EPPLER 546 AIRFOIL (e546-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 546 AIRFOIL (e546-il) Reynolds number: 100,000 Max Cl/Cd: 43.3 at α=8.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e546-il-100000-n5.txt Download as CSV file: xf-e546-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 546 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.750 -0.4633 0.10632 0.10111 -0.0491 1.0000 0.0178
-12.500 -0.4710 0.09941 0.09422 -0.0523 1.0000 0.0176
-12.250 -0.4849 0.09144 0.08625 -0.0568 1.0000 0.0173
-12.000 -0.5047 0.08394 0.07870 -0.0611 1.0000 0.0171
-11.750 -0.5273 0.07754 0.07223 -0.0640 1.0000 0.0169
-11.500 -0.5507 0.07210 0.06668 -0.0656 1.0000 0.0167
-11.250 -0.5732 0.06750 0.06195 -0.0659 1.0000 0.0166
-11.000 -0.5938 0.06361 0.05793 -0.0653 1.0000 0.0165
-10.750 -0.6138 0.06017 0.05434 -0.0637 1.0000 0.0164
-10.500 -0.6330 0.05711 0.05112 -0.0613 1.0000 0.0163
-10.250 -0.6513 0.05450 0.04835 -0.0581 1.0000 0.0162
-10.000 -0.6701 0.05229 0.04599 -0.0540 1.0000 0.0162
-9.750 -0.6896 0.05059 0.04416 -0.0488 1.0000 0.0162
-9.500 -0.7073 0.04898 0.04242 -0.0434 1.0000 0.0161
-9.250 -0.7228 0.04740 0.04069 -0.0381 1.0000 0.0161
-9.000 -0.7119 0.04407 0.03694 -0.0380 0.9936 0.0162
-8.750 -0.6903 0.04070 0.03311 -0.0390 0.9848 0.0164
-8.500 -0.6654 0.03775 0.02973 -0.0399 0.9765 0.0168
-8.250 -0.6344 0.03513 0.02658 -0.0411 0.9703 0.0178
-8.000 -0.6065 0.03356 0.02497 -0.0424 0.9614 0.0191
-7.750 -0.5723 0.03181 0.02299 -0.0440 0.9556 0.0208
-7.500 -0.5386 0.03006 0.02107 -0.0452 0.9489 0.0222
-7.250 -0.5029 0.02856 0.01952 -0.0471 0.9439 0.0235
-7.000 -0.4730 0.02722 0.01807 -0.0479 0.9355 0.0256
-6.750 -0.4402 0.02595 0.01679 -0.0497 0.9288 0.0296
-6.500 -0.4146 0.02484 0.01563 -0.0502 0.9184 0.0342
-6.250 -0.3852 0.02367 0.01441 -0.0513 0.9105 0.0427
-6.000 -0.3624 0.02266 0.01345 -0.0511 0.8993 0.0556
-5.750 -0.3409 0.02160 0.01258 -0.0507 0.8887 0.0876
-5.500 -0.3218 0.02029 0.01176 -0.0502 0.8792 0.1678
-5.250 -0.3216 0.01874 0.01093 -0.0469 0.8654 0.2973
-5.000 -0.3029 0.01970 0.01351 -0.0416 0.8564 0.5872
-4.750 -0.2845 0.01977 0.01330 -0.0402 0.8465 0.6375
-4.500 -0.2653 0.02028 0.01360 -0.0381 0.8350 0.6637
-4.250 -0.2313 0.02138 0.01449 -0.0373 0.8274 0.6822
-4.000 -0.1965 0.02276 0.01572 -0.0360 0.8187 0.7008
-3.750 -0.1556 0.02407 0.01684 -0.0357 0.8118 0.7174
