EPPLER 502 AIRFOIL (e502-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 502 AIRFOIL (e502-il) Reynolds number: 200,000 Max Cl/Cd: 58.79 at α=8° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e502-il-200000.txt Download as CSV file: xf-e502-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 502 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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.3284 0.08980 0.08688 -0.0645 0.9490 0.0504
-11.250 -0.3381 0.08197 0.07905 -0.0697 0.9410 0.0511
-11.000 -0.3617 0.07186 0.06889 -0.0770 0.9338 0.0505
-10.750 -0.3899 0.06384 0.06074 -0.0819 0.9254 0.0503
-10.500 -0.4187 0.05777 0.05450 -0.0839 0.9151 0.0497
-10.250 -0.4430 0.05335 0.04991 -0.0839 0.9050 0.0496
-10.000 -0.4651 0.04980 0.04618 -0.0824 0.8959 0.0506
-9.750 -0.5839 0.04652 0.04166 -0.0744 0.9011 0.0229
-9.500 -0.5942 0.04352 0.03836 -0.0703 0.8909 0.0223
-9.250 -0.5993 0.04034 0.03481 -0.0665 0.8832 0.0215
-9.000 -0.6001 0.03705 0.03109 -0.0629 0.8757 0.0212
-8.750 -0.5931 0.03407 0.02765 -0.0601 0.8704 0.0208
-8.500 -0.5793 0.03125 0.02440 -0.0582 0.8654 0.0207
-8.250 -0.5637 0.03020 0.02308 -0.0564 0.8601 0.0229
-8.000 -0.5450 0.02887 0.02139 -0.0549 0.8558 0.0240
-7.750 -0.5123 0.02563 0.01792 -0.0562 0.8525 0.0264
-7.500 -0.4845 0.02427 0.01647 -0.0564 0.8484 0.0293
-7.250 -0.4503 0.02261 0.01465 -0.0577 0.8453 0.0352
-7.000 -0.4269 0.02178 0.01376 -0.0572 0.8421 0.0423
-6.750 -0.4017 0.02064 0.01263 -0.0572 0.8389 0.0504
-6.500 -0.3810 0.01995 0.01187 -0.0562 0.8348 0.0581
-6.250 -0.3645 0.01925 0.01121 -0.0547 0.8309 0.0662
-6.000 -0.3513 0.01858 0.01052 -0.0524 0.8277 0.0747
-5.750 -0.3392 0.01808 0.01001 -0.0499 0.8245 0.0849
-5.500 -0.3308 0.01747 0.00952 -0.0469 0.8205 0.1033
-5.250 -0.3281 0.01655 0.00897 -0.0430 0.8167 0.1591
-5.000 -0.3389 0.01520 0.00842 -0.0369 0.8131 0.3154
-4.750 -0.3548 0.01389 0.00817 -0.0294 0.8091 0.5104
-4.500 -0.3407 0.01458 0.00953 -0.0252 0.8059 0.6902
-4.250 -0.3227 0.01525 0.01006 -0.0227 0.8029 0.7255
-4.000 -0.2906 0.01632 0.01106 -0.0220 0.8005 0.7464
-3.750 -0.2582 0.01732 0.01194 -0.0213 0.7985 0.7631
-3.500 -0.2218 0.01846 0.01301 -0.0214 0.7962 0.7764
-3.250 -0.1732 0.01999 0.01447 -0.0230 0.7941 0.7899
-3.000 -0.1018 0.02173 0.01608 -0.0283 0.7929 0.8024
-2.750 -0.0787 0.02200 0.01627 -0.0272 0.7899 0.8131
-2.500 -0.0052 0.02261 0.01672 -0.0348 0.7885 0.8163
-2.250 0.0085 0.02269 0.01674 -0.0323 0.7859 0.8269
-2.000 0.0632 0.02284 0.01677 -0.0371 0.7845 0.8294
-1.750 0.0958 0.02288 0.01675 -0.0384 0.7821 0.8331
-1.500 0.0888 0.02278 0.01668 -0.0329 0.7775 0.8413
-1.250 0.1266 0.02273 0.01658 -0.0352 0.7750 0.8430
-1.000 0.1588 0.02267 0.01646 -0.0365 0.7728 0.8455
-0.750 0.1815 0.02255 0.01629 -0.0359 0.7704 0.8489
-0.500 0.1641 0.02251 0.01629 -0.0285 0.7659 0.8557
-0.250 0.1971 0.02254 0.01631 -0.0302 0.7626 0.8572
0.000 0.2263 0.02249 0.01624 -0.0309 0.7597 0.8590
0.250 0.2504 0.02238 0.01610 -0.0306 0.7572 0.8615
0.500 0.2088 0.02227 0.01602 -0.0187 0.7518 0.8692
0.750 0.2337 0.02237 0.01614 -0.0189 0.7476 0.8707
1.000 0.2620 0.02230 0.01606 -0.0193 0.7446 0.8722
1.250 0.2888 0.02213 0.01586 -0.0194 0.7423 0.8736
1.500 0.2968 0.02227 0.01604 -0.0166 0.7373 0.8767
