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May 6, 2014, 11:27 
SU2 cfg file and runtime problems

#1 
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Hedley
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Does anyone have an explanation of the .cfg file and what each option does in clear plain English ?
When running I get the error message "The solution contains .......... nonphysical points " followed by " terminate called after throwing instance of int " what do these mean ? any help greatly appreciated. 

May 18, 2014, 22:57 

#2  
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Francisco Palacios
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Quote:
CFL_NUMBER= 0.75 MGLEVEL= 0 Thanks for using SU2, Best Francisco 

May 20, 2014, 09:09 

#3 
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Concerning your question about an explanation of the config file. I have made a start with some other students and until now, this is what we got. It is still work in progress, keep that in mind.
Common parameters defining the numerical method Spatial gradient calculation The current version of SU2 (3.0) has the possibility to use GreenGauss (GG), Leastsquares (LS) and Weightedleastsquares (WLS) methods to compute the spatial gradients. In the case of well ordered meshes, both GG as WLS as LS can be used. [1] In principle all the methods should converge to the same result, the only difference being convergence speed. In the case of distorted meshes, GG is preferred. Distorted meshes often appear in meshes that represent boundary layers for high Reynolds numbers. Linear solver definition BCGSTAB or FGMRES In case an implicit or discrete adjoint formulation is chosen to solve a problem, one can chose two linear solvers. FGMRES is the more stable one from the two options, working for both Euler and RANS flows. In case Euler is used, one could also use BCGSTAB. Preconditioner of the krylov linear solver If one wants to use a preconditioner, a choice can be made between Jacobi, Linelet and LU_SGS. From these three preconditioners, Jacobi is the most simple one and works well for Euler but not for RANS. In case of Jacobi, RANS simulations will not converge. LU_SGS should be used if one uses RANS flow. Flow numerical method definition Convective numerical method Numerical Flow Methods are used to calculate the convective fluxes. The current version of SU2 (3.0) has the possibility to use a large number of different methods. A few are mentioned here. JST (JamesonSchmidtTurkel) JST is a central scheme with second order accuracy making it more stable but less accurate than ROE. To obtain similar results for JST as those of ROE, the mesh should be increased. Lax  Friedrich The Lax  Friedrich method (centered scheme) requires very fine grids in order to achieve correct results. In case the grids are not very fine, the numerical results will be badly smeared. This is due to the fact that the Lax  Friedrich method uses more diffusion than necessary. This method is stable for CFL numbers up to 1. (Finite Volume Methods for Hyperbolic Problems, LeVeque R.J., 2002) ROE (Roe’s Approximate Riemann Solver) ROE has better accuracy than JST and is particularly useful when examining boundary layers. Time discretization The current version of SU2 (3.0) has the possibility to use Eulerimplicit, Eulerexplicit and RungeKuttaexplicit. Difference between Implicit and Explicit Implicit schemes use both the current and latter state of a system to find a solution while explicit schemes only use a latter system. Implicit schemes are known to be more stable over a wide range of time steps at the cost of higher cost per time step since they have to perform an extra calculation. If the differential equation features a rapid decaying solution, it is called a stiff differential equation. When a differential equation is not stiff, one can use explicit methods. Difference between RungeKutta and Euler Where Euler connects points with a straight line (first order), RungeKutta methods uses parabolas (second order) and quartic curves (fourth order) to connects points with a curved line. Derivatives at multiple points are calculated in the evaluation interval after which weighted averages are calculated for the change of the function between these evaluation points. This results in a smaller error compared to Euler. RungeKutta methods are preferred for dynamic time varying chaotic systems where very large accuracy is necessary. The room for error in such systems is very small, leading to rapid changing behavior and potential nonconvergence. 

May 20, 2014, 10:49 

#4  
Senior Member

Quote:
Quote:
You need more evaluation for making a comparison between the Roe and JST schemes and their accuracy. Considering the external compressible inviscid flow over an airfoil with AoA ranging from 0 to 6 degree, the resulted aerodynamic coefficients using different schemes and their comparison demonstrate that the JST obtains more accurate aerodynamic coefficients rather than the Roe, in particular moment coefficient. The drag and lift coefficient for both schemes agree very well. However, in the AoA bigger than 3 degree, the moment coefficient resulted from 2ndorder Roe have a considerable vertical offset from the actual trend. Already, I have compared them for the aforementioned case. Their CmCl comparison has been attached. As the figure demonstrates, 3 schemes are considered for 5 cases. The 2ndorder Roe's test cases differ in their artificial dissipation coefficients, which are (0.25, 0.5, 0.5) and (0.65, 0.5, 0.02). The results from the JST scheme match with the actual trend for the test case. This is not limited to the SU2 or the inviscid twodimensional condition. Actually, I had this situation long ago with the 2ndorder Roe scheme for the viscous or inviscid flow around a wing, wingbody or airfoils using the Fluent. I have no idea yet why the 2ndorder Roe originally results in such a vertical offset from the actual trend by increasing the AoA. What I do know is JST obtains accurate coefficients. Cheers, PDP 

May 20, 2014, 14:52 
Basic lift and drag on a 3d wing

#5 
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Hedley
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Still trying to figure out why the solver on each iteration increases cl and cd and the values written to the history file are larger than expected . At a basic level if I do an analisys of a 2d airfoil or look up in abbot et al the max cl may be in the order of 1.2 . I then process a wing in SU2 expecting + 7080% of section cl and yet get values of > 3 or 4 for cl and cd > 1.
If anyone has done a flow over a 3d wing and got meaningful results ie similar to expected in a wind tunnel please post eg cfd file . Regards 

