Thesis Abstract
The main aim of this thesis work is to numerically investigate the cooling performance of the GCH4−LOX rocket engine for future human Mars missions. The thesis works broadly divided into three sections of studies. These sections are 1. Non-premixed Combustion of GCH4−LOX at Trans-critical and Supercritical pressure, 2. Regenerative cooling analysis of GCH4−LOX rocket engine, 3. Film cooling analysis of GCH4−LOX rocket engine. In the combustion Study, a steady-state chemically equilibrium flamelet approach is adapted to account for real-gas thermodynamics effects which are a prominent feature of flames at near-critical injection conditions. The thermodynamic model is based on the Peng Robinson equation of state in conjunction with a novel volume-translation method to correct deficiencies in the trans-critical regime. k−ω SST Turbulence model employed to calculate the turbulent fluctuating quantities near the solid wall boundaries for all the cases. Combustion results show that the maximum attainable flame temperature is 3525K. By looking at this result for all cooling analyses maximum temperature is set to 3500 K. Flame Temperature varied from 2000 K to 3500 K for all cooling analyses.
In regenerative cooling analysis heat flux, Nusselt number, Thrust Coefficient had been intuitively studied w.r.t Flame temperature, Coolant Inlet Temperature, Coolant Velocity, Coolant Pressure, Types of Coolant. The result shows that due to heavy regenerative cooling, the losses occur owing to combustion and heat had been utilized optimally which leads to an increase in the propulsive parameter of a rocket engine. In the film cooling analysis heat flux, film cooling effectiveness was studied w.r.t flame temperature, types of fuels, and NPR.
Both cooling shows promising results, but comparative analysis between regenerative cooling and film cooling shows that regenerative cooling is more efficient than film cooling. During regenerative cooling fuel absorbed the heat (combustion and heat loss) from the nozzle wall and its specific enthalpy increases which leads to an increase in propulsive performance of the rocket engine.