Change all tutorial image paths in the tutorial markdown files.

Author: TobiKattmann

_tutorials/compressible_flow/Inviscid_Bump/Inviscid_Bump.md

@@ -11,8 +11,8 @@ requires: SU2_CFD
 complexity: basic
 follows: 
 ---
-
-![Channel Mach](../../Inviscid_Bump/images/channel_mach.png)
+[//]: # (Example of a currently working link to a picture)
+![Channel Mach](../../tutorials_files/compressible_flow/Inviscid_Bump/images/channel_mach.png)
 
 ## Goals
 
@@ -52,7 +52,7 @@ There is also a set of inlet/outlet conditions for transonic flow available in t
 
 The channel is of length 3L with a height L and a circular bump centered along the lower wall with height 0.1L. For the SU2 mesh, L = 1.0 was chosen, as seen in the figure of the mesh below. The mesh is composed of quadrilaterals with 256 nodes along the length of the channel and 128 nodes along the height. The following figure contains a view of the mesh topology (a coarser mesh is shown for clarity).
 
-![Channel Mesh](../../Inviscid_Bump/images/channel_mesh_bcs.png)
+![Channel Mesh](../../tutorials_files/compressible_flow/Inviscid_Bump/images/channel_mesh_bcs.png)
 Figure (1): The computational mesh with boundary conditions highlighted.
 
 The boundary conditions for the channel are also highlighted in the figure. Inlet, outlet, and Euler wall boundary conditions are used. The Euler wall boundary condition enforces flow tangency at the upper and lower walls. 
@@ -152,8 +152,8 @@ The channel simulation for the 256x128 node mesh is relatively small, so this ca
 
 The following images show some SU2 results for the inviscid channel problem.
 
-![Channel Mach](../../Inviscid_Bump/images/channel_mach.png)
+![Channel Mach](../../tutorials_files/compressible_flow/Inviscid_Bump/images/channel_mach.png)
 Figure (2): Mach number contours for the 2D channel.
 
-![Channel Pressure](../../Inviscid_Bump/images/channel_pressure.png)
+![Channel Pressure](../../tutorials_files/compressible_flow/Inviscid_Bump/images/channel_pressure.png)
 Figure (3): Pressure contours for the 2D channel.

_tutorials/compressible_flow/Inviscid_ONERAM6/Inviscid_ONERAM6.md

@@ -12,7 +12,7 @@ complexity: basic
 follows: 
 ---
 
-![ONERA M6 Cp](../../Inviscid_ONERAM6/images/oneram6_cp.png)
+![ONERA M6 Cp](../../tutorials_files/compressible_flow/Inviscid_ONERAM6/images/oneram6_cp.png)
 
 ## Goals
 
@@ -52,10 +52,10 @@ These transonic flow conditions will cause the typical "lambda" shock along the
 
 The computational domain is a large parallelepiped with the wing half-span mounted on one boundary in the x-z plane. The mesh consists of 582,752 tetrahedral elements and 108,396 nodes. Three boundary conditions are employed: Euler wall on the wing surface, a far-field characteristic-based condition on the far-field markers, and a symmetry boundary condition for the marker where the wing half-span is attached. The symmetry condition acts to mirror the flow about the x-z plane, reducing the complexity of the mesh and the computational cost. Images of the entire domain and the triangular elements on the wing surface are shown below.
 
-![ONERA M6 Mesh](../../Inviscid_ONERAM6/images/oneram6_mesh_bcs.png)
+![ONERA M6 Mesh](../../tutorials_files/compressible_flow/Inviscid_ONERAM6/images/oneram6_mesh_bcs.png)
 Figure (1): Far-field view of the computational mesh with boundary conditions.
 
-![ONERA M6 Surface Mesh](../../Inviscid_ONERAM6/images/oneram6_wing_mesh.png)
+![ONERA M6 Surface Mesh](../../tutorials_files/compressible_flow/Inviscid_ONERAM6/images/oneram6_wing_mesh.png)
 Figure (2): Close-up view of the unstructured mesh on the top surface of the ONERA M6 wing.
 
