GOE 393 AIRFOIL (goe393-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 393 AIRFOIL (goe393-il) Reynolds number: 500,000 Max Cl/Cd: 101.32 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe393-il-500000.txt Download as CSV file: xf-goe393-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 393 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4792 0.10111 0.09896 0.0176 1.0000 0.0221
-8.000 -0.4725 0.09793 0.09581 0.0158 1.0000 0.0228
-7.750 -0.4661 0.09459 0.09248 0.0136 1.0000 0.0239
-7.500 -0.4578 0.09050 0.08842 0.0052 1.0000 0.0252
-7.250 -0.4422 0.08604 0.08395 -0.0018 1.0000 0.0254
-7.000 -0.4255 0.08141 0.07931 -0.0073 1.0000 0.0255
-6.750 -0.4204 0.07620 0.07411 -0.0078 1.0000 0.0261
-6.500 -0.4056 0.07354 0.07145 -0.0086 1.0000 0.0266
-6.250 -0.3882 0.07060 0.06851 -0.0110 1.0000 0.0273
-6.000 -0.3683 0.06730 0.06520 -0.0144 1.0000 0.0285
-5.500 -0.3008 0.05760 0.05528 -0.0303 1.0000 0.0323
-5.250 -0.2882 0.05210 0.04977 -0.0325 1.0000 0.0335
-5.000 -0.2581 0.04970 0.04734 -0.0354 0.9694 0.0344
-4.750 -0.2330 0.04734 0.04471 -0.0369 0.8608 0.0360
-4.500 -0.1943 0.04400 0.04079 -0.0409 0.7847 0.0407
-4.250 -0.1745 0.03859 0.03509 -0.0433 0.7511 0.0422
-4.000 -0.1518 0.03724 0.03357 -0.0435 0.7256 0.0432
-3.750 -0.1264 0.03558 0.03175 -0.0442 0.7062 0.0448
-3.500 -0.0869 0.03302 0.02875 -0.0461 0.6915 0.0505
-3.250 -0.0589 0.02067 0.01548 -0.0488 0.6835 0.0409
-3.000 -0.0327 0.01518 0.00920 -0.0493 0.6721 0.0390
-2.750 -0.0051 0.01380 0.00748 -0.0491 0.6599 0.0395
-2.500 0.0226 0.01299 0.00643 -0.0489 0.6487 0.0402
-2.250 0.0503 0.01225 0.00545 -0.0486 0.6385 0.0405
-2.000 0.0781 0.01161 0.00467 -0.0484 0.6284 0.0409
-1.750 0.1056 0.01077 0.00365 -0.0481 0.6192 0.0418
-1.500 0.1331 0.01030 0.00308 -0.0478 0.6098 0.0429
-1.250 0.1609 0.00998 0.00272 -0.0476 0.6007 0.0443
-1.000 0.1886 0.00979 0.00248 -0.0474 0.5924 0.0462
-0.750 0.2166 0.00961 0.00227 -0.0472 0.5838 0.0487
-0.500 0.2445 0.00944 0.00203 -0.0469 0.5760 0.0517
-0.250 0.2725 0.00929 0.00192 -0.0467 0.5676 0.0598
0.000 0.3006 0.00931 0.00200 -0.0465 0.5595 0.0790
0.250 0.3287 0.00943 0.00209 -0.0464 0.5507 0.0908
0.500 0.3569 0.00953 0.00216 -0.0462 0.5424 0.0978
0.750 0.3847 0.00953 0.00213 -0.0461 0.5342 0.1032
1.000 0.4127 0.00955 0.00217 -0.0459 0.5257 0.1095
1.500 0.4683 0.00949 0.00208 -0.0456 0.5086 0.1169
1.750 0.4961 0.00946 0.00206 -0.0455 0.4998 0.1219
2.000 0.5238 0.00948 0.00206 -0.0453 0.4890 0.1269
2.250 0.5516 0.00941 0.00201 -0.0452 0.4777 0.1318
2.500 0.5794 0.00940 0.00201 -0.0450 0.4675 0.1365
2.750 0.6071 0.00943 0.00201 -0.0449 0.4568 0.1416
3.000 0.6348 0.00942 0.00204 -0.0448 0.4446 0.1503
3.250 0.6625 0.00942 0.00208 -0.0446 0.4325 0.1656
3.500 0.6904 0.00776 0.00222 -0.0449 0.4197 1.0000
3.750 0.7179 0.00790 0.00229 -0.0447 0.4072 1.0000
4.000 0.7453 0.00806 0.00238 -0.0445 0.3971 1.0000
