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

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

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

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

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

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

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

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

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

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

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

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

Structure Concept 5/11/2019

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

Injection Concept 1 5/11/2019

Injection Concept 2 5/11/2019

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

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

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

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

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

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

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

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

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