EPPLER 625 AIRFOIL (e625-il) Xfoil prediction polar at RE=50,000 Ncrit=5
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Airfoil: EPPLER 625 AIRFOIL (e625-il) Reynolds number: 50,000 Max Cl/Cd: 28.82 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e625-il-50000-n5.txt Download as CSV file: xf-e625-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 625 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.4261 0.10850 0.10157 -0.0192 1.0000 0.0698
-10.000 -0.4179 0.10483 0.09792 -0.0193 1.0000 0.0685
-9.500 -0.4497 0.09153 0.08473 -0.0306 1.0000 0.0593
-9.250 -0.4478 0.08790 0.08115 -0.0309 1.0000 0.0589
-9.000 -0.4498 0.08423 0.07752 -0.0313 1.0000 0.0583
-8.750 -0.4560 0.08059 0.07392 -0.0314 1.0000 0.0576
-8.500 -0.4653 0.07719 0.07054 -0.0306 1.0000 0.0569
-8.250 -0.4763 0.07389 0.06724 -0.0291 1.0000 0.0561
-8.000 -0.4860 0.07043 0.06374 -0.0276 1.0000 0.0553
-7.750 -0.4943 0.06709 0.06031 -0.0256 1.0000 0.0545
-7.500 -0.5020 0.06377 0.05689 -0.0232 1.0000 0.0537
-7.250 -0.5091 0.06058 0.05357 -0.0202 1.0000 0.0530
-7.000 -0.5162 0.05762 0.05045 -0.0167 1.0000 0.0524
-6.750 -0.5243 0.05493 0.04757 -0.0126 1.0000 0.0518
-6.500 -0.5328 0.05269 0.04515 -0.0081 1.0000 0.0515
-6.250 -0.5411 0.05069 0.04293 -0.0035 0.9994 0.0511
-6.000 -0.5146 0.04681 0.03854 -0.0051 0.9810 0.0510
-5.750 -0.4854 0.04375 0.03509 -0.0066 0.9643 0.0516
-5.500 -0.4528 0.04120 0.03221 -0.0085 0.9492 0.0535
-5.250 -0.4181 0.03856 0.02908 -0.0102 0.9355 0.0557
-5.000 -0.3806 0.03596 0.02597 -0.0119 0.9224 0.0572
-4.750 -0.3411 0.03356 0.02307 -0.0136 0.9088 0.0586
-4.500 -0.3005 0.03155 0.02082 -0.0156 0.8946 0.0608
-4.250 -0.2601 0.03012 0.01927 -0.0177 0.8799 0.0651
-4.000 -0.2127 0.02859 0.01748 -0.0204 0.8657 0.0712
-3.750 -0.1635 0.02723 0.01605 -0.0236 0.8512 0.0789
-3.500 -0.1240 0.02620 0.01495 -0.0252 0.8354 0.0913
-3.250 -0.0924 0.02529 0.01397 -0.0253 0.8190 0.1115
-3.000 -0.0676 0.02430 0.01322 -0.0245 0.8028 0.1505
-2.750 0.0100 0.02235 0.01436 -0.0289 0.7912 0.8443
-2.500 0.1141 0.02391 0.01513 -0.0387 0.7729 0.9520
-2.250 0.2203 0.02286 0.01344 -0.0531 0.7510 1.0000
-2.000 0.2404 0.02282 0.01314 -0.0518 0.7332 1.0000
-1.750 0.2606 0.02279 0.01287 -0.0505 0.7163 1.0000
-1.500 0.2809 0.02279 0.01264 -0.0492 0.7000 1.0000
-1.250 0.3014 0.02279 0.01243 -0.0479 0.6843 1.0000
-1.000 0.3219 0.02281 0.01226 -0.0467 0.6691 1.0000
-0.750 0.3426 0.02285 0.01212 -0.0454 0.6543 1.0000
-0.500 0.3634 0.02289 0.01199 -0.0442 0.6401 1.0000
-0.250 0.3843 0.02294 0.01186 -0.0429 0.6267 1.0000
0.000 0.4052 0.02303 0.01181 -0.0418 0.6127 1.0000
