Hybrid Propulsion

Background:

The Hybrid Rocket Engine project began in 2013 to explore complex rocketry systems. In 2018, the project was revamped and an initiative to build our current engine began. A hybrid rocket engine utilizes rocket propellants in two different phases, one solid and the other tends to be a gas; in our case, we are using a nitrous oxide/oxygen blend (Nytrox) and 40% Al-Paraffin. Our 4kN hybrid rocket engine is entering the final stages of testing before a full-scale hot fire can be attempted.

Leadership

  • Aakash Shah

    Hybrid Propulsion Lead

    aakash4@illinois.edu

  • Nathan Zhu

    Electronics Lead

    nyzhu2@illinois.edu

  • Drew Eimer

    Manufacturing Co-Lead

    deeimer2@illinois.edu

  • Maddie Conrad

    Manufacturing Co-Lead

    mconrad5@illinois.edu

  • Zana Serbest

    Safety and Logistics Lead

    zserb2@illinois.edu

  • Dillon W Mulrooney

    Engine Testing Lead

    dillionm6@illinois.edu

Skills YOU will learn:

As a member of the Hybrid Propulsion team, you’ll be able to develop and acquire a lot of different workplace skills. This past year, new and current members were able to learn a lot about:

  1. Engine Testing:

    1. Designing and building a brand new test stand

    2. Creating piping and instrumentation diagrams (P&ID’s)

    3. Drafting and finalizing testing procedures

    4. Repairing electronic systems

  2. Manufacturing:

    1. Assembling full-scale manufacturing set-ups

    2. Rethinking old igniter and fuel grain manufacturing methods

    3. Combining highly-purified grades of nitrous oxide and gaseous oxide to create contaminant/catalytic-free Nytrox mixtures.

    4. 3D printing

System Architecture:

The hybrid rocket engine consists of the following components:

  • Oxidizer Tank

    • Contains Nytrox-96 at 45 bar, 0℃

      1. Flat end-caps

        1. Low weight penalty

        2. Cost-effective

      2. Successfully pressure tested to 69 bar

  • Combustion Chamber

    • Contains the fuel grain, nozzle, thermal protections, and pyrotechnic ignition system.

    • Excellent specific tensile strength

    • Body tube & end caps were successfully pressure tested to 45 bar

  • Feed System

    • Supplies oxidizer to the chamber to sustain the combustion

    • Self-pressurized

    • Successfully verified a required flow rate of 1.543 kg/s

  • Nozzle

    • Monolithic graphite nozzle

    • Provides effective erosion minimization

    • Computation fluid dynamics (CFD) analysis to verify design

  • Fuel Grain

    • Solid rocket propellant (40% Al-Paraffin)

    • Using a spin-casting mechanism to manufacture

  • Injector

    • Atomizes and disperes the oxidizer into the combustion chamber

    • Showerhead injector with straight, sharp-edged orifices

    • Atomization will be verified with cold-flow testing

  • Thermal Protection System

    • Keeps the combustion chamber structure below 100℃ during burn time

    • Canvas-phenolic composite

    • Two layers of protection