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transcript
1 Long Beach Rocketry | CDR 2017 - 2018
Table of Content
1) General Information ...................................................................................................... 10
1.1 Student Leader ........................................................................................................ 10
1.2 Safety Officer .......................................................................................................... 10
1.3 Team Structure ........................................................................................................ 10
1.4 NAR/TRA Sections ................................................................................................ 11
2) Summary of CDR Report.............................................................................................. 12
2.1 Team Summary ....................................................................................................... 12
2.2 Launch Vehicle Summary ....................................................................................... 12
2.3 Payload Summary ................................................................................................... 12
3) Changes Made Since PDR ............................................................................................ 13
3.1 Changes to Launch Vehicle .................................................................................... 13
3.2 Changes to RDM..................................................................................................... 13
3.3 Changes to DORITO............................................................................................... 13
4) Safety ............................................................................................................................ 15
4.1 Safety Officer Duty ................................................................................................. 15
4.2 Compliance with Safety Plan .................................................................................. 16
4.3 Safety Equipment .................................................................................................... 16
4.4 Facilities .................................................................................................................. 16
4.5 Injury and Emergency ............................................................................................. 16
4.6 NAR Safety Code Compliance ............................................................................... 17
4.7 Compliance with Federal, State, and Local Laws ................................................... 19
4.8 Handling Rocket Motor .......................................................................................... 20
4.9 Range Safety Regulation......................................................................................... 20
4.10 Risk Assessment Codes ........................................................................................ 21
4.11 Personal Hazard Risk Assessment ........................................................................ 23
4.12 Failure Mode Hazard Risk Assessment ................................................................ 27
4.13 Environmental Hazard Risk Assessment .............................................................. 38
4.14 Launch Operation Procedure ................................................................................ 42
4.15 Troubleshooting .................................................................................................... 47
5) Launch Vehicle Criteria ................................................................................................ 48
5.1 Mission Statement ................................................................................................... 48
5.2 Mission Success Criteria ......................................................................................... 48
5.3 Launch Vehicle Overview ...................................................................................... 48
5.4 Launch Vehicle Design ........................................................................................... 50
5.5 Subscale Flight Results ........................................................................................... 68
5.6 Recovery System .................................................................................................... 75
5.7 Mission Performance Prediction ............................................................................. 86
5.8 Team-Derived Requirements .................................................................................. 99
6) Payload Criteria – Rover Deployment Mechanism (RDM) ...................................... 105
6.1 System Overview .................................................................................................. 105
6.2 RDM System Design ............................................................................................ 107
6.3 Subscale Prototype ................................................................................................ 123
6.4 Team Derived Requirement-RDM ....................................................................... 128
7) Payload Criteria – DORITO ....................................................................................... 134
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7.1 System Overview .................................................................................................. 134
7.2 Rover Design ........................................................................................................ 135
7.3 Electronics............................................................................................................. 150
7.4 Control System...................................................................................................... 154
7.5 Subscale Prototype ................................................................................................ 155
7.6 Mass Budget.......................................................................................................... 156
7.7 Team Derived Requirement – DORITO ............................................................... 157
8) Testing Plan ................................................................................................................ 160
8.1 Approach to Testing .............................................................................................. 160
8.2 Launch Vehicle Testing Plan ................................................................................ 160
8.3 RDM Testing Plan ................................................................................................ 164
8.4 DORITO Testing Plan .......................................................................................... 168
9) Project Plan ................................................................................................................. 170
9.1 Requirement Verification...................................................................................... 170
9.2 Timeline ................................................................................................................ 193
9.3 Budget ................................................................................................................... 196
9.4 Funding Plan ......................................................................................................... 201
9.5 Educational Engagement ...................................................................................... 203
10) Appendices .................................................................................................................. 206
10.1 References ........................................................................................................... 206
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List of Figures
Figure 1.1: 2017-2018 NASA Student Launch team structure
Figure 5.1: Full-Scale Launch Vehicle CAD
Figure 5.2: Full Scale CAD Model Split into Sections
Figure 5.3: CAD of the nose cone
Figure 5.4: Ogive Nose Cone Flow Simulation results using SolidWorks
Figure 5.5: Full Scale Launch Vehicle Simulation Results
Figure 5.6 CAD Model of Payload Section
Figure 5.7 CAD Model of Payload Section with Inside View
Figure 5.8: Full-Scale Launch Vehicle Recovery and Avionics Dimensions
Figure 5.9: CAD of Avionic Bay (Collapsed View)
Figure 5.10: CAD of Avionics Bay (Exploded View) Figure 5.11: 3D printed AV tray
Figure 5.12: Avionics Tray (Collapsed View)
Figure 5.13: Avionics Tray (Exploded View)
Figure 5.14: U-Bolt and Aluminum Bulkhead Attachment (Exploded View)
Figure 5.15: U-Bolt and Aluminum Bulkhead Attachment (Collapsed View)
Figure 5.16: CAD Model of Propulsion Section with Inside View
Figure 5.17: Fin Drawing
Figure 5.18: Fin Slot Alignment Jig
Figure 5.19: CAD of Motor Casing
Figure 5.20: Stress Analysis on Centering Ring Using SolidWorks Simulation
Figure 5.21: Stress Analysis on Thruster Plate Using SolidWorks Simulation
Figure 5.22: Stress Analysis on Fin Using SolidWorks Simulation
Figure 5.23: Results from Fin Analysis producing a Divergence Velocity of 3287.16 ft/s and
Flutter Velocity of 4471.68 ft/s
Figure 5.24: Fin Stress Analysis confirming stress experienced by fin does not exceed maximum
allowable stress
Figure 5.25: Subscale Launch Vehicle OpenRocket Design
Figure 5.27: Experimental Data for Subscale Launch
Figure 5.28: Simulation for Subscale Drift
Figure 5.29: Subscale Launch Vehicle post-flight recovery
Figure 5.30: RDM Performance During Subscale Launch
Figure 5.31: Launch Vehicle Dual Deployment Attachment Layout
Figure 5.32: Diagram of the Launch Sequence
Figure 5.33: Block Diagram of Altimeter Electrical Connections
Figure 5.34: Electrical Schematic of Altimeters in Avionics Bay
Figure 5.35: GPS Tracking of Sub-Scale Launch
Figure 5.36: Ground Ejection Test for Drogue Parachute
Figure 5.37: Ground Ejection Test for Main Parachute
Figure 5.38: MATLAB/Simulink simulation (beta phase)
Figure 5.39: Graph of Launch Vehicle Motion vs. Time with 0 mph Wind Conditions
Figure 5.40: Motor thrust curve for an AeroTech L1390G motor
Figure 5.41: OpenRocket side view of launch vehicle
Figure 5.42: Hand calculations of the static stability
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Figure 5.43 OpenRocket Simulation of Drift in 5 mph Winds
Figure 6.1: CAD of full integrated RDM design.
Figure 6.2.1: Half section of RDM subscale
Figure 6.2.2: Half section of RDM full scale
Figure 6.3: Figure of rod assembly of rods in full scale
Figure 6.4.1: Assembly of rover-pushing-plate (Exploded on left)
Figure 6.4.2: Drawing of front half of rover-pushing-plate
Figure 6.4.3: Drawing of back half of rover-pushing-plate.
Figure 6.5: Complete full scale RDM assembly with rover
Figure 6.6: Assembly of full scale nose cone bulkhead (Exploded on left)
Figure 6.7.1: Drawing of the front half of nose cone bulkhead.
Figure 6.7.2: Drawing of back half of nose cone bulkhead.
Figure 6.8: Assembly of full scale nose cone and nose cone bulkhead
Figure 6.9: Screw and nut diagram with parameter designation
Figure 6.10.1: Assembly and exploded view of full scale electronics bay
Figure 6.10.2: Drawing of full scale electronics bay components.
Figure 6.11: Full scale electronics assembly with components
Figure 6.12: Graph of current drawn to break three shear pins using a 32 RPM motor
Figure 6.13: Electric schematic of full scale RDM
Figure 6.14.1: Location of the RDM section relative to the rocket
Figure 6.14.2: Internal components of RDM system: Labelled
Figure 6.15: Assembly of subscale nose cone bulkhead (Exploded on left)
Figure 6.16.1: Assembly of top of subscale motor bulkhead
Figure 6.16.2: Exploded view of top of subscale motor bulkhead
Figure 6.17.1: Bottom view of subscale RDM electronics bay
Figure 6.17.2: Top view of subscale RDM electronics bay
Figure 6.18.1: Top view of subscale electronics bay with components
Figure 6.18.2: Side views of subscale electronics bay with components
Figure 7.1: LBR rover design
Figure 7.2: Rover center driveshaft
Figure 7.3: Rover gearbox
Figure 7.4: Rover exterior body
Figure 7.5: Rover bogie system
Figure 7.6: Bogie system rod
Figure 7.7: Bogie system ball joint
Figure 7.8: Integrated bogie system and components
Figure 7.9: Wheel design
Figure 7.10: Tire selection
Figure 7.11: Wheels and tires on rover
Figure 7.12: Rover exterior body
Figure 7.13: Rover gearbox system
Figure 7.14: Rover motor mount
Figure 7.15: Rover motor mount
Figure 7.16: Rover bogie system
Figure 7.17: Solar panel
Figure 7.18: Deployment mechanism before and after deployment side view
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Figure 7.19: Deployment mechanism before and after deployment
Figure 7.20: Rover side panels which house solar cells
Figure 7.21: Rover with side panel assembly
Figure 7.22: Rover Circuit Diagram
Figure 7.23: Arduino Nano
Figure 7.24: L298n Motor Driver
Figure 7.25:116 RPM planetary gear motor
Figure 7.26: System Response Flowchart
Figure 7.27: Rover subscale model
Figure 9.1: Gantt Chart for 2017-2018 Competition
Figure 9.2: Budget and Present Expenditures
Figure 9.3: Fulfillment Percentages of Various Funding Methods
Figure 9.4: Sponsorship tier system
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List of Tables
Table 1.1: Student Leader Information
Table 1.2: Safety Officer Information
Table 1.3: NAR/TRA sections
Table 4.1: NAR Safety Code Compliance
Table 4.2: Risk Assessment Code (RAC)
Table 4.3: Risk Definition
Table 4.4: Severity Definition
Table 4.5: Probability Definition
Table 4.6: Personnel Hazard Risk Assessment
Table 4.7: Failure Mode Hazard Risk Assessment – Structure and Propulsion
Table 4.8: Failure Mode Hazard Risk Assessment – Recovery
Table 4.9: Failure Mode Hazard Risk Assessment – Rover
Table 4.10: Failure Mode Hazard Risk Assessment – Rover Deployment (RDM)
Table 4.11: Environmental Risk Assessment – Environment Impact on Rocket
Table 4.12: Environmental Risk Assessment – Rocket Impact on Environment
Table 4.13: Troubleshooting
Table 5.1: Launch Vehicle Section Lengths and Weight
Table 5.2: Launch Vehicle Flight Specifications
Table 5.3: Coefficient of Drag Data from the Simulation
Table 5.4: Material Properties of Carbon Fiber
Table 5.5: Experimental Data vs. Simulation Data
Table 5.6: Subscale Launch Drift Data
Table 5.7: Full-Scale Black Powder Calculations
Table 5.8: Subscale Black Powder Calculations
Table 5.9: Subscale Black Powder Calculations
Table 5.10: Projected Apogee at Different Wind Speeds Using OpenRocket
Table 5.11: Definition of Hand Calculation Symbols
Table 5.12: Full-Scale Descent Velocity Calculations
Table 5.13: Kinetic Energy Calculations for Each Independent Section
Table 5.14: Drift Distance Calculations
Table 5.15: Team-Derived Requirements – Launch Vehicle
Table 6.1: Comparison of features: subscale design (Option 1 vs. Option 2)
Table 6.2: Estimated mass of mechanical components: full scale
Table 6.3: Power screw calculator parameters and results
Table 6.4: Estimated mass of full scale electronics
Table 6.5: Mass of subscale electronic components
Table 6.7: NASA Experiment Requirements
Table 6.8: Team derived requirements verification status and plan (RDM)
Table 7.1: Mass of DORITO
Table 7.2: Team Experiment Requirements
Table 8.1: RDM full scale testing comparison chart
Table 9.1: General Requirements
Table 9.2: Launch Vehicle Requirements
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Table 9.3: Recovery Requirements
Table 9.4: Deployable Rover Requirements
Table 9.5: Safety Requirements
Table 9.6: Expected Development Schedule
Table 9.7: Expected Development on Each Team
Table 9.8: Projected Expense
Table 9.9: Projected Income
Table 9.10: Educational Engagement Schedule
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Acronyms
CSULB: California State University, Long Beach
LBR: Long Beach Rocketry
ASI: Associated Students Inc.
AESB: Associated Engineering Student Body
AIAA: American Institute of Aeronautics and Astronautics
SGT: Sigma Gamma Tau, National Aerospace Engineering Honor Society
PDR: Preliminary Design Review LRR: Launch Readiness Review
FRR: Flight Readiness Review CDR: Critical Design Review
RIOS: Robotic arm Interactive Operating System ROC: Rocketry Organization of California
NAR: National Association of Rocketry
TRA: Tripoli Rocket Association
MSDS: Material Safety Data Sheets
ROC: Rocket Organization of California
CAD: Computer Aided Design
CNC: Computer Numerically Controlled
CATIA: Computer Aided Three-dimensional Interactive Application
MATLAB: Matrix Laboratory
NVEDMS: Noise and Vibration Engineering Database Management System
GPS: Global Positioning System
PVC: Polyvinyl Chloride
AGL: Above ground level
APCP: Ammonium Perchlorate Composite Propellant
FEA: Finite Element Analysis
MAES: Latinos in Science and Mathematics
SWE: Society of Women Engineers
CALVEIN: California Launch Vehicle Education Initiative IRA: Instructionally Related
Activities STEM: Science Technology Engineering and Mathematics
ABS - Acrylonitrile Butadiene Styrene
CAD - Computer Aided Design
CNC - Computer Numerical Control
DC- Direct Current
FOS - Factor of Safety
IMU - Inertial Measurement Unit
LBR- Long Beach Rocketry
LiPo - Lithium Polymer
MCU - Microcontroller Unit
PLA - Polylactic Acid
PWM - Pulse Width Modulation
RC - Remote Control
RDM - Rover Deployment Mechanism
RPM - Revolutions Per Minute
SAE - Society of Automotive Engineers
DORITO – Dynamically Oriented Rocket Integrated Triangular Object
LV- Launch Vehicle
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AV- Avionic
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Section 1: General Information
1.1 Student Team Leader
The Team Leader is Nam Nguyen (California State University, Long Beach, class of 2019). The
2017-2018 academic year is Nam’s third year on LBR.
Table 1.1: Student Leader Information
Name Nam Nguyen
Title of Long Beach Rocketry Team Leader
Contact nguyentransinam1717@yahoo.com
(714) 797-6784
1.2 Safety Officer
The team Safety Officer is Shawn Everts (California State University, Long Beach, class of
2018). The 2017-2018 academic year is Shawn’s first year on LBR, and he has gone through all
the mandatory safety trainings during the summer for engineering teams.
Table 1.2: Safety Officer Information
Name Shawn Everts
Title of Long Beach Rocketry Safety Officer
Contact shawneverts@gmail.com
(951) 440-6953
1.3 Team Structure
The 2017-2018 Long Beach Rocketry Team will consist of approximately 20 students from a
variety of backgrounds. The team consists of students from the aerospace engineering,
mechanical engineering, electrical engineering, and computer science departments.
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Figure 1.1: 2017-2018 NASA Student Launch team structure
1.4 NAR/TRA Sections
The team will work with the NAR/TRA sections listed in Table 3 for purposes of mentoring
review of designs and documentation, and/or launch assistance.
Table 1.3: NAR/TRA sections
Section NAR Number TRA Number Launch Field
Location
Rocketry Organization of California
(ROC) 538 48 Lucerne Valley, CA
Friends of Amateur Rocketry, Inc.
