EPPLER 68 AIRFOIL (e68-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 68 AIRFOIL (e68-il) Reynolds number: 200,000 Max Cl/Cd: 78.43 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e68-il-200000-n5.txt Download as CSV file: xf-e68-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 68 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
------ -------- --------- --------- -------- -------- --------
-12.000 -0.5776 0.07174 0.06786 -0.0686 1.0000 0.0141
-11.750 -0.6127 0.06325 0.05920 -0.0734 1.0000 0.0138
-11.500 -0.6392 0.05803 0.05390 -0.0751 1.0000 0.0138
-11.250 -0.6575 0.05203 0.04772 -0.0803 0.9975 0.0138
-11.000 -0.6635 0.04603 0.04139 -0.0882 0.9906 0.0138
-10.750 -0.6645 0.04184 0.03688 -0.0930 0.9834 0.0138
-10.500 -0.6632 0.03849 0.03322 -0.0943 0.9746 0.0140
-10.250 -0.6514 0.03564 0.03012 -0.0959 0.9688 0.0142
-10.000 -0.6379 0.03343 0.02770 -0.0964 0.9626 0.0144
-9.750 -0.6160 0.03141 0.02548 -0.0980 0.9592 0.0148
-9.500 -0.6006 0.02979 0.02366 -0.0976 0.9529 0.0151
-9.250 -0.5781 0.02829 0.02198 -0.0983 0.9486 0.0156
-9.000 -0.5509 0.02690 0.02038 -0.0997 0.9459 0.0165
-8.750 -0.5293 0.02564 0.01890 -0.0996 0.9412 0.0172
-8.500 -0.5065 0.02442 0.01748 -0.0996 0.9364 0.0178
-8.250 -0.4793 0.02304 0.01595 -0.1005 0.9335 0.0183
-8.000 -0.4499 0.02174 0.01456 -0.1019 0.9314 0.0191
-7.750 -0.4287 0.02089 0.01362 -0.1013 0.9257 0.0198
-7.500 -0.4023 0.02003 0.01266 -0.1017 0.9215 0.0208
-7.250 -0.3721 0.01917 0.01166 -0.1028 0.9187 0.0223
-7.000 -0.3404 0.01831 0.01068 -0.1041 0.9165 0.0241
-6.750 -0.3193 0.01772 0.01006 -0.1032 0.9101 0.0271
-6.500 -0.2911 0.01704 0.00931 -0.1037 0.9060 0.0318
-6.250 -0.2598 0.01634 0.00856 -0.1048 0.9032 0.0402
-6.000 -0.2309 0.01575 0.00797 -0.1054 0.8995 0.0522
-5.750 -0.2070 0.01533 0.00756 -0.1049 0.8934 0.0651
-5.500 -0.1763 0.01486 0.00711 -0.1058 0.8898 0.0832
-5.250 -0.1440 0.01435 0.00668 -0.1071 0.8870 0.1100
-5.000 -0.1216 0.01401 0.00643 -0.1063 0.8802 0.1342
-4.750 -0.0921 0.01364 0.00608 -0.1068 0.8758 0.1576
-4.500 -0.0600 0.01325 0.00573 -0.1079 0.8725 0.1848
-4.250 -0.0362 0.01300 0.00551 -0.1073 0.8659 0.2095
-4.000 -0.0074 0.01269 0.00526 -0.1077 0.8610 0.2401
-3.750 0.0243 0.01235 0.00499 -0.1086 0.8574 0.2766
-3.500 0.0477 0.01214 0.00486 -0.1079 0.8503 0.3106
-3.250 0.0768 0.01188 0.00468 -0.1083 0.8453 0.3480