May 20, 2014, 15:08 

#6  
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Quote:


May 20, 2014, 15:17 
here is the .cfg file

#7 
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Hedley
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%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
% % % Stanford University Unstructured (SU2) configuration file % % Case description: __________________________________________________ _______ % % Author: __________________________________________________ _________________ % % Institution: __________________________________________________ ____________ % % Date: __________ % % File Version 1.1.2 October 1st, 2012 % % % %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %  DIRECT, ADJOINT, AND LINEARIZED PROBLEM DEFINITION % % % Physical governing equations (EULER, NAVIER_STOKES, % PLASMA_EULER, PLASMA_NAVIER_STOKES, % FREE_SURFACE_EULER, FREE_SURFACE_NAVIER_STOKES, % FLUID_STRUCTURE_EULER, FLUID_STRUCTURE_NAVIER_STOKES, % AEROACOUSTIC_EULER, AEROACOUSTIC_NAVIER_STOKES, % WAVE_EQUATION, HEAT_EQUATION, LINEAR_ELASTICITY) PHYSICAL_PROBLEM= NAVIER_STOKES % % Specify turbulence model (NONE, SA, SST) KIND_TURB_MODEL= SA % % Mathematical problem (DIRECT, ADJOINT, LINEARIZED) MATH_PROBLEM= DIRECT % % Restart solution (NO, YES) RESTART_SOL= NO %  COMPRESSIBLE FREESTREAM DEFINITION % % % Mach number (nondimensional, based on the freestream values) MACH_NUMBER= 0.1147 % % Angle of attack (degrees, only for compressible flows) AoA= 3 % % Sideslip angle (degrees, only for compressible flows) SIDESLIP_ANGLE= 0.0 % % Freestream temperature (273.15 K by default) FREESTREAM_TEMPERATURE= 273.15 % this is 15 degrees C HEDLEY % % Reynolds number (nondimensional, based on the freestream values) REYNOLDS_NUMBER= 3.282E6 % % Reynolds length (1 m by default) REYNOLDS_LENGTH= 1 %  COMPRESSIBLE AND INCOMPRESSIBLE FLUID CONSTANTS % % % Ratio of specific heats (1.4 (air), only for compressible flows) GAMMA_VALUE= 1.4 % % Specific gas constant (287.87 J/kg*K (air), only for compressible flows) GAS_CONSTANT= 287.87 % % Laminar Prandtl number (0.72 (air), only for compressible flows) PRANDTL_LAM= 0.72 % % Turbulent Prandtl number (0.9 (air), only for compressible flows) PRANDTL_TURB= 0.9 %  REFERENCE VALUE DEFINITION % % % Conversion factor for converting the grid to meters CONVERT_TO_METER= 1 % % Write a new mesh converted to meters (NO, YES) WRITE_CONVERTED_MESH = NO % % Reference origin for moment computation REF_ORIGIN_MOMENT_X = 0.25 REF_ORIGIN_MOMENT_Y = 0.00 REF_ORIGIN_MOMENT_Z = 0.00 % % Reference length for pitching, rolling, and yawing nondimensional moment REF_LENGTH_MOMENT= 1 % % Reference area for force coefficients (0 implies automatic calculation) REF_AREA= 0 % % Reference pressure (101325.0 N/m^2 by default, only for compressible flows) REF_PRESSURE= 1 % % Reference temperature (273.15 K by default, only for compressible flows) REF_TEMPERATURE= 1 % % Reference density (1.2886 Kg/m^3 by default, only for compressible flows) REF_DENSITY= 1 % % Reference element length for computing the slope limiter epsilon REF_ELEM_LENGTH= 0.1 %  BOUNDARY CONDITION DEFINITION % % % NavierStokes wall boundary marker(s) (NONE = no marker) MARKER_HEATFLUX= ( wing, 0.0 ) % % Farfield boundary marker(s) (NONE = no marker) MARKER_FAR= ( farfield ) % % Symmetry boundary marker(s) (NONE = no marker) MARKER_SYM= ( symmetry ) %was symmetry hedley % Adiabatic wall boundary condition (YES, NO) ADIABATIC_WALL= YES % % Marker(s) of the surface to be plotted or designed MARKER_PLOTTING= ( wing ) % was wing hd % Marker(s) of the surface where the functional (Cd, Cl, etc.) will be evaluated MARKER_MONITORING= ( wing) %  COMMON PARAMETERS DEFINING THE NUMERICAL METHOD % % % Numerical method for spatial gradients (GREEN_GAUSS, WEIGHTED_LEAST_SQUARES) NUM_METHOD_GRAD= GREEN_GAUSS % % CourantFriedrichsLewy condition of the finest grid CFL_NUMBER= .5 % % CFL ramp (factor, number of iterations, CFL limit) CFL_RAMP= ( 1.0, 50, 2.0 ) % % RungeKutta alpha coefficients RK_ALPHA_COEFF= ( 0.66667, 0.66667, 1.000000 ) % % Number of total iterations EXT_ITER= 100 %  LINEAR SOLVER DEFINITION % % % Linear solver for the implicit (or discrete adjoint) formulation (BCGSTAB, FGMRES) LINEAR_SOLVER= FGMRES % % Preconditioner of the Krylov linear solver (NONE, JACOBI, LINELET) LINEAR_SOLVER_PREC= LU_SGS % % Min error of the linear solver for the implicit formulation LINEAR_SOLVER_ERROR= 1E6 % % Max number of iterations of the linear solver for the implicit formulation LINEAR_SOLVER_ITER= 20 % % Linear solver history LINEAR_SOLVER_HIST= NO %  MULTIGRID PARAMETERS % % % MultiGrid Levels (0 = no multigrid) MGLEVEL= 0 % % MultiGrid Cycle (0 = V cycle, 1 = W Cycle) MGCYCLE= 0 % % CFL reduction factor on the coarse levels MG_CFL_REDUCTION= 0.5 % % Maximum number of children in the agglomeration stage MAX_CHILDREN= 50 % % Maximum length of an agglomerated element (relative to the domain) MAX_DIMENSION= 0.1 % % Multigrid presmoothing level MG_PRE_SMOOTH= ( 