def get_cm_derivatives(weight, height, v_t, delta_cg, delta_de, dd_dalpha):
"""
:param weight:
:param height:
:param v_t: True airspeed
:param delta_cg: Change in CG location
:param delta_de: Change in elevator deflection
:param delta_alpha: Change in angle of attack
:return: cm_alpha, cm_delta
"""
averaged_weight = np.mean(weight)
temperature = isa.get_t_at_height(height)
pressure = isa.get_p_at_temperature(temperature)
density = isa.get_rho(pressure, temperature)
averaged_density = np.mean(density)
averaged_v_t = np.mean(v_t)
cm_delta = (
-1 / delta_de * averaged_weight /
(0.5 * averaged_density * np.power(averaged_v_t, 2) * cessna.S) *
delta_cg / cessna.c)
cm_alpha = -dd_dalpha * cm_delta
return cm_alpha, cm_delta
def get_thrust_coefficient(height, thrust, velocity):
"""
:param height:
:param thrust:
:param rps: Revolutions per second
:return: Thrust coefficient
"""
temperature = isa.get_t_at_height(height)
pressure = isa.get_p_at_temperature(temperature)
density = isa.get_rho(pressure, temperature)
return thrust / (density * np.power(velocity, 2) * np.power(cessna.D, 2))
def get_reduced_equivalent_airspeed(weight, height, v_t):
"""
:param weight:
:param height:
:param v_t: True airspeed
:return: Reduced equivalent airspeed
"""
temperature = isa.get_t_at_height(height)
pressure = isa.get_p_at_temperature(temperature)
density = isa.get_rho(pressure, temperature)
v_e = v_t * np.sqrt(density / isa.rho_0)
return v_e * np.sqrt(cessna.W / weight)
def get_mach_number(height, v_c):
"""
:param height:
:param v_c: Calibrated airspeed
:return: Mach number
"""
pressure = isa.get_p_at_temperature(isa.get_t_at_height(height))
innermost = np.power(
1 + (isa.gamma - 1) / (2 * isa.gamma) *
(isa.rho_0 / isa.p_0) * np.power(v_c, 2),
isa.gamma / (isa.gamma - 1),
)
middle = np.power(1 + isa.p_0 / pressure * (innermost - 1),
(isa.gamma - 1) / isa.gamma)
M = np.sqrt(2 / (isa.gamma - 1) * (middle - 1))
return M
def get_cl_params(alpha, weight, height, true_airspeed):
"""
Inputs should be single values or numpy arrays
:param alpha: angle of attack
:param weight:
:param height:
:param true_airspeed:
:return: cn_alpha, alpha_0, cn from flight
"""
temperature = isa.get_t_at_height(height)
pressure = isa.get_p_at_temperature(temperature)
density = isa.get_rho(pressure, temperature)
cn = weight / (0.5 * density * np.power(true_airspeed, 2) * cessna.S)
popt, _ = opt.curve_fit(cl_curve, alpha, cn)
cn_alpha, alpha_0 = popt
return cn_alpha, alpha_0, cn
def run_thrust_exe(height, calibrated_airspeed, fuel_flow_left,
fuel_flow_right, measured_temperature):
"""[summary]
:param height: from stationary measurements
:param calibrated_airspeed: from stationary measurements
:param fuel_flow_left: from stationary measurements
:param fuel_flow_right: from stationary measurements
:param measured_temperature: from stationary measurements
:return: array of thrust values for both engines for each stationary measurement from thrust.exe program
"""
mach_number = get_mach_number(height, calibrated_airspeed)
temperature_delta = measured_temperature - isa.get_t_at_height(height)
matlab_data = np.vstack([
height, mach_number, temperature_delta, fuel_flow_left, fuel_flow_right
]).T
np.savetxt("matlab.dat", matlab_data, delimiter=" ")
subprocess.call("thrust.exe")
thrust = np.genfromtxt("thrust.dat")
thrust = thrust[:, 0] + thrust[:, 1]
return thrust
def get_cd_params(alpha, weight, height, true_airspeed, thrust):
"""
Inputs should be single values or numpy arrays
:param alpha: angle of attack
:param weight:
:param height:
:param true_airspeed:
:param thrust:
:return: cd_0, e, cd from flight
"""
temperature = isa.get_t_at_height(height)
pressure = isa.get_p_at_temperature(temperature)
density = isa.get_rho(pressure, temperature)
cn_alpha, alpha_0, _ = get_cl_params(alpha, weight, height, true_airspeed)
cl = cn_alpha * (alpha - alpha_0)
cd = thrust / (0.5 * density * np.power(true_airspeed, 2) * cessna.S)
popt, _ = opt.curve_fit(cd_curve, cl, cd)
cd_0, e = popt
return cd_0, e, cd