Thrust Chamber Design and Cooling
Background
A liquid-fueled rocket depends on many different factors to perform efficiently. Since the rocket depends on its propulsion system to achieve its goal, ensuring the efficiency of all the components is paramount to a successful rocket. Any loss of efficiencies in these components can be detrimental to the overall performance of the rocket. The goal of this project is to design a propulsion system consisting of a combustion chamber, injector, nozzle, and efficient cooling system that is able to deliver 2000 pounds of thrust and a specific impulse of 299 seconds using liquid oxygen as the oxidizer and liquid methane as the fuel. The thrust chamber should also be able to withstand a chamber pressure of 500 psi. Engine parameters such as Oxidizer:Fuel Ratio, exit pressure, exit Mach number, throat area, propellant mass flow rate, etc. must be analytically found.
Objectives
- Conceptualize and design an injector that can thoroughly mix the fuel and oxidizer to ensure stable combustion while also being relatively simple to manufacture.
- Determine combustion chamber dimensions to achieve given chamber pressure and guarantee complete combustion of the propellants.
- Calculate nozzle geometries to achieve the given thrust and exit pressure.
- Design a cooling system that is capable of keeping the combustion chamber and nozzle within safe operating temperatures.
- Select materials for each component that are capable of handling the high pressures and temperatures while also remaining lightweight and cost-effective.
Team Contacts
Anika Hussain (anikaih@uci.edu)
Francisco Menjivar (fmenjiva@uci.edu)
Nareg Haladjian (haladjin@uci.edu)
Rigoberto Collazo-Rocha (collazor@uci.edu)
Yuchen Song (yuchs10@uci.edu)
Project Sponsor
