Oregon State University Hybrid Rocket Final Presentation

1. June 7, 2016
Final Presentation

2. Overview
Project description
Design process
Sub-team presentations:
•Hot End 1
•Hot End 2
•Cold End
•Vehicle Engineering
•Avionics
•Launch Rail
Budget
Overall results and recommendations
Conclusion, Q&A

3. Project Description
What is a hybrid rocket?
spacesafetymagazine.com
Project background and purpose:
• Sponsored by OSU chapter of AIAA
• Design and build a flight-viable hybrid rocket
• Research and optimize for the future
• Collaborative design experience

4. Multidisciplinary Collaboration
Fuel and Energetics - Chemical
Engineers
• Fuel composition
optimization
• Combustion reaction
energetics
Data Acquisition - Electrical
Engineers
• DAQ system for propulsion
test stand
bbc.co.uk
Hybrid Rocket Design - Mechanical Engineers
● Hot End 1 - injector, fuel, igniter
● Hot End 2 - combustion chamber, injector
manifold, nozzle
● Cold End - oxidizer feed system, remote
priming system
● Vehicle Engineering - recovery system,
aerodynamics, structures, integration
Launch Rail -
Mechanical
Engineers
● Large
enough for
hybrid
rocket
● Collapsible
for travel

5. General Design Process
Team Customer Requirements:
1. Safe 4. Can be
assembled quickly
2. Reliable 5. Flight viable
3. Cost-effective 6. Lightweight
House of
Quality:
CRs and ESs
Simulation:
Solidworks
ANSYS
OpenRocket
NASA CEA
Sub-system
testing:
Sub-team
TPs
Full-scale
testing:
Test fires
Dry Run
Assemblies

6. Hot End 1
314-1
Max Flansberg
Anthony Harteloo
David Ha

7. Subsystem Layout
Post Combustion
Chamber
Fuel Pre Combustion
Chamber Injector

8. Fuel Composition
• Function
– Provide the chemical energy
to propel the rocket
• Rocket Fuel Composition
– Composed of paraffin wax, corn starch, and aluminum powder
– Corn starch plasticizes the paraffin wax
– Aluminum powder increases the combustion temperature
• Customer Requirements
– Thrust to weight ratio
– Operating Temperatures
– Specific impulse of motor

9. Igniter
• Function
– Preheat the combustion chamber
– Decompose oxidizer
• Powderless solid igniter grain
– 65% potassium nitrate, 25% sugar, 10% corn syrup by mass
– Ignited using commercial E-match
– Located in pre-combustion chamber
• Customer Requirements
– Suitable chamber temperature
– Highly reliable
– Increase specific impulse
Source:
https://www.youtube.com/watch?v=Bgy
C1jXTY4c

10. Injector
• Function
– Inject oxidizer into combustion chamber
– Strong effect on motor performance
• Dictates flow field in rocket
• Swirling Injector
– Increases burning rate of fuel
– Increased combustion efficiency over
design alternative
• Customer requirements satisfied
– Increase specific impulse
– Combustion efficiency
– Reliable
– Flight Viable

11. Results
• Fuel
– Specific impulse 210 s
• Igniter
– Chamber temperature > 823K
– Burn time > 2s
• Injector
– 54% increase in thrust
– 32% increase in specific impulse
– Substantial improvement to combustion efficiency
– Subsystem weight within tolerance

12. Recommendations
• Increase melting temperature of the fuel
• Scale up motor for competition
• Lengthen the post combustion chamber
• Shorten the precombustion chamber
• Continue researching high energy fuels
• Refine igniter grain manufacturing process

13. Hot End 2
314-4
Ben Smucker
Frank Huynh
Kyle Fox

14. Overview
• Injector Integration (Ben)
• Combustion Chamber (Frank)
• Nozzle (Kyle)
• Relevant Customer Requirements
– Lightweight
– Durability of parts
– Conform to size restraints

15. Injector Integration
• Manifold /Combustion Chamber
– 4-40 socket head cap screws
– Loaded in tension
• Injector Attachment
– Lip
– No fasteners in injector

16. Combustion Chamber
• Aluminum Chamber
• Initial design failed during testing
• Wall thickness: increase from 3.6
mm to 8.0 mm
• Maximum temperature: 600F
• Tensile Yield Strength: 4640 psi
• Design Pressure of 400 psi
Initial Design
Final Design

17. Nozzle
• Accelerate flow of combustion
products
• Conical nozzle for simplicity
• Rocket may not leave rail safely
• Optimized to help meet speed
requirement off launch rail
• Nozzle Design has uncertainty
– Combustion gases require mixing
– Design does not change much

18. Results and Recommendations
Results
• Maintained integrity during testing
• Nozzle produced phenomena indicative of flow acceleration
• Assembly fits in the rocket
• Some tests failed in order to fit the motor in the rocket.
Recommendations
• Design for manufacturing (Avoid boring!)
• More heat transfer analysis in the combustion chamber walls
• Further optimize nozzle for launch altitude

19. Cold End
314-3
Joshua Laas
Luis Mendoza
Nigel Swehla

20. Overview
• Driving CRs
– Safe pressurization
– Non-corrosive
– High Isp
• Proposed problems
– Safety
– Spatial constraints
– Acceptable pressure &
mass flow rate
• Subsystem solutions
– Oxidizer/pressurant
– Feed system components
– Remote oxidizer priming

