What is the pressure distribution around a supersonic wing?
Normally, supersonic fighter wings for which airfoil information is published use a very thin NACA 6-digit section with very little camber, such as
Aircraft                root airfoil      tip airfoil
McDonnell Douglas F-15  NACA 64A006.6     NACA 64A203
General Dynamics F-16   NACA 64A204       NACA 64A204
Lockheed-Martin F-22    NACA 64A?05.92    NACA 64A?04.29

This information is from The Incomplete Guide to Airfoil Usage by Dave Lednicer. As you can see, the F-16 uses camber throughout, indicated by the 0.2 design lift coefficient of the airfoil, while the F-15 uses an uncambered root airfoil.
Those airfoils are chosen because flow around them at supersonic flight speed is still heavily influenced by subsonic flow characteristics as long as wing sweep allows for a subsonic leading edge. This produces nose suction, and flow around the wingtips disturbes the pure supersonic flow characteristics even when the leading edge is supersonic.
How the idealised flow around a supersonic airfoil looks is explained in this answer.
The angle of attack in supersonic flight depends on air density and load factor, but generally is only a few degrees because drag increases with the square of angle of attack. In other words, if the plane flies at high angle of attack, it decelerates to subsonic speed quickly.
The angle of attack in straight flight depends on altitude, wing loading and aspect ratio and grows with decreasing air density. Normally, at low level the maximum dynamic pressure limit and engine thrust will only allow very low supersonic speeds around Mach 1.1 to 1.3. Only above 20.000 ft will the envelope expend enough to make flight at Mach 1.8 possible.
For a typical wing loading of $\frac{m}{S}$ = 300 kg/m² and Mach 1.3 at 10.000 ft (air density $\rho$ = 0.9 kg/m³, speed of sound $a$ = 330 m/s) the angle of attack is $$\alpha = \frac{2\cdot m\cdot g}{\rho\cdot (Mach\cdot a)^2\cdot S\cdot c_{L\alpha}} \approx 0.5°$$ with a typical lift curve slope of a supersonic wing of $c_{L\alpha}$ = 4.