-3.000 0.1056 0.01166 0.00453 -0.1086 0.8405 0.3851
-2.750 0.1311 0.01150 0.00444 -0.1081 0.8337 0.4191
-2.500 0.1611 0.01131 0.00430 -0.1086 0.8290 0.4536
-2.250 0.1869 0.01120 0.00425 -0.1082 0.8224 0.4852
-2.000 0.2148 0.01107 0.00417 -0.1082 0.8165 0.5160
-1.750 0.2433 0.01096 0.00409 -0.1082 0.8110 0.5450
-1.500 0.2691 0.01088 0.00408 -0.1077 0.8038 0.5727
-1.250 0.2989 0.01078 0.00399 -0.1080 0.7987 0.6001
-1.000 0.3235 0.01075 0.00401 -0.1073 0.7909 0.6257
-0.750 0.3520 0.01067 0.00397 -0.1073 0.7850 0.6508
-0.500 0.3774 0.01065 0.00400 -0.1067 0.7775 0.6747
-0.250 0.4049 0.01061 0.00398 -0.1064 0.7710 0.6984
0.000 0.4304 0.01060 0.00402 -0.1058 0.7635 0.7210
0.250 0.4572 0.01058 0.00403 -0.1054 0.7565 0.7431
0.500 0.4824 0.01059 0.00407 -0.1047 0.7488 0.7646
0.750 0.5085 0.01059 0.00409 -0.1041 0.7415 0.7849
1.000 0.5327 0.01061 0.00415 -0.1032 0.7334 0.8050
1.250 0.5588 0.01063 0.00417 -0.1026 0.7259 0.8245
1.500 0.5817 0.01066 0.00424 -0.1014 0.7171 0.8433
1.750 0.6068 0.01067 0.00425 -0.1006 0.7093 0.8619
2.000 0.6297 0.01071 0.00432 -0.0993 0.6998 0.8815
2.250 0.6538 0.01074 0.00436 -0.0984 0.6906 0.9013
2.500 0.6801 0.01076 0.00438 -0.0978 0.6813 0.9211
2.750 0.7079 0.01080 0.00445 -0.0978 0.6707 0.9428
3.000 0.7420 0.01086 0.00450 -0.0991 0.6600 0.9639
3.250 0.7788 0.01093 0.00455 -0.1011 0.6486 0.9914
3.500 0.8048 0.01105 0.00464 -0.1009 0.6369 1.0000
3.750 0.8285 0.01119 0.00475 -0.1002 0.6242 1.0000
4.000 0.8524 0.01134 0.00488 -0.0995 0.6113 1.0000
4.250 0.8762 0.01150 0.00502 -0.0988 0.5980 1.0000
4.500 0.8997 0.01168 0.00518 -0.0981 0.5840 1.0000
4.750 0.9227 0.01188 0.00535 -0.0972 0.5687 1.0000
5.000 0.9452 0.01209 0.00553 -0.0962 0.5524 1.0000
5.250 0.9670 0.01233 0.00573 -0.0951 0.5352 1.0000
5.500 0.9881 0.01260 0.00597 -0.0939 0.5172 1.0000
5.750 1.0085 0.01288 0.00622 -0.0925 0.4973 1.0000
6.000 1.0276 0.01320 0.00648 -0.0910 0.4762 1.0000
6.250 1.0459 0.01355 0.00678 -0.0893 0.4537 1.0000
6.500 1.0624 0.01396 0.00712 -0.0873 0.4298 1.0000
6.750 1.0778 0.01439 0.00748 -0.0851 0.4045 1.0000
7.000 1.0907 0.01487 0.00787 -0.0825 0.3784 1.0000
7.250 1.1022 0.01543 0.00833 -0.0797 0.3505 1.0000
7.500 1.1129 0.01606 0.00884 -0.0769 0.3217 1.0000
7.750 1.1234 0.01675 0.00942 -0.0742 0.2936 1.0000
8.000 1.1338 0.01750 0.01007 -0.0716 0.2670 1.0000
8.250 1.1450 0.01825 0.01075 -0.0691 0.2422 1.0000
