EPPLER 342 AIRFOIL (e342-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 342 AIRFOIL (e342-il) Reynolds number: 200,000 Max Cl/Cd: 61.86 at α=10° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e342-il-200000-n5.txt Download as CSV file: xf-e342-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 342 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.3600 0.09792 0.09459 -0.0275 1.0000 0.0248
-9.750 -0.3588 0.09428 0.09098 -0.0288 1.0000 0.0249
-9.500 -0.3594 0.09035 0.08711 -0.0305 1.0000 0.0250
-9.250 -0.2751 0.06855 0.06501 -0.0464 0.7839 0.0171
-8.750 -0.2871 0.06082 0.05708 -0.0498 0.7397 0.0165
-8.500 -0.3030 0.05636 0.05251 -0.0509 0.7246 0.0162
-8.250 -0.3968 0.05840 0.05419 -0.0493 0.7663 0.0140
-8.000 -0.3997 0.05608 0.05167 -0.0470 0.7415 0.0139
-7.750 -0.3996 0.05385 0.04925 -0.0448 0.7209 0.0137
-7.500 -0.3993 0.05124 0.04643 -0.0424 0.7037 0.0135
-7.000 -0.3960 0.04498 0.03964 -0.0369 0.6746 0.0128
-6.750 -0.4065 0.03627 0.03007 -0.0305 0.6676 0.0113
-6.500 -0.3962 0.03389 0.02736 -0.0279 0.6544 0.0112
-6.250 -0.3842 0.03137 0.02445 -0.0253 0.6419 0.0111
-6.000 -0.3692 0.02911 0.02180 -0.0230 0.6305 0.0111
-5.750 -0.3515 0.02700 0.01928 -0.0210 0.6191 0.0110
-5.500 -0.3309 0.02492 0.01677 -0.0194 0.6080 0.0111
-5.250 -0.3074 0.02317 0.01460 -0.0183 0.5973 0.0112
-5.000 -0.2821 0.02177 0.01289 -0.0175 0.5858 0.0113
-4.750 -0.2571 0.02078 0.01175 -0.0169 0.5753 0.0115
-4.500 -0.2318 0.01995 0.01077 -0.0163 0.5657 0.0120
-4.250 -0.2059 0.01913 0.00982 -0.0158 0.5561 0.0125
-4.000 -0.1807 0.01837 0.00888 -0.0151 0.5474 0.0134
-3.750 -0.1560 0.01769 0.00809 -0.0143 0.5382 0.0139
-3.500 -0.1326 0.01724 0.00757 -0.0135 0.5299 0.0148
-3.250 -0.1085 0.01685 0.00710 -0.0127 0.5212 0.0172
-3.000 -0.0854 0.01649 0.00669 -0.0117 0.5138 0.0191
-2.750 -0.0624 0.01606 0.00621 -0.0107 0.5065 0.0215
-2.500 -0.0395 0.01571 0.00577 -0.0096 0.4996 0.0249
-2.250 -0.0166 0.01540 0.00545 -0.0086 0.4933 0.0321
-2.000 0.0058 0.01505 0.00511 -0.0074 0.4863 0.0446
-1.750 0.0272 0.01470 0.00485 -0.0062 0.4803 0.0763
-1.500 0.0935 0.01289 0.00606 -0.0127 0.4719 0.8145
-1.250 0.0926 0.01313 0.00623 -0.0064 0.4673 0.8576
-1.000 0.1242 0.01396 0.00686 -0.0058 0.4621 0.8859
-0.750 0.1610 0.01470 0.00748 -0.0065 0.4559 0.9079
-0.500 0.2274 0.01541 0.00795 -0.0133 0.4485 0.9219
-0.250 0.2690 0.01556 0.00794 -0.0159 0.4430 0.9270
0.000 0.2999 0.01555 0.00784 -0.0166 0.4377 0.9293
0.250 0.3281 0.01557 0.00774 -0.0167 0.4332 0.9321
0.500 0.3515 0.01562 0.00767 -0.0159 0.4294 0.9358
0.750 0.3729 0.01565 0.00764 -0.0148 0.4254 0.9393
1.000 0.4045 0.01563 0.00756 -0.0157 0.4205 0.9406
1.250 0.4345 0.01565 0.00749 -0.0163 0.4163 0.9423
1.500 0.4631 0.01570 0.00743 -0.0166 0.4126 0.9441
1.750 0.4903 0.01572 0.00743 -0.0166 0.4089 0.9462
2.000 0.5153 0.01577 0.00745 -0.0162 0.4051 0.9487
2.250 0.5353 0.01584 0.00749 -0.0148 0.4016 0.9516
2.500 0.5616 0.01589 0.00747 -0.0147 0.3980 0.9532
2.750 0.5913 0.01595 0.00747 -0.0153 0.3947 0.9544
3.000 0.6196 0.01599 0.00753 -0.0156 0.3912 0.9557
3.250 0.6467 0.01605 0.00759 -0.0157 0.3875 0.9570