(FAR) Not NAR/TRA sponsored Mojave Desert, CA
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Section 2: Summary of PDR report
2.1 Team Summary
School Name: California State University, Long Beach
Organization: Long Beach Rocketry
Mailing Address: 1250 Bellflower Blvd.
Long Beach, CA 92841
LBR will have a Team Advisor and a Team Mentor:
Team Advisor
Dr. Praveen Shankar
Associate Professor
Department of Mechanical and AeroSpace Engineering
praveen.shankar@csulb.edu
562-985-1518
Team Mentor
David Alexander Roy
Visual Artist, High Power Rocketry enthusiast
Otis College of Art and Design
kj6zfv@gmail.com
323-807-9980
2.2 Launch Vehicle Summary
The launch vehicle will be 6 inches’ diameter and 103 inches in length. The launch vehicle is
planned to weight 38.7 lb. with 13% extra mass contingency. The launch vehicle will have a 75-
mm diameter motor mount tube and launch on a AeroTech L1390R-P motor off of a 12-ft long
1515 rail. LBR’s recovery system involves three separate sections landing with the drogue and
the main parachutes. The Milestone Flysheet can be found on LBR’s website at
longbeachrocketry.com/documents.
2.3 Payload Summary
The payload of the Long Beach Rocketry launch vehicle will consist of the deployable rover
design experiment selected from the NASA Handbook. The team will design and build an
autonomous rover capable of deploying solar panels after moving at least 5 feet, as well as a
rover deployment mechanism which will deploy the rover after the team remotely activates the
process. Several designs have been considered but have been narrowed down through trade
studies to ensure the best outcome of the mission.
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Section 3: Changes Made Since PDR
3.1 Changes to Launch Vehicle
A major change in the launch vehicle is the removal of the airbrake subsystem. Due to some
complications and safety concerns, LBR decided to remove the airbrake subsystem from the full-
scale launch vehicle. The general design and layout of the of the full-scale launch vehicle is
similar to the design the team had during the proposal phase with a few adjustments. LBR also
changed the motor from the Cesaroni L1350 CS-P to the AeroTech L1390G.
Another change to the LV is the addition of 3 inches to the payload section and 5 inches to the
main compartment. Because this space previously belonged to the airbrake subsystem, the
overall length of the full-scale launch vehicle did not change significantly.
Another significant change made since the Preliminary Design Review includes swapping the
location of the main and drogue parachutes. Previously, the drogue chute was located between
the Propulsion Bay front end and the Avionics Bay aft end, while the main parachute was located
between the Avionics Bay front end and the Payload Bay aft end. The parachute locations are
now switched, with the drogue being located where the main parachute was and vice versa. The
reason for this change is to increase the stability of the launch vehicle to yield better flight
results, as NASA had previously suggested during the PDR teleconference.
3.2 Changes to RDM
1. Change in the motor speed and torque specifications so that if speed variability is applied
according to the value of the rotary encoder, the speed of the deployment process will not
be severely affected; and at the same time improving the allowable torque capacity to
avoid stalling the motor.
2. Addition of a movable pushing-plate with an embedded nut to push the rover outward
instead of having two hex nuts installed on the rover itself. The main benefit of the rover-
pushing-plate will ensure that the rover will not get caught from any protruding rods.
3. Change in the length configuration of the rods due to the addition of the pushing-plate;
i.e. before the threaded rod is longer than the un-threaded rods. New design will be the
threaded rod is shorter than the two un-threaded rods.
4. Addition of two limit switches: one for the rover assembly into the RDM and another
when the rover is fully released from the RDM.
3.3 Changes to DORITO
Since the PDR submission and subscale launch, LBR has considered several different options in
completing the mission and found room for improvement through the subscale tests which has
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resulted in minor design changes and further development for both the rover and the rover
deployment mechanism.
Although many different rover designs have been considered, ultimately the team has decided to
settle on the triangular rover design and has proceeded with detailed development. The triangular
rover was lengthened to allow for efficient use of payload space and will allow for deployment
on any side of the rover. A gearbox has been created to allow a single motor to control all three
wheels on each side, which also allows the three wheels to pivot on a bogie system and gives
additional space for electronics inside the rover. A tank tread design has been considered through
a trade study due to its additional grip due to more surface area in contact with the ground and
reduced weight compared to the original 3D printed design which did not perform as anticipated
during subscale testing. Also, to reduce complexity in manufacturing the rover, the center hex
coupler rods will be unthreaded since the RDM will have a threaded plate that can push the rover
out of the payload bay. These slight changes have addressed the challenges that arose during the
proposal which will increase the likelihood of mission success for the rover.
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Section 4: Safety
4.1 Safety Officer Duties
Shawn Everts is the official Safety Officer of the Long Beach Rocketry team. The Safety Officer
is responsible for ensuring the safety of persons participating in the NASA University Student
Launch Initiative activities. Shawn will assure that the team adheres to all regulations pertaining
to the construction, assembly, testing, flight, and recovery of the launch vehicle.
Required Training for Safety Officer
The Safety Officer will receive training to gain a thorough understanding of the National
Association of Rocketry (NAR) safety code, the Tripoli Rocketry Association (TRA) safety
code, and the Federal Aviation Administration's (FAA) code pertaining to High Power Rocketry.
The Safety Officer will obtain MSDS sheets for all chemicals used and will keep copies of them
for reference for all team members.
Safety Officer Responsibilities
The Safety Officer will be responsible for the following:
• Ensuring that all team members understand and comply with the NAR high power safety
code, the TRA safety code, and federal, state, and local laws regarding unmanned rocket
launches and motor handling.
• Provide a team safety manual to all members which includes safety plans, precautionary
procedures, NAR regulations, TRA regulations, FAA laws, Material Safety Data Sheets
(MSDS), and operator manuals.
• Verify that all team members have read and adhere to guidelines, laws, and regulations
set forth by the team safety manual.
• Revise the team safety manual throughout the season as the Safety Officer sees fit.
• Ensure team leaders develop and maintain their respective hazard analysis, failure mode
analysis and MSDS/chemical inventory data.
• Ensure that all MSDS information is complete and readily accessible to all team members
while working in the assembly and test area.
• Determine safety violations and take pertinent action to reduce any threat.
• Prepare the team with a safety plan for the various materials used, test procedures, and
settings.
• Institute risk assessment tables that determine the risk level of each threat based on the
likelihood of each event occurring and the severity of each event.
• Ensure that a risk assessment table is created for all possible threats.
• Ensure that all team members understand the risks associated with working on the vehicle
and launcher.
• Ensure team leaders implement and update any procedures to ensure safety in the
construction, assembly, launch and recovery of the vehicle and launcher.
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• Participate in the design, construction, assembly and ground testing of the vehicle and
launcher in order to identify any potential safety risks and to ensure the team adheres to
the team safety plan.
• Participate in the subscale test, full-scale launch test, competition launch and recovery in
order to identify any potential safety risks and guarantee that the team satisfies the team
safety plan.
• Monitor safety during educational engagement activities.
• Enforce the use of safety equipment throughout all stages of design, construction,
assembly, testing, launch, operation, and recovery phases of the launch vehicle.
• Work with the University safety officer Michael Hom located in building ESC-620 to
ensure all procedures are safe and in compliance with California State University, Long
Beach policies.
4.2 Compliance with Safety Plan
A written safety statement has been signed by all members who will be working on the rocket
design and construction. The Safety Officer, will be responsible for communication of the safety
plan to every team member and will ensure that all team members follow the safety protocols at
all times. Anyone who has not signed the safety statement will not be allowed to work on the
project.
4.3 Safety Equipment
Personal Protective Equipment (PPE) such as nitrile gloves, protective glasses and respirator
masks will be available during the construction of the rocket. Material Safety Data Sheets
(MSDS) will be provided and referenced for proper handling, storage and PPE of materials
during the construction of the rocket. Proper PPE required by the MSDS will be enforced when
any team member is working with said material.
4.4 Facilities
Qualified personnel will be present to supervise during the construction of the rocket in all
facilities. All facilities being worked in have shop tools and machinery that are used by students.
University personnel are aware of the shop safety practices and will ensure that students follow
proper operation when using any of the tools or equipment in the facilities.
4.5 Injury and Emergency
First aid kits are readily available in the lab and machine shop with the location of these being
known by all team members. Also located in the lab and machine shop are fire extinguishers, fire
sprinklers and fire alarms that are all periodically checked to ensure that they comply to
regulation. If a chemical related injury occurs, the first aid measures listed in the MSDS will be
followed. Emergency contacts are available at all time on campus through the campus police
department.
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4.6 NAR Safety Code Compliance
Table 4.1: NAR Safety Code Compliance
NAR Code Compliance
Certification. I will only fly high power rockets
or possess high power rocket motors that are
within the scope of my user certification and
required licensing.
David Roy, NAR/TRA personnel, has a level 2
certification from the TRA and he will solely
handle the rocket motors.
Materials. I will use only lightweight materials
such as paper, wood, rubber, plastic, fiberglass, or
when necessary ductile metal, for the construction
of my rocket.
Materials in the launch vehicle consist of
fiberglass, carbon fiber, other lightweight
materials, and a minimal amount of aluminum for
bulkheads.
Motors. I will use only certified, commercially
made rocket motors, and will not tamper with
these motors or use them for any purposes except
those recommended by the manufacturer. I will
not allow smoking, open flames, nor heat sources
within 25 feet of these motors.
Motor will be purchased and shipped to Huntsville
before Launch week. NAR/TAR personnel will be
responsible for storing and handling the motor
before, during and after launch week.
Ignition System. I will launch my rockets with an
electrical launch system, and with electrical motor
igniters that are installed in the motor only after
my rocket is at the launch pad or in a designated
prepping area. My launch system will have a
safety interlock that is in series with the launch
switch that is not installed until my rocket is ready
for launch, and will use a launch switch that
returns to the “off” position when released. The
function of onboard energetics and firing circuits
will be inhibited except when my rocket is in the
launching position.
The rocket will be launched with an electrical
launch system using electrical motor igniters
installed only when the vehicle is in launch ready
configuration on the launch pad. The launch
system has a safety interlock that is installed only
when the rocket is ready for launch. The switch
returns to the “off” position upon release. Onboard
energetics and firing circuits will be turned on
only when the launch vehicle is in launch ready
configuration.
Misfires. If my rocket does not launch when I
press the button of my electrical launch system, I
will remove the launcher’s safety interlock or
disconnect its battery, and will wait 60 seconds
after the last launch attempt before allowing
anyone to approach the rocket.
In the event the rocket does not launch, the Safety
Officer and David Roy will ensure the power
supply to the electrical launch system is off, and
not allow anyone to approach the vehicle for 60
seconds. During a misfire the Range Safety
Officer will have the final say.
Launch Safety. I will use a 5-second countdown
before launch. I will ensure that a means is
available to warn participants and spectators in the
event of a problem. I will ensure that no person is
closer to the launch pad than allowed by the
accompanying Minimum Distance Table. When
arming onboard energetics and firing circuits I
will ensure that no person is at the pad except
safety personnel and those required for arming and
disarming operations. I will check the stability of
my rocket before flight and will not fly it if it
The launch safety requirement will be follow. The
Safety Officer will ensure the minimum distance
table is enforced.
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cannot be determined to be stable. When
conducting a simultaneous launch of more than
one high power rocket I will observe the
additional requirements of NFPA 1127.
Launcher. I will launch my rocket from a stable
device that provides rigid guidance until the rocket
has attained a speed that ensures a stable flight,
and that is pointed to within 20 degrees of vertical.
If the wind speed exceeds 5 miles per hour I will
use a launcher length that permits the rocket to
attain a safe velocity before separation from the
launcher. I will use a blast deflector to prevent the
motor’s exhaust from hitting the ground. I will
ensure that dry grass is cleared around each launch
pad in accordance with the accompanying
Minimum Distance table, and will increase this
distance by a factor of 1.5 and clear that area of all
combustible material if the rocket motor being
launched uses titanium sponge in the propellant.
The team will comply with this NAR code. At the
launch field the Range Safety Officer will
determine if it is safe to launch.
Size. My rocket will not contain any combination
of motors that total more than 40,960 N-sec (9208
pound-seconds) of total impulse. My rocket will
not weigh more at liftoff than one-third of the
certified average thrust of the high-power rocket
motor(s) intended to be ignited at launch.
The team will follow this requirement. Leads will
be responsible for design the rocket with the
constraint.
Flight Safety. I will not launch my rocket at
targets, into clouds, near airplanes, nor on
trajectories that take it directly over the heads of
spectators or beyond the boundaries of the launch
site, and will not put any flammable or explosive
payload in my rocket. I will not launch my rockets
if wind speeds exceed 20 miles per hour. I will
comply with Federal Aviation Administration
airspace regulations when flying, and will ensure
that my rocket will not exceed any applicable
altitude limit in effect at that launch site.
The Safety Officer and NAR/TRA personnel will
ensure that this requirement is followed.
19 Long Beach Rocketry | CDR 2017 - 2018
Launch Site. I will launch my rocket outdoors, in
an open area where trees, power lines, occupied
buildings, and persons not involved in the launch
do not present a hazard, and that is at least as large
on its smallest dimension as one-half of the
maximum altitude to which rockets are allowed to
be flown at that site or 1500 feet, whichever is
greater, or 1000 feet for rockets with a combined
total impulse of less than 160 N-sec, a total liftoff
weight of less than 1500 grams, and a maximum
expected altitude of less than 610 meters (2000
feet).
The test site is NAR/TRA certified and the team
will comply with the what the Range Safety
Officer says.
Launcher Location. My launcher will be 1500 feet
from any occupied building or from any public
highway on which traffic flow exceeds 10 vehicles
per hour, not including traffic flow related to the
launch. It will also be no closer than the
appropriate Minimum Personnel Distance from
the accompanying table from any boundary of the
launch site.
The launch field is set up with the appropriate
distance from any buildings and highway. The
team will comply with the minimum distance
table and follow the instructions from the Range
Safety Officer.
Recovery System. I will use a recovery system
such as a parachute in my rocket so that all parts
of my rocket return safely and undamaged and can
be flown again, and I will use only flame-resistant
or fireproof recovery system wadding in my
rocket.
The Recovery lead and the Safety Officer will
make sure that the design adheres to this
requirement.
Recovery Safety. I will not attempt to recover my
rocket from power lines, tall trees, or other
dangerous places, fly it under conditions where it
is likely to recover in spectator areas or outside the
launch site, nor attempt to catch it as it approaches
the ground.
The team will follow this requirement during
every launch.
4.7 Compliance with Federal, State and Local Laws
The Long Beach Rocketry Team will comply with all federal, state and local laws on unmanned
rocket launches and motor handling.
The launch vehicle is classified as a class 2 high power rocket. The team will be launching with
the Rocketry Organization of California at Lucerne Dry Lake launch site and with the Friends of
Amateur Rocketry in Cantil, California. The Rocketry Organization of California ensures that all
rocket launches comply with FAA regulations and in strict compliance with all regulations of the
California State Fire Marshal. By launching with ROC, the team complies with Federal Aviation
20 Long Beach Rocketry | CDR 2017 - 2018
Regulations 14 CFR, Subchapter F, Part 101, Subpart C and fire prevention, NFPA 1127 “Code
for High Power Rocket Motors”.
To comply with Amateur Rockets, Code of Federal Regulation 27 Part 55: Commerce in
Explosives, the mentor will either acquire a federal low explosives user permit or find a person
that has a federal low explosives user permit to purchase or transport the motor to Huntsville,
Alabama.
4.8 Handling of Rocket Motors
Purchase and Storage
The Long Beach Rocketry Team recognizes that rocket motors will be purchased by the
NAR/TRA certified team mentor. The team mentor will be in possession of the motor until the
team's launch at the at the Lucerne Dry Lake Bed or the FAR (Friends of Amateur Rocketry)
sites. Prior to the launch all motors will be disassembled and stay in the original packaging.
Handling and Use
The rocket motors will only be handled by the team members while under the supervision of the
team mentor. The team mentor will oversee the preparation of the rocket motor for all launches.
