GOE 279 (DAIMLER X) AIRFOIL (goe279-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 279 (DAIMLER X) AIRFOIL (goe279-il) Reynolds number: 1,000,000 Max Cl/Cd: 105.02 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe279-il-1000000-n5.txt Download as CSV file: xf-goe279-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 279 (DAIMLER X) 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3345 0.10906 0.10739 -0.0267 1.0000 0.0054
-9.500 -0.3301 0.10619 0.10453 -0.0274 1.0000 0.0052
-9.250 -0.3274 0.10335 0.10171 -0.0278 1.0000 0.0051
-9.000 -0.3163 0.09949 0.09785 -0.0310 0.9972 0.0050
-8.750 -0.3044 0.09556 0.09393 -0.0344 0.9935 0.0049
-8.500 -0.2926 0.09161 0.08998 -0.0379 0.9878 0.0049
-8.250 -0.2811 0.08769 0.08606 -0.0415 0.9800 0.0050
-8.000 -0.2707 0.08380 0.08216 -0.0450 0.9696 0.0051
-7.750 -0.2622 0.08001 0.07837 -0.0482 0.9566 0.0052
-7.500 -0.2503 0.07581 0.07414 -0.0528 0.9443 0.0053
-7.250 -0.2362 0.07138 0.06967 -0.0580 0.9329 0.0054
-7.000 -0.2195 0.06615 0.06438 -0.0644 0.9221 0.0056
-6.750 -0.1998 0.06126 0.05943 -0.0705 0.9116 0.0058
-6.500 -0.1777 0.05675 0.05484 -0.0761 0.9021 0.0058
-6.250 -0.1538 0.05270 0.05069 -0.0809 0.8933 0.0059
-6.000 -0.1283 0.04887 0.04677 -0.0852 0.8845 0.0061
-5.000 -0.0090 0.01111 0.00634 -0.1043 0.8548 0.0116
-4.750 0.0183 0.01042 0.00551 -0.1043 0.8466 0.0126
-4.500 0.0458 0.00960 0.00452 -0.1043 0.8380 0.0137
-4.250 0.0738 0.00950 0.00438 -0.1042 0.8302 0.0148
-4.000 0.1021 0.00972 0.00461 -0.1041 0.8215 0.0157
-3.750 0.1302 0.00960 0.00443 -0.1041 0.8124 0.0168
-3.500 0.1580 0.00916 0.00385 -0.1040 0.8031 0.0180
-3.250 0.1861 0.00912 0.00377 -0.1040 0.7909 0.0187
-3.000 0.2141 0.00933 0.00395 -0.1039 0.7753 0.0194
-2.750 0.2418 0.00945 0.00399 -0.1037 0.7556 0.0202
-2.500 0.2694 0.00932 0.00375 -0.1036 0.7317 0.0212
-2.250 0.2959 0.00932 0.00353 -0.1032 0.6867 0.0224
-2.000 0.3221 0.00941 0.00338 -0.1029 0.6410 0.0231
-1.750 0.3492 0.00942 0.00324 -0.1027 0.6138 0.0234
-1.500 0.3764 0.00906 0.00275 -0.1026 0.5925 0.0246
-1.250 0.4037 0.00901 0.00260 -0.1026 0.5681 0.0252
-1.000 0.4311 0.00899 0.00247 -0.1025 0.5449 0.0259
-0.750 0.4587 0.00894 0.00234 -0.1024 0.5259 0.0266
-0.500 0.4864 0.00891 0.00224 -0.1024 0.5089 0.0276
-0.250 0.5140 0.00887 0.00211 -0.1023 0.4890 0.0283
0.000 0.5413 0.00888 0.00201 -0.1023 0.4636 0.0290
0.250 0.5681 0.00898 0.00195 -0.1021 0.4278 0.0295
0.500 0.5947 0.00911 0.00191 -0.1020 0.3926 0.0299
0.750 0.6217 0.00921 0.00189 -0.1019 0.3670 0.0302
1.000 0.6488 0.00930 0.00189 -0.1018 0.3458 0.0305
1.250 0.6756 0.00947 0.00194 -0.1016 0.3215 0.0309
1.500 0.7027 0.00954 0.00190 -0.1015 0.3033 0.0320
1.750 0.7301 0.00960 0.00191 -0.1015 0.2911 0.0337
2.000 0.7577 0.00966 0.00195 -0.1015 0.2820 0.0357
2.250 0.7850 0.00976 0.00201 -0.1014 0.2730 0.0376
2.500 0.8124 0.00984 0.00207 -0.1013 0.2648 0.0398
2.750 0.8396 0.00995 0.00215 -0.1013 0.2544 0.0438
3.000 0.8666 0.01008 0.00225 -0.1012 0.2434 0.0567
3.500 0.9147 0.00871 0.00269 -0.1003 0.2191 1.0000
3.750 0.9410 0.00896 0.00284 -0.1001 0.2022 1.0000
