Source code for RCAIDE.Library.Methods.Powertrain.Propulsors.Turbojet.design_turbojet
# RCAIDE/Methods/Energy/Propulsors/Turbojet/design_turbojet.py
#
#
# Created: Jul 2023, M. Clarke
# ----------------------------------------------------------------------------------------------------------------------
# IMPORT
# ----------------------------------------------------------------------------------------------------------------------
# RCAIDE Imports
import RCAIDE
from RCAIDE.Framework.Mission.Common import Conditions
from RCAIDE.Library.Methods.Powertrain.Converters.Ram import compute_ram_performance
from RCAIDE.Library.Methods.Powertrain.Converters.Combustor import compute_combustor_performance
from RCAIDE.Library.Methods.Powertrain.Converters.Compressor import compute_compressor_performance
from RCAIDE.Library.Methods.Powertrain.Converters.Turbine import compute_turbine_performance
from RCAIDE.Library.Methods.Powertrain.Converters.Supersonic_Nozzle import compute_supersonic_nozzle_performance
from RCAIDE.Library.Methods.Powertrain.Converters.Compression_Nozzle import compute_compression_nozzle_performance
from RCAIDE.Library.Methods.Powertrain.Propulsors.Turbojet import size_core
from RCAIDE.Library.Methods.Powertrain import setup_operating_conditions
# Python package imports
import numpy as np
# ----------------------------------------------------------------------------------------------------------------------
# Design Turbojet
# ----------------------------------------------------------------------------------------------------------------------
[docs]
def design_turbojet(turbojet):
"""
Designs a turbojet engine by computing performance properties and sizing components based on design conditions.
Parameters
----------
turbojet : Turbojet
Turbojet engine object containing design parameters and components
- design_mach_number : float
Design point Mach number [-]
- design_altitude : float
Design point altitude [m]
- design_isa_deviation : float
ISA temperature deviation [K]
- working_fluid : Gas
Working fluid object for gas properties
- Components:
- ram : Ram
- inlet_nozzle : Compression_Nozzle
- low_pressure_compressor : Compressor
- high_pressure_compressor : Compressor
- combustor : Combustor
- high_pressure_turbine : Turbine
- low_pressure_turbine : Turbine
- core_nozzle : Supersonic_Nozzle
Returns
-------
None
Updates turbojet object attributes in-place:
- mass_flow_rate_design : float
Design core mass flow rate [kg/s]
- design_core_massflow : float
Core mass flow at design point [kg/s]
Notes
-----
This function performs the following steps:
1. Computes atmospheric conditions at design point
2. Sets up freestream conditions
3. Links and analyzes flow through each component:
- Ram inlet
- Inlet nozzle
- Low pressure compressor
- High pressure compressor
- Combustor
- High pressure turbine
- Low pressure turbine
- Core nozzle
4. Sizes the core based on design thrust requirements
5. Computes static sea level performance
**Major Assumptions**
* Quasi-one-dimensional flow
* Each component operates in steady state
* Perfect gas behavior in non-combustion sections
* US Standard Atmosphere 1976 model
* Earth gravity model
* Design point defines core sizing
**Theory**
The design process follows standard gas turbine design principles:
.. math::
\\text{Mass flow continuity: } \\dot{m}_{in} = \\dot{m}_{out}
\\text{Power Balance: } W_{compressor} = W_{turbine}
\\text{Core sizing: } \\dot{m}_{core} = \\frac{F_{design}}{F_{sp} a_0}
where:
- :math:`F_{design}` is the design thrust
- :math:`F_{sp}` is the specific thrust
- :math:`a_0` is the freestream speed of sound
References
----------
[1] Mattingly, J. D., "Elements of Gas Turbine Propulsion", McGraw-Hill, 1996
[2] Walsh, P. P., Fletcher, P., "Gas Turbine Performance", Blackwell Science, 2004
See Also
--------
RCAIDE.Library.Methods.Powertrain.Propulsors.Turbojet.compute_turbojet_performance
RCAIDE.Library.Methods.Powertrain.Propulsors.Turbojet.size_core
RCAIDE.Library.Methods.Powertrain.Propulsors.Common.compute_static_sea_level_performance
"""
#check if mach number and temperature are passed
if(turbojet.design_mach_number==None or turbojet.design_altitude==None):
#raise an error
raise NameError('The sizing conditions require an altitude and a Mach number')
else:
#call the atmospheric model to get the conditions at the specified altitude
atmosphere = RCAIDE.Framework.Analyses.Atmospheric.US_Standard_1976()
atmo_data = atmosphere.compute_values(turbojet.design_altitude,turbojet.design_isa_deviation)
planet = RCAIDE.Library.Attributes.Planets.Earth()
p = atmo_data.pressure
T = atmo_data.temperature
rho = atmo_data.density
a = atmo_data.speed_of_sound
