Aero Chapter 02, High-Speed Flight

T-45 Aerodynamics Student Workbook

standard day temperature of -48F (-44.4C). Compressibility effects will be experienced at lower

airspeeds as altitude is increased. Mach Number is used to describe the relationship of airspeed to the

speed of sound…M = TAS /SOS where:

M = Mach Number

TAS = True Airspeed

SOS = Speed of Sound

In our discussion we will consider the effect of airflow only on an airfoil section. Keep in mind that airflow

about an aircraft is considerably more complicated.

Initially all airflow about the airfoil is less than

the speed of sound (Figure 17). The airfoil is

in the subsonic flight regime. However,

velocity of the airflow about the airfoil is

greater than the velocity of the free airstream.

At some flight velocity, less than Mach 1,

there will be an area of local sonic flow on the

airfoil (Figure 18). This is the “Critical Mach

Number” (Mcrit) of the airfoil. Critical Mach

Number is defined as the flight Mach number

Sonic Flow (m = 1)

where there is first evidence of sonic flow on

Flow Accelerates to above Flight Mach

the aircraft. It is an important reference point

because it is the beginning of the transonic

flight regime. All compression waves and

phenomenon occur at a Mach number

greater than Mcrit. As the airflow accelerates,

MCRIT = .80 Mach for T-45

the area of supersonic airflow increases.

Pressure waves created in the area of

supersonic airflow begin to pile up against the

aft moving airflow. This “piling up” forms a

weak normal compression wave. “NORMAL”

Normal Shock Wave

is a mathematical term meaning

Supersonic

Flow

perpendicular. Thus, a normal compression

wave is perpendicular to the airflow. A

Possible Separation

Subsonic

supersonic airstream traversing a normal

shock wave experiences a rapid reduction in

velocity to subsonic speed. If the Mach

number ahead of the compression wave is

Force Divergent Mach Number = 0.85 for T-45

Mach 1.2 or less, the velocity of the airflow

after the normal wave is approximately the

reciprocal of Mach 0.85 (Figure 19). With the

great reduction in velocity comes an increase in static pressure, density, and temperature. A great deal of

kinetic energy is converted into unusable heat.

The possibility of boundary layer separation from the trailing edge of the airfoil and the degree of that

separation relates directly to the aerodynamic design of the airfoil and the existent angle of attack.

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