GOE 566 AIRFOIL (goe566-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 566 AIRFOIL (goe566-il) Reynolds number: 500,000 Max Cl/Cd: 79.33 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe566-il-500000-n5.txt Download as CSV file: xf-goe566-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 566 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.8313 0.02929 0.02591 -0.0558 1.0000 0.0106
-10.250 -0.8310 0.02586 0.02208 -0.0532 1.0000 0.0110
-10.000 -0.8148 0.02472 0.02082 -0.0520 0.9995 0.0114
-9.750 -0.7846 0.02406 0.02011 -0.0532 0.9967 0.0120
-9.500 -0.7559 0.02306 0.01897 -0.0543 0.9933 0.0126
-9.250 -0.7287 0.02159 0.01725 -0.0552 0.9892 0.0134
-9.000 -0.7007 0.01971 0.01499 -0.0564 0.9856 0.0144
-8.750 -0.6696 0.01957 0.01487 -0.0573 0.9810 0.0152
-8.500 -0.6362 0.01971 0.01502 -0.0586 0.9774 0.0158
-8.250 -0.6033 0.01949 0.01475 -0.0599 0.9745 0.0168
-8.000 -0.5757 0.01857 0.01364 -0.0603 0.9689 0.0180
-7.750 -0.5458 0.01754 0.01234 -0.0611 0.9648 0.0191
-7.500 -0.5160 0.01702 0.01173 -0.0619 0.9604 0.0199
-7.250 -0.4858 0.01699 0.01171 -0.0626 0.9550 0.0206
-7.000 -0.4549 0.01673 0.01139 -0.0634 0.9504 0.0215
-6.750 -0.4265 0.01630 0.01086 -0.0637 0.9444 0.0226
-6.500 -0.3981 0.01580 0.01023 -0.0639 0.9381 0.0238
-6.250 -0.3698 0.01534 0.00962 -0.0641 0.9324 0.0247
-6.000 -0.3430 0.01498 0.00914 -0.0639 0.9254 0.0253
-5.500 -0.2919 0.01339 0.00734 -0.0632 0.9119 0.0273
-5.250 -0.2653 0.01298 0.00685 -0.0630 0.9056 0.0280
-5.000 -0.2392 0.01265 0.00647 -0.0626 0.8991 0.0289
-4.750 -0.2131 0.01224 0.00599 -0.0622 0.8924 0.0297
-4.500 -0.1872 0.01180 0.00548 -0.0618 0.8860 0.0303
-4.250 -0.1613 0.01140 0.00501 -0.0613 0.8790 0.0309
-4.000 -0.1353 0.01104 0.00457 -0.0609 0.8729 0.0315
-3.750 -0.1092 0.01073 0.00422 -0.0604 0.8659 0.0323
-3.500 -0.0829 0.01048 0.00390 -0.0600 0.8593 0.0330
-3.250 -0.0570 0.01021 0.00358 -0.0595 0.8507 0.0335
-3.000 -0.0310 0.00997 0.00329 -0.0590 0.8416 0.0338
-2.750 -0.0055 0.00966 0.00291 -0.0584 0.8318 0.0345
-2.500 0.0201 0.00937 0.00257 -0.0578 0.8204 0.0356
-2.250 0.0459 0.00917 0.00231 -0.0573 0.8080 0.0367
-2.000 0.0720 0.00901 0.00210 -0.0568 0.7945 0.0378
-1.750 0.0979 0.00889 0.00192 -0.0562 0.7784 0.0391
-1.500 0.1239 0.00881 0.00176 -0.0557 0.7608 0.0408
-1.250 0.1499 0.00874 0.00161 -0.0551 0.7434 0.0432
-1.000 0.1759 0.00868 0.00148 -0.0546 0.7263 0.0476
-0.750 0.2017 0.00857 0.00138 -0.0541 0.7091 0.0618
-0.500 0.2270 0.00837 0.00132 -0.0536 0.6947 0.1187
-0.250 0.2526 0.00823 0.00128 -0.0531 0.6793 0.1682
0.000 0.2773 0.00807 0.00125 -0.0525 0.6589 0.2375
0.250 0.2946 0.00711 0.00122 -0.0509 0.6381 0.5704
0.500 0.3139 0.00672 0.00128 -0.0489 0.6166 0.7281
0.750 0.3347 0.00648 0.00133 -0.0471 0.5970 0.8342
1.250 0.4392 0.00674 0.00152 -0.0574 0.5203 0.9859
1.500 0.4850 0.00719 0.00160 -0.0617 0.4454 1.0000
1.750 0.5066 0.00748 0.00170 -0.0605 0.4063 1.0000
2.000 0.5286 0.00776 0.00180 -0.0593 0.3722 1.0000
2.250 0.5508 0.00804 0.00192 -0.0583 0.3387 1.0000