Transportation
If the team mentor possesses a federal low explosive user permit he may purchases a rocket
motor in Huntsville, Alabama. He will coordinate the purchases and shipping of a rocket motor
to Huntsville, Alabama through a third party possessing a federal low explosives user permit, if
the team mentor does not have said permit.
4.9 Range Safety Regulations
All members of the Long Beach Rocketry Team will follow the range safety regulation as stated
in the Long Beach Rocketry Team Safety Agreement. The team members understand and will
abide by the following regulations:
• Range safety inspections of each rocket before it is flown. Each team shall comply with
the determination of the safety inspection or may be removed from the program.
• The Range Safety Officer has the final say on all rocket safety issues. Therefore, the
Range Safety Officer has the right to deny the launch of any rocket for safety reasons.
Any team that does not comply with the safety requirements will not be allowed to launch their
rocket
21 Long Beach Rocketry | CDR 2017 - 2018
4.10 Risk Assessment Codes
The following tables explain the Risk Assessment Codes (RAC) used to evaluate the hazards in
this report. The RAC is based on the severity of hazard and the probability of that hazard
occurring. The severity of the hazard is based on a number scale of 1 to 4, with 1 being the worst
classified as catastrophic and 4 being the lowest classified as negligible. Probability of the hazard
occurring is based on a letter scale of A to E, with A being the highest being classified as
frequent and E being the lowest classified as improbable. Once severity and probability are
defined the RAC can be used to define a risk level from high to minimal. Tables 4.2 – 4.5 more
clearly show the criteria of severity, probability and risk definition.
Table 4.2: Risk Assessment Code (RAC)
Probability
Severity
1
Catastrophic
2
Critical
3
Marginal
4
Negligible
A - Frequent 1A 2A 3A 4A
B - Probable 1B 2B 3B 4B
C - Occasional 1C 2C 3C 4C
D - Remote 1D 2D 3D 4D
E - Improbable 1E 2E 3E 4E
Table 4.3: Risk Definition
Level of Risk Level of Approval
High Risk Unacceptable. Documented approval from the RSO and NASA SL
officials.
Moderate Risk Undesirable. Document approval from team mentor, team lead and
safety officer.
Low Risk Acceptable. Document approval by lead of subsystem.
Minimal Risk Acceptable. Document approval not required.
22 Long Beach Rocketry | CDR 2017 - 2018
Table 4.4: Severity Definition
Description Personnel
Safety and
Health
Facility and
Hardware
Environmental Project Plan
1 - Catastrophic Loss of life or a
permanent
disabling injury.
Loss of facility,
systems or
associated
hardware.
Irreversible
severe
environmental
damage that
violates law and
regulation.
Delay of critical
components or
budget overruns
that results in
termination of
project
2 - Critical Severe injury or
occupational
related illness.
Major damage
to facilities,
systems, or
equipment.
Reversible
environmental
damage causing
a violation of
law or
regulations.
Delay of mission
critical
components or
budget overruns
that compromise
mission scope.
3 - Marginal Minor injury or
occupational
related illness.
Minor damage
to facilities,
system, or
equipment.
Mitigatable
environmental
damage without
violation of law
or regulation
where restoration
activities can be
accomplished.
Minor delays of
non – critical
components or
budget increase.
4 - Negligible First aid injury
or occupational
related illness.
Minimal
damage to
facility, systems,
or hardware.
Minimal
environmental
damage not
violating law or
regulation.
Minor delays of
non – critical
components.
Table 4.5: Probability Definition
Description Qualitative Definition Quantitative Definition
A - Frequent High Likelihood to occur immediately or
expected to be continuously experienced.
60% < Probability < 100%
B - Probable Likely to occur or expected to occur
frequently within time.
30% < Probability < 60%
C - Occasional Expected to occur several times or
occasionally within time.
10% < Probability < 30%
D - Remote Unlikely to occur, but can be reasonably
expected to occur at some point within
time.
5% < Probability < 10%
E - Improbable Very unlikely to occur and an occurrence
is not expected to be experienced within
time.
Probability < 5%
23 Long Beach Rocketry | CDR 2017 - 2018
4.11 Personnel Hazard Risk Assessment
Table 4.6 below indicates the possible hazard to personnel while working on the project including, risk from
working in the machine shop or any of the labs on campus.
Table 4.6: Personnel Hazard Risk Assessment
Hazard Cause Effect Pre-
RAC
Mitigation Verification Post-
RAC
Getting
caught in
machine
Hair not tied
back or loose-
fitting
clothing
Potential for
death or
serious injury
1C 1. When using
machinery proper
PPE will be required
1. Per section 4.4
“Facilities” students
will be supervised
when using shop
equipment
1D
Personnel
exposure to
fiberglass
dust, epoxy
fumes, or
paint fumes
Proper
precaution not
taken for
working with
fiberglass,
epoxy, or
paint
Personnel
injuries
including
irritation,
infections
and
pulmonary
disease
1C 1. When working
with hazardous
material personnel
will wear proper PPE
2. When working
with hazardous
material personnel
will work in a well-
ventilated area or
fume hood
1. Section 4.3 “Safety
Equipment” requires
that all team member’s
wear proper PPE in
accordance to the
MSDS
2. Section 4.3 “Safety
Equipment” states that
proper handling of
material will be
followed according to
the MSDS
1E
Explosion
while
handling
black powder
Unexpected
connection to
voltage
Personnel
injuries
including
blast injuries
and hearing
damage
1C 1. Black powder will
only be handled by
NAR/TRA personnel
2. Black powder will
only be handled in
small amounts at a
time.
1. Per section 4.13.1
“Ejection Charge
Integration Procedure”
the NAR/TRA will be
responsible for all
handling of black
powder
2. Per section 4.13.1
“Ejection Charge
Integration Procedure”
black powder will only
be handled in small
amounts
1E
24 Long Beach Rocketry | CDR 2017 - 2018
Cuts from
machines and
tools
Incorrect use
of machines
and tools
Personnel
injuries
requiring
medical
attention
1C 1. Team members
using machines and
tools will be trained
and supervised on
how to use those
tools.
1. Section 4.4
“Facilities” states that
students will be
supervised during the
use of machinery and
tools during
construction of the
rocket
1E
Electrical
shock from
electrical
components
Improper
handling of
electrical
system, such
as getting the
system wet or
improperly
connecting
electrical
components
Depending
on voltage of
source
injuries could
include slight
shock or
death
1C 1. When handling
high voltage
components, the
handler must be
grounded
1. Section 4.4
“Facilities” states that
students will be
supervised so that they
follow proper
operations when
working with shop
tools
1E
Personnel
exposure to
dust and
fumes from
cutting
carbon fiber
Safety
equipment not
used while
cutting carbon
fiber
Temporary
damage or
irritation to
lungs, skin or
eyes
2B 1. Team members
working with
hazardous materials
will do so in a well-
ventilated area like a
fume hood
2. When working
with hazardous
materials team
members will use
gloves and safety
respirator masks
1. Section 4.3 “Safety
Equipment” states that
proper handling of
material will be
followed according to
the MSDS
2. Section 4.3 “Safety
Equipment” requires
that all team member’s
wear proper PPE in
accordance to the
MSDS
2E
Skin contact
with
chemicals
Improper
handling of
materials
Mild to
severe burns
of the skin
2B 1. Nitrile gloves will
be worn while
working with any
chemicals
2. If skin contact
with chemical
happens effected area
will be flushed with
water for 20 minutes
or until help arrives
1. Section 4.3 “Safety
Equipment” requires
that all team member’s
wear proper PPE in
accordance to the
MSDS
2. As stated in section
4.5 “Injury and
Emergency” if the
injury is chemical
related first aid
precautions will be
followed according to
the chemicals MSDS
2E
25 Long Beach Rocketry | CDR 2017 - 2018
Fire outbreak
in the
machine
shop or lab
Short circuitry
of electrical
components
Mild to
severe burn
to personnel,
damage to
facility,
damage to
rocket
components
1D 1. Team member will
learn the proper way
to handle the
electrical
components so that a
fire is not created
2. If fire is
unpreventable fire
extinguishers will be
available to use in all
the rooms that we
work in
1. Section 4.4
“Facilities” states that
students will be
supervised so that they
follow proper
operations when
working with shop
tools
2. Section 4.5 “Injury
and Emergency” states
that fire extinguishers,
fire sprinklers and fire
alarm are available in
the lab and machine
shop
1E
Eye contact
with
chemicals
Improper
handling of
materials
Vision
impairment
and eye
irritation
2C 1. When working
with chemicals
appropriate eye
protection must be
worn
1. Section 4.3 “Safety
Equipment” requires
that proper PPE be
used in reference of
the MSDS
2E
Inhalation of
fumes
created form
soldering
Use of
soldering iron
can produce
toxic fumes
Breath in
toxic fumes
causing
irritation to
nose, throat
and
respiratory
organs
2C 1. Member will wear
protective face mask
when soldering
1. Section 4.4
“Facilities” states that
students will be
supervised so that they
follow proper
operations when
working with shop
tools
2E
Flying debris
from
machining
operation
Incorrect use
of machine
Irritation of
eyes or skin
2C 1. Long pants, closed
toed shoes and eye
protection will be
worn while in the
machine shop
1. Per section 4.4
“Facilities” students
will be supervised by
university personnel
that are aware of the
shop safety practices
such as the proper PPE
for the shop
2E
26 Long Beach Rocketry | CDR 2017 - 2018
Personnel
injured from
contact with
hot surfaces
while
soldering
components.
Accidental
contact with
hot soldering
iron that was
left on
Personnel
sustains
burns
3B 1. While using heat
producing tools they
will be turned off
when not in use
2. Members will be
properly trained in
how to safely handle
a soldering iron.
1. Section 4.4
“Facilities” states that
students will be
supervised so that they
follow proper
operations when
working with shop
tools
2. Per section 4.4
“Facilities” students
will be supervised by
university personnel
that are aware of the
shop safety practices
3D
Contact with
falling parts
or tools
Improper
storage of
parts and tools
Personnel
injuries
3C 1. Team members
will wear proper PPE
such as closed toed
shoes and long pants
while working in the
lab or machine shop
1. Per section 4.4
“Facilities” students
will be supervised by
university personnel
that are aware of the
shop safety practices
such as the proper PPE
for the shop
3E
27 Long Beach Rocketry | CDR 2017 - 2018
4.12 Failure Mode Hazard Risk Assessment
Tables 4.7 – 4.10 below indicate the possible hazards to all the subsystems on the rocket. This includes structure
and propulsion, recovery, rover deployment (RDM) and rover.
Table 4.7: Failure Mode Hazard Risk Assessment – Structure and Propulsion
Structure and Propulsion Risk Analysis
Hazard Cause Effect Pre-
RAC
Mitigation Verification Post-
RAC
Motor fails
to ignite
Delayed
Ignition,
Faulty motor,
damage
during
transportation
Rocket could
launch
unexpectedly
or not at all
1C 1. Follow proper
procedure setting up
igniter
2. Motors will be
purchased from
vendors with a good
reputation
1. See section 4.13.4
“Igniter Instillation”
for proper procedure
to follow when setting
up igniter
2. Section 4.8
“Handling of Rocket
Motors” states that
purchasing of the
motor will done by
NAR/TRA team
mentor who will
purchase from
reputable vendor
1E
Motor
explodes on
launch pad
Faulty motor,
improper
motor packing
Rocket will be
highly
damaged,
damage to
motor casing,
potential
injury to
personnel
1C 1. Proper motor
assembly
2. Motor preparation
overseen by
NAR/TRA team
mentor
1. Procedure available
in section 4.13.2
“Motor Preparation”
2. Section 4.8
“Handling of Rocket
Motors” states that
preparation of the
motor will done by
NAR/TRA team
mentor
1E
Rocket
Velocity not
high enough
leaving the
launch pad
Rocket to
heavy. Thrust
of the motor
is not large
enough
Launch is
unstable
1C 1. Run simulations to
verify that the motor
we have selected will
provide a sufficient
velocity leaving the
launch pad
1. In the “Motor
Selection and
Alternatives” section
you can see that the
Cesaroni L1350 CS-P
provides a sufficient
velocity off the rod
1E
28 Long Beach Rocketry | CDR 2017 - 2018
Motor
centering
ring fails
Adhesive not
properly
applied to the
centering
rings,
centering
rings material
not strong
enough
Motor to
launch
through the
rocket
1C 1. That there are no
cracks to the
adhesive before
launch
2. Ensure that
centering rings are
made of strong
enough material
1. Section 4.13.3
“Setup on Launcher”
requires that the rocket
is check for any
cracking of the
adhesive before it is
launched
2. In the launch
vehicle “Propulsion
Bay” section it can be
seen that the centering
will be made from
6061 Aluminum which
will be testing during
the full-scale launch
1E
Buckling of
airframe
during flight
Structure
cannot
properly
handle stress
of flight
Rocket will
lose control
and become
unstable and
unpredictable
2B 1. Use proper
material to ensure
that it can handle the
stress of the launch
and flight
1. As stated in the
launch vehicle section
“Material Selection”
the material that will
be used for the rocket
is fiberglass which
will be able to handle
all the forces of the
flight
2E
Tail fins
shear off
during flight
Fins not
properly
secured to
airframe
Rocket takes
unpredictable
flight path and
becomes
unstable
2C
1. Ensure that
adhesive used to
secure fins is strong
enough to handle the
force of flight
2. Check the
adhesive for cracks
before launch
1. As seen in the
launch vehicle “Fin
Analysis” section the
epoxy being used has a
allowable tensile stress
of 3960psi, and the
analysis results show
that the fins will
remain securely in
position throughout
the entire flight
2. In 4.13.3 “Launcher
Setup Procedure” the
adhesive will be
checked before the
rocket is launched
2E
29 Long Beach Rocketry | CDR 2017 - 2018
Structure
Damage
during
transport
Improper
storage during
transportation
Rocket will be
unstable and
unpredictable
2D 1. Proper handling
and storage overseen
by NAR/TRA team
mentor
1. Section 4.8
“Handling of Rocket
Motors” states that all
handling and storage
of the motor will be
done by NAR/TRA
team mentor
2E
Fins not
properly
aligned
Fins not
assembled
correctly
Rocket
becomes
unstable and
spins
uncontrollably
2D 1. Proper procedure
is followed when
assembling the fins
1. Safety officer will
oversee all aspects of
construction of the
rocket
2E
Improper
assemble of
motor
Proper motor
preparation or
procedure not
followed
Rocket flight
will be
unstable,
target altitude
not reached,
damage to
rocket
2C 1. NAR/TRA team
mentor will oversee
motor preparation
2. Follow procedure
to preparing motor
1. Section 4.8
“Handling of Rocket
Motors” states that
handling of the motor
will done by
NAR/TRA team
mentor
2. Procedure available
in section 4.13.2
“Motor Preparation”
1E
30 Long Beach Rocketry | CDR 2017 - 2018
Table 4.8: Failure Mode Hazard Risk Assessment – Recovery
Recovery Risk Analysis
Hazard Cause Effect Pre-
RAC
Mitigation Verification Post-
RAC
Parachute
does not
deploy
Parachute gets
tangled
around rocket,
Rocket does
not split open
Rocket will
fall to the
ground at
high velocity
and become
damaged
upon impact
1B 1. Parachute will be
properly integrated
into rocket to reduce
risk of getting
tangled
2. Team mentor will
ensure proper amount
of black powder
charge to split rocket
open
1. Section 4.13.1
“Recovery
Preparation” the
parachute integration
procedure is listed step
by step with warnings
listed at critical steps
2. In the Recovery
section it discusses the
proper amount of
black powder that
should be used to
separate the rocket
1D
Rocket fails
to separate
Black powder
fails to ignite,
Black powder
fails to break
shear pins
Parachute
will not
deploy, and
rocket will
fall to the
ground at
high velocity
and become
damaged
upon impact
1B 1. Black powder
amount will be
picked properly to
ensure that rocket
separated
2. Ensure that black
