Source code for RCAIDE.Library.Methods.Powertrain.Propulsors.Turboprop.compute_turboprop_performance
# RCAIDE/Methods/Energy/Propulsors/Networks/Turboprop/compute_turboprop_performance.py
#
#
# Created: Jul 2023, M. Clarke
# ----------------------------------------------------------------------------------------------------------------------
# IMPORT
# ----------------------------------------------------------------------------------------------------------------------
# RCAIDE imports
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 compute_thrust
# python imports
from copy import deepcopy
import numpy as np
# ----------------------------------------------------------------------------------------------------------------------
# compute_turboprop_performance
# ----------------------------------------------------------------------------------------------------------------------
[docs]
def compute_turboprop_performance(turboprop, state, center_of_gravity=[[0.0, 0.0, 0.0]]):
"""
Computes the performance of a turboprop engine by analyzing the thermodynamic cycle.
Parameters
----------
turboprop : RCAIDE.Library.Components.Propulsors.Turboprop
Turboprop engine component with the following attributes:
- tag : str
Identifier for the turboprop
- working_fluid : Data
Working fluid properties object
- ram : Data
Ram component
- tag : str
Identifier for the ram
- inlet_nozzle : Data
Inlet nozzle component
- tag : str
Identifier for the inlet nozzle
- compressor : Data
Compressor component
- tag : str
Identifier for the compressor
- motor : Data, optional
Electric motor component
- generator : Data, optional
Electric generator component
- design_angular_velocity : float
Design angular velocity [rad/s]
- combustor : Data
Combustor component
- tag : str
Identifier for the combustor
- fuel_data : Data
Fuel properties
- specific_energy : float
Fuel specific energy [J/kg]
- high_pressure_turbine : Data
High pressure turbine component
- tag : str
Identifier for the high pressure turbine
- low_pressure_turbine : Data
Low pressure turbine component
- tag : str
Identifier for the low pressure turbine
- core_nozzle : Data
Core nozzle component
- tag : str
Identifier for the core nozzle
- reference_temperature : float
Reference temperature for mass flow scaling [K]
- reference_pressure : float
Reference pressure for mass flow scaling [Pa]
- compressor_nondimensional_massflow : float
Non-dimensional mass flow parameter [kg·√K/(s·Pa)]
- origin : list of lists
Origin coordinates [[x, y, z]] [m]
state : RCAIDE.Framework.Mission.Common.State
State object containing:
- conditions : Data
Flight conditions
- freestream : Data
Freestream properties
- velocity : numpy.ndarray
Freestream velocity [m/s]
- temperature : numpy.ndarray
Freestream temperature [K]
- pressure : numpy.ndarray
Freestream pressure [Pa]
- noise : Data
Noise conditions
- propulsors : dict
Propulsor noise conditions indexed by tag
- energy : Data
Energy conditions
- propulsors : dict
Propulsor energy conditions indexed by tag
- converters : dict
Converter energy conditions indexed by tag
- hybrid_power_split_ratio : float
Ratio of power split for hybrid systems
- numerics : Data
Numerical properties
- time : Data
Time properties
- differentiate : list
List of differentiation methods
center_of_gravity : list of lists, optional
Center of gravity coordinates [[x, y, z]] [m]
Default: [[0.0, 0.0, 0.0]]
Returns
-------
thrust_vector : numpy.ndarray
Thrust force vector [N]
moment : numpy.ndarray
Moment vector [N·m]
power : numpy.ndarray
Shaft power output [W]
power_elec : numpy.ndarray
Electrical power input/output [W]
stored_results_flag : bool
Flag indicating if results are stored
stored_propulsor_tag : str
Tag of the turboprop with stored results
Notes
-----
This function computes the performance of a turboprop engine by sequentially analyzing
each component in the engine's thermodynamic cycle. It links the output conditions of
each component to the input conditions of the next component in the flow path.