21. Oxidizer/Pressurant
• Nitrous Oxide
– Non-cryogenic
– Common
– Non-toxic
– Easy storage
– Better performance
– Cooling
• Self-pressurizing
– Single tank
– No pressurant
– Acceptable pressures

22. Feed System Components
• Single tank configuration
– Light weight
– Smaller size
• Material selection
– Stainless steel
– Brass
– Aluminum
• Physical properties
– Length: 44.90 in
– Width: 4.65 in
– Weight: 12.30 lbs (dry)

23. Remote Oxidizer Priming
• Collapsible arm
– Quick disconnect fittings
– Worm drive and gear motor
– Servo to Initiate
• Accessibility
– Side of rocket
– Door
• Material selection
– Stainless steel
– Electronic disconnect

24. Results
Passed: 10 out of 12 testing procedures
Failed:
Specific Impulse Performance Test
Target: 221-270 s
Achieved: 210 s
Reason: Ambient temperature of 56℉ during test, optimal
oxidizer temperature between 70 and 74℉
Oxidizer Pressure Test
Target: 760-1500 psi
Achieved: 550 psi
Reason: Ambient temperature of 56℉ during test, optimal
oxidizer temperature between 70 and 74℉

25. Recommendations
For better results:
• Test rocket motor vertically
• Use double valve carbon fiber tank
• Pressurize with inert gas
• Decrease plumbing pressure drop
– Test larger check valve
• Temperature control storage tank
For ease of data analysis:
• Ensure all sensors work for every test
• Advance method for cleaning up data
• Create a venturi flow meter
https://www.youtube.com/watch?v=ZyfvJF529no
For safety:
• Keep impressing other hybrid teams by enhancing
existing safety mechanisms, preventing spontaneous
nitrous oxide reactions http://www.simmonsmfg.com/wp-
content/uploads/2012/11/PAGE-5-
CC1b.jpg

26. Vehicle Engineering
314-2
Rodney Fischer
Krissy Kellogg
Parker Weide

27. Structures
•Rolled carbon fiber body tubes (ICE)
•Fiberglass couplers
•Aircraft plywood bulkheads, centering rings
–Steel dowel pins where necessary
•Blue tube motor tube
Relevant Customer Requirements:
• Integrates with motor assembly, recovery, avionics
• Stable throughout all stages of flight
• Easy to manufacture

28. Recovery
Parachute:
• Modified cruciform, Nylon webbing shroud
lines, Kevlar shock cord, stainless steel
hardware
• Total weight 0.362 kg
• Provides 43.4 lb drag force, for a landing
velocity of 7.3 m/s
Deployment:
• Piston
• Aluminum charge holder
• 1.2 g Triple 7 ejection
charge, e-match ignitionRelevant Customer Requirements:
• Recoverable with minimal damage
• Easy to manufacture
• Lightweight

29. Aerodynamics
Fins
• Trapezoidal with Airfoil
• T6 6061 Aluminum
• Manufactured on CNC
Relevant Customer
Requirements:
• Stable throughout all stages of flight
• Recoverable with minimal damage
Nose Cone
• Von Karman with 5:1 Fineness Ratio
• Fiberglass with Aluminum tip
• Male mold manufacturing process
Relevant Customer Requirements:
• Stable throughout all stages of flight
• Lightweight

30. Integration
Motor assembly → Plumbing → Tank+Avionics → Recovery → Nosecone
Relevant Customer Requirements:
• Integrates with motor assembly, recovery, avionics
• Requires reasonable assembly time
• Easy to manufacture

31. Results and Recommendations
Results
• Solutions provided good balance of simplicity and
reliable functionality
• Passed 10 of 14 testing procedures
– True failure: stability margin, pin shear force
– Failed body tube bending, body tube buckling due to
underestimation of performance (rocket is stronger than
anticipated)
Recommendations:
• Recovery bulkhead e-match connectors
• RF transparency to simplify wireless communication
• Female mold for nosecone

32. Avionics
Beaglebone Black embedded
computer to control ignition,
oxidizer valve
Redundant Stratologger
systems to deploy parachute
Xbee wireless module for
remote control and
communication

33. Launch Rail
322
Moises Higgins
Michael Arthur
Loren Valiente
Nathan Leendertse
Sam Monroe

34. Launch Rail: Project Description

35. Launch Rail: Design Development

36. Launch Rail: Design Solution

37. Launch Rail: Results
Budget

38. Launch Rail: Recommendations

39. Budget
Rocket Sub-Team Amount
Hot End 1 $1406.46
Hot End 2 $1020.95
Cold End $1824.07
Vehicle Engineering $336.14
Data Acquisition $321.19 (approx.
394.71)
Avionics $158.99
Total (goal $5000) $ 5046.61
Launch Rail
Total (goal $2000) $1990

40. Overall Results and Recommendations
Results:
• Flight viable rocket
• Launch scrubbed due to time constraints
– Avionics communication issue
– Ball valve battery issue
• Excellent teamwork and collaborationRecommendations:
• Divide up responsibilities differently
• More flight-like testing
• Larger motor

41. Conclusion
Sincere thanks to OSU AIAA, John Lyngdahl,
Steve Cutonilli and Dr. Squires for their support!

42. Questions?