EPPLER 325 AIRFOIL (e325-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 325 AIRFOIL (e325-il) Reynolds number: 200,000 Max Cl/Cd: 58.34 at α=9° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e325-il-200000.txt Download as CSV file: xf-e325-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 325 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5600 0.09292 0.08963 -0.0073 1.0000 0.0401
-9.750 -0.5731 0.08661 0.08332 -0.0122 1.0000 0.0404
-9.500 -0.5891 0.08144 0.07812 -0.0146 1.0000 0.0408
-9.250 -0.6059 0.07714 0.07378 -0.0152 1.0000 0.0408
-9.000 -0.6226 0.07351 0.07009 -0.0143 1.0000 0.0412
-8.750 -0.6417 0.07035 0.06683 -0.0116 1.0000 0.0423
-8.500 -0.6582 0.06753 0.06386 -0.0081 1.0000 0.0433
-8.250 -0.6914 0.06674 0.06249 -0.0012 1.0000 0.0445
-8.000 -0.6884 0.06062 0.05643 -0.0004 1.0000 0.0454
-7.750 -0.6777 0.05732 0.05320 0.0005 1.0000 0.0464
-7.000 -0.6253 0.03791 0.03378 0.0038 1.0000 0.0567
-6.750 -0.6162 0.03547 0.03130 0.0058 1.0000 0.0603
-6.250 -0.6076 0.02954 0.02501 0.0117 1.0000 0.0683
-6.000 -0.5983 0.02743 0.02283 0.0142 1.0000 0.0721
-5.750 -0.6305 0.03904 0.03357 0.0236 1.0000 0.0787
-5.250 -0.5584 0.02570 0.01855 0.0268 0.9796 0.0293
-5.000 -0.5066 0.02227 0.01472 0.0230 0.9513 0.0267
-4.750 -0.4480 0.01982 0.01197 0.0178 0.9165 0.0263
-4.500 -0.4011 0.01825 0.01027 0.0146 0.8732 0.0271
-4.250 -0.3728 0.01731 0.00916 0.0151 0.8338 0.0293
-4.000 -0.3523 0.01653 0.00830 0.0168 0.8023 0.0333
-3.750 -0.3323 0.01606 0.00768 0.0187 0.7755 0.0370
-3.500 -0.3173 0.01518 0.00676 0.0214 0.7535 0.0471
-3.250 -0.3070 0.01400 0.00590 0.0250 0.7343 0.1126
-3.000 -0.3062 0.01169 0.00676 0.0315 0.7185 0.7811
-2.750 -0.2852 0.01361 0.00850 0.0361 0.7002 0.8459
-2.500 -0.1407 0.01799 0.01233 0.0196 0.6729 0.8835
-2.250 -0.0703 0.01891 0.01295 0.0132 0.6509 0.9008
-2.000 -0.0135 0.01925 0.01305 0.0085 0.6319 0.9166
-1.750 0.0169 0.01920 0.01282 0.0079 0.6166 0.9237
-1.500 0.0452 0.01905 0.01251 0.0076 0.6022 0.9295
-1.250 0.0811 0.01880 0.01208 0.0059 0.5882 0.9334
-1.000 0.1080 0.01873 0.01185 0.0058 0.5759 0.9389
-0.750 0.1353 0.01858 0.01160 0.0056 0.5630 0.9443
-0.500 0.1706 0.01828 0.01119 0.0039 0.5504 0.9477
-0.250 0.2002 0.01815 0.01094 0.0032 0.5397 0.9521
0.000 0.2221 0.01821 0.01090 0.0041 0.5296 0.9577
0.250 0.2583 0.01782 0.01043 0.0021 0.5182 0.9605
0.500 0.2902 0.01763 0.01014 0.0009 0.5086 0.9641
0.750 0.3169 0.01755 0.01001 0.0008 0.4990 0.9680
1.000 0.3442 0.01749 0.00989 0.0005 0.4902 0.9721
1.250 0.3778 0.01720 0.00952 -0.0010 0.4815 0.9748
1.500 0.4077 0.01704 0.00935 -0.0018 0.4723 0.9777
1.750 0.4346 0.01703 0.00925 -0.0020 0.4649 0.9810
2.000 0.4621 0.01696 0.00920 -0.0023 0.4565 0.9840
2.250 0.4950 0.01672 0.00887 -0.0037 0.4491 0.9864
2.500 0.5252 0.01656 0.00878 -0.0046 0.4406 0.9890
2.750 0.5525 0.01652 0.00867 -0.0049 0.4339 0.9910
3.000 0.5808 0.01650 0.00870 -0.0054 0.4259 0.9938
3.250 0.6130 0.01629 0.00845 -0.0067 0.4188 0.9959
3.500 0.6411 0.01624 0.00846 -0.0072 0.4116 0.9975
