EPPLER 335 AIRFOIL (e335-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 335 AIRFOIL (e335-il) Reynolds number: 1,000,000 Max Cl/Cd: 90.13 at α=9.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e335-il-1000000.txt Download as CSV file: xf-e335-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 335 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 -0.5189 0.08138 0.07990 -0.0138 1.0000 0.0078
-9.250 -0.5376 0.07518 0.07367 -0.0181 1.0000 0.0077
-9.000 -0.5562 0.07058 0.06902 -0.0187 1.0000 0.0078
-8.750 -0.5741 0.06692 0.06531 -0.0170 1.0000 0.0078
-8.500 -0.5917 0.06371 0.06203 -0.0136 1.0000 0.0077
-8.250 -0.5829 0.05859 0.05659 -0.0161 0.8825 0.0079
-8.000 -0.5914 0.05586 0.05347 -0.0122 0.8038 0.0080
-7.750 -0.5952 0.05263 0.04998 -0.0090 0.7687 0.0081
-7.500 -0.5951 0.04937 0.04649 -0.0059 0.7423 0.0083
-7.250 -0.5925 0.04610 0.04298 -0.0029 0.7211 0.0084
-5.250 -0.5046 0.01782 0.01168 0.0231 0.6066 0.0073
-4.750 -0.4499 0.01428 0.00771 0.0241 0.5813 0.0059
-4.500 -0.4244 0.01334 0.00663 0.0248 0.5688 0.0059
-4.250 -0.4000 0.01268 0.00585 0.0257 0.5569 0.0059
-3.750 -0.3521 0.01172 0.00471 0.0275 0.5349 0.0063
-3.500 -0.3278 0.01137 0.00429 0.0284 0.5248 0.0068
-3.250 -0.3047 0.01092 0.00376 0.0294 0.5144 0.0077
-3.000 -0.2799 0.01067 0.00348 0.0301 0.5044 0.0083
-2.750 -0.2548 0.01048 0.00323 0.0308 0.4948 0.0092
-2.500 -0.2305 0.01019 0.00289 0.0316 0.4857 0.0103
-2.250 -0.2053 0.01002 0.00268 0.0323 0.4772 0.0120
-2.000 -0.1812 0.00977 0.00243 0.0331 0.4683 0.0215
-1.750 -0.1588 0.00937 0.00224 0.0342 0.4606 0.0735
-1.500 -0.1370 0.00900 0.00210 0.0354 0.4527 0.1480
-1.250 -0.1251 0.00803 0.00188 0.0383 0.4466 0.3614
-1.000 -0.1283 0.00673 0.00158 0.0444 0.4413 0.6334
-0.750 -0.1293 0.00600 0.00174 0.0511 0.4356 0.8655
-0.500 -0.1054 0.00633 0.00206 0.0526 0.4290 0.8952
-0.250 -0.0751 0.00688 0.00258 0.0528 0.4218 0.9098
0.000 -0.0407 0.00736 0.00301 0.0520 0.4153 0.9171
0.250 0.0019 0.00790 0.00348 0.0494 0.4078 0.9210
0.500 0.0376 0.00833 0.00384 0.0481 0.4011 0.9261
0.750 0.0775 0.00873 0.00418 0.0460 0.3952 0.9302
1.000 0.1157 0.00891 0.00428 0.0438 0.3888 0.9308
1.250 0.1485 0.00894 0.00427 0.0428 0.3835 0.9313
1.500 0.1812 0.00898 0.00427 0.0417 0.3781 0.9318
1.750 0.2131 0.00905 0.00427 0.0407 0.3727 0.9323
2.000 0.2453 0.00908 0.00428 0.0397 0.3686 0.9329
2.250 0.2766 0.00912 0.00428 0.0390 0.3636 0.9335
2.500 0.3067 0.00918 0.00429 0.0384 0.3586 0.9342
2.750 0.3361 0.00922 0.00432 0.0380 0.3545 0.9351
3.000 0.3643 0.00923 0.00432 0.0378 0.3505 0.9361
3.250 0.3904 0.00927 0.00433 0.0381 0.3461 0.9374
3.500 0.3740 0.00912 0.00418 0.0475 0.3436 0.9445
3.750 0.4028 0.00913 0.00418 0.0472 0.3396 0.9450
4.000 0.4330 0.00914 0.00419 0.0465 0.3355 0.9455
4.250 0.4628 0.00918 0.00421 0.0460 0.3312 0.9459
4.500 0.4912 0.00926 0.00426 0.0457 0.3261 0.9464
