GOE 287 AIRFOIL (goe287-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 287 AIRFOIL (goe287-il) Reynolds number: 1,000,000 Max Cl/Cd: 106.61 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe287-il-1000000-n5.txt Download as CSV file: xf-goe287-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 287 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 -0.4351 0.11332 0.11102 0.0078 0.7372 0.0051
-9.500 -0.4322 0.10905 0.10673 0.0058 0.7280 0.0053
-9.000 -0.4377 0.09761 0.09525 0.0003 0.7097 0.0060
-8.750 -0.4314 0.09446 0.09206 -0.0017 0.6967 0.0062
-8.500 -0.4247 0.09139 0.08895 -0.0037 0.6825 0.0063
-8.250 -0.4182 0.08825 0.08576 -0.0060 0.6682 0.0065
-8.000 -0.4123 0.08514 0.08262 -0.0085 0.6544 0.0067
-7.750 -0.4019 0.08111 0.07855 -0.0131 0.6418 0.0070
-7.500 -0.3932 0.07397 0.07135 -0.0216 0.6321 0.0078
-7.250 -0.3793 0.06614 0.06345 -0.0312 0.6237 0.0085
-7.000 -0.3585 0.06279 0.06002 -0.0354 0.6134 0.0087
-6.750 -0.3365 0.05924 0.05639 -0.0396 0.6044 0.0091
-6.500 -0.3130 0.05468 0.05172 -0.0447 0.5971 0.0097
-6.250 -0.2862 0.04443 0.04122 -0.0542 0.5918 0.0115
-6.000 -0.2603 0.04231 0.03898 -0.0561 0.5847 0.0118
-5.750 -0.2312 0.02735 0.02328 -0.0637 0.5817 0.0152
-5.500 -0.2040 0.02677 0.02261 -0.0641 0.5751 0.0154
-5.250 -0.1767 0.02629 0.02204 -0.0645 0.5695 0.0158
-5.000 -0.1492 0.02490 0.02051 -0.0651 0.5645 0.0163
-4.750 -0.1212 0.01730 0.01206 -0.0664 0.5602 0.0196
-4.500 -0.0928 0.01662 0.01121 -0.0666 0.5545 0.0198
-4.250 -0.0649 0.01485 0.00915 -0.0672 0.5488 0.0206
-4.000 -0.0364 0.01443 0.00862 -0.0675 0.5436 0.0209
-3.750 -0.0077 0.01398 0.00807 -0.0677 0.5392 0.0212
-3.500 0.0210 0.01352 0.00751 -0.0680 0.5352 0.0216
-3.250 0.0497 0.01318 0.00709 -0.0682 0.5307 0.0221
-3.000 0.0785 0.01283 0.00663 -0.0685 0.5271 0.0226
-2.750 0.1074 0.01233 0.00604 -0.0687 0.5240 0.0231
-2.500 0.1363 0.01187 0.00549 -0.0689 0.5206 0.0236
-2.250 0.1652 0.01146 0.00499 -0.0691 0.5170 0.0240
-2.000 0.1941 0.01108 0.00453 -0.0693 0.5127 0.0243
-1.750 0.2230 0.01073 0.00411 -0.0696 0.5091 0.0246
-1.500 0.2520 0.01038 0.00371 -0.0698 0.5053 0.0248
-1.250 0.2809 0.01009 0.00337 -0.0700 0.5019 0.0250
-1.000 0.3098 0.00982 0.00306 -0.0702 0.4985 0.0251
-0.750 0.3388 0.00963 0.00283 -0.0704 0.4952 0.0255
-0.500 0.3678 0.00944 0.00262 -0.0707 0.4923 0.0258
-0.250 0.3968 0.00923 0.00239 -0.0709 0.4888 0.0259
0.000 0.4259 0.00907 0.00220 -0.0712 0.4845 0.0261
0.250 0.4549 0.00896 0.00205 -0.0714 0.4806 0.0263
0.500 0.4840 0.00885 0.00195 -0.0717 0.4772 0.0266
0.750 0.5130 0.00879 0.00188 -0.0719 0.4725 0.0269
1.000 0.5421 0.00865 0.00169 -0.0722 0.4670 0.0281
1.250 0.5711 0.00857 0.00160 -0.0724 0.4619 0.0293
1.500 0.6001 0.00856 0.00156 -0.0727 0.4532 0.0303
1.750 0.6289 0.00858 0.00154 -0.0730 0.4385 0.0311
2.000 0.6574 0.00869 0.00154 -0.0732 0.4095 0.0319
2.250 0.6847 0.00925 0.00174 -0.0736 0.3322 0.0326
2.500 0.7126 0.00955 0.00190 -0.0739 0.2999 0.0339