3.750 0.6717 0.01614 0.00835 -0.0081 0.4044 0.9997
4.000 0.6956 0.01622 0.00846 -0.0077 0.3976 1.0000
4.250 0.7183 0.01629 0.00858 -0.0071 0.3907 1.0000
4.500 0.7409 0.01643 0.00870 -0.0064 0.3844 1.0000
4.750 0.7634 0.01651 0.00887 -0.0058 0.3770 1.0000
5.000 0.7860 0.01665 0.00898 -0.0051 0.3710 1.0000
5.250 0.8083 0.01676 0.00923 -0.0044 0.3636 1.0000
5.500 0.8307 0.01688 0.00934 -0.0036 0.3570 1.0000
5.750 0.8528 0.01705 0.00961 -0.0029 0.3498 1.0000
6.000 0.8749 0.01715 0.00975 -0.0021 0.3426 1.0000
6.250 0.8968 0.01733 0.01000 -0.0014 0.3353 1.0000
6.500 0.9186 0.01744 0.01018 -0.0005 0.3277 1.0000
6.750 0.9401 0.01761 0.01043 0.0003 0.3198 1.0000
7.000 0.9615 0.01771 0.01058 0.0012 0.3117 1.0000
7.250 0.9824 0.01787 0.01086 0.0021 0.3027 1.0000
7.500 1.0034 0.01801 0.01099 0.0031 0.2945 1.0000
7.750 1.0237 0.01815 0.01131 0.0041 0.2846 1.0000
8.000 1.0438 0.01832 0.01154 0.0051 0.2753 1.0000
8.250 1.0635 0.01849 0.01177 0.0063 0.2653 1.0000
8.500 1.0827 0.01868 0.01211 0.0074 0.2538 1.0000
8.750 1.1012 0.01891 0.01244 0.0087 0.2418 1.0000
9.000 1.1190 0.01918 0.01279 0.0100 0.2288 1.0000
9.250 1.1358 0.01951 0.01320 0.0114 0.2145 1.0000
9.500 1.1514 0.01993 0.01368 0.0130 0.1990 1.0000
9.750 1.1654 0.02045 0.01423 0.0148 0.1826 1.0000
10.000 1.1767 0.02113 0.01491 0.0168 0.1657 1.0000
10.250 1.1858 0.02193 0.01573 0.0192 0.1487 1.0000
10.500 1.1921 0.02283 0.01665 0.0219 0.1326 1.0000
10.750 1.1944 0.02385 0.01767 0.0251 0.1188 1.0000
11.000 1.1919 0.02495 0.01876 0.0290 0.1079 1.0000
11.250 1.1841 0.02604 0.01985 0.0336 0.0994 1.0000
11.500 1.1717 0.02696 0.02085 0.0391 0.0929 1.0000
11.750 1.1465 0.02766 0.02161 0.0465 0.0906 1.0000
12.000 1.1197 0.02861 0.02259 0.0534 0.0893 1.0000
12.250 1.0938 0.02992 0.02391 0.0591 0.0875 1.0000
12.500 1.0762 0.03173 0.02575 0.0626 0.0837 1.0000
12.750 1.0657 0.03394 0.02802 0.0643 0.0786 1.0000
13.000 1.0514 0.03689 0.03090 0.0654 0.0745 1.0000
13.250 1.0460 0.03952 0.03366 0.0656 0.0693 1.0000
13.500 1.0362 0.04268 0.03681 0.0655 0.0653 1.0000
13.750 1.0283 0.04579 0.03998 0.0655 0.0612 1.0000
14.000 1.0221 0.04893 0.04318 0.0650 0.0576 1.0000
14.250 1.0136 0.05214 0.04633 0.0649 0.0544 1.0000
14.500 1.0080 0.05541 0.04974 0.0642 0.0511 1.0000
14.750 1.0023 0.05872 0.05310 0.0633 0.0481 1.0000
15.000 0.9995 0.06134 0.05561 0.0635 0.0451 1.0000
15.250 0.9949 0.06489 0.05934 0.0624 0.0430 1.0000
15.500 0.9916 0.06825 0.06279 0.0613 0.0407 1.0000
15.750 0.9900 0.07135 0.06588 0.0604 0.0387 1.0000
16.000 0.9896 0.07415 0.06871 0.0602 0.0365 1.0000
16.250 0.9849 0.07805 0.07276 0.0586 0.0349 1.0000
16.500 0.9822 0.08167 0.07647 0.0572 0.0334 1.0000
16.750 0.9837 0.08450 0.07928 0.0563 0.0319 1.0000
17.000 0.9844 0.08734 0.08217 0.0561 0.0305 1.0000
17.250 0.9765 0.09211 0.08712 0.0539 0.0297 1.0000
17.500 0.9696 0.09679 0.09199 0.0517 0.0289 1.0000
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Polar data table (+)
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