GOE 565 AIRFOIL (goe565-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 565 AIRFOIL (goe565-il) Reynolds number: 200,000 Max Cl/Cd: 69.05 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe565-il-200000-n5.txt Download as CSV file: xf-goe565-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 565 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 -0.4585 0.08820 0.08454 -0.0334 1.0000 0.0181
-9.250 -0.4653 0.08363 0.08003 -0.0350 1.0000 0.0184
-9.000 -0.4748 0.07905 0.07551 -0.0365 1.0000 0.0185
-8.750 -0.4901 0.07440 0.07095 -0.0376 1.0000 0.0186
-8.500 -0.5114 0.06990 0.06656 -0.0384 1.0000 0.0186
-8.250 -0.5325 0.06296 0.05968 -0.0427 1.0000 0.0184
-8.000 -0.5791 0.04545 0.04189 -0.0522 1.0000 0.0180
-7.750 -0.5950 0.02914 0.02420 -0.0578 0.9965 0.0194
-7.500 -0.5718 0.02449 0.01872 -0.0595 0.9930 0.0216
-7.250 -0.5402 0.02479 0.01905 -0.0607 0.9900 0.0231
-7.000 -0.5096 0.02356 0.01753 -0.0621 0.9872 0.0256
-6.750 -0.4814 0.02150 0.01496 -0.0629 0.9841 0.0286
-6.500 -0.4507 0.02156 0.01504 -0.0639 0.9807 0.0309
-6.250 -0.4190 0.02074 0.01391 -0.0650 0.9780 0.0346
-6.000 -0.3895 0.01983 0.01277 -0.0658 0.9747 0.0375
-5.750 -0.3593 0.01951 0.01242 -0.0667 0.9709 0.0404
-5.500 -0.3268 0.01892 0.01162 -0.0678 0.9683 0.0438
-5.250 -0.2958 0.01856 0.01103 -0.0685 0.9650 0.0461
-5.000 -0.2679 0.01730 0.00969 -0.0690 0.9610 0.0492
-4.750 -0.2353 0.01667 0.00899 -0.0701 0.9583 0.0517
-4.500 -0.2013 0.01599 0.00820 -0.0715 0.9563 0.0536
-4.250 -0.1747 0.01543 0.00756 -0.0713 0.9509 0.0551
-4.000 -0.1420 0.01494 0.00696 -0.0723 0.9478 0.0573
-3.750 -0.1075 0.01450 0.00645 -0.0737 0.9455 0.0590
-3.500 -0.0802 0.01388 0.00583 -0.0737 0.9400 0.0611
-3.250 -0.0481 0.01342 0.00536 -0.0745 0.9356 0.0636
-3.000 -0.0137 0.01302 0.00493 -0.0758 0.9312 0.0663
-2.750 0.0155 0.01269 0.00457 -0.0759 0.9234 0.0694
-2.500 0.0501 0.01235 0.00421 -0.0772 0.9192 0.0739
-2.250 0.0761 0.01213 0.00402 -0.0767 0.9112 0.0822
-2.000 0.1081 0.01186 0.00377 -0.0775 0.9063 0.0981
-1.750 0.1354 0.01156 0.00355 -0.0773 0.8991 0.1206
-1.500 0.1643 0.01108 0.00336 -0.0776 0.8926 0.1876
-1.250 0.1918 0.01072 0.00328 -0.0776 0.8854 0.2816
-1.000 0.2201 0.01038 0.00317 -0.0775 0.8770 0.3477
-0.750 0.2457 0.00997 0.00308 -0.0770 0.8652 0.4378
-0.500 0.2709 0.00954 0.00303 -0.0762 0.8529 0.5546
-0.250 0.2961 0.00919 0.00296 -0.0752 0.8402 0.6509
0.000 0.3211 0.00881 0.00287 -0.0740 0.8271 0.7420
0.500 0.4181 0.00836 0.00268 -0.0818 0.8003 1.0000
0.750 0.4443 0.00840 0.00263 -0.0812 0.7852 1.0000
1.000 0.4701 0.00846 0.00259 -0.0805 0.7681 1.0000
1.250 0.4958 0.00853 0.00257 -0.0798 0.7502 1.0000
1.500 0.5214 0.00863 0.00258 -0.0791 0.7316 1.0000
1.750 0.5466 0.00875 0.00259 -0.0783 0.7109 1.0000
2.000 0.5713 0.00889 0.00264 -0.0774 0.6871 1.0000
2.250 0.5958 0.00904 0.00271 -0.0765 0.6621 1.0000
