EPPLER 341 AIRFOIL (e341-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: EPPLER 341 AIRFOIL (e341-il) Reynolds number: 200,000 Max Cl/Cd: 60.13 at α=10.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e341-il-200000-n5.txt Download as CSV file: xf-e341-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 341 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4790 0.09350 0.09037 -0.0141 1.0000 0.0275
-9.500 -0.4868 0.08707 0.08392 -0.0198 1.0000 0.0275
-9.250 -0.4977 0.08191 0.07871 -0.0226 1.0000 0.0275
-9.000 -0.5099 0.07752 0.07426 -0.0234 1.0000 0.0275
-8.750 -0.5191 0.07338 0.06999 -0.0237 0.9106 0.0275
-8.250 -0.5473 0.05925 0.05524 -0.0193 0.8064 0.0146
-8.000 -0.5470 0.05674 0.05250 -0.0173 0.7750 0.0144
-7.750 -0.5453 0.05388 0.04940 -0.0153 0.7509 0.0142
-7.500 -0.5416 0.05073 0.04597 -0.0132 0.7308 0.0140
-7.250 -0.5363 0.04731 0.04225 -0.0110 0.7133 0.0136
-7.000 -0.5298 0.04358 0.03815 -0.0086 0.6979 0.0132
-6.750 -0.5227 0.03914 0.03325 -0.0056 0.6842 0.0127
-6.250 -0.5006 0.02941 0.02198 0.0012 0.6601 0.0116
-6.000 -0.4804 0.02734 0.01950 0.0027 0.6462 0.0116
-5.500 -0.4338 0.02397 0.01545 0.0047 0.6198 0.0117
-5.250 -0.4087 0.02269 0.01396 0.0053 0.6076 0.0119
-5.000 -0.3830 0.02161 0.01268 0.0058 0.5957 0.0122
-4.750 -0.3569 0.02066 0.01155 0.0063 0.5840 0.0126
-4.500 -0.3307 0.01977 0.01051 0.0068 0.5728 0.0130
-4.250 -0.3049 0.01898 0.00957 0.0073 0.5628 0.0135
-4.000 -0.2798 0.01825 0.00872 0.0080 0.5529 0.0142
-3.750 -0.2553 0.01763 0.00800 0.0087 0.5438 0.0150
-3.500 -0.2307 0.01722 0.00754 0.0093 0.5345 0.0167
-3.250 -0.2062 0.01676 0.00700 0.0100 0.5256 0.0188
-3.000 -0.1822 0.01634 0.00650 0.0108 0.5171 0.0207
-2.750 -0.1580 0.01591 0.00600 0.0116 0.5089 0.0235
-2.500 -0.1339 0.01554 0.00558 0.0124 0.5015 0.0292
-2.250 -0.1104 0.01513 0.00521 0.0133 0.4945 0.0456
-2.000 -0.0877 0.01470 0.00492 0.0142 0.4873 0.0866
-1.750 -0.0696 0.01392 0.00463 0.0158 0.4811 0.2076
-1.500 -0.0327 0.01315 0.00644 0.0166 0.4733 0.8566
-1.250 -0.0098 0.01382 0.00693 0.0186 0.4674 0.8834
-1.000 0.0304 0.01459 0.00755 0.0174 0.4604 0.9034
-0.750 0.0806 0.01526 0.00801 0.0139 0.4534 0.9186
-0.500 0.1247 0.01558 0.00816 0.0112 0.4470 0.9282
-0.250 0.1555 0.01559 0.00806 0.0105 0.4409 0.9319
0.000 0.1793 0.01562 0.00794 0.0112 0.4360 0.9358
0.250 0.2059 0.01561 0.00784 0.0114 0.4312 0.9386
0.500 0.2383 0.01558 0.00773 0.0103 0.4261 0.9401
0.750 0.2693 0.01558 0.00761 0.0095 0.4214 0.9417
1.000 0.2990 0.01559 0.00752 0.0089 0.4170 0.9437
1.250 0.3272 0.01559 0.00748 0.0087 0.4120 0.9458
1.500 0.3529 0.01562 0.00745 0.0090 0.4077 0.9484
1.750 0.3743 0.01568 0.00743 0.0101 0.4043 0.9512
2.000 0.4056 0.01570 0.00738 0.0092 0.4004 0.9523
2.250 0.4361 0.01570 0.00737 0.0084 0.3959 0.9534
2.500 0.4653 0.01573 0.00736 0.0079 0.3917 0.9547
2.750 0.4940 0.01578 0.00735 0.0075 0.3883 0.9562
3.000 0.5214 0.01585 0.00738 0.0073 0.3849 0.9577
