GOE 415 AIRFOIL (goe415-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 415 AIRFOIL (goe415-il) Reynolds number: 200,000 Max Cl/Cd: 68.93 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe415-il-200000.txt Download as CSV file: xf-goe415-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 415 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4264 0.10421 0.10092 -0.0089 1.0000 0.0247
-8.250 -0.4212 0.10059 0.09734 -0.0116 1.0000 0.0248
-8.000 -0.4221 0.09452 0.09134 -0.0131 1.0000 0.0252
-7.750 -0.4166 0.09029 0.08712 -0.0112 1.0000 0.0260
-7.500 -0.4104 0.08696 0.08382 -0.0115 1.0000 0.0266
-7.250 -0.4035 0.08356 0.08046 -0.0131 1.0000 0.0273
-7.000 -0.3948 0.07991 0.07683 -0.0159 1.0000 0.0280
-6.750 -0.3848 0.07605 0.07299 -0.0195 1.0000 0.0289
-6.500 -0.3732 0.07198 0.06893 -0.0235 1.0000 0.0299
-6.250 -0.3598 0.06776 0.06471 -0.0279 1.0000 0.0315
-6.000 -0.3413 0.06307 0.05993 -0.0339 1.0000 0.0340
-5.750 -0.3146 0.05889 0.05543 -0.0411 1.0000 0.0357
-5.500 -0.3010 0.05586 0.05216 -0.0422 1.0000 0.0360
-5.250 -0.2995 0.04874 0.04517 -0.0423 1.0000 0.0374
-5.000 -0.2900 0.04606 0.04253 -0.0413 1.0000 0.0387
-4.750 -0.2780 0.04348 0.03989 -0.0406 1.0000 0.0403
-4.500 -0.2634 0.04070 0.03698 -0.0404 1.0000 0.0425
-4.250 -0.2452 0.03788 0.03390 -0.0403 1.0000 0.0461
-4.000 -0.1737 0.01808 0.01406 -0.0487 0.9840 0.0544
-3.750 -0.1809 0.03099 0.02647 -0.0457 0.9938 0.0562
-3.500 -0.1369 0.02782 0.02293 -0.0500 0.9866 0.0683
-3.250 -0.0939 0.02552 0.02030 -0.0542 0.9788 0.0931
-2.750 0.0250 0.00632 0.00023 -0.0610 0.9363 0.0762
-2.500 0.0461 0.01809 0.01119 -0.0598 0.9517 0.0496
-2.250 0.0845 0.01552 0.00847 -0.0611 0.9376 0.0449
-2.000 0.1196 0.01430 0.00709 -0.0615 0.9184 0.0459
-1.750 0.1498 0.01354 0.00621 -0.0609 0.8944 0.0486
-1.500 0.1754 0.01259 0.00520 -0.0596 0.8647 0.0500
-1.250 0.1993 0.01176 0.00430 -0.0581 0.8283 0.0551
-1.000 0.2240 0.01141 0.00373 -0.0567 0.7834 0.0627
-0.750 0.2489 0.01110 0.00315 -0.0554 0.7404 0.0763
-0.500 0.2823 0.00832 0.00301 -0.0563 0.7040 1.0000
-0.250 0.3080 0.00856 0.00284 -0.0555 0.6722 1.0000
0.000 0.3339 0.00879 0.00275 -0.0548 0.6452 1.0000
0.250 0.3601 0.00900 0.00271 -0.0543 0.6214 1.0000
0.500 0.3864 0.00921 0.00269 -0.0539 0.6006 1.0000
0.750 0.4127 0.00943 0.00269 -0.0534 0.5811 1.0000
1.000 0.4391 0.00966 0.00272 -0.0530 0.5632 1.0000
1.250 0.4655 0.00990 0.00278 -0.0527 0.5466 1.0000
1.500 0.4921 0.01013 0.00287 -0.0523 0.5303 1.0000
1.750 0.5187 0.01036 0.00297 -0.0520 0.5148 1.0000
