RCAIDE.Library.Methods.Powertrain.Converters.Combustor.compute_combustor_performance

compute_combustor_performance#

compute_combustor_performance(combustor, conditions)[source]#

Computes the thermodynamic performance of a combustor in a gas turbine engine.

Parameters:

  • combustor (RCAIDE.Library.Components.Converters.Combustor) –

    Combustor component with the following attributes:
    • tagstr

      Identifier for the combustor

    • working_fluidData

      Working fluid properties object

    • turbine_inlet_temperaturefloat

      Target turbine inlet temperature [K]

    • pressure_ratiofloat

      Pressure ratio across the combustor (typically < 1.0 due to losses)

    • efficiencyfloat

      Combustion efficiency

    • area_ratiofloat

      Exit to inlet area ratio

    • fuel_dataData
      Fuel properties
      • specific_energyfloat

        Fuel specific energy [J/kg]

  • conditions (RCAIDE.Framework.Mission.Common.Conditions) –

    Flight conditions with:
    • energyData
      Energy conditions
      • convertersdict

        Converter energy conditions indexed by tag

Returns:

Results are stored in conditions.energy.converters[combustor.tag].outputs:
  • stagnation_temperaturenumpy.ndarray

    Stagnation temperature at combustor exit [K]

  • stagnation_pressurenumpy.ndarray

    Stagnation pressure at combustor exit [Pa]

  • stagnation_enthalpynumpy.ndarray

    Stagnation enthalpy at combustor exit [J/kg]

  • fuel_to_air_rationumpy.ndarray

    Fuel-to-air ratio

  • static_temperaturenumpy.ndarray

    Static temperature at combustor exit [K]

  • static_pressurenumpy.ndarray

    Static pressure at combustor exit [Pa]

  • mach_numbernumpy.ndarray

    Mach number at combustor exit

Return type:

None

Notes

This function computes the thermodynamic properties at the combustor exit based on the inlet conditions, combustor characteristics, and fuel properties. It calculates the fuel-to-air ratio required to achieve the specified turbine inlet temperature, accounting for combustion efficiency and pressure losses.

The computation follows these steps:

  1. Extract inlet conditions (temperature, pressure, Mach number)

  2. Compute working fluid properties (gamma, Cp)

  3. Calculate stagnation pressure at exit using pressure ratio

  4. Set exit stagnation temperature to the specified turbine inlet temperature

  5. Compute stagnation enthalpies at inlet and exit

  6. Calculate fuel-to-air ratio required to achieve the temperature rise

  7. Compute exit static conditions (temperature, pressure) based on exit Mach number

  8. Store all results in the conditions data structure

Major Assumptions

  • Constant efficiency and pressure ratio

  • Turbine inlet temperature is controlled to a specified value

  • Mach number is preserved from inlet to exit

References

[1] Cantwell, B., “AA283 Course Notes”, Stanford University https://web.stanford.edu/~cantwell/AA283_Course_Material/

See also

RCAIDE.Library.Methods.Powertrain.Converters.Turbine.compute_turbine_performance