Turbofan calculator. The formulae correspond to P.G.Hill and C.R. Peterson, "Mechanics and Thermodynamics of Propulsion", Addison-Wesley, 1970. The sketch explains the notation.  

Diffuser input:
Specific heat ratio γ= (for air 1.4)
Diffuser adiabatic efficiency η_d=
Flight Mach number M =

   


 
Diffuser output:
Stagnation temperature/ambient temperature T_0d/T_a=
Stagnation pressure/ambient pressure p_0d/p_a=

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Compressor input:
Compressor pressure ratio p_rc=
Compressor specific heat ratio γ_c=
Compressor adiabatic efficiency η_c=

 
Compressor output:
Stagnation temperature/ambient temperature T_0c/T_aa =
Stagnation pressure/ambient pressure p_0c/p_a=

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Fan input:
Fan pressure ratio p_rf=
Fan specific heat ratio γ_f=
Fan adiabatic efficiency η_f=
Fan nozzle adiabatic efficiency η_fn=
Gas constant (fan nozzle) R_f= in J/(Kg K). (For dry air it is around 287.)
Ambient temperature T_a= in grad K




Fan output:
Stagnation temperature/ambient temperature T_0f/T_a=
Stagnation pressure/ambient pressure p_0f/p_a=
Fan nozzle exhaust velocity u_ef=m/s

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Burner input:
Burner pressure ratio r_b=
Stagnation temperature T_0b= in grad K
Effective heat of reaction η_b Q_R= in kJ/kg (for typical aviation kerosene Q_R=45000 kJ/kg)
Specific heat ratio of the combustion products c_p= in J/(kg K) (for dry air it is around 1005)




Burner output:
Fuel-to-air ratio f=
Stagnation temperature/ambient temperature T_0b/T_a=
Stagnation pressure/ambient pressure p_0b/p_a=

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Turbine input:
Specific heat ratio (turbine) γ_t=
Turbine adiabatic efficiency η_t=
Bypass ratio β=





Turbine output:
Stagnation temperature/ambient temperature T_0t/T_a=
Stagnation pressure/ambient pressure p_0t/p_a=

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Nozzle input:
Specific heat ratio (nozzle) γ_n=
Nozzle adiabatic efficiency η_n=
Gas constant (nozzle) R_n= in J/(kg K)
Gas constant (air) R= in J/(kg K)





Nozzle output:
Stagnation pressure/ambient pressure p_0e/p_a=
Exhaust stagnation temperature T_0e= in grad K
Exhaust velocity u_e= m/s
Flight velocity u= m/s
Specific Thrust = kN/(kg/s)
TSFC= (kg/s)/kN
Propulsion efficiency η_p=
Thermal efficiency η_th=η_b x
Overall efficiency η_o=η_b x

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