GOE 9K AIRFOIL (goe09k-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
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Airfoil: GOE 9K AIRFOIL (goe09k-il) Reynolds number: 1,000,000 Max Cl/Cd: 77.65 at α=1° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe09k-il-1000000.txt Download as CSV file: xf-goe09k-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 9K AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.000 -0.5682 0.07949 0.07793 -0.0048 1.0000 0.0019
-6.750 -0.5652 0.07539 0.07383 -0.0064 1.0000 0.0021
-6.500 -0.5523 0.07098 0.06940 -0.0104 1.0000 0.0021
-6.250 -0.5403 0.06677 0.06516 -0.0138 1.0000 0.0023
-6.000 -0.5250 0.06246 0.06080 -0.0171 1.0000 0.0025
-5.750 -0.5078 0.05797 0.05626 -0.0200 1.0000 0.0026
-5.500 -0.4881 0.05357 0.05177 -0.0225 1.0000 0.0031
-5.250 -0.4654 0.04904 0.04713 -0.0246 1.0000 0.0032
-5.000 -0.4423 0.04469 0.04263 -0.0259 1.0000 0.0032
-4.750 -0.4197 0.04062 0.03842 -0.0264 1.0000 0.0033
-4.500 -0.3979 0.03683 0.03446 -0.0263 1.0000 0.0033
-4.250 -0.3800 0.03259 0.03004 -0.0265 1.0000 0.0034
-4.000 -0.3600 0.02957 0.02684 -0.0261 1.0000 0.0036
-3.750 -0.3379 0.02648 0.02354 -0.0253 1.0000 0.0039
-3.500 -0.3148 0.02347 0.02029 -0.0242 1.0000 0.0043
-3.250 -0.2894 0.02068 0.01722 -0.0224 1.0000 0.0049
-2.000 -0.1697 0.01065 0.00580 -0.0165 1.0000 0.0109
-1.500 -0.1182 0.00773 0.00254 -0.0141 1.0000 0.0042
-1.250 -0.0932 0.00696 0.00164 -0.0132 1.0000 0.0047
-1.000 -0.0690 0.00660 0.00121 -0.0124 1.0000 0.0075
-0.750 -0.0400 0.00512 0.00090 -0.0133 0.9989 0.4119
-0.500 -0.0001 0.00349 0.00096 -0.0163 1.0000 1.0000
0.000 0.0789 0.00354 0.00091 -0.0219 0.9952 1.0000
0.250 0.1190 0.00350 0.00086 -0.0248 0.9910 1.0000
0.500 0.1672 0.00339 0.00078 -0.0295 0.9834 1.0000
0.750 0.2079 0.00330 0.00072 -0.0324 0.9740 1.0000
1.000 0.2710 0.00349 0.00050 -0.0400 0.8012 1.0000
1.250 0.2731 0.00576 0.00075 -0.0350 0.2326 1.0000
1.500 0.2925 0.00701 0.00130 -0.0334 0.0066 1.0000
1.750 0.3164 0.00758 0.00195 -0.0323 0.0042 1.0000
5.750 0.6629 0.05108 0.04895 -0.0119 0.0021 1.0000
6.000 0.6794 0.05563 0.05365 -0.0122 0.0020 1.0000
6.250 0.6942 0.06023 0.05837 -0.0129 0.0020 1.0000
6.500 0.7085 0.06509 0.06334 -0.0143 0.0020 1.0000
6.750 0.7207 0.07004 0.06838 -0.0164 0.0020 1.0000
7.000 0.7304 0.07501 0.07340 -0.0191 0.0020 1.0000
7.250 0.7375 0.07976 0.07819 -0.0221 0.0019 1.0000
7.500 0.7409 0.08427 0.08274 -0.0252 0.0020 1.0000
7.750 0.7412 0.08855 0.08702 -0.0285 0.0019 1.0000
8.000 0.7365 0.09256 0.09101 -0.0313 0.0019 1.0000
8.250 0.7362 0.09663 0.09506 -0.0337 0.0019 1.0000
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