powder is integrated
properly so that it
does not leak and
have weaker ejection
1. In the Recovery
section it discusses the
proper amount of
black powder that
should be used to
separate the rocket
2. In “Recovery
Preparation” section
4.13.1 of launch
procedures it lists the
steps to integrating the
black powder in a way
that none of it will leak
1D
31 Long Beach Rocketry | CDR 2017 - 2018
Parachute
has rip or
tear
Parachute gets
rip while
packaging in
rocket,
Parachute gets
ripped while
deploying
Rocket will
descend
quickly and
become
damaged
upon impact
1C 1. Parachute will be
carefully inspected
before it is packaged
2.Team members
will be careful during
packaging of
parachute
1. The section 4.13.1
“Recovery
Preparation” lists steps
for parachute
integration stating to
inspect the parachute
before putting it in the
rocket
2. The section 4.13.1
“Recovery
Preparation” lists steps
for parachute
integration with
warning of things to be
careful with during
packaging of the
parachute
1E
Altimeter
failure
Faultily
altimeter,
Altimeter gets
damaged
during launch
Parachute
will not
deploy, and
rocket will
fall to the
ground at
high velocity
and become
damaged
upon impact
1C 1. Use more than one
altimeter for
redundancy
1. The avionic section
goes into details about
the two altimeters
being used
1E
Avionics
malfunction
Low power
supply or
incorrect
assemble of
avionics
Early or no
deployment
of parachute
causing
rocket to
descend
quickly and
become
damaged
upon impact
1C 1. Two avionics
systems will be used
for redundancy to
reduce the chance of
malfunction
2. Testing will be
done before launch to
ensure avionics is
functioning properly
1. In the avionics part
of the paper discusses
in detail the design of
the avionics system
2. The section 4.13.3
“Setup on Launcher”
ensures that the
avionics system
altimeter beeps are
checked and give the
correct results before
the rocket is launched
1E
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Rocket
separates
from
recovery
system
Parachute
disconnects
from U-bolt
Rocket will
descend
quickly and
become
damaged
upon impact
1C 1. Parachute cables
and the U-bolts are
designed to handle
large loads
2. Parachute will be
properly connected to
the U-bolt
1. The recovery
section discusses the
U-bolt selection and
how much force it can
handle
2. Section 4.13.1
“Recovery
Preparation” goes over
procedure to properly
connect the parachute
to the U-bolt
1E
Rocket
descends to
quickly
Parachute not
sized properly
Rocket will
descend
quickly and
become
damaged
upon impact
2C 1. Select parachutes
based on predicted
weight of rocket to
ensure it descends at
a reasonable speed
1. In the recovery
section you can find
the calculations that
were done to
determine the size of
the parachute
2E
33 Long Beach Rocketry | CDR 2017 - 2018
Table 4.9: Failure Mode Hazard Risk Assessment – Rover
Rover Risk Analysis
Hazard Cause Effect Pre-
RAC
Mitigation Verification Post-
RAC
Rover
damaged on
landing
Faster than
normal
landing,
Payload not
secure in
place
Rover
becomes
damaged and
inoperable
2C 1. Testing will be
done to verify that
the rover can handle
all forces of landing
1. Section 6.6.1.1
“Impact Testing”
discusses testing and
results that have been
done to test the rover’s
ability to handle an
impact
2E
Rover
damaged
during flight
Payload not
secure in
place, rover
material not
strong enough
Rover
becomes
damaged and
inoperable
2C 1. Pick a material
that is strong enough
and test it during the
subscale launch
1. In section 6.2.6
“Material” the material
for the rover is ABS
plastic even though
this material has a
lower tensile strength
then carbon fiber and
aluminum it proved to
be strong enough to
survive to force of
flight based on the
subscale launch
2E
Rover flips
over
Rover not
designed to
handle the
terrain
Rover
becomes
stuck and
unable to
make
distance
requirement
3B 1. Design rover in a
way so that is
incapable of flipping
over
1. Section 6.2 “Rover
Design” states the
rover will use the
triangular design
which means that it
will be capable of
driving from any
orientation
3E
34 Long Beach Rocketry | CDR 2017 - 2018
Rover gets
stuck on
terrain
Rover not
designed to
handle various
objects that
could be in
the terrain
Rover will be
stuck and
potential
damage itself
making it
unable to
make the
distance
requirement
3B 1. Motor selection
must be selected
properly so wheels
are strong enough
2. Wheels designed
to handle the terrain
properly
2.Verify through
testing that the rover
can handle the terrain
of the launch field
1. Section 6.3.1
“Design and
Components” talks
about the motor
selection and how the
team will be focusing
on high torque over
high power to combat
the terrain
2. In section 6.2.5
“Wheel and Tire
Design” the team has
picked to use
expandable wheels due
to the non-expandable
wheels in the subscale
launch being
ineffective
2. Section 6.6.1.3
“Environmental
Condition” discusses
the test planned for the
rover for various
environments
3E
Rover veers
of course
Electronics
failure
Rover will
spin in circles
and not get
the distance
requirement
3D 1. Provide sensors in
rover that will
analyze rovers
motion
1. Section 6.3.1
“Design and
Components” talks
about the IMU that
will be aboard the
rover provide
gyroscopic and
acceleration data to
detect when data is off
course
3E
Rover battery
dies before it
can deploy
Battery
capacity for
rover not
large enough
Rover will
not deploy
and make the
distance
requirement
4B 1. Battery needs to
have enough capacity
to handle the time of
flight and any idle
time once the rocket
is assembled
1. As seen in section
6.3.1 “Design and
Components” it was
calculated that we
need a battery with a
capacity of 1642mAh,
so the selected power
source was an 11.1V
battery with 1800mAh
4E
35 Long Beach Rocketry | CDR 2017 - 2018
Table 4.10: Failure Mode Hazard Risk Assessment – Rover Deployment (RDM)
Rover Deployment (RDM) Risk Analysis
Hazard Cause Effect Pre-
RAC
Mitigation Verification Post-
RAC
RDM
deploys
during flight
Programing or
electronic
failure
Nose cone
will separate
from rocket
during flight
causing
rocket to
become
unstable and
lose control
2B 1. Verify through
testing that the
electronics will only
deploy when they are
wanted to
1. In the RDM part of
the paper the section
“Subscale Prototype”
it dives into the details
of the electronics of
the RDM system that
was successful at the
subscale launch
2E
RDM
deployment
damages the
rover
Poor design of
RDM system
Rover will
become
damaged and
not function
2C 1. Properly design
the RDM system and
test it with a model to
ensure that it works
1. In the RDM part of
the paper the section
“Subscale Prototype”
goes into details of the
design that was used
for the subscale launch
and any of the
problems that it had
2E
RDM
becomes
damaged
during
launch or
flight
Materials for
RDM system
cannot handle
the force of
launch or
flight
RDM will
become
damaged and
rover will not
deploy
2C 1. Ensure that the
materials chosen for
the RDM system is
capable of handling
of force or launch
and flight
1. In section RDM
section “Subscale
Prototype” the
material of RDM
system are all listed
and was able to
survive all the forces
of the subscale launch
2E
36 Long Beach Rocketry | CDR 2017 - 2018
RDM does
not deploy
when
activated
Programing
failure or dead
battery
Rover will
not deploy
3C 1. Testing electronics
to ensure deployment
upon activation
2. Easy access to
battery on/off switch
of receiver battery
1. In the RDM part of
the paper the section
“Subscale Prototype”
it dives into the details
of the electronics of
the RDM system that
was successful at the
subscale launch
2. As discussed in
RDM section
“Subscale Prototype”
the RDM team found a
way to be able to turn
on and off the battery
easily so that it does
not die before rocket
assembly
3E
Rover gets
caught on
rods of RDM
system
RDM not
designed
properly to
deploy rover
Rover will
not deploy
3C 1. RDM system will
be designed so rover
does not get caught
on rods
2. Test the RDM
design to ensure that
it works
1. In the RDM section
“RDM Changes Since
PDR” it discusses why
the rover got caught on
the subscale launch
and how the design
was changed to
prevent it from
happening again
2. Section “RDM
Testing Plan” of the
RDM section
discusses the test
planned for checking
the design
3E
37 Long Beach Rocketry | CDR 2017 - 2018
RDM motor
not strong
enough to
break shear
pins
RDM system
motor not
selected
properly
RDM system
will not work
and rover will
not deploy
3D 1. Pick a motor with
enough force to
break the shear pins
2.Test RDM system
to ensure they can
break the shear pins
1. In the RDM section
“RDM Changes since
PDR” it discusses the
issues that the motor
had with breaking the
shear pins during the
subscale launch and
discusses solution to
this problem
2. Section “RDM
Testing Plan” of the
RDM section
discusses the test
planned for breaking
the shear pins
3E
38 Long Beach Rocketry | CDR 2017 - 2018
4.13 Environmental Hazard Risk Assessment
Tables 4.11 and 4.12 below indicate the possible hazards to and from the environment for this project. This
includes the environmental hazards on rocket and the hazards that the rocket has on the environment.
Table 4.11: Environmental Risk Assessment – Environment Impact on Rocket
Environment Impact on Rocket
Hazard Cause Effect Pre-
RAC
Mitigation Verification Post-
RAC
Black
powder
exposure to
humid
weather
Rocket left in
humid
weather for an
extended
period
Black powder
becomes
damp and
does not
deploy
1B 1. Black powder will
be store such that
they are not subject
to humid weather
2. Launches are
planned to avoid
humid weather
1. Section 4.3 “Safety
Equipment” states that
proper storage of
material will be
followed according to
the MSDS
2. FAR site is in the
desert where humidity
is usually low and will
be verified in section
4.13.3 “Setup on
Launcher” procedure
to check for humidity
before launch
1D
Extended UV
exposure
Rocket left in
the sun for an
extended
period
Weakening
of adhesives,
potential
damage to
electrical
components
2A 1. Rocket will not be
exposed to the sun
for more than it what
is necessary
2. A shelter should
be used if the rocket
needs to be worked
on in the sun
1. The rocket will only
be in the sun to move
the rocket to the
launch rail and the
short duration that it
sits on the rail
2. FAR launch site
provides shaded areas
for assembly of the
rocket
2D
Trees in
landing sight
of rocket
Launch site
has trees in
drift distance
of rocket
Damage to
rocket or
parachute and
possible
unable to
retrieve
rocket
2B 1. Ensure that trees
are out of range of
the drift distance so
that the rocket does
not hit the trees or
get stuck in the trees.
1. Section 4.13.3
“Setup on Launcher”
the procedure will
ensure that launch site
is check for trees
before the rocket is
launched
2E
39 Long Beach Rocketry | CDR 2017 - 2018
High wind
speeds
during
launch
Rocket
launched in
conditions of
high winds
Rocket will
become
unstable
1D 1. Wind speed will
be measured prior to
rocket launch
2. Rocket will not
launch with wind
speeds 20 mph or
higher
1. As stated in section
4.13.3 “Setup on
Launcher” wind speed
will be measured
before the launch
occurs.
2. Section 4.13.3
“Setup on Launcher”
list a warning in the
procedure to not
launch if wind speeds
are greater than 20
mph
1E
Battery
exposed to
low
temperatures
during
launch or
storage
Rocket left in
in low
temperature
for an
extended
period
Increase the
internal
resistance
and lower the
capacity,
quicker
discharge of
battery
2C 1. Batteries shall be
tested for voltage and
discharge rate prior
to launch
2. Batteries will be
stored in ideal
condition
1. The procedure in
section 4.13.3 “Setup
on Launcher” will
ensure that all batteries
have been checked
prior to launch
2. Section 4.3 “Safety
Equipment” states that
proper storage of
material will be
followed according to
the MSDS
2E
Rain during
launch
Rocket
launched in
conditions of
rain
Electronics of
rocket
become
damaged
2C 1. Rocket will not
launch in raining
conditions
2. Keep electrical
equipment away
from the rain to avoid
water damage
1. Launch days will be
plan according to the
forecast report for a
day with no rain and
will be checked up to
the day we launch
2. Team member will
use caution to ensure
that no components get
damaged from the rain
2E
Excessive
cloud
coverage
during
launch
Rocket
launched in
conditions of
excessive
clouds
Rocket is not
visible so
there no way
to tell if
rocket launch
is going well
3C 1. Rocket will not be
launched in condition
of excessive clouds
1. Section 4.13.3
“Setup on Launcher”
procedure states to
check the sky before
launch to look for
clouds and abort
launch if conditions
are excessively cloudy
3E
40 Long Beach Rocketry | CDR 2017 - 2018
Table 4.12: Environmental Risk Assessment – Rocket Impact on Environment
Rocket Impact on Environment
Hazard Cause Effect Pre-
RAC
Mitigation Verification Post-
RAC
Rocket
recovery
system failure
Altimeter
failure or
parachute
failure
Rocket
becomes free
falling object
upon impact
body may
shatter
leaving the
material in
the
environment
1C 1. Parachute will be
properly integrated
into rocket to reduce
risk of getting
tangled
2. Team mentor will
ensure proper
amount of black
powder charge to
split rocket open
3. Use more than one
altimeter for
redundancy
1. Section 4.13.1
“Recovery
Preparation” the
parachute integration
procedure is listed step
by step with warnings
listed at critical steps
2. In the Recovery
section it discusses the
proper amount of
black powder that
should be used to
separate the rocket
3. The avionic section
goes into details about
the two altimeters
being used
1E
Motor exhaust
cause fire on
ground
Exhaust
allowed to
flow onto the
surrounding
ground
Fire can
occur at the
grounds
around the
launch rails
2B 1. Blast deflector
will be used, and
grass will be cleared
from launch pad
1. As stated in section
4.6 “NAR Safety
Code Compliance” the
team has agreed to
follow this
requirement listed in
the NAR Safety Codes
2E
Battery
rupture/damage
spreads
hazardous
chemical
Impact of
launch or
landing
damages the
batteries
Batteries
burst and
spread
hazardous
chemicals on
to launch site
damaging the
environment
2B 1. Batteries will be
tightly secure to
reduce chance of
movement
2. Batteries will be
enclosed to prevent
punctures
1. All subsystems that
use a battery have
designed them in such
a way that they are
tightly secure
2. All batteries will be
placed strategically
inside the rocket so
that nothing can
puncture them
2E
41 Long Beach Rocketry | CDR 2017 - 2018
Battery
explosion
breaks rocket
and scatters
parts
Batteries may
explode
unpredictably
if not charged
properly
Explosion
cause parts
of the
destroyed
rocket to be
spread all
over the
launch field
2C 1. Team members
will follow correct
procedure when
charging batteries
1. Section 4.3 “Safety
Equipment” states that
proper handling of
material will be
followed according to
the MSDS
2E
Harm to
environment
from litter
Team
members not
properly
disposing of
trash
Danger to
wildlife and
environment
3D 1. Team members
will carefully check
the launch site before
leaving to ensure that
no trash is left at the
site
2. Deployment of
trash bags at launch
site
1. In the section 4.13.5
“Post-flight
Inspection” the last
step in the procedure
instructs the team to
clear all trash before
they leave
2. Trash bags will be
pack and brought to
the launch site and be
one of the first thing
that the team sets up
3E
42 Long Beach Rocketry | CDR 2017 - 2018
4.14 Launch Operation Procedures
4.14.1 Recovery Preparation To be checked and signed by Recover Lead and Safety Officer
1. _______________________________ 2. _______________________________
Required Equipment Checklist:
Recovery Avionics Bay
Electronic matches (x4)
Masking Tape
Scissors
Black powder
Wadding
Spoon
Paper
Main Parachute
Drogue Parachute
Main Deployment bag
Drogue Blanket
Shock Cord (2x)
Shock Cord Protective sleeves (2x)
Quick Links (6x)
U - Bolts (4x)
Precision flathead screwdriver
Digital Scale
E-matches
43 Long Beach Rocketry | CDR 2017 - 2018
Required Personal Protective Equipment (PPE):
Nitrile gloves
Safety glasses
Ejection Charge Integration Procedure:
WARNING Black powder must only be handled by a Tripoli or NAR certified member.
