1. Project METEOR Hybrid Rocket Motor Team Members: Marc Balaban
Ken Court
Chad Eberhart
Patrick Haus
Chris Natoli
Nohl Schluntz
5/11/2019

2. Contents Overview of METEOR Hybrid Background Team Organization
Constraints/Risks
Deliverables
Needs Assessment
Objective Tree
Specifications List
Concept Selection Criteria
Hybrid Rocket Concept Strategy
Concepts
Overall Hybrid Rocket Motor
Structure
Injection
Combustion/Exhaust
Nozzles
Ignition
SD1 Project Plan
Questions/Discussion
5/11/2019

3. Project METEOR Overview
Projected Flight Pattern
~ 24 km
5/11/2019

4. Current Test Chamber Setup
Hybrid Background
Classified as utilizing a liquid oxidizer and solid propellant to achieve thrust
Current Oxidizer: Nitrous Oxide (NOX)
Current Propellant: Hydroxyl Terminated Poly-Butadiene (HTPB)
Current Test Chamber Setup
Hydroxyl-Terminated Polybutadiene (HTPB) Fuel Grain
Snap Ring
Chamber Wall
Injector Plate
Garolite Pre & Post Combustion Chambers
2-D Nozzle
5/11/2019

5. Team P08105 Organization Patrick Haus – Project Leader
Chris Natoli – Lead Engineer
Marc Balaban – Fluids/Combustion Specialist
Ken Court – Fuel Injection Specialist
Chad Eberhart – Fluids/Combustion Specialist
Nohl Schluntz – Fuel Injection Specialist
5/11/2019

6. Constraints – Risks Scope of Work Budget = $5000
Time = 22 Weeks (16 Remaining)
Integration to other METEOR Teams
Safety
Internal Temperature and Pressure
Titanium Shell (Process & Strength)
Lead Time
Machine Shop Availability
External Resources
5/11/2019

7. Project Deliverables
Provide a liquid/hybrid rocket motor that will support a launchable test flight
This will be accomplished by improving the following subsystems:
Safety
Improve safety procedures based on incident in the Mojave desert
Utilize factors of safety appropriately in motor design
Structure
Integration with P08106 (Flying Rocket Body)
Injection
Injector Plate – Reduce pressure losses and increase atomization
Feed system – Improve overall efficiency and safety
Combustion & Exhaust
Ignition – Built in Redundancy
Fuel Grain Geometry and Composition – Change geometry and additives to increase thrust predictability and specific impulse
Supersonic Nozzle – Utilize Brendan Denton’s contoured nozzles to increase efficiency
Experimental Design & Data Analysis
Water jet testing through Injector Plate
CFD
ANSYS
Structural Testing
Vertical/Horizontal Test Stand
Ignition testing
Fuel Grain Regression
5/11/2019

8. Project Needs Overall: Launch capability at 80,000 ft.
Provide a specific impulse of 220s.
Deliver a mass to propellant ratio of 1:10.
Offer repeatability.
Ensure safety to all participants.
Injection:
Minimize pressure loss across the feed system and injector plate.
Provide a consistent atomization of the Nitrous Oxide.
Resistance to high temperature
Withstand high pressures (roughly 400 psi).
Integrate easily with the combustion chamber.
Provide mixing with the oxidizer at reasonable cost. (Outside help from Delphi)
Combustion/Exhaust:
Provide a reliable ignition system.
Withstand high pressures and temperatures.
Provide predictable regression rates and thrust.
Integrate the nozzle design from Brandon Denton.
Structure:
Integrate into the rocket body and be structurally sound.
Limit complexity of fabrication.
Should be easily assembled and disassembled.
Withstand high temperature and pressure.
5/11/2019

9. Project Specifications
To be handed out as a separate sheet
Concept Grading Criteria
To be handed out as a separate sheet
5/11/2019

10. Hybrid Rocket Concept Strategy
Test Flight
Design Flying Rocket
- Structure
- Injector
- Combustion
- Ignition
- Nozzle
Test Steel Rocket
Design Iteration
5/11/2019

11. Overall Hybrid Rocket Motor Concept
Supersonic Nozzle (Graphite)
Post-Combustion (Garolite)
HTPB Fuel Grain
Pre-Combustion (Garolite)
Injector Plate
Brackets
Titanium Shell (t = 1/16’’)
Support Rod
5/11/2019

