! ------------ AirfoilInfo v1.01.x Input File ---------------------------------- ! DU30 airfoil with an aspect ratio of 17. Original -180 to 180deg Cl, Cd, and Cm versus AOA data taken from Appendix A of DOWEC document 10046_009.pdf (numerical values obtained from Koert Lindenburg of ECN). ! Cl and Cd values corrected for rotational stall delay and Cd values corrected using the Viterna method for 0 to 90deg AOA by Jason Jonkman using AirfoilPrep_v2p0.xls. ! note that this file uses Marshall Buhl's new input file processing; start all comment lines with ! ! ------------------------------------------------------------------------------ "DEFAULT" InterpOrd ! Interpolation order to use for quasi-steady table lookup {1=linear; 3=cubic spline; "default"} [default=1] 0.30 RelThickness ! The non-dimensional thickness of the airfoil (thickness/chord) [only used if UAMod=7] [default=0.2] (-) 1 NonDimArea ! The non-dimensional area of the airfoil (area/chord^2) (set to 1.0 if unsure or unneeded) @"DU30_A17_coords.txt" NumCoords ! The number of coordinates in the airfoil shape file. Set to zero if coordinates not included. "unused" BL_file ! The file name including the boundary layer characteristics of the profile. Unused if the aeroacoustic module is not called. 1 NumTabs ! Number of airfoil tables in this file. ! ------------------------------------------------------------------------------ ! data for table 1 ! ------------------------------------------------------------------------------ 0.75 Re ! Reynolds number in millions 0 UserProp ! User property (control) setting True InclUAdata ! Is unsteady aerodynamics data included in this table? If TRUE, then include UA coefficients below this line. Any missing coefficients will be calculated internally. !........................................ > start of UA coefficients included only when InclUAdata=TRUE !Note: If any of the following lines are missing, they will be calculated internally (they must be read in the following order, though) -2.2 alpha0 ! 0-lift angle of attack, depends on airfoil. ! THIS IS AN OPTIONAL LINE; if omitted, it will be calculated from the polar data 9 alpha1 ! Angle of attack at f=0.7, (approximately the stall angle) for AOA>alpha0. (deg) [unused when UAMod=4|5] ! THIS IS AN OPTIONAL LINE; if omitted, it will be calculated from the polar data -9 alpha2 ! Angle of attack at f=0.7, (approximately the stall angle) for AOA1] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to 0 0 S2 ! Constant in the f curve best-fit for AOA> alpha1; by definition it depends on the airfoil. [unused if UAMod<>1] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to 0 0 S3 ! Constant in the f curve best-fit for alpha2<=AOA< alpha0; by definition it depends on the airfoil. [unused if UAMod<>1] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to 0 0 S4 ! Constant in the f curve best-fit for AOA< alpha2; by definition it depends on the airfoil. [unused if UAMod<>1] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to 0 1.449 Cn1 ! Critical value of C0n at leading edge separation. It should be extracted from airfoil data at a given Mach and Reynolds number. [unused when UAMod=4] ! THIS IS AN OPTIONAL LINE; if omitted, it will be calculated from the polar data -0.6138 Cn2 ! As Cn1 for negative AOAs. [unused when UAMod=4] ! THIS IS AN OPTIONAL LINE; if omitted, it will be calculated from the polar data 0.19 St_sh ! Strouhal's shedding frequency constant. [default = 0.19; unused when UAMod = 4|5] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to its default value 0.008 Cd0 ! 