GOE 492 AIRFOIL (goe492-il) Xfoil prediction polar at RE=500,000 Ncrit=9
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Airfoil: GOE 492 AIRFOIL (goe492-il) Reynolds number: 500,000 Max Cl/Cd: 111.96 at α=2.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe492-il-500000.txt Download as CSV file: xf-goe492-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 492 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
------ -------- --------- --------- -------- -------- --------
-14.500 -0.3772 0.15657 0.15427 -0.0173 1.0000 0.0097
-14.250 -0.3760 0.15490 0.15260 -0.0187 1.0000 0.0098
-12.250 -0.4594 0.14014 0.13768 -0.0181 1.0000 0.0104
-7.250 -0.3562 0.06295 0.06106 -0.0360 0.9952 0.0166
-7.000 -0.3457 0.05684 0.05494 -0.0421 0.9913 0.0172
-6.750 -0.3293 0.05265 0.05074 -0.0465 0.9879 0.0180
-6.500 -0.3097 0.04725 0.04531 -0.0537 0.9850 0.0187
-6.250 -0.2857 0.04128 0.03928 -0.0623 0.9829 0.0203
-6.000 -0.2641 0.03498 0.03288 -0.0705 0.9779 0.0223
-5.000 -0.1769 0.01915 0.01472 -0.0935 0.9705 0.0156
-4.750 -0.1429 0.01591 0.01092 -0.0952 0.9693 0.0159
-4.500 -0.1199 0.01293 0.00742 -0.0945 0.9642 0.0175
-4.250 -0.0874 0.01218 0.00658 -0.0956 0.9619 0.0200
-4.000 -0.0525 0.01186 0.00618 -0.0971 0.9601 0.0236
-3.750 -0.0188 0.01075 0.00495 -0.0984 0.9584 0.0285
-3.500 0.0164 0.01051 0.00466 -0.1000 0.9568 0.0352
-3.250 0.0454 0.00998 0.00408 -0.1004 0.9530 0.0425
-3.000 0.0749 0.00969 0.00374 -0.1007 0.9487 0.0475
-2.750 0.1069 0.00931 0.00330 -0.1016 0.9455 0.0514
-2.500 0.1392 0.00882 0.00279 -0.1026 0.9427 0.0578
-2.250 0.1652 0.00860 0.00251 -0.1021 0.9364 0.0628
-2.000 0.1950 0.00825 0.00216 -0.1024 0.9317 0.0714
-1.750 0.2231 0.00787 0.00191 -0.1025 0.9262 0.1092
-1.500 0.2494 0.00742 0.00189 -0.1023 0.9192 0.2603
-1.250 0.2777 0.00733 0.00181 -0.1022 0.9130 0.2841
-1.000 0.3046 0.00724 0.00172 -0.1019 0.9047 0.3044
-0.750 0.3313 0.00714 0.00163 -0.1015 0.8960 0.3217
-0.250 0.3845 0.00686 0.00148 -0.1008 0.8784 0.3627
0.000 0.4086 0.00627 0.00151 -0.1003 0.8697 0.5664
0.250 0.4477 0.00524 0.00151 -0.1025 0.8625 1.0000
0.500 0.4736 0.00526 0.00149 -0.1019 0.8514 1.0000
0.750 0.4995 0.00529 0.00146 -0.1014 0.8385 1.0000
1.000 0.5254 0.00533 0.00146 -0.1008 0.8244 1.0000
1.250 0.5514 0.00538 0.00146 -0.1002 0.8099 1.0000
1.500 0.5773 0.00545 0.00148 -0.0997 0.7948 1.0000
1.750 0.6028 0.00553 0.00151 -0.0991 0.7756 1.0000
2.000 0.6277 0.00564 0.00154 -0.0983 0.7501 1.0000
2.250 0.6516 0.00582 0.00161 -0.0973 0.7157 1.0000
2.500 0.6723 0.00619 0.00168 -0.0956 0.6569 1.0000
2.750 0.6947 0.00652 0.00183 -0.0944 0.6140 1.0000
3.000 0.7168 0.00689 0.00201 -0.0932 0.5624 1.0000
3.250 0.7314 0.00787 0.00230 -0.0907 0.4104 1.0000
3.500 0.7334 0.01070 0.00341 -0.0868 0.0430 1.0000
3.750 0.7558 0.01140 0.00414 -0.0856 0.0249 1.0000
4.000 0.7795 0.01187 0.00467 -0.0847 0.0215 1.0000
4.250 0.8022 0.01247 0.00530 -0.0837 0.0183 1.0000
4.500 0.8198 0.01372 0.00667 -0.0817 0.0164 1.0000
4.750 0.8387 0.01488 0.00793 -0.0799 0.0155 1.0000
5.000 0.8597 0.01598 0.00909 -0.0784 0.0150 1.0000
5.250 0.8813 0.01754 0.01077 -0.0770 0.0150 1.0000
5.500 0.9055 0.01872 0.01203 -0.0761 0.0142 1.0000
5.750 0.9301 0.01996 0.01337 -0.0753 0.0132 1.0000
6.000 0.9557 0.02213 0.01573 -0.0744 0.0131 1.0000
8.500 1.0389 0.04931 0.04626 -0.0474 0.0135 1.0000
8.750 1.0335 0.05359 0.05075 -0.0446 0.0131 1.0000
9.000 1.0223 0.05773 0.05506 -0.0415 0.0129 1.0000
9.250 1.0027 0.06117 0.05864 -0.0376 0.0128 1.0000
9.500 0.9799 0.06500 0.06259 -0.0349 0.0128 1.0000
9.750 0.9545 0.06967 0.06738 -0.0337 0.0129 1.0000
10.000 0.9279 0.07520 0.07304 -0.0341 0.0132 1.0000
10.250 0.8999 0.08189 0.07984 -0.0364 0.0135 1.0000
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Polar data table (+)
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