4.750 0.5200 0.00926 0.00429 0.0453 0.3228 0.9469
5.000 0.5484 0.00929 0.00432 0.0450 0.3185 0.9474
5.250 0.5772 0.00937 0.00437 0.0446 0.3136 0.9479
5.500 0.6055 0.00944 0.00445 0.0443 0.3095 0.9486
5.750 0.6330 0.00946 0.00449 0.0442 0.3052 0.9492
6.000 0.6587 0.00952 0.00454 0.0445 0.3004 0.9500
6.250 0.6824 0.00960 0.00462 0.0451 0.2954 0.9509
6.500 0.7052 0.00961 0.00466 0.0460 0.2914 0.9521
6.750 0.7253 0.00965 0.00471 0.0474 0.2862 0.9539
7.000 0.7339 0.00966 0.00471 0.0513 0.2816 0.9568
7.250 0.7487 0.00963 0.00473 0.0539 0.2777 0.9585
7.500 0.7751 0.00967 0.00477 0.0539 0.2717 0.9591
7.750 0.8013 0.00977 0.00487 0.0539 0.2653 0.9597
8.000 0.8284 0.00985 0.00498 0.0538 0.2591 0.9602
8.250 0.8550 0.01001 0.00513 0.0536 0.2519 0.9609
8.500 0.8824 0.01012 0.00527 0.0534 0.2447 0.9615
8.750 0.9078 0.01030 0.00545 0.0535 0.2365 0.9622
9.000 0.9332 0.01046 0.00561 0.0536 0.2278 0.9629
9.250 0.9576 0.01064 0.00580 0.0539 0.2192 0.9638
9.500 0.9804 0.01089 0.00603 0.0544 0.2086 0.9649
9.750 1.0022 0.01112 0.00625 0.0551 0.1975 0.9661
10.000 1.0217 0.01139 0.00650 0.0563 0.1851 0.9676
10.250 1.0373 0.01168 0.00676 0.0582 0.1718 0.9697
10.500 1.0488 0.01199 0.00704 0.0609 0.1583 0.9719
10.750 1.0711 0.01250 0.00748 0.0611 0.1417 0.9730
11.000 1.0942 0.01302 0.00796 0.0611 0.1269 0.9740
11.250 1.1163 0.01357 0.00849 0.0612 0.1136 0.9750
11.500 1.1367 0.01421 0.00908 0.0615 0.0993 0.9762
11.750 1.1557 0.01485 0.00969 0.0620 0.0876 0.9777
12.000 1.1719 0.01556 0.01036 0.0629 0.0759 0.9797
12.250 1.1761 0.01615 0.01094 0.0663 0.0686 0.9838
12.500 1.1953 0.01711 0.01188 0.0659 0.0576 0.9848
12.750 1.2140 0.01813 0.01289 0.0653 0.0494 0.9860
13.000 1.2302 0.01935 0.01411 0.0646 0.0417 0.9874
13.250 1.2434 0.02086 0.01563 0.0636 0.0350 0.9894
13.500 1.2534 0.02276 0.01756 0.0623 0.0292 0.9918
13.750 1.2628 0.02486 0.01971 0.0607 0.0258 0.9943
14.000 1.2742 0.02733 0.02224 0.0581 0.0229 0.9958
14.250 1.2835 0.03004 0.02503 0.0555 0.0204 0.9975
14.500 1.2884 0.03330 0.02835 0.0527 0.0183 0.9990
14.750 1.2928 0.03651 0.03165 0.0502 0.0167 1.0000
15.000 1.2681 0.03914 0.03435 0.0537 0.0161 1.0000
15.250 1.2514 0.04206 0.03733 0.0554 0.0157 1.0000
15.500 1.2397 0.04536 0.04070 0.0557 0.0154 1.0000
15.750 1.2276 0.04904 0.04445 0.0553 0.0148 1.0000
16.000 1.2158 0.05282 0.04831 0.0547 0.0138 1.0000
16.250 1.2068 0.05647 0.05203 0.0540 0.0129 1.0000
16.500 1.1982 0.06025 0.05589 0.0531 0.0128 1.0000
16.750 1.1873 0.06442 0.06012 0.0519 0.0120 1.0000
17.000 1.1754 0.06883 0.06460 0.0506 0.0114 1.0000
17.250 1.1679 0.07276 0.06860 0.0494 0.0108 1.0000
17.500 1.1591 0.07696 0.07289 0.0480 0.0108 1.0000
17.750 1.1514 0.08106 0.07705 0.0466 0.0100 1.0000
18.000 1.1428 0.08530 0.08136 0.0451 0.0095 1.0000
18.250 1.1340 0.08968 0.08581 0.0435 0.0092 1.0000
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