Functions
def	SUAVE.Methods.Flight_Dynamics.Dynamic_Stability.Approximations.dutch_roll.dutch_roll (velocity, Cn_Beta, S_gross_w, density, span, I_z, Cn_r)

def	SUAVE.Methods.Flight_Dynamics.Dynamic_Stability.Approximations.phugoid.phugoid (g, velocity, CD, CL)

def	SUAVE.Methods.Flight_Dynamics.Dynamic_Stability.Approximations.roll.roll (I_x, S_gross_w, density, velocity, span, Cl_p)

def	SUAVE.Methods.Flight_Dynamics.Dynamic_Stability.Approximations.short_period.short_period (velocity, density, S_gross_w, mac, Cm_q, Cz_alpha, mass, Cm_alpha, Iy, Cm_alpha_dot)

def	SUAVE.Methods.Flight_Dynamics.Dynamic_Stability.Approximations.spiral.spiral (mass, velocity, density, S_gross_w, Cl_p, Cn_Beta, Cy_phi, Cl_Beta, Cn_r, Cl_r)
	spiral.py More...

Detailed Description

Function Documentation

◆ dutch_roll()

def SUAVE.Methods.Flight_Dynamics.Dynamic_Stability.Approximations.dutch_roll.dutch_roll	(	velocity,
		Cn_Beta,
		S_gross_w,
		density,
		span,
		I_z,
		Cn_r
	)

This calculates the natural frequency and damping ratio for the 
approximate dutch roll characteristics       

Assumptions:
    Major effect of rudder deflection is the generation of the Dutch roll mode.
    Dutch roll mode only consists of sideslip and yaw
    Beta = -Psi
    Phi and its derivatives are zero
    consider only delta_r input and Theta = 0
    Neglect Cy_r
    X-Z axis is plane of symmetry
    Constant mass of aircraft
    Origin of axis system at c.g. of aircraft
    Aircraft is a rigid body
    Earth is inertial reference frame
    Perturbations from equilibrium are small
    Flow is Quasisteady
  
Source:
  J.H. Blakelock, "Automatic Control of Aircraft and Missiles" Wiley & Sons, Inc. New York, 1991, p 132-134.      
  
Inputs:
    velocity - flight velocity at the condition being considered          [meters/seconds]
    Cn_Beta - coefficient for change in yawing moment due to sideslip     [dimensionless]
    S_gross_w - area of the wing                                          [meters**2]
    density - flight density at condition being considered                [kg/meters**3]
    span - wing span of the aircraft                                      [meters]
    I_z - moment of interia about the body z axis                         [kg * meters**2]
    Cn_r - coefficient for change in yawing moment due to yawing velocity [dimensionless]

Outputs:
    output - a data dictionary with fields:
    dutch_w_n - natural frequency of the dutch roll mode                  [radian/second]
    dutch_zeta - damping ratio of the dutch roll mode                     [dimensionless]
 
Properties Used:
    N/A

◆ phugoid()

def SUAVE.Methods.Flight_Dynamics.Dynamic_Stability.Approximations.phugoid.phugoid	(	g,
		velocity,
		CD,
		CL
	)

This calculates the natural frequency and damping ratio for the approximate 
phugoid characteristics       

Assumptions:
    constant angle of attack
    theta changes very slowly
    Inertial forces are neglected
    Neglect Cz_q
    Theta = 0
    X-Z axis is plane of symmetry
    Constant mass of aircraft
    Origin of axis system at c.g. of aircraft
    Aircraft is a rigid body
    Earth is inertial reference frame
    Perturbations from equilibrium are small
    Flow is Quasisteady 
    
Source:
    J.H. Blakelock, "Automatic Control of Aircraft and Missiles" Wiley & Sons, Inc. New York, 1991, p 50-53.
    
Inputs:
    g - gravitational constant                                   [meters/second**2]
    velocity - flight velocity at the condition being considered [meters/seconds]
    CD - coefficient of drag                                     [dimensionless]
    CL - coefficient of lift                                     [dimensionless]

Outputs:
    output - a data dictionary with fields:
        phugoid_w_n - natural frequency of the phugoid mode      [radian/second]
        phugoid_zeta - damping ratio of the phugoid mode         [dimensionless]
               
Properties Used:
    N/A

◆ roll()

def SUAVE.Methods.Flight_Dynamics.Dynamic_Stability.Approximations.roll.roll	(	I_x,
		S_gross_w,
		density,
		velocity,
		span,
		Cl_p
	)

This calculates the approximate time constant for the roll mode       

Assumptions:
   Only the rolling moment equation is needed from the Lateral-Directional equations
   Sideslip and yaw angle are being neglected and thus set to be zero.
   delta_r = 0
   X-Z axis is plane of symmetry
   Constant mass of aircraft
   Origin of axis system at c.g. of aircraft
   Aircraft is a rigid body
   Earth is inertial reference frame
   Perturbations from equilibrium are small
   Flow is Quasisteady
   
Source:
   J.H. Blakelock, "Automatic Control of Aircraft and Missiles" Wiley & Sons, Inc. New York, 1991, p 134-135.
   