1. Measure needed black powder for PRIMARY ejection charge for drogue and main
a. Primary main ejection charge: TBD
b. Primary drogue ejection charge: TBD
WARNING Work with small amounts of black powder in case of accidental ejection.
2. Cut off excess wire off the E-match as needed.
3. Place the E-match inside the PVC cap. Make sure the E-match is lying flat on the
bottom of the cap
4. Place the specific measured amount of black powder into the PVC cap.
5. Place wadding on top of the black powder and compress the wadding until the PVC
cap is full.
6. Place tape over the top and around the PVC cap so the wadding and black powder and
wadding does not fall out.
WARNING If black powder leaks, the resulting ejection charge may weaken or fail to
detonate. This could cause fatal failure in the recovery system and personnel injury.
7. Mark the PVC "P" and the amount of black powder in the cap
8. Measure needed black powder for SECONDARY ejection charge for drogue and main
a. Secondary main ejection charge: TBD
b. Secondary drogue ejection charge: TBD
9. Repeat step 2-6
10. Mark the PVC "S" and the amount of black powder in the cap.
11. Wait for avionics assembly to be complete and check that the altimeters are off.
12. Connect each the E-match to its corresponding terminal blocks.
13. Check that all connections are correct and secure.
Parachute Integration Procedure:
1. Verify all cell phones are off at launch site.
2. Check for any damage including burns, cut, fraying, or any other visible damage for
the following parts:
a. Parachutes
b. Shock Cords
c. Blankets
d. Deployment Bag
e. Quick Links
f. Eyebolts
WARNING If damage is identified, abort launch.
3. Lay drogue and main parachute canopy flat on the ground.
44 Long Beach Rocketry | CDR 2017 - 2018
4. Check line for any entanglement, if entanglement is found, untangles the line as
needed.
WARNING If entanglement is left it can lead to failure of the recovery system. Leading
to damage to rocket or personnel.
5. Link all shock cord segments with quick links.
6. Fold up both parachutes following the procedures form Fruity Chutes.
7. Connect the main parachute to the shock cord connecting the avionics to the motor
section of the rocket.
8. Attach the drogue parachute and harness into the deployment bag.
9. Slide the main parachute and harness into the deployment bag.
10. Connect one quick link to the U-bolt that is connected to the payload.
11. Slide the deployment bag with the main parachute inside into the main parachute bay.
12. Connect the other end of the shock cord to the U-bolt connected to the avionics bay.
13. Slide the avionics into the main parachute bay.
14. Screw in the shear pin into shear pin holes for the main parachute bay.
15. Connect drogue parachute to harness and fold flame shield blanket over drogue.
16. Connect one end of the harness to the eye U-bolts connected to the motor.
17. Slide the covered drogue parachute into the drogue bay.
18. Connect the other end of the harness to the U-bolts connected to the avionics bay.
19. Slide avionics bay into the drogue bay.
20. Screw in shear pin into shear pin holes for drogue bay.
WARNING Failure to properly pack parachute can cause recovery system failure.
4.14.2 Motor Preparation To be checked and signed by NAR/TAR personnel and Safety Officer
1. _______________________________ 2. _______________________________
Required Equipment Checklist:
Grease
Cesaroni L1350 CS-P Motor
Forward Extended Plugged Tapped Closure
Motor Casing
Aft Closure
54mm Retainer
End Cap
Required Personal Protective Equipment (PPE):
Nitrile gloves
Safety glasses
Motor Assembly Procedure:
WARNING Motor Assembly must only be handled by a Tripoli or NAR certified member.
Improper assemble of the motor may lead to motor failure and damage to rocket or personnel.
1. The Cesaroni L1350 CS-P motor is taken out of its antistatic bag.
45 Long Beach Rocketry | CDR 2017 - 2018
2. Inspect the motor to ensure that no damage has occurred during transportation.
WARNING If damage to the motor is found abort launch.
3. The instructions provided with the motor is followed to prepare the motor to be placed
into the motor casing.
4. The forward extended plugged tapped closure and aft closure is screwed on the motor
casing.
5. Motor casing is slid into the propulsion bay of the rocket.
6. 54mm retainer cap is screwed onto the end of the propulsion bay, securing the motor
casing to the propulsion bay.
7. The eyebolt along with its bulkhead is screwed on the forward extended plugged
tapped closure.
4.14.3 Setup on Launcher To be checked and signed by Team Lead and Safety Officer
1. _______________________________ 2. _______________________________
Required Equipment Checklist:
Pen or pencil
Level 2 Certification card.
Rocket
Featherweight Screw Switch (Arming Switch) (2x)
Required Personal Protective Equipment (PPE):
Closed toe shoes
Pants
Launcher Setup Procedure:
1. Check all batteries voltage and discharge rate. Verify that all batteries have the
expected voltage and discharge rate.
2. Verify that none of the adhesive holding the fins or centering rings to the rocket have
any cracks.
3. Find the center of gravity of the launch vehicle and calculate static stability margin.
Verify that the stability is within 0.25cal of the expected stability margin.
WARNING Failure to measure the stability margin could result in unstable flight.
4. Check wind speed before launch. Verify that it is below 20 mph.
46 Long Beach Rocketry | CDR 2017 - 2018
WARNING Abort launch if wind speeds are greater than 20mph. If rocket is launch in
wind speed greater than 20 mph rocket could become unstable and become a danger to
the environment and everyone at the launch site.
5. Check to see that they sky is clear of clouds and that the weather has low humidity.
Launch will be aborted if it conditions are excessively cloudy or high humidity.
6. Confirm that there are no trees in the radius of the calculated drift distance of the
rocket.
WARNING If rocket is launch with trees in the landing range it could potential get stuck
in the tree and be unrecoverable or lead to damage of the rocket.
7. Verify that permission has been granted by the RSO to launch.
8. Slide the rocket onto the launch pad rail and raise to vertical.
9. Arm all electronics. Check for correct LED readout, beeping pattern, etc.
10. Before leaving launch pad area, double check that all electronics are still operating
correctly.
11. Clear and Leave the launch pad area.
WARNING Failure to ensure NAR minimum distance is observed before launch may
result in personnel injury.
4.14.4 Igniter Installation To be checked and signed by NAR/TRA Personnel
1. _______________________________
Required Personal Protective Equipment (PPE):
Closed toe shoes
Pants
Igniter Installation Procedure:
WARNING Igniter installation must only be handled by a Tripoli or NAR certified member.
1. Insert igniter into the rocket motor.
2. Attach the leads that connect to the igniter to the ignition trigger.
3. Ensure that the ignition system is wired to the power source.
WARNING If the leads or ignition system is wired incorrectly motor will not ignite.
4.13.5 Post-flight Inspection To be checked and signed by Team Lead and Safety Officer
1. _______________________________ 2. _______________________________
Required Personal Protective Equipment (PPE):
Closed toe shoes
47 Long Beach Rocketry | CDR 2017 - 2018
Pants
Post-flight Inspection Procedure:
1. Disarm the rocket recovery system.
WARNING If ejection charges are unblown use caution. Disarm recovery altimeters
immediately. Failure to do so may result in personnel injury.
2. Take photos of how the rocket landed recording any damage.
3. Inspect the parachutes and the shock cords for damage.
4. Return the launch to the RSO for the official altimeter reading.
5. Remove motor from the motor casing and dispose of it.
6. Clean all parts of the motor casing.
7. Disassemble the rocket looking for any signs of damage or fatigue.
8. Clear launch site of all trash and pack up to leave.
WARNING If trash site is left at launch site it could cause danger to the environment or
wildlife.
4.15 Troubleshooting
Table 4.13: Troubleshooting
Problem Solution
Mass of rocket after assembling is different
than planned mass of rocket.
Disassemble rocket and ensure that all
components that are supposed to be inside the
rocket are there. If nothing is missing
reassemble rocket and weight it again with a
different scale.
Shear pins not aligning with the original holes
drilled for the shear pins.
Remove parachute from rocket then realign
the frame of the rocket with the coupler and
remark new holes for the shear pins then drill
those holes then try to assemble again.
Parachutes do not fit in the rocket sections. Refold the parachute more tightly then again
attempt to package in the rocket.
Ignition system does not activate when
triggered. WARNING Wait 30 seconds to confirm it is
not a delayed response. Approach ignition
system and double check wiring and make
sure leads are attached properly then try
again.
Recovery altimeters do not return the expect
beeping pattern informing that they are ready
to launch.
WARNING Turn off altimeter to prevent
charges form igniting. Return rocket to
preparation area. Disassemble parts of the
rocket so that the avionics bay can be
accessed. Then inspect the wires for missing
connection and reattach any loose wires.
48 Long Beach Rocketry | CDR 2017 - 2018
Section 5: Launch Vehicle Criteria
5.1 Mission Statement
The Long Beach Rocketry 2017-2018 team mission is to successfully build, test, and fly a launch
vehicle carrying a rover which will deploy upon landing. The airframe of the launch vehicle must
safely house the interior components of the launch vehicle throughout the duration of the flight,
and withstand the forces induced during takeoff and ascent. In order to achieve an apogee of
5280 feet, the launch vehicle will be optimized to minimize mass. Furthermore, the launch
vehicle must successfully deploy its recovery parachutes and land without causing any damage
or safety hazards.
5.2 Mission Success Criteria
1. The launch vehicle shall be successfully departed from the launch rail.
2. The launch vehicle shall carry a payload up to an apogee of 5280 feet ±100 feet.
3. All the recovery events shall successfully occur at the programmed altitude.
4. The launch vehicle shall have a stable takeoff and ascent.
5. The kinetic energy of any independent section upon landing shall be below 75 ft-lb.
6. The launch vehicle shall be successfully recovered in a reusable condition.
5.3 Launch Vehicle Overview
Figure 5.1: Full-Scale Launch Vehicle CAD
The launch vehicle will have a 6-inch diameter airframe because additional space is needed for
this year’s scientific payload. To achieve an apogee of 5280 feet, the main goal of the launch
vehicle this year will be to minimize mass while optimizing for maximum efficiency. The total
length of the launch vehicle is 103 inches.
49 Long Beach Rocketry | CDR 2017 - 2018
Table 5.1: Launch Vehicle Section Lengths and Weight
Section Length (in) Weight (lb.)
Nose Cone 24 3
Payload Bay 16 11.82
Avionics Bay (AV) 13 4.77
Propulsion Bay 42 19.17
Total Length and Weight 103 38.7
Using estimated payload masses and known material densities, the assembled weight of the
launch vehicle will be approximately 38.7 lb. with the motor, or 30.15 lb. without the motor. The
launch vehicle has a center of pressure at 80.675 inches and a center of gravity at 65.258 inches
measured from the nose cone, which yield a static stability margin at rail exit of 2.5 calibers. An
OpenRocket model was created to verify the locations of the center of gravity, center of pressure,
and apogee of the full-scale launch vehicle. The specifications of the OpenRocket Simulation of
the launch vehicle are shown in Table 5.2.
Table 5.2: Launch Vehicle Flight Specifications
Specifications of the launch vehicle Numerical Value
Center of Gravity (in. from nose cone) 65.258
Center of Gravity after the motor burnout 61.299
Center of Pressure (in. from nose cone) 80.675
Static Stability Margin (cal) 2.5
Static Stability Margin after the motor burnout (cal) 3.01
Rail exit velocity (ft./s) 75.9
Max acceleration (ft./s^2) 283
Predicted Apogee (ft.) 5295
The launch vehicle will have three sections to allow for dual parachute deployments. Deploying
the drogue parachute at apogee and the main parachute at 500 ft. will significantly reduce drift.
Shown in the image below is the nose cone/payload section, the recovery/avionics section, and
the propulsion section, shown respectively from left to right in Figure 5.3.
50 Long Beach Rocketry | CDR 2017 - 2018
Figure 5.2: Full Scale CAD Model Split into Sections
The nose cone/payload section contains communication tracking electronics, the rover
deployment mechanism, the DORITOS and the drogue parachute. The AV bay contains recovery
electronics that control black powder ejections. Lastly, the propulsion bay contains the four
carbon fiber fins, the motor tube, and the main parachute.
A coupler tube will be used to connect separation points of the launch vehicle, such as the nose
cone/payload section to the AV section and the AV section to the propulsion bay section. This
coupler tube is 13 inches in length, with 6 inches going into each side of the connection. The
coupler tube will have a diameter of 6 inches, slightly smaller than the airframe to allow a tight
fit.
5.4 Launch Vehicle Design
5.4.1 Launch Vehicle Material Selection
The airframe, coupler, and nose cone of the launch vehicle are constructed using G12 fiberglass.
G12 was selected for its high strength and non-conductive properties. All bulkheads and
centering rings, with exception of the RDM bulkhead, are constructed out of 6061 Aluminum
because it has good mechanical properties and is low-cost. The RDM bulkhead is constructed of
3D printed ABS material. Lastly, the fins are constructed of carbon fiber because carbon fiber is
a high strength material with relatively low weight.
5.4.2 Nose cone
Figure 5.3: CAD of the nose cone
51 Long Beach Rocketry | CDR 2017 - 2018
The chosen nose cone design for the LBR 2017-2018 launch vehicle is the Ogive 4:1. The nose
cone weighs approximately 3 lb, has a length of 24 inches, a shoulder length of 3.5 inches, and
an outer diameter of 6.17 inches. This nose cone was selected based on the successful use of this
design in the previous years of NSL competition, and for its combination of low mass and
relatively low efficient of drag when compared to other nose cone profiles. It is necessary that
the nose cone will be secured to the launch vehicle via three 4-40 nylon screws to ensure no
separation occurs during flight.
Since the full-scale launch vehicle will be flying below the speed of sound (Mach 0.6), the nose
cone pressure drag is essentially zero (𝐶𝐷 = 0.04) and most of the drag comes from the friction
drag. The friction drag is dependent upon the surface roughness, the whetted area, and
discontinuities in sections.
Figure 5.4: Ogive Nose Cone Flow Simulation results using SolidWorks
SolidWorks flow simulations with various parameters were performed on both the nose cone and
the launch vehicle. A pressure surface plot for each simulation was created to view the different
pressure values on each part of the nose cone and launch vehicle. Each simulation computed the
drag force which was then used to calculate the drag coefficient.
Table 5.3: Coefficient of Drag Data from the Simulation
Components Coefficient of Drag
Nose Cone 0.04
Full Scale Launch Vehicle 0.4
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Figure 5.5: Full Scale Launch Vehicle Simulation Results
The secondary purpose of the nose cone is to store communications and tracking electronics. The
communications bay that is housed within the nose cone of the launch vehicle is made from 3D
printed ABS plastic attached to the nose cone’s bulkhead via steel screw. LBR chose to use 3D
printed ABS plastic material because it is light weighted and easy to manufacture. Additionally,
it also provides a sturdy frame for the communications equipment during flight and landing.
5.4.3 Payload Section
Figure 5.6 CAD Model of Payload Section
The payload bay is responsible for housing the DORITO, RDM and the drogue parachute. The
payload bay features a 36-inch G12 fiberglass airframe. The airframe material was chosen over
other alternatives because of its high strength and non-conductive properties to withstand several
dynamic launches and landing, perfect for storing the DORITO and RDM. Further details about
the design of the DORITO and RDM are discussed in the Payload Section.
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Figure 5.7 CAD Model of Payload Section with Inside View
A bulkhead with a diameter of 5.998 inches installs into the nose cone shoulder for the RDM to
attach to. The nose cone will be secured by the RDM’s threaded rod and three 4-40 nylon screws
to ensure safety. Another aluminum bulkhead with the same dimensions is attached to the
airframe securing the opposite end of the RDM Bay. The bulkhead is secured via 4-40 steel
screws installed from outside the launch vehicle into the bulkhead.
The payload section failure would occur at the aluminum bulkhead between the RDM bay and
the drogue compartment. The aluminum bulkhead could shear during maximum loading in flight
and compromise the safety of the launch vehicle. In order to mitigate the failure mode, the
maximum possible forces on the bulkhead must be analyzed. The scenario in which the
maximum load is imposed is when the launch vehicle is descending under the drogue parachute.