2.000 0.5453 0.01059 0.00309 -0.0517 0.4998 1.0000
2.250 0.5719 0.01082 0.00322 -0.0514 0.4851 1.0000
2.500 0.5985 0.01105 0.00336 -0.0511 0.4705 1.0000
2.750 0.6251 0.01128 0.00352 -0.0509 0.4565 1.0000
3.000 0.6516 0.01151 0.00367 -0.0506 0.4430 1.0000
3.250 0.6781 0.01173 0.00383 -0.0503 0.4294 1.0000
3.500 0.7045 0.01196 0.00399 -0.0500 0.4161 1.0000
3.750 0.7309 0.01220 0.00419 -0.0497 0.4037 1.0000
4.000 0.7573 0.01245 0.00439 -0.0495 0.3918 1.0000
4.250 0.7837 0.01269 0.00462 -0.0492 0.3802 1.0000
4.500 0.8101 0.01297 0.00490 -0.0490 0.3699 1.0000
4.750 0.8364 0.01327 0.00520 -0.0487 0.3603 1.0000
5.000 0.8626 0.01356 0.00548 -0.0485 0.3511 1.0000
5.250 0.8889 0.01386 0.00584 -0.0482 0.3422 1.0000
5.500 0.9149 0.01422 0.00617 -0.0480 0.3341 1.0000
5.750 0.9408 0.01447 0.00651 -0.0477 0.3237 1.0000
6.000 0.9659 0.01465 0.00670 -0.0473 0.3087 1.0000
6.250 0.9908 0.01486 0.00692 -0.0469 0.2937 1.0000
6.500 1.0160 0.01506 0.00719 -0.0465 0.2789 1.0000
6.750 1.0413 0.01524 0.00750 -0.0462 0.2615 1.0000
7.000 1.0663 0.01547 0.00779 -0.0458 0.2424 1.0000
7.250 1.0903 0.01582 0.00807 -0.0454 0.2075 1.0000
7.500 1.1114 0.01671 0.00861 -0.0447 0.1409 1.0000
7.750 1.1285 0.01832 0.00979 -0.0436 0.0739 1.0000
8.000 1.1427 0.02034 0.01148 -0.0421 0.0296 1.0000
8.250 1.1630 0.02138 0.01276 -0.0411 0.0260 1.0000
8.500 1.1816 0.02260 0.01419 -0.0399 0.0237 1.0000
8.750 1.1979 0.02400 0.01582 -0.0386 0.0225 1.0000
9.000 1.2108 0.02565 0.01768 -0.0369 0.0217 1.0000
9.250 1.2200 0.02753 0.01971 -0.0349 0.0212 1.0000
9.500 1.2267 0.02952 0.02182 -0.0327 0.0209 1.0000
9.750 1.2330 0.03154 0.02393 -0.0304 0.0208 1.0000
10.000 1.2379 0.03378 0.02623 -0.0279 0.0206 1.0000
10.250 1.2453 0.03580 0.02838 -0.0258 0.0201 1.0000
10.500 1.2539 0.03745 0.03018 -0.0240 0.0196 1.0000
10.750 1.2636 0.03963 0.03248 -0.0222 0.0194 1.0000
11.000 1.2746 0.04221 0.03520 -0.0206 0.0196 1.0000
11.250 1.2867 0.04541 0.03857 -0.0191 0.0199 1.0000
11.500 1.3080 0.04889 0.04219 -0.0184 0.0207 1.0000
11.750 1.3115 0.05024 0.04374 -0.0165 0.0212 1.0000
12.000 1.3094 0.05211 0.04593 -0.0147 0.0222 1.0000
12.250 1.3012 0.05592 0.05024 -0.0134 0.0243 1.0000
12.500 1.2917 0.06054 0.05518 -0.0129 0.0260 1.0000
12.750 1.2805 0.06572 0.06061 -0.0130 0.0273 1.0000
13.000 1.2798 0.06958 0.06491 -0.0130 0.0321 1.0000
13.250 1.2562 0.07576 0.07139 -0.0159 0.0329 1.0000
13.500 1.2343 0.08236 0.07824 -0.0195 0.0334 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 415 AIRFOIL (goe415-il)