Project Northstar

Objectives

Project Name: Northstar – Variable impulse hybrid rocket engine

  1. Developed by: Saint Louis University Rocket Propulsion Lab (SLU RPL)

  2. Purpose: Sounding rocket propulsion system designed for research and IREC competition

  3. Motor Classification: O-4200 hybrid engine

  4. Fuel and Oxidizer:

    • Fuel: PMMA (acrylic)

    • Oxidizer: Liquid nitrous oxide

  5. Structural Design:

    • Combustion chamber: 6061 aluminum

    • Nozzle: Graphite

    • Thermal liner: Phenolic

    • Oxidizer tank: 6061 aluminum, rated for 700 psi operating pressure

  6. Physical Integration:

    • Fits standard 4-inch motor mount tubes

    • Compatible with legacy SLU airframes

  7. Performance:

    • Average thrust: 1000 lbf

    • Total impulse: 5000 lbf·s

  8. Control System:

    • Electronically adjustable oxidizer flow

    • Microcontroller-based real-time thrust modulation

    • Enables altitude control during flight

  9. Safety Features:

    • Automated motor shutdown for overpressure, over-temperature, and off-nominal trajectory

    • Redundant safety systems

  10. Reusability:

    • Designed for reuse comparable to commercial off-the-shelf motors

  11. Significance:

    • Advances SLU’s propulsion research capabilities

    • Demonstrates the viability of hybrid propulsion in collegiate rocketry

Design

The hybrid rocket engine developed for Project Northstar is a high-thrust propulsion system utilizing liquid nitrous oxide (N₂O) as the oxidizer and potassium nitrate-coated acrylic as the solid fuel. Pottasium Nitrate serves as an initial ignition method. The chamber is constructed from 6061-T6 aluminum body with a 4.0-inch outer diameter and incorporates a 25-inch phenolic thermal liner to mitigate heat transfer to the structural casing. The fuel grain measures 23.8 inches in length and is ported to optimize oxidizer flow and regression rate. The nozzle is machined from graphite with a 1.33-inch throat diameter and custom-engineered converging/diverging angles to maximize exhaust velocity and maintain structural resilience under high-temperature conditions. The oxidizer tank stores approximately 20.96 pounds of nitrous oxide and is sealed via dual EPDM O-ring aluminum bulkheads, pressure relief valves, and instrumentation-grade pressure transducers. The system includes an onboard microcontroller-based avionics suite with a Kalman filter for real-time state estimation and thrust prediction. During initial static testing, the engine produced a peak thrust of 725 lbf with a total impulse exceeding 2,000 lbf·s.

Testing

Cold Flow Tests:

  • Purpose: Characterize the performance of the injector and plumbing system without combustion.

  • Outcomes: Verified pressure vessel integrity, valve timing, and fill times under simulated operating conditions.

  1. Hot Fire Tests:

    • Conducted to validate full-system operation, including ignition reliability and sustained combustion.

    • Metrics observed included thrust profile, chamber pressure, injector performance, and grain regression.

  2. Sensor Calibration and Validation:

    • Pressure and temperature sensors were calibrated before integration.

    • Data collected during tests was cross-validated with expected values from simulations.

  3. Iterative Design:

    • Data from each test informed refinements in injector design, thermal liner selection, and grain geometry.

    • Resulted in improved ignition consistency and better control over oxidizer flow rates.

  4. Safety Protocols:

    • Testing was performed remotely with emergency cutoff and blowout systems in place.

    • Procedures for pre-test checklists and post-test inspections were followed strictly.

Test Discussion

During testing, the intended oxidizer tank operating pressure was not fully achieved, resulting in a lower than expected pressure differential across the injector.

  1. This limited oxidizer flow rate, reducing the overall thrust and altering the oxidizer-to-fuel (O/F) ratio from the design target.

  2. Despite this, the engine still achieved stable ignition and sustained combustion, allowing valuable data collection.

  3. The test confirmed mechanical and structural integrity of the propulsion system, validating core subsystems including plumbing, injector mounting, and thermal protection.

  4. Post-test analysis suggested the pressurization system was the limiting factor, leading to plans for increasing pressurant bottle pressure or flow capacity in future iterations.

  5. SLU’s first ever operational hybrid rocket engine

Benjamin Khotsyphom, James Winkelhoch, Hao Lai Jin, Jack Herlihy