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High drag for airfoil compared to XFOIL and wind tunnel data 

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February 22, 2016, 12:11 
High drag for airfoil compared to XFOIL and wind tunnel data

#1 
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Ryan Barrett
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I am trying to get the lift and drag coefficients out of a 2D airfoil using SU2. I am doing a steady 2D incompressible RANS simulation with the SA turbulence model. I have attached two graphs that I compare the results from XFOIL(panel method) and known wind tunnel data for different angles of attack. The lift matches pretty well but the drag is very high. Is there something wrong with my config file that would cause this? I tried to attach my su2 mesh but it was too big, but let me know if you need it. Thanks!
Code:
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% % % % SU2 configuration file % % Case description: 2D NACA 0012 Airfoil Validation Case (incompressible) % % http://turbmodels.larc.nasa.gov/naca0012_val_sa.html % % Author: Francisco Palacios % % Institution: Stanford University % % Date: Feb 18th, 2013 % % File Version 4.0.0 "Cardinal" % % % %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %  DIRECT, ADJOINT, AND LINEARIZED PROBLEM DEFINITION % % % Physical governing equations (EULER, NAVIER_STOKES, % TNE2_EULER, TNE2_NAVIER_STOKES, % WAVE_EQUATION, HEAT_EQUATION, LINEAR_ELASTICITY, % POISSON_EQUATION) PHYSICAL_PROBLEM= NAVIER_STOKES % % Specify turbulent model (NONE, SA, SA_NEG, SST) KIND_TURB_MODEL= SA % % Mathematical problem (DIRECT, ADJOINT, LINEARIZED) MATH_PROBLEM= DIRECT % % Regime type (COMPRESSIBLE, INCOMPRESSIBLE, FREESURFACE) REGIME_TYPE= INCOMPRESSIBLE % % Restart solution (NO, YES) RESTART_SOL= NO %  UNSTEADY SIMULATION % % % Unsteady simulation (NO, TIME_STEPPING, DUAL_TIME_STEPPING1ST_ORDER, % DUAL_TIME_STEPPING2ND_ORDER, TIME_SPECTRAL) UNSTEADY_SIMULATION= NO % CONSOLE = CONCISE %  INCOMPRESSIBLE FREESTREAM DEFINITION % % % Mach number % MACH_NUMBER= 0.15 % Angle of attack (deg) % AoA= 5.0 % Freestream density (1.2886 Kg/m^3 (air), 998.2 Kg/m^3 (water)) FREESTREAM_DENSITY= 2.13163 % % Freestream velocity (m/s) FREESTREAM_VELOCITY= ( 10.0, 0.00, 0.00 ) % % Freestream viscosity (1.853E5 Ns/m^2 (air), 0.798E3 Ns/m^2 (water)) FREESTREAM_VISCOSITY= 1.853e05 %  REFERENCE VALUE DEFINITION % % % 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= 1.0 % % Reference area for force coefficients (0 implies automatic calculation) REF_AREA= 1.0 %%%%% % MOTION %GRID_MOVEMENT = YES %GRID_MOVEMENT_KIND = NONE %GRID_MOVEMENT_KIND = ROTATING_FRAME %MACH_MOTION = 0.7958 %MARKER_MOVING = NONE %MOTION_ORIGIN_X = 0.5 %MOTION_ORIGIN_Y = 32.0 %MOTION_ORIGIN_Z = 0.0 %ROTATION_RATE_X = 0.0 %ROTATION_RATE_Y = 0.0 %ROTATION_RATE_Z = 0.0 %8.25 %  BOUNDARY CONDITION DEFINITION % % % NavierStokes wall boundary marker(s) (NONE = no marker) MARKER_HEATFLUX= ( airfoil, 0.0 ) % % Farfield boundary marker(s) (NONE = no marker) MARKER_FAR= ( farfield ) % % Marker(s) of the surface to be plotted or designed MARKER_PLOTTING= ( airfoil ) % % Marker(s) of the surface where the functional (Cd, Cl, etc.) will be evaluated MARKER_MONITORING= ( airfoil ) %  COMMON PARAMETERS DEFINING THE NUMERICAL METHOD % % % Numerical method for spatial gradients (GREEN_GAUSS, WEIGHTED_LEAST_SQUARES) NUM_METHOD_GRAD= WEIGHTED_LEAST_SQUARES % % CourantFriedrichsLewy condition of the finest grid CFL_NUMBER= 10.0 % % Adaptive CFL number (NO, YES) CFL_ADAPT= NO % % Parameters of the adaptive CFL number (factor down, factor up, CFL min value, % CFL max value ) CFL_ADAPT_PARAM= ( 1.5, 0.5, 1.0, 100.0 ) % % Number of total iterations EXT_ITER= 75000 %  SLOPE LIMITER DEFINITION % % % Reference element length for computing the slope and sharp edges limiters. REF_ELEM_LENGTH= 0.1 % % Coefficient for the limiter LIMITER_COEFF= 0.1 % % % Reference coefficient (sensitivity) for detecting sharp edges. REF_SHARP_EDGES= 3.0 % % Remove sharp edges from the sensitivity evaluation (NO, YES) SENS_REMOVE_SHARP= NO %  LINEAR SOLVER DEFINITION % % % Linear solver for implicit formulations (BCGSTAB, FGMRES) LINEAR_SOLVER= FGMRES % % Preconditioner of the Krylov linear solver (JACOBI, LINELET, LU_SGS) LINEAR_SOLVER_PREC= LU_SGS % % Minimum error of the linear solver for implicit formulations LINEAR_SOLVER_ERROR= 1E4 % % Max number of iterations of the linear solver for the implicit formulation LINEAR_SOLVER_ITER= 5 %  MULTIGRID PARAMETERS % % % MultiGrid Levels (0 = no multigrid) MGLEVEL= 0 % % Multigrid cycle (V_CYCLE, W_CYCLE, FULLMG_CYCLE) MGCYCLE= V_CYCLE % % Multigrid presmoothing level MG_PRE_SMOOTH= ( 1, 2, 3, 3 ) % % 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= 0.75 % % Damping factor for the correction prolongation MG_DAMP_PROLONGATION= 0.75 %  FLOW NUMERICAL METHOD DEFINITION % % % Convective numerical method (JST, LAXFRIEDRICH, CUSP, ROE, AUSM, HLLC, % TURKEL_PREC, MSW) CONV_NUM_METHOD_FLOW= ROE % % 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 % % 1st, 2nd and 4th order artificial dissipation coefficients AD_COEFF_FLOW= ( 0.15, 0.5, 0.02 ) % % Time discretization (RUNGEKUTTA_EXPLICIT, EULER_IMPLICIT, EULER_EXPLICIT) TIME_DISCRE_FLOW= EULER_IMPLICIT %  ADJOINTFLOW NUMERICAL METHOD DEFINITION % % Adjoint problem boundary condition (DRAG, LIFT, SIDEFORCE, MOMENT_X, % MOMENT_Y, MOMENT_Z, EFFICIENCY, % EQUIVALENT_AREA, NEARFIELD_PRESSURE, % FORCE_X, FORCE_Y, FORCE_Z, THRUST, % TORQUE, FREE_SURFACE, TOTAL_HEAT, % MAXIMUM_HEATFLUX, INVERSE_DESIGN_PRESSURE, % INVERSE_DESIGN_HEATFLUX) OBJECTIVE_FUNCTION= DRAG % % Convective numerical method (JST, LAXFRIEDRICH, ROE) CONV_NUM_METHOD_ADJFLOW= ROE % % Spatial numerical order integration (1ST_ORDER, 2ND_ORDER, 2ND_ORDER_LIMITER) % SPATIAL_ORDER_ADJFLOW= 2ND_ORDER % % Slope limiter (VENKATAKRISHNAN, SHARP_EDGES) SLOPE_LIMITER_ADJFLOW= VENKATAKRISHNAN % % Coefficient for the sharp edges limiter SHARP_EDGES_COEFF= 3.0 % % 1st, 2nd, and 4th order artificial dissipation coefficients AD_COEFF_ADJFLOW= ( 0.15, 0.0, 0.002 ) % % Reduction factor of the CFL coefficient in the adjoint problem CFL_REDUCTION_ADJFLOW= 0.25 % % Time discretization (RUNGEKUTTA_EXPLICIT, EULER_IMPLICIT) TIME_DISCRE_ADJFLOW= EULER_IMPLICIT % % Adjoint frozen viscosity (NO, YES) FROZEN_VISC= YES %  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 % % Time discretization (EULER_IMPLICIT) TIME_DISCRE_TURB= EULER_IMPLICIT %  CONVERGENCE PARAMETERS % % % Convergence criteria (CAUCHY, RESIDUAL) % CONV_CRITERIA= CAUCHY % % 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= 10 % % 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= 8E6 % % Function to apply the criteria (LIFT, DRAG, NEARFIELD_PRESS, SENS_GEOMETRY, % SENS_MACH, DELTA_LIFT, DELTA_DRAG) CAUCHY_FUNC_FLOW= LIFT %CAUCHY_FUNC_LIN= DELTA_DRAG %  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 % % Before each computation do an implicit smoothing of the nodes coord (NO, YES) SMOOTH_GEOMETRY= NO %  INPUT/OUTPUT INFORMATION % % % Mesh input file %MESH_FILENAME= mesh_NACA0012_turb_897x257.su2 MESH_FILENAME = mesh_AIRFOIL.su2 % % Mesh input file format (SU2, CGNS, NETCDF_ASCII) MESH_FORMAT= SU2 % % Mesh output file MESH_OUT_FILENAME= mesh_out.su2 % % Restart flow input file SOLUTION_FLOW_FILENAME= solution_flow.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 restart flow RESTART_FLOW_FILENAME= solution_flow.dat % % Output file restart adjoint RESTART_ADJ_FILENAME= restart_adj.dat % % Output file flow (w/o extension) variables VOLUME_FLOW_FILENAME= flow % % Output file adjoint (w/o extension) variables VOLUME_ADJ_FILENAME= adjoint % % 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 % % Writing solution file frequency WRT_SOL_FREQ= 5000 % % Writing convergence history frequency WRT_CON_FREQ= 1000 DV_KIND = HICKS_HENNE DV_MARKER = ( airfoil ) DV_PARAM = ( 1, 0.5 ) DEFINITION_DV= ( 1, 1.0  airfoil  0, 0.05 ); ( 1, 1.0  airfoil  0, 0.10 ); ( 1, 1.0  airfoil  0, 0.15 ); ( 1, 1.0  airfoil  0, 0.20 ); ( 1, 1.0  airfoil  0, 0.25 ); ( 1, 1.0  airfoil  0, 0.30 ); ( 1, 1.0  airfoil  0, 0.35 ); ( 1, 1.0  airfoil  0, 0.40 ); ( 1, 1.0  airfoil  0, 0.45 ); ( 1, 1.0  airfoil  0, 0.50 ); ( 1, 1.0  airfoil  0, 0.55 ); ( 1, 1.0  airfoil  0, 0.60 ); ( 1, 1.0  airfoil  0, 0.65 ); ( 1, 1.0  airfoil  0, 0.70 ); ( 1, 1.0  airfoil  0, 0.75 ); ( 1, 1.0  airfoil  0, 0.80 ); ( 1, 1.0  airfoil  0, 0.85 ); ( 1, 1.0  airfoil  0, 0.90 ); ( 1, 1.0  airfoil  0, 0.95 ); ( 1, 1.0  airfoil  1, 0.05 ); ( 1, 1.0  airfoil  1, 0.10 ); ( 1, 1.0  airfoil  1, 0.15 ); ( 1, 1.0  airfoil  1, 0.20 ); ( 1, 1.0  airfoil  1, 0.25 ); ( 1, 1.0  airfoil  1, 0.30 ); ( 1, 1.0  airfoil  1, 0.35 ); ( 1, 1.0  airfoil  1, 0.40 ); ( 1, 1.0  airfoil  1, 0.45 ); ( 1, 1.0  airfoil  1, 0.50 ); ( 1, 1.0  airfoil  1, 0.55 ); ( 1, 1.0  airfoil  1, 0.60 ); ( 1, 1.0  airfoil  1, 0.65 ); ( 1, 1.0  airfoil  1, 0.70 ); ( 1, 1.0  airfoil  1, 0.75 ); ( 1, 1.0  airfoil  1, 0.80 ); ( 1, 1.0  airfoil  1, 0.85 ); ( 1, 1.0  airfoil  1, 0.90 ); ( 1, 1.0  airfoil  1, 0.95 ) % DEFINITION_DV = (( 1, 1.0  airfoil  0, 0.961538461538 ); ( 1, 1.0  airfoil  0, 0.923076923077 ); ( 1, 1.0  airfoil  0, 0.884615384615 ); ( 1, 1.0  airfoil  0, 0.846153846154 ); ( 1, 1.0  airfoil  0, 0.807692307692 ); ( 1, 1.0  airfoil  1, 0.846153846154 ); ( 1, 1.0  airfoil  1, 0.884615384615 ); ( 1, 1.0  airfoil  1, 0.923076923077 ); ( 1, 1.0  airfoil  1, 0.961538461538 ) 

