GOE 155 (SSW D.1) AIRFOIL (goe155-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 155 (SSW D.1) AIRFOIL (goe155-il) Reynolds number: 100,000 Max Cl/Cd: 61.57 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe155-il-100000-n5.txt Download as CSV file: xf-goe155-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 155 (SSW D.1) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3070 0.10456 0.09999 -0.0287 1.0000 0.0339
-8.000 -0.3092 0.10316 0.09869 -0.0312 1.0000 0.0341
-7.750 -0.3122 0.10171 0.09734 -0.0329 1.0000 0.0342
-7.500 -0.3003 0.09546 0.09111 -0.0279 1.0000 0.0357
-7.250 -0.3036 0.09360 0.08932 -0.0265 0.9985 0.0365
-7.000 -0.2823 0.08957 0.08525 -0.0317 0.9898 0.0383
-6.750 -0.2585 0.08550 0.08115 -0.0384 0.9812 0.0405
-6.500 -0.2208 0.08164 0.07721 -0.0538 0.9695 0.0434
-6.250 -0.1944 0.07670 0.07222 -0.0605 0.9625 0.0441
-6.000 -0.1833 0.07321 0.06876 -0.0577 0.9562 0.0471
-5.750 -0.1314 0.06978 0.06506 -0.0754 0.9451 0.0538
-5.500 -0.1229 0.06534 0.06071 -0.0727 0.9391 0.0551
-5.250 -0.1038 0.06224 0.05756 -0.0738 0.9315 0.0580
-5.000 -0.0490 0.05881 0.05371 -0.0880 0.9228 0.0645
-4.750 -0.0445 0.05512 0.05018 -0.0844 0.9152 0.0666
-4.500 0.0028 0.05280 0.04739 -0.0931 0.9081 0.0757
-4.250 0.0207 0.04869 0.04331 -0.0936 0.9006 0.0767
-4.000 0.0424 0.04565 0.04023 -0.0943 0.8941 0.0784
-3.750 0.0798 0.02489 0.01945 -0.0922 0.8639 0.0885
-3.250 0.1441 0.03425 0.02765 -0.1021 0.8740 0.0530
-3.000 0.1704 0.03186 0.02506 -0.1026 0.8651 0.0516
-2.750 0.2002 0.02938 0.02214 -0.1031 0.8560 0.0525
-2.500 0.2290 0.02723 0.01959 -0.1032 0.8481 0.0525
-2.250 0.2567 0.02543 0.01746 -0.1033 0.8391 0.0516
-2.000 0.2856 0.02376 0.01542 -0.1032 0.8333 0.0509
-1.750 0.3137 0.02242 0.01375 -0.1032 0.8254 0.0506
-1.500 0.3420 0.02123 0.01223 -0.1029 0.8191 0.0506
-1.250 0.3698 0.02036 0.01107 -0.1027 0.8114 0.0520
-1.000 0.3977 0.01967 0.01009 -0.1022 0.8047 0.0539
-0.750 0.4247 0.01892 0.00924 -0.1019 0.7968 0.0544
-0.500 0.4518 0.01827 0.00849 -0.1014 0.7899 0.0547
-0.250 0.4786 0.01777 0.00795 -0.1010 0.7816 0.0554
0.000 0.5055 0.01730 0.00744 -0.1005 0.7743 0.0563
0.250 0.5320 0.01698 0.00711 -0.1001 0.7654 0.0576
0.500 0.5586 0.01666 0.00670 -0.0994 0.7578 0.0595
0.750 0.5854 0.01648 0.00645 -0.0990 0.7469 0.0623
1.000 0.6123 0.01628 0.00618 -0.0985 0.7352 0.0665
1.250 0.6389 0.01605 0.00590 -0.0978 0.7217 0.0791
1.500 0.6655 0.01568 0.00578 -0.0973 0.7066 0.1675
2.000 0.7193 0.01406 0.00557 -0.0965 0.6775 1.0000
2.250 0.7456 0.01417 0.00555 -0.0959 0.6640 1.0000
2.500 0.7719 0.01428 0.00557 -0.0955 0.6496 1.0000
2.750 0.7980 0.01441 0.00563 -0.0950 0.6339 1.0000
3.000 0.8240 0.01454 0.00569 -0.0945 0.6168 1.0000
