Source code for RCAIDE.Library.Methods.Powertrain.Propulsors.Turboprop.design_turboprop
# RCAIDE/Library/Methods/Energy/Powertrain/Propulsors/Turboprop/design_turboprop.py
#
#
# Created: Jul 2023, M. Clarke
# ----------------------------------------------------------------------------------------------------------------------
# IMPORT
# ----------------------------------------------------------------------------------------------------------------------
# RCAIDE Imports
import RCAIDE
from RCAIDE.Framework.Core import Data
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.Propulsors.Turboprop import size_core
from RCAIDE.Library.Methods.Powertrain import setup_operating_conditions
from RCAIDE.Library.Methods.Powertrain.Converters.Motor import design_optimal_motor
from RCAIDE.Library.Methods.Mass_Properties.Weight_Buildups.Electric.Common import compute_motor_weight
# Python package imports
import numpy as np
# ----------------------------------------------------------------------------------------------------------------------
# Design Turboshaft
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[docs]
def design_turboprop(turboprop):
"""
Sizes a turboprop engine based on design point conditions and computes its performance characteristics.
Parameters
----------
turboprop : Turboprop
Turboprop engine object containing all component definitions and design parameters
- 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 the turboprop object attributes:
- design_thrust_specific_fuel_consumption : float
TSFC at design point [kg/N/s]
- design_non_dimensional_thrust : float
Non-dimensional thrust at design point [-]
- design_core_mass_flow_rate : float
Core mass flow rate at design point [kg/s]
- design_fuel_flow_rate : float
Fuel flow rate at design point [kg/s]
- design_power : float
Power output at design point [W]
- design_specific_power : float
Specific power at design point [W/kg]
- design_power_specific_fuel_consumption : float
Power specific fuel consumption [kg/W/s]
- design_thermal_efficiency : float
Thermal efficiency at design point [-]
- design_propulsive_efficiency : float
Propulsive efficiency at design point [-]
Notes
-----
The function performs the following steps:
1. Computes atmospheric conditions at design altitude
2. Sets up freestream conditions
3. Analyzes flow through each component sequentially
4. Sizes the core based on design point requirements
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
**Theory**
The design process follows standard gas turbine cycle analysis, with each
component modeled using appropriate thermodynamic relations. The core sizing
is based on achieving the required power output while maintaining component
matching throughout the engine.
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 turboprop.design_altitude==None:
if turboprop.design_mach_number==None and turboprop.design_freestream_velocity ==None:
raise NameError('The sizing conditions require an altitude and a Mach number or Velocity ')
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(turboprop.design_altitude,turboprop.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
if turboprop.design_mach_number==None:
turboprop.design_mach_number = turboprop.design_freestream_velocity / a
# setup conditions
conditions = RCAIDE.Framework.Mission.Common.Results()
# freestream conditions
conditions.freestream.altitude = np.atleast_1d(turboprop.design_altitude)
conditions.freestream.mach_number = np.atleast_1d(turboprop.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(turboprop.design_altitude))
conditions.freestream.isentropic_expansion_factor = np.atleast_1d(turboprop.working_fluid.compute_gamma(T,p))
conditions.freestream.Cp = np.atleast_1d(turboprop.working_fluid.compute_cp(T,p))
conditions.freestream.R = np.atleast_1d(turboprop.working_fluid.gas_specific_constant)
conditions.freestream.speed_of_sound = np.atleast_1d(a)
conditions.freestream.velocity = np.atleast_1d(a*turboprop.design_mach_number)
segment = RCAIDE.Framework.Mission.Segments.Segment()
segment.state.conditions = conditions
turboprop.append_operating_conditions(segment,conditions.energy,conditions.noise)
ram = turboprop.ram
inlet_nozzle = turboprop.inlet_nozzle
compressor = turboprop.compressor
combustor = turboprop.combustor
high_pressure_turbine = turboprop.high_pressure_turbine
low_pressure_turbine = turboprop.low_pressure_turbine
core_nozzle = turboprop.core_nozzle
# unpack component conditions
turboprop_conditions = conditions.energy.propulsors[turboprop.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 = turboprop.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.nondimensional_massflow = turboprop.compressor_nondimensional_massflow
compressor.reference_temperature = turboprop.reference_temperature
compressor.reference_pressure = turboprop.reference_pressure
# Step 6: Compute flow through the low pressure compressor
compute_compressor_performance(compressor,conditions)
# Step 7: 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
# Step 8: 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
high_pressure_turbine.working_fluid = combustor.working_fluid
hpt_conditions.inputs.bypass_ratio = 0.0
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
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.bypass_ratio = 0.0
lpt_conditions.inputs.fuel_to_air_ratio = combustor_conditions.outputs.fuel_to_air_ratio
low_pressure_turbine.working_fluid = high_pressure_turbine.working_fluid
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
turboprop_conditions.total_temperature_reference = compressor_conditions.inputs.stagnation_temperature
turboprop_conditions.total_pressure_reference = compressor_conditions.inputs.stagnation_pressure
# Step 25: Size the core of the turboprop
size_core(turboprop,conditions)
# Step 26: 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(turboprop, altitude = 0,velocity_range=np.array([V]))
operating_state.conditions.energy.propulsors[turboprop.tag].throttle[:,0] = 1.0
sls_T,_,sls_P,_,_,_ = turboprop.compute_performance(operating_state)
turboprop.sealevel_static_thrust = sls_T[0][0]
turboprop.sealevel_static_power = sls_P[0][0]
turboprop.design_thrust_specific_fuel_consumption = turboprop_conditions.thrust_specific_fuel_consumption
turboprop.design_non_dimensional_thrust = turboprop_conditions.non_dimensional_thrust
turboprop.design_core_mass_flow_rate = turboprop_conditions.core_mass_flow_rate
turboprop.design_fuel_flow_rate = turboprop_conditions.fuel_flow_rate
turboprop.design_specific_power = turboprop_conditions.specific_power
turboprop.design_power_specific_fuel_consumption = turboprop_conditions.power_specific_fuel_consumption
turboprop.design_thermal_efficiency = turboprop_conditions.thermal_efficiency
turboprop.design_propulsive_efficiency = turboprop_conditions.propulsive_efficiency
if compressor.motor != None:
V = turboprop.design_freestream_velocity
operating_state = setup_operating_conditions(turboprop, altitude = turboprop.design_altitude,velocity_range=np.array([V]))
operating_state.conditions.energy.propulsors[turboprop.tag].throttle[:,0] = 1.0
T,_,P,_,_,_ = turboprop.compute_performance(operating_state)
motor = compressor.motor
motor.design_torque = P[0][0] /compressor.design_angular_velocity
motor.design_angular_velocity = compressor.design_angular_velocity
motor.mass_properties.mass = compute_motor_weight(motor)
design_optimal_motor(motor)
return