February 22, 2016, 20:03 
Experienced similar things

#2 
Disabled
Join Date: Jul 2015
Posts: 7
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Hi,
For what it's worth, I have noticed the same thing. The drag is always higher than the test data I have from an old NACA report. Cl and Cm vary from too high to too low, depending on the solver choice and CFL number. I was using SU2_EDU. 

February 23, 2016, 00:07 

#3 
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Eduardo Molina
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Hi...
Let me ask you some questions. Did your simulation converged? Why you are using Cauchy converge criteria? Can you show us the density residual? Cheers Molina 

February 23, 2016, 17:38 

#4  
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Ryan Barrett
Join Date: May 2015
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Quote:
Res[Press] Res[nu] 7.827376 10.677632 

February 24, 2016, 11:01 

#5 
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Mike
Join Date: Oct 2015
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Hi
Is the reference area you've set definitely correct/the same as you've used for the wind tunnel/other experiments? You currently have it set to the default 1m^2 (REF_AREA= 1.0) Cheers, M. 

February 24, 2016, 15:54 

#6 
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Ryan Barrett
Join Date: May 2015
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Yes. Everything was normalized to 1.0 for all the calculations that are compared. My lift matches pretty well, it is my drag that is quite high in comparison.


October 11, 2016, 04:58 

#7 
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luigi
Join Date: Oct 2016
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Hi,
I have to calculate Cl Cd for different type of airfoils to vary the angle of attack (alpha), I do in this way: naca0018 pane oper visc 100000 iter 1000 alfa 0 pacc naca0018.pol aseq 30.0 30.0 2.0 pacc quit The results are different from the datas, maybe cause the large variation of alpha. what do you think? 

October 13, 2016, 16:35 

#8 
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Heather Kline
Join Date: Jun 2013
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It looks like this hasn't been mentioned on this thread yet:
Since the panel method and data match well, but the turbulent viscous simulation does not, it could be that the issue is with the turbulence  either that the flow in the experiment was closer to laminar, or that the grid is not refined enough near the wall to accurately capture the boundary layer. On the angle of attack sweep from 30 to 30 degrees  that is a wide enough range that the flow would certainly be separated and not likely to be providing an accurate solution towards the ends of that range. (unless using very high fidelity CFD  ie unsteady LES methods) 

October 14, 2016, 03:58 

#9 
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Just some other points:
1. did you set the transition point in XFOIL? If not the code will predict a "free" transition, conversely SU2 is fully turbulent. A comparison in terms of cf could be interesting 2. the turbulence model is solved as 1st order, did you check with 2nd order accuracy? 

October 14, 2016, 06:29 

#10 
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luigi
Join Date: Oct 2016
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Hi,
I didn't set the transition point and I didn't check with 2nd order accuracy, the only things that I write in the prompt are: naca0018 pane oper visc 100000 iter 1000 alfa 0 pacc naca0018.pol aseq 30.0 30.0 2.0 pacc quit I read in a document that is better using the command "init" for a large range of alpha and I don't know if the command "alpha 0" (after iter) is necessary. What do you think? I tryed the same for different Reynold's Numbers and I have the same problems, maybe the issues are linked to an error in my script. Thank you very much! 

October 14, 2016, 07:53 

#11  
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Quote:
The script itself is not wrong, anyway you are performing a free transition analysis (so keep in mind that when comparing the results). Could you post the details about SU2 simulation, like y+ and the input file? 

October 15, 2016, 03:15 

#12 
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luigi
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I don't have an SU2 simulation; I had experimental datas (maybe from Sandia Lab) to be compared with my xfoil's results.


October 19, 2016, 13:16 

#13 
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luigi
Join Date: Oct 2016
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Hi,
I have another question, like you can see on my script there is the command alpha 0 (after iter 1000); I used this command cause there was in a tutorial but I don't know the utility, I thought that was only to see the Cp at alpha 0 but then I noticed that, with this command (alpha 0), my final results are less correct. Do you know its function? thank you very much! 

October 19, 2016, 13:25 

#14 
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Heather Kline
Join Date: Jun 2013
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For more information on running and interpreting xfoil data, I would suggest seeking out a similar support forum for that code, or seeing what their contact/support suggestions are on their website.


October 20, 2016, 14:03 

#15 
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luigi
Join Date: Oct 2016
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Hi,
I have another question, like you can see on my script there is the command alpha 0 (after iter 1000); I used this command cause there was in a tutorial but I don't know the utility, I thought that was only to see the Cp at alpha 0 but then I noticed that, with this command (alpha 0), my final results are less correct. Do you know its function? thank you very much! About this what do you think? 

October 30, 2016, 17:27 

#16 
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C
Join Date: Apr 2013
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Hello all,
The problem here is possibly because that the flow is transitional (Re = 1.1M and freestream turbulence intensity is around 0.1%). You should solve the case with a transition model. Recently, I have added my own transition model (named BC (BasCakmakcioglu) Model) to the code, and I have opened a pull request (#326). Maybe you can test this case with the BC model, and see the results. Hopefully, the BC transition model will be merged into the master branch after some testing. Best, S. Cakmakcioglu 

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drag, su2, xfoil 
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