mu = atmo_data.dynamic_viscosity
# setup conditions
conditions = RCAIDE.Framework.Mission.Common.Results()
# freestream conditions
conditions.freestream.altitude = np.atleast_1d(turbojet.design_altitude)
conditions.freestream.mach_number = np.atleast_1d(turbojet.design_mach_number)
conditions.freestream.pressure = np.atleast_1d(p)
conditions.freestream.temperature = np.atleast_1d(T)
conditions.freestream.density = np.atleast_1d(rho)
conditions.freestream.dynamic_viscosity = np.atleast_1d(mu)
conditions.freestream.gravity = np.atleast_1d(planet.compute_gravity(turbojet.design_altitude))
conditions.freestream.isentropic_expansion_factor = np.atleast_1d(turbojet.working_fluid.compute_gamma(T,p))
conditions.freestream.Cp = np.atleast_1d(turbojet.working_fluid.compute_cp(T,p))
conditions.freestream.R = np.atleast_1d(turbojet.working_fluid.gas_specific_constant)
conditions.freestream.speed_of_sound = np.atleast_1d(a)
conditions.freestream.velocity = np.atleast_1d(a*turbojet.design_mach_number)
segment = RCAIDE.Framework.Mission.Segments.Segment()
segment.state.conditions = conditions
turbojet.append_operating_conditions(segment,conditions.energy,conditions.noise)
ram = turbojet.ram
inlet_nozzle = turbojet.inlet_nozzle
low_pressure_compressor = turbojet.low_pressure_compressor
high_pressure_compressor = turbojet.high_pressure_compressor
combustor = turbojet.combustor
high_pressure_turbine = turbojet.high_pressure_turbine
low_pressure_turbine = turbojet.low_pressure_turbine
core_nozzle = turbojet.core_nozzle
# unpack component conditions
turbojet_conditions = conditions.energy.propulsors[turbojet.tag]
ram_conditions = conditions.energy.converters[ram.tag]
inlet_nozzle_conditions = conditions.energy.converters[inlet_nozzle.tag]
core_nozzle_conditions = conditions.energy.converters[core_nozzle.tag]
lpc_conditions = conditions.energy.converters[low_pressure_compressor.tag]
hpc_conditions = conditions.energy.converters[high_pressure_compressor.tag]
lpt_conditions = conditions.energy.converters[low_pressure_turbine.tag]
hpt_conditions = conditions.energy.converters[high_pressure_turbine.tag]
combustor_conditions = conditions.energy.converters[combustor.tag]
# Step 1: Set the working fluid to determine the fluid properties
ram.working_fluid = turbojet.working_fluid
# Step 2: Compute flow through the ram , this computes the necessary flow quantities and stores it into conditions
compute_ram_performance(ram,conditions)
# Step 3: link inlet nozzle to ram
inlet_nozzle_conditions.inputs.stagnation_temperature = ram_conditions.outputs.stagnation_temperature
inlet_nozzle_conditions.inputs.stagnation_pressure = ram_conditions.outputs.stagnation_pressure
inlet_nozzle_conditions.inputs.static_temperature = ram_conditions.outputs.static_temperature
inlet_nozzle_conditions.inputs.static_pressure = ram_conditions.outputs.static_pressure
inlet_nozzle_conditions.inputs.mach_number = ram_conditions.outputs.mach_number
inlet_nozzle.working_fluid = ram.working_fluid
# Step 4: Compute flow through the inlet nozzle
compute_compression_nozzle_performance(inlet_nozzle,conditions)
# Step 5: Link low pressure compressor to the inlet nozzle
lpc_conditions.inputs.stagnation_temperature = inlet_nozzle_conditions.outputs.stagnation_temperature
lpc_conditions.inputs.stagnation_pressure = inlet_nozzle_conditions.outputs.stagnation_pressure
lpc_conditions.inputs.static_temperature = inlet_nozzle_conditions.outputs.static_temperature
lpc_conditions.inputs.static_pressure = inlet_nozzle_conditions.outputs.static_pressure
lpc_conditions.inputs.mach_number = inlet_nozzle_conditions.outputs.mach_number
low_pressure_compressor.working_fluid = inlet_nozzle.working_fluid
low_pressure_compressor.reference_temperature = turbojet.reference_temperature
low_pressure_compressor.reference_pressure = turbojet.reference_pressure
# Step 6: Compute flow through the low pressure compressor
compute_compressor_performance(low_pressure_compressor,conditions)
# Step 7: Link the high pressure compressor to the low pressure compressor
hpc_conditions.inputs.stagnation_temperature = lpc_conditions.outputs.stagnation_temperature
hpc_conditions.inputs.stagnation_pressure = lpc_conditions.outputs.stagnation_pressure
hpc_conditions.inputs.static_temperature = lpc_conditions.outputs.static_temperature
hpc_conditions.inputs.static_pressure = lpc_conditions.outputs.static_pressure
hpc_conditions.inputs.mach_number = lpc_conditions.outputs.mach_number
high_pressure_compressor.working_fluid = low_pressure_compressor.working_fluid
high_pressure_compressor.reference_temperature = turbojet.reference_temperature
high_pressure_compressor.reference_pressure = turbojet.reference_pressure
