A rocket needs to deliver 5000 N thrust with a chamber pressure of 2.039 MPa, chamber temperature of 2800 K. The specific heat ratio is 1.2 and gas constant is 360 J/kg-K. The atmospheric pressure is 2.549 kPa. Assuming perfectly expanded, determine the areas of the throat and nozzle exit. Hint: you may need to find other flow/thermodynamic quantities at each location.