Configuring MATLAB

The configuration should have been automatically completed when you obtained your account on conductor, so whenever you start a matlab session, you will be able to run SOUSSA.  However, SOUSSA requires a large disk space to run; normally, your account will not have enough disk space.  Therefore, you must make a temporary directory in a location where there is ample room.

For conductor-am or engc:

mkdir   /tmp/<username>
cd         /tmp/<username>
matlab531

Now, you are all set to run SOUSSA.  Remember to change to the temporary directory before starting matlab each time you log in before running SOUSSA.

Executing SOUSSA

Run MATLAB version 5.3.1 by typing matlab531 at the prompt of one of the unix windows.  Now at the MATLAB prompt type

soussa

An input panel will pop up.  The program that creates this panel is  called a GUI.  If you are interested in running a rectangular wing geometry with a basic airfoil shape you can do this right from the input panel!   For specifying non-rectangular wing geometries, see the section on Geometry Specification.  For more advanced SOUSSA usage, see the Input Specification section.

Info on running SOUSSA in MATLAB can be viewed at any time by typing 'help soussa' from the MATLAB prompt.
 

GUI Input:

To run a rectangular wing,  change the parameters as desired.  The "Airfoil Type" describes the profile used for the hub and tip of the wing and assumes a  4- or 5-series NACA profile.  For example, if you want to run a NACA4412 enter 4412 in the box denoted AFtype.

The definitions and default values for the input parameters are:

Number of wings: will start as 0, but when it is computing it should always be 1 (if you are doing a flapped airfoil then it will be 2 or 3 depending on how many flaps you have)

Nx           Number of streawise panels
Ny           Number of Spanwise panels
AFtype       Airfoil section definition
Mach         Mach number (U_o / c)
             (c - speed of sound)
AOA          Angle of attack of entire wing
             (in degrees)
Nose         Leading edge position, '[x, z]' (important for flap)
Size         x and z airfoil scaling factor
             (Always 1 for single wing) (important for flap)
Flap ang.    Deflection angle of current element
             (in degrees) (only for flapped wings etc.)
Span         Span (normalized by root chord length)
                Works together with size to create flaps
                of various chord lengths (based on root
                chord length) and span.
Dihedral     Spanwise elevation angle in degrees. Positive
                means wing tip is higher than wing root
Taper ratio  Ratio of tip chord length to root chord length
Sweep        Measured in degrees from perpendicular to
                leading edge of root chord.  Positive means
                that wing tip is further back than wing root.
 

Defaults are:
      Nx           16
      Ny           11
      Nwake        1
      AFtype       '0012'
      Mach         0.0
      AOA          0
      Nose         '[0 0]'
      Size         1
      Flap ang.    0
      Span         6
      Dihedral     0
      Taper ratio  1
      Sweep        0

The default values for Nx and Ny  should not need to be changed.  You can experiment with Nx and Ny to see if the answers change much.

When you add the wing element, you will see a graphical representation of the wing that you created.  If it is not the wing you desire, delete the element and try again.  Additional wing elements such as flaps may be added to a main wing element.  Set the flap geometry making sure to state where the nose of the flap root chord lies and specify the size of the flap by using size (scales chord and span based on main wing element) and span (will affect only the flap element).  You can note that the default arguments are set to the previous element added.

AFtype can be either Parabolic or NACA (default).  Parabolic airfoils must include one argument for percent thickness.  First argument of NACA airfoils is the standard four or five digit tag.  If 'cut' is included in the argument list of AFtype, a flap "cutout" is effected between 60% and 80% chord stations; these values may be changed via two arguments following 'cut' token.

Airfoils that you can choose are shown below.  For the NACA 4- and 5-series enter numbers for the x's into the box denoted AFtype.

 

SOUSSA name	Description
NACA xxxx	NACA 4-series
NACA 2x0xx	NACA 5-series
N63215	NACA 63(2)-215
NLR7301	NLR-7301 transonic airfoil

Run the flow around the specified geometry and interpret the results.