0.250 0.4261 0.02313 0.01180 -0.0406 0.5988 1.0000
0.500 0.4471 0.02326 0.01181 -0.0395 0.5856 1.0000
0.750 0.4683 0.02338 0.01182 -0.0384 0.5732 1.0000
1.000 0.4899 0.02350 0.01178 -0.0372 0.5616 1.0000
1.250 0.5112 0.02368 0.01191 -0.0362 0.5488 1.0000
1.500 0.5324 0.02389 0.01205 -0.0352 0.5366 1.0000
1.750 0.5536 0.02408 0.01213 -0.0340 0.5254 1.0000
2.000 0.5747 0.02426 0.01223 -0.0328 0.5146 1.0000
2.250 0.5953 0.02455 0.01249 -0.0317 0.5028 1.0000
2.500 0.6161 0.02480 0.01267 -0.0306 0.4924 1.0000
2.750 0.6370 0.02503 0.01282 -0.0294 0.4824 1.0000
3.000 0.6570 0.02540 0.01320 -0.0282 0.4715 1.0000
3.250 0.6777 0.02567 0.01340 -0.0270 0.4623 1.0000
3.500 0.6976 0.02603 0.01376 -0.0258 0.4523 1.0000
3.750 0.7173 0.02643 0.01418 -0.0246 0.4429 1.0000
4.000 0.7375 0.02675 0.01444 -0.0233 0.4344 1.0000
4.250 0.7560 0.02726 0.01499 -0.0220 0.4247 1.0000
4.500 0.7766 0.02756 0.01521 -0.0207 0.4173 1.0000
4.750 0.7939 0.02819 0.01596 -0.0193 0.4079 1.0000
5.000 0.8145 0.02848 0.01617 -0.0180 0.4010 1.0000
5.250 0.8304 0.02921 0.01703 -0.0165 0.3917 1.0000
5.500 0.8508 0.02953 0.01730 -0.0152 0.3852 1.0000
5.750 0.8652 0.03036 0.01828 -0.0135 0.3765 1.0000
6.000 0.8850 0.03071 0.01857 -0.0121 0.3699 1.0000
6.250 0.8982 0.03160 0.01961 -0.0103 0.3617 1.0000
6.500 0.9166 0.03207 0.02007 -0.0088 0.3553 1.0000
6.750 0.9296 0.03295 0.02109 -0.0069 0.3480 1.0000
7.000 0.9452 0.03359 0.02178 -0.0052 0.3412 1.0000
7.250 0.9604 0.03429 0.02252 -0.0034 0.3350 1.0000
7.500 0.9703 0.03531 0.02369 -0.0012 0.3279 1.0000
7.750 0.9902 0.03566 0.02401 0.0002 0.3226 1.0000
8.000 0.9932 0.03710 0.02567 0.0030 0.3156 1.0000
8.250 1.0066 0.03783 0.02646 0.0049 0.3097 1.0000
8.500 1.0193 0.03865 0.02732 0.0069 0.3044 1.0000
8.750 1.0180 0.04024 0.02913 0.0101 0.2980 1.0000
9.000 1.0349 0.04074 0.02962 0.0117 0.2928 1.0000
9.250 1.0340 0.04230 0.03133 0.0148 0.2874 1.0000
9.500 1.0272 0.04412 0.03330 0.0183 0.2818 1.0000
9.750 1.0470 0.04445 0.03362 0.0197 0.2770 1.0000
10.000 1.0310 0.04680 0.03614 0.0237 0.2722 1.0000
10.250 0.9936 0.05010 0.03954 0.0291 0.2677 1.0000
10.500 0.9990 0.05119 0.04068 0.0312 0.2632 1.0000
10.750 1.0049 0.05243 0.04195 0.0331 0.2591 1.0000
11.000 0.8468 0.07061 0.06019 0.0315 0.2491 1.0000
11.250 0.8810 0.06867 0.05831 0.0338 0.2465 1.0000
11.750 0.8111 0.08460 0.07426 0.0294 0.2312 1.0000
12.250 0.7627 0.09887 0.08854 0.0249 0.2176 1.0000
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Polar data table (+)
Polar graphs
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