4.000 0.9670 0.00925 0.00301 -0.0998 0.1825 1.0000
4.250 0.9920 0.00965 0.00325 -0.0995 0.1539 1.0000
4.500 1.0165 0.01010 0.00355 -0.0990 0.1263 1.0000
4.750 1.0424 0.01039 0.00377 -0.0988 0.1164 1.0000
5.000 1.0687 0.01059 0.00395 -0.0986 0.1117 1.0000
5.250 1.0947 0.01084 0.00418 -0.0984 0.1063 1.0000
5.500 1.1207 0.01107 0.00439 -0.0981 0.1013 1.0000
5.750 1.1467 0.01129 0.00460 -0.0979 0.0962 1.0000
6.000 1.1719 0.01161 0.00487 -0.0976 0.0873 1.0000
6.250 1.1924 0.01247 0.00545 -0.0966 0.0423 1.0000
6.500 1.2171 0.01282 0.00578 -0.0961 0.0371 1.0000
6.750 1.2422 0.01310 0.00607 -0.0958 0.0348 1.0000
7.000 1.2673 0.01337 0.00638 -0.0954 0.0342 1.0000
7.250 1.2921 0.01365 0.00668 -0.0951 0.0337 1.0000
7.500 1.3167 0.01395 0.00702 -0.0946 0.0332 1.0000
7.750 1.3410 0.01427 0.00737 -0.0942 0.0328 1.0000
8.000 1.3650 0.01460 0.00774 -0.0937 0.0323 1.0000
8.250 1.3885 0.01497 0.00813 -0.0931 0.0311 1.0000
8.500 1.4112 0.01540 0.00858 -0.0924 0.0293 1.0000
8.750 1.4337 0.01584 0.00905 -0.0917 0.0277 1.0000
9.000 1.4557 0.01630 0.00957 -0.0909 0.0264 1.0000
9.250 1.4780 0.01670 0.01001 -0.0902 0.0257 1.0000
9.500 1.5009 0.01701 0.01036 -0.0896 0.0253 1.0000
9.750 1.5234 0.01734 0.01074 -0.0889 0.0246 1.0000
10.000 1.5453 0.01770 0.01115 -0.0882 0.0232 1.0000
10.250 1.5661 0.01814 0.01160 -0.0873 0.0205 1.0000
10.500 1.5853 0.01869 0.01214 -0.0862 0.0173 1.0000
10.750 1.6026 0.01936 0.01276 -0.0848 0.0126 1.0000
11.000 1.6185 0.02008 0.01349 -0.0831 0.0098 1.0000
11.250 1.6335 0.02079 0.01424 -0.0813 0.0083 1.0000
11.500 1.6456 0.02152 0.01503 -0.0789 0.0071 1.0000
11.750 1.6576 0.02226 0.01583 -0.0766 0.0064 1.0000
12.000 1.6680 0.02312 0.01675 -0.0742 0.0057 1.0000
12.250 1.6783 0.02401 0.01771 -0.0719 0.0052 1.0000
12.500 1.6889 0.02490 0.01869 -0.0698 0.0049 1.0000
12.750 1.6983 0.02591 0.01977 -0.0676 0.0045 1.0000
13.000 1.7063 0.02704 0.02098 -0.0654 0.0042 1.0000
13.250 1.7128 0.02832 0.02234 -0.0632 0.0040 1.0000
13.500 1.7175 0.02981 0.02393 -0.0610 0.0037 1.0000
13.750 1.7236 0.03122 0.02544 -0.0591 0.0036 1.0000
14.000 1.7287 0.03280 0.02712 -0.0573 0.0035 1.0000
14.250 1.7324 0.03456 0.02898 -0.0557 0.0033 1.0000
14.500 1.7346 0.03652 0.03104 -0.0541 0.0032 1.0000
14.750 1.7358 0.03869 0.03332 -0.0527 0.0031 1.0000
15.000 1.7354 0.04111 0.03585 -0.0515 0.0030 1.0000
15.250 1.7330 0.04386 0.03871 -0.0505 0.0029 1.0000
15.500 1.7291 0.04697 0.04194 -0.0499 0.0028 1.0000
15.750 1.7231 0.05053 0.04562 -0.0496 0.0027 1.0000
16.000 1.7146 0.05464 0.04986 -0.0499 0.0026 1.0000
16.250 1.7036 0.05943 0.05478 -0.0507 0.0025 1.0000
16.500 1.6907 0.06474 0.06024 -0.0519 0.0024 1.0000
16.750 1.6781 0.07014 0.06578 -0.0534 0.0024 1.0000
17.000 1.6632 0.07607 0.07186 -0.0553 0.0024 1.0000
17.250 1.6471 0.08236 0.07828 -0.0574 0.0024 1.0000
17.500 1.6273 0.08937 0.08544 -0.0599 0.0024 1.0000
17.750 1.6064 0.09673 0.09295 -0.0627 0.0024 1.0000
18.000 1.5844 0.10446 0.10082 -0.0658 0.0024 1.0000
18.250 1.5617 0.11237 0.10886 -0.0691 0.0024 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 279 (DAIMLER X) AIRFOIL (goe279-il)