-3.500 -0.1345 0.02410 0.01672 -0.0344 0.8018 0.7278
-3.250 -0.1036 0.02407 0.01650 -0.0347 0.7936 0.7310
-3.000 -0.0796 0.02387 0.01613 -0.0342 0.7846 0.7356
-2.750 -0.0665 0.02338 0.01549 -0.0326 0.7753 0.7424
-2.500 -0.0370 0.02326 0.01522 -0.0328 0.7674 0.7444
-2.250 -0.0118 0.02313 0.01497 -0.0325 0.7588 0.7471
-2.000 0.0128 0.02291 0.01461 -0.0322 0.7510 0.7502
-1.750 0.0301 0.02258 0.01416 -0.0310 0.7425 0.7549
-1.500 0.0510 0.02223 0.01368 -0.0305 0.7349 0.7586
-1.250 0.0760 0.02212 0.01349 -0.0301 0.7270 0.7603
-1.000 0.1016 0.02197 0.01325 -0.0299 0.7196 0.7622
-0.750 0.1255 0.02182 0.01301 -0.0295 0.7123 0.7645
-0.500 0.1483 0.02164 0.01275 -0.0291 0.7047 0.7672
-0.250 0.1718 0.02141 0.01242 -0.0289 0.6981 0.7698
0.000 0.1915 0.02115 0.01209 -0.0283 0.6903 0.7734
0.250 0.2192 0.02102 0.01188 -0.0285 0.6846 0.7749
0.500 0.2414 0.02096 0.01182 -0.0278 0.6766 0.7764
0.750 0.2678 0.02085 0.01165 -0.0278 0.6704 0.7778
1.000 0.2917 0.02080 0.01158 -0.0275 0.6636 0.7797
1.250 0.3161 0.02073 0.01149 -0.0272 0.6568 0.7819
1.500 0.3430 0.02063 0.01133 -0.0275 0.6513 0.7839
1.750 0.3650 0.02058 0.01130 -0.0271 0.6438 0.7860
2.000 0.3920 0.02048 0.01115 -0.0275 0.6380 0.7880
2.250 0.4160 0.02046 0.01114 -0.0273 0.6314 0.7901
2.500 0.4407 0.02047 0.01119 -0.0270 0.6249 0.7916
2.750 0.4687 0.02046 0.01114 -0.0272 0.6198 0.7932
3.000 0.4902 0.02053 0.01129 -0.0265 0.6122 0.7949
3.250 0.5171 0.02052 0.01129 -0.0267 0.6065 0.7965
3.500 0.5416 0.02057 0.01139 -0.0265 0.6000 0.7983
3.750 0.5663 0.02061 0.01148 -0.0264 0.5932 0.8003
4.000 0.5951 0.02061 0.01146 -0.0270 0.5878 0.8023
4.250 0.6170 0.02073 0.01169 -0.0266 0.5798 0.8050
4.500 0.6441 0.02074 0.01172 -0.0266 0.5736 0.8066
4.750 0.6657 0.02086 0.01196 -0.0258 0.5657 0.8084
5.000 0.6916 0.02089 0.01204 -0.0257 0.5586 0.8101
5.250 0.7141 0.02100 0.01225 -0.0251 0.5505 0.8120
5.500 0.7397 0.02104 0.01235 -0.0249 0.5428 0.8140
5.750 0.7623 0.02116 0.01260 -0.0244 0.5344 0.8163
6.000 0.7885 0.02120 0.01268 -0.0245 0.5262 0.8186
6.250 0.8099 0.02135 0.01296 -0.0238 0.5165 0.8213
6.500 0.8343 0.02137 0.01305 -0.0232 0.5078 0.8234
6.750 0.8530 0.02151 0.01334 -0.0219 0.4970 0.8259
7.000 0.8734 0.02163 0.01357 -0.0208 0.4862 0.8284
7.250 0.8960 0.02169 0.01373 -0.0201 0.4754 0.8310
7.500 0.9159 0.02183 0.01398 -0.0191 0.4631 0.8339
7.750 0.9344 0.02202 0.01429 -0.0179 0.4498 0.8369