The function follows this sequence:
1. Set working fluid properties
2. Compute ram performance
3. Compute inlet nozzle performance
4. Compute compressor performance
5. Compute combustor performance
6. Compute high pressure turbine performance
7. Compute low pressure turbine performance
8. Compute core nozzle performance
9. Compute thrust and power output
10. Calculate efficiencies
11. Handle electrical power generation/consumption if applicable
**Major Assumptions**
* Steady state operation
* One-dimensional flow through components
* Adiabatic components except for the combustor
* Perfect gas behavior with variable properties
References
----------
[1] Mattingly, J.D., "Elements of Gas Turbine Propulsion", 2nd Edition, AIAA Education Series, 2005. https://soaneemrana.org/onewebmedia/ELEMENTS%20OF%20GAS%20TURBINE%20PROPULTION2.pdf
See Also
--------
RCAIDE.Library.Methods.Powertrain.Propulsors.Turboprop.compute_thrust
"""
conditions = state.conditions
noise_conditions = conditions.noise.propulsors[turboprop.tag]
turboprop_conditions = conditions.energy.propulsors[turboprop.tag]
U0 = conditions.freestream.velocity
T = conditions.freestream.temperature
P = conditions.freestream.pressure
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
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_conditions.reference_temperature = turboprop.reference_temperature
compressor_conditions.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
# Compute the power
compute_thrust(turboprop,conditions)
# Compute forces and moments
moment_vector = 0*state.ones_row(3)
thrust_vector = 0*state.ones_row(3)
thrust_vector[:,0] = turboprop_conditions.thrust[:,0]
moment_vector[:,0] = turboprop.origin[0][0] - center_of_gravity[0][0]
moment_vector[:,1] = turboprop.origin[0][1] - center_of_gravity[0][1]
moment_vector[:,2] = turboprop.origin[0][2] - center_of_gravity[0][2]
M = np.cross(moment_vector, thrust_vector)
moment = M
power = turboprop_conditions.power
# compute efficiencies
mdot_air_core = turboprop_conditions.core_mass_flow_rate
fuel_enthalpy = combustor.fuel_data.specific_energy
mdot_fuel = turboprop_conditions.fuel_flow_rate
h_e_c = core_nozzle_conditions.outputs.static_enthalpy
h_0 = turboprop.working_fluid.compute_cp(T,P) * T
h_t4 = combustor_conditions.outputs.stagnation_enthalpy
h_t3 = compressor_conditions.outputs.stagnation_enthalpy
turboprop_conditions.overall_efficiency = thrust_vector* U0 / (mdot_fuel * fuel_enthalpy)
turboprop_conditions.thermal_efficiency = 1 - ((mdot_air_core + mdot_fuel)*(h_e_c - h_0) + mdot_fuel *h_0)/((mdot_air_core + mdot_fuel)*h_t4 - mdot_air_core *h_t3)
compressor_conditions.omega = compressor.design_angular_velocity * turboprop_conditions.throttle
# compute electrical power if generated/supplied
power_elec = 0*state.ones_row(1)
if compressor.motor != None and len(state.numerics.time.differentiate) > 0:
compressor_motor_conditions = conditions.energy.converters[compressor.motor.tag]
compressor_motor_conditions.outputs.power = power *conditions.energy.hybrid_power_split_ratio
compressor_motor_conditions.outputs.omega = compressor_conditions.omega
compressor_motor_conditions.outputs.torque = compressor_motor_conditions.outputs.power / compressor_motor_conditions.outputs.omega
power_elec = compressor_motor_conditions.outputs.power
if compressor.generator != None and len(state.numerics.time.differentiate) > 0:
compressor_generator_conditions = conditions.energy.converters[compressor.generator.tag]
compressor_generator_conditions.inputs.power = power *conditions.energy.hybrid_power_split_ratio
compressor_generator_conditions.inputs.omega = compressor_conditions.omega
compressor_generator_conditions.outputs.torque = compressor_generator_conditions.outputs.power / compressor_generator_conditions.outputs.omega
power_elec = compressor_generator_conditions.inputs.power
# Store data
core_nozzle_res = Data(
exit_static_temperature = core_nozzle_conditions.outputs.static_temperature,
exit_static_pressure = core_nozzle_conditions.outputs.static_pressure,
exit_stagnation_temperature = core_nozzle_conditions.outputs.stagnation_temperature,
exit_stagnation_pressure = core_nozzle_conditions.outputs.static_pressure,
exit_velocity = core_nozzle_conditions.outputs.velocity
)
noise_conditions.core_nozzle = core_nozzle_res
# Pack results
stored_results_flag = True
stored_propulsor_tag = turboprop.tag
return thrust_vector,moment,power,power_elec,stored_results_flag,stored_propulsor_tag
[docs]
def reuse_stored_turboprop_data(turboprop,state,network,stored_propulsor_tag,center_of_gravity= [[0.0, 0.0,0.0]]):
'''Reuses results from one turboprop for identical propulsors
Assumptions:
N/A
Source:
N/A
Inputs:
conditions - operating conditions data structure [-]
fuel_line - fuelline [-]
turboprop - turboprop data structure [-]
total_power - power of turboprop group [W]
Outputs:
total_power - power of turboprop group [W]
Properties Used:
N.A.
