BARO POWER N AV XPDR  LCL 18:19:09 1 1200 ALT RAT 10°C  ISA -0°C NORTH UP 2 NM T OFF SC ALE GPS ENR 360 ° HDG 360° CRS 360° 10 10 5 5 5 5 4 2 2 4 000 30800 30700 30900 31100 31200 31300 20 80 31 000 _ _ _ _ 180 190 200 220 230 240 9 210 1 M .579 TRAFFIC NAV1 108.00 117.95 NAV2 108.00 136.975 118.000 COM1 136.975  COM2 DIS  BRG ___° __._NM ADF/DME /TFC SENSOR TRFC/MAP MSG 180 190 200 220 230 240 9 210 1 .579 Mach Air Speed V mo  and M mo $_{ }$Barber Pole Airspeed Indicator Figure 15-8.  Jet airspeed indicator. High-speed airplanes designed for subsonic flight are limited  to some Mach number below the speed of sound to avoid the  formation of shock waves that begin to develop as the airplane  nears Mach 1.0. These shock waves (and the adverse effects  associated with them) can occur when the airplane speed is  substantially below Mach 1.0. The Mach speed at which some  portion of the airflow over the wing first equals Mach 1.0  is termed the critical Mach number (Mcr). This is also the  speed at which a shock wave first appears on the airplane. As the graph in  Figure 15-10  illustrates, initially  as speed is  increased up to Mach .72, the wing develops an increasing  amount of lift requiring a nose-down force or trim to maintain  level flight. Wit increased speed and the aft movement of  the shock wave, the wing’s center of pressure also moves  aft causing the start of a nose-down tendency or “tuck.” By  Mach .9, the nose-down forces are well developed  to a point  where a total of 70 pounds of back pressure are required to  There is no particular problem associated with the acceleration  of the airflow up to Mach Crit, the point where Mach 1.0 is  encountered; however, a shock wave is formed at the point  where the airflow suddenly returns to subsonic flow. This  shock wave becomes more severe and moves aft on the wing  as speed of the wing is increased and eventually flow separation  occurs behind the well-developed shock wave.  [Figure 15-9]  S e p ar ation P o s i b l e s e p a r a tion M = 0.72 (Critical mach number) M = 0.77 M = 0.82 Maximum local velocity is less than sonic Normal shock  Normal shock  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 15-8 Figure 15-9.  Transonic flow patterns. If allowed to progress well beyond the M$_{MO}$ for the airplane,  this separation of air behind the shock wave can result in  severe buffeting and possible loss of control or “upset.”  Because of the changing center of lift of the wing resulting  from the movement of the shock wave, the pilot experiences  pitch change tendencies as the airplane moves through the  transonic speeds up to and exceeding M$_{MO}$.  [Figure 15-10]