2.750 0.7408 0.00974 0.00202 -0.0741 0.2815 0.0358
3.250 0.7977 0.00984 0.00229 -0.0747 0.2605 0.1692
3.500 0.8203 0.00826 0.00254 -0.0742 0.2537 1.0000
3.750 0.8485 0.00843 0.00266 -0.0745 0.2472 1.0000
4.000 0.8768 0.00859 0.00279 -0.0747 0.2413 1.0000
4.250 0.9049 0.00877 0.00293 -0.0749 0.2357 1.0000
4.500 0.9330 0.00892 0.00307 -0.0751 0.2309 1.0000
4.750 0.9608 0.00915 0.00325 -0.0753 0.2213 1.0000
5.000 0.9886 0.00933 0.00341 -0.0755 0.2140 1.0000
5.250 1.0161 0.00959 0.00361 -0.0757 0.2037 1.0000
5.500 1.0437 0.00979 0.00378 -0.0759 0.1953 1.0000
5.750 1.0707 0.01009 0.00400 -0.0760 0.1789 1.0000
6.000 1.0934 0.01117 0.00468 -0.0759 0.1079 1.0000
6.250 1.1142 0.01248 0.00562 -0.0754 0.0315 1.0000
6.500 1.1397 0.01292 0.00602 -0.0753 0.0198 1.0000
6.750 1.1655 0.01328 0.00639 -0.0753 0.0152 1.0000
7.000 1.1912 0.01362 0.00675 -0.0752 0.0135 1.0000
7.250 1.2164 0.01403 0.00715 -0.0751 0.0115 1.0000
7.500 1.2413 0.01444 0.00758 -0.0749 0.0101 1.0000
7.750 1.2663 0.01482 0.00799 -0.0747 0.0094 1.0000
8.000 1.2907 0.01523 0.00843 -0.0745 0.0086 1.0000
8.250 1.3145 0.01571 0.00892 -0.0742 0.0078 1.0000
8.500 1.3373 0.01627 0.00952 -0.0738 0.0071 1.0000
8.750 1.3607 0.01672 0.01000 -0.0734 0.0067 1.0000
9.000 1.3834 0.01721 0.01053 -0.0730 0.0062 1.0000
9.250 1.4054 0.01775 0.01112 -0.0725 0.0058 1.0000
9.500 1.4266 0.01833 0.01172 -0.0719 0.0055 1.0000
9.750 1.4462 0.01902 0.01246 -0.0711 0.0051 1.0000
10.000 1.4637 0.01986 0.01336 -0.0700 0.0048 1.0000
10.250 1.4817 0.02056 0.01413 -0.0691 0.0047 1.0000
10.500 1.4980 0.02134 0.01498 -0.0679 0.0045 1.0000
10.750 1.5119 0.02219 0.01591 -0.0664 0.0043 1.0000
11.000 1.5213 0.02311 0.01690 -0.0643 0.0042 1.0000
11.250 1.5305 0.02416 0.01802 -0.0625 0.0040 1.0000
11.500 1.5403 0.02531 0.01923 -0.0611 0.0038 1.0000
11.750 1.5498 0.02659 0.02057 -0.0599 0.0037 1.0000
12.000 1.5579 0.02812 0.02217 -0.0589 0.0036 1.0000
12.250 1.5643 0.02993 0.02405 -0.0581 0.0034 1.0000
12.500 1.5673 0.03218 0.02639 -0.0574 0.0033 1.0000
12.750 1.5663 0.03495 0.02928 -0.0568 0.0032 1.0000
13.000 1.5670 0.03761 0.03204 -0.0563 0.0032 1.0000
13.250 1.5686 0.04026 0.03480 -0.0561 0.0031 1.0000
13.500 1.5681 0.04321 0.03786 -0.0559 0.0031 1.0000
13.750 1.5667 0.04630 0.04106 -0.0558 0.0030 1.0000
14.000 1.5641 0.04961 0.04447 -0.0558 0.0030 1.0000
14.250 1.5598 0.05314 0.04811 -0.0559 0.0030 1.0000
14.500 1.5540 0.05690 0.05198 -0.0561 0.0029 1.0000
14.750 1.5482 0.06076 0.05595 -0.0564 0.0029 1.0000
15.000 1.5411 0.06489 0.06018 -0.0569 0.0028 1.0000
15.250 1.5338 0.06916 0.06456 -0.0575 0.0028 1.0000
15.500 1.5264 0.07354 0.06905 -0.0582 0.0028 1.0000
15.750 1.5181 0.07811 0.07372 -0.0591 0.0027 1.0000
16.000 1.5091 0.08289 0.07861 -0.0601 0.0027 1.0000
16.250 1.5001 0.08780 0.08362 -0.0612 0.0027 1.0000
16.500 1.4911 0.09278 0.08871 -0.0624 0.0026 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 287 AIRFOIL (goe287-il)