2.500 0.6197 0.00922 0.00278 -0.0755 0.6320 1.0000
2.750 0.6434 0.00943 0.00287 -0.0745 0.5996 1.0000
3.000 0.6665 0.00968 0.00298 -0.0733 0.5671 1.0000
3.250 0.6891 0.00998 0.00314 -0.0722 0.5344 1.0000
3.500 0.7105 0.01037 0.00333 -0.0708 0.4938 1.0000
3.750 0.7312 0.01082 0.00355 -0.0694 0.4456 1.0000
4.000 0.7520 0.01128 0.00380 -0.0680 0.3983 1.0000
4.250 0.7737 0.01171 0.00407 -0.0669 0.3619 1.0000
4.500 0.7959 0.01212 0.00437 -0.0658 0.3303 1.0000
4.750 0.8177 0.01258 0.00469 -0.0648 0.2960 1.0000
5.000 0.8391 0.01309 0.00505 -0.0637 0.2580 1.0000
5.250 0.8599 0.01366 0.00543 -0.0625 0.2163 1.0000
5.500 0.8816 0.01417 0.00583 -0.0615 0.1910 1.0000
5.750 0.9036 0.01466 0.00625 -0.0606 0.1741 1.0000
6.000 0.9260 0.01511 0.00672 -0.0597 0.1599 1.0000
6.250 0.9484 0.01557 0.00718 -0.0588 0.1440 1.0000
6.500 0.9705 0.01605 0.00764 -0.0578 0.1171 1.0000
6.750 0.9886 0.01696 0.00824 -0.0564 0.0761 1.0000
7.000 1.0079 0.01776 0.00896 -0.0551 0.0563 1.0000
7.250 1.0275 0.01851 0.00971 -0.0538 0.0472 1.0000
7.500 1.0469 0.01926 0.01048 -0.0525 0.0420 1.0000
7.750 1.0656 0.02007 0.01135 -0.0511 0.0381 1.0000
8.000 1.0846 0.02080 0.01221 -0.0497 0.0351 1.0000
8.250 1.1015 0.02171 0.01319 -0.0481 0.0318 1.0000
8.500 1.1150 0.02290 0.01444 -0.0461 0.0290 1.0000
8.750 1.1326 0.02369 0.01541 -0.0445 0.0270 1.0000
9.000 1.1486 0.02460 0.01643 -0.0428 0.0246 1.0000
9.250 1.1634 0.02555 0.01745 -0.0411 0.0225 1.0000
9.500 1.1712 0.02708 0.01902 -0.0384 0.0209 1.0000
9.750 1.1845 0.02809 0.02019 -0.0363 0.0199 1.0000
10.000 1.1964 0.02924 0.02148 -0.0342 0.0187 1.0000
10.250 1.2076 0.03038 0.02275 -0.0320 0.0175 1.0000
10.500 1.2183 0.03143 0.02390 -0.0300 0.0163 1.0000
10.750 1.2275 0.03259 0.02518 -0.0280 0.0155 1.0000
11.000 1.2326 0.03427 0.02694 -0.0257 0.0148 1.0000
11.250 1.2381 0.03618 0.02900 -0.0235 0.0143 1.0000
11.500 1.2440 0.03807 0.03110 -0.0215 0.0138 1.0000
11.750 1.2481 0.04015 0.03340 -0.0196 0.0134 1.0000
12.000 1.2506 0.04234 0.03579 -0.0178 0.0129 1.0000
12.250 1.2512 0.04472 0.03838 -0.0162 0.0126 1.0000
12.500 1.2506 0.04709 0.04094 -0.0149 0.0121 1.0000
12.750 1.2489 0.04963 0.04366 -0.0139 0.0117 1.0000
13.000 1.2466 0.05228 0.04645 -0.0133 0.0114 1.0000
13.250 1.2419 0.05545 0.04979 -0.0131 0.0112 1.0000
13.500 1.2354 0.05906 0.05358 -0.0133 0.0110 1.0000
13.750 1.2285 0.06286 0.05752 -0.0141 0.0108 1.0000
14.000 1.2189 0.06733 0.06216 -0.0154 0.0107 1.0000
14.250 1.2073 0.07241 0.06743 -0.0173 0.0105 1.0000
14.500 1.1932 0.07830 0.07351 -0.0200 0.0104 1.0000
14.750 1.1772 0.08501 0.08043 -0.0235 0.0104 1.0000
15.000 1.1602 0.09236 0.08797 -0.0277 0.0104 1.0000
15.250 1.1420 0.10050 0.09629 -0.0326 0.0104 1.0000
15.500 1.1198 0.11021 0.10622 -0.0386 0.0106 1.0000
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Polar data table (+)
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