GOE 55 AIRFOIL (goe55-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 55 AIRFOIL (goe55-il) Reynolds number: 100,000 Max Cl/Cd: 43.3 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe55-il-100000-n5.txt Download as CSV file: xf-goe55-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 55 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 -0.6092 0.10560 0.10055 0.0330 1.0000 0.0615
-8.000 -0.6048 0.10207 0.09705 0.0312 1.0000 0.0630
-7.750 -0.6016 0.09849 0.09351 0.0285 1.0000 0.0647
-7.500 -0.5958 0.09363 0.08863 0.0145 1.0000 0.0678
-7.000 -0.5747 0.08457 0.07959 0.0113 1.0000 0.0697
-6.750 -0.5619 0.08110 0.07612 0.0110 1.0000 0.0710
-6.250 -0.5181 0.06646 0.06108 -0.0067 1.0000 0.0563
-6.000 -0.4975 0.06171 0.05621 -0.0101 1.0000 0.0560
-5.750 -0.4755 0.05715 0.05152 -0.0132 1.0000 0.0557
-5.500 -0.4530 0.05300 0.04724 -0.0156 1.0000 0.0544
-5.250 -0.4268 0.04832 0.04234 -0.0189 1.0000 0.0530
-5.000 -0.3981 0.04351 0.03721 -0.0221 1.0000 0.0520
-4.750 -0.3687 0.03949 0.03285 -0.0247 1.0000 0.0529
-4.500 -0.3384 0.03571 0.02868 -0.0268 1.0000 0.0538
-4.250 -0.3081 0.03226 0.02483 -0.0283 1.0000 0.0535
-4.000 -0.2777 0.02931 0.02148 -0.0295 1.0000 0.0535
-3.750 -0.2474 0.02680 0.01856 -0.0303 1.0000 0.0538
-3.500 -0.2172 0.02471 0.01608 -0.0309 1.0000 0.0549
-3.250 -0.1869 0.02305 0.01398 -0.0312 1.0000 0.0567
-3.000 -0.1578 0.02150 0.01223 -0.0314 1.0000 0.0576
-2.750 -0.1291 0.02024 0.01087 -0.0315 1.0000 0.0584
-2.500 -0.1004 0.01920 0.00977 -0.0315 1.0000 0.0595
-2.250 -0.0719 0.01828 0.00881 -0.0314 1.0000 0.0609
-2.000 -0.0434 0.01744 0.00795 -0.0313 1.0000 0.0627
-1.750 -0.0151 0.01677 0.00727 -0.0311 1.0000 0.0660
-1.500 0.0132 0.01610 0.00664 -0.0309 1.0000 0.0693
-1.250 0.0520 0.01563 0.00616 -0.0327 0.8377 0.0729
-1.000 0.0734 0.01594 0.00577 -0.0300 0.6411 0.0766
-0.750 0.0989 0.01599 0.00536 -0.0290 0.5431 0.0816
-0.500 0.1262 0.01598 0.00511 -0.0287 0.4977 0.0912
-0.250 0.1542 0.01585 0.00484 -0.0285 0.4711 0.1053
0.000 0.1823 0.01559 0.00458 -0.0283 0.4531 0.1275
0.250 0.2106 0.01530 0.00449 -0.0283 0.4398 0.1852
0.500 0.2306 0.01297 0.00433 -0.0261 0.4305 1.0000
0.750 0.2591 0.01322 0.00432 -0.0259 0.4209 1.0000
1.000 0.2876 0.01347 0.00437 -0.0257 0.4123 1.0000
1.250 0.3159 0.01373 0.00447 -0.0255 0.4044 1.0000
1.500 0.3443 0.01399 0.00460 -0.0254 0.3957 1.0000
1.750 0.3724 0.01427 0.00471 -0.0252 0.3839 1.0000
2.000 0.4005 0.01452 0.00483 -0.0250 0.3708 1.0000
2.250 0.4287 0.01475 0.00501 -0.0249 0.3595 1.0000
2.500 0.4568 0.01503 0.00520 -0.0247 0.3512 1.0000
2.750 0.4851 0.01528 0.00544 -0.0246 0.3423 1.0000
3.000 0.5132 0.01557 0.00570 -0.0244 0.3346 1.0000
3.250 0.5413 0.01582 0.00595 -0.0243 0.3242 1.0000
3.500 0.5694 0.01602 0.00617 -0.0241 0.3111 1.0000
3.750 0.5973 0.01620 0.00636 -0.0240 0.2961 1.0000
4.000 0.6253 0.01634 0.00653 -0.0238 0.2794 1.0000
4.250 0.6532 0.01647 0.00669 -0.0237 0.2611 1.0000
4.500 0.6813 0.01663 0.00690 -0.0236 0.2416 1.0000
4.750 0.7092 0.01683 0.00713 -0.0234 0.2180 1.0000
5.000 0.7367 0.01713 0.00734 -0.0233 0.1900 1.0000
5.250 0.7638 0.01764 0.00771 -0.0233 0.1621 1.0000
5.500 0.7908 0.01829 0.00828 -0.0233 0.1392 1.0000
5.750 0.8175 0.01906 0.00899 -0.0232 0.1156 1.0000
6.000 0.8442 0.01990 0.00982 -0.0232 0.0915 1.0000
6.250 0.8705 0.02082 0.01068 -0.0232 0.0754 1.0000
6.500 0.8968 0.02169 0.01160 -0.0230 0.0679 1.0000
6.750 0.9223 0.02268 0.01258 -0.0229 0.0636 1.0000
7.000 0.9478 0.02360 0.01366 -0.0227 0.0609 1.0000
7.250 0.9728 0.02458 0.01477 -0.0224 0.0586 1.0000
7.500 0.9972 0.02564 0.01592 -0.0221 0.0568 1.0000
7.750 1.0209 0.02681 0.01714 -0.0218 0.0550 1.0000
8.000 1.0439 0.02810 0.01853 -0.0214 0.0533 1.0000
8.250 1.0672 0.02937 0.01999 -0.0209 0.0517 1.0000
8.500 1.0898 0.03079 0.02159 -0.0204 0.0505 1.0000
8.750 1.1117 0.03236 0.02333 -0.0199 0.0495 1.0000
9.000 1.1330 0.03406 0.02524 -0.0194 0.0486 1.0000
9.250 1.1534 0.03588 0.02724 -0.0189 0.0477 1.0000
9.500 1.1727 0.03775 0.02926 -0.0184 0.0466 1.0000
9.750 1.1907 0.03992 0.03151 -0.0180 0.0454 1.0000
10.000 1.2060 0.04255 0.03444 -0.0175 0.0445 1.0000
10.250 1.2186 0.04546 0.03780 -0.0170 0.0439 1.0000
10.500 1.2269 0.04883 0.04161 -0.0166 0.0434 1.0000
10.750 1.2295 0.05267 0.04590 -0.0165 0.0428 1.0000
11.000 1.2250 0.05701 0.05065 -0.0167 0.0423 1.0000
11.250 1.2111 0.06198 0.05598 -0.0177 0.0419 1.0000
11.500 1.1843 0.06846 0.06276 -0.0213 0.0421 1.0000
11.750 1.1455 0.08096 0.07554 -0.0336 0.0428 1.0000
12.000 1.0909 0.09814 0.09287 -0.0474 0.0437 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 55 AIRFOIL (goe55-il)