3.500 0.6732 0.01612 0.00765 -0.0157 0.3842 0.9587
3.750 0.6978 0.01623 0.00770 -0.0153 0.3810 0.9604
4.000 0.7207 0.01634 0.00780 -0.0145 0.3782 0.9621
4.250 0.7404 0.01646 0.00798 -0.0131 0.3751 0.9644
4.500 0.7686 0.01653 0.00807 -0.0134 0.3715 0.9653
4.750 0.7953 0.01662 0.00817 -0.0135 0.3681 0.9661
5.000 0.8219 0.01675 0.00826 -0.0136 0.3653 0.9671
5.250 0.8487 0.01689 0.00842 -0.0137 0.3623 0.9683
5.500 0.8742 0.01701 0.00862 -0.0136 0.3588 0.9697
5.750 0.8979 0.01714 0.00881 -0.0130 0.3552 0.9708
6.000 0.9208 0.01728 0.00899 -0.0124 0.3521 0.9720
6.250 0.9435 0.01745 0.00915 -0.0117 0.3492 0.9735
6.500 0.9642 0.01765 0.00937 -0.0106 0.3463 0.9751
6.750 0.9883 0.01781 0.00965 -0.0102 0.3428 0.9759
7.000 1.0132 0.01797 0.00989 -0.0101 0.3389 0.9767
7.250 1.0376 0.01813 0.01009 -0.0098 0.3354 0.9776
7.500 1.0619 0.01832 0.01028 -0.0095 0.3322 0.9785
7.750 1.0860 0.01852 0.01061 -0.0093 0.3282 0.9799
8.000 1.1088 0.01873 0.01093 -0.0088 0.3241 0.9812
8.250 1.1300 0.01892 0.01119 -0.0079 0.3202 0.9823
8.500 1.1507 0.01914 0.01141 -0.0070 0.3166 0.9835
8.750 1.1694 0.01939 0.01180 -0.0057 0.3122 0.9848
9.000 1.1901 0.01962 0.01216 -0.0049 0.3074 0.9859
9.250 1.2130 0.01983 0.01241 -0.0045 0.3029 0.9869
9.500 1.2360 0.02010 0.01278 -0.0042 0.2982 0.9882
9.750 1.2575 0.02037 0.01321 -0.0037 0.2923 0.9897
10.000 1.2767 0.02064 0.01352 -0.0027 0.2872 0.9911
10.250 1.2947 0.02097 0.01399 -0.0015 0.2814 0.9926
10.500 1.3115 0.02131 0.01443 -0.0002 0.2753 0.9942
10.750 1.3306 0.02167 0.01487 0.0005 0.2687 0.9955
11.000 1.3483 0.02207 0.01539 0.0014 0.2614 0.9972
11.250 1.3640 0.02252 0.01593 0.0026 0.2543 0.9990
11.500 1.3701 0.02299 0.01650 0.0056 0.2471 1.0000
11.750 1.3588 0.02341 0.01693 0.0118 0.2424 1.0000
12.000 1.3503 0.02391 0.01754 0.0173 0.2367 1.0000
12.250 1.3382 0.02456 0.01823 0.0229 0.2312 1.0000
12.500 1.3267 0.02535 0.01909 0.0279 0.2260 1.0000
12.750 1.3159 0.02633 0.02013 0.0323 0.2202 1.0000
13.000 1.3045 0.02764 0.02145 0.0360 0.2144 1.0000
13.250 1.2983 0.02914 0.02303 0.0385 0.2072 1.0000
13.500 1.2901 0.03111 0.02502 0.0403 0.2005 1.0000
13.750 1.2852 0.03322 0.02722 0.0414 0.1928 1.0000
14.000 1.2767 0.03583 0.02985 0.0422 0.1856 1.0000
14.250 1.2686 0.03864 0.03273 0.0426 0.1777 1.0000
14.500 1.2591 0.04172 0.03585 0.0427 0.1710 1.0000
14.750 1.2485 0.04506 0.03924 0.0426 0.1637 1.0000
15.000 1.2377 0.04853 0.04276 0.0424 0.1572 1.0000
15.250 1.2260 0.05222 0.04649 0.0419 0.1505 1.0000
15.500 1.2149 0.05602 0.05034 0.0413 0.1441 1.0000
15.750 1.2033 0.06002 0.05438 0.0404 0.1375 1.0000
16.000 1.1932 0.06398 0.05838 0.0395 0.1315 1.0000
16.250 1.1832 0.06804 0.06249 0.0384 0.1253 1.0000
16.500 1.1734 0.07216 0.06665 0.0372 0.1195 1.0000
16.750 1.1642 0.07632 0.07085 0.0359 0.1135 1.0000
17.000 1.1558 0.08045 0.07501 0.0346 0.1083 1.0000
17.250 1.1490 0.08445 0.07908 0.0333 0.1029 1.0000
17.500 1.1401 0.08878 0.08342 0.0317 0.0979 1.0000
17.750 1.1343 0.09283 0.08755 0.0302 0.0927 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 342 AIRFOIL (e342-il)