#Naca supersonic airfoil generator series#
#Naca supersonic airfoil generator download#

In the right hand graph the moment curve shown is for the moment at the вЂњaerodynamic centerвЂќ rather than the quarter chord but since it is also zero in this plot it confirms the theoretical prediction that for a symmetrical airfoil the center of pressure (where the moment is zero) coincides with the aerodynamic center.ДЇigure A-2 gives similar data for the NACA 2412 airfoil, another 12% thick shape but one with camber. Also note that the pitching moment coefficient at c/4 (in the left hand graph) is essentially zero from -12 degrees to+ 14 degrees angle of attack and then goes negative in stall at positive angle of attack. Note that the flap deflection shifts the lift curve far to the left giving a zero lift angle of attack of roughly minus 12 degrees while it increases the maximum lift coefficient (Re= 6 x 106) from just under 1.6 to 2.4, a huge increase in lifting capability that can contribute to large decreases in takeoff and landing distances. They also show what happens when a 20% chord flap is deflected 40 degrees. These graphs show test results for several different Reynolds numbers and for вЂњstandard roughnessвЂќ on the surface. Note that for the symmetrical shape the lift coefficient is zero at zero angle of attack. The airfoils presented represent a cross section of airfoil shapes selected to illustrate why one would select one airfoil over another for any given aircraft design or performance requirement.ДЇigure A-1 shows data for the NACA 0012 airfoil, a classic symmetrical shape that is used for everything from airplane stabilizers and canards to helicopter rotors to submarine вЂњsailsвЂќ. In the following appendix material a selection of airfoil graphical data is presented which can be found in the Theory of Wing Sections and in the non-copyrighted NACA publications which are the source of the Dover publicationвЂ™s data. While the date of original publication might lead one to think this material must be out of date, that is simply not true and the Theory of Wing Sections is one of the most valuable references in any aerospace engineerвЂ™s personal library. Many of the more important airfoil shapes have their test results summarized in the Theory of Wing Sections, a Dover paperback publication authored by Ira Abbott and Albert Von Doenhoff and first published in 1949. This data is most conveniently presented in plots of lift coefficient versus angle of attack, pitching moment coefficient versus angle of attack, drag coefficient versus lift coefficient, and pitching moment coefficient versus lift coefficient and is found in literally hundreds of NACA and NASA Reports, Notes, and Memoranda published since the 1920s. Lift, drag, and pitching moment data for hundreds of such airfoil shapes was determined in wind tunnel tests by the National Advisory Committee for Aeronautics (NACA) and later by NASA, the National Aeronautics and Space Administration.

#Naca supersonic airfoil generator series#

2.In Chapter 3 of this text we discussed many of the aspects of airfoil design as well as the NACA designations for several series of airfoils. The evolution of pressure and shape during the optimization is as follows.ДЇig. According to вЂњlogOpt.txtвЂќ and вЂњopt_SLSQP.txtвЂќ, the initial drag is 0.015226161 and the optimized drag is 0.013617362 with a drag reduction of 10.6%. The case ran for 50 steps and took about 25 minutes using Intel 3.0 GHz CPU with 4 cores. Mpirun -np 4 python runScript.py 2>&1 | tee logOpt.txt Then, use the following command to run the optimization with 4 CPU cores: Then go to tutorials-master/NACA0012_Airfoil/subsonic and run the вЂњpreProcessing.shвЂќ script to generate the mesh:

#Naca supersonic airfoil generator download#

To run this case, first download tutorials and untar it. This is done by setting lower and upper bounds in вЂњprimalVarBoundsвЂќ for all variables. To ensure a robust convergence, we need to bound the variables when solving the primal. Accordingly, we set вЂњflowConditionвЂќ to вЂњCompressibleвЂќ. We use вЂњDARhoSimpleFoamвЂќ, which is an OpenFOAM built-in compressible flow solver suitable for subsonic conditions. In вЂњprimalBCвЂќ, we need to set boundary condition for temperature вЂњT0вЂќ. In addition, we provide the absolute value of pressure вЂњpвЂќ, instead of the relative value use in the low speed case. In the global parameters, we provide additional variable such as вЂњT0вЂќ (far field temperature) and вЂњrho0вЂќ (a reference density to normalize CD and CL). The вЂњrunScript.pyвЂќ is similar to the one used in the NACA0012 low speed case with the following exceptions: Mesh and FFD points for the NACA0012 airfoil Objective function: Drag coefficient (CD)Дesign variables: 20 free-form deformation (FFD) points moving in the y direction, one angle of attackД¬onstraints: Symmetry, volume, thickness, and lift constraints (total number: 34)ДЇig. The following is an aerodynamic shape optimization case for the NACA0012 airfoil in subsonic conditions. Note: We recommend going through the tutorial in Get started before running this case.