1, 2, 2, 2 ) % % Multigrid postsmoothing level MG_POST_SMOOTH= ( 0, 0, 0, 0 ) % % Jacobi implicit smoothing of the correction MG_CORRECTION_SMOOTH= ( 0, 0, 0, 0 ) % % Damping factor for the residual restriction MG_DAMP_RESTRICTION= 0.85 % % Damping factor for the correction prolongation MG_DAMP_PROLONGATION= 0.85 % % Full Multigrid (NO, YES) FULLMG= NO % % Start up iterations using the fine grid START_UP_ITER= 0 %  FLOW NUMERICAL METHOD DEFINITION % % % Convective numerical method (JST, LAXFRIEDRICH, CUSP, ROE, AUSM, HLLC, % TURKEL_PREC, MSW) CONV_NUM_METHOD_FLOW= JST %was ROE hedley % Spatial numerical order integration (1ST_ORDER, 2ND_ORDER, 2ND_ORDER_LIMITER) % SPATIAL_ORDER_FLOW= 2ND_ORDER_LIMITER % % Slope limiter (VENKATAKRISHNAN, MINMOD) SLOPE_LIMITER_FLOW= VENKATAKRISHNAN % % Coefficient for the limiter (smooth regions) LIMITER_COEFF= 0.3 % % 1st, 2nd and 4th order artificial dissipation coefficients AD_COEFF_FLOW= ( 0.15, 0.5, 0.02 ) % % Viscous numerical method (AVG_GRAD, AVG_GRAD_CORRECTED, GALERKIN) VISC_NUM_METHOD_FLOW= AVG_GRAD_CORRECTED % % Source term numerical method (PIECEWISE_CONSTANT) SOUR_NUM_METHOD_FLOW= PIECEWISE_CONSTANT % % Time discretization (RUNGEKUTTA_EXPLICIT, EULER_IMPLICIT, EULER_EXPLICIT) TIME_DISCRE_FLOW= EULER_IMPLICIT %  TURBULENT NUMERICAL METHOD DEFINITION % % % Convective numerical method (SCALAR_UPWIND) CONV_NUM_METHOD_TURB= SCALAR_UPWIND % % Spatial numerical order integration (1ST_ORDER, 2ND_ORDER, 2ND_ORDER_LIMITER) % SPATIAL_ORDER_TURB= 1ST_ORDER % % Slope limiter (VENKATAKRISHNAN, MINMOD) SLOPE_LIMITER_TURB= VENKATAKRISHNAN % % Viscous numerical method (AVG_GRAD, AVG_GRAD_CORRECTED) VISC_NUM_METHOD_TURB= AVG_GRAD_CORRECTED % % Source term numerical method (PIECEWISE_CONSTANT) SOUR_NUM_METHOD_TURB= PIECEWISE_CONSTANT % % Time discretization (EULER_IMPLICIT) TIME_DISCRE_TURB= EULER_IMPLICIT %  PARTITIONING STRATEGY % % Write a tecplot/paraview file for each partition (NO, YES) VISUALIZE_PART= YES %  GEOMETRY EVALUATION PARAMETERS % % % Geometrical evaluation mode (FUNCTION, GRADIENT) GEO_MODE= FUNCTION % % Marker(s) of the surface where geometrical based func. will be evaluated GEO_MARKER= ( airfoil ) % % Number of airfoil sections GEO_NUMBER_SECTIONS= 5 % % Orientation of airfoil sections (X_AXIS, Y_AXIS, Z_AXIS) GEO_ORIENTATION_SECTIONS= Y_AXIS % % Location (coordinate) of the airfoil sections (MinValue, MaxValue) GEO_LOCATION_SECTIONS= (1.5, 3.5) % % Plot loads and Cp distributions on each airfoil section GEO_PLOT_SECTIONS= YES % % Number of section cuts to make when calculating internal volume GEO_VOLUME_SECTIONS= 101 %  GRID ADAPTATION STRATEGY % % % Percentage of new elements (% of the original number of elements) NEW_ELEMS= 5 % % Kind of grid adaptation (NONE, FULL, FULL_FLOW, GRAD_FLOW, FULL_ADJOINT, % GRAD_ADJOINT, GRAD_FLOW_ADJ, ROBUST, % FULL_LINEAR, COMPUTABLE, COMPUTABLE_ROBUST, % REMAINING, WAKE, SMOOTHING, SUPERSONIC_SHOCK, % TWOPHASE) KIND_ADAPT= FULL_FLOW % % Scale factor for the dual volume DUALVOL_POWER= 0.5 % % Use analytical definition for surfaces (NONE, NACA0012_AIRFOIL, BIPARABOLIC, % NACA4412_AIRFOIL, CYLINDER) ANALYTICAL_SURFDEF= NONE % % Before each computation do an implicit smoothing of the nodes coord (NO, YES) SMOOTH_GEOMETRY= NO % % Adapt the boundary elements (NO, YES) ADAPT_BOUNDARY= YES %  GRID DEFORMATION PARAMETERS % % % Kind of deformation (FFD_SETTING, HICKS_HENNE, PARABOLIC, NACA_4DIGITS, % DISPLACEMENT, ROTATION, FFD_CONTROL_POINT, % FFD_DIHEDRAL_ANGLE, FFD_TWIST_ANGLE, % FFD_ROTATION, FFD_CAMBER, FFD_THICKNESS, FFD_VOLUME) DV_KIND= FFD_SETTING % % Marker of the surface in which we are going apply the shape deformation DV_MARKER= ( airfoil ) % % Parameters of the shape deformation %  HICKS_HENNE ( Lower Surface (0)/Upper Surface (1)/Only one Surface (2), x_Loc ) %  NACA_4DIGITS ( 1st digit, 2nd digit, 3rd and 4th digit ) %  PARABOLIC ( Center, Thickness ) %  DISPLACEMENT ( x_Disp, y_Disp, z_Disp ) %  ROTATION ( x_Orig, y_Orig, z_Orig, x_End, y_End, z_End ) %  OBSTACLE ( Center, Bump size ) %  FFD_CONTROL_POINT ( Chunk ID, i_Ind, j_Ind, k_Ind, x_Disp, y_Disp, z_Disp ) %  FFD_DIHEDRAL_ANGLE ( Chunk ID, x_Orig, y_Orig, z_Orig, x_End, y_End, z_End ) %  FFD_TWIST_ANGLE ( Chunk ID, x_Orig, y_Orig, z_Orig, x_End, y_End, z_End ) %  FFD_ROTATION ( Chunk ID, x_Orig, y_Orig, z_Orig, x_End, y_End, z_End ) %  FFD_CAMBER ( Chunk ID, i_Ind, j_Ind ) %  FFD_THICKNESS ( Chunk ID, i_Ind, j_Ind ) %  FFD_VOLUME ( Chunk ID, i_Ind, j_Ind ) DV_PARAM= ( 1, 0.5 ) % % New value of the shape deformation DV_VALUE= 0.01 % % Grid deformation technique (SPRING, FEA) GRID_DEFORM_METHOD= SPRING % % Hold the grid fixed in a region (NO, YES) HOLD_GRID_FIXED= NO % % Coordinates of the box where the grid will be deformed (Xmin, Ymin, Zmin, % Xmax, Ymax, Zmax) HOLD_GRID_FIXED_COORD= ( 0.5, 0.49, 0.0, 2.5, 0.49, 0.0 ) % % Visualize the deformation (NO, YES) VISUALIZE_DEFORMATION= NO %  CONVERGENCE PARAMETERS % % % Convergence criteria (CAUCHY, RESIDUAL) % CONV_CRITERIA= RESIDUAL % % Residual reduction (order of magnitude with respect to the initial value) RESIDUAL_REDUCTION= 5 % % Min value of the residual (log10 of the residual) RESIDUAL_MINVAL= 8 % % Start convergence criteria at iteration number STARTCONV_ITER= 10 % % Number of elements to apply the criteria CAUCHY_ELEMS= 100 % % Epsilon to control the series convergence CAUCHY_EPS= 1E10 % % Function to apply the criteria (LIFT, DRAG, NEARFIELD_PRESS, SENS_GEOMETRY, % SENS_MACH, DELTA_LIFT, DELTA_DRAG) CAUCHY_FUNC_FLOW= DRAG CAUCHY_FUNC_ADJFLOW= SENS_GEOMETRY CAUCHY_FUNC_LIN= DELTA_DRAG % % Epsilon for full multigrid method evaluation FULLMG_CAUCHY_EPS= 1E4 %  INPUT/OUTPUT INFORMATION % % % Mesh input file MESH_FILENAME= wingscaled.su2 % was wing.su2 HEDLEY % Mesh input file format (SU2, CGNS, NETCDF_ASCII) MESH_FORMAT= SU2 % % Divide rectangles into triangles (NO, YES) DIVIDE_ELEMENTS= NO % % Convert a CGNS mesh to SU2 format (YES, NO) CGNS_TO_SU2= NO % % Mesh output file MESH_OUT_FILENAME= mesh_out.su2 % % Restart flow input file SOLUTION_FLOW_FILENAME= solution_flow.dat % % Restart linear flow input file SOLUTION_LIN_FILENAME= solution_lin.dat % % Restart adjoint input file SOLUTION_ADJ_FILENAME= solution_adj.dat % % Output file format (PARAVIEW, TECPLOT, STL) OUTPUT_FORMAT= PARAVIEW % % Output file convergence history (w/o extension) CONV_FILENAME= history % % Output file linear solver history (w/o extension) LIN_CONV_FILENAME= lin_history % % Output file restart flow RESTART_FLOW_FILENAME= restart_flow.dat % % Output file restart adjoint RESTART_ADJ_FILENAME= restart_adj.dat % % Output file linear flow RESTART_LIN_FILENAME= restart_lin.dat % % Output file flow (w/o extension) variables VOLUME_FLOW_FILENAME= flow % % Output file adjoint (w/o extension) variables VOLUME_ADJ_FILENAME= adjoint % % Output file linearized (w/o extension) variables VOLUME_LIN_FILENAME= linearized % % Output objective function gradient (using continuous adjoint) GRAD_OBJFUNC_FILENAME= of_grad.dat % % Output file surface flow coefficient (w/o extension) SURFACE_FLOW_FILENAME= surface_flow % % Output file surface adjoint coefficient (w/o extension) SURFACE_ADJ_FILENAME= surface_adjoint % % Output file surface linear coefficient (w/o extension) SURFACE_LIN_FILENAME= surface_linear % % Writing solution file frequency WRT_SOL_FREQ= 100 % % Writing convergence history frequency WRT_CON_FREQ= 1 % % Writing linear solver history frequency WRT_LIN_CON_FREQ= 1 % % Write unsteady data adding headers and prefixes (NO, YES) WRT_UNSTEADY= NO % % Write mass averaged solution file (plasma solver only, NO by default) WRT_MASS_AVG_FLOW= NO %  OPTIMAL SHAPE DESIGN DEFINITION % % % Objective function (DRAG, LIFT, SIDEFORCE, MOMENT_X, MOMENT_Y, % MOMENT_Z, EFFICIENCY, EQUIVALENT_AREA, NEARFIELD_PRESSURE, % FORCE_X, FORCE_Y, FORCE_Z, THRUST, TORQUE, FREESURFACE) OBJFUNC= DRAG % % Scale objective function. OBJFUNC_SCALE= 0.001 % % Inequality constraints list separated by comma (DRAG, LIFT, SIDEFORCE, % MOMENT_X, MOMENT_Y, MOMENT_Z, EFFICIENCY, EQUIVALENT_AREA, % NEARFIELD_PRESSURE, FORCE_X, FORCE_Y, FORCE_Z, THRUST, TORQUE, % FREESURFACE) CONST_IEQ= LIFT, MOMENT_Z % % Scale inequality constraints (separated by comma) CONST_IEQ_SCALE= 0.001, 0.001 % % Min value inequality constraints list (NONE, LESS, GREATER) CONST_IEQ_SIGN= GREATER, GREATER % % Max value inequality constraints list (separated by comma) CONST_IEQ_VALUE= 0.328188, 0.034068 % % Equality constraints list separated by comma (DRAG, LIFT, SIDEFORCE, MOMENT_X, MOMENT_Y, % MOMENT_Z, EFFICIENCY, EQUIVALENT_AREA, NEARFIELD_PRESSURE, % FORCE_X, FORCE_Y, FORCE_Z, THRUST, TORQUE, FREESURFACE) CONST_EQ= NONE % % Scale equality constraints (separated by comma) CONST_EQ_SCALE= 0.0 % % Value equality constraints list (separated by comma) CONST_EQ_VALUE= 0.0 % % List of design variables (Design variables are separated by semicolons) % From 1 to 99, Geometrycal design variables. %  HICKS_HENNE ( 1, Scale  Mark. List  Lower(0)/Upper(1) side, x_Loc ) %  NACA_4DIGITS ( 4, Scale  Mark. List  1st digit, 2nd digit, 3rd and 4th digit ) %  DISPLACEMENT ( 5, Scale  Mark. List  x_Disp, y_Disp, z_Disp ) %  ROTATION ( 6, Scale  Mark. List  x_Axis, y_Axis, z_Axis, x_Turn, y_Turn, z_Turn ) %  FFD_CONTROL_POINT ( 7, Scale  Mark. List  Chunk, i_Ind, j_Ind, k_Ind, x_Mov, y_Mov, z_Mov ) %  FFD_DIHEDRAL_ANGLE ( 8, Scale  Mark. List  Chunk, x_Orig, y_Orig, z_Orig, x_End, y_End, z_End ) %  FFD_TWIST_ANGLE ( 9, Scale  Mark. List  Chunk, x_Orig, y_Orig, z_Orig, x_End, y_End, z_End ) %  FFD_ROTATION ( 10, Scale  Mark. List  Chunk, x_Orig, y_Orig, z_Orig, x_End, y_End, z_End ) %  FFD_CAMBER ( 11, Scale  Mark. List  Chunk, i_Ind, j_Ind ) %  FFD_THICKNESS ( 12, Scale  Mark. List  Chunk, i_Ind, j_Ind ) %  FFD_VOLUME ( 13, Scale  Mark. List  Chunk, i_Ind, j_Ind ) % From 100 to 199, Flow solver design variables. %  MACH_NUMBER ( 101, Scale  Markers List ) %  AOA ( 102, Scale  Markers List ) DEFINITION_DV= ( 1, 0.001  airfoil  0, 0.1 ); ( 1, 0.001  airfoil  0, 0.2 ) 

