GOE 474 AIRFOIL (goe474-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 474 AIRFOIL (goe474-il) Reynolds number: 50,000 Max Cl/Cd: 38.3 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe474-il-50000-n5.txt Download as CSV file: xf-goe474-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 474 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4245 0.09969 0.09253 -0.0270 1.0000 0.0563
-8.500 -0.4268 0.09595 0.08888 -0.0284 1.0000 0.0545
-8.000 -0.4447 0.08800 0.08121 -0.0335 1.0000 0.0503
-7.750 -0.4439 0.08444 0.07771 -0.0342 1.0000 0.0499
-7.500 -0.4437 0.08071 0.07404 -0.0353 1.0000 0.0496
-7.250 -0.4433 0.07685 0.07023 -0.0366 1.0000 0.0494
-6.750 -0.4411 0.06809 0.06144 -0.0402 1.0000 0.0498
-6.500 -0.4385 0.06310 0.05637 -0.0421 1.0000 0.0502
-6.250 -0.4342 0.05786 0.05096 -0.0437 1.0000 0.0507
-6.000 -0.4264 0.05591 0.04905 -0.0422 1.0000 0.0525
-5.750 -0.4181 0.05191 0.04488 -0.0424 1.0000 0.0531
-5.500 -0.4079 0.04771 0.04043 -0.0426 1.0000 0.0539
-5.250 -0.3955 0.04393 0.03634 -0.0425 1.0000 0.0554
-5.000 -0.3808 0.04020 0.03219 -0.0424 1.0000 0.0591
-4.750 -0.3632 0.03514 0.02627 -0.0426 1.0000 0.0628
-4.500 -0.3455 0.03322 0.02417 -0.0416 1.0000 0.0657
-4.250 -0.3260 0.03142 0.02202 -0.0406 1.0000 0.0718
-4.000 -0.3056 0.02978 0.02003 -0.0398 1.0000 0.0792
-3.750 -0.2836 0.02794 0.01772 -0.0389 1.0000 0.0859
-3.500 -0.2620 0.02678 0.01632 -0.0379 1.0000 0.0942
-3.250 -0.2408 0.02606 0.01538 -0.0370 1.0000 0.1080
-3.000 -0.2177 0.02497 0.01403 -0.0361 1.0000 0.1218
-2.750 -0.1943 0.02416 0.01296 -0.0354 1.0000 0.1401
-2.500 -0.1726 0.02374 0.01227 -0.0345 1.0000 0.1613
-2.250 -0.1532 0.02368 0.01219 -0.0336 1.0000 0.1880
-2.000 -0.1319 0.02338 0.01182 -0.0328 1.0000 0.2083
-1.750 -0.0969 0.02302 0.01123 -0.0345 0.9952 0.2235
-1.500 -0.0622 0.02271 0.01075 -0.0361 0.9898 0.2363
-1.250 -0.0278 0.02248 0.01043 -0.0377 0.9845 0.2507
-1.000 0.0057 0.02227 0.01014 -0.0392 0.9788 0.2642
-0.750 0.0383 0.02206 0.00996 -0.0405 0.9729 0.2819
-0.500 0.0713 0.02183 0.00988 -0.0420 0.9671 0.3117
-0.250 0.1021 0.02139 0.00983 -0.0432 0.9611 0.3780
0.000 0.1467 0.01989 0.00979 -0.0462 0.9573 1.0000
0.250 0.1817 0.02027 0.00988 -0.0480 0.9512 1.0000
0.500 0.2114 0.02059 0.01000 -0.0487 0.9430 1.0000
0.750 0.2463 0.02096 0.01021 -0.0505 0.9368 1.0000
1.000 0.2756 0.02129 0.01044 -0.0512 0.9282 1.0000
1.250 0.3096 0.02160 0.01067 -0.0526 0.9190 1.0000
