First aircraft

src/pyfme/aircrafts/__init__.py

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+
+

src/pyfme/aircrafts/cessna_310.py

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+# -*- coding: utf-8 -*-
+"""
+Created on Sun Jan 3 18:44:39 2016
+
+@author:olrosales@gmail.com
+
+@AeroPython
+"""
+
+# Aircraft = Cessna 310
+
+
+import numpy as np
+
+from pyfme.utils.anemometry import calculate_dynamic_pressure
+
+
+def geometric_Data():
+
+ """ Provides the value of some geometric data.
+
+ Returns
+ ----
+
+ Sw : float
+ Wing surface (ft2 * 0.3048 * 0.3048 = m2)
+ c : foat
+ Mean aerodynamic Chord (ft * 0.3048 = m)
+ span : float
+ Wing span (ft * 0.3048 = m)
+
+ References
+ ----------
+ AIRCRAFT DYNAMICS From modelling to simulation (Marcello R. Napolitano)
+ page 589
+ """
+
+ Sw = 175 * 0.3048 * 0.3048 # m2
+ c = 4.79 * 0.3048 # m
+ span = 36.9 * 0.3048 # m
+
+ return Sw, c, span
+
+
+def mass_and_Inertial_Data():
+
+ """ Provides the value of some mass and inertial data.
+
+ Returns
+ ------
+ mass : float
+ mass (lb * 0.453592 = kg)
+ Ixx_b : float
+ Moment of Inertia x-axis ( slug * ft2 * 1.3558179 = Kg * m2)
+ Iyy_b : float
+ Moment of Inertia y-axis ( slug * ft2 * 1.3558179 = Kg * m2)
+ Izz_b : float
+ Moment of Inertia z-axis ( slug * ft2 * 1.3558179 = Kg * m2)
+ Ixz_b : float
+ Product of Inertia xz-plane ( slug * ft2 * 1.3558179 = Kg * m2)
+ inertia : array_like
+ I_xx_b,I_yy_b,I_zz_b]
+
+ References
+ ----------
+ AIRCRAFT DYNAMICS From modelling to simulation (Marcello R. Napolitano)
+ page 589
+ """
+
+ mass = 4600 * 0.453592 # kg
+ Ixx_b = 8884 * 1.3558179 # Kg * m2
+ Iyy_b = 1939 * 1.3558179 # Kg * m2
+ Izz_b = 11001 * 1.3558179 # Kg * m2
+ Ixz_b = 0 * 1.3558179 # Kg * m2
+
+ inertia = np.diag([Ixx_b, Iyy_b, Izz_b])
+
+ return mass, inertia
+
+
+def long_Aero_Coefficients():
+
+ """Assigns the value of the coefficients
+ of stability in cruise conditions and order them in a matrix.
+
+
+ CL_0 is the lift coefficient evaluated at the initial condition
+ CL_a is the lift stability derivative with respect to the angle of attack
+ CL_de is the lift stability derivative with respect to the elevator
+ deflection
+ CL_dih is the lift stability derivative with respect to the stabilator
+ deflection
+
+ CD_0 is the drag coefficient evaluated at the initial condition
+ CD_a is the drag stability derivative with respect to the angle of attack
+
+ Cm_0 is the pitching moment coefficient evaluated at the condition
+ (alpha0 = deltaE = deltaih = 0º)
+ Cm_a is the pitching moment stability derivative with respect to the angle
+ of attack
+ Cm_de is the pitching moment stability derivative with respect to the
+ elevator deflection
+ Cm_dih is the pitching moment stability derivative with respect to the
+ stabilator deflection
+
+ Returns
+ -------
+
+ Long_coef_matrix : array_like
+ [
+ [CL_0, CL_a, CL_de, CL_dih],
+ [CD_0, CD_a, 0, 0],
+ [Cm_0, Cm_a, Cm_de, Cm_dih]
+ ]
+
+
+ References
+ ----------
+ AIRCRAFT DYNAMICS From modelling to simulation (Marcello R. Napolitano)
+ page 589
+ """
+ CL_0 = 0.288
+ CL_a = 4.58
+ CL_de = 0.81
+ CL_dih = 0
+
+ CD_0 = 0.029
+ CD_a = 0.160
+
+ Cm_0 = 0.07
+ Cm_a = -0.137
+ Cm_de = -2.26
+ Cm_dih = 0
+
+ long_coef_matrix = np.array([
+ [CL_0, CL_a, CL_de, CL_dih],
+ [CD_0, CD_a, 0, 0],
+ [Cm_0, Cm_a, Cm_de, Cm_dih]
+ ])
+
+ return long_coef_matrix
+
+
+def lat_Aero_Coefficients():
+
+ """Assigns the value of the coefficients
+ of stability in cruise conditions and order them in a matrix.
+
+ CY_b is the side force stability derivative with respect to the
+ angle of sideslip
+ CY_da is the side force stability derivative with respect to the
+ aileron deflection
+ CY_dr is the side force stability derivative with respect to the
+ rudder deflection
+
+ Cl_b is the rolling moment stability derivative with respect to
+ angle of sideslip
+ Cl_da is the rolling moment stability derivative with respect to
+ the aileron deflection
+ Cl_dr is the rolling moment stability derivative with respect to
+ the rudder deflection
+
+ Cn_b is the yawing moment stability derivative with respect to the
