High speed flows Flashcards
Compressibility effects
As an aircraft approaches the speed of sound, significant density changes take place in the air flowing past the airframe
Mach number
Variation of lift and drag with mach number
Pressure waves when M=0
pressure waves radiate from the body in concentric circles when the body moves, the pressure waves infront of the body are closer together and the waves velocity relative to the body of a-V pressure waves behind the body are further apart and the waves velocity relative to the body of a+V
Pressure waves when M = 1
the body is moving at the same speed as the pressure waves the waves collect into a single pressure wave known as a Mach wave when M > 1, the mach wave trails back from the body
Shock wave
a wave formed due to a sudden pressure change between the ambient air and the aircraft’s pressure field
air flowing through a shockwave undergoes rapid changes in density, pressure, temperature and velocity
Critical mach number
the aircraft mach number at which local velocities first reach M=1
What happens when the freestream mach number = critical mach number?
no shock wave formed the local mach number only equals 1 at one point
What happens when the critical mach number < freestream mach number < 1?
flow accelerates to supersonic conditions
supersonic flow starts to decelerate over the rear of the aerofoil
pressure waves from flow downstream of the supersonic region, cannot move forward, therefore, a terminating shock is produced
Define terminating shock
the shock wave at the downstream end of the supersonic region
it terminates the supersonic region and slows the flow abruptly to below the speed of sound
Shock induced separation
separation of the boundary layer from the aerofoil due to the slowing of the supersonic flow as a result of the terminating shock
This causes a significant increase in drag and decrease in lift
Drag divergence Mach number
the mach number at which there is a rapid increase in drag
What happens when the freestream mach number > 1 (Blunt nose)?
bow shockwave lies forward of the leading edge
flow decelerates across the shock and compresses
flow downstream of the shock will accelerate and possibly become supersonic
loss of pressure in flow over the rear of the aerofoil results in pressure or wave drag
terminating shock is at the trailing edge no longer slowing the flow to subsonic
the bow wave is perpendicular to the flow directly ahead of the body but its angle to the flow is the same as the Mach angle off to the sides of the body’s path
What happens when the freestream mach number > 1 (Sharp nose)?
shockwave attached at the leading edge
M>1 across shockwave
shockwaves are oblique, not perpendicular to the flow
shock angle same as the mach angle further from the body
in oblique shockwaves, loss of velocity and total pressure is reduced
Wing sweep
used to delay the onset of critical flow
