bug

Author: juancamarotti

potential_flow/validation/naca0012_potential_incompressible_flow/README.md

@@ -36,21 +36,21 @@ Figure 2 presents the pressure coefficient ($C_p$) distribution along the chord
 
 <p align="center">
   <img src="data/Cp_distribution_x.png" alt="Pressure coefficient distribution along the airfoil chord." style="width: 600px;"/>
-  <figcaption align="center"> Figure 2: Pressure coefficient ($C_p$) distribution along the chord of the NACA 0012 airfoil. </figcaption>
+  <figcaption align="center"> Figure 2: Pressure coefficient (Cp) distribution along the chord of the NACA 0012 airfoil. </figcaption>
 </p>
 
 Figure 3 shows the spatial distribution of the pressure coefficient over the entire computational domain.
 
 <p align="center">
   <img src="data/Cp_distribution_airfoil.png" alt="Pressure coefficient distribution around the NACA 0012 airfoil." style="width: 600px;"/>
-  <figcaption align="center"> Figure 3: Pressure coefficient ($C_p$) distribution in the computational domain around the NACA 0012 airfoil. </figcaption>
+  <figcaption align="center"> Figure 3: Pressure coefficient (Cp) distribution in the computational domain around the NACA 0012 airfoil. </figcaption>
 </p>
 
 The aerodynamic coefficients obtained from the simulation can be compared with reference data using the lift curve shown in Figure 4.
 
 <p align="center">
   <img src="data/naca0012_lift_curve.png" alt="NACA 0012 Lift vs AOA Curve" style="width: 600px;"/>
-  <figcaption align="center"> Figure 4: Reference lift coefficient ($C_l$) as a function of the angle of attack for the NACA 0012 airfoil. </figcaption>
+  <figcaption align="center"> Figure 4: Reference lift coefficient (Cl) as a function of the angle of attack for the NACA 0012 airfoil. </figcaption>
 </p>
 
 For an angle of attack of $\alpha = 5^\circ$, the computed lift coefficient is $C_l = 0.569$. This result shows very good agreement with the reference data presented in Figure 4, which is used here for validation purposes.