1.750 0.2800 0.02213 0.01592 -0.0093 0.7311 0.8819
2.000 0.3026 0.02186 0.01563 -0.0086 0.7285 0.8839
2.250 0.3357 0.02163 0.01538 -0.0096 0.7266 0.8847
2.500 0.3378 0.02209 0.01594 -0.0062 0.7181 0.8872
2.750 0.3677 0.02185 0.01571 -0.0066 0.7151 0.8887
3.000 0.4012 0.02148 0.01533 -0.0076 0.7129 0.8900
3.250 0.3953 0.02179 0.01572 -0.0026 0.7040 0.8938
3.500 0.4231 0.02138 0.01532 -0.0027 0.7008 0.8956
3.750 0.4551 0.02090 0.01484 -0.0033 0.6985 0.8974
4.000 0.4645 0.02120 0.01525 -0.0008 0.6896 0.8997
4.250 0.5013 0.02070 0.01477 -0.0022 0.6863 0.9006
4.500 0.5411 0.02012 0.01422 -0.0039 0.6837 0.9013
4.750 0.5513 0.02024 0.01446 -0.0014 0.6745 0.9041
5.000 0.5892 0.01963 0.01387 -0.0029 0.6710 0.9052
5.250 0.6062 0.01950 0.01383 -0.0012 0.6631 0.9083
5.500 0.6399 0.01883 0.01321 -0.0020 0.6577 0.9103
5.750 0.6642 0.01845 0.01292 -0.0013 0.6499 0.9125
6.000 0.6998 0.01777 0.01230 -0.0023 0.6430 0.9137
6.250 0.7234 0.01741 0.01208 -0.0015 0.6333 0.9160
6.500 0.7570 0.01674 0.01147 -0.0022 0.6250 0.9177
6.750 0.7827 0.01620 0.01101 -0.0016 0.6136 0.9203
7.000 0.8020 0.01573 0.01064 0.0001 0.5999 0.9240
7.250 0.8205 0.01533 0.01035 0.0019 0.5825 0.9275
7.500 0.8399 0.01504 0.01018 0.0036 0.5598 0.9312
7.750 0.8559 0.01482 0.00998 0.0058 0.5313 0.9361
8.000 0.8678 0.01476 0.00982 0.0086 0.4906 0.9414
8.250 0.8757 0.01510 0.00991 0.0119 0.4367 0.9473
8.500 0.8759 0.01580 0.01037 0.0159 0.3886 0.9562
8.750 0.8812 0.01698 0.01132 0.0180 0.3348 0.9651
9.000 0.8896 0.01859 0.01268 0.0183 0.2787 0.9742
9.250 0.9002 0.02051 0.01433 0.0175 0.2227 0.9843
9.500 0.9129 0.02260 0.01617 0.0161 0.1713 0.9963
9.750 0.9001 0.02349 0.01692 0.0204 0.1499 1.0000
10.000 0.8966 0.02494 0.01818 0.0230 0.1234 1.0000
10.250 0.8965 0.02656 0.01962 0.0250 0.0995 1.0000
10.500 0.8960 0.02835 0.02124 0.0269 0.0809 1.0000
10.750 0.8981 0.03006 0.02285 0.0286 0.0664 1.0000
11.000 0.8995 0.03189 0.02462 0.0302 0.0547 1.0000
11.250 0.9046 0.03350 0.02623 0.0315 0.0460 1.0000
11.500 0.9086 0.03527 0.02804 0.0330 0.0388 1.0000
11.750 0.9078 0.03749 0.03019 0.0347 0.0331 1.0000
12.000 0.9172 0.03892 0.03174 0.0355 0.0292 1.0000
12.250 0.9209 0.04088 0.03367 0.0366 0.0256 1.0000
12.500 0.9293 0.04273 0.03564 0.0380 0.0231 1.0000
12.750 0.9399 0.04438 0.03740 0.0390 0.0212 1.0000
13.000 0.9493 0.04610 0.03918 0.0397 0.0197 1.0000
13.500 0.9699 0.05049 0.04377 0.0415 0.0171 1.0000
13.750 0.9736 0.05265 0.04613 0.0421 0.0160 1.0000
14.000 0.9761 0.05509 0.04874 0.0426 0.0147 1.0000
14.250 0.9798 0.05782 0.05167 0.0431 0.0144 1.0000
14.500 0.9808 0.06104 0.05515 0.0436 0.0149 1.0000
14.750 0.9764 0.06450 0.05882 0.0438 0.0143 1.0000
15.000 0.9709 0.06809 0.06258 0.0436 0.0139 1.0000
15.250 0.9592 0.07277 0.06754 0.0432 0.0141 1.0000
15.500 0.9476 0.07741 0.07239 0.0423 0.0140 1.0000
15.750 0.9283 0.08347 0.07873 0.0407 0.0143 1.0000
16.000 0.9119 0.08930 0.08478 0.0385 0.0143 1.0000
16.250 0.8876 0.09695 0.09269 0.0352 0.0146 1.0000
16.500 0.8653 0.10476 0.10071 0.0312 0.0146 1.0000
16.750 0.8405 0.11377 0.10993 0.0262 0.0150 1.0000
17.000 0.8124 0.12441 0.12075 0.0199 0.0152 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 502 AIRFOIL (e502-il)