Source code for RCAIDE.Library.Methods.Powertrain.Converters.Turboshaft.design_turboshaft
# RCAIDE/Library/Methods/Energy/Converters/Turboshaft/design_turboshaft.py
#
#
# Created: Jul 2023, M. Clarke
# ----------------------------------------------------------------------------------------------------------------------
# IMPORT
# ----------------------------------------------------------------------------------------------------------------------
# RCAIDE Imports
import RCAIDE
from RCAIDE.Framework.Core import Data
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.Expansion_Nozzle import compute_expansion_nozzle_performance
from RCAIDE.Library.Methods.Powertrain.Converters.Compression_Nozzle import compute_compression_nozzle_performance
from RCAIDE.Library.Methods.Powertrain.Converters.Turboshaft import size_core
from RCAIDE.Library.Methods.Powertrain import setup_operating_conditions
# Python package imports
import numpy as np
# ----------------------------------------------------------------------------------------------------------------------
# Design Turboshaft
# ----------------------------------------------------------------------------------------------------------------------
[docs]
def design_turboshaft(turboshaft):
"""
Designs and sizes a turboshaft engine based on design point conditions and performance requirements.
Parameters
----------
turboshaft : Turboshaft
Turboshaft 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 : FluidProperties
Working fluid properties object
- Components:
- ram : Ram
- inlet_nozzle : Compression_Nozzle
- compressor : Compressor
- combustor : Combustor
- high_pressure_turbine : Turbine
- low_pressure_turbine : Turbine
- core_nozzle : Expansion_Nozzle
Returns
-------
None
Results are stored in turboshaft attributes:
- design_thrust_specific_fuel_consumption : float
TSFC at design point [kg/N/s]
- design_non_dimensional_thrust : float
Non-dimensional thrust [-]
- design_core_mass_flow_rate : float
Core mass flow rate [kg/s]
- design_fuel_flow_rate : float
Fuel flow rate [kg/s]
- design_power : float
Power output [W]
- design_specific_power : float
Specific power [W/kg]
- design_power_specific_fuel_consumption : float
Power specific fuel consumption [kg/W/s]
- design_thermal_efficiency : float
Thermal efficiency [-]
- design_propulsive_efficiency : float
Propulsive efficiency [-]
Notes
-----
The function performs the following steps:
1. Computes atmospheric conditions at design altitude
2. Initializes design point conditions
3. Analyzes flow through each component sequentially:
- Ram inlet
- Inlet nozzle
- Compressor
- Combustor
- High pressure turbine
- Low pressure turbine
- Core nozzle
4. Sizes core components
5. Computes sea level static performance
**Major Assumptions**
* Standard atmospheric conditions (with possible ISA deviation)
* Steady state operation
* Perfect gas behavior
* Adiabatic component processes except combustor
* No bleed air extraction
* Design point sizing determines component characteristics
**Theory**
The design process follows standard gas turbine cycle analysis principles.
Component sizing is based on achieving required power output while maintaining
appropriate component matching throughout the engine.
**Extra modules required**
* RCAIDE.Framework.Analyses.Atmospheric.US_Standard_1976
* RCAIDE.Library.Attributes.Planets.Earth
See Also
--------
RCAIDE.Library.Methods.Powertrain.Converters.Ram.compute_ram_performance
RCAIDE.Library.Methods.Powertrain.Converters.Combustor.compute_combustor_performance
RCAIDE.Library.Methods.Powertrain.Converters.Compressor.compute_compressor_performance
RCAIDE.Library.Methods.Powertrain.Converters.Turbine.compute_turbine_performance
"""
#check if mach number and temperature are passed
if(turboshaft.design_mach_number==None or turboshaft.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(turboshaft.design_altitude,turboshaft.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(turboshaft.design_altitude)
conditions.freestream.mach_number = np.atleast_1d(turboshaft.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(turboshaft.design_altitude))
conditions.freestream.isentropic_expansion_factor = np.atleast_1d(turboshaft.working_fluid.compute_gamma(T,p))
conditions.freestream.Cp = np.atleast_1d(turboshaft.working_fluid.compute_cp(T,p))
conditions.freestream.R = np.atleast_1d(turboshaft.working_fluid.gas_specific_constant)
conditions.freestream.speed_of_sound = np.atleast_1d(a)
conditions.freestream.velocity = np.atleast_1d(a*turboshaft.design_mach_number)
fuel_line = RCAIDE.Library.Components.Powertrain.Distributors.Fuel_Line() # may not need
segment = RCAIDE.Framework.Mission.Segments.Segment()
segment.state.conditions = conditions
turboshaft.append_operating_conditions(segment,conditions.energy,conditions.noise)
ram = turboshaft.ram
inlet_nozzle = turboshaft.inlet_nozzle
compressor = turboshaft.compressor
combustor = turboshaft.combustor
high_pressure_turbine = turboshaft.high_pressure_turbine
low_pressure_turbine = turboshaft.low_pressure_turbine
core_nozzle = turboshaft.core_nozzle
