EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il) Reynolds number: 1,000,000 Max Cl/Cd: 76.02 at α=9° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ea61009-il-1000000-n5.txt Download as CSV file: xf-ea61009-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER EA 6(-1)-009 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.500 -1.2783 0.06855 0.06583 -0.0090 1.0000 0.0044
-16.250 -1.3102 0.05842 0.05549 -0.0162 1.0000 0.0044
-16.000 -1.3365 0.05084 0.04771 -0.0210 1.0000 0.0044
-15.750 -1.3520 0.04579 0.04250 -0.0233 1.0000 0.0044
-15.500 -1.3660 0.04151 0.03806 -0.0246 1.0000 0.0044
-15.250 -1.3755 0.03817 0.03458 -0.0247 1.0000 0.0045
-15.000 -1.3844 0.03523 0.03148 -0.0241 1.0000 0.0046
-14.750 -1.3886 0.03296 0.02909 -0.0229 1.0000 0.0047
-14.500 -1.3917 0.03102 0.02702 -0.0212 1.0000 0.0048
-14.250 -1.3927 0.02945 0.02534 -0.0189 1.0000 0.0048
-14.000 -1.3903 0.02812 0.02391 -0.0163 1.0000 0.0050
-13.750 -1.3810 0.02684 0.02252 -0.0148 1.0000 0.0051
-13.500 -1.3684 0.02569 0.02127 -0.0135 1.0000 0.0053
-13.250 -1.3539 0.02460 0.02007 -0.0123 1.0000 0.0054
-13.000 -1.3376 0.02360 0.01897 -0.0114 1.0000 0.0056
-12.750 -1.3198 0.02268 0.01796 -0.0105 1.0000 0.0058
-12.500 -1.3009 0.02183 0.01703 -0.0097 1.0000 0.0060
-12.250 -1.2811 0.02103 0.01614 -0.0090 1.0000 0.0062
-12.000 -1.2603 0.02030 0.01532 -0.0084 1.0000 0.0064
-11.750 -1.2393 0.01954 0.01448 -0.0078 1.0000 0.0066
-11.500 -1.2189 0.01863 0.01347 -0.0071 1.0000 0.0072
-11.250 -1.1971 0.01787 0.01265 -0.0065 1.0000 0.0076
-11.000 -1.1748 0.01716 0.01189 -0.0060 1.0000 0.0081
-10.750 -1.1518 0.01653 0.01120 -0.0056 1.0000 0.0085
-10.500 -1.1282 0.01596 0.01057 -0.0052 1.0000 0.0090
-10.250 -1.1040 0.01545 0.00999 -0.0048 1.0000 0.0094
-10.000 -1.0798 0.01490 0.00938 -0.0045 1.0000 0.0103
-9.750 -1.0558 0.01431 0.00877 -0.0041 1.0000 0.0118
-9.500 -1.0315 0.01376 0.00821 -0.0038 1.0000 0.0135
-9.250 -1.0065 0.01330 0.00774 -0.0035 1.0000 0.0168
-9.000 -0.9810 0.01290 0.00739 -0.0033 1.0000 0.0215
-8.750 -0.9551 0.01257 0.00703 -0.0031 1.0000 0.0235
-8.500 -0.9286 0.01231 0.00679 -0.0030 1.0000 0.0257
-8.250 -0.9017 0.01213 0.00663 -0.0029 1.0000 0.0279
-8.000 -0.8666 0.01196 0.00641 -0.0046 0.9420 0.0298
-7.750 -0.8409 0.01186 0.00618 -0.0042 0.9161 0.0307
-7.500 -0.8151 0.01174 0.00599 -0.0038 0.9024 0.0312
-7.250 -0.7894 0.01147 0.00568 -0.0034 0.8924 0.0325
