More on Aerofoils Flashcards
Cl vs α graph for flapped aerofoils
Same shape as a standard Cl vs α graph however with increasing δ (flap deflection angle) the graph moves vertically upwards
Eq for Cl for a flapped aerofoil
Cl = Clα (α - αL=0 + (τ * δf))
τ = flap effectiveness
δf = flap deflection angle
Order the 8 possibilities of aerofoil configuration in terms of max Cl
8 Possibilities:
Aerofoil only, Plain flap, Split flap, Leading edge slat, Single-slotted flap, Double-slotted flap, Double-slotted flap w/ leading edge slat, Double-slotted flap w/ leading edge slat and boundary layer suction
Ranked in order of highest to lowest Cl max
1 - Double-slotted flap w/ leading edge slat and boundary layer suction
2 - Double-slotted flap w/ leading edge slat
3 - Double-slotted flap
4 - Single-slotted flap
5 - Split flap
6 - Plain flap
7 - Leading edge slat
8 - Aerofoil only
Local Mach No. Vs Global Mach No. w/ Eqs
Local Mach No. is based on the local velocity defined in every point of the space surrounding the aeroplane
Ma = v/a
Global Mach No. is based on the free stream velocity and characterises the flight regime
Ma∞ = v∞/a
The three Mach numbers
Macr = (critical Mach No.) The smallest value of Ma∞ for which there is at least one point of the flow field where the local Mach No. = 1
Ma∞ = Global Mach Number
Ma-dragdivergence = The smallest value of Ma∞ for which a shockwave is generated in the flow field
4 Possible flight states subsonic → supersonic
State 1 - Subsonic Flow Field
Ma < 1
Ma∞ < Macr
State 2- Supersonic region on upper surface, subsonic everywhere else, no shockwave
Macr < Ma∞ < Ma-dragdivergence
Ma < 1 except for a small region on the upper surface where Ma > 1
State 3 - Supersonic region on upper surface, subsonic elsewhere, one shockwave
Ma∞ > Ma- dragdivergence
Ma > 1 on upper surface before shockwave
State 4 - Subsonic region at leading edge, supersonic elsewhere, several shockwaves
Ma∞ > 1
Ma < 1 at leading edge
Ma > 1 everywhere else