 ### Configuration File Options
@@ -189,14 +189,14 @@ If SU2 has been built with parallel support, the recommended method for running
 
 Results are here given for the SU2 solution of inviscid flow over the ONERA M6 wing.
 
-![ONERA M6 Cp](../../Inviscid_ONERAM6/images/oneram6_cp.png)
+![ONERA M6 Cp](../../tutorials_files/compressible_flow/Inviscid_ONERAM6/images/oneram6_cp.png)
 Figure (3): Cp contours on the upper surface of the ONERA M6.
 
-![ONERA M6 Mach](../../Inviscid_ONERAM6/images/oneram6_mach.png)
+![ONERA M6 Mach](../../tutorials_files/compressible_flow/Inviscid_ONERAM6/images/oneram6_mach.png)
 Figure (4): Mach number contours on the upper surface of the ONERA M6 wing. Notice the "lambda" shock pattern typically seen on the upper surface.
 
-![ONERA M6 Coefficients](../../Inviscid_ONERAM6/images/oneram6_coefficients.png)
+![ONERA M6 Coefficients](../../tutorials_files/compressible_flow/Inviscid_ONERAM6/images/oneram6_coefficients.png)
 Figure (5): Convergence of the non-dimensional coefficients.
 
-![ONERA M6 Convergence](../../Inviscid_ONERAM6/images/oneram6_convergence.png)
+![ONERA M6 Convergence](../../tutorials_files/compressible_flow/Inviscid_ONERAM6/images/oneram6_convergence.png)
 Figure (6): Convergence of the density residual (speed up x20, iteration based).

_tutorials/compressible_flow/Inviscid_Wedge/Inviscid_Wedge.md

@@ -12,7 +12,7 @@ complexity: basic
 follows: 
 ---
 
-![Wedge Mach](../../Inviscid_Wedge/images/wedge_mach.png)
+![Wedge Mach](../../tutorials_files/compressible_flow/Inviscid_Wedge/images/wedge_mach.png)
 
 ## Goals
 
@@ -49,7 +49,7 @@ This problem will solve for the flow over the wedge with these conditions:
 
 The wedge mesh is a structured mesh (75x50) of rectangular elements with a total of 3,750 nodes. The upper and lower wall of the geometry are solid (`MARKER_EULER`), and the lower wall has a 10 degree wedge starting at x = 0.5. Figure (1) shows the mesh with the boundary markers and flow conditions highlighted.
 
-![Wedge Mach](../../Inviscid_Wedge/images/wedge_mesh_bcs.png)
+![Wedge Mach](../../tutorials_files/compressible_flow/Inviscid_Wedge/images/wedge_mesh_bcs.png)
 Figure (1): The computational mesh with boundary conditions highlighted.
 
 For this test case, the inlet marker will be set to a `MARKER_SUPERSONIC_INLET` boundary condition, while the outlet marker will be set to the `MARKER_OUTLET` condition. In supersonic flow, all characteristics are incoming to the domain at the entrance (inlet marker), and therefore, all flow quantities can be specified, i.e., no information travels upstream. 
@@ -174,8 +174,8 @@ The wedge simulation is small and will execute quickly on a single workstation o
 
 The following images show some SU2 results for the supersonic wedge problem.
 
-![Wedge Mach](../../Inviscid_Wedge/images/wedge_mach.png)
+![Wedge Mach](../../tutorials_files/compressible_flow/Inviscid_Wedge/images/wedge_mach.png)
 Figure (2): Mach contours showing the oblique shock for supersonic flow over a wedge.
 
-![Wedge Pressure](../../Inviscid_Wedge/images/wedge_pressure.png)
+![Wedge Pressure](../../tutorials_files/compressible_flow/Inviscid_Wedge/images/wedge_pressure.png)
 Figure (3): Pressure contours (N/m2) for supersonic flow over a wedge.

