From b1f5c3f8d6671b4331cef1dcebdf63b7a43a3a2b Mon Sep 17 00:00:00 2001 From: priyanka Date: Wed, 24 Jun 2015 15:03:17 +0530 Subject: initial commit / add all books --- 2223/CH18/EX18.13/Ex18_13.sav | Bin 0 -> 56928 bytes 2223/CH18/EX18.13/Ex18_13.sce | 55 ++++++++++++++++++++++++++++++++++++++++++ 2 files changed, 55 insertions(+) create mode 100755 2223/CH18/EX18.13/Ex18_13.sav create mode 100755 2223/CH18/EX18.13/Ex18_13.sce (limited to '2223/CH18/EX18.13') diff --git a/2223/CH18/EX18.13/Ex18_13.sav b/2223/CH18/EX18.13/Ex18_13.sav new file mode 100755 index 000000000..9641ce3e5 Binary files /dev/null and b/2223/CH18/EX18.13/Ex18_13.sav differ diff --git a/2223/CH18/EX18.13/Ex18_13.sce b/2223/CH18/EX18.13/Ex18_13.sce new file mode 100755 index 000000000..9169fc8bd --- /dev/null +++ b/2223/CH18/EX18.13/Ex18_13.sce @@ -0,0 +1,55 @@ +// scilab Code Exa 18.13 Turbojet Gas Turbine Engine + +T1=223.15; // in Kelvin +n_C=0.75; // Compressor Efficiency +c1=85; // entry velocity in m/s +m=50; // mass flow rate of air in kg/s +R=287; +n_B=0.98; // Combustion chamber Efficiency +Qf=43*1e3; // Calorific Value of fuel in kJ/kg; +T03=1220; // Turbine inlet stagnation temp in Kelvin +n_T=0.8; // Turbine Efficiency +gamma=1.4; // Specific Heat Ratio +n_m=0.98; // Mechanical efficiency +sigma=0.5; // flight to jet speed ratio(u/ce) +n_N=0.98; // exhaust nozzle efficiency +cp=1.005; // Specific Heat at Constant Pressure in kJ/(kgK) +u=886/3.6; // flight speed of a turbo prop aircraft in m/s +delT=200; // temperature rise through the compressor(T02-T01) in K +pi=.701; // Initial Pressure in bar +n_D=0.88; // inlet diffuser efficiency +a1=sqrt(gamma*R*T1); +M1=u/a1; // Mach number at the compressor inlet +T1_01=0.881; // (T1/T01)from isentropic flow gas tables at M1 and gamma values +T01=T1/T1_01; +T1=T01-(0.5*(c1^2)/(cp*1e3)); + +// part(a) compressor pressure ratio +T02s=T01+(delT*n_C); +r_oc=(T02s/T01)^(gamma/(gamma-1)); //compressor pressure ratio(p02/p01) +disp(r_oc,"(a)compressor pressure ratio is") + +// part(b) +T02=T01+delT; +f=((cp*T03)-(cp*T02))/((Qf*n_B)-(cp*T03)); // f=(ma/mf);energy balance in the combustion chamber +disp(1/f,"(b)Air-Fuel Ratio is") + +// part(c) turbine pressure ratio +// turbine power input P_T=n_m*(ma+mf)*cp*(T03-T01) +// power input to the compressor P_C=ma*cp*(T02-T01) +T04s=T03-(delT/(n_m*n_T*(1+f))); // from energy balance P_T=P_C +r_ot=(T03/T04s)^(gamma/(gamma-1)); //turbine pressure ratio(p03/p04) +disp(r_ot,"(c)turbine pressure ratio is") + +// part(d)exhaust nozzle pressure ratio +ce=u/sigma; // jet velocity at the exit of the exhaust nozzle +T04=T03-(delT/(n_m*(1+f))); +Te=T04-(0.5*(ce^2)/(cp*1e3)); +Tes=T04-((T04-Te)/n_N); +r_N=(T04/Tes)^(gamma/(gamma-1)); //exhaust nozzle pressure ratio(p04/pe) +disp(r_N,"(d)exhaust nozzle pressure ratio is") +ae=sqrt(gamma*R*Te); +Me=ce/ae; // Mach number +disp(Me,"and the Mach Number is") + + -- cgit