'''
# unpack
conditions = state.conditions
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
ram_0 = network.propulsors[stored_propulsor_tag].ram
inlet_nozzle_0 = network.propulsors[stored_propulsor_tag].inlet_nozzle
compressor_0 = network.propulsors[stored_propulsor_tag].compressor
combustor_0 = network.propulsors[stored_propulsor_tag].combustor
high_pressure_turbine_0 = network.propulsors[stored_propulsor_tag].high_pressure_turbine
low_pressure_turbine_0 = network.propulsors[stored_propulsor_tag].low_pressure_turbine
core_nozzle_0 = network.propulsors[stored_propulsor_tag].core_nozzle
# deep copy results
conditions.energy.propulsors[turboprop.tag] = deepcopy(conditions.energy.propulsors[stored_propulsor_tag])
conditions.noise.propulsors[turboprop.tag] = deepcopy(conditions.noise.propulsors[stored_propulsor_tag])
conditions.energy.converters[ram.tag] = deepcopy(conditions.energy.converters[ram_0.tag] )
conditions.energy.converters[inlet_nozzle.tag] = deepcopy(conditions.energy.converters[inlet_nozzle_0.tag] )
conditions.energy.converters[compressor.tag] = deepcopy(conditions.energy.converters[compressor_0.tag] )
conditions.energy.converters[combustor.tag] = deepcopy(conditions.energy.converters[combustor_0.tag] )
conditions.energy.converters[low_pressure_turbine.tag] = deepcopy(conditions.energy.converters[low_pressure_turbine_0.tag] )
conditions.energy.converters[high_pressure_turbine.tag] = deepcopy(conditions.energy.converters[high_pressure_turbine_0.tag] )
conditions.energy.converters[core_nozzle.tag] = deepcopy(conditions.energy.converters[core_nozzle_0.tag] )
# compute moment
moment_vector = 0*state.ones_row(3)
thrust_vector = 0*state.ones_row(3)
thrust_vector[:,0] = conditions.energy.propulsors[turboprop.tag].thrust[:,0]
moment_vector[:,0] = turboprop.origin[0][0] - center_of_gravity[0][0]
moment_vector[:,1] = turboprop.origin[0][1] - center_of_gravity[0][1]
moment_vector[:,2] = turboprop.origin[0][2] - center_of_gravity[0][2]
moment = np.cross(moment_vector,thrust_vector)
power = conditions.energy.propulsors[turboprop.tag].power
conditions.energy.propulsors[turboprop.tag].moment = moment
power_elec = 0*state.ones_row(1)
if compressor.motor != None and len(state.numerics.time.differentiate) > 0:
conditions.energy.converters[compressor.motor.tag] = deepcopy(conditions.energy.converters[compressor_0.motor.tag])
power_elec = conditions.energy.converters[compressor.motor.tag].outputs.power
if compressor.generator != None and len(state.numerics.time.differentiate) > 0:
conditions.energy.converters[compressor.generator.tag] = deepcopy(conditions.energy.converters[compressor_0.generator.tag])
power_elec = conditions.energy.converters[compressor.generator.tag].inputs.power
return thrust_vector,moment,power, power_elec