_tutorials/compressible_flow/Laminar_Cylinder/Laminar_Cylinder.md

@@ -12,7 +12,7 @@ complexity: basic
 follows: 
 ---
 
-![Cylinder Mach](../../Laminar_Cylinder/images/cylinder_mach.png)
+![Cylinder Mach](../../tutorials_files/compressible_flow/Laminar_Cylinder/images/cylinder_mach.png)
 
 ## Goals
 
@@ -53,7 +53,7 @@ This problem will solve the for the external, compressible flow over the cylinde
 
 The problem geometry is 2D. The mesh has 26,192 triangular elements and 13,336 points. It is fine near the surface of the cylinder to resolve the boundary layer. The exterior boundary is approximately 15 diameters away from the cylinder surface to avoid interaction between the boundary conditions. Far-field boundary conditions are used at the outer boundary. No-slip boundary conditions are placed on the surface of the cylinder. 
 
-![Cylinder Mesh](../../Laminar_Cylinder/images/cylinder_mesh.png)
+![Cylinder Mesh](../../tutorials_files/compressible_flow/Laminar_Cylinder/images/cylinder_mesh.png)
 Figure (1): The computational mesh for the 2D cylinder test case. 
 
 The outer boundary in red is the far-field, and the small circle in the center is the cylinder which uses the Navier-Stokes Wall boundary condition.
@@ -105,11 +105,11 @@ The cylinder simulation for the 13,336 node mesh is small and will execute relat
 
 The following results show the flow around the cylinder as calculated by SU2 (note that these were for a slightly higher Mach number of 0.3).
 
-![Cylinder Pressure](../../Laminar_Cylinder/images/cylinder_pressure.png)
+![Cylinder Pressure](../../tutorials_files/compressible_flow/Laminar_Cylinder/images/cylinder_pressure.png)
 Figure (2): Pressure contours around the cylinder.
 
-![Cylinder Viscosity](../../Laminar_Cylinder/images/cylinder_lam_visc.png)
+![Cylinder Viscosity](../../tutorials_files/compressible_flow/Laminar_Cylinder/images/cylinder_lam_visc.png)
 Figure (3): Laminar viscosity contours for this steady, low Reynolds number flow.
 
-![Cylinder Mach](../../Laminar_Cylinder/images/cylinder_mach.png)
+![Cylinder Mach](../../tutorials_files/compressible_flow/Laminar_Cylinder/images/cylinder_mach.png)
 Figure (4): Mach number contours around the cylinder with streamlines. Note the large laminar separation region behind the cylinder at Re = 40.

_tutorials/compressible_flow/Laminar_Flat_Plate/Laminar_Flat_Plate.md

@@ -12,7 +12,7 @@ complexity: basic
 follows: 
 ---
 
-![Lam Plate Profile](../../Laminar_Flat_Plate/images/lam_plate_velocity_profile.png)
+![Lam Plate Profile](../../tutorials_files/compressible_flow/Laminar_Flat_Plate/images/lam_plate_velocity_profile.png)
 
 ## Goals
 
@@ -41,11 +41,11 @@ The following tutorial will walk you through the steps required when solving for
 
 In his PhD dissertation in 1908, H. Blasius obtained what is now referred to as the Blasius equation for incompressible, laminar flow over a flat plate:
 
-![Blasius Equation](../../Laminar_Flat_Plate/images/blasius.png)
+![Blasius Equation](../../tutorials_files/compressible_flow/Laminar_Flat_Plate/images/blasius.png)
 
 The third-order, ordinary differential equation can be solved numerically using a shooting method resulting in the well-known laminar boundary layer profile. Using the numerical solution, an expression for the skin friction coefficient along the flat plate can also be derived:
 