turboshaft_conditions = conditions.energy.converters[turboshaft.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]
compressor_conditions = conditions.energy.converters[compressor.tag]
combustor_conditions = conditions.energy.converters[combustor.tag]
lpt_conditions = conditions.energy.converters[low_pressure_turbine.tag]
hpt_conditions = conditions.energy.converters[high_pressure_turbine.tag]
# Step 1: Set the working fluid to determine the fluid properties
ram.working_fluid = turboshaft.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
compressor_conditions.inputs.stagnation_temperature = inlet_nozzle_conditions.outputs.stagnation_temperature
compressor_conditions.inputs.stagnation_pressure = inlet_nozzle_conditions.outputs.stagnation_pressure
compressor_conditions.inputs.static_temperature = inlet_nozzle_conditions.outputs.static_temperature
compressor_conditions.inputs.static_pressure = inlet_nozzle_conditions.outputs.static_pressure
compressor_conditions.inputs.mach_number = inlet_nozzle_conditions.outputs.mach_number
compressor.working_fluid = inlet_nozzle.working_fluid
compressor.reference_temperature = turboshaft.reference_temperature
compressor.reference_pressure = turboshaft.reference_pressure
# Step 6: Compute flow through the low pressure compressor
compute_compressor_performance(compressor,conditions)
# Step 11: Link the combustor to the high pressure compressor
combustor_conditions.inputs.stagnation_temperature = compressor_conditions.outputs.stagnation_temperature
combustor_conditions.inputs.stagnation_pressure = compressor_conditions.outputs.stagnation_pressure
combustor_conditions.inputs.static_temperature = compressor_conditions.outputs.static_temperature
combustor_conditions.inputs.static_pressure = compressor_conditions.outputs.static_pressure
combustor_conditions.inputs.mach_number = compressor_conditions.outputs.mach_number
combustor.working_fluid = compressor.working_fluid
compressor.reference_temperature = turboshaft.reference_temperature
compressor.reference_pressure = turboshaft.reference_pressure
# Step 12: Compute flow through the high pressor compressor
compute_combustor_performance(combustor,conditions)
#link the high pressure turbione to the combustor
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 = compressor_conditions.outputs
hpt_conditions.inputs.bypass_ratio = 0.0
hpt_conditions.inputs.fan = Data()
hpt_conditions.inputs.fan.work_done = 0.0
high_pressure_turbine.working_fluid = combustor.working_fluid
compute_turbine_performance(high_pressure_turbine,conditions)
#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 = Data()
lpt_conditions.inputs.compressor.work_done = 0.0
lpt_conditions.inputs.compressor.external_shaft_work_done = 0.0
lpt_conditions.inputs.fuel_to_air_ratio = combustor_conditions.outputs.fuel_to_air_ratio
lpt_conditions.inputs.bypass_ratio = 0.0
lpt_conditions.inputs.fan = Data()
lpt_conditions.inputs.fan.work_done = 0.0
compute_turbine_performance(low_pressure_turbine,conditions)
#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
#flow through the core nozzle
compute_expansion_nozzle_performance(core_nozzle,conditions)
# compute the thrust using the thrust component
#link the thrust component to the core nozzle
turboshaft_conditions.core_exit_velocity = core_nozzle_conditions.outputs.velocity
turboshaft_conditions.core_area_ratio = core_nozzle_conditions.outputs.area_ratio
turboshaft_conditions.core_nozzle = core_nozzle_conditions.outputs
#link the thrust component to the combustor
turboshaft_conditions.fuel_to_air_ratio = combustor_conditions.outputs.fuel_to_air_ratio
#link the thrust component to the low pressure compressor
turboshaft_conditions.combustor_stagnation_temperature = combustor_conditions.outputs.stagnation_temperature
turboshaft_conditions.stag_temp_lpt_exit = compressor_conditions.inputs.stagnation_temperature
turboshaft_conditions.stag_press_lpt_exit = compressor_conditions.inputs.stagnation_pressure
turboshaft_conditions.total_temperature_reference = compressor_conditions.inputs.stagnation_temperature
turboshaft_conditions.total_pressure_reference = compressor_conditions.inputs.stagnation_pressure
#compute the power
turboshaft_conditions.fan_nozzle = Data()
turboshaft_conditions.fan_nozzle.velocity = 0.0
turboshaft_conditions.fan_nozzle.area_ratio = 0.0
turboshaft_conditions.fan_nozzle.static_pressure = 0.0
turboshaft_conditions.bypass_ratio = 0.0
turboshaft_conditions.flow_through_core = 1.0 #scaled constant to turn on core power computation
turboshaft_conditions.flow_through_fan = 0.0 #scaled constant to turn on fan power computation
# Step 25: Size the core of the turboshaft
size_core(turboshaft,conditions)
# Step 26: Static Sea Level Thrust
atmosphere = RCAIDE.Framework.Analyses.Atmospheric.US_Standard_1976()
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(turboshaft, altitude = 0,velocity_range=np.array([V]))
operating_state.conditions.energy.converters[turboshaft.tag].throttle[:,0] = 1.0
sls_P,_,_ = turboshaft.compute_performance(operating_state,fuel_line)
turboshaft.sealevel_static_power = sls_P[0][0]
return