At this moment, the payload bay is moving close to 90.67 𝑓𝑡/𝑠. The upward force on the
payload section is equal to the force that the 20-inch drogue parachute pulls onto the payload
section. The force that the drogue parachute produces can be calculated below:
𝐹 =1
2𝜌𝑣2𝐶𝑑𝐴 (1)
where F is the drag force in lb, 𝜌 is the air density which is 0.00238 𝑠𝑙𝑢𝑔/𝑓𝑡3, 𝐶𝑑 is the
coefficient of drag of the drogue parachute which is approximately 1.5, A is the surface area of
the drogue parachute which is 8.38 𝑓𝑡2, and 𝑣 is the drogue descent velocity of the launch
vehicle which is 90.67 𝑓𝑡/𝑠.
Given the above equation, the force that the drogue parachute produces during the descent is
approximately 123 lb.
LBR is using four 6-32 stainless steel screws to screw in this bulkhead from outside of the
airframe. These screws are placed perpendicular to the axis of the launch vehicle and they shall
withstand the shearing stress. This means that each screws will experience 30.75 lb across its
cross-sectional area. LBR chose the 6-32 stainless steel screw with a diameter of 0.138 inch. This
results in a shear stress of 2055.88 psi.
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30.75 𝑙𝑏
𝜋0.0692 𝑖𝑛2= 2055.88 𝑝𝑠𝑖 (2)
Using the factor of safety of 3 for flight hardware and a shear stress of 6-32 stainless steel screw
of 70000 psi [1], the margin of safety value is 10.35 as shown from the equation below.
𝑀𝑂𝑆 =𝜎𝑦
𝜎𝑙𝑜𝑎𝑑𝐹𝑂𝑆− 1 =
70000 𝑝𝑠𝑖
2055.88 × 3 𝑝𝑠𝑖− 1 = 10.35 (3)
This is only an estimate of the maximum shear force that the screws experience. LBR will
perform further testing and analysis to ensure that these screws are sufficient to hold onto the
bulkhead despite the force that is experienced by the drogue parachute during descent.
5.4.4 Recovery/Avionic Bay Section
Figure 5.8: Full-Scale Launch Vehicle Recovery and Avionics Dimensions
The AV bay section consists of the AV bay and the two sections of airframe on both ends of the
AV bay. The main parachute is in the bottom section of the launch vehicle which is between the
Propulsion Bay front end and the Avionics Bay aft end. With the removal of the airbrakes, the
main parachute is now allotted additional spacing within the launch vehicle, allowing plenty of
room for the total packed length of 14.5 inches for the main chute. The main will have a packed
diameter of approximately 5.9 inches as to secure a snug but easy-sliding package. Doing so
would provide the main parachute with a higher probability for deploying successfully. The
drogue parachute is located in the upper section of the launch vehicle which is between the
Avionics Bay front end and the Payload Bay aft end. The drogue parachute will have a total
packed length of 8 inches, and will be packed with a maximum allotted packing diameter of 5.9
inches. Both the main and drogue parachutes have enough spacing and dimension allotments for
a successful exit upon deployment.
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Figure 5.9: CAD of Avionic Bay (Collapsed View)
All avionics electronics and its accompanying components will be housed in the avionics bay.
The avionics bay will be manufactured out of a 6-inch diameter fiberglass coupler with a total
length of 13 inches. A 1-inch switch band will slide over the coupler piece and will be epoxied at
the midpoint location of the coupler. The switch band is used to activate the primary and backup
arming switches, and will also serve as the location for the four static pressure sampling holes
that the barometric altimeters will utilize during the launch phase. The pressure sampling holes
have a diameter of 0.37”.
Figure 5.10: CAD of Avionics Bay (Exploded View)
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The avionics bay will be enclosed on either side with 6061 aluminum bulkheads. The bulkheads
are held together against the coupler through the application of two 5/16-inch threaded rods.
These rods lay directly parallel to the coupler length and are inserted through the two holes
drilled on both bulkheads; the threaded rods will also hold tension loading during the descent
phase. Nuts and washers will be used to fasten the threaded rods to the bulkheads and coupler
assembly. Additionally, these aluminum bulkheads will also act as the hardpoint for the
attachment of the recovery harnesses. A 5/16-inch U-Bolt will be mounted on the bulkhead to
serve as the hardpoint. The aluminum bulkhead will also act as a surface for mounting PVC
charge holders as well as the terminal blocks.
All of the avionics electronics will be mounted on the avionics tray. The avionics tray will be
constructed of 3D printed spacers that are enclosed between two 12.5 x 5.5-inch wooden plates.
The 3D printed spacers and the two wooden boards will be fastened using machine screws and
nuts.
The 3D printed spacers will be ¾-inch thick and will also have guide holes for the threaded rods
to slide through. In addition, the 3D printed AV tray will have compartments for the standard 9V
batteries to power the barometric altimeters. The batteries will be held in place between the AV
tray and the two wooden boards, ensuring the batteries remain secure during launch.
Figure 5.11: 3D printed AV tray
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Figure 5.12: Avionics Tray (Collapsed View) Figure 5.13: Avionics Tray (Exploded View)
Each parachute will be banded together through 38” flat nylon webbed shock cords with a tested
breaking strength of 1,400 lbs. Zinc-plated quick links (5/16-inch) will be utilized as the hard
point to connect the parachutes to the shock cord and the shock cords to the rest of the sections
of the launch vehicle. These quick links allow for fast and straightforward attachment or
detachment of the nylon shock cords, and have a capacity 2,200 lbs. The chosen quick links will
be able to withstand the forces created during launch. 5/16-inch Galvanized Steel U-Bolts with a
maximum capacity of 600 lbs will be utilized as the hard point for the attachment of the recovery
harness to the rest of the launch vehicle sections. The U-Bolts will be mounted on the 6061
aluminum bulkheads which have a thickness of ¼-inch. The U-Bolt and aluminum bulkhead
arrangement will enclose the avionics coupler and acts as the hard point for the recovery harness
to attach to. Terminal blocks and PVC ejection charge holders will be secured with their
respective machine screws, washers, and hex nuts.
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Figure 5.14: U-Bolt and Aluminum Bulkhead Attachment (Exploded View)
Figure 5.15: U-Bolt and Aluminum Bulkhead Attachment (Collapsed View)
One end of the drogue parachute harness will be connected to a U-Bolt located on the motor tube
and the other end of the harness will be connected to another U-Bolt on the avionics coupler aft
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end. For the main parachute harness, one end will be connected to the second U-Bolt located on
the avionics coupler forward end, and the other end will be connected to another U-Bolt on the
RDM bay aft end. Four 4-40 nylon shear pins will be used to fasten the Payload Bay and
Propulsion Bay sections to the avionics coupler.
Extreme protective measures will be taken to ensure both the drogue and main parachutes are
safe from potential damage that could occur from the black powder ejection charges. The drogue
parachute will have a fireproof Nomex blanket that will shield the drogue chute from the black
powder ejection charges. The main chute will be enclosed in a deployment bag along with a
fireproof Nomex blanket to protect the parachute from the black powder ejection charge while
ensuring the main chute inflates properly.
The drogue parachute will have two opportunities to deploy accordingly – there will be a
detonated charge once the launch vehicle reaches apogee and then there will be a secondary,
redundant charge detonated on a two second delay after the launch vehicle has reached apogee. If
the first black powder ejection charge fails, the secondary charge will ensure that the drogue
chute inflates outwards into the atmosphere from the airframe. Similarly, there will be a primary
and backup charge for the main parachute as well. The primary main ejection charge is
configured to deploy at an altitude of 500 feet, and the backup main ejection charge is set to
deploy at 450 feet should the primary main ejection charge fail.
5.4.5 Propulsion Bay
Figure 5.16: CAD Model of Propulsion Section with Inside View
The material of the propulsion bay will consist of G12 fiberglass airframe tubing with four slots,
6-in. in length by 0.15-in. in width, where the fins will be inserted. The length of the airframe
will be 42-in. with a 6-in. inner diameter and a 6.17-in. outer diameter. The exterior of the
airframe will include three 1515 rail buttons for stable positioning of the vehicle on the launch
rail. The interior will also contain three aluminum centering rings, and a thruster plate. Each
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centering ring will be fastened to the airframe by four 4-40 screws. The most forward centering
ring will be positioned 14 inches from the aft of the airframe, the middle centering ring will be
placed 7.5 inches from the aft, and the last centering ring (thruster mount) will be placed 1.5
inches from the aft. The thruster plate will be secured to the most aft centering ring by three 10-
32 screws and will secure the motor at the base of the propulsion bay. The centering rings and
thruster plate are designed to hold the motor in place to prevent it from rattling and disturbing the
smooth flight path. As for the main parachute, it will be attached to the top of the motor casing.
Further design and analysis of the fins, motor tube and centering rings are detailed below.
The AeroTech L1390G motor produces a maximum thrust of 1650 N. With an outer diameter of
6.17 inches and inner diameter of 6 inches of the full-scale launch vehicle, the compressive stress
in the airframe during maximum thrust can be calculated using the compressive stress equation
below:
𝜎 =𝐹𝑇
𝐴(4)
Where 𝜎 is the compressive stress through the airframe, 𝐹𝑇 is the maximum thrust produced by
the motor, and A is the cross-sectional area of the airframe.
𝜎 =370.93 𝑙𝑏
𝜋(3.0852 − 32)= 228.28 𝑝𝑠𝑖 (5)
The maximum average compressive stress is 228.28 psi.
𝑀𝑂𝑆 =5000 𝑝𝑠𝑖
228.28 × 3 𝑝𝑠𝑖− 1 = 6.30 (6)
The maximum compressive strength of fiberglass is approximately 5000 psi [], which gives a
margin of safety of 7.30, when using the factor of safety of 3. Therefore, the G12 fiberglass
airframe is capable of withstanding loads due to the thrust from the motor.
Fins
After reviewing the possible fin designs, LBR decided to select a trapezoidal fin shape with
tapered edges. The trapezoidal shape was chosen to maximize the stability and minimize the
trailing edge contacting the ground upon landing, thus avoiding damage to the fins. LBR has
chosen to utilize four trapezoidal fins composed of twelve layers of Hexcel carbon fiber, which
will be heated in the oven for 24 hours. The team has selected carbon fiber due to its lightweight
and durable properties. Trapezoidal fins will be integrated since this shape has a relatively
optimized area and will provide more stability during flight. Because the launch vehicle will fly
within the subsonic regime, the fins will include airfoiled edges to reduce pressure drag and
induced drag; the leading edge will be rounded, and the trailing edge will be tapered. Dimensions
of the fins consist of a 7-in. root chord, 3.5-in. tip chord, and 6-in. fin height.
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Figure 5.17: Fin Drawing
The fin placement is designed to redirect the center of pressure aft of the center of gravity, which
allows perturbing forces on the center of gravity (such as wind) to be balanced. Each fin will be
inserted into the airframe and securely adhered with J-B Weld Epoxy. The manufacturer of J-B
Weld Epoxy states that this type of epoxy can withstand temperatures up to 550° F when fully
cured and has a tensile lap shear at 3960 psi []. To ensure a smooth flight and to reduce
interference drag, a 0.5-in. radius fillet will be created using Aeropoxy Light Epoxy.
Fins Alignment
To ensure exact positioning of the fins when mounting onto the airframe, a 3D printed jig has
been created to align the fins into the airframe slots. It is necessary for each fin to be aligned with
precision when applying the epoxy to the fins. This jig is 3D printed using ABS material and will
hold all four fins in place perpendicular to the airframe until the epoxy is completely cured. The
jig will first be placed on the airframe, and the fins with epoxy on the root edge will be placed in
the jig and secured in place. After epoxy is cured, jig will be removed, and fin fillets will be
created.
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Figure 5.18: Fin Slot Alignment Jig
Motor Tube
Figure 5.19: CAD of Motor Casing
The motor tube will be fitted to suit an Aerotech L1390 engine. The official motor casing will be
a RMS-75/5120 casing including a forward disk that will be manufactured by Aerotech. The
motor tube will be constructed from aluminum with a weight of 2.24 lbs. and dimensions of
23.72-in. length and 3-in. width. As previously stated, the motor will be secured with three
centering rings and a thruster plate.
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Stress Analysis and Calculation of Propulsion Components
Stress analysis was performed on several propulsion components to ensure no deformation or
fractures occur during launch. SolidWorks Simulation was used to analyze the stress experienced
by the centering rings and thruster plate due to the forces of the motor, as well as the force
experienced by a fin upon landing.
Centering Ring
The outer radius edge has a fixed geometry and is represented by the orange arrows, which are
positioned where the centering ring will be screwed into the airframe. The purple arrows indicate
the force experienced from the motor which is normal to the inner radius edge and is measured to
be 371 lbf, representing the maximum motor thrust.
Figure 5.20: Stress Analysis on Centering Ring Using SolidWorks Simulation
Thruster Plate
The green arrows indicate a fixed geometry and are positioned inside each hole where the
thruster plate is screwed into the thruster mount. The purple arrows indicate the force
experienced from the motor and is normal to the inner radius edge of the plate. The total force
exerted on these faces is 371 lbf which represents the maximum motor thrust.
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Figure 5.21: Stress Analysis on Thruster Plate Using SolidWorks Simulation
Fin
The green arrows indicate a fixed geometry and are positioned on all faces of the fin tab that are
inserted into the airframe and secured by epoxy. The purple arrows indicate the force which is
normal to the outboard edge of the fin representing a scenario in which the fin lands on that edge.
The value of this landing force is calculated to be 97.5 lbf using the following equation.
𝐹 =1
2𝜌𝐶𝐷𝐴𝑉2 (7)
F = force, 𝜌 = 𝑎𝑖𝑟 𝑑𝑒𝑛𝑠𝑖𝑡𝑦 = .0023769𝑠𝑙𝑢𝑔
𝑓𝑡3, 𝐶𝐷 = 𝐷𝑟𝑎𝑔 𝑐𝑜𝑒𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑡 = 2.2,
𝐴 = 𝑠𝑢𝑟𝑓𝑎𝑐𝑒 𝑎𝑟𝑒𝑎 𝑜𝑓 𝑚𝑎𝑖𝑛 𝑝𝑎𝑟𝑎𝑐ℎ𝑢𝑡𝑒 = 116.9 𝑓𝑡2, 𝑉 = 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦 𝑜𝑓 𝑝𝑟𝑜𝑝𝑢𝑙𝑠𝑖𝑜𝑛 𝑏𝑎𝑦 = 17.86 𝑓𝑡/𝑠
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Figure 5.22: Stress Analysis on Fin Using SolidWorks Simulation
To ensure that the launch vehicle’s carbon fiber fins have sufficient aeroelastic and dynamic
stability for flight, a fin flutter and stress analysis program, AeroFinSim was used to analyze all
four fins. This program created by AeroRocket/John Cipolla uses the Pines Approximation
Method to calculate both divergence and flutter velocities, as well as fin stress at given
velocities.
Divergence Velocity
The critical speed at which the elastic stiffness becomes insufficient to hold the fin in position is
known as the divergence velocity. At this speed, steady-state aeroelastic instability occurs.