Inputs:
   I_x -  moment of interia about the body x axis [kg * meters**2]
   S_gross_w - area of the wing [meters**2]
   density - flight density at condition being considered [kg/meters**3]
   span - wing span of the aircraft [meters]
   velocity - flight velocity at the condition being considered [meters/seconds]
   Cl_p - change in rolling moment due to the rolling velocity [dimensionless]
   
Outputs:
   roll_tau - approximation of the time constant of the roll mode of an aircraft [seconds] (positive values are bad)

Properties Used:
   N/A

◆ short_period()

def SUAVE.Methods.Flight_Dynamics.Dynamic_Stability.Approximations.short_period.short_period	(	velocity,
		density,
		S_gross_w,
		mac,
		Cm_q,
		Cz_alpha,
		mass,
		Cm_alpha,
		Iy,
		Cm_alpha_dot
	)

This calculates the natural frequency and damping ratio for the approximate short
period characteristics        
        
Assumptions:
    X-Z axis is plane of symmetry
    Constant mass of aircraft
    Origin of axis system at c.g. of aircraft
    Aircraft is a rigid body
    Earth is inertial reference frame
    Perturbations from equilibrium are small
    Flow is Quasisteady
    Constant forward airspeed
    Neglect Cz_alpha_dot and Cz_q
    Theta = 0
    
Source:
    J.H. Blakelock, "Automatic Control of Aircraft and Missiles" Wiley & Sons, Inc. New York, 1991, p 46-50.
    
Inputs:
    velocity - flight velocity at the condition being considered                                          [meters/seconds]
    density - flight density at condition being considered                                                [kg/meters**3]
    S_gross_w - area of the wing                                                                          [meters**2]
    mac - mean aerodynamic chord of the wing                                                              [meters]
    Cm_q - coefficient for the change in pitching moment due to pitch rate                                [dimensionless]
    Cz_alpha - coefficient for the change in Z force due to the angle of attack                           [dimensionless]
    mass - mass of the aircraft                                                                           [kilograms]
    Cm_alpha - coefficient for the change in pitching moment due to angle of attack                       [dimensionless]
    Iy - moment of interia about the body y axis                                                          [kg * meters**2]
    Cm_alpha_dot - coefficient for the change in pitching moment due to rate of change of angle of attack [dimensionless]

Outputs:
    output - a data dictionary with fields:
    w_n - natural frequency of the short period mode                                                      [radian/second]
    zeta - damping ratio of the short period mode                                                         [dimensionless]

Properties Used:
    N/A

◆ spiral()

def SUAVE.Methods.Flight_Dynamics.Dynamic_Stability.Approximations.spiral.spiral	(	mass,
		velocity,
		density,
		S_gross_w,
		Cl_p,
		Cn_Beta,
		Cy_phi,
		Cl_Beta,
		Cn_r,
		Cl_r
	)

spiral.py

Created: Apr 2014, A. Wendorff Modified: Jan 2016, E. Botero

This calcualtes the approximate time constant for the spiral mode         
  
Assumptions:
    Linearized equations of motion
    X-Z axis is plane of symmetry
    Constant mass of aircraft
    Origin of axis system at c.g. of aircraft
    Aircraft is a rigid body
    Earth is inertial reference frame
    Perturbations from equilibrium are small
    Flow is Quasisteady      
               
Source:
    J.H. Blakelock, "Automatic Control of Aircraft and Missiles" Wiley & Sons, Inc. New York, 1991, p 142.        
    
Inputs:
    mass - mass of the aircraft                                            [kilograms]
    velocity - flight velocity at the condition being considered           [meters/seconds]
    density - flight density at condition being considered                 [kg/meters**3]
    S_gross_w - area of the wing                                           [meters**2]
    Cl_p - change in rolling moment due to the rolling velocity            [dimensionless]
    Cn_Beta - coefficient for change in yawing moment due to sideslip      [dimensionless]
    Cy_phi - coefficient for change in sideforce due to aircraft roll      [dimensionless] (Usually equals C_L)
    Cl_Beta - coefficient for change in rolling moment due to sideslip     [dimensionless]
    Cn_r - coefficient for change in yawing moment due to yawing velocity  [dimensionless]
    Cl_r - coefficient for change in rolling moment due to yawing velocity [dimensionless] (Usually equals C_L/4)

Outputs:
    spiral_tau - time constant for the spiral mode                         [seconds] (positive values are bad)
 
Properties Used:
    N/A

Functions

Detailed Description

Function Documentation

◆ dutch_roll()

◆ phugoid()

◆ roll()

◆ short_period()

◆ spiral()