Longitudinal Static Stability Flashcards
Define trim
control surface deflection required to get aircraft into the steady state
Define stability
if aircraft will settle into steady state after a disturbance
Define dynamic response
motion caused by atmospheric disturbances or changes to the control inputs
Describe body fixed reference frame
origin centre of gravity
x axis - along body out of nose
y axis - along starboard side
x axis - down
Why use CG as origin, what are the difficulties with it
why - makes equations of motion simpler
difficulties - CG moves with mass change
Define quasi-static
unsteady aerodynamic and inertial effects are ignored
Define statically stable
after a disturbance restoring moments generated to tilt aircraft back into equilibrium
Define statically unstable
after a disturbance moments generated to tilt aircraft away from equilibrium
Define neutrally stable
no restoring moments generated
Define aerodynamic centre
where changes in angle of attack do not change the pitching moment
not typically 0, just a constant
Why does stability increase as CG moves forward
- moment arm of horizontal stabiliser increases
- contribution of wings lift to pitching moment is also stabilising
Why is it desirable to have trim alpha as positive
the wing will produce more lift
List all contributions for stability of an aircraft
- wing
- tail
- fuselage
- propulsion system
List method to derive expression for the contribution of the wing
- resolve lift and drag into normal and chord wise forces
- sum moments around CG
- substitute Lw and D
- divide by 0.5pV^2Sc to get moments
- assume alpha is small (cosx = 1, sinx = x)
- assume Clw»_space; Cdw and Zgcw = 0
- apply conditions for statically stable
Describe wing contribution
- indicates ac must lie aft of cg
- requires negative-cambered aerofoil section
- none of these fulfilled due to other considerations (Aerodynamics)
- wing contribution is always destabilising
List method to derive expression for the contribution of the tail
- total lift L = Lw + Lt
- use lift equation
- i.e. 0.5pV^2SCl = 0.5pVw^2SClw + 0.5pVt^2SClt
- divide by QS, assume V = Vw
- substitute n = tail efficiency = 0.5pVt^2/0.5pV^2
- sum pitching moments around CG
- assume small angles and Lt»_space; Dt
- nondimensionalise (divide by QSc)
- Substitute Vh
- substitute Clt = Cla x at
- substitute at = aw - iw - e - it
- split into Cmot + Cmat x a
List advantages and disadvantages of canard
A:
- free from wing/propulsion system flow interference
- more attractive from trimming nose-down moment of high-lift devices
D
- destabilises contribution to static stability
Define stick fixed neutral point
- the rear most position to which the CG can be moved aft before it becomes unstable. At neutral point, the aircraft is neutrally stable
- stick-fixed related to the fact the elevator angle is controlled by the pilot through the control circuit (stick deflection constant at this point hence ‘stick-fixed’
define static margin
the distance through which CG can move rearwards before reaching the neutral point. It is a measure of longitudinal stability.
How to find stick fixed neutral point
- substitute Xnp = Xcg
- make Cma = 0
- solve for Xnp
List primary aerodynamic controls
roll - differential deflection of ailerons
Pitch - elevators
Yaw - rudder
high lift devices - flaps, slats and spoilers
Factors effecting control surface design
- control effectiveness (flap size/tail volume)
- hinge moments (force required to overcome aerodynamic moments to rotate control surfaces)
- aerodynamic and mass balancing (ensure stick force within acceptable range)