During flight conditions above this speed, the fin will begin to deform and create an angle of
twist, causing a torsional divergence of the vehicle. AeroFinSim uses the following equation to
calculate divergence velocity:
𝑞𝐷 =𝑘𝛼
𝑠𝑒𝜕𝐶𝐿𝜕𝛼
(8)
Where 𝑞𝐷 = Divergence Velocity, 𝑘𝛼 = Torsion Spring Stiffness, 𝑠 = Fin Surface Area,
𝑒 = 𝑋𝑒𝑎 − 𝑋𝑎𝑐, 𝜕𝐶𝐿
𝜕𝛼 = Fin Lift Slope
𝑋𝑒𝑎: 𝐸𝑙𝑎𝑠𝑡𝑖𝑐 𝑎𝑥𝑖𝑠 𝑙𝑜𝑐𝑎𝑡𝑖𝑜𝑛 𝑚𝑒𝑎𝑠𝑢𝑟𝑒𝑑 𝑓𝑟𝑜𝑚 𝑎𝑖𝑟𝑓𝑜𝑖𝑙𝑒𝑑 𝑙𝑒𝑎𝑑𝑖𝑛𝑔 𝑒𝑑𝑔𝑒 𝑑𝑖𝑣𝑖𝑑𝑒𝑑 𝑏𝑦 𝑐ℎ𝑜𝑟𝑑 𝑙𝑒𝑛𝑔𝑡ℎ (𝑐)
𝑋𝑎𝑐: 𝐴𝑒𝑟𝑜𝑑𝑦𝑛𝑎𝑚𝑖𝑐 𝑐𝑒𝑛𝑡𝑒𝑟 𝑙𝑜𝑐𝑎𝑡𝑖𝑜𝑛 𝑚𝑒𝑎𝑠𝑢𝑟𝑒𝑑 𝑓𝑟𝑜𝑚 𝑙𝑒𝑎𝑑𝑖𝑛𝑔 𝑒𝑑𝑔𝑒 𝑑𝑖𝑣𝑖𝑑𝑒𝑑 𝑏𝑦 𝑐ℎ𝑜𝑟𝑑 𝑙𝑒𝑛𝑔𝑡ℎ (𝑐)
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Flutter Velocity
The lowest speed at which an elastic body flying at given conditions will exhibit sustained
harmonic oscillation maintaining steady amplitude is known as the flutter velocity. At this speed,
dynamic instability of the fin in an airstream occurs. The forces that produce flutter are from the
deflection of the fin due to deformation which occurs at the divergence velocity. At speeds above
the flutter velocity, even the slightest disturbance to the fin, such as wind, can cause violent
oscillations that can rip the fin off the airframe. AeroFinSim uses the following equation to
calculate flutter velocity:
𝑈
𝑏𝜔𝛼= 𝜌∞√(
2𝑚
𝜌∞𝑏𝑠)
𝑟∞2
𝜕𝐶𝐿
𝜕𝛼[𝑥𝛼 +
𝑒𝑏
](9)
Where 𝑈 = Flutter Velocity, 𝑏 = Average Fin Half-Chord, 𝜔𝛼 = Uncoupled Torsion Frequency,
𝑚 = Fin Mass, 𝑟∞ = Fin Radius of Gyration, 𝑥𝛼 = 𝑋𝐶𝐺 − 𝑋𝑒𝑎
𝑋𝐶𝐺: 𝐶𝑒𝑛𝑡𝑒𝑟 𝑜𝑓 𝑔𝑟𝑎𝑣𝑖𝑡𝑦 𝑙𝑜𝑐𝑎𝑡𝑖𝑜𝑛 𝑚𝑒𝑎𝑠𝑢𝑟𝑒𝑑 𝑓𝑟𝑜𝑚 𝑙𝑒𝑎𝑑𝑖𝑛𝑔 𝑒𝑑𝑔𝑒 𝑑𝑖𝑣𝑖𝑑𝑒𝑑 𝑏𝑦 𝑐ℎ𝑜𝑟𝑑 (𝑐)
Inputs
Several values must be input to AeroFinSim for proper analysis to be performed. All fin
dimensions were input as well as a User Defined Adhesive which is the epoxy used for fin
attachment with an allowable tensile stress of 3960 psi. Additional values input and represented
below specify the User Defined Material that is the Hexcel Carbon Fiber used to make the fins.
Table 5.4: Material Properties of Carbon Fiber
Modulus of Elasticity 33503717.48 lb/in2
Material Density .0643 lb/in3
Poisson’s Ratio .1
Yield Strength - Tensile 644982.8 lb/in2
Results
To ensure no fin deformation or separation from the airframe occurs, both divergence and flutter
velocity shall never be met at any point during flight. From the OpenRocket simulation, the
maximum velocity the launch vehicle will experience is 645 ft/s. AeroFinSim produced
divergence and flutter velocity values that are significantly greater than the maximum velocity,
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thus ensuring all fins will remain undeformed and securely in position throughout the entire
flight.
AeroFinSim was used to perform fin stress analysis to determine the stress experienced by the
fin at the maximum velocity to ensure this value does not exceed the maximum allowable fin
stress. At each velocity up until the flutter velocity, the fin stress was plotted below.
Figure 5.23: Results from Fin Analysis producing a Divergence Velocity of 3287.16 ft/s and Flutter Velocity of 4471.68 ft/s
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5.5 Subscale Flight Results
5.5.1 Subscale Design
The subscale flight in December gave the team valuable data as well as considerations to keep in
mind as the full scale vehicle and payload designs were finalized. Many features from the
subscale design were proven functional and fit for the full scale design. These include the
Figure 5.24: Fin Stress Analysis confirming stress experienced by fin does not exceed maximum allowable stress
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fiberglass airframe that successfully protected the DORITO and RDM, the carbon fiber fins and
attachment epoxy, and the bolted and fully removable centering rings and bulkheads.
To test the new features in the full-scale design, the subscale launch vehicle was designed to be
very similar to the full-scale launch vehicle. The configuration of the full-scale vehicle sections
are identical to that of the subscale launch vehicle, starting with the Ogive Nose Cone, then
Payload Bay, AV bay, and the propulsion bay. The only major difference in the assembly of the
subscale launch vehicle and the full-scale launch vehicle is the airbrake subsystem. The subscale
launch vehicle had an airbrake subsystem that was placed in front for the motor tube. Due to
some complications and safety concerns, LBR decided to not place the airbrake subsystem in the
full-scale launch vehicle as mentioned previously in Section 3.
Figure 5.25: Subscale Launch Vehicle OpenRocket Design
The subscale launch vehicle was scaled to 66.7% of the full-scale launch vehicle. The scaling
factor was enforced both in the diameter of the launch vehicle which is 4.024 inches and the
length of the launch vehicle which is 70 inches, compared to the planned launch vehicle
dimensions of 103 inches in length, and 6.17 inches in diameter for full-scale. The subscale
length was 0.679 times the length of the full scale, and the subscale diameter was 0.652 times the
diameter of the full scale. Therefore, any effects of the full scale launch would appear in the
subscale launch.
Figure 5.26 below shows the various aspects of the launch vehicle that will see a scaling subscale
to full scale. Included is the comparison in length, diameter, mass, rail exit velocity, acceleration,
altitude, landing drift and motor impulse.
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Figure 5.26: Scaling Factors from Subscale and Full scale
The stability of the subscale launch vehicle was 2.62 calibers. This was designed to be close to
the planned stability of the full scale launch vehicle, which is approximated to be 2.5 calibers.
The only parameter not constant between the subscale and full-scale launch vehicles is the fin
dimensions. The fins of the subscale launch vehicle were not scaled down to 66.7% because the
fin design change occurred after the subscale fins were manufactured.
5.5.2 Launch Day Simulation and Recorded Data
The subscale test launch was performed on December 2nd, 2017 at the Friends of Amateur
Rocketry Launch Site near Mojave, CA at an elevation of 2125 feet above sea level. On the day
of launch, the sky was clear, there were 5 mph winds, and the air temperature was approximately
68° F.
Altitude
Mathematically modeling, OpenRocket, produced a predicted apogee altitude of 4165 feet. The
subscale launch vehicle was flown with two altimeters. The average recorded apogee from both
the altimeters on board was 4111 feet. Additionally, the altimeter found the average maximum
acceleration was 9.8 G. The maximum velocity reached was 534 ft/s, and the time the subscale
launch vehicle reached apogee was 17 seconds after launch. It is noted that the variation between
predicted and experimental values could be due to a 1.5% variance in motor impulse, a 2.5%
variance in coefficient of drag, or a variance in wind speed during the flight duration. A data plot
of altitude over the duration of the flight, and quantitative results are shown below.
00.10.20.30.40.50.60.70.80.9
11.11.2
Length Diameter Mass MaxAcceleration
Rail ExitVelocity
Altitude Drift MotorImpulse
Comparison Between Subscale and Full Scale Launch Vehicle
Subscale Full Scale
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Table 5.5: Experimental Data vs. Simulation Data
Measurement Experimental Data Simulation
Apogee 4111 feet 4165 feet
Time to Apogee 17 seconds 16.2 seconds
Maximum Acceleration 9.5 G 9.87 G
Maximum Velocity 534 ft/s 569 ft/s
Coefficient of Drag 0.58 0.56
Figure 5.27: Experimental Data for Subscale Launch
Using the experimental apogee of 4111 feet from the subscale flight, the coefficient of drag was
calculated to be 0.58 using MATLAB.
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Drift Distance
Drift distance was factored into a simulated trajectory of the subscale launch vehicle. Based on
the conditions on the day of the subscale launch, the drift distance was predicted to be 800 ft,
launching into a 5 mph headwind with 5 degrees of the launch rail. Actual subscale launch drift
was paced to be approximately 700 feet. A simulated trajectory from OpenRocket and data can
be seen below.
Table 5.6: Subscale Launch Drift Data
Drift Distance (feet)
Actual Launch 700
OpenRocket Simulation 800
Figure 5.28: Simulation for Subscale Drift
5.5.3 Subscale Flight Analysis
Much of the error between the experimental and simulated values came from the fact that LBR
was testing the airbrake subsystem during the subscale flight. Since LBR has yet to develop a
simulation that can integrate an airbrake subsystem to the launch vehicle, simulation of the
subscale launch vehicle was done assuming the airbrake does not deploy. By removing the
airbrake subsystem for the full scale launch vehicle, LBR should expect that the simulated data
will be similar to the experimental data.
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Figure 5.29: Subscale Launch Vehicle post-flight recovery
From observation, the subscale flight has a stable flight. Successful recovery events verified the
soundness of the launch vehicle recovery design. However, when the drogue parachute deployed,
the team noticed that the subscale launch vehicle maintained an unpredictably spiral maneuver
upon descending. When LBR examined the landing site, the shock cords were twisted and some
shock cords from the AV bay were tangled. From the footage of the onboard camera and from
the ground, the team concluded that there was not enough space when packing the main and the
drogue parachute in the launch vehicle. Because it was very tight in the drogue and the main
compartment, these shock cords had a higher risk for overlapping with each other with the quick
links, causing these shock cords to twist and tangle. In addition to the tightness of the packaging,
another observation made was with the shock cord layout of the recovery system. LBR
overlooked the location of the longest shock cord to be used for the Propulsion Bay front end and
drogue parachute quick link connection and ultimately used a shock cord setup that was inverse
to the actual layout desired. Previous ground testing proved that the parachutes and cords
deployed successfully without tangling even with the layout used during launch; however, the
ground test could not confirm those results on launch day. With the removal of the airbrake
subsystem from the full-scale launch vehicle, there will be an additional space for the main and
the drogue compartment. The desired shock cord layout will be analyzed again and utilized to
prevent any source of error. In addition, LBR will perform several ground tests to confirm that
this problem will not happen again during the full-scale test flight.
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Overall, the subscale test flight was a success. None of the launch vehicle sections or the internal
electronics were damaged, and the subscale launch vehicle was recoverable and reusable. The
team also managed to test the RDM subsystem as seen in figure below. As the result, LBR is
moving forward with the general configuration of the launch vehicle as planned in the Launch
Vehicle Design section. Since the fiberglass airframe successfully protected the DORITO and
RDM, the carbon fiber fins and its attachment epoxy were not damaged, and the bolted and fully
removable centering rings and bulkheads did not shear, LBR will continue using these designs
while still looking for improvements that will reduce the chance of failure.
Figure 5.30: RDM Performance During Subscale Launch
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5.6 Recovery System
5.6.1 Design Overview
Figure 5.31: Launch Vehicle Dual Deployment Attachment Layout
The launch vehicle recovery system will employ a dual deployment recovery method, utilizing
two separate parachutes at different periods of time during the descent phase. A compact drogue
chute will deploy at apogee and the main parachute will deploy at a lower altitude of
approximately 500 feet. Using this recovery at these current configurations minimizes the drift
distance and ensures that the launch vehicle recovers within the maximum allocated recovery
radius of 2,500 feet.
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To deploy said parachutes at their respective times and altitudes, the launch vehicle will integrate
a black powder charge ejection system. Using black powder charges are ideal because of its low
mass system, minimal construction and assembly, low cost, and straightforward operation.
Therefore, using black powder charge ejection is a highly effective and reliable system for
LBR’s launch vehicle. Because of the heated gasses and highly corrosive atmosphere that black
powder ejection creates once it is ignited, the following will be used to protect the parachutes
and shock cords from the burst of the black powder charges: fireproof protective blankets,
protector sleeves, and deployment bags.
Figure 5.32: Diagram of the Launch Sequence
The launch sequence will consist of four main stages: Launch, Coast, Apogee-Drogue
Deployment, and Main Parachute Deployment. The first stage is the Launch stage and consists of
the launch vehicle initiating its propulsion start. The second stage is the Coast stage and
illustrates the launch vehicle as it coasts upwards in altitude after motor burnout. The third stage
is the Apogee-Drogue Deployment stage, where the launch vehicle will execute its first event
and deploy the drogue parachute once the vehicle reaches apogee. The fourth and final stage is
the Main Parachute Deployment stage, where the launch vehicle will deploy its main parachute
once the vehicle reaches an approximate altitude of 500 ft.
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5.6.2 Altimeter
The selected altimeter for the launch vehicle is the PerfectFlite StratoLoggerCF. The
StratoLoggerCF altimeter was chosen as the competition altimeter because of its dual
deployment capabilities, reliability, and cost efficiency. Furthermore, several members on the
team have prior experience and knowledge using the StratoLoggerCF altimeter and will be able
to provide additional troubleshooting should there be any problems that arise during launch day.
The StratoLoggerCF altimeter will adapt the Schurter rotary switch as the primary and backup
arming switch. This rotary arming switch was chosen because it can be mounted flush with the
airframe switch band and can be easily armed or disarmed using a flat blade screwdriver.
Two PerfectFlite StratoLoggerCF altimeters will be used for the launch vehicle; one will serve as
the primary altimeter and the other will serve as a backup altimeter. Using two altimeters
provides the recovery system with a necessary degree of redundancy. The e-matches will be
located on the outside of the avionics bay and will be connected to the altimeter through terminal
blocks that are mounted and secured on the bulkheads. E-matches will be used to ignite the black
powder. The block diagram and schematic in Figure 8 and Figure 9 demonstrate the electrical
connections of the altimeters and how it is integrated in the avionics bay.
Figure 5.33: Block Diagram of Altimeter Electrical Connections
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Figure 5.34: Electrical Schematic of Altimeters in Avionics Bay
As seen from the figures above, each altimeter will have its own arming switch, dedicated power
supply, e-match, and ejection charge to ensure a successful parachute deployment. The primary
altimeter will deploy the drogue chute at apogee and the main chute at approximately 500 feet.
The backup altimeter will be programmed to deploy the drogue parachute two seconds after
apogee is reached and the main parachute at approximately 450 feet. The backup altimeter will
be connected to ejection charges that will carry an increased amount of black powder by 10% of
the amount housed in the primary ejection charge. This will ensure that the backup altimeter
ejection charges will deploy the parachutes for the event that the primary ejection charges fail to
ignite or fail to create enough driving force to deploy the necessary parachutes.
5.6.3 Tracking Devices
The chosen tracking device for the launch vehicle is the BRB900 GPS transmitter from
BigRedBee. The BRB900 was chosen because of its easy integration and simple user interface. It
also does not require a HAM radio license which allows for an overall simpler setup. The
BRB900 is a GPS receiver with a 250MW 900Mhz spread spectrum transmitter; a receiver with
a rubber duck antenna receives the GPS coordinates sent back from the transmitter which then
can be displayed on the LCD screen that is on the receiver. The coordinates and data are sent
back to the LCD receiver every five seconds. The transmitter also carries an integrated LiPo
battery power supply which gives the transmitter an overall lasting life of more than three hours.