May 20, 2014, 15:45 
and here is some output with Cl and Cd increasing each loop

#8 
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Hedley
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Initialize jacobian structure (NavierStokes). MG level: 0.
Initialize jacobian structure (SA model).  Integration and Numerics Preprocessing  Integration Preprocessing. Numerics Preprocessing.  Begin Solver  Maximum residual: 1.89096, located at point 1648. Iter Time(s) Res[Rho] Res[nu] CLift(Total) CDrag(Total) 0 25.741053 0.323718 4.513804 0.839490 0.345001 1 24.044417 0.187468 4.448524 1.183776 0.487726 2 21.945658 0.126626 4.477596 1.464226 0.607455 3 21.030338 0.076769 4.502995 1.693231 0.705232 4 20.238298 0.043514 4.522412 1.876937 0.785252 5 20.570025 0.016620 4.534918 2.026783 0.852054 6 20.145987 0.005051 4.541085 2.147346 0.907573 7 19.977807 0.021602 4.542675 2.242479 0.953327 8 19.858203 0.034791 4.541540 2.316662 0.990674 9 20.289774 0.046206 4.539163 2.374541 1.021008 

May 20, 2014, 15:48 

#9 
Senior Member

In what dimension the grid and geometry is created? Inch or m?
Further, your geometry is a wing, right? what is the root and tip chord length? Besides, what is the MAC (Mean Aerodynamic Chord) length? 