-6.750 -0.7360 0.01122 0.00539 -0.0030 0.8767 0.0346
-6.500 -0.7090 0.01112 0.00526 -0.0028 0.8701 0.0356
-6.250 -0.6814 0.01105 0.00518 -0.0027 0.8643 0.0367
-6.000 -0.6544 0.01089 0.00497 -0.0026 0.8583 0.0375
-5.750 -0.6276 0.01066 0.00469 -0.0024 0.8530 0.0379
-5.500 -0.6005 0.01043 0.00443 -0.0022 0.8477 0.0383
-5.250 -0.5735 0.01022 0.00418 -0.0021 0.8424 0.0386
-5.000 -0.5463 0.01002 0.00394 -0.0020 0.8369 0.0389
-4.750 -0.5190 0.00983 0.00372 -0.0018 0.8303 0.0391
-4.500 -0.4918 0.00966 0.00350 -0.0017 0.8242 0.0393
-4.250 -0.4650 0.00928 0.00309 -0.0015 0.8182 0.0407
-4.000 -0.4376 0.00905 0.00283 -0.0014 0.8125 0.0420
-3.750 -0.4100 0.00889 0.00266 -0.0013 0.8074 0.0434
-3.500 -0.3824 0.00875 0.00250 -0.0012 0.7981 0.0446
-3.250 -0.3549 0.00862 0.00232 -0.0010 0.7835 0.0454
-3.000 -0.3273 0.00849 0.00216 -0.0009 0.7701 0.0463
-2.750 -0.2996 0.00838 0.00201 -0.0008 0.7596 0.0472
-2.500 -0.2718 0.00828 0.00187 -0.0007 0.7469 0.0480
-2.250 -0.2440 0.00818 0.00174 -0.0006 0.7318 0.0489
-2.000 -0.2162 0.00811 0.00162 -0.0005 0.7124 0.0496
-1.750 -0.1887 0.00810 0.00150 -0.0004 0.6786 0.0503
-1.500 -0.1613 0.00814 0.00140 -0.0002 0.6371 0.0508
-1.250 -0.1341 0.00823 0.00131 -0.0001 0.5788 0.0520
-1.000 -0.1076 0.00853 0.00125 0.0000 0.4639 0.0543
-0.750 -0.0804 0.00873 0.00124 0.0000 0.3945 0.0565
-0.500 -0.0539 0.00916 0.00125 0.0000 0.2620 0.0588
-0.250 -0.0266 0.00937 0.00127 0.0000 0.1975 0.0634
0.000 0.0000 0.00920 0.00120 0.0000 0.1494 0.1498
0.250 0.0266 0.00937 0.00127 0.0000 0.0634 0.1973
0.500 0.0539 0.00916 0.00125 0.0000 0.0588 0.2620
0.750 0.0804 0.00874 0.00124 0.0000 0.0565 0.3939
1.000 0.1076 0.00853 0.00125 0.0000 0.0543 0.4627
1.250 0.1340 0.00822 0.00131 0.0001 0.0520 0.5798
1.500 0.1612 0.00814 0.00140 0.0002 0.0508 0.6374
1.750 0.1886 0.00809 0.00150 0.0004 0.0503 0.6793
2.000 0.2161 0.00811 0.00162 0.0005 0.0496 0.7123
2.250 0.2439 0.00818 0.00174 0.0006 0.0489 0.7323
2.500 0.2718 0.00827 0.00187 0.0007 0.0480 0.7472
2.750 0.2995 0.00838 0.00201 0.0008 0.0472 0.7596
3.000 0.3273 0.00849 0.00216 0.0009 0.0463 0.7700
3.250 0.3549 0.00862 0.00232 0.0010 0.0455 0.7832
3.500 0.3823 0.00874 0.00250 0.0012 0.0446 0.7982
3.750 0.4100 0.00889 0.00266 0.0013 0.0434 0.8073
4.000 0.4376 0.00905 0.00283 0.0014 0.0420 0.8124
4.250 0.4649 0.00928 0.00309 0.0015 0.0407 0.8181
4.500 0.4917 0.00967 0.00351 0.0017 0.0393 0.8242
4.750 0.5189 0.00983 0.00372 0.0018 0.0391 0.8304
