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Pilot's Handbook of Aeronautical Knowledge
Aerodynamics of Flight
High Speed Flight

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Pilot's Handbook of Aeronautical Knowledge



Table of Contents

Chapter 1, Introduction To Flying
Chapter 2, Aircraft Structure
Chapter 3, Principles of Flight
Chapter 4, Aerodynamics of Flight
Chapter 5, Flight Controls
Chapter 6, Aircraft Systems
Chapter 7, Flight Instruments
Chapter 8, Flight Manuals and Other Documents
Chapter 9, Weight and Balance
Chapter 10, Aircraft Performance
Chapter 11, Weather Theory
Chapter 12, Aviation Weather Services
Chapter 13, Airport Operation
Chapter 14, Airspace
Chapter 15, Navigation
Chapter 16, Aeromedical Factors
Chapter 17, Aeronautical Decision Making




Boundary layer.
Figure 4-58. Boundary layer.

Vortex generators are used to delay or prevent shock wave
induced boundary layer separation encountered in transonic
flight. They are small low aspect ratio airfoils placed at a 12°
to 15° AOA to the airstream. Usually spaced a few inches
apart along the wing ahead of the ailerons or other control
surfaces, vortex generators create a vortex which mixes the
boundary airflow with the high energy airflow just above the
surface. This produces higher surface velocities and increases
the energy of the boundary layer. Thus, a stronger shock wave
is necessary to produce airflow separation.

Shock Waves
When an airplane flies at subsonic speeds, the air ahead is
"warned" of the airplane's coming by a pressure change
transmitted ahead of the airplane at the speed of sound.
Because of this warning, the air begins to move aside before
the airplane arrives and is prepared to let it pass easily. When
the airplane's speed reaches the speed of sound, the pressure
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.

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.

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:

• 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.

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
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 angle of attack, the boundary layer may reattach.

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 4-59]

Shock waves.
Figure 4-59. Shock waves.

Associated with "drag rise" are buffet (known as Mach
buffet), trim and stability changes, and a decrease in control
force effectiveness. The loss of lift due to airflow separation
results in a loss of downwash, and a change in the position of
the center pressure on the wing. Airflow separation produces
a turbulent wake behind the wing, which causes the tail
surfaces to buffet (vibrate). The nose-up and nose-down pitch
control provided by the horizontal tail is dependent on the
downwash behind the wing. Thus, an increase in downwash
decreases the horizontal tail's pitch control effectiveness
since it effectively increases the angle of attack that the tail
surface is seeing. Movement of the wing CP affects the wing
pitching moment. If the CP moves aft, a diving moment
referred to as "Mach tuck" or "tuck under" is produced,
and if it moves forward, a nose-up moment is produced.
This is the primary reason for the development of the T-tail
configuration on many turbine-powered aircraft, which
places the horizontal stabilizer as far as practical from the
turbulence of the wings.