GOE 456 AIRFOIL (goe456-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 456 AIRFOIL (goe456-il) Reynolds number: 50,000 Max Cl/Cd: 42.53 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe456-il-50000-n5.txt Download as CSV file: xf-goe456-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 456 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
------ -------- --------- --------- -------- -------- --------
-7.250 -0.3313 0.10529 0.09940 -0.0287 1.0000 0.0939
-7.000 -0.3465 0.10514 0.09940 -0.0279 1.0000 0.0947
-6.750 -0.3574 0.10465 0.09903 -0.0291 1.0000 0.0952
-6.500 -0.3531 0.10002 0.09447 -0.0255 1.0000 0.0967
-6.250 -0.3518 0.09697 0.09148 -0.0230 1.0000 0.0987
-6.000 -0.3537 0.09479 0.08936 -0.0217 1.0000 0.1009
-5.750 -0.3560 0.09270 0.08734 -0.0213 1.0000 0.1035
-5.500 -0.3486 0.09120 0.08586 -0.0292 0.9975 0.1089
-5.250 -0.3234 0.08608 0.08072 -0.0350 0.9906 0.1104
-4.750 -0.2553 0.07257 0.06686 -0.0506 0.9770 0.0532
-4.500 -0.2274 0.06799 0.06220 -0.0554 0.9698 0.0500
-4.000 -0.1354 0.05504 0.04864 -0.0761 0.9564 0.0417
-3.750 -0.1027 0.05048 0.04386 -0.0807 0.9490 0.0410
-3.500 -0.0615 0.04549 0.03851 -0.0867 0.9435 0.0398
-3.250 -0.0216 0.04078 0.03328 -0.0918 0.9371 0.0389
-3.000 0.0208 0.03643 0.02822 -0.0963 0.9316 0.0387
-2.750 0.0636 0.03293 0.02383 -0.0998 0.9269 0.0407
-2.500 0.0975 0.03054 0.02086 -0.1014 0.9197 0.0450
-2.250 0.1380 0.02854 0.01830 -0.1037 0.9151 0.0491
-2.000 0.1684 0.02722 0.01649 -0.1038 0.9069 0.0556
-1.750 0.2097 0.02596 0.01502 -0.1062 0.9020 0.0743
-1.500 0.2424 0.02522 0.01442 -0.1070 0.8938 0.1436
-1.250 0.2789 0.02524 0.01455 -0.1090 0.8876 0.2843
-1.000 0.3078 0.02499 0.01407 -0.1093 0.8784 0.3167
-0.750 0.3475 0.02455 0.01339 -0.1114 0.8734 0.3466
-0.500 0.3746 0.02432 0.01313 -0.1115 0.8640 0.3822
-0.250 0.4089 0.02361 0.01279 -0.1128 0.8582 0.4759
0.000 0.4384 0.02267 0.01248 -0.1128 0.8495 1.0000
0.250 0.4679 0.02293 0.01237 -0.1132 0.8412 1.0000
0.500 0.5007 0.02310 0.01225 -0.1140 0.8342 1.0000
0.750 0.5263 0.02343 0.01237 -0.1138 0.8248 1.0000
1.000 0.5609 0.02355 0.01228 -0.1150 0.8191 1.0000
1.250 0.5837 0.02397 0.01257 -0.1143 0.8090 1.0000
1.500 0.6141 0.02419 0.01267 -0.1148 0.8022 1.0000
2.000 0.6657 0.02490 0.01327 -0.1144 0.7854 1.0000
2.250 0.6951 0.02514 0.01348 -0.1147 0.7784 1.0000
2.500 0.7179 0.02562 0.01400 -0.1141 0.7694 1.0000
2.750 0.7492 0.02578 0.01418 -0.1146 0.7632 1.0000
3.000 0.7700 0.02634 0.01479 -0.1137 0.7535 1.0000
3.250 0.8034 0.02640 0.01493 -0.1144 0.7480 1.0000
3.750 0.8485 0.02718 0.01598 -0.1126 0.7268 1.0000
4.000 0.8781 0.02693 0.01586 -0.1119 0.7138 1.0000
4.250 0.9072 0.02658 0.01572 -0.1110 0.6987 1.0000
4.500 0.9354 0.02631 0.01563 -0.1101 0.6836 1.0000
4.750 0.9625 0.02628 0.01582 -0.1092 0.6705 1.0000
5.000 0.9895 0.02632 0.01612 -0.1084 0.6580 1.0000
5.250 1.0163 0.02638 0.01654 -0.1076 0.6450 1.0000
5.500 1.0410 0.02572 0.01608 -0.1050 0.6146 1.0000
5.750 1.0564 0.02494 0.01522 -0.1004 0.5552 1.0000
6.000 1.0645 0.02503 0.01517 -0.0957 0.4820 1.0000
6.250 1.0691 0.02583 0.01523 -0.0909 0.3495 1.0000
6.500 1.0636 0.02814 0.01648 -0.0862 0.2333 1.0000
7.000 1.0523 0.03366 0.02057 -0.0785 0.0449 1.0000
7.250 1.0600 0.03547 0.02254 -0.0760 0.0390 1.0000
7.500 1.0664 0.03744 0.02467 -0.0735 0.0352 1.0000
7.750 1.0733 0.03937 0.02685 -0.0713 0.0318 1.0000
8.000 1.0774 0.04156 0.02924 -0.0690 0.0298 1.0000
8.250 1.0793 0.04401 0.03189 -0.0668 0.0288 1.0000
8.500 1.0828 0.04634 0.03443 -0.0647 0.0282 1.0000
8.750 1.0880 0.04860 0.03693 -0.0628 0.0276 1.0000
9.000 1.0957 0.05076 0.03932 -0.0610 0.0270 1.0000
9.250 1.1084 0.05276 0.04157 -0.0591 0.0257 1.0000
9.500 1.1299 0.05455 0.04359 -0.0574 0.0240 1.0000
9.750 1.1684 0.05652 0.04590 -0.0562 0.0229 1.0000
10.000 1.2073 0.05942 0.04917 -0.0557 0.0228 1.0000
10.250 1.2296 0.06281 0.05298 -0.0547 0.0230 1.0000
10.500 1.2396 0.06632 0.05691 -0.0531 0.0232 1.0000
10.750 1.2415 0.06990 0.06088 -0.0513 0.0235 1.0000
11.000 1.2380 0.07358 0.06492 -0.0495 0.0238 1.0000
11.250 1.2305 0.07747 0.06916 -0.0481 0.0240 1.0000
11.500 1.2200 0.08157 0.07357 -0.0470 0.0242 1.0000
11.750 1.2073 0.08595 0.07825 -0.0466 0.0244 1.0000
12.000 1.1933 0.09064 0.08321 -0.0468 0.0246 1.0000
12.250 1.1777 0.09573 0.08854 -0.0477 0.0248 1.0000
12.500 1.1611 0.10122 0.09426 -0.0494 0.0249 1.0000
12.750 1.1440 0.10718 0.10043 -0.0519 0.0251 1.0000
13.000 1.1269 0.11360 0.10703 -0.0551 0.0254 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 456 AIRFOIL (goe456-il)