EPPLER 332 AIRFOIL (e332-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 332 AIRFOIL (e332-il) Reynolds number: 1,000,000 Max Cl/Cd: 107.88 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e332-il-1000000.txt Download as CSV file: xf-e332-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 332 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4133 0.08377 0.08227 -0.0193 1.0000 0.0081
-8.500 -0.4180 0.07892 0.07744 -0.0234 1.0000 0.0081
-8.250 -0.4251 0.07476 0.07328 -0.0264 1.0000 0.0081
-8.000 -0.4362 0.07109 0.06959 -0.0266 1.0000 0.0081
-7.750 -0.4086 0.06420 0.06230 -0.0363 0.8643 0.0081
-7.500 -0.4132 0.06073 0.05853 -0.0348 0.8055 0.0081
-7.250 -0.4118 0.05753 0.05513 -0.0337 0.7729 0.0081
-7.000 -0.4085 0.05407 0.05147 -0.0324 0.7469 0.0081
-6.750 -0.4026 0.05054 0.04776 -0.0310 0.7272 0.0081
-6.500 -0.4083 0.04491 0.04189 -0.0287 0.7104 0.0083
-6.250 -0.3994 0.04236 0.03916 -0.0272 0.6926 0.0084
-6.000 -0.3874 0.04009 0.03672 -0.0258 0.6764 0.0086
-5.750 -0.3737 0.03780 0.03425 -0.0243 0.6607 0.0087
-5.500 -0.3588 0.03543 0.03169 -0.0228 0.6466 0.0089
-5.250 -0.3428 0.03303 0.02909 -0.0211 0.6339 0.0092
-5.000 -0.3260 0.03051 0.02635 -0.0192 0.6216 0.0095
-4.750 -0.3079 0.02802 0.02360 -0.0173 0.6101 0.0100
-4.500 -0.2841 0.02561 0.02082 -0.0148 0.5984 0.0110
-4.250 -0.2721 0.02141 0.01615 -0.0118 0.5886 0.0116
-4.000 -0.2497 0.02028 0.01490 -0.0110 0.5773 0.0119
-3.750 -0.2268 0.01925 0.01371 -0.0101 0.5661 0.0124
-3.500 -0.2241 0.11791 0.11434 0.0317 0.5583 0.0114
-3.250 -0.1752 0.01848 0.01249 -0.0081 0.5446 0.0149
-3.000 0.0038 0.16966 0.16639 -0.0072 0.5064 0.0072
-2.750 -0.1240 0.01192 0.00527 -0.0052 0.5266 0.0072
-2.500 -0.0997 0.01101 0.00421 -0.0041 0.5178 0.0067
-2.250 -0.0753 0.01047 0.00357 -0.0031 0.5086 0.0065
-2.000 -0.0504 0.01008 0.00310 -0.0023 0.5000 0.0065
-1.500 0.0005 0.00953 0.00243 -0.0010 0.4826 0.0069
-1.250 0.0264 0.00935 0.00218 -0.0005 0.4746 0.0076
-1.000 0.0527 0.00917 0.00195 0.0000 0.4671 0.0087
-0.750 0.0772 0.00885 0.00172 0.0008 0.4596 0.0440
-0.500 0.1024 0.00860 0.00165 0.0013 0.4524 0.1022
-0.250 0.1189 0.00767 0.00152 0.0032 0.4457 0.3749
0.000 0.1119 0.00595 0.00157 0.0106 0.4416 0.8855
0.250 0.1335 0.00624 0.00191 0.0128 0.4353 0.9213
0.500 0.1574 0.00650 0.00212 0.0142 0.4289 0.9343
0.750 0.1854 0.00672 0.00230 0.0146 0.4224 0.9412
1.000 0.2150 0.00704 0.00254 0.0148 0.4160 0.9498
1.250 0.2612 0.00739 0.00284 0.0113 0.4092 0.9542
1.500 0.3005 0.00763 0.00299 0.0090 0.4024 0.9575
1.750 0.3133 0.00761 0.00295 0.0123 0.3982 0.9615
2.000 0.3445 0.00763 0.00292 0.0115 0.3922 0.9621
2.250 0.3751 0.00769 0.00292 0.0108 0.3862 0.9626
2.500 0.4065 0.00771 0.00292 0.0100 0.3816 0.9631
2.750 0.4375 0.00775 0.00292 0.0092 0.3759 0.9636
3.000 0.4677 0.00782 0.00295 0.0086 0.3705 0.9642
3.250 0.4984 0.00785 0.00297 0.0079 0.3659 0.9648
3.500 0.5282 0.00791 0.00300 0.0074 0.3605 0.9654
3.750 0.5568 0.00798 0.00303 0.0071 0.3555 0.9662
