Aerodynamics of Flight

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