# Step 8: Compute flow through the high pressure compressor
compute_compressor_performance(high_pressure_compressor,conditions)
# Step 9: Link the combustor to the high pressure compressor
combustor_conditions.inputs.stagnation_temperature = hpc_conditions.outputs.stagnation_temperature
combustor_conditions.inputs.stagnation_pressure = hpc_conditions.outputs.stagnation_pressure
combustor_conditions.inputs.static_temperature = hpc_conditions.outputs.static_temperature
combustor_conditions.inputs.static_pressure = hpc_conditions.outputs.static_pressure
combustor_conditions.inputs.mach_number = hpc_conditions.outputs.mach_number
combustor.working_fluid = high_pressure_compressor.working_fluid
# Step 10: Compute flow through the high pressor compressor
compute_combustor_performance(combustor,conditions)
hpt_conditions.inputs.stagnation_temperature = combustor_conditions.outputs.stagnation_temperature
hpt_conditions.inputs.stagnation_pressure = combustor_conditions.outputs.stagnation_pressure
hpt_conditions.inputs.fuel_to_air_ratio = combustor_conditions.outputs.fuel_to_air_ratio
hpt_conditions.inputs.static_temperature = combustor_conditions.outputs.static_temperature
hpt_conditions.inputs.static_pressure = combustor_conditions.outputs.static_pressure
hpt_conditions.inputs.mach_number = combustor_conditions.outputs.mach_number
hpt_conditions.inputs.compressor = hpc_conditions.outputs
hpt_conditions.inputs.bypass_ratio = 0.0
high_pressure_turbine.working_fluid = combustor.working_fluid
# Step 11: Compute flow through the high pressure turbine
compute_turbine_performance(high_pressure_turbine,conditions)
# Step 12: Link the low pressure turbine to the high pressure turbine
lpt_conditions.inputs.stagnation_temperature = hpt_conditions.outputs.stagnation_temperature
lpt_conditions.inputs.stagnation_pressure = hpt_conditions.outputs.stagnation_pressure
lpt_conditions.inputs.static_temperature = hpt_conditions.outputs.static_temperature
lpt_conditions.inputs.static_pressure = hpt_conditions.outputs.static_pressure
lpt_conditions.inputs.mach_number = hpt_conditions.outputs.mach_number
low_pressure_turbine.working_fluid = high_pressure_turbine.working_fluid
lpt_conditions.inputs.compressor = lpc_conditions.outputs
lpt_conditions.inputs.fuel_to_air_ratio = combustor_conditions.outputs.fuel_to_air_ratio
lpt_conditions.inputs.bypass_ratio = 0.0
# Step 16: Compute flow through the low pressure turbine
compute_turbine_performance(low_pressure_turbine,conditions)
# Step 17: Link the core nozzle to the low pressure turbine
core_nozzle_conditions.inputs.stagnation_temperature = lpt_conditions.outputs.stagnation_temperature
core_nozzle_conditions.inputs.stagnation_pressure = lpt_conditions.outputs.stagnation_pressure
core_nozzle_conditions.inputs.static_temperature = lpt_conditions.outputs.static_temperature
core_nozzle_conditions.inputs.static_pressure = lpt_conditions.outputs.static_pressure
core_nozzle_conditions.inputs.mach_number = lpt_conditions.outputs.mach_number
core_nozzle.working_fluid = low_pressure_turbine.working_fluid
# Step 18: Compute flow through the core nozzle
compute_supersonic_nozzle_performance(core_nozzle,conditions)
# Step 19: link the thrust component to the core nozzle
turbojet_conditions.core_nozzle_area_ratio = core_nozzle_conditions.outputs.area_ratio
turbojet_conditions.core_nozzle_static_pressure = core_nozzle_conditions.outputs.static_pressure
turbojet_conditions.core_nozzle_exit_velocity = core_nozzle_conditions.outputs.velocity
turbojet_conditions.fuel_to_air_ratio = combustor_conditions.outputs.fuel_to_air_ratio
turbojet_conditions.stag_temp_lpt_exit = lpc_conditions.outputs.stagnation_temperature
turbojet_conditions.stag_press_lpt_exit = lpc_conditions.outputs.stagnation_pressure
turbojet_conditions.total_temperature_reference = lpc_conditions.outputs.stagnation_temperature
turbojet_conditions.total_pressure_reference = lpc_conditions.outputs.stagnation_pressure
turbojet_conditions.flow_through_core = 1.0 #scaled constant to turn on core thrust computation
turbojet_conditions.flow_through_fan = 0.0 #scaled constant to turn on fan thrust computation
# Step 20: Size the core of the turbojet
size_core(turbojet,conditions)
# Step 21: Static Sea Level Thrust
atmo_data_sea_level = atmosphere.compute_values(0.0,0.0)
V = atmo_data_sea_level.speed_of_sound[0][0]*0.01
operating_state = setup_operating_conditions(turbojet, altitude = 0,velocity_range=np.array([V]))
operating_state.conditions.energy.propulsors[turbojet.tag].throttle[:,0] = 1.0
sls_T,_,sls_P,_,_,_ = turbojet.compute_performance(operating_state)
turbojet.sealevel_static_thrust = sls_T[0][0]
turbojet.sealevel_static_power = sls_P[0][0]
return
