GOE 802 AIRFOIL (goe802-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 802 AIRFOIL (goe802-il) Reynolds number: 1,000,000 Max Cl/Cd: 122.98 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe802-il-1000000-n5.txt Download as CSV file: xf-goe802-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 802 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.2407 0.10683 0.10483 -0.0406 0.8863 0.0110
-9.750 -0.2343 0.10375 0.10170 -0.0419 0.8785 0.0113
-9.500 -0.2361 0.09807 0.09600 -0.0446 0.8709 0.0121
-9.250 -0.2268 0.09579 0.09367 -0.0457 0.8633 0.0122
-9.000 -0.2168 0.09355 0.09143 -0.0469 0.8563 0.0124
-8.750 -0.2070 0.09125 0.08910 -0.0482 0.8488 0.0125
-8.500 -0.1972 0.08889 0.08672 -0.0496 0.8420 0.0127
-8.250 -0.1880 0.08631 0.08411 -0.0511 0.8345 0.0130
-8.000 -0.1798 0.08342 0.08120 -0.0530 0.8276 0.0133
-7.750 -0.1719 0.08034 0.07811 -0.0551 0.8200 0.0136
-7.500 -0.1697 0.07370 0.07141 -0.0622 0.8127 0.0146
-7.250 -0.1535 0.07142 0.06912 -0.0648 0.8059 0.0147
-7.000 -0.1362 0.06913 0.06679 -0.0676 0.7988 0.0149
-6.750 -0.1176 0.06663 0.06427 -0.0710 0.7919 0.0151
-6.500 -0.0977 0.06384 0.06144 -0.0748 0.7844 0.0153
-6.250 -0.0763 0.06047 0.05803 -0.0795 0.7774 0.0156
-6.000 -0.0530 0.05686 0.05436 -0.0845 0.7692 0.0159
-5.750 -0.0277 0.05297 0.05039 -0.0899 0.7615 0.0162
-5.500 0.0066 0.04513 0.04236 -0.1004 0.7527 0.0175
-5.250 0.0319 0.04299 0.04012 -0.1027 0.7403 0.0176
-5.000 0.0580 0.04080 0.03782 -0.1050 0.7260 0.0178
-4.750 0.0856 0.03824 0.03513 -0.1076 0.7114 0.0180
-4.500 0.1143 0.03534 0.03207 -0.1104 0.6987 0.0182
-4.250 0.1440 0.03220 0.02875 -0.1130 0.6862 0.0185
-4.000 0.1740 0.02890 0.02525 -0.1152 0.6735 0.0188
-3.750 0.2048 0.02504 0.02110 -0.1172 0.6616 0.0194
-3.500 0.2363 0.01923 0.01471 -0.1190 0.6504 0.0200
-3.250 0.2652 0.01580 0.01075 -0.1196 0.6369 0.0203
-3.000 0.2931 0.01426 0.00887 -0.1200 0.6220 0.0206
-2.750 0.3211 0.01348 0.00789 -0.1201 0.6071 0.0208
-2.500 0.3492 0.01294 0.00720 -0.1203 0.5941 0.0209
-2.250 0.3774 0.01250 0.00662 -0.1204 0.5815 0.0211
-2.000 0.4055 0.01210 0.00608 -0.1204 0.5684 0.0213
-1.500 0.4619 0.01136 0.00508 -0.1205 0.5437 0.0219
-1.250 0.4901 0.01102 0.00462 -0.1206 0.5340 0.0221
-1.000 0.5184 0.01070 0.00420 -0.1206 0.5242 0.0224
-0.750 0.5467 0.01042 0.00384 -0.1206 0.5155 0.0226
-0.250 0.6030 0.00999 0.00326 -0.1207 0.4964 0.0231
0.000 0.6311 0.00989 0.00310 -0.1207 0.4872 0.0234
0.250 0.6593 0.00979 0.00295 -0.1207 0.4783 0.0237
0.500 0.6873 0.00969 0.00281 -0.1207 0.4702 0.0239
0.750 0.7155 0.00956 0.00264 -0.1207 0.4615 0.0241
1.000 0.7435 0.00943 0.00249 -0.1208 0.4534 0.0246
1.250 0.7715 0.00939 0.00243 -0.1208 0.4440 0.0250
1.500 0.7994 0.00939 0.00240 -0.1208 0.4348 0.0254
1.750 0.8270 0.00943 0.00240 -0.1208 0.4241 0.0259
2.000 0.8545 0.00948 0.00240 -0.1208 0.4105 0.0264
2.250 0.8815 0.00959 0.00243 -0.1206 0.3941 0.0269
2.500 0.9086 0.00969 0.00248 -0.1205 0.3794 0.0274
2.750 0.9359 0.00979 0.00254 -0.1205 0.3688 0.0281
3.000 0.9631 0.00987 0.00258 -0.1204 0.3593 0.0292
3.250 0.9903 0.00997 0.00265 -0.1203 0.3491 0.0301
3.500 1.0173 0.01009 0.00274 -0.1202 0.3399 0.0312
