SUAVE.Components.Energy.Converters.Combustor.Combustor Class Reference

Inheritance diagram for SUAVE.Components.Energy.Converters.Combustor.Combustor:

Public Member Functions
def	__defaults__ (self)
def	compute (self, conditions)
def	compute_rayleigh (self, conditions)
def	compute_supersonic_combustion (self, conditions)

Public Attributes
	tag
	fuel_data
	alphac
	turbine_inlet_temperature
	area_ratio
	axial_fuel_velocity_ratio
	fuel_velocity_ratio
	burner_drag_coefficient
	absolute_sensible_enthalpy
	fuel_equivalency_ratio
Public Attributes inherited from SUAVE.Components.Energy.Energy_Component.Energy_Component
	inputs
	outputs
Public Attributes inherited from SUAVE.Components.Physical_Component.Physical_Component
	tag
	mass_properties
	origin
	symmetric
Public Attributes inherited from SUAVE.Components.Component.Component
	tag
	origin
	generative_design_max_per_vehicle
	generative_design_characteristics
	generative_design_special_parent

Detailed Description

This provides output values for a combustor
Calling this class calls the compute function.

Assumptions:
None

Source:
https://web.stanford.edu/~cantwell/AA283_Course_Material/AA283_Course_Notes/

Member Function Documentation

__defaults__()

def SUAVE.Components.Energy.Converters.Combustor.Combustor.__defaults__

(

self

)

This sets the default values for the component to function.

Assumptions:
None

Source:
N/A

Inputs:
None

Outputs:
None

Properties Used:
None

Reimplemented from SUAVE.Components.Energy.Energy_Component.Energy_Component.

compute()

def SUAVE.Components.Energy.Converters.Combustor.Combustor.compute	(		self,
			conditions
	)

This computes the output values from the input values according to
equations from the source.

Assumptions:
Constant efficiency and pressure ratio

Source:
https://web.stanford.edu/~cantwell/AA283_Course_Material/AA283_Course_Notes/

Inputs:
conditions.freestream.
  isentropic_expansion_factor         [-]
  specific_heat_at_constant_pressure  [J/(kg K)]
  temperature                         [K]
  stagnation_temperature              [K]
self.inputs.
  stagnation_temperature              [K]
  stagnation_pressure                 [Pa]
  nondim_mass_ratio                   [-]

Outputs:
self.outputs.
  stagnation_temperature              [K]  
  stagnation_pressure                 [Pa]
  stagnation_enthalpy                 [J/kg]
  fuel_to_air_ratio                   [-]

Properties Used:
self.
  turbine_inlet_temperature           [K]
  pressure_ratio                      [-]
  efficiency                          [-]
  area_ratio                          [-]
  fuel_data.specific_energy           [J/kg]

compute_rayleigh()

def SUAVE.Components.Energy.Converters.Combustor.Combustor.compute_rayleigh	(		self,
			conditions
	)

This combutes the temperature and pressure change across the
the combustor using Rayleigh Line flow; it checks for themal choking.

Assumptions:
Constant efficiency and pressure ratio

Source:
https://web.stanford.edu/~cantwell/AA283_Course_Material/AA283_Course_Notes/

Inputs:
conditions.freestream.
  isentropic_expansion_factor         [-]
  specific_heat_at_constant_pressure  [J/(kg K)]
  temperature                         [K]
  stagnation_temperature              [K]
self.inputs.
  stagnation_temperature              [K]
  stagnation_pressure                 [Pa]

Outputs:
self.outputs.
  stagnation_temperature              [K]  
  stagnation_pressure                 [Pa]
  stagnation_enthalpy                 [J/kg]
  fuel_to_air_ratio                   [-]

Properties Used:
self.
  turbine_inlet_temperature           [K]
  pressure_ratio                      [-]
  efficiency                          [-]
  area_ratio                          [-]
  fuel_data.specific_energy           [J/kg]

compute_supersonic_combustion()

def SUAVE.Components.Energy.Converters.Combustor.Combustor.compute_supersonic_combustion	(		self,
			conditions
	)

This function computes the output values for supersonic  
combustion (Scramjet).  This will be done using stream thrust 
analysis. 

Assumptions: 
Constant Pressure Combustion      
Flow is in axial direction at all times 
Flow properities at exit are 1-Da averages 

Source: 
Heiser, William H., Pratt, D. T., Daley, D. H., and Unmeel, B. M., 
"Hypersonic Airbreathing Propulsion", 1994 
Chapter 4 - pgs. 175-180

Inputs: 
conditions.freestream. 
   isentropic_expansion_factor          [-] 
   specific_heat_at_constant_pressure   [J/(kg K)] 
   temperature                          [K] 
   stagnation_temperature               [K]
   universal_gas_constant               [J/(kg K)]  
self.inputs. 
   stagnation_temperature               [K] 
   stagnation_pressure                  [Pa] 
   inlet_nozzle                         [-] 
  
Outputs: 
self.outputs. 
   stagnation_temperature               [K] 
   stagnation_pressure                  [Pa] 
   stagnation_enthalpy                  [J/kg] 
   fuel_to_air_ratio                    [-] 
   static_temperature                   [K] 
   static_pressure                      [Pa] 
   velocity                             [m/s] 
   mach_number                          [-]          

       Properties Used: 
  self.fuel_data.specific_energy       [J/kg] 
  self.efficiency                      [-]
  self.axial_fuel_velocity_ratio       [-] 
  self.fuel_velocity_ratio             [-] 
  self.burner_drag_coefficient         [-] 
  self.temperature_reference           [K] 
  self.absolute_sensible_enthalpy      [J/kg] 
  self.specific_heat_constant_pressure [J/(kg K)] 

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The documentation for this class was generated from the following file:

• /Users/emiliobotero/Dropbox/SUAVE/SUAVE/trunk/SUAVE/Components/Energy/Converters/Combustor.py