-![Blasius Cf](../../Laminar_Flat_Plate/images/blasius_cf.png)
+![Blasius Cf](../../tutorials_files/compressible_flow/Laminar_Flat_Plate/images/blasius_cf.png)
 
 where Re_x is the Reynolds number along the plate. In this tutorial, we will perform a solution of nearly incompressible (low Mach number) laminar flow over a flat plate and compare our results against the analytical Blasius solutions for the profile shape and skin friction coefficient along the plate. This problem has become a classic test case for viscous flow solvers. More detail on the Blasius solution and the similarity variables can be found in Chapter 18 of Fundamentals of Aerodynamics (Fourth Edition) by John D. Anderson, Jr. and most other texts on aerodynamics.
 
@@ -63,7 +63,7 @@ This problem will solve the for the flow over the flat plate with these conditio
 
 The computational mesh for the flat plate is composed of quadrilaterals with 65 nodes in both the x- and y-directions. The flat plate is along the lower boundary of the domain (y = 0) starting at x = 0 m and is of length 0.3048 m (1 ft). In the figure of the mesh, this corresponds to the Navier-Stokes (no-slip) boundary condition highlighted in green. The domain extends a distance upstream of the flat plate, and a symmetry boundary condition is used to simulate a free-stream approaching the plate in this region (highlighted in purple). Axial stretching of the mesh is used to aid in resolving the region near the start of the plate where the no-slip boundary condition begins at x = 0 m, as shown in Figure (1).
 
-![Lam Plate Mesh](../../Laminar_Flat_Plate/images/lam_plate_mesh_bcs.png)
+![Lam Plate Mesh](../../tutorials_files/compressible_flow/Laminar_Flat_Plate/images/lam_plate_mesh_bcs.png)
 Figure (1): Figure of the computational mesh with boundary conditions.
 
 Because the flow is subsonic and disturbances caused by the presence of the plate can propagate both upstream and downstream, characteristic-based, subsonic inlet and outlet boundary conditions are used for the flow entrance plane (red) and the outflow regions along the upper region of the domain and the exit plane at x = 0.3048 m (blue). 
@@ -131,11 +131,11 @@ The flat plate simulation for the 65x65 node mesh is small and will execute rela
 
 Results are given here for the SU2 solution of laminar flow over the flat plate. The results show excellent agreement with the closed-form Blasius solution.
 
-![Lam Plate Mach](../../Laminar_Flat_Plate/images/lam_plate_mach.png)
+![Lam Plate Mach](../../tutorials_files/compressible_flow/Laminar_Flat_Plate/images/lam_plate_mach.png)
 Figure (2): Mach contours for the laminar flat plate.
 
-![Lam Plate Profile](../../Laminar_Flat_Plate/images/lam_plate_velocity_profile.png)
+![Lam Plate Profile](../../tutorials_files/compressible_flow/Laminar_Flat_Plate/images/lam_plate_velocity_profile.png)
 Figure (3):  Velocity data was extracted from the exit plane of the mesh (x = 0.3048 m) near the wall, and the boundary layer velocity profile was plotted compared to and using the similarity variables from the Blasius solution.
 
-![Lam Plate Cf](../../Laminar_Flat_Plate/images/lam_plate_skin_friction.png)
+![Lam Plate Cf](../../tutorials_files/compressible_flow/Laminar_Flat_Plate/images/lam_plate_skin_friction.png)
 Figure (4): A plot of the skin friction coefficient along the plate created using the values written in the surface_flow.csv file and compared to Blasius.