May 20, 2014, 15:53 
sizing

#10 
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Hedley
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The wing is in meters  chord 1.6 meters  tip chord 800mm  half span 3.5 meters . The root airfoil is NACA 64212.5 and tip NACA 64210 .
The grid and volume were created in pointwise which used inches and then after reading previous post I converted down to meters in pointwise by scaling by 0.0254 regards 

May 20, 2014, 15:58 
images attached

#11 
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MAC is + 400mm


May 20, 2014, 17:46 

#12 
Senior Member

Okay, thank you for providing extra information.
The main point is, your Mach number is below the 0.3, so the flow is incompressible. However, you have created a compressible domain which includes hemisphere farfield, symmetric plane and the wall. Typically, for the incompressible external flow we are defining an inlet, outlet with a cylinderlike or rectangularlike domain. Nevertheless, by solving the compressible flow with the incompressible flow condition or initial condition, the results will be obtained same as using the incompressible solver. For this purpose, you need to set your config file for solving the compressible flow, albeit with incompressible flow condition. Therefore, from the aforementioned points, I have changed some part of your Config file which is stated as follow: First, your REFERENCE VALUE DEFINITION should be changed for the compressible flow. Code:
% Reference pressure (101325.0 N/m^2 by default, only for compressible flows) REF_PRESSURE= 101325.0 % % Reference temperature (273.15 K by default, only for compressible flows) REF_TEMPERATURE= 273.15 % % Reference density (1.2886 Kg/m^3 by default, only for compressible flows) REF_DENSITY= 1.292319138159549 MAC calculation for trapezoid wing: Code:
((2/3)*Croot*(1+taper+(taper)^2))/(1+taper) Code:
M=0.1147 T=273.15 R=8314.4621 Rs=286.959 MAC=1.1022 ro=(101325)/((R/28.966)*T) mo=1.79*10e05 Re=ro*(M*((gama*Rs*T)^0.5))*MAC/mo Code:
%COMPRESSIBLE FREESTREAM DEFINITION % Reynolds number (nondimensional, based on the freestream values) REYNOLDS_NUMBER= 3.023535797839057E5 % % Reynolds length (1 m by default) REYNOLDS_LENGTH= 1.1022 %REFERENCE VALUE DEFINITION % Reference origin for moment computation REF_ORIGIN_MOMENT_X = 0.77335 REF_ORIGIN_MOMENT_Y = 0.00 REF_ORIGIN_MOMENT_Z = 0.00 % Reference length for pitching, rolling, and yawing nondimensional moment REF_LENGTH_MOMENT= 1.1022 Third, in this case, you need more than 100 iterations for obtaining suitable convergence, so I changed the max iteration number. Code:
% Number of total iterations EXT_ITER= 100000 Code:
% MultiGrid Levels (0 = no multigrid) MGLEVEL= 2 % % MultiGrid Cycle (0 = V cycle, 1 = W Cycle) MGCYCLE= 1 % % CFL reduction factor on the coarse levels MG_CFL_REDUCTION= 0.75 % Multigrid presmoothing level MG_PRE_SMOOTH= ( 1, 2, 3, 3 ) Code:
% Spatial numerical order integration (1ST_ORDER, 2ND_ORDER, 2ND_ORDER_LIMITER) % SPATIAL_ORDER_TURB= 2ND_ORDER In this way, you will create a Hybrid grid, which definitely improves your results. Then, you could use the SST turbulence model in your Config file, which gives better results in comparison with the SA model. Note: The modified Config file by me has been attached, however I couldn't find time to check it with my laptop. In case it didn't work, please let me know. Besides, it didn't let me to attach ".cfg" file, after downloading the file, change the extension to the ".cfg" Good Luck, PDP Last edited by pdp.aero; May 20, 2014 at 20:04. Reason: attachment 

May 20, 2014, 23:15 

#14 
Member
Hedley
Join Date: May 2014
Posts: 52
Rep Power: 9 
Dear pdp.aero
Thank you for a comprehensive and very well explained and informative reply . It is 4am in South Africa so I will start up the computer and try out when spousal unit awakens to avoid domestic violence  much appreciated . Hedley Last edited by hedley; May 20, 2014 at 23:24. Reason: Wrong name spelling 

