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+clc
+// Example 1.6.py
+// Consider the flow through a rocket engine nozzle. Assume that the gas flow 
+// through the nozzle in an isentropic expansion of a calorically perfect gas.
+// In the combustion chamber, the gas which results from the combustion of the 
+// rocket fuel and oxidizer is at a pressure and temperature of 15 atm and 
+// 2500 K, respectively, the molecular weight and specific heat at constant 
+// pressure of the combustion gas are 12 and 4157 J/kg K, respectively. The gas
+// expands to supersonic speed through the nozzle, with temperature of 1350 K at
+// the nozzle exit. Calculate the pressure at the exit.
+
+
+// Variable declaration
+pc = 15.0          // pressure combustion chamber (atm)
+Tc = 2500.0        // temperature combustion chamber (K)
+mol_wt = 12.0      // molecular weight (gm)
+cp = 4157.0        // specific heat at constant pressure (J/Kg/K)
+
+Tn  = 1350.0       // temperature at nozzle exit (K)
+
+// Calculations 
+R = 8314.0/mol_wt  // gas constant = R_prime/mo_wt, R_prime = 8314 J/K
+cv = cp - R        // specific heat at constant volume (J/Kg/K)
+gamma1 = cp/cv      // ratio of specific heat
+
+pn_by_pc = (Tn/Tc** gamma1/(gamma1-1)) // ratio of pressure for isentropic process** pn/pc
+
+pn = pc * pn_by_pc // pn = pc * pn/pc
+
+// Result 
+printf("\n Pressure at the exit is %.3f atm", pn)
+