1.500 0.3508 0.02181 0.01084 -0.0552 0.9086 1.0000
1.750 0.3892 0.02194 0.01096 -0.0571 0.8960 1.0000
2.000 0.4237 0.02208 0.01112 -0.0582 0.8833 1.0000
2.250 0.4573 0.02223 0.01134 -0.0592 0.8713 1.0000
2.500 0.4916 0.02234 0.01152 -0.0602 0.8593 1.0000
2.750 0.5257 0.02240 0.01167 -0.0609 0.8464 1.0000
3.000 0.5605 0.02229 0.01170 -0.0614 0.8303 1.0000
3.250 0.6005 0.02182 0.01136 -0.0621 0.8104 1.0000
3.500 0.6295 0.02152 0.01118 -0.0610 0.7861 1.0000
3.750 0.6604 0.02118 0.01100 -0.0601 0.7625 1.0000
4.000 0.6888 0.02093 0.01089 -0.0589 0.7375 1.0000
4.250 0.7146 0.02071 0.01081 -0.0572 0.7070 1.0000
4.500 0.7381 0.02056 0.01080 -0.0551 0.6673 1.0000
4.750 0.7616 0.02044 0.01070 -0.0529 0.6151 1.0000
5.000 0.7836 0.02051 0.01065 -0.0505 0.5393 1.0000
5.250 0.8040 0.02099 0.01059 -0.0479 0.4455 1.0000
5.500 0.8197 0.02203 0.01120 -0.0455 0.3769 1.0000
5.750 0.8358 0.02315 0.01206 -0.0435 0.3266 1.0000
6.000 0.8532 0.02425 0.01302 -0.0418 0.2886 1.0000
6.250 0.8716 0.02536 0.01401 -0.0403 0.2584 1.0000
6.500 0.8915 0.02644 0.01507 -0.0391 0.2335 1.0000
6.750 0.9132 0.02758 0.01625 -0.0381 0.2150 1.0000
7.000 0.9360 0.02872 0.01746 -0.0373 0.1966 1.0000
7.250 0.9580 0.02990 0.01874 -0.0364 0.1784 1.0000
7.500 0.9778 0.03109 0.02001 -0.0353 0.1602 1.0000
7.750 0.9971 0.03238 0.02137 -0.0342 0.1431 1.0000
8.000 1.0141 0.03370 0.02290 -0.0327 0.1240 1.0000
8.250 1.0270 0.03495 0.02434 -0.0308 0.1046 1.0000
8.500 1.0389 0.03639 0.02577 -0.0290 0.0880 1.0000
8.750 1.0521 0.03820 0.02764 -0.0273 0.0737 1.0000
9.000 1.0679 0.04038 0.03013 -0.0257 0.0630 1.0000
9.250 1.0819 0.04278 0.03284 -0.0241 0.0550 1.0000
9.500 1.0923 0.04478 0.03492 -0.0223 0.0501 1.0000
9.750 1.1036 0.04799 0.03849 -0.0206 0.0468 1.0000
10.000 1.1097 0.05113 0.04206 -0.0186 0.0438 1.0000
10.250 1.1122 0.05361 0.04476 -0.0165 0.0411 1.0000
10.500 1.1128 0.05597 0.04719 -0.0144 0.0392 1.0000
10.750 1.1073 0.05943 0.05092 -0.0121 0.0384 1.0000
11.000 1.0952 0.06338 0.05525 -0.0100 0.0379 1.0000
11.250 1.0800 0.06759 0.05978 -0.0086 0.0377 1.0000
11.500 1.0617 0.07232 0.06479 -0.0083 0.0376 1.0000
11.750 1.0419 0.07758 0.07027 -0.0092 0.0377 1.0000
12.000 1.0212 0.08350 0.07638 -0.0114 0.0379 1.0000
12.250 1.0001 0.09030 0.08333 -0.0149 0.0381 1.0000
12.500 0.9790 0.09807 0.09120 -0.0196 0.0384 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 474 AIRFOIL (goe474-il)