May 21, 2014, 02:13 
try of new cfg file

#15 
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Hedley
Join Date: May 2014
Posts: 52
Rep Power: 9 
All goes well and then Killed ?
 Config file boundary information  Farfield boundary marker(s): farfield. Symmetry plane boundary marker(s): symmetry. Constant heat flux wall boundary marker(s): wing.  Flow & Nondimensionalization information  Viscous flow: Computing pressure using the ideal gas law based on the freestream temperature and a density computed from the Reynolds number. Input conditions: Ratio of specific heats: 1.4 Specific gas constant (J/(kg.K)): 287.87 Freestream pressure (N/m^2): 9763.69 Freestream temperature (K): 273.15 Freestream density (kg/m^3): 0.12417 Freestream velocity (m/s): (38.0041,0,1.99171) Freestream velocity magnitude (m/s): 38.0563 Freestream energy (kg.m/s^2): 197303 Freestream viscosity (N.s/m^2): 1.72261e05 Reference values: Reference pressure (N/m^2): 101325 Reference temperature (K): 273.15 Reference energy (kg.m/s^2): 78405.6 Reference density (kg/m^3): 1.29232 Reference velocity (m/s): 280.01 Reference viscosity (N.s/m^2): 361.862 Resulting nondimensional state: Mach number (nondimensional): 0.1147 Reynolds number (nondimensional): 302354 Reynolds length (m): 1.1022 Specific gas constant (nondimensional): 287.87 Freestream temperature (nondimensional): 1 Freestream pressure (nondimensional): 0.0963602 Freestream density (nondimensional): 0.0960831 Freestream velocity (nondimensional): (0.135724,0,0.007113) Freestream velocity magnitude (nondimensional): 0.13591 Freestream turb. kinetic energy (nondimensional): 6.92687e05 Freestream specific dissipation (nondimensional): 13.981 Freestream energy (nondimensional): 2.51645 Freestream viscosity (nondimensional): 4.76042e08 Force coefficients computed using freestream values. Note: Negative pressure, temperature or density is not allowed!  Read grid file information  Three dimensional problem. 2819141 interior elements. 2819141 tetrahedra. 485190 points. 3 surface markers. 18306 boundary elements in index 0 (Marker = farfield). 20743 boundary elements in index 1 (Marker = symmetry). 27467 boundary elements in index 2 (Marker = wing).  Geometry Preprocessing  Setting local point and element connectivity. Checking the numerical grid orientation. Identifying edges and vertices. Computing centers of gravity. Setting the control volume structure. Volume of the computational grid: 7.41088e+06. Searching for the closest normal neighbors to the surfaces. Compute the surface curvature. Max K: 90362.6. Mean K: 324.818. Standard deviation K: 3003.99. Setting the multigrid structure. CVs of the MG level: 82163. Agglom. rate 1/5.90521. MG level: 1. CVs of the MG level: 12560. Agglom. rate 1/6.54164. MG level: 2.  Solver Preprocessing  Computing wall distances. Area projection in the zplane = 4.02345. Initialize jacobian structure (NavierStokes). MG level: 0. Initialize jacobian structure (SA model). Initialize jacobian structure (NavierStokes). MG level: 1. Killed 

May 21, 2014, 04:02 
test run

#16 
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Hedley
Join Date: May 2014
Posts: 52
Rep Power: 9 
1. Ok the only way I can stop it crashing is to set MGLEVEL=0
2. I thought taper ratio was tip chord / root chord 3. I fixed the one typo changing 2st to 2ND. Ran the suggested .cfg file which crashes with multi grid and runs with multigrid set to 0 BUT then back to where I started with CLift and Cdrag increasing on each iteration . There seem to be many input variables and I am sure that there are dependencies between them so unlike my airplane with has a few control inputs SU2 is like a black box with many many knobs that one tweeks without any real idea of how this will affect the solver. If anyone has ever run a .cfg file on a 3d wing section at anywhere between 60KTS and 175KTS and got outputs for Clift and CDrag that make sense please assist. Regards Hedley 

May 21, 2014, 04:31 

#17 
Senior Member

Please, turn off the AMG solver.
Code:
% MultiGrid Levels (0 = no multigrid) MGLEVEL= 0 % % MultiGrid Cycle (0 = V cycle, 1 = W Cycle) MGCYCLE= 0 Okay, after I posted this, I have seen your last post. Yes, I have noticed that using AMG solver with parallel RANS equations is not compatible, and they didn't use them for their tutorial test cases. You are right, taper ratio is Ctip/Croot, typically it is somewhere between 0.2 to 0.3 for general aviation. Your wing is similar to the ONERA M6 geometry, one of the SU2's tutorial, but your Mach number is pretty small, and your Reynolds number indicates the Laminar flow (e.g. below the 5e+05). Therefore, my suggestion is change your Mach number to 0.83 and calculate the Reynolds number, then run the code. In this way you are running test case similar to the tutorial 7, turbulent ONERA M6. I already run this tutorial, and I got the convergence. Continuing this, if you got the convergence, it means you need to change your solver into the incompressible. For checking the convergence, after your iteration finished, plot the residual and see they are going down or not. If they come down 4 order of magnitude, you are obtaining the convergence hopefully. After this, when you concluded that you got the convergence with 0.83 Mach, you need to change into incompressible solver. For this purpose, create a rectangularlike unstructured block around your wing. Consequently, define velocity inlet at the entry, pressure outlet at the outlet, wall or pressure outlet for the surrounding domain and the symmetric boundary condition. For setting your configuration, please refer to the tutorial 3, laminar flat plat, in order to get more information on implementing laminar NS equations for your case. Run the code on your new configuration. Sorry if couldn't help you any more. Good Luck, Payam Last edited by pdp.aero; May 21, 2014 at 05:21. Reason: Updating 

