GOE 553 AIRFOIL (goe553-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 553 AIRFOIL (goe553-il) Reynolds number: 200,000 Max Cl/Cd: 74.95 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe553-il-200000-n5.txt Download as CSV file: xf-goe553-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 553 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.2824 0.07358 0.07006 -0.0615 0.9632 0.0340
-8.500 -0.3102 0.04256 0.03826 -0.1016 0.9002 0.0341
-8.250 -0.2939 0.03810 0.03340 -0.1059 0.8765 0.0343
-8.000 -0.2764 0.03500 0.02994 -0.1079 0.8514 0.0346
-7.750 -0.2582 0.03240 0.02697 -0.1088 0.8267 0.0350
-7.500 -0.2386 0.03044 0.02467 -0.1091 0.8032 0.0356
-7.250 -0.2191 0.02851 0.02234 -0.1090 0.7814 0.0364
-7.000 -0.1988 0.02661 0.02002 -0.1088 0.7624 0.0370
-6.750 -0.1772 0.02493 0.01792 -0.1085 0.7458 0.0374
-6.500 -0.1543 0.02353 0.01613 -0.1081 0.7315 0.0378
-6.250 -0.1304 0.02236 0.01461 -0.1076 0.7186 0.0382
-6.000 -0.1056 0.02139 0.01333 -0.1073 0.7071 0.0386
-5.750 -0.0806 0.02051 0.01221 -0.1069 0.6972 0.0393
-5.500 -0.0552 0.01975 0.01138 -0.1066 0.6875 0.0400
-5.250 -0.0294 0.01913 0.01063 -0.1063 0.6792 0.0407
-5.000 -0.0033 0.01850 0.00988 -0.1059 0.6707 0.0412
-4.750 0.0229 0.01793 0.00918 -0.1055 0.6634 0.0417
-4.500 0.0492 0.01739 0.00855 -0.1052 0.6556 0.0424
-4.250 0.0755 0.01692 0.00797 -0.1048 0.6486 0.0431
-4.000 0.1021 0.01647 0.00746 -0.1045 0.6416 0.0438
-3.750 0.1286 0.01610 0.00700 -0.1041 0.6345 0.0446
-3.500 0.1550 0.01574 0.00656 -0.1038 0.6284 0.0457
-3.250 0.1814 0.01533 0.00619 -0.1035 0.6215 0.0471
-3.000 0.2081 0.01503 0.00585 -0.1032 0.6150 0.0484
-2.750 0.2349 0.01478 0.00554 -0.1030 0.6093 0.0496
-2.500 0.2620 0.01453 0.00526 -0.1027 0.6026 0.0511
-2.250 0.2892 0.01434 0.00500 -0.1025 0.5964 0.0527
-2.000 0.3163 0.01412 0.00474 -0.1023 0.5909 0.0552
-1.750 0.3436 0.01395 0.00455 -0.1021 0.5841 0.0593
-1.500 0.3708 0.01377 0.00437 -0.1019 0.5778 0.0659
-1.250 0.3980 0.01357 0.00426 -0.1017 0.5718 0.0867
-1.000 0.4250 0.01328 0.00415 -0.1016 0.5652 0.1411
-0.750 0.4505 0.01263 0.00406 -0.1016 0.5591 0.3164
-0.500 0.4759 0.01231 0.00420 -0.1012 0.5526 0.4692
-0.250 0.5023 0.01227 0.00428 -0.1007 0.5454 0.5283
0.000 0.5285 0.01228 0.00434 -0.1001 0.5390 0.5728
0.250 0.5543 0.01227 0.00447 -0.0995 0.5315 0.6209
0.500 0.5800 0.01229 0.00453 -0.0988 0.5248 0.6547
0.750 0.6067 0.01231 0.00456 -0.0983 0.5178 0.6713
1.000 0.6333 0.01235 0.00458 -0.0979 0.5100 0.6862
1.250 0.6597 0.01239 0.00460 -0.0974 0.5021 0.7024
1.500 0.6858 0.01241 0.00463 -0.0970 0.4930 0.7198
1.750 0.7115 0.01244 0.00467 -0.0964 0.4844 0.7390
2.000 0.7367 0.01247 0.00472 -0.0957 0.4761 0.7603
2.250 0.7623 0.01250 0.00478 -0.0950 0.4683 0.7862
2.500 0.7879 0.01249 0.00484 -0.0944 0.4605 0.8245
2.750 0.8249 0.01241 0.00487 -0.0961 0.4516 1.0000
3.000 0.8512 0.01260 0.00497 -0.0958 0.4427 1.0000
3.250 0.8775 0.01280 0.00509 -0.0956 0.4342 1.0000
