def test_01(self):
# Truth values from NASA RP 1046
Value = SA.alt2density(8000)
Truth = 0.06011
self.assertLessEqual(RE(Value, Truth), 1e-5)
def __init__(self, h):
# Cantera Solution object
self.gas = ct.Solution('air.xml')
# Discretised altitude steps
self.h = h
# Average molecular diameter of gas
self.d = 4E-10
self.steps = len(h)
self.rho = np.zeros(self.steps)
self.p = np.zeros(self.steps)
self.T = np.zeros(self.steps)
self.a = np.zeros(self.steps)
self.k = np.zeros(self.steps)
self.mu = np.zeros(self.steps)
for index, alt in enumerate(self.h):
self.rho[index] = atm.alt2density(alt, alt_units='m', density_units='kg/m**3')
self.p[index] = atm.alt2press(alt, press_units='pa', alt_units='m')
self.T[index] = atm.alt2temp(alt, alt_units='m', temp_units='K')
self.a[index] = atm.temp2speed_of_sound(self.T[index], temp_units='K', speed_units='m/s')
for index, alt in enumerate(self.h):
self.gas.TP = self.T[index], self.p[index]
self.k[index] = self.gas.cp / self.gas.cv
self.mu[index] = self.gas.viscosity
print 'ATMOSPHERIC MODEL COMPUTED (US76)'
return None
def test_02(self):
# test psf
# Truth values from NASA RP 1046
Value = SA.alt2density(49000)
Truth = 0.012215
self.assertLessEqual(RE(Value, Truth), 2e-5)
def test_06(self):
# truth value calculated from program at:
# http://www.sworld.com.au/steven/space/atmosphere/
Value = SA.alt2density(65322, density_units='kg/m**3',
alt_units='m')
Truth = 1.4296e-4
self.assertLessEqual(RE(Value, Truth), 5e-5)
def test_03(self):
# test pa and m
# Truth values from NASA RP 1046
Value = SA.alt2density(25000, alt_units='m',
density_units='kg/m**3')
Truth = 0.039466
self.assertLessEqual(RE(Value, Truth), 1e-5)
def test_07(self):
# test units in other order
# truth value calculated from program at:
# http://www.sworld.com.au/steven/space/atmosphere/
Value = SA.alt2density(80956, density_units='kg/m**3',
alt_units='m')
Truth = 1.3418e-5
self.assertLessEqual(RE(Value, Truth), 3e-5)
def __init__(self, h, T_thermosphere):
# Cantera Solution object
self.gas = ct.Solution('air.xml')
# Discretised altitude steps
self.h = h
# Average molecular diameter of gas
self.d = 4E-10
# Ratio of specific heats
self.steps = len(h)
self.rho = np.zeros(self.steps)
self.p = np.zeros(self.steps)
self.T = np.zeros(self.steps)
self.a = np.zeros(self.steps)
self.k = np.zeros(self.steps)
self.mu = np.zeros(self.steps)
# Call Jacchia77 model
data = j77.j77sri(np.max(h), T_thermosphere)
data_np = np.array(data)
h_int = spint.griddata
T_int = spint.griddata
mw_int = spint.griddata
n = spint.griddata
for index, alt in enumerate(self.h):
self.rho[index] = atm.alt2density(alt, alt_units='m', density_units='kg/m**3')
self.p[index] = atm.alt2press(alt, press_units='pa', alt_units='m')
self.T[index] = atm.alt2temp(alt, alt_units='m', temp_units='K')
self.a[index] = atm.temp2speed_of_sound(self.T[index], temp_units='K', speed_units='m/s')
for index, alt in enumerate(self.h):
self.gas.TP = self.T[index], self.p[index]
self.k[index] = self.gas.cp / self.gas.cv
self.mu[index] = self.gas.viscosity
return None
def Aerodynamic_Acceleration(self, body):
'''Computes the spacecraft's acceleration due to aerodynamic drag whilst encountering the body's atmosphere.'''
# Altitude of spacecraft above the body [m]
H = self.Altitude(body)
# Atmospheric density at this altitude [kg/m^3]
try:
rho = alt2density(H, alt_units='m', density_units='kg/m**3')
except:
return 0.0
# Coefficient of drag
CD = 2.2
# Planaform area of satellite perpendicular to velocity
A = self.area
# Velocity vector with respect to body
v_vec = self.Position_and_Velocity_WRT(body)[1]
# Velocity magnitude with respect to body
v = np.linalg.norm(v_vec)
# Velocity vector direction from body to spacecraft
v_hat = np.divide(v_vec, v)
# Most recent mass of spacecraft
m = self.masses[-1]
# Aerodynamic acceleration of spacecraft due to body's atmosphere
aa = (rho * CD * A * v**2 * v_hat) / (2 * m)
return aa
def Aerodynamic_Acceleration(self, body):
'''Computes the spacecraft's acceleration due to aerodynamic drag whilst encountering the body's atmosphere.'''
# Altitude of spacecraft above the body [m]
H = self.Altitude(body)
# Atmospheric density at this altitude [kg/m^3]
try:
rho = alt2density(H, alt_units='m', density_units='kg/m**3')
except:
return 0.0
# Coefficient of drag
CD = 2.2
# Planaform area of satellite perpendicular to velocity
A = self.area
# Velocity vector with respect to body
v_vec = self.Position_and_Velocity_WRT(body)[1]
# Velocity magnitude with respect to body
v = np.linalg.norm(v_vec)
# Velocity vector direction from body to spacecraft
v_hat = np.divide(v_vec, v)
# Most recent mass of spacecraft
m = self.masses[-1]
# Aerodynamic acceleration of spacecraft due to body's atmosphere
aa = (rho * CD * A * v**2 * v_hat) / (2 * m)
return aa
# In[193]:
Resolution = 2000
Start_Pa = 0.1
# ## Preliminary Calculations
# ### The operating environment
# The environment in which the aircraft is expected to operate plays a very important role in many of the conceptual design calculations to follow. The conditions corresponding to the current design brief are computed as follows:
# In[194]:
SeaLevelDens_kgm3 = ISA.alt2density(0, alt_units='ft', density_units='kg/m**3')
print('ISA density at Sea level elevation: {:0.3f} kg/m^3'.format(
SeaLevelDens_kgm3))
# In[195]:
TakeOffDens_kgm3 = ISA.alt2density(TakeOffElevation_m,
alt_units='m',
density_units='kg/m**3')
print('ISA density at take-off elevation: {:0.3f} kg/m^3'.format(
TakeOffDens_kgm3))
# In[196]:
ClimbAltDens_kgm3 = ISA.alt2density(ROCAlt_m,
alt_units='m',
