Are these pressure and velocity distributions calculated by CFD simulation correct?
Qualitatively they seem to be correct:


Highest pressure at stagnation point and maximal pressure in diagram close to total pressure:
$$101843.84\ Pa \approx P_{tot}= P_{\infty}+\frac{\rho}{2}\cdot U_\infty^2 =101853\ Pa$$
Speed drop at stagnation point.
The slight flow deflection downwards behind the airfoil shows that your airfoil is generating lift (I am assuming your free stream flow is in x-direction).
Low pressure (dark blue) region above airfoil is bigger than below.


As @ROIMaison wrote the correct way to validate the results is looking at the aerodynamic coefficients and pressure ditribution $c_p$. In a quick search I found some polars calculated with xfoil you should be able to replicate. With xfoil you can calculate the pressure distribution too.

If you are only interested in lift at small angles of attack laminar flow simulation should be enough. If you also want to calculate drag I would set turbulent air. To validate your setup (model, boundary conditions,...) for turbulent flow there is a great NASA webpage.