Python atmosphere_calc Examples

Example #1

File: main_conventional_72.py Project: araquila/sinkorporated

print_payloadrange_jet = False
print_payloadrange_tbp = True
weight_fractions = False

## INPUTS AND CONSTANTS
# fuel efficiency
chosen_fuel_energy_density = energy_density_kerosene
fuel_efficiency_factor = energy_density_kerosene/chosen_fuel_energy_density

# Flight parameters
s_landing = 1400                    #[m]
altitude = 7600
V_landing = 48.93                   #[m/s] maximum landing speed that is allowed on a runway of 1400 m this is set for all aircraft

# Atmospherical parameters at cruise altitude
temperature, pressure, rho, speed_of_sound = atmosphere_calc(altitude, temperature0, temperature_gradient, g, R, gamma)

# Initial jet and tbp aircraft parameters
C_fe = 0.003
S = 1
S_wet = 5 * S
c = 6.88                                  #[m/s]

# Other jet parameters
A_jet = 10
e_jet = 0.8
M_cruise_jet = 0.8
V_cruise_jet =  M_cruise_jet*speed_of_sound                 # [m/s]
S_jet = 61
TOP_jet = 6698
C_L_cruise_jet = 0.4

Example #2

import parameters as p
import numpy as np
import matplotlib.pyplot as plt

from atmosphere import atmosphere_calc

# Atmosphere Input Parameters
t0 = p.temperature0
t_gradient = p.temperature_gradient
atR = p.R
atgamma = p.gamma
g = p.g
rho0 = p.rho0
altitude1 = 8000
temperature, pressure, rho, speed_of_sound = atmosphere_calc(
    altitude1, t0, t_gradient, g, atR, atgamma)
rho = rho * rho0

# Aircraft Input Parameters
S = p.S
W = p.MTOW
A = p.A
e = p.e
CD0 = p.Cd0
CLmaxprofile = 1.5
CLmaxto = (p.C_L_max_TO + CLmaxprofile) / 2 * 0.85
CLmaxld = (p.C_L_max_land + CLmaxprofile) / 2 * 0.85
CLmaxclean = CLmaxprofile * 0.85
#CDcruise = CD0 + CLcruise**2/(np.pi*A*e)

V_S = np.sqrt((2 * W) / (rho * CLmaxclean * S))

Example #3

File: takeoff_validation.py Project: araquila/sinkorporated

import parameters as p
import numpy as np
import matplotlib.pyplot as plt

from atmosphere import atmosphere_calc

# Runway Input Parameters
g = p.g
rho0 = p.rho0
mu = 0.02

altrange = np.linspace(0, 1500, 100)
dist_TO = []

for altitude in altrange:
    temperature, pressure, rho, speed_of_sound = atmosphere_calc(
        altitude, p.temperature0, p.temperature_gradient, p.g, p.R, p.gamma)
    rho = 1.225 * rho
    pressure = 101325 * pressure

    # Aircraft Input Parameters
    S = 61
    CL = 1.6
    CD0 = 0.02
    CD = CD0 + CL**2 / (np.pi * p.A * p.e)
    P = 3.689e6
    W = 223668

    # Take-Off Input Parameters
    V_stall = np.sqrt((W / S) * (2 / rho) * (1 / CL))
    gamma = 4 * (np.pi / 180)
    n_rotation = 1.15