3.250 0.5475 0.01590 0.00745 0.0075 0.3809 0.9593
3.500 0.5715 0.01598 0.00754 0.0080 0.3771 0.9613
3.750 0.5961 0.01607 0.00759 0.0084 0.3739 0.9627
4.000 0.6248 0.01616 0.00763 0.0079 0.3709 0.9635
4.250 0.6542 0.01623 0.00774 0.0073 0.3672 0.9645
4.500 0.6828 0.01630 0.00787 0.0069 0.3633 0.9656
4.750 0.7099 0.01640 0.00797 0.0067 0.3597 0.9666
5.000 0.7359 0.01652 0.00806 0.0068 0.3566 0.9676
5.250 0.7618 0.01665 0.00822 0.0068 0.3533 0.9688
5.500 0.7869 0.01678 0.00843 0.0071 0.3497 0.9703
5.750 0.8099 0.01692 0.00861 0.0077 0.3462 0.9716
6.000 0.8343 0.01706 0.00878 0.0081 0.3428 0.9725
6.250 0.8617 0.01722 0.00891 0.0078 0.3396 0.9732
6.500 0.8896 0.01736 0.00916 0.0073 0.3357 0.9741
6.750 0.9164 0.01751 0.00941 0.0071 0.3316 0.9751
7.000 0.9417 0.01766 0.00961 0.0072 0.3279 0.9760
7.250 0.9664 0.01785 0.00979 0.0075 0.3246 0.9768
7.500 0.9903 0.01804 0.01011 0.0078 0.3203 0.9778
7.750 1.0135 0.01822 0.01040 0.0083 0.3158 0.9788
8.000 1.0363 0.01841 0.01063 0.0088 0.3117 0.9803
8.250 1.0598 0.01863 0.01089 0.0092 0.3079 0.9813
8.500 1.0845 0.01884 0.01126 0.0093 0.3028 0.9820
8.750 1.1085 0.01903 0.01153 0.0095 0.2976 0.9828
9.000 1.1318 0.01926 0.01180 0.0099 0.2931 0.9837
9.250 1.1543 0.01951 0.01222 0.0103 0.2872 0.9847
9.500 1.1758 0.01974 0.01252 0.0109 0.2815 0.9859
9.750 1.1969 0.02003 0.01291 0.0116 0.2754 0.9873
10.000 1.2180 0.02032 0.01332 0.0122 0.2687 0.9887
10.250 1.2402 0.02063 0.01372 0.0125 0.2621 0.9897
10.500 1.2610 0.02097 0.01417 0.0130 0.2542 0.9909
10.750 1.2804 0.02137 0.01467 0.0137 0.2464 0.9922
11.000 1.2978 0.02182 0.01519 0.0146 0.2380 0.9937
11.250 1.3145 0.02233 0.01582 0.0156 0.2289 0.9953
11.500 1.3320 0.02295 0.01650 0.0162 0.2189 0.9967
11.750 1.3461 0.02371 0.01731 0.0171 0.2086 0.9986
12.000 1.3533 0.02449 0.01819 0.0192 0.1986 1.0000
12.250 1.3370 0.02520 0.01893 0.0257 0.1933 1.0000
12.500 1.3209 0.02606 0.01986 0.0314 0.1875 1.0000
12.750 1.3020 0.02729 0.02112 0.0367 0.1824 1.0000
13.000 1.2879 0.02872 0.02261 0.0406 0.1762 1.0000
13.250 1.2745 0.03069 0.02462 0.0432 0.1695 1.0000
13.500 1.2679 0.03304 0.02703 0.0440 0.1616 1.0000
13.750 1.2595 0.03606 0.03008 0.0439 0.1542 1.0000
14.000 1.2542 0.03911 0.03321 0.0433 0.1462 1.0000
14.250 1.2451 0.04269 0.03684 0.0424 0.1395 1.0000
14.500 1.2363 0.04638 0.04059 0.0415 0.1325 1.0000
14.750 1.2253 0.05036 0.04462 0.0404 0.1265 1.0000
15.000 1.2149 0.05434 0.04867 0.0392 0.1201 1.0000
15.250 1.2026 0.05863 0.05299 0.0379 0.1149 1.0000
15.500 1.1937 0.06270 0.05714 0.0366 0.1089 1.0000
16.000 1.1742 0.07133 0.06589 0.0336 0.0984 1.0000
16.250 1.1635 0.07589 0.07048 0.0320 0.0935 1.0000
16.500 1.1561 0.08008 0.07474 0.0305 0.0889 1.0000
16.750 1.1483 0.08439 0.07910 0.0289 0.0842 1.0000
17.000 1.1388 0.08898 0.08372 0.0272 0.0802 1.0000
17.250 1.1336 0.09307 0.08789 0.0256 0.0758 1.0000
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