2.500 0.5734 0.00830 0.00204 -0.0573 0.3097 1.0000
2.750 0.5965 0.00853 0.00216 -0.0563 0.2872 1.0000
3.000 0.6199 0.00874 0.00229 -0.0555 0.2706 1.0000
3.250 0.6435 0.00894 0.00243 -0.0546 0.2582 1.0000
3.500 0.6672 0.00914 0.00257 -0.0538 0.2482 1.0000
3.750 0.6913 0.00931 0.00272 -0.0530 0.2401 1.0000
4.000 0.7152 0.00951 0.00288 -0.0523 0.2317 1.0000
4.500 0.7623 0.00995 0.00324 -0.0506 0.2084 1.0000
4.750 0.7862 0.01015 0.00342 -0.0499 0.2000 1.0000
5.000 0.8100 0.01037 0.00361 -0.0491 0.1911 1.0000
5.250 0.8340 0.01057 0.00380 -0.0484 0.1805 1.0000
5.500 0.8576 0.01081 0.00402 -0.0476 0.1672 1.0000
5.750 0.8806 0.01111 0.00424 -0.0467 0.1476 1.0000
6.000 0.9017 0.01158 0.00456 -0.0456 0.1151 1.0000
6.250 0.9226 0.01209 0.00495 -0.0445 0.0946 1.0000
6.500 0.9445 0.01251 0.00532 -0.0435 0.0851 1.0000
6.750 0.9676 0.01281 0.00565 -0.0427 0.0791 1.0000
7.000 0.9899 0.01319 0.00602 -0.0418 0.0713 1.0000
7.250 1.0127 0.01351 0.00635 -0.0410 0.0612 1.0000
7.750 1.0531 0.01466 0.00732 -0.0386 0.0334 1.0000
8.000 1.0739 0.01516 0.00784 -0.0375 0.0291 1.0000
8.250 1.0957 0.01556 0.00830 -0.0366 0.0266 1.0000
8.500 1.1162 0.01607 0.00884 -0.0355 0.0237 1.0000
8.750 1.1363 0.01660 0.00942 -0.0343 0.0211 1.0000
9.000 1.1570 0.01704 0.00992 -0.0333 0.0188 1.0000
9.250 1.1763 0.01760 0.01049 -0.0320 0.0167 1.0000
9.500 1.1948 0.01821 0.01115 -0.0306 0.0151 1.0000
9.750 1.2134 0.01876 0.01177 -0.0293 0.0139 1.0000
10.000 1.2310 0.01936 0.01242 -0.0279 0.0127 1.0000
10.250 1.2463 0.02012 0.01322 -0.0261 0.0116 1.0000
10.500 1.2615 0.02079 0.01397 -0.0243 0.0109 1.0000
10.750 1.2744 0.02149 0.01475 -0.0221 0.0104 1.0000
11.000 1.2861 0.02221 0.01555 -0.0198 0.0099 1.0000
11.250 1.2974 0.02298 0.01640 -0.0176 0.0095 1.0000
11.500 1.3073 0.02386 0.01735 -0.0153 0.0090 1.0000
11.750 1.3145 0.02494 0.01851 -0.0129 0.0086 1.0000
12.000 1.3203 0.02616 0.01984 -0.0105 0.0081 1.0000
12.250 1.3268 0.02735 0.02114 -0.0083 0.0081 1.0000
12.500 1.3352 0.02847 0.02237 -0.0066 0.0077 1.0000
12.750 1.3403 0.02988 0.02389 -0.0047 0.0074 1.0000
13.000 1.3475 0.03118 0.02529 -0.0032 0.0071 1.0000
13.250 1.3517 0.03279 0.02701 -0.0017 0.0069 1.0000
13.500 1.3526 0.03478 0.02911 -0.0003 0.0067 1.0000
13.750 1.3545 0.03678 0.03123 0.0007 0.0066 1.0000
14.000 1.3557 0.03896 0.03351 0.0015 0.0064 1.0000
14.250 1.3531 0.04168 0.03635 0.0021 0.0063 1.0000
14.500 1.3484 0.04479 0.03958 0.0022 0.0061 1.0000
14.750 1.3426 0.04824 0.04318 0.0019 0.0061 1.0000
15.000 1.3318 0.05256 0.04764 0.0011 0.0059 1.0000
15.250 1.3220 0.05701 0.05222 -0.0001 0.0059 1.0000
15.500 1.3139 0.06156 0.05692 -0.0017 0.0058 1.0000
15.750 1.3033 0.06672 0.06223 -0.0038 0.0058 1.0000
16.000 1.2899 0.07265 0.06831 -0.0065 0.0058 1.0000
16.250 1.2776 0.07872 0.07453 -0.0095 0.0057 1.0000
16.500 1.2649 0.08499 0.08095 -0.0126 0.0057 1.0000
16.750 1.2492 0.09194 0.08803 -0.0161 0.0057 1.0000
17.000 1.2337 0.09902 0.09525 -0.0197 0.0057 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 566 AIRFOIL (goe566-il)