
13.750 1.4226 0.02820 0.02242 0.0099 0.0146 1.0000
14.000 1.4237 0.02969 0.02402 0.0120 0.0141 1.0000
14.250 1.4217 0.03162 0.02606 0.0135 0.0136 1.0000
14.500 1.4182 0.03395 0.02852 0.0143 0.0133 1.0000
14.750 1.4117 0.03696 0.03166 0.0142 0.0130 1.0000
15.000 1.4021 0.04076 0.03560 0.0131 0.0128 1.0000
15.250 1.3895 0.04544 0.04042 0.0111 0.0127 1.0000
15.500 1.3721 0.05129 0.04644 0.0080 0.0125 1.0000
15.750 1.3505 0.05825 0.05358 0.0040 0.0125 1.0000
16.000 1.3260 0.06607 0.06158 -0.0006 0.0126 1.0000
16.250 1.2990 0.07462 0.07033 -0.0055 0.0128 1.0000
16.500 1.2653 0.08457 0.08046 -0.0112 0.0130 1.0000
16.750 1.2364 0.09368 0.08972 -0.0163 0.0132 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 426 AIRFOIL (e426-il)