12. Structure Details
Brackets support Injector Plate and Nozzle while under compression
Combustion chamber easily attaches to support rods implemented by the Rocket Body Team
Advantages
Assembly
Lightweight
Simple to machine/replace
Easily Adaptable for Redesign
Disadvantages
Thin sections
Possibility for Shear
Leakage/Sealing
5/11/2019

13. Structure Concept
5/11/2019

14. Injection Concepts Assumptions
Fully Developed, Turbulent, Steady, Viscous Flow
Liquid Phase
Current test setup
Nine Hole Straight Injector (1-Piece)
Bolts to Steel Test Chamber
Improvements
Determine Phase
Interchangeable Injector Inserts
Snap Ring/Thread In
Geometry of Inserts (Flow Characteristics)
Working with DELPHI (Bill Humphrey)
Atomization
Relocation of Pressure Transducer
Weight Reduction
CFD Modeling/Dupont?
5/11/2019

15. Injection Concept 1
5/11/2019

16. Injection Concept 2
5/11/2019

17. Combustion/Exhaust Background
Assumptions
Locally Isentropic
Compressible Flow
Ratio of Specific Heats (γ) = 1.258
Complete Combustion
Current test setup
HTPB (Hydroxyl-Terminated Polybutadiene) as the solid fuel grain
Nitrous Oxide (NOX) as the oxidizer
Cylindrical fuel grain geometry
Linear (Laval) Convergent-Divergent Supersonic Nozzle
Improvements
Fuel Grain Type (Paraffin Wax, HTPB)
Geometry
Possibility of additives to fuel grain (ie. Aluminum Powder)
Contoured Supersonic Nozzles (Brandon Denton – Thesis Student)
5/11/2019

18. Fuel Grain Geometries Cylindrical (Current) Circular-Cross
6-Point Star
10-Point Star
5/11/2019

19. Supersonic Nozzles Linear
- Advantages: Simpler machineability, simple design
- Disadvantages: Lower efficiency
Contoured
Advantages: Higher thrust achievement
Disadvantages: More difficult to machine
Aerospike
Advantages: Maximum thrust at all altitudes
Disadvantages: Experimental, most difficult to machine
May implement contoured supersonic nozzles
Brandon Denton is a graduate student of Dr. Kozak
Will determine most thrust efficient nozzles at 80,000 ft
Cost-performance Analysis
To be performed by Test Stand team
Most efficient (Cost and Performance) will be integrated into final design
5/11/2019

20. Supersonic Nozzles Annular - Mach 4 Annular - Mach 3
8° Conical - Mach 4
12° Conical - Mach 4
5/11/2019

21. Ignition Concepts - Igniters
Ultra-Low Current Igniter
Materials
1.2 V, 20mA light bulb
Black powder
Advantages
requires only 25mA to fire
highly shock proof
Disadvantages
very rapid burn rate
Pyromix Igniter
Materials
Nichrome bridgewire-30 AWG
Pyromix
potassium perchlorate
sulfur
high grade epoxy
Advantages
requires very low voltage to fire
epoxy offers hotter and more sustained
burn rates
5/11/2019

22. Ignition Concepts – Igniters continued
Thermex Igniter
Materials
1.2V, 20mA light bulb
Thermex powder
65% potassium perchlorate
20% charcoal
10% aluminum powder
5% red Ferric Oxide
Advantages
Requires very low voltage to fire
- Thermex offers much hotter and more
sustained burn rates
5/11/2019

23. Ignition Concepts - Delivery
Electric charge will be delivered via wires running from launch platform through rocket nozzle
Lithium Ion batteries operate nominally from -40 to 60 degrees Celsius
Batteries will be delivering charge from platform to save weight on rocket
Wires will detach from rocket via burn off from ignition
Built-in redundancy
If batteries fail, rocket will have onboard altimeter with a second set of wires running through nozzle, which will fire at given altitude
Wires from platform will have quick connects that will disengage with 1-2 pound force
5/11/2019

24. Senior Design I Project Plan
Concept Prove Out
ANSYS
Structural
Thermal
FLUENT
Injector
Combustion Chamber
Nozzle
Pro Engineer
Final Design Models
Testing Methods
Parts/Material Ordering
*See Project Plan Handout for Timeline
5/11/2019

25. Questions/Comments Please feel free give suggestions and critiques
5/11/2019