_tutorials/compressible_flow/NICFD_nozzle/NICFD_nozzle.md

@@ -12,7 +12,7 @@ complexity: advanced
 follows: 
 ---
 
-![NICFD nozzle Mach](../../NICFD_nozzle/images/mach_isolines.png)
+![NICFD nozzle Mach](../../tutorials_files/compressible_flow/NICFD_nozzle/images/mach_isolines.png)
 
 ## Goals
 
@@ -60,7 +60,7 @@ In design conditions, the total to exhaust pressure ratio of the nozzle is 3.125
 
 The total length of the nozzle is 0.123 m, with an inlet height of 0.036 m and a throat height of 0.0084 m. The mesh is composed of quadrilateral elements, with 3,540 elements and 3,660 nodes. The figure shows the mesh topology and an indication of the boundary conditions. Characteristic-based Riemann boundary conditions are used on the INFLOW and OUTFLOW boundaries. The Navier-Stokes adiabatic wall condition is imposed on the WALL boundary. The symmetry boundary condition is used at the SYMMETRY boundary. The symmetry condition mirrors the flow about the x axis, thus allowing to reduce the size of the mesh and the computational cost.
 
-![NICFD nozzle mesh](../../NICFD_nozzle/images/mesh.png)
+![NICFD nozzle mesh](../../tutorials_files/compressible_flow/NICFD_nozzle/images/mesh.png)
 Figure (1): Computational mesh.
 
 ### Configuration File Options
@@ -217,8 +217,8 @@ The nozzle simulation is relatively small and will execute quickly on a single w
 Results are given here for the SU2 solution of supersonic non-ideal compressible flow in the converging-diverging nozzle. As part of this tutorial, a coarse mesh was provided, but for comparison, results obtained by using a refined mesh (80,223 elements and 80,840 points) as well as experimental results are shown.
 The figures below compare pressure and Mach number trends along the nozzle axis obtained from SU2 flow solutions and experimental data. Numerical results agree with the experimental data from the TROVA wind tunnel.
 
-![NICFD nozzle results A](../../NICFD_nozzle/images/nozzle_geometry_schlieren.png)
-![NICFD nozzle results B](../../NICFD_nozzle/images/Pressure_SU2_experiments.png)
-![NICFD nozzle results C](../../NICFD_nozzle/images/Mach_SU2_experiments.png)
+![NICFD nozzle results A](../../tutorials_files/compressible_flow/NICFD_nozzle/images/nozzle_geometry_schlieren.png)
+![NICFD nozzle results B](../../tutorials_files/compressible_flow/NICFD_nozzle/images/Pressure_SU2_experiments.png)
+![NICFD nozzle results C](../../tutorials_files/compressible_flow/NICFD_nozzle/images/Mach_SU2_experiments.png)
 
 Figure (2): (a) Geometry of the test section and schlieren image of the nozzle flow. (b, c) Comparison Pressure (b) and Mach number (c) profiles of the experimental results of Spinelli *et al* (black dots with error bars) against SU2 computational results for the test case mesh (red lines) and a reference fine mesh (blue lines).

_tutorials/compressible_flow/Transitional_Flat_Plate/Transitional_Flat_Plate.md

@@ -12,7 +12,7 @@ complexity: basic
 follows: 
 ---
 
-![lam_to_turb](../../Transitional_Flat_Plate/images/lam_to_turb.png)
+![lam_to_turb](../../tutorials_files/compressible_flow/Transitional_Flat_Plate/images/lam_to_turb.png)
 
 ## Goals
 
@@ -48,7 +48,7 @@ The length of the flat plate is 1.5 meters, and it is represented by an adiabati
 
 The mesh used for this tutorial, which consists of 41,412 quadrilaterals, is shown below.
 
-![Flat Plate](../../Transitional_Flat_Plate/images/FlatPMesh.png)
+![Flat Plate](../../tutorials_files/compressible_flow/Transitional_Flat_Plate/images/FlatPMesh.png)
 
 Figure (1): Mesh with boundary conditions (red: inlet, blue:outlet, orange:symmetry, green:wall)
 
@@ -114,15 +114,15 @@ To run this test case, follow these steps at a terminal command line:
 
 The figure below compares the skin friction results obtained by the B-C transition model to the experimental data. 
 