Once you have the input panel set the way you want, press the add button at the bottom of the panel, and then the compute button.  The flow past the wing will be computed and an output screen will appear.  You will see one half of the wing on the top and a plot of the coefficient of pressure along the centerspan on the bottom.  The colors on the wing figure indicate the coefficient of pressure.  To view the velocity you can click on the top right button and choose the velocity component you would like to view.  To just see the geometry you can choose geometry.  The airfoil section coefficient of pressure plot shows the coefficient of pressure on the top of the airfoil in blue and on the bottom of the airfoil in red.  To see the coefficient of pressure at different locations along the span use the slider bar below the lift and drag values.

The lift and drag values denoted CL_A and CD_i_A are the coefficient of wing lift and wing induced drag, where, CL_A is defined by

CL = L / ( 0.5 rho U^2 A)

with 'A' being the total surface area (not the same as projected planform area normally used to define CL) which is also displayed in the force window.  For very thin and flat wings, 'A' would be approximately twice the planform area.  In general, the commonly defined Lift Coefficient would be obtained by

CL_normal = CL_A * A / S

where S is the projected planform area (based on a chord of 1).  If the chord of the real wing you are analyzing is not 1, then you can obtain CL_normal = CL_A * A / (S / c_root^2)

Note then that the results for CL for a chord length of 1 and span length of 6 will be identical to the results for a chord of 4m and span of 24m, for example, because of the dimensionless scaling of the pressure data.

Also remember that the CD_i_A  is only the inviscid drag.  You must still account for parasite drag.  Again you can recover the usual value of CD_i in the same manner as described above for CL.

If you want to change the wing specifications, press the delete button at the bottom of the input panel, make the appropriate changes to the input panel, press the add button, then press the compute button.  If you do not press the delete button, the number of wings (indicated at the top fo the input panel) will start to increase and time for calculation will increase and the results will be erroneous.

NOTE:

SOUSSA can also be run in batch mode by supplying multiple input arguments.  This is achieved by calling soussa with 2 arguments for each parameter to be changed.  The first argument is a string flag and the second is the input parameter value.  It is never necessary to run SOUSSA in this manner if you are only interested in analyzing a rectangular wing.  The following examples show what you would type at the MATLAB prompt to calculate the flow past wings if you did not want to enter values into the input panel.

EXAMPLES:

Simple NACA 23012 wing (AR=6) at 5 degrees angle of attack using 21 panels along the chord and 7 panels along the span for both the upper and lower surfaces...

soussa('AFtype','NACA23012','AOA',5,'AR',6,'Nx',21,'Ny',7)

Set the main element to the default and add a 30 % chord flap at 15 degree flap deflection with nose positioned (0.05,-0.1) relative to trailing edge of main wing...

soussa('flap','Nose','[1.05,-.1]','alpha',15,'size',0.3)

Substitute main element in 1) with a NACA23012 section...

soussa('AFtype','23012','Nx',21,'flap','Nose','[1,-.03]',...
           'AFtype','6212', 'Nx',15,'alpha',15,'size',0.3)    soussa('AFtype','9616','Nx',13,'size',0.8,'flap','Nose','[0.76,-.045]',...
           'AFtype','9616','Nx',9 , 'size',0.3,'alpha',32, 'AOA',10)   soussa('AFtype','0012 cutout','Nx',41,'flap','Nose','[0.77,-0.01]',...
           'AFtype','6214',       'Nx',25,'alpha',15,'size',0.3,'AOA',5)

Thin, symmetric (apprx. Thoedorsen)

soussa('SOUSSAmode','f','AFtype','para .1','Nx',21,'Ny',11,'AOA',0,...
            'AR',20,'Nwake',20,'Lwake',5,'freq',0.47,'BCmode',2)

Symmetric Parabolic

soussa('SOUSSAmode','f','AFtype','PARA 5','Nx',10,'Ny',10,...
        'AOA',0,'Mach',0.24,'AR',3,'freq','[0.42]','Nwake',20,'Lwake',2.5)

Split Flap

soussa('AFtype','0012','Nx',21,'Ny',10,'AR',6,...
            'Span','[.5 1]','flap',...
            'AFtype','0012 cutout','Nx',21,'Ny',10,'Span','[0 .5]',...
            'flap','AFtype','6214','Nx',15,'Ny',10,'Span','[0 .5]',...
            'size',0.3,'Nose','[0.77,-.01]','alpha',15)

For alternate specification of input geometry, direct editting of the file 'soussa.inp' is required.  See  the section on Input Specification  for further details.