+ angle of sideslip
+ Cn_da is the yawing moment stability derivative with respect to the
+ aileron deflection
+ Cn_dr is the yawing moment stability derivative with respect to the
+ rudder deflection
+
+ returns
+ -------
+ Long_coef_matrix : array_like
+ [
+ [CY_b, CY_da, CY_dr],
+ [Cl_b, Cl_da, Cl_dr],
+ [Cn_b, Cn_da, Cn_dr]
+ ]
+
+
+ References
+ ----------
+ AIRCRAFT DYNAMICS From modelling to simulation (Marcello R. Napolitano)
+ page 590
+ """
+
+ CY_b = -0.698
+ CY_da = 0
+ CY_dr = 0.230
+
+ Cl_b = -0.1096
+ Cl_da = 0.172
+ Cl_dr = 0.0192
+
+ Cn_b = 0.1444
+ Cn_da = -0.0168
+ Cn_dr = -0.1152
+
+ lat_coef_matrix = np.array([
+ [CY_b, CY_da, CY_dr],
+ [Cl_b, Cl_da, Cl_dr],
+ [Cn_b, Cn_da, Cn_dr]
+ ])
+
+ return lat_coef_matrix
+
+
+def get_aerodynamic_forces(TAS, rho, alpha, beta, delta_e, ih, delta_ail,
+ delta_r):
+
+ """ Calculates forces
+
+ Parameters
+ ----------
+
+ rho : float
+ density (kg/(m3))
+ TAS : float
+ velocity (m/s)
+
+ alpha : float
+ attack angle (rad).
+ beta : float
+ sideslip angle (rad).
+ delta_e : float
+ elevator deflection (rad).
+ ih : float
+ stabilator deflection (rad).
+ delta_ail : float
+ aileron deflection (rad).
+ delta_r : float
+ rudder deflection (rad).
+
+ Returns
+ -------
+
+ forces : array_like
+ 3 dimensional vector with (F_x_s, F_y_s, F_z_s) forces
+ in stability axes.
+
+ References
+ ----------
+ AIRCRAFT DYNAMICS From modelling to simulation (Marcello R. Napolitano)
+ chapter 3 and 4
+ """
+
+ long_coef_matrix = long_Aero_Coefficients()
+ lat_coef_matrix = lat_Aero_Coefficients()
+
+ CL_0, CL_a, CL_de, CL_dih = long_coef_matrix[0, :]
+ CD_0, CD_a, CD_de, CD_dih = long_coef_matrix[1, :]
+ CY_b, CY_da, CY_dr = lat_coef_matrix[0, :]
+
+ CL_full = CL_0 + CL_a * alpha + CL_de * delta_e + CL_dih * ih
+ CD_full = CD_0 + CD_a * alpha + CD_de * delta_e + CD_dih * ih
+ CY_full = CY_b * beta + CY_da * delta_ail + CY_dr * delta_r
+
+ Sw = geometric_Data()[0]
+
+ aerodynamic_forces = calculate_dynamic_pressure(rho, TAS) *\
+ Sw * np.array([-CD_full, CY_full, -CL_full]) # N
+
+ return aerodynamic_forces
+
+
+def get_aerodynamic_moments(TAS, rho, alpha, beta, delta_e, ih, delta_ail,
+ delta_r):
+
+ """ Calculates forces
+
+ Parameters
+ ----------
+
+ rho : float
+ density (kg/m3)
+ TAS : float
+ velocity (m/s)
+
+ alpha : float
+ attack angle (rad).
+ beta : float
+ sideslip angle (rad).
+ delta_e : float
+ elevator deflection (rad).
+ ih : float
+ stabilator deflection (rad).
+ delta_ail : float
+ aileron deflection (rad).
+ delta_r : float
+ rudder deflection (rad).
+
+ returns
+ -------
+
+ moments : array_like
+ 3 dimensional vector with (Mxs, Mys, Mzs) forces
+ in stability axes.
+
+ References
+ ----------
+ AIRCRAFT DYNAMICS From modelling to simulation (Marcello R. Napolitano)
+ chapter 3 and 4
+ """
+
+ long_coef_matrix = long_Aero_Coefficients()
+ lat_coef_matrix = lat_Aero_Coefficients()
+
+ Cm_0, Cm_a, Cm_de, Cm_dih = long_coef_matrix[2, :]
+ Cl_b, Cl_da, Cl_dr = lat_coef_matrix[1, :]
+ Cn_b, Cn_da, Cn_dr = lat_coef_matrix[2, :]
+
+ Cm_full = Cm_0 + Cm_a * alpha + Cm_de * delta_e + Cm_dih * ih
+ Cl_full = Cl_b * beta + Cl_da * delta_ail + Cl_dr * delta_r
+ Cn_full = Cn_b * beta + Cn_da * delta_ail + Cn_dr * delta_r
+
+ span = geometric_Data()[2]
+ c = geometric_Data()[1]
+ Sw = geometric_Data()[0]
+
+ aerodynamic_moments = calculate_dynamic_pressure(rho, TAS) * Sw\
+ * np.array([Cl_full * span, Cm_full * c, Cn_full * span])
+
+ return aerodynamic_moments
+
+
+def get_engine_force(delta_t):
+
+ """ Calculates forces
+
+ Parameters
+ ----------
+
+ delta_t : float
+ trust_lever (0 = 0 Newton, 1 = CT)
+
+
+ returns
+ -------
+
+ engine_force : float
+ Trust (N)
+
+ References
+ ----------
+ Airplane Flight Dyanamics and Automatic Flight Controls part I - Jan Roskam
+ """
+ if delta_t >= 0 and delta_t <= 1:
+
+ Ct = 0.031 * delta_t
+
+ q_cruise = 91.2 * 47.880172 # Pa
+
+ Sw = geometric_Data()[0]
+
+ engine_force = Ct * Sw * q_cruise # N
+
+ else:
+ raise ValueError('delta_t must be between 0 and 1')
+
+ return engine_force