4.250 0.7727 0.00820 0.00249 -0.0444 0.3871 1.0000
4.500 0.8001 0.00836 0.00263 -0.0442 0.3776 1.0000
4.750 0.8273 0.00855 0.00278 -0.0440 0.3686 1.0000
5.000 0.8546 0.00870 0.00294 -0.0438 0.3586 1.0000
5.250 0.8818 0.00887 0.00311 -0.0436 0.3496 1.0000
5.500 0.9088 0.00908 0.00331 -0.0435 0.3396 1.0000
5.750 0.9358 0.00928 0.00349 -0.0433 0.3258 1.0000
6.000 0.9625 0.00950 0.00368 -0.0431 0.3041 1.0000
6.250 0.9889 0.00981 0.00391 -0.0429 0.2767 1.0000
6.500 1.0142 0.01034 0.00422 -0.0427 0.2228 1.0000
6.750 1.0368 0.01144 0.00490 -0.0424 0.1500 1.0000
7.000 1.0614 0.01209 0.00544 -0.0421 0.1251 1.0000
7.250 1.0866 0.01260 0.00590 -0.0418 0.1086 1.0000
7.500 1.1117 0.01310 0.00637 -0.0416 0.0956 1.0000
7.750 1.1366 0.01359 0.00685 -0.0413 0.0832 1.0000
8.000 1.1606 0.01425 0.00739 -0.0409 0.0611 1.0000
8.250 1.1820 0.01537 0.00829 -0.0404 0.0340 1.0000
8.500 1.2049 0.01620 0.00918 -0.0398 0.0281 1.0000
8.750 1.2259 0.01730 0.01034 -0.0391 0.0244 1.0000
9.000 1.2487 0.01801 0.01114 -0.0385 0.0226 1.0000
9.250 1.2704 0.01884 0.01203 -0.0379 0.0209 1.0000
9.500 1.2872 0.02033 0.01360 -0.0369 0.0191 1.0000
9.750 1.3072 0.02126 0.01464 -0.0361 0.0182 1.0000
10.000 1.3265 0.02225 0.01572 -0.0353 0.0173 1.0000
10.250 1.3444 0.02333 0.01689 -0.0343 0.0165 1.0000
10.500 1.3610 0.02449 0.01813 -0.0333 0.0158 1.0000
10.750 1.3751 0.02586 0.01957 -0.0321 0.0152 1.0000
11.000 1.3829 0.02785 0.02165 -0.0304 0.0146 1.0000
11.250 1.3877 0.03013 0.02408 -0.0285 0.0142 1.0000
11.500 1.3984 0.03130 0.02538 -0.0270 0.0138 1.0000
11.750 1.4054 0.03279 0.02700 -0.0255 0.0134 1.0000
12.000 1.4104 0.03461 0.02894 -0.0243 0.0131 1.0000
12.250 1.4141 0.03679 0.03125 -0.0236 0.0127 1.0000
12.500 1.4171 0.03925 0.03383 -0.0234 0.0125 1.0000
12.750 1.4194 0.04195 0.03667 -0.0236 0.0122 1.0000
13.000 1.4207 0.04487 0.03971 -0.0240 0.0120 1.0000
13.250 1.4209 0.04798 0.04294 -0.0245 0.0118 1.0000
13.500 1.4198 0.05128 0.04637 -0.0251 0.0116 1.0000
13.750 1.4171 0.05480 0.05001 -0.0258 0.0114 1.0000
14.000 1.4127 0.05861 0.05394 -0.0266 0.0113 1.0000
14.250 1.4063 0.06270 0.05815 -0.0275 0.0112 1.0000
14.500 1.3978 0.06716 0.06275 -0.0286 0.0111 1.0000
14.750 1.3870 0.07203 0.06777 -0.0299 0.0110 1.0000
15.000 1.3741 0.07737 0.07326 -0.0316 0.0109 1.0000
15.250 1.3600 0.08300 0.07906 -0.0337 0.0109 1.0000
15.500 1.3446 0.08916 0.08538 -0.0363 0.0109 1.0000
15.750 1.3289 0.09573 0.09212 -0.0394 0.0108 1.0000
16.000 1.3132 0.10263 0.09918 -0.0429 0.0108 1.0000
16.250 1.2982 0.10973 0.10644 -0.0467 0.0109 1.0000
16.500 1.2827 0.11721 0.11407 -0.0509 0.0109 1.0000
16.750 1.2663 0.12522 0.12223 -0.0556 0.0109 1.0000
17.000 1.2493 0.13378 0.13095 -0.0607 0.0110 1.0000
17.250 1.2290 0.14379 0.14114 -0.0669 0.0111 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 393 AIRFOIL (goe393-il)