3.500 0.9035 0.01301 0.00522 -0.0952 0.4265 1.0000
3.750 0.9297 0.01323 0.00539 -0.0950 0.4204 1.0000
4.000 0.9558 0.01344 0.00557 -0.0947 0.4139 1.0000
4.250 0.9813 0.01370 0.00575 -0.0943 0.4080 1.0000
4.500 1.0071 0.01391 0.00597 -0.0939 0.4012 1.0000
4.750 1.0322 0.01417 0.00617 -0.0935 0.3942 1.0000
5.000 1.0573 0.01443 0.00640 -0.0930 0.3882 1.0000
5.250 1.0827 0.01467 0.00665 -0.0926 0.3824 1.0000
5.500 1.1071 0.01495 0.00690 -0.0921 0.3762 1.0000
5.750 1.1317 0.01522 0.00717 -0.0916 0.3700 1.0000
6.000 1.1562 0.01549 0.00745 -0.0911 0.3640 1.0000
6.250 1.1799 0.01581 0.00773 -0.0904 0.3586 1.0000
6.500 1.2041 0.01608 0.00806 -0.0899 0.3530 1.0000
6.750 1.2277 0.01638 0.00838 -0.0892 0.3470 1.0000
7.000 1.2504 0.01673 0.00870 -0.0885 0.3416 1.0000
7.250 1.2738 0.01703 0.00907 -0.0878 0.3361 1.0000
7.500 1.2964 0.01735 0.00943 -0.0871 0.3304 1.0000
7.750 1.3179 0.01773 0.00980 -0.0861 0.3251 1.0000
8.000 1.3399 0.01806 0.01021 -0.0853 0.3191 1.0000
8.250 1.3602 0.01844 0.01062 -0.0842 0.3119 1.0000
8.500 1.3794 0.01884 0.01104 -0.0829 0.3040 1.0000
8.750 1.3973 0.01927 0.01149 -0.0815 0.2950 1.0000
9.000 1.4151 0.01970 0.01197 -0.0801 0.2867 1.0000
9.250 1.4295 0.02019 0.01246 -0.0781 0.2779 1.0000
9.500 1.4442 0.02066 0.01299 -0.0762 0.2691 1.0000
9.750 1.4557 0.02125 0.01359 -0.0739 0.2603 1.0000
10.000 1.4687 0.02184 0.01423 -0.0719 0.2512 1.0000
10.250 1.4798 0.02252 0.01495 -0.0698 0.2429 1.0000
10.500 1.4892 0.02332 0.01576 -0.0676 0.2341 1.0000
10.750 1.4989 0.02416 0.01665 -0.0656 0.2254 1.0000
11.000 1.5058 0.02519 0.01769 -0.0634 0.2177 1.0000
11.250 1.5130 0.02628 0.01883 -0.0615 0.2094 1.0000
11.500 1.5180 0.02758 0.02014 -0.0595 0.2023 1.0000
11.750 1.5230 0.02897 0.02158 -0.0578 0.1955 1.0000
12.000 1.5261 0.03058 0.02321 -0.0562 0.1899 1.0000
12.250 1.5315 0.03214 0.02485 -0.0549 0.1846 1.0000
12.500 1.5342 0.03398 0.02675 -0.0537 0.1799 1.0000
12.750 1.5362 0.03599 0.02880 -0.0526 0.1756 1.0000
13.000 1.5413 0.03782 0.03074 -0.0518 0.1711 1.0000
13.250 1.5422 0.04010 0.03308 -0.0511 0.1661 1.0000
13.500 1.5401 0.04273 0.03575 -0.0504 0.1611 1.0000
13.750 1.5440 0.04490 0.03805 -0.0500 0.1559 1.0000
14.000 1.5422 0.04768 0.04089 -0.0497 0.1503 1.0000
14.250 1.5403 0.05056 0.04386 -0.0495 0.1440 1.0000
14.500 1.5375 0.05365 0.04704 -0.0495 0.1355 1.0000
14.750 1.5341 0.05692 0.05039 -0.0496 0.1245 1.0000
15.000 1.5250 0.06098 0.05446 -0.0501 0.1084 1.0000
15.250 1.5083 0.06616 0.05958 -0.0509 0.0940 1.0000
15.500 1.4914 0.07159 0.06498 -0.0520 0.0858 1.0000
15.750 1.4767 0.07692 0.07033 -0.0533 0.0805 1.0000
16.000 1.4649 0.08201 0.07548 -0.0546 0.0766 1.0000
16.250 1.4555 0.08689 0.08045 -0.0559 0.0733 1.0000
16.500 1.4445 0.09210 0.08574 -0.0575 0.0706 1.0000
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Polar data table (+)
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