Transition region  Laminar sublayer  Laminar boundary layer  Turbulent boundary layer  Figure 5-66.  Boundary layer.  change can no longer warn the air ahead because the airplane  is keeping up with its own pressure waves. Rather, the air  particles pile up in front of the airplane causing a sharp  decrease in the flow velocity directly in front of the airplane  with a corresponding increase in air pressure and density.  part of the velocity energy of the airstream is converted to  heat as it flows through the wave, is a contributing factor  in the drag increase, but the drag resulting from airflow  separation is much greater. If the shock wave is strong,  the boundary layer may not have sufficient kinetic energy  to withstand airflow separation. The drag incurred in the  transonic region due to shock wave formation and airflow  separation is known as “wave drag.” When speed exceeds  the critical Mach number by about 10 percent, wave drag  increases sharply. A considerable increase in thrust (power)  is required to increase flight speed beyond this point into the  supersonic range where, depending on the airfoil shape and  the AOA, the boundary layer may reattach.  As the airplane’s speed increases beyond the speed of sound,  the pressure and density of the compressed air ahead of it  increase, the area of compression extending some distance  ahead of the airplane. At some point in the airstream, the air  particles are completely undisturbed, having had no advanced  warning of the airplane’s approach, and in the next instant the  same air particles are forced to undergo sudden and drastic  changes in temperature, pressure, density, and velocity.  The boundary between the undisturbed air and the region  of compressed air is called a shock or “compression” wave.  This same type of wave is formed whenever a supersonic  airstream is slowed to subsonic without a change in direction,  such as when the airstream is accelerated to sonic speed  over the cambered portion of a wing, and then decelerated  to subsonic speed as the area of maximum camber is passed.  A shock wave forms as a boundary between the supersonic  and subsonic ranges.  Normal shock waves form on the wing’s upper surface and  form an additional area of supersonic flow and a normal shock  wave on the lower surface. As flight speed approaches the  speed of sound, the areas of supersonic flow enlarge and the  shock waves move nearer the trailing edge.  [Figure 5-67]  Whenever a shock wave forms perpendicular to the airflow, it  is termed a “normal” shock wave, and the flow immediately  behind the wave is subsonic. A supersonic airstream passing  through a normal shock wave experiences these changes:  M = 0.82  M = 0.95  M = 1.05  Normal shock wave  S u p e r s o n i c f l o w  Normal shock wave  S u p e r s o n i c f l o w  Subsonic airflow  Bow wave  Figure 5-67.  Shock waves.  Shock wave formation causes an increase in drag. One of  the principal effects of a shock wave is the formation of a  dense high pressure region immediately behind the wave.  The instability of the high pressure region, and the fact that  •  The airstream is slowed to subsonic.  •  The airflow immediately behind the shock wave does  not change direction.  • The static pressure and density of the airstream behind the wave is greatly increased.   •  The energy of the airstream (indicated by total pressure—dynamic plus static) is greatly reduced.  5-47 