Deriving the longitudinal equations of motion for an aircraft

Question

I'm trying to create a state-space model for the pitch control of an aircraft as described here: https://ctms.engin.umich.edu/CTMS/index.php?example=AircraftPitch&section=SystemModeling

The following equations are given for change in angle of attack and pitch rate:

Here's the translation of what any of those symbols mean:

$q$ = Pitch rate.

$\theta$ = Pitch angle.

$\delta$ = Elevator deflection angle.

$\mu = \frac{\rho S \bar{c}}{4 m}$.

$\rho$ = Density of air.

$S$ = Platform area of the wing.

$\bar{c}$ = Average chord length.

$m$ = Mass of the aircraft.

$\Omega = \frac{2 U}{\bar{c}}$.

$U$ = Equilibrium flight speed.

$C_T$ = Coefficient of thrust.

$C_D$ = Coefficient of drag.

$C_L$ = Coefficient of lift.

$C_W$ = Coefficient of weight.

$C_M$ = Coefficient of pitch moment.

$\gamma$ = Flight path angle.

$\sigma=\frac{1}{1+\mu C_L}$ = Constant.

$i_{yy}$ = Normalized moment of inertia.

$\eta=\mu \sigma C_M$ = Constant

But how do I get to these equations?

"Please refer to any aircraft-related textbooks for the explanation of how to derive these equations."

Not very helpful.