GOE 428 AIRFOIL (goe428-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 428 AIRFOIL (goe428-il) Reynolds number: 100,000 Max Cl/Cd: 58.41 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe428-il-100000.txt Download as CSV file: xf-goe428-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 428 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3059 0.10018 0.09570 -0.0311 1.0000 0.0616
-7.500 -0.3167 0.09945 0.09510 -0.0312 1.0000 0.0618
-7.250 -0.3237 0.09848 0.09424 -0.0333 1.0000 0.0621
-7.000 -0.3286 0.09719 0.09302 -0.0353 1.0000 0.0623
-6.750 -0.3197 0.09006 0.08594 -0.0262 1.0000 0.0643
-6.500 -0.3219 0.08790 0.08385 -0.0233 1.0000 0.0662
-6.250 -0.3291 0.08633 0.08237 -0.0213 1.0000 0.0677
-6.000 -0.3367 0.08486 0.08097 -0.0199 1.0000 0.0698
-5.750 -0.3427 0.08333 0.07950 -0.0199 1.0000 0.0723
-5.500 -0.3408 0.08227 0.07844 -0.0259 1.0000 0.0749
-5.250 -0.3318 0.07984 0.07594 -0.0323 1.0000 0.0760
-5.000 -0.3339 0.07572 0.07195 -0.0253 0.9983 0.0781
-4.750 -0.2729 0.07152 0.06746 -0.0424 0.9891 0.0888
-4.500 -0.2503 0.06591 0.06192 -0.0440 0.9837 0.0915
-4.250 -0.2182 0.06222 0.05816 -0.0482 0.9753 0.0988
-4.000 -0.1734 0.05740 0.05318 -0.0571 0.9681 0.1069
-3.750 -0.1269 0.05310 0.04863 -0.0655 0.9592 0.1192
-3.500 -0.0883 0.04940 0.04478 -0.0706 0.9503 0.1339
-3.250 -0.0416 0.04541 0.04063 -0.0765 0.9431 0.1498
-3.000 -0.0016 0.04202 0.03705 -0.0808 0.9324 0.1661
-2.750 0.0427 0.03866 0.03344 -0.0857 0.9235 0.1906
-2.500 0.0830 0.03591 0.03059 -0.0890 0.9161 0.2340
-2.250 0.1609 0.02772 0.02019 -0.0957 0.9113 0.0888
-2.000 0.2019 0.02520 0.01740 -0.0982 0.9035 0.0927
-1.750 0.2453 0.02328 0.01508 -0.1004 0.8965 0.0961
-1.500 0.2878 0.02180 0.01315 -0.1022 0.8889 0.1034
-1.250 0.3241 0.02061 0.01190 -0.1033 0.8795 0.1129
-1.000 0.3673 0.01932 0.01051 -0.1053 0.8728 0.1276
-0.750 0.3988 0.01839 0.00964 -0.1054 0.8616 0.1564
-0.500 0.4313 0.01682 0.00857 -0.1058 0.8514 0.2630
-0.250 0.4853 0.01448 0.00780 -0.1099 0.8444 1.0000
0.000 0.5149 0.01447 0.00750 -0.1095 0.8313 1.0000
0.250 0.5437 0.01447 0.00727 -0.1091 0.8177 1.0000
0.500 0.5717 0.01450 0.00709 -0.1085 0.8036 1.0000
0.750 0.5990 0.01455 0.00697 -0.1077 0.7890 1.0000
1.000 0.6257 0.01462 0.00688 -0.1069 0.7741 1.0000
1.250 0.6521 0.01470 0.00680 -0.1061 0.7588 1.0000
1.500 0.6783 0.01477 0.00673 -0.1051 0.7429 1.0000
1.750 0.7023 0.01490 0.00676 -0.1040 0.7245 1.0000
2.000 0.7275 0.01500 0.00674 -0.1029 0.7065 1.0000
2.250 0.7534 0.01510 0.00673 -0.1020 0.6891 1.0000
2.500 0.7797 0.01524 0.00673 -0.1012 0.6721 1.0000
2.750 0.8038 0.01549 0.00690 -0.1002 0.6527 1.0000