8.500 1.1555 0.01907 0.01149 -0.0667 0.2180 1.0000
8.750 1.1646 0.01999 0.01230 -0.0642 0.1931 1.0000
9.000 1.1731 0.02100 0.01319 -0.0618 0.1681 1.0000
9.250 1.1810 0.02209 0.01415 -0.0594 0.1424 1.0000
9.500 1.1887 0.02323 0.01520 -0.0571 0.1192 1.0000
9.750 1.1968 0.02440 0.01627 -0.0550 0.1006 1.0000
10.000 1.2062 0.02553 0.01736 -0.0530 0.0870 1.0000
10.250 1.2156 0.02668 0.01849 -0.0512 0.0751 1.0000
10.500 1.2248 0.02788 0.01969 -0.0494 0.0641 1.0000
10.750 1.2330 0.02919 0.02099 -0.0477 0.0537 1.0000
11.000 1.2411 0.03054 0.02236 -0.0460 0.0454 1.0000
11.250 1.2483 0.03201 0.02386 -0.0444 0.0383 1.0000
11.500 1.2537 0.03366 0.02553 -0.0427 0.0325 1.0000
11.750 1.2605 0.03525 0.02720 -0.0412 0.0282 1.0000
12.000 1.2641 0.03716 0.02913 -0.0397 0.0250 1.0000
12.250 1.2696 0.03896 0.03105 -0.0383 0.0227 1.0000
12.500 1.2739 0.04092 0.03310 -0.0371 0.0209 1.0000
12.750 1.2745 0.04329 0.03551 -0.0359 0.0193 1.0000
13.000 1.2793 0.04533 0.03769 -0.0350 0.0181 1.0000
13.250 1.2824 0.04760 0.04009 -0.0341 0.0171 1.0000
13.500 1.2843 0.05007 0.04270 -0.0335 0.0163 1.0000
13.750 1.2853 0.05274 0.04548 -0.0329 0.0158 1.0000
14.000 1.2840 0.05574 0.04858 -0.0325 0.0152 1.0000
14.250 1.2810 0.05906 0.05200 -0.0323 0.0148 1.0000
14.500 1.2793 0.06236 0.05543 -0.0323 0.0145 1.0000
14.750 1.2788 0.06559 0.05881 -0.0324 0.0141 1.0000
15.000 1.2777 0.06898 0.06235 -0.0327 0.0139 1.0000
15.250 1.2763 0.07253 0.06604 -0.0331 0.0135 1.0000
15.500 1.2746 0.07620 0.06985 -0.0337 0.0132 1.0000
15.750 1.2726 0.08000 0.07378 -0.0345 0.0128 1.0000
16.000 1.2703 0.08395 0.07786 -0.0355 0.0125 1.0000
16.250 1.2676 0.08806 0.08210 -0.0368 0.0122 1.0000
16.500 1.2642 0.09235 0.08651 -0.0382 0.0119 1.0000
16.750 1.2599 0.09688 0.09115 -0.0398 0.0116 1.0000
17.000 1.2550 0.10157 0.09597 -0.0416 0.0113 1.0000
17.250 1.2502 0.10630 0.10081 -0.0435 0.0112 1.0000
17.500 1.2449 0.11116 0.10578 -0.0455 0.0109 1.0000
17.750 1.2405 0.11609 0.11088 -0.0477 0.0108 1.0000
18.000 1.2355 0.12116 0.11611 -0.0501 0.0107 1.0000
18.250 1.2296 0.12653 0.12166 -0.0529 0.0106 1.0000
18.500 1.2233 0.13208 0.12738 -0.0558 0.0105 1.0000
18.750 1.2164 0.13784 0.13330 -0.0591 0.0104 1.0000
19.000 1.2081 0.14408 0.13971 -0.0627 0.0103 1.0000
19.250 1.1994 0.15054 0.14634 -0.0667 0.0103 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 68 AIRFOIL (e68-il)