-![SK_Cf_Rex](../../Transitional_Flat_Plate/images/Cf_Rex_SK.png)
+![SK_Cf_Rex](../../tutorials_files/compressible_flow/Transitional_Flat_Plate/images/Cf_Rex_SK.png)
 
 Figure (2): Comparison of the skin friction coefficients for the Schubauer & Klebanoff case.
 
 ## Notes
 
 By changing the freestream velocity and turbulence intensity options in the config file with the values given in the table below, you may also simulate other very popular zero pressure gradient transitional flat plate test cases. You may use the same grid file for these test cases.
 
-![other_cases_table](../../Transitional_Flat_Plate/images/other_transition_cases.png)
+![other_cases_table](../../tutorials_files/compressible_flow/Transitional_Flat_Plate/images/other_transition_cases.png)
 
 ## References
 

_tutorials/compressible_flow/Turbulent_Flat_Plate/Turbulent_Flat_Plate.md

@@ -12,7 +12,7 @@ complexity: basic
 follows: 
 ---
 
-![Turb Plate Uplus v Yplus](../../Turbulent_Flat_Plate/images/turb_plate_uplus_vs_yplus.png)
+![Turb Plate Uplus v Yplus](../../tutorials_files/compressible_flow/Turbulent_Flat_Plate/images/turb_plate_uplus_vs_yplus.png)
 
 ## Goals
 
@@ -50,7 +50,7 @@ The length of the flat plate is 2 meters, and it is represented by an adiabatic
 
 The mesh used for this tutorial, which consists of 13,056 quadrilateral elements (the 137x97 grid), is shown below. A finer 545x385 grid file is also available. Additional grids for the flat plate in this same family can be obtained from the NASA Turbulence Modeling resource page.
 
-![Turb Plate Mesh](../../Turbulent_Flat_Plate/images/turb_plate_mesh_bcs.png)
+![Turb Plate Mesh](../../tutorials_files/compressible_flow/Turbulent_Flat_Plate/images/turb_plate_mesh_bcs.png)
 Figure (1): Mesh with boundary conditions: inlet (red), outlet (blue), symmetry (purple), wall (green).
 
 ### Configuration File Options
@@ -109,11 +109,11 @@ To run this test case, follow these steps at a terminal command line:
 
 The figures below show results obtained from SU2 and compared to several results from NASA codes. Note that the SU2 results for the skin friction correspond to the coarser mesh ([mesh_flatplate_turb_137x97.su2](https://github.com/su2code/Tutorials/tree/master/compressible_flow/Turbulent_Flat_Plate/mesh_flatplate_turb_137x97.su2)) while the NASA results are based on the finer mesh ([mesh_flatplate_turb_545x385.su2](https://github.com/su2code/Tutorials/tree/master/compressible_flow/Turbulent_Flat_Plate/mesh_flatplate_turb_545x385.su2)). SU2 still matches very closely.
 
-![Turb Plate Nu Tilde](../../Turbulent_Flat_Plate/images/turb_plate_nu_tilde.png)
+![Turb Plate Nu Tilde](../../tutorials_files/compressible_flow/Turbulent_Flat_Plate/images/turb_plate_nu_tilde.png)
 Figure (2): Contour of turbulence variable (nu-hat).
 
-![Turb Plate Cf](../../Turbulent_Flat_Plate/images/turb_plate_skin_friction.png)
+![Turb Plate Cf](../../tutorials_files/compressible_flow/Turbulent_Flat_Plate/images/turb_plate_skin_friction.png)
 Figure (3): Profile for the skin friction coefficient. 
 
-![Turb Plate Uplus v Yplus](../../Turbulent_Flat_Plate/images/turb_plate_uplus_vs_yplus.png)
+![Turb Plate Uplus v Yplus](../../tutorials_files/compressible_flow/Turbulent_Flat_Plate/images/turb_plate_uplus_vs_yplus.png)
 Figure (4): Velocity profile comparison against law of the wall.