GOE 6K AIRFOIL (goe06k-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 6K AIRFOIL (goe06k-il) Reynolds number: 200,000 Max Cl/Cd: 76.99 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe06k-il-200000-n5.txt Download as CSV file: xf-goe06k-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 6K 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4422 0.11549 0.11201 -0.0273 1.0000 0.0137
-9.250 -0.4441 0.11286 0.10941 -0.0268 1.0000 0.0140
-9.000 -0.4467 0.11035 0.10693 -0.0262 1.0000 0.0147
-8.750 -0.4408 0.10675 0.10335 -0.0283 0.9988 0.0145
-8.500 -0.4314 0.10287 0.09948 -0.0317 0.9967 0.0153
-8.250 -0.4224 0.09925 0.09588 -0.0375 0.9936 0.0167
-8.000 -0.4138 0.09571 0.09236 -0.0438 0.9895 0.0171
-7.500 -0.3957 0.08753 0.08419 -0.0533 0.9808 0.0172
-7.000 -0.3779 0.07665 0.07329 -0.0618 0.9719 0.0178
-6.750 -0.3675 0.07258 0.06919 -0.0632 0.9694 0.0185
-6.500 -0.3568 0.06873 0.06530 -0.0656 0.9647 0.0190
-6.250 -0.3432 0.06471 0.06121 -0.0684 0.9600 0.0195
-6.000 -0.3233 0.06010 0.05650 -0.0726 0.9570 0.0204
-5.750 -0.3098 0.05614 0.05242 -0.0742 0.9519 0.0209
-5.500 -0.2931 0.05217 0.04830 -0.0758 0.9471 0.0213
-5.250 -0.2702 0.04767 0.04359 -0.0782 0.9441 0.0214
-5.000 -0.2419 0.04016 0.03557 -0.0800 0.9418 0.0121
-4.750 -0.2331 0.03726 0.03246 -0.0776 0.9353 0.0116
-4.500 -0.2122 0.03371 0.02859 -0.0772 0.9321 0.0108
-4.250 -0.1877 0.02975 0.02416 -0.0768 0.9298 0.0107
-4.000 -0.1608 0.02636 0.02024 -0.0764 0.9282 0.0117
-3.750 -0.1432 0.02390 0.01734 -0.0740 0.9238 0.0118
-3.500 -0.1218 0.02149 0.01441 -0.0722 0.9204 0.0120
-3.250 -0.0954 0.01949 0.01193 -0.0714 0.9184 0.0124
-3.000 -0.0676 0.01832 0.01057 -0.0715 0.9166 0.0143
-2.750 -0.0382 0.01722 0.00927 -0.0717 0.9153 0.0160
-2.500 -0.0091 0.01602 0.00790 -0.0718 0.9140 0.0178
-2.250 0.0112 0.01549 0.00732 -0.0703 0.9099 0.0227
-2.000 0.0326 0.01499 0.00670 -0.0690 0.9060 0.0282
-1.750 0.0599 0.01463 0.00631 -0.0689 0.9036 0.0441
-1.500 0.0885 0.01420 0.00588 -0.0692 0.9017 0.0704
-1.250 0.1175 0.01368 0.00556 -0.0698 0.9003 0.1301
-0.750 0.2602 0.01168 0.00586 -0.0905 0.9078 1.0000
-0.500 0.2867 0.01168 0.00575 -0.0904 0.9045 1.0000
-0.250 0.3165 0.01166 0.00562 -0.0909 0.9023 1.0000
0.000 0.3478 0.01163 0.00552 -0.0919 0.9007 1.0000
0.250 0.3805 0.01159 0.00542 -0.0931 0.8994 1.0000
0.750 0.4230 0.01178 0.00555 -0.0906 0.8897 1.0000
1.000 0.4533 0.01178 0.00555 -0.0914 0.8875 1.0000
1.250 0.4859 0.01175 0.00554 -0.0926 0.8859 1.0000
1.500 0.5218 0.01168 0.00550 -0.0945 0.8844 1.0000
1.750 0.5347 0.01181 0.00568 -0.0913 0.8758 1.0000
2.000 0.5722 0.01161 0.00555 -0.0933 0.8724 1.0000
2.250 0.6017 0.01144 0.00546 -0.0936 0.8650 1.0000
2.500 0.6386 0.01116 0.00528 -0.0953 0.8579 1.0000
2.750 0.6623 0.01107 0.00529 -0.0942 0.8480 1.0000
3.000 0.6931 0.01082 0.00517 -0.0945 0.8351 1.0000
3.250 0.7270 0.01040 0.00485 -0.0951 0.8094 1.0000
3.500 0.7784 0.01011 0.00393 -0.0987 0.6335 1.0000
3.750 0.7603 0.01159 0.00428 -0.0886 0.4348 1.0000
4.000 0.7473 0.01321 0.00494 -0.0806 0.2452 1.0000
4.250 0.7458 0.01474 0.00565 -0.0752 0.0861 1.0000
4.750 0.7683 0.01656 0.00719 -0.0686 0.0184 1.0000
5.000 0.7837 0.01722 0.00800 -0.0661 0.0156 1.0000
5.250 0.7966 0.01814 0.00900 -0.0632 0.0134 1.0000
5.500 0.8117 0.01896 0.00992 -0.0607 0.0116 1.0000
5.750 0.8272 0.01997 0.01102 -0.0583 0.0108 1.0000
6.000 0.8454 0.02144 0.01252 -0.0567 0.0100 1.0000
6.250 0.8751 0.02333 0.01460 -0.0571 0.0095 1.0000
6.500 0.9037 0.02494 0.01645 -0.0572 0.0081 1.0000
6.750 0.9313 0.02688 0.01858 -0.0574 0.0075 1.0000
7.000 0.9594 0.02985 0.02186 -0.0575 0.0073 1.0000
7.250 0.9806 0.03284 0.02532 -0.0560 0.0072 1.0000
7.500 0.9954 0.03598 0.02887 -0.0534 0.0073 1.0000
8.000 1.0093 0.04481 0.03897 -0.0439 0.0084 1.0000
8.250 1.0092 0.04856 0.04307 -0.0393 0.0087 1.0000
8.500 1.0066 0.05198 0.04679 -0.0347 0.0089 1.0000
8.750 1.0014 0.05537 0.05045 -0.0300 0.0090 1.0000
9.000 0.9937 0.05875 0.05409 -0.0252 0.0085 1.0000
9.250 0.9820 0.06164 0.05717 -0.0202 0.0085 1.0000
9.500 0.9659 0.06488 0.06059 -0.0152 0.0087 1.0000
9.750 0.9514 0.06783 0.06369 -0.0113 0.0082 1.0000
10.000 0.9369 0.07120 0.06718 -0.0082 0.0105 1.0000
10.250 0.9208 0.07438 0.07048 -0.0059 0.0103 1.0000
10.500 0.8979 0.07971 0.07598 -0.0046 0.0106 1.0000
10.750 0.8832 0.08339 0.07974 -0.0046 0.0104 1.0000
11.000 0.8675 0.08795 0.08438 -0.0058 0.0102 1.0000
11.250 0.8544 0.09238 0.08885 -0.0076 0.0101 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 6K AIRFOIL (goe06k-il)