8.000 0.9506 0.02219 0.01459 -0.0162 0.4360 0.8394
8.250 0.9658 0.02238 0.01490 -0.0143 0.4208 0.8425
8.500 0.9795 0.02262 0.01523 -0.0122 0.4043 0.8463
8.750 0.9910 0.02290 0.01556 -0.0099 0.3865 0.8506
9.000 0.9993 0.02323 0.01595 -0.0070 0.3678 0.8547
9.250 1.0058 0.02375 0.01654 -0.0040 0.3468 0.8593
9.500 1.0122 0.02440 0.01717 -0.0014 0.3242 0.8643
9.750 1.0167 0.02521 0.01798 0.0013 0.3006 0.8694
10.000 1.0185 0.02618 0.01892 0.0042 0.2782 0.8751
10.250 1.0204 0.02735 0.02007 0.0066 0.2559 0.8819
10.500 1.0200 0.02865 0.02137 0.0091 0.2354 0.8892
11.000 1.0191 0.03171 0.02446 0.0131 0.1976 0.9106
11.250 1.0199 0.03346 0.02626 0.0144 0.1795 0.9276
11.500 1.0217 0.03534 0.02818 0.0147 0.1624 0.9810
11.750 1.0231 0.03754 0.03036 0.0150 0.1468 1.0000
12.000 1.0247 0.03987 0.03267 0.0151 0.1326 1.0000
12.250 1.0261 0.04227 0.03506 0.0151 0.1197 1.0000
12.500 1.0274 0.04475 0.03756 0.0151 0.1081 1.0000
12.750 1.0279 0.04735 0.04016 0.0150 0.0980 1.0000
13.000 1.0268 0.05016 0.04294 0.0149 0.0896 1.0000
13.250 1.0286 0.05278 0.04565 0.0147 0.0813 1.0000
13.500 1.0286 0.05565 0.04854 0.0143 0.0743 1.0000
13.750 1.0275 0.05872 0.05164 0.0139 0.0683 1.0000
14.000 1.0284 0.06163 0.05463 0.0134 0.0629 1.0000
14.250 1.0271 0.06486 0.05791 0.0128 0.0583 1.0000
14.500 1.0272 0.06803 0.06116 0.0121 0.0538 1.0000
14.750 1.0269 0.07132 0.06453 0.0113 0.0500 1.0000
15.000 1.0252 0.07482 0.06806 0.0104 0.0469 1.0000
15.250 1.0258 0.07815 0.07154 0.0095 0.0434 1.0000
15.500 1.0240 0.08185 0.07528 0.0082 0.0409 1.0000
15.750 1.0235 0.08543 0.07898 0.0071 0.0383 1.0000
16.000 1.0221 0.08930 0.08297 0.0057 0.0357 1.0000
16.250 1.0197 0.09329 0.08700 0.0041 0.0339 1.0000
16.500 1.0182 0.09727 0.09113 0.0026 0.0319 1.0000
16.750 1.0154 0.10160 0.09561 0.0007 0.0300 1.0000
17.000 1.0123 0.10595 0.10004 -0.0014 0.0285 1.0000
17.250 1.0093 0.11036 0.10453 -0.0034 0.0271 1.0000
17.500 1.0043 0.11538 0.10978 -0.0058 0.0257 1.0000
17.750 0.9996 0.12037 0.11490 -0.0084 0.0245 1.0000
18.000 0.9956 0.12522 0.11981 -0.0112 0.0234 1.0000
18.250 0.9928 0.12985 0.12451 -0.0137 0.0226 1.0000
18.500 0.9830 0.13630 0.13120 -0.0173 0.0218 1.0000
18.750 0.9735 0.14284 0.13794 -0.0211 0.0213 1.0000
19.000 0.9616 0.15017 0.14544 -0.0255 0.0207 1.0000
19.250 0.9479 0.15822 0.15366 -0.0304 0.0205 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 546 AIRFOIL (e546-il)