3.750 1.0440 0.01024 0.00285 -0.1201 0.3303 0.0325
4.000 1.0708 0.01038 0.00296 -0.1199 0.3200 0.0344
4.250 1.0971 0.01055 0.00310 -0.1197 0.3083 0.0458
5.250 1.1987 0.00977 0.00396 -0.1187 0.2786 1.0000
5.500 1.2249 0.00996 0.00413 -0.1185 0.2706 1.0000
5.750 1.2508 0.01019 0.00431 -0.1182 0.2626 1.0000
6.000 1.2761 0.01045 0.00451 -0.1179 0.2518 1.0000
6.250 1.3009 0.01074 0.00474 -0.1175 0.2381 1.0000
6.500 1.3244 0.01114 0.00503 -0.1169 0.2187 1.0000
6.750 1.3449 0.01178 0.00547 -0.1159 0.1862 1.0000
7.000 1.3650 0.01243 0.00595 -0.1148 0.1599 1.0000
7.250 1.3867 0.01292 0.00634 -0.1140 0.1442 1.0000
7.500 1.4090 0.01333 0.00670 -0.1132 0.1344 1.0000
7.750 1.4316 0.01371 0.00703 -0.1125 0.1258 1.0000
8.000 1.4535 0.01411 0.00740 -0.1117 0.1172 1.0000
8.250 1.4745 0.01456 0.00779 -0.1108 0.1084 1.0000
8.500 1.4948 0.01503 0.00821 -0.1097 0.0979 1.0000
8.750 1.5120 0.01568 0.00876 -0.1082 0.0819 1.0000
9.000 1.5261 0.01647 0.00943 -0.1062 0.0647 1.0000
9.250 1.5422 0.01700 0.00994 -0.1044 0.0589 1.0000
9.500 1.5568 0.01757 0.01050 -0.1024 0.0543 1.0000
9.750 1.5725 0.01811 0.01104 -0.1007 0.0503 1.0000
10.000 1.5857 0.01881 0.01173 -0.0986 0.0450 1.0000
10.250 1.5997 0.01949 0.01242 -0.0968 0.0397 1.0000
10.500 1.5998 0.02109 0.01387 -0.0932 0.0178 1.0000
10.750 1.6103 0.02206 0.01488 -0.0912 0.0144 1.0000
11.000 1.6222 0.02298 0.01585 -0.0894 0.0130 1.0000
11.250 1.6349 0.02387 0.01678 -0.0879 0.0121 1.0000
11.500 1.6471 0.02481 0.01778 -0.0863 0.0114 1.0000
11.750 1.6581 0.02588 0.01890 -0.0847 0.0107 1.0000
12.000 1.6676 0.02708 0.02015 -0.0831 0.0101 1.0000
12.250 1.6774 0.02831 0.02143 -0.0816 0.0096 1.0000
12.500 1.6874 0.02954 0.02273 -0.0802 0.0093 1.0000
12.750 1.6965 0.03089 0.02414 -0.0789 0.0090 1.0000
13.000 1.7043 0.03236 0.02567 -0.0775 0.0086 1.0000
13.250 1.7113 0.03397 0.02735 -0.0762 0.0083 1.0000
13.500 1.7170 0.03574 0.02918 -0.0750 0.0080 1.0000
13.750 1.7211 0.03769 0.03120 -0.0737 0.0077 1.0000
14.000 1.7231 0.03992 0.03350 -0.0725 0.0074 1.0000
14.250 1.7276 0.04196 0.03562 -0.0716 0.0073 1.0000
14.500 1.7310 0.04418 0.03792 -0.0708 0.0071 1.0000
14.750 1.7335 0.04656 0.04038 -0.0700 0.0069 1.0000
15.000 1.7350 0.04915 0.04304 -0.0694 0.0067 1.0000
15.250 1.7356 0.05192 0.04590 -0.0690 0.0066 1.0000
15.500 1.7351 0.05490 0.04897 -0.0686 0.0064 1.0000
15.750 1.7335 0.05808 0.05224 -0.0684 0.0063 1.0000
16.000 1.7308 0.06148 0.05572 -0.0683 0.0061 1.0000
16.250 1.7269 0.06512 0.05945 -0.0684 0.0060 1.0000
16.500 1.7217 0.06900 0.06342 -0.0687 0.0059 1.0000
16.750 1.7150 0.07312 0.06764 -0.0690 0.0058 1.0000
17.000 1.7068 0.07752 0.07215 -0.0695 0.0056 1.0000
17.250 1.6971 0.08223 0.07696 -0.0702 0.0055 1.0000
17.500 1.6863 0.08716 0.08201 -0.0710 0.0055 1.0000
17.750 1.6774 0.09184 0.08679 -0.0719 0.0054 1.0000
18.000 1.6681 0.09666 0.09171 -0.0729 0.0054 1.0000
18.250 1.6581 0.10163 0.09679 -0.0741 0.0053 1.0000
18.500 1.6480 0.10669 0.10196 -0.0754 0.0053 1.0000
18.750 1.6368 0.11194 0.10732 -0.0769 0.0052 1.0000
19.000 1.6259 0.11724 0.11272 -0.0785 0.0052 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 802 AIRFOIL (goe802-il)