Source code for RCAIDE.Library.Components.Powertrain.Propulsors.Turbojet
# RCAIDE/Library/Components/Propulsors/Turbojet.py
#
#
# Created: Mar 2024, M. Clarke
# ----------------------------------------------------------------------------------------------------------------------
# IMPORT
# ----------------------------------------------------------------------------------------------------------------------
## RCAIDE imports
from RCAIDE.Framework.Core import Data
from . import Propulsor
from RCAIDE.Library.Methods.Powertrain.Propulsors.Turbojet .append_turbojet_conditions import append_turbojet_conditions
from RCAIDE.Library.Methods.Powertrain.Propulsors.Turbojet .compute_turbojet_performance import compute_turbojet_performance, reuse_stored_turbojet_data
# ----------------------------------------------------------------------
# Turbojet Propulsor
# ----------------------------------------------------------------------
[docs]
class Turbojet(Propulsor):
"""
A turbojet propulsion system model that simulates the performance of a turbojet engine.
Attributes
----------
tag : str
Identifier for the turbojet engine. Default is 'Turbojet'.
nacelle : Component
Nacelle component of the engine. Default is None.
ram : Component
Ram inlet component. Default is None.
inlet_nozzle : Component
Inlet nozzle component. Default is None.
low_pressure_compressor : Component
Low pressure compressor component. Default is None.
high_pressure_compressor : Component
High pressure compressor component. Default is None.
low_pressure_turbine : Component
Low pressure turbine component. Default is None.
high_pressure_turbine : Component
High pressure turbine component. Default is None.
combustor : Component
Combustor component. Default is None.
afterburner : Component
Afterburner component. Default is None.
core_nozzle : Component
Core exhaust nozzle component. Default is None.
length : float
Length of the engine [m]. Default is 0.0.
bypass_ratio : float
Engine bypass ratio. Default is 0.0.
design_isa_deviation : float
ISA temperature deviation at design point [K]. Default is 0.0.
design_altitude : float
Design altitude of the engine [m]. Default is 0.0.
afterburner_active : bool
Flag indicating if afterburner is in use. Default is False.
specific_fuel_consumption_reduction_factor : float
Specific fuel consumption adjustment factor (Less than 1 is a reduction). Default is 0.0.
compressor_nondimensional_massflow : float
Non-dimensional mass flow through the compressor. Default is 0.0.
reference_temperature : float
Reference temperature for calculations [K]. Default is 288.15.
reference_pressure : float
Reference pressure for calculations [Pa]. Default is 101325.0.
design_thrust : float
Design thrust of the engine [N]. Default is 0.0.
design_mass_flow_rate : float
Design mass flow rate [kg/s]. Default is 0.0.
OpenVSP_flow_through : bool
Flag for OpenVSP flow-through analysis. Default is False.
areas : Data
Collection of engine areas
- wetted : float
Wetted area [m²]. Default is 0.0.
- maximum : float
Maximum cross-sectional area [m²]. Default is 0.0.
- exit : float
Exit area [m²]. Default is 0.0.
- inflow : float
Inflow area [m²]. Default is 0.0.
Notes
-----
The Turbojet class inherits from the Propulsor base class and implements
methods for computing turbojet engine performance. Unlike a turbofan engine,
a turbojet does not have a bypass flow and all air goes through the core.
**Definitions**
'ISA'
International Standard Atmosphere - standard atmospheric model
'SFC'
Specific Fuel Consumption - fuel efficiency metric
'OpenVSP'
Open Vehicle Sketch Pad - open-source parametric aircraft geometry tool
See Also
--------
RCAIDE.Library.Components.Powertrain.Propulsors.Propulsor
RCAIDE.Library.Components.Powertrain.Propulsors.Turbofan
"""
def __defaults__(self):
# setting the default values
self.tag = 'Turbojet'
self.nacelle = None
self.ram = None
self.inlet_nozzle = None
self.low_pressure_compressor = None
self.high_pressure_compressor = None
self.low_pressure_turbine = None
self.high_pressure_turbine = None
self.combustor = None
self.afterburner = None
self.core_nozzle = None
self.length = 0.0
self.diameter = 0.0
self.height = 0.0
self.bypass_ratio = 0.0
self.design_isa_deviation = 0.0
self.design_altitude = 0.0
self.afterburner_active = False
self.specific_fuel_consumption_reduction_factor = 0.0
self.compressor_nondimensional_massflow = 0.0
self.reference_temperature = 288.15
self.reference_pressure = 1.01325*10**5
self.design_thrust = 0.0
self.design_mass_flow_rate = 0.0
self.OpenVSP_flow_through = False
#areas needed for drag; not in there yet
self.areas = Data()
self.areas.wetted = 0.0
self.areas.maximum = 0.0
self.areas.exit = 0.0
self.areas.inflow = 0.0
[docs]
def append_operating_conditions(self,segment,energy_conditions,noise_conditions=None):
append_turbojet_conditions(self,segment,energy_conditions,noise_conditions)
return
[docs]
def unpack_propulsor_unknowns(self,segment):
return
[docs]
def pack_propulsor_residuals(self,segment):
return
[docs]
def append_propulsor_unknowns_and_residuals(self,segment):
return
[docs]
def reuse_stored_data(turbojet,state,network,stored_propulsor_tag = None,center_of_gravity = [[0, 0, 0]]):
thrust,moment,power_mech,power_elec = reuse_stored_turbojet_data(turbojet,state,network,stored_propulsor_tag,center_of_gravity)
return thrust,moment,power_mech,power_elec