3.250 0.8497 0.01469 0.00576 -0.0939 0.5983 1.0000
3.500 0.8754 0.01485 0.00583 -0.0933 0.5792 1.0000
3.750 0.9008 0.01505 0.00598 -0.0928 0.5587 1.0000
4.000 0.9261 0.01529 0.00614 -0.0922 0.5386 1.0000
4.250 0.9511 0.01557 0.00632 -0.0916 0.5192 1.0000
4.500 0.9760 0.01589 0.00661 -0.0910 0.4996 1.0000
4.750 1.0005 0.01625 0.00691 -0.0904 0.4807 1.0000
5.000 1.0249 0.01665 0.00725 -0.0898 0.4627 1.0000
5.250 1.0490 0.01708 0.00764 -0.0892 0.4457 1.0000
5.500 1.0731 0.01753 0.00810 -0.0886 0.4297 1.0000
5.750 1.0971 0.01800 0.00858 -0.0881 0.4146 1.0000
6.000 1.1205 0.01849 0.00909 -0.0874 0.3975 1.0000
6.250 1.1433 0.01899 0.00962 -0.0867 0.3777 1.0000
6.500 1.1659 0.01951 0.01020 -0.0860 0.3596 1.0000
6.750 1.1886 0.02002 0.01079 -0.0853 0.3438 1.0000
7.000 1.2094 0.02058 0.01136 -0.0843 0.3203 1.0000
7.250 1.2290 0.02122 0.01197 -0.0833 0.2951 1.0000
7.500 1.2462 0.02203 0.01272 -0.0820 0.2586 1.0000
7.750 1.2604 0.02318 0.01365 -0.0804 0.2103 1.0000
8.000 1.2590 0.02608 0.01551 -0.0774 0.0711 1.0000
8.250 1.2672 0.02808 0.01741 -0.0750 0.0484 1.0000
8.500 1.2777 0.02966 0.01909 -0.0729 0.0406 1.0000
8.750 1.2845 0.03142 0.02099 -0.0705 0.0362 1.0000
9.000 1.2900 0.03298 0.02279 -0.0678 0.0338 1.0000
9.250 1.2920 0.03473 0.02476 -0.0650 0.0315 1.0000
9.500 1.2914 0.03678 0.02695 -0.0624 0.0295 1.0000
9.750 1.2854 0.03938 0.02968 -0.0599 0.0279 1.0000
10.000 1.2835 0.04181 0.03226 -0.0578 0.0267 1.0000
10.250 1.2845 0.04411 0.03473 -0.0561 0.0259 1.0000
10.500 1.2860 0.04648 0.03724 -0.0544 0.0250 1.0000
10.750 1.2895 0.04878 0.03966 -0.0528 0.0241 1.0000
11.000 1.2950 0.05099 0.04198 -0.0513 0.0228 1.0000
11.250 1.3009 0.05321 0.04431 -0.0499 0.0214 1.0000
11.500 1.3067 0.05551 0.04668 -0.0487 0.0201 1.0000
11.750 1.3158 0.05781 0.04906 -0.0472 0.0193 1.0000
12.000 1.3315 0.06036 0.05167 -0.0453 0.0185 1.0000
12.250 1.3450 0.06342 0.05492 -0.0437 0.0180 1.0000
12.500 1.3481 0.06650 0.05827 -0.0426 0.0178 1.0000
12.750 1.3463 0.06985 0.06192 -0.0418 0.0175 1.0000
13.000 1.3412 0.07349 0.06585 -0.0412 0.0171 1.0000
13.250 1.3338 0.07745 0.07010 -0.0411 0.0168 1.0000
13.500 1.3244 0.08175 0.07468 -0.0413 0.0164 1.0000
13.750 1.3135 0.08639 0.07958 -0.0419 0.0162 1.0000
14.000 1.3010 0.09142 0.08487 -0.0431 0.0160 1.0000
14.250 1.2873 0.09683 0.09053 -0.0447 0.0158 1.0000
14.500 1.2726 0.10267 0.09660 -0.0470 0.0157 1.0000
14.750 1.2574 0.10901 0.10318 -0.0499 0.0157 1.0000
15.000 1.2418 0.11585 0.11023 -0.0535 0.0157 1.0000
15.250 1.2257 0.12323 0.11781 -0.0577 0.0158 1.0000
15.500 1.2095 0.13115 0.12593 -0.0626 0.0159 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 155 (SSW D.1) AIRFOIL (goe155-il)