4.000 0.5853 0.00802 0.00307 0.0068 0.3509 0.9670
4.250 0.6136 0.00809 0.00312 0.0066 0.3455 0.9680
4.500 0.6408 0.00817 0.00318 0.0066 0.3408 0.9692
4.750 0.6664 0.00820 0.00321 0.0069 0.3360 0.9704
5.000 0.6879 0.00826 0.00326 0.0081 0.3312 0.9723
5.250 0.7042 0.00833 0.00333 0.0105 0.3267 0.9749
5.500 0.7352 0.00837 0.00339 0.0096 0.3217 0.9753
5.750 0.7651 0.00847 0.00347 0.0090 0.3159 0.9757
6.000 0.7950 0.00856 0.00356 0.0083 0.3108 0.9761
6.250 0.8248 0.00864 0.00365 0.0076 0.3050 0.9766
6.500 0.8539 0.00878 0.00377 0.0071 0.2988 0.9772
6.750 0.8834 0.00885 0.00388 0.0065 0.2935 0.9778
7.000 0.9120 0.00900 0.00401 0.0060 0.2866 0.9784
7.250 0.9406 0.00910 0.00414 0.0056 0.2803 0.9792
7.500 0.9684 0.00925 0.00429 0.0053 0.2731 0.9801
7.750 0.9962 0.00939 0.00444 0.0050 0.2660 0.9811
8.000 1.0228 0.00959 0.00462 0.0049 0.2565 0.9824
8.250 1.0486 0.00972 0.00479 0.0050 0.2490 0.9837
8.500 1.0696 0.00992 0.00498 0.0061 0.2402 0.9857
8.750 1.0967 0.01017 0.00520 0.0058 0.2281 0.9865
9.250 1.1537 0.01077 0.00575 0.0044 0.2024 0.9877
9.500 1.1811 0.01114 0.00609 0.0038 0.1865 0.9885
9.750 1.2075 0.01159 0.00647 0.0034 0.1689 0.9895
10.000 1.2329 0.01209 0.00691 0.0031 0.1521 0.9907
10.250 1.2570 0.01266 0.00742 0.0030 0.1348 0.9921
10.500 1.2793 0.01331 0.00799 0.0032 0.1156 0.9938
10.750 1.3017 0.01397 0.00858 0.0033 0.1007 0.9954
11.000 1.3261 0.01478 0.00929 0.0027 0.0820 0.9966
11.250 1.3502 0.01551 0.00998 0.0022 0.0697 0.9980
11.500 1.3715 0.01648 0.01086 0.0020 0.0549 0.9996
11.750 1.3824 0.01694 0.01137 0.0044 0.0511 1.0000
12.000 1.3761 0.01744 0.01187 0.0102 0.0468 1.0000
12.250 1.3662 0.01791 0.01234 0.0166 0.0421 1.0000
12.500 1.3665 0.01843 0.01291 0.0208 0.0403 1.0000
12.750 1.3670 0.01932 0.01381 0.0242 0.0357 1.0000
13.000 1.3662 0.02053 0.01500 0.0271 0.0298 1.0000
13.250 1.3702 0.02171 0.01624 0.0289 0.0273 1.0000
13.500 1.3714 0.02330 0.01786 0.0303 0.0241 1.0000
13.750 1.3714 0.02524 0.01983 0.0312 0.0206 1.0000
14.000 1.3742 0.02716 0.02183 0.0316 0.0194 1.0000
14.250 1.3722 0.02970 0.02443 0.0318 0.0173 1.0000
14.500 1.3721 0.03220 0.02701 0.0317 0.0159 1.0000
14.750 1.3700 0.03501 0.02989 0.0314 0.0146 1.0000
15.000 1.3625 0.03846 0.03340 0.0309 0.0129 1.0000
15.250 1.3583 0.04163 0.03666 0.0304 0.0124 1.0000
15.500 1.3537 0.04486 0.03998 0.0298 0.0116 1.0000
15.750 1.3464 0.04848 0.04368 0.0290 0.0108 1.0000
16.000 1.3392 0.05212 0.04742 0.0282 0.0108 1.0000
16.250 1.3268 0.05645 0.05180 0.0271 0.0095 1.0000
16.500 1.3163 0.06072 0.05616 0.0259 0.0091 1.0000
16.750 1.3109 0.06447 0.05999 0.0248 0.0086 1.0000
17.000 1.3023 0.06870 0.06431 0.0234 0.0082 1.0000
17.250 1.2945 0.07288 0.06858 0.0220 0.0080 1.0000
17.500 1.2850 0.07743 0.07321 0.0204 0.0077 1.0000
17.750 1.2749 0.08214 0.07801 0.0187 0.0075 1.0000
18.000 1.2643 0.08698 0.08291 0.0169 0.0069 1.0000
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Polar data table (+)
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