Source code for RCAIDE.Library.Methods.Powertrain.Converters.Turboshaft.compute_power
# RCAIDE/Library/Methods/Energy/Converters/Turboshaft/compute_power.py
#
#
# Created: Jul 2023, M. Clarke
# Modified: Jun 2024, M. Guidotti
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# IMPORT
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# Python package imports
import numpy as np
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# compute_power
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[docs]
def compute_power(turboshaft,conditions):
"""
Computes power and other performance properties for a turboshaft engine.
Parameters
----------
turboshaft : RCAIDE.Library.Components.Converters.Turboshaft
Turboshaft engine component with the following attributes:
- tag : str
Identifier for the turboshaft
- fuel_type : Data
Fuel properties object
- lower_heating_value : float
Fuel lower heating value [J/kg]
- working_fluid : Data
Working fluid properties object
- reference_temperature : float
Reference temperature for mass flow scaling [K]
- reference_pressure : float
Reference pressure for mass flow scaling [Pa]
- compressor : Data
Compressor component
- pressure_ratio : float
Compressor pressure ratio
- mass_flow_rate : float
Design mass flow rate [kg/s]
- conversion_efficiency : float
Efficiency of converting thermal energy to shaft power
- inverse_calculation : bool
Flag for inverse calculation mode (power to throttle)
conditions : RCAIDE.Framework.Mission.Common.Conditions
Flight conditions with:
- freestream : Data
Freestream properties
- isentropic_expansion_factor : numpy.ndarray
Ratio of specific heats (gamma)
- speed_of_sound : numpy.ndarray
Speed of sound [m/s]
- mach_number : numpy.ndarray
Flight Mach number
- gravity : numpy.ndarray
Gravitational acceleration [m/s²]
- energy.converters[turboshaft.tag] : Data
Turboshaft operating conditions
- throttle : numpy.ndarray
Throttle setting [0-1]
- total_temperature_reference : numpy.ndarray
Reference total temperature [K]
- total_pressure_reference : numpy.ndarray
Reference total pressure [Pa]
- combustor_stagnation_temperature : numpy.ndarray
Combustor exit stagnation temperature [K]
- power : numpy.ndarray
Required power output (for inverse calculation) [W]
Returns
-------
None
Notes
-----
This function implements a thermodynamic model for a turboshaft engine with a free power turbine.
It can operate in two modes: direct (throttle to power) or inverse (power to throttle).
**Major Assumptions**
* Perfect gas behavior
* Turboshaft engine with free power turbine
* Constant component efficiencies
**Theory**
The turboshaft performance is calculated using gas turbine cycle analysis. The power output
is determined by the temperature rise in the combustor, the compressor pressure ratio, and
the component efficiencies.
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
[2] Stuyvenberg, L., "Helicopter Turboshafts", University of Colorado, 2015. https://www.colorado.edu/faculty/kantha/sites/default/files/attached-files/70652-116619_-_luke_stuyvenberg_-_dec_17_2015_1258_pm_-_stuyvenberg_helicopterturboshafts.pdf
See Also
--------
RCAIDE.Library.Methods.Powertrain.Converters.Turbine.compute_turbine_performance
"""
#unpack the values
working_fluid = turboshaft.working_fluid
Tref = turboshaft.reference_temperature
Pref = turboshaft.reference_pressure
eta_c = turboshaft.conversion_efficiency
SFC_adjustment = turboshaft.specific_fuel_consumption_reduction_factor
pi_c = turboshaft.compressor.pressure_ratio
m_dot_compressor = turboshaft.compressor.mass_flow_rate
LHV = turboshaft.fuel_type.lower_heating_value
gamma = conditions.freestream.isentropic_expansion_factor
a0 = conditions.freestream.speed_of_sound
M0 = conditions.freestream.mach_number
turboshaft_conditions = conditions.energy.converters[turboshaft.tag]
Tt4 = turboshaft_conditions.combustor_stagnation_temperature
total_temperature_reference = turboshaft_conditions.total_temperature_reference
total_pressure_reference = turboshaft_conditions.total_pressure_reference
Power = turboshaft_conditions.power
Cp = working_fluid.compute_cp(total_temperature_reference,total_pressure_reference)
#unpacking from turboshaft
tau_lambda = Tt4/total_temperature_reference
tau_r = 1 + ((gamma - 1)/2)*M0**2
tau_c = pi_c**((gamma - 1)/gamma)
tau_t = (1/(tau_r*tau_c)) + ((gamma - 1)*M0**2)/(2*tau_lambda*eta_c**2)
tau_tH = 1 - (tau_r/tau_lambda)*(tau_c - 1)
tau_tL = tau_t/tau_tH
x = tau_t*tau_r*tau_c
# Computing Specifc Thrust and Power
Tsp = a0*(((2/(gamma - 1))*(tau_lambda/(tau_r*tau_c))*(tau_r*tau_c*tau_t - 1))**eta_c - M0)
Psp = Cp*total_temperature_reference*tau_lambda*tau_tH*(1 - tau_tL)*eta_c
if turboshaft.inverse_calculation == False:
m_dot_air = m_dot_compressor*turboshaft_conditions.throttle*np.sqrt(Tref/total_temperature_reference)*(total_pressure_reference/Pref)
Power = Psp*m_dot_air
else:
m_dot_air = Power / Psp
turboshaft_conditions.throttle = m_dot_air / (m_dot_compressor*np.sqrt(Tref/total_temperature_reference)*(total_pressure_reference/Pref) )
#fuel to air ratio
f = (Cp*total_temperature_reference/LHV)*(tau_lambda - tau_r*tau_c)
fuel_flow_rate = (1 - SFC_adjustment) *f*m_dot_air
#Computing the PSFC
PSFC = f/Psp
#Computing the thermal efficiency
eta_T = 1 - (tau_r*(tau_c - 1))/(tau_lambda*(1 - x/(tau_r*tau_c)))
#pack outputs
turboshaft_conditions.power_specific_fuel_consumption = PSFC
turboshaft_conditions.fuel_flow_rate = fuel_flow_rate
turboshaft_conditions.power = Power
turboshaft_conditions.non_dimensional_power = Psp
turboshaft_conditions.non_dimensional_thrust = Tsp
turboshaft_conditions.thermal_efficiency = eta_T
return