May 21, 2014, 12:26 
final attempt

#18 
Member
Hedley
Join Date: May 2014
Posts: 52
Rep Power: 9 
Thank you for the time spent on this
1. I turned off AMG 2. Am running Latest version see below 3.1.0 Eagle 3. Still comes up with strange values. I will put this to one side and realize that CFD is not for the faint hearted  regards and thanks again .   _____ _ _ ___   / ____     __ \ Web: su2.stanford.edu    (___     )  Twitter: @su2code   \___ \     / / Forum: www.cfdonline.com/Forums/su2/   ____)   __  / /_   _____/ \____/ ____ Suite (Computational Fluid Dynamic Code)   Release 3.1.0 "eagle"    SU2, Copyright (C) 20122014 Aerospace Design Laboratory (ADL).   SU2 is distributed in the hope that it will be useful,   but WITHOUT ANY WARRANTY; without even the implied warranty of   MERCHANTABILITY or FITNESS FOR A PARTICULAR PURPOSE. See the GNU   Lesser General Public License (version 2.1) for more details.    Physical case definition  Compressible RANS equations. Turbulence model: Spalart Allmaras Mach number: 0.1147. Angle of attack (AoA): 3 deg, and angle of sideslip (AoS): 0 deg. Reynolds number: 302354. No restart solution, use the values at infinity (freestream). The reference length/area will be computed using y(2D) or z(3D) projection. The reference length (moment computation) is 1.2444. Reference origin (moment computation) is (0.77335, 0, 0). Surface(s) where the force coefficients are evaluated: wing. Surface(s) plotted in the output file: wing. Surface(s) affected by the design variables: airfoil. Input mesh file name: smallwing.su2  Space numerical integration  Roe solver for the flow inviscid terms. Second order integration with slope limiter. Venkatakrishnan slopelimiting method, with constant: 0.3. The reference element size is: 0.1. Scalar upwind solver (first order) for the turbulence model. Second order integration. Average of gradients with correction (viscous flow terms). Piecewise constant integration of the flow source terms. Average of gradients with correction (viscous turbulence terms). Piecewise constant integration of the turbulence model source terms. Gradient computation using GreenGauss theorem.  Time numerical integration  Local time stepping (steady state simulation). Euler implicit method for the flow equations. FGMRES is used for solving the linear system. Convergence criteria of the linear solver: 1e06. Max number of iterations: 20. Relaxation coefficient: 1. No CFL ramp. CourantFriedrichsLewy number: 0.75 Euler implicit time integration for the turbulence model.  Convergence criteria  Maximum number of iterations: 100. Reduce the density residual 5 orders of magnitude. The minimum bound for the density residual is 10^(8). Start convergence criteria at iteration 10.  Output information  Writing a flow solution every 100 iterations. Writing the convergence history every 1 iterations. The output file format is Paraview ASCII (.vtk). Convergence history file name: history. Surface flow coefficients file name: surface_flow. Flow variables file name: flow. Restart flow file name: restart_flow.dat.  Config file boundary information  Farfield boundary marker(s): farfield. Symmetry plane boundary marker(s): symmetry. Constant heat flux wall boundary marker(s): wing.  Flow & Nondimensionalization information  Viscous flow: Computing pressure using the ideal gas law based on the freestream temperature and a density computed from the Reynolds number. Input conditions: Ratio of specific heats: 1.4 Specific gas constant (J/(kg.K)): 287.87 Freestream pressure (N/m^2): 8647.98 Freestream temperature (K): 273.15 Freestream density (kg/m^3): 0.109981 Freestream velocity (m/s): (38.0041,0,1.99171) Freestream velocity magnitude (m/s): 38.0563 Freestream energy (kg.m/s^2): 197303 Freestream viscosity (N.s/m^2): 1.72261e05 Reference values: Reference pressure (N/m^2): 101325 Reference temperature (K): 273.15 Reference energy (kg.m/s^2): 78405.6 Reference density (kg/m^3): 1.29232 Reference velocity (m/s): 280.01 Reference viscosity (N.s/m^2): 361.862 Resulting nondimensional state: Mach number (nondimensional): 0.1147 Reynolds number (nondimensional): 302354 Reynolds length (m): 1.2444 Specific gas constant (nondimensional): 287.87 Freestream temperature (nondimensional): 1 Freestream pressure (nondimensional): 0.0853489 Freestream density (nondimensional): 0.0851035 Freestream velocity (nondimensional): (0.135724,0,0.007113) Freestream velocity magnitude (nondimensional): 0.13591 Freestream turb. kinetic energy (nondimensional): 6.92687e05 Freestream specific dissipation (nondimensional): 12.3834 Freestream energy (nondimensional): 2.51645 Freestream viscosity (nondimensional): 4.76042e08 Force coefficients computed using freestream values. Note: Negative pressure, temperature or density is not allowed!  Read grid file information  Three dimensional problem. 1573830 interior elements. 946959 tetrahedra. 618440 prisms. 8431 pyramids. 485190 points. 3 surface markers. 18306 boundary elements in index 0 (Marker = farfield). 17850 boundary elements in index 1 (Marker = symmetry). 27467 boundary elements in index 2 (Marker = wing).  Geometry Preprocessing  Setting local point and element connectivity. Checking the numerical grid orientation. Identifying edges and vertices. Computing centers of gravity. Setting the control volume structure. Volume of the computational grid: 7.41088e+06. Searching for the closest normal neighbors to the surfaces. Compute the surface curvature. Max K: 90362.6. Mean K: 345.171. Standard deviation K: 3074.93.  Solver Preprocessing  Computing wall distances. Area projection in the zplane = 4.02345. Initialize jacobian structure (NavierStokes). MG level: 0. Initialize jacobian structure (SA model).  Integration and Numerics Preprocessing  Integration Preprocessing. Numerics Preprocessing.  Begin Solver  Maximum residual: 4.9885, located at point 1648. Iter Time(s) Res[Rho] Res[nu] CLift(Total) CDrag(Total) 0 20.581689 6.555742 10.821365 0.840691 0.411578 1 18.891452 6.682332 10.750260 1.252222 0.573674 2 19.230903 6.742428 10.774174 1.593348 0.706413 

January 25, 2016, 08:00 
su2 shape optimization using hickshenne bump fution

#19 
New Member
balaji
Join Date: Jan 2016
Posts: 14
Rep Power: 7 
Any having shape optimization cfg file or please tell me how to do shape optimization in su2 using hickshenne bup function please


January 26, 2016, 05:17 

#20 
Super Moderator
Tim Albring
Join Date: Sep 2015
Posts: 195
Rep Power: 8 
Hi bala732,
you can find a tutorial using HicksHenne functions here: https://github.com/su2code/SU2/wiki/...nsonicAirfoil 

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