Propulsion and Motor Systems
STATUS: Currently constructing 1/4-scale test motor, static testing to be completed soon.
The Propulsion division of Project: Daedalus is responsible for designing, developing, and testing the hybrid motor that will be used for Daedalus. In order to develop the engine systems, the propulsion team will initially create a ¼ scale flight motor to test various aspects of the engine design, followed by a ½ scale model. After the testing motors are optimized the first flight motor, to be used on Daedalus III, will begin development, followed by the full scale Daedalus IV motor.
Hybrid Engine
A hybrid engine is created by combining a liquid oxidizer source with a solid fuel inside of a combustion chamber. The reaction between the oxidizer is controlled by valves that are placed between the oxidizer tank and the combustion chamber, which contains the fuel grain. In small hybrid engines, aluminum or steel is used to contain both the pressurized oxidizer and the combustion chamber, but for large scale engines, composite materials are used to lighten the weight of the engine. At the end of the combustion chamber is the nozzle, which is used to regulate the amount of pressure that is released from the combustion chamber and choke the flow of gas and then expand it to supersonic speeds.
Basic Design
The Oxidizer tank will be constructed from interlays of composite materials in a similar manner to the construction of industrial gas tanks. Since industrial gas tanks already have a reputation of reliability, a minimal amount of testing will be required before integrating the tanks into the engine systems. The engine systems required inside of the tank will include an anti-swash baffle system, which dampens the effects of the liquid oxidizer moving up and down as a result of jolt. The tank itself will be supported by a structural system that is integrated into the rocket body and will have 2 valves, one downward facing valve that meters oxidizer into the combustion chamber, and an unrestricted blowout valve that will enable the tank to be purged of oxidizer in the case of an emergency, malfunction, or misfire.
FEM Analysis of bottom bulkhead for custom N2O Tank
The Combustion chamber is located underneath the oxidizer tank and is connected to the lower valve on the oxidizer tank. The chamber contains the solid fuel for the hybrid, and is designed to burn the plastic fuel grain (made of HDPE) at 600 psi so that the oxidizer tank, which is pressurized by N2O at 750 psi, has a pressure difference of 150 psi with which to move the oxidizer (N2O) into the combustion chamber.
Propulsion Sub-Divisions
The Propulsion Team is compromised of three sub-divisions:
- Systems Integration and Modeling
- Oxidizer Tank/Valves/Injector
- Combustion Chamber/Nozzle
The Integration and Modeling team is responsible for all the CAD models of the Engine, and is also responsible for integrating the engine assembly into the motor casing. The Oxidizer tank, valves and injector team will be responsible for the design and construction of the oxidizer tank as well as the slosh reduction system. The oxidizer tank team is also responsible for the construction and wiring of the valve systems on the tank, as well as the integration of the emergency dump valve into the outer rocket casing. The Combustion Chamber team is responsible for the construction of the burn chamber, as well as design of all of the thermal insulation of the fuel grain and motor casing. The secondary responsibility of the chamber team is to manufacture the solid fuel grains for all of the rocket motors using extrusion or coring, or to alternatively acquire the fuel grain from an industrial supplier. They are also responsible for the design of the nozzle and for the optimization of compression and expansion throughout the combustion process. The final system the Chamber team is responsible for is the Pyrogen starting system that will ignite the oxidizer fuel mixture in the combustion chamber.
Team Members and Contact Info
Nick Spinler, Division Lead
Aerospace Engineering
nick.spinler@gmail.com
Systems Integration and Modeling
Joe Coverston, Aerospace Engineering (Sub-Division Lead)
Moise D'haiti, Aerospace Engineering
Juan Barredo, Aerospace Engineering
Jeff Cunningham, Aerospace Engineering
Oxidizer Tank/Valves/Injector
Brian Zollman, ET Space Systems (Sub-Division Lead)
Kevin Schillo, Aerospace Engineering
Doug Churaman, Aerospace Engineering
Kristyn Englar, Physics
Ekaterina Gatons, Aerospace Engineering
Daniel Jakob, Aerospace Engineering
Matt Hardin, Physics
Combustion Chamber/Nozzle
Richard Bertrand, Aerospace Engineering (Sub-Division Lead)
Matt Cole, Aerospace Engineering
Tyler Maddox, Aerospace Engineering
Travis Riley, Aerospace Engineering
Jaclyn Hislop, Aerospace Engineering
Johnny Brooks, Aerospace Engineering
Matt Leili, Aerospace Engineering