5.000 0.5462 0.01002 0.00394 0.0020 0.0389 0.8369
5.250 0.5734 0.01022 0.00418 0.0021 0.0386 0.8424
5.500 0.6004 0.01043 0.00443 0.0023 0.0383 0.8476
5.750 0.6275 0.01066 0.00469 0.0024 0.0379 0.8529
6.000 0.6542 0.01090 0.00497 0.0026 0.0375 0.8583
6.250 0.6813 0.01105 0.00518 0.0028 0.0368 0.8643
6.500 0.7088 0.01112 0.00526 0.0028 0.0356 0.8701
6.750 0.7358 0.01122 0.00539 0.0030 0.0346 0.8766
7.250 0.7893 0.01147 0.00568 0.0034 0.0325 0.8924
7.500 0.8149 0.01174 0.00599 0.0038 0.0312 0.9024
7.750 0.8407 0.01185 0.00618 0.0042 0.0307 0.9164
8.000 0.8663 0.01196 0.00641 0.0047 0.0298 0.9427
8.250 0.9014 0.01213 0.00663 0.0030 0.0279 1.0000
8.500 0.9283 0.01231 0.00679 0.0031 0.0257 1.0000
8.750 0.9548 0.01256 0.00703 0.0032 0.0235 1.0000
9.000 0.9807 0.01290 0.00739 0.0034 0.0215 1.0000
9.250 1.0061 0.01329 0.00774 0.0036 0.0169 1.0000
9.500 1.0311 0.01375 0.00821 0.0039 0.0136 1.0000
9.750 1.0554 0.01430 0.00876 0.0042 0.0119 1.0000
10.000 1.0793 0.01490 0.00938 0.0046 0.0103 1.0000
10.250 1.1035 0.01544 0.00999 0.0049 0.0094 1.0000
10.500 1.1276 0.01595 0.01056 0.0053 0.0090 1.0000
10.750 1.1512 0.01653 0.01120 0.0057 0.0085 1.0000
11.000 1.1742 0.01715 0.01188 0.0062 0.0081 1.0000
11.250 1.1965 0.01786 0.01265 0.0067 0.0076 1.0000
11.500 1.2182 0.01862 0.01346 0.0072 0.0072 1.0000
11.750 1.2387 0.01952 0.01446 0.0079 0.0066 1.0000
12.000 1.2596 0.02029 0.01531 0.0085 0.0064 1.0000
12.250 1.2802 0.02104 0.01614 0.0092 0.0062 1.0000
12.500 1.3001 0.02182 0.01701 0.0099 0.0060 1.0000
12.750 1.3190 0.02267 0.01795 0.0106 0.0058 1.0000
13.000 1.3367 0.02359 0.01897 0.0115 0.0056 1.0000
13.250 1.3531 0.02457 0.02005 0.0125 0.0055 1.0000
13.500 1.3676 0.02565 0.02123 0.0136 0.0053 1.0000
13.750 1.3799 0.02685 0.02253 0.0150 0.0051 1.0000
14.000 1.3896 0.02807 0.02385 0.0165 0.0050 1.0000
14.250 1.3919 0.02941 0.02530 0.0190 0.0049 1.0000
14.500 1.3911 0.03097 0.02698 0.0213 0.0048 1.0000
14.750 1.3888 0.03285 0.02898 0.0231 0.0047 1.0000
15.000 1.3830 0.03527 0.03153 0.0243 0.0045 1.0000
15.250 1.3728 0.03835 0.03477 0.0248 0.0044 1.0000
15.500 1.3675 0.04126 0.03781 0.0247 0.0045 1.0000
15.750 1.3501 0.04594 0.04266 0.0233 0.0044 1.0000
16.000 1.3345 0.05105 0.04793 0.0209 0.0044 1.0000
16.250 1.3130 0.05786 0.05492 0.0167 0.0044 1.0000
16.500 1.2821 0.06771 0.06497 0.0097 0.0045 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER EA 6(-1)-009 AIRFOIL (ea61009-il)