2D drag coefficient value at 0-lift. ! THIS IS AN OPTIONAL LINE; if omitted, it will be calculated from the polar data -0.09 Cm0 ! 2D pitching moment coefficient about 1/4-chord location, at 0-lift, positive if nose up. [unused when UAMod = 4|5] ! THIS IS AN OPTIONAL LINE; if omitted, it will be calculated from the polar data 0 k0 ! Constant in the \hat(x)_cp curve best-fit; = (\hat(x)_AC-0.25). [unused if UAMod<>1] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to 0 0 k1 ! Constant in the \hat(x)_cp curve best-fit. [unused if UAMod<>1] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to 0 0 k2 ! Constant in the \hat(x)_cp curve best-fit. [unused if UAMod<>1] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to 0 0 k3 ! Constant in the \hat(x)_cp curve best-fit. [unused if UAMod<>1] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to 0 0 k1_hat ! Constant in the expression of Cc due to leading edge vortex effects. [unused if UAMod<>1] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to 0 0.2 x_cp_bar ! Constant in the expression of \hat(x)_cp^v. [unused if UAMod<>1, default = 0.2] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to its default value "DEFAULT" UACutout ! Angle of attack above which unsteady aerodynamics are disabled (deg). [Default = 45 degrees] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to its default value "DEFAULT" UACutout_delta ! Delta angle of attack below UACutout where unsteady aerodynamics begin to turn off/blend with steady solution (deg) [Default = 5 degrees] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to its default value "DEFAULT" filtCutOff ! Reduced frequency cut-off for low-pass filtering the AoA input to UA, as well as the 1st and 2nd derivatives (-) [default = 0.5; unused when UAMod = 4|5] ! THIS IS AN OPTIONAL LINE; if omitted, it will be set to its default value !........................................ < end of UA coefficients included only when InclUAdata=TRUE ! Table of aerodynamics coefficients 143 NumAlf ! Number of data lines in the following table ! Alpha Cl Cd Cm ! (deg) (-) (-) (-) -180.00 0.000 0.0267 0.0000 -175.00 0.274 0.0370 0.1379 -170.00 0.547 0.0968 0.2778 -160.00 0.685 0.2876 0.2740 -155.00 0.766 0.4025 0.3118 -150.00 0.816 0.5232 0.3411 -145.00 0.836 0.6454 0.3631 -140.00 0.832 0.7656 0.3791 -135.00 0.804 0.8807 0.3899 -130.00 0.756 0.9882 0.3965 -125.00 0.690 1.0861 0.3994 -120.00 0.609 1.1730 0.3992 -115.00 0.515 1.2474 0.3964 -110.00 0.411 1.3084 0.3915 -105.00 0.300 1.3552 0.3846 -100.00 0.182 1.3875 0.3761 -95.00 0.061 1.4048 0.3663 -90.00 -0.061 1.4070 0.3551 -85.00 -0.183 1.3941 0.3428 -80.00 -0.302 1.3664 0.3295 -75.00 -0.416 1.3240 0.3153 -70.00 -0.523 1.2676 0.3001 -65.00 -0.622 1.1978 0.2841 -60.00 -0.708 1.1156 0.2672 -55.00 -0.781 1.0220 0.2494 -50.00 -0.838 0.9187 0.2308 -45.00 -0.877 0.8074 0.2113 -40.00 -0.895 0.6904 0.1909 -35.00 -0.889 0.5703 0.1696 -30.00 -0.858 0.4503 0.1475 -25.00 -0.832 0.3357 0.1224 -24.00 -0.852 0.3147 0.1156 -23.00 -0.882 0.2946 0.1081 -22.00 -0.919 0.2752 0.1000 -21.00 -0.963 0.2566 0.0914 -20.00 -1.013 0.2388 0.0823 -19.00 -1.067 0.2218 0.0728 -18.00 -1.125 0.2056 0.0631 -17.00 -1.185 0.1901 0.0531 -16.00 -1.245 0.1754 0.0430 -15.25 -1.290 0.1649 0.0353 -14.24 -1.229 0.1461 0.0240 -13.24 -1.148 0.1263 0.0100 -12.22 -1.052 0.1051 -0.0090 -11.22 -0.965 0.0886 -0.0230 -10.19 -0.867 0.0740 -0.0336 -9.70 -0.822 0.0684 -0.0375 -9.18 -0.769 0.0605 -0.0440 -8.18 -0.756 0.0270 -0.0578 -7.19 -0.690 0.0180 -0.0590 -6.65 -0.616 0.0166 -0.0633 -6.13 -0.542 0.0152 -0.0674 -6.00 -0.525 0.0117 -0.0732 -5.50 -0.451 0.0105 -0.0766 -5.00 -0.382 0.0097 -0.0797 -4.50 -0.314 0.0092 -0.0825 -4.00 -0.251 0.0091 -0.0853 -3.50 -0.189 0.0089 -0.0884 -3.00 -0.120 0.0089 -0.0914 -2.50 -0.051 0.0088 -0.0942 -2.00 0.017 0.0088 -0.0969 -1.50 0.085 0.0088 -0.0994 -1.00 0.152 0.0088 -0.1018 -0.50 0.219 0.0088 -0.1041 0.00 0.288 0.0087 -0.1062 0.50 0.354 0.0087 -0.1086 1.00 0.421 0.0088 -0.1107 1.50 0.487 0.0089 -0.1129 2.00 0.554 0.0090 -0.1149 2.50 0.619 0.0091 -0.1168 3.00 0.685 0.0092 -0.1185 3.50 0.749 0.0093 -0.1201 4.00 0.815 0.0095 -0.1218 4.50 0.879 0.0096 -0.1233 5.00 0.944 0.0097 -0.1248 5.50 1.008 0.0099 -0.1260 6.00 1.072 0.0101 -0.1270 6.50 1.135 0.0103 -0.1280 7.00 1.197 0.0107 -0.1287 7.50 1.256 0.0112 -0.1289 8.00 1.305 0.0125 -0.1270 9.00 1.390 0.0155 -0.1207 9.50 1.424 0.0171 -0.1158 10.00 1.458 0.0192 -0.1116 10.50 1.488 0.0219 -0.1073 11.00 1.512 0.0255 -0.1029 11.50 1.533 0.0307 -0.0983 12.00 1.549 0.0370 -0.0949 12.50 1.558 0.0452 -0.0921 13.00 1.470 0.0630 -0.0899 13.50 1.398 0.0784 -0.0885 14.00 1.354 0.0931 -0.0885 14.50 1.336 0.1081 -0.0902 15.00 1.333 0.1239 -0.0928 15.50 1.326 0.1415 -0.0963 16.00 1.329 0.1592 -0.1006 16.50 1.326 0.1743 -0.1042 17.00 1.321 0.1903 -0.1084 17.50 1.331 0.2044 -0.1125 18.00 1.333 0.2186 -0.1169 18.50 1.340 0.2324 -0.1215 19.00 1.362 0.2455 -0.1263 19.50 1.382 0.2584 -0.1313 20.00 1.398 0.2689 -0.1352 20.50 1.426 0.2814 -0.1406 21.00 1.437 0.2943 -0.1462 22.00 1.418 0.3246 -0.1516 23.00 1.397 0.3557 -0.1570 24.00 1.376 0.3875 -0.1623 25.00 1.354 0.4198 -0.1676 26.00 1.332 0.4524 -0.1728 28.00 1.293 0.5183 -0.1832 30.00 1.265 0.5843 -0.1935 32.00 1.253 0.6492 -0.2039 35.00 1.264 0.7438 -0.2193 40.00 1.258 0.8970 -0.2440 45.00 1.217 1.0402 -0.2672 50.00 1.146 1.1686 -0.2891 55.00 1.049 1.2779 -0.3097 60.00 0.932 1.3647 -0.3290 65.00 0.799 1.4267 -0.3471 70.00 0.657 1.4621 -0.3641 75.00 0.509 1.4708 -0.3799 80.00 0.362 1.4544 -0.3946 85.00 0.221 1.4196 -0.4081 90.00 0.092 1.3938 -0.4204 95.00 -0.030 1.3943 -0.4313 100.00 -0.150 1.3798 -0.4408 105.00 -0.267 1.3504 -0.4486 110.00 -0.379 1.3063 -0.4546 115.00 -0.483 1.2481 -0.4584 120.00 -0.578 1.1763 -0.4597 125.00 -0.660 1.0919 -0.4582 130.00 -0.727 0.9962 -0.4532 135.00 -0.777 0.8906 -0.4441 140.00 -0.807 0.7771 -0.4303 145.00 -0.815 0.6581 -0.4109 150.00 -0.797 0.5364 -0.3848 155.00 -0.750 0.4157 -0.3508 160.00 -0.673 0.3000 -0.3074 170.00 -0.547 0.1051 -0.2786 175.00 -0.274 0.0388 -0.1380 180.00 0.000 0.0267 0.0000