Changed 2/1/04 by kchen
span
root chord
tip chord
mean chord
root thickness for structures
tip thickness
average thickness
wetted area
Reynolds Number
Length of tail
Horizontal Tail Twist
assumed dynamic pressure at the tail / dynamic pressure, qh/q
incidence angle of the horizontal tail
Elevator, fraction of chord,
Elevator, fraction of span,
span
root chord
tip chord
mean chord
root thickness, for structures
tip thickness
average thickness
wetted area
volume ratio
Reynolds number
tail vol coeff
This value (C.v) is the y input on Fig. 9-24 in Torenbeek.
0.038 or greater for a single engine propeller A/C with little normal
force on the vertical tail.
Look at the "Checking the Vertical Tail Size" section.
the analysis shows that a C.vt of 0.04 is required.
assumed dynamic pressure at the tail / dynamic pressure, qh/q
Rudder, fraction of chord,
Rudder, fraction of span,
Note that Ltail is not entirely correct as the term should really
be L.v, the length to the quarter chord of the vertical tail.
span
root chord
tip chord
mean chord
root thickness, for structures
tip thickness
average thickness
wetted area
rudder chord
Reynold's number
Tau is the flap effectiveness parameter and is
found in Fig. 2.21 on pg. 64 in Flight Stability and Control.modeled here
Flap Fraction defined here
Needs to be in Layout
rudder effectiveness
elevator effectiveness
Change in Lift of Vertical Tail Due to Rudder deflection
Needs to Change use 2D CLa
Change in Lift of Horizontal Tail Due to
Diameter of the fuselage
Length of Fuselage
Vertical Tail Volume Coefficient
Vertical Distance from Wing cc to cg
NOT used in calculations, used to look up k.n factor (Nelson, 74-75)
Fuselage height at 1/4 length
Fuselage height at 3/4 length
Fuselage side angle
k.n factor (Nelson 74)
Reynolds number of the fuselage
k.rl factor, correction for Re.fuse (Nelson 74)
Coeficient of yaw moment from fuselage
PER DEGREE
Coeficient of yaw from vertical tail, per radian
This value should be between .05 and .1
From the wing lift vs. alpha plot
Need to check what input to C.M.w
Moment of wing at 0 angle of attack
Also assumed
Should be 5 - 10%
Scaled from a general aviation plane
Units on this should be s^-2 if correct.
This is in Table 3.6 of Flight Stability and Control.
This parameter is a stability check.
The side force corresponding to a change in yawing moment created by rudder deflection is calculated as follows:
And the corresponding yawing moment:
The force exerted on the ailerons is due to the rolling rate, p,
and the change in moment that roll creates.
The control power of the ailerons can be calculated as follows:
Eqn. 2.96 Flight Stability and Control
X.a is equal to the distance from the root chord to the closest edge of the aileron in the spanwise direction this value will come from layout
y.2 is equal to the half span of the wing minus 10%
y.1 is equal to the distance from the root chord to near edge of the aileron in the spanwise direction
CL.
Cn.
The servos must be powerful enough to actuate the elevator during all loading conditions.
The most critical flight segments are take-off and climb-out.
as all others) the force the servos exert is equal to the force felt by the elevator.
quantified by multiplying the coefficient of the normal force on the elevator by the weight of
the aircraft.
The elevator changes the longitudinal orientation of the airplane, i.e. produces a change
in lift as a function of the pitching velocity, q.
in lift, neglecting the contribution of the wing because it's quite small in comparison to that
of the horizontal tail.
The next equation calculates the corresponding change in pitching moment.
(Both of these equations can be found on pgs. 112-113 of "Flight Stability and Automatic
Control", 2nd ed., Robert C. Nelson.)