1. The problem statement, all variables and given/known data A rocket operating with combustion chamber pressure and temperatures of 14MPa and 2500K respectively, has a throat diameter of 0.3m, and a nozzle area ratio of 50:1. Find the thrust and specific impulse developed by the motor with back pressures of 1 bar (10^5 Pa) and zero. At what back pressure would the nozzle be correctly expanded? Assume the combustion products behave as a perfect gas, with a constant specific heat ratio (y) of 1.4, and a constant specific gas constant (R) of 300 J/kg/K 2. Relevant equations throat P/chamber P = ((y+1)/2)^(y/(1-y)) (for M=1) M = Mach number throat T/chamber T = y/(y+1) (for M=1) mdot = [(area of throat * throat P)/sqrt(throat t)]*[sqrt(y/R)]*((y + 1)/2)*((y + 1)/2(1 - y)) area of throat = pi * throat radius * throat radius 3. The attempt at a solution chamber P = 14MPa chamber T = 2,500K throat diameter = 0.3m; throat diameter = 0.15m y = 1.4 (constant specific heat ratio) R = 300 J/kg/K Calculate mdot --> through conditions @ throat M=M*=1 throat P/chamber P = ((y+1)/2)^(y/(1-y)) throat P = 7.4MPa throat T/chamber T = y/(y+1) throat T = 2,083.3K mdot = [(area of throat * throat P)/sqrt(throat t)]*[sqrt(y/R)]*((y + 1)/2)*((y + 1)/2(1 - y)) mdot = 782.47 area of throat = pi * throat radius * throat radius area of throat = 0.0707 This is where I get stuck, I need to incorporate these back pressures of 10^5 Pa and 0 Pa somehow. I know that for a correctly expanded nozzle, pressure of exit = ambient pressure Any help would be greatly appreciated!