3.000 0.8289 0.01575 0.00705 -0.0994 0.6347 1.0000
3.250 0.8543 0.01604 0.00726 -0.0987 0.6179 1.0000
3.500 0.8797 0.01638 0.00751 -0.0980 0.6023 1.0000
3.750 0.9049 0.01677 0.00783 -0.0974 0.5880 1.0000
4.000 0.9297 0.01718 0.00823 -0.0968 0.5746 1.0000
4.250 0.9545 0.01760 0.00868 -0.0962 0.5618 1.0000
4.500 0.9792 0.01801 0.00910 -0.0955 0.5497 1.0000
4.750 1.0042 0.01842 0.00952 -0.0950 0.5387 1.0000
5.000 1.0306 0.01881 0.00987 -0.0946 0.5292 1.0000
5.250 1.0540 0.01927 0.01044 -0.0938 0.5182 1.0000
5.500 1.0781 0.01973 0.01103 -0.0932 0.5081 1.0000
5.750 1.1037 0.02014 0.01147 -0.0927 0.4991 1.0000
6.000 1.1277 0.02057 0.01199 -0.0920 0.4889 1.0000
6.250 1.1511 0.02107 0.01265 -0.0912 0.4791 1.0000
6.500 1.1746 0.02140 0.01309 -0.0903 0.4669 1.0000
6.750 1.1965 0.02119 0.01273 -0.0886 0.4442 1.0000
7.000 1.2139 0.02125 0.01290 -0.0865 0.4217 1.0000
7.250 1.2339 0.02139 0.01305 -0.0849 0.4024 1.0000
7.500 1.2490 0.02148 0.01329 -0.0825 0.3766 1.0000
7.750 1.2634 0.02163 0.01356 -0.0799 0.3490 1.0000
8.000 1.2727 0.02189 0.01385 -0.0766 0.3047 1.0000
8.250 1.2730 0.02309 0.01451 -0.0724 0.1850 1.0000
8.500 1.2606 0.02642 0.01685 -0.0675 0.0842 1.0000
8.750 1.2603 0.02858 0.01895 -0.0638 0.0651 1.0000
9.000 1.2624 0.03016 0.02064 -0.0602 0.0579 1.0000
9.250 1.2605 0.03197 0.02249 -0.0566 0.0528 1.0000
9.500 1.2607 0.03368 0.02434 -0.0535 0.0493 1.0000
9.750 1.2591 0.03563 0.02642 -0.0506 0.0466 1.0000
10.000 1.2561 0.03782 0.02876 -0.0480 0.0453 1.0000
10.250 1.2529 0.04025 0.03126 -0.0458 0.0441 1.0000
10.500 1.2523 0.04277 0.03379 -0.0436 0.0430 1.0000
10.750 1.2655 0.04478 0.03579 -0.0416 0.0421 1.0000
11.000 1.2913 0.04660 0.03770 -0.0401 0.0412 1.0000
11.250 1.3160 0.04880 0.04008 -0.0391 0.0394 1.0000
11.500 1.3426 0.05157 0.04307 -0.0384 0.0382 1.0000
11.750 1.3677 0.05524 0.04709 -0.0377 0.0386 1.0000
12.000 1.3769 0.05916 0.05143 -0.0362 0.0393 1.0000
12.250 1.3780 0.06319 0.05581 -0.0346 0.0401 1.0000
12.500 1.3746 0.06738 0.06032 -0.0331 0.0408 1.0000
12.750 1.3690 0.07200 0.06521 -0.0319 0.0416 1.0000
13.000 1.3762 0.07814 0.07158 -0.0316 0.0428 1.0000
13.250 1.3587 0.08058 0.07432 -0.0302 0.0434 1.0000
13.500 1.3338 0.08473 0.07880 -0.0301 0.0442 1.0000
13.750 1.3052 0.09002 0.08439 -0.0314 0.0446 1.0000
14.000 1.2758 0.09627 0.09093 -0.0340 0.0450 1.0000
14.250 1.2478 0.10317 0.09807 -0.0375 0.0453 1.0000
14.500 1.2194 0.11091 0.10602 -0.0422 0.0455 1.0000
14.750 1.1907 0.11961 0.11486 -0.0480 0.0458 1.0000
15.000 1.1605 0.12978 0.12518 -0.0553 0.0460 1.0000
15.250 1.1261 0.14254 0.13805 -0.0646 0.0465 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 428 AIRFOIL (goe428-il)