finals2 Flashcards
(30 cards)
PRANDTL- GLAUERT THEORY
A theoretical approach based on a number of simplyfing assumptions theory. It gives useful
results provided that the wing is not too thick, and some of these results are
quoted below.
for low mach numbers the glauert factor is
close to 1
why there is adjustment in lift and in drag in Prandtl- Glauert theory
for lift coefficient: ensures the analysis in the wings in the compressible flows
For drag: the form drag changes proportional to the CP. but at low angles of attack the profile drag remains unaffected due to the domination of the skin friction
limitations of adjustment’s formula
for cpc: infinite as Mach goes to one which is unrealistic
does not account boundary layer separation
only applies for 2d airfoils
High subsonic airfoils tend to have
Lower thickness- chord ratio
smaller leading edge radii of curvature
maximum thickness points after aft than in the case of conventional low- speed sections
pressure distribution on such a wing at low incidence tends to be as
at low aoa pressure dist temnds to be smoother and controled
but at a0a increases, the suction peak, region where lowest pressure occur will be sharper and move forward, forming leading edge separation bubble
which is more difficult
transonic flow regimes constitutes a more difficult problem than either subsonic or supersonic flow. because the flow is mixed it s much less amenable
the lowest mach number at which the airflow over any part of the aircraft reaches the speed of sound
critical mach number
aoa roles in critical mach number
at high incidence: acceleration on the upper surface will be greater than lower incidence in the same velocity
increasing aoa could reduce the critical mach number
as mach increases
sonic line moves forward and the shockwaves moves backwards
both shockwaves moves forward but more rapid on the lower surface
lambda pattern causes the schock wave to become oblique as it approaches to the trailing edge
The upper surface shock eventually also reaches the trailing edge, forming the so-
called
bifurcated trailing edge shock pattern.
normal shock wave is also called___ happens when a supersonic stream approaches a rounded-nosed airfoil
bow shock
at supersonic speed, rounded airfoil nose causes
detached bow shock
even if the leading edge is sharp, @ low supersonic speed; a detached shock may occur if____
flow deviation required is greater than the maximum flow deviation for a given mach number
it defines
the upper limit of the transonic flow regime. Beyond this, fully developed
supersonic flow occurs,
shock attachment Mach number,
as free stream mach numbers increases above M crit, the upper surface shock waves moves backwards and strengthens. soon it will become strong enough to cause the flow to separate
shock stall
pressure distribution:
sharp drop in the negative pressure when the schock is terminated (happpens when above crit mach)
before the ctrit mach shock wave is apparent
pressure distribution when shock pass through the trailing edge
less evere on the effect of pressre distribution
the average location of all of the pressure acting upon a body moving through a fluid
center of pressure position
in transonic regime , aircraft experience a signifiicant increase in drag, this phenomenon
transonic drag rise
wing section features to increase M crit and reduce transonic drag:
Low thickness to chord ratio
max thick point well aft at about half chord
small leading edge radius of curvature
symmetrical section
small wig loading
2 wing planfrom features to which have a significant effect on M crt
swwp back
aspect ratio