A range test was conducted for the BRB900 GPS Transmitter to ensure that the GPS Transmitter
can send accurate coordinates of its location within a specified range. For the test to be
successful, the transmitter must be able to receive and transmit coordinates up to at least 2,500
feet away from the handheld LCD receiver. The test involves one team member that is stationary
while holding the handheld LCD receiver and another team member positioned at the necessary
distance to test the range of the transmitter. During the preliminary testing of the transmitter, the
GPS was able to transmit location coordinates at that given condition and thus, passes the range
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test for the GPS transmitter and receiver. Shown below is the GPS working to locate the
Subscale launch vehicle.
Figure 5.35: GPS Tracking of Sub-Scale Launch
From the photos above of the previous subscale test launch, the BRB900 GPS Transmitter was
able to successfully locate the launch vehicle and succeeds as a viable and efficient GPS tracking
device.
To ensure that the BRB900 GPS Transmitter can send back location coordinates over a length of
time, a power duration test was conducted. The test involves charging the battery on the
transmitter until it has reached maximum power and then leaving it powered on while timed.
With a fully charged LiPo battery, the transmitter was still able to transmit data for over three
hours. After three hours, the transmitter was still operational and proves itself to be efficient for
use during launch day.
The BRB900 GPS Transmitter was also tested for radio frequency interference. Although
competition guidelines require the GPS device to be in separate compartments from the avionics
bay as to not interfere with the altimeters, tests were conducted to ensure that even when placed
in the Nose Cone compartment of the launch vehicle, the GPS still did not interfere with the
altimeters while next to the altimeters or in the Nose Cone compartment. The test involved
placing both the transmitter and altimeters within proximity to each other to observe any type of
interference when powered on. The altimeter did not show any signs of interference throughout
the experiment. Regardless, the GPS was housed in the Nose Cone compartment to deter any risk
for interference on the altimeters. The added safeguard and observation with the GPS and
altimeters close together provided a sense of additional security once placed in the Nose Cone.
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5.6.4 Shock Cord
For the launch vehicle shock cords, LBR will use 3/8-inch Flat Nylon Webbing from
Strapworks. Strapworks provide cost efficient and high-strength webbing that suits all purposes
for LBR’s launch vehicle. With a thickness range of 0.070 to 0.075 of an inch, the 3/8-inch Flat
Nylon provides a breaking strength of 1,400 lbs. with a melting point of 380 degrees Fahrenheit.
These qualifications far exceed the strength required to keep the launch vehicle intact after both
events have initiated. Furthermore, previous testing with past subscale and full-scale launches
solidify using 3/8-inch Flat Nylon Webbing as the priority shock cord tethering. The total length
of the shock cords being used for the drogue compartment will be 34 feet; a 14-ft shock cord will
be used to connect the Avionics Bay front end to the drogue chute via zinc-plated steel quick
links. Additionally, the 20-ft shock cord will be attached from the drogue parachute through the
same quick link to the Payload bay aft end. For the main compartment, two 14-ft shock cords
equal in length will be used as the tethering cords for the main. One cord will attach from the
Propulsion Bay front end to the main chute via quick link, and the other cord will utilize the
same quick link and attach from the main parachute to the Avionics Bay aft end.
5.6.5 U-Bolt
5/16-inch Galvanized Steel U-Bolts with a maximum capacity of 600 lbs. will be utilized as the
hard point for the attachment of the recovery harness to the rest of the launch vehicle sections.
The U-Bolts will be mounted on 6061 aluminum bulkheads; the bulkheads will have a thickness
of ¼-inch. The U-Bolt and aluminum bulkhead arrangement will enclose the avionics coupler
and acts as the hard point for the recovery harness to attach to.
5.6.6 Quick Link
LBR will utilize 3/8-inch thick Zinc-Plated Steel oval shaped threaded connecting links for any
connecting points throughout the recovery subsystem. These connecting links allow ½-inch
openings that have a maximum capacity of 2,200 lbs. These quick links have a capacity that
exceeds the necessary requirement of the launch vehicle and will be able to secure the
connection points accordingly without a high risk for failure.
5.6.7 Black Powder
The Launch Vehicle utilizes a black powder ejection charge system when dealing with the
recovery procedures. Using black powder enables the drogue and main parachute to be deployed
and therefore, allows the launch vehicle to recover safely. Necessary calculations for the precise
amount of black powder to be used is crucial for ensuring that the 4-40 shear pins will
successfully shear as well as ensuring that the main and drogue parachutes deploy at their
respective altitudes. The method and equations used to calculate ejection charges is shown
below:
The Ideal Gas Law equation is used to estimate the amount of black powder necessary to shear
the 4-40 shear pins.
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𝑃𝑉 = 𝑁𝑅𝑇 (10)
Where R = 266 𝑖𝑛 − 𝑙𝑏𝑓/𝑙𝑏𝑚 is the Ideal Gas Constant, T = 3307 °𝑅 is the combustion
temperature of black powder, P = is the pressure in 𝑝𝑠𝑖, V is the volume of the tube in 𝑖𝑛3, and N
is the mass of black powder in 𝑙𝑏.
A pressure differential of 15 psi is desired to fully shear the pins and ensure successful
deployment of the parachutes. Therefore, the equation can be simplified as follows:
𝑁 = 0.006 × 𝐷2 × 𝐿 (11)
N = Amount of Black Powder (g)
D = Diameter of Combustion Compartment (in)
L = Length of Combustion Compartment (in)
The equation provides a baseline for gauging the amount of black powder to be used. Actual
results of how much black powder to use will vary depending on future pop test. Using the
equation and knowing the diameter and length of the combustion compartment, the following
data is calculated:
Main Combustion Compartment
𝑁 = (0.006) × (6 𝑖𝑛)2 × (14.5 𝑖𝑛) = 3.13 𝑔𝑟𝑎𝑚𝑠 (12)
Drogue Combustion Compartment
𝑁 = (0.006) × (6 𝑖𝑛)2 × (8 𝑖𝑛) = 1.73 𝑔𝑟𝑎𝑚𝑠 (13)
Table 5.7: Full-Scale Black Powder Calculations Section Compartment
Diameter (in)
Compartment
Length (in)
Estimated
Black Powder
(g)
Actual Black
Powder (g)
Actual Black
Powder +
10% (g)
Main 6 14.5 3.13 To be
determined
To be
determined
Drogue 6 8 1.73 To be
determined
To be
determined
The outer diameter of the launch vehicle is 6.17 inches and the inner diameter of the launch
vehicle is 6 inches. The internal pressure of the launch vehicle has a desired value of 15 psi and
knowing the thickness of the airframe to be 0.17-inch, the Hoop Stress equation allows an
approximate value of the stress created by the internal pressures of the airframe. The Hoop Stress
equation is shown below.
σ𝐻 =𝑃𝐷𝑚
2𝑡(14)
𝜎𝐻= Hoop Stress caused by Internal Pressure
P = Internal Pressure = 15 psi
𝐷𝑚= Mean Diameter of Airframe = 6.085 in
t = Thickness (in) = 0.170 in
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The Hoop Stress caused by the internal pressure of the airframe is calculated below.
σ𝐻 =(15 𝑝𝑠𝑖 ∗ 6.085 𝑖𝑛)
2 ∗ 0.170 𝑖𝑛= 268.46 𝑝𝑠𝑖 (15)
A maximum pressure differential of 15 psi, a mean diameter of 6.0 inches, and wall thickness of
0.17 inches yields a hoop stress of 268.46 psi. Using the hoop stress and the G12 fiberglass
tensile strength of 40,000 psi results in a margin of safety of 48.66, as seen in the equation
below.
𝑀𝑂𝑆 =𝜎𝑦
𝜎𝑙𝑜𝑎𝑑𝐹𝑂𝑆− 1 =
40000 𝑝𝑠𝑖
268.46 × 3 𝑝𝑠𝑖− 1 = 268.46 (16)
Therefore, the G12 fiberglass airframe can withstand all expected pressures without causing
failure to the launch vehicle.
5.6.8 Ground Ejection Test: Subscale Recovery System
Ground ejection testing was conducted for the recovery system prior to subscale launches. To
ensure successful deployment of the parachutes and successful shearing of the pins, each
parachute bay had two black powder ejection charges. This adds additional redundancy to the
deployment system. The first ejection charge is for the primary altimeter and the second ejection
charge is for the backup altimeter. Backup ejection charges carry 10% more black powder than
the primary ejection charge holders. Again, this ensures successful deployment of the parachutes
should the primary charges fail. Using the Ideal Gas Law equation, the black powder charges for
the subscale launch vehicle was calculated.
Table 5.8: Subscale Black Powder Calculations
Section Compartment
Diameter (in)
Compartment
Length (in)
Estimated
Black
Powder (g)
Actual
Black
Powder (g)
Actual
Black
Powder +
10% (g)
Main 4 10 0.96 3.08 3.39
Drogue 4 5 0.48 3.41 3.75
The black powder amount calculated provides the team with a solid foundation of where to start
when gauging the amount of black powder to be used come subscale or full-scale launch days.
However, to confirm a successful separation of the parachute compartments, an executive
decision between the team leads was made to increase the total amount of black powder to be
used in the ejection charge holders. Shown below are photos of a semi-successful ground
ejection test for both the main and drogue using 3.08 grams of black powder for the main
compartment and 3.41 grams of black powder for the drogue compartment. The drogue required
more black powder than the main because the drogue needs to split the launch vehicle in half,
while the main separates less than half of the launch vehicle.
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Figure 5.36: Ground Ejection Test for Drogue Parachute
Figure 5.37: Ground Ejection Test for Main Parachute
As seen above, the drogue compartment did not have trouble shearing the pins. The pop test was
considered semi-successful because the amount of black powder used was able to shear the shear
pins successfully, but was unable to push the main parachute out of the airframe successfully.
However, it is believed this is mainly due to the packaging of the main parachute. Further testing
proved that the main parachute needed to be packaged with an overall width that is less than
what it was when packaged into the airframe as shown in the photo above. After committing to
the packaging width decrease, the main parachute was able to securely deploy, and the previous
subscale launch was able to recover safely with both parachutes deployed at their respective
altitudes. Using 3.08 grams and 3.41 grams of black powder for the main and drogue
respectively, along with a more condensed package for the main chute proved successful as the
subscale launch went according to plan and was able to recover safely and successfully.
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5.6.9 Shear Pin
The team will be using four 4-40 nylon screws as the shear pins that will secure the airframes
together during launch, and will also serve as the pins to be sheared once the launch vehicle
reaches apogee and when the launch vehicle descends to approximately 500 ft. To determine the
number of shear pins to connect the sections of the launch vehicle, LBR calculated the maximum
force that the shear pins need to withstand. Since there are no other external forces once the
launch vehicle descends, the only force that acts on the shear pins are the weight of the
Propulsion Bay and the Payload Bay. The weight of the Payload Bay is 14.819 lbs., the
Propulsion Bay is 18.958 lbs., and the 4-40 nylon shear pins break at 71 lbs. []. Using four shear
pins gives the most security when assembling the launch vehicle together and will ensure that
there will be less risk for errors caused by launch vehicle assembly. While there are other forces
that may affect the shear pins during flight that cannot be calculated, it is very unlikely that any
of those forces will exceed the maximum strength of the shear pins.
Assuming the desired pressure differential to shear off the 4-40 pins is approximately 10 psi and
the internal diameter of the airframe is 6 inches, the force exerted by the black powder to
separate the airframes is calculated using the equation below:
𝐹 = 𝑃𝐴 = 15 𝑝𝑠𝑖 × 𝜋𝑟2 = 424.11 𝑙𝑏 (17)
The force from the separation charges is 424 lbs. and the force on each of the nylon screws is
106.03 lbs. From the calculation, this force by the separation charges is greater than the strength
of the nylon screws; thus, they will break when the charges go off, separating the launch vehicle.
Additionally, ground testing will be performed to confirm that the charges are sufficient, and the
shear pins will successfully separate. The full scale test launch in February will also verify that
the shear pins can hold the launch vehicle together during the flight duration.
5.6.10 Sizing Parachutes
The launch vehicle is approximately projected to weigh 38.7 lbs.; after motor depletion and
excluding the main and drogue parachutes during descension from launch, the approximate
weight to be considered when deciding which parachute to use becomes 32.1 lbs. Knowing that
the maximum allotted kinetic energy at landing of each section is 75 ft-lb, the required parachute
sizing for the main and drogue chutes can be calculated. The following equations are used to
determine the descent velocity and parachute diameter:
𝐾. 𝐸. = 1
2𝑚𝑣2 (18)
K.E. = Kinetic Energy (ft-lb)
m = mass of section (slugs)
V = Descent Velocity (ft/s)
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𝐷 = √8𝑚𝑔
𝜌𝑉2𝜋𝐶𝑑
(19)
D = Diameter (ft)
m = mass (lbs.)
g =acceleration due to gravity = 32.2 ft/s2
ρ = air density at Launch Site = 0.0023769 slugs/ft3
V = Descent Velocity (ft/s)
Cd = Coefficient of Drag
The following equation calculates the maximum descent velocity for the heaviest section of the
launch vehicle (Payload Section = 14.30 lbs.) using the maximum allotted kinetic energy.
𝑉 = √2 × 75
0.444= 18.38
𝑓𝑡
𝑠
Using the now known maximum descent velocity, the equation calculates the minimum main
parachute diameter to achieve a safe recovery process. The calculation assumes a desired 2.20
drag coefficient.
𝐷 = √8 × (0.9968) × 32.2
(0.0023769) × (18.38)2 × 𝜋 × (2.20)= 6.80 𝑓𝑡 = 81.63 𝑖𝑛
To remain within kinetic energy guidelines and with a desired drag coefficient of 2.20, the
minimum diameter for the main parachute must be at or larger than 81.63 inches. This will allow
the launch vehicle to descend at approximately 18.38 ft/s with a kinetic energy of 75 ft-lbs.
Using the same equations, the drogue chute can be calculated with an assumed desired maximum
descent velocity of equal to or less than 95 ft/s. Assuming a desired drag coefficient of 1.50, the
following calculation demonstrates the minimum diameter required for the drogue chute to
descend at a rate of 95 ft/s.
𝐷 = √8(0.9968) × 32.2
(0.0023769) × (95)2 × 𝜋 × (1.50)= 1.59 𝑓𝑡 = 19.13 𝑖𝑛
Therefore, the minimum diameter for the drogue parachute must be equal to or larger than 19.13
inches. This will give the launch vehicle a descent rate of approximately 95 ft/s.
Lastly, LBR is limited in the choice of its parachutes because vendors only sell parachutes of
certain sizes. The team cannot use parachutes exactly of the sizes it calculates because those are
not sold commercially, but instead chooses parachutes similar in size.
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Table 5.9: Subscale Black Powder Calculations
Section Calculated Size
for Drogue (in)
Calculate Size
for Main (in)
Full scale LV
Drogue (in)
Full Scale LV
Main (in)
Payload Bay, AV
Bay and Propulsion
Bay
19.13 81.63 20 84
5.7 Mission Performance Predictions
The launch vehicle and payload performance should be indicative of successful implementation
of the design, build and test process. LBR understands that a safe and stable rocket flight is a pre-
requisite to any innovation in payload design.
5.7.1 OpenRocket and MATLAB/Simulink Flight Simulation
Production of an overall flight profile is crucial in understanding the role various launch
conditions play on the flight targets. Theoretical results can be used to validate design choices
and provide a benchmark for comparison with experimental data. For these reasons, LBR chose
to use an open source simulation software, OpenRocket, to simulate the entire rocket flight from
launch to touchdown. In addition, LBR is currently developing a MATLAB/Simulink flight
simulation to present data from a different method to verify that the OpenRocket results are
accurate.
Figure 5.38: MATLAB/Simulink simulation (beta phase)
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5.7.2 Basic Rocket Equations
To successfully launch a rocket of any size, it is important to understand how to calculate center
of gravity, center of pressure, static stability, and peak altitude of the rocket. The static stability
is a dimensionless number found by dividing the distance between the center of gravity and the
center of pressure by the body tube diameter.
𝑆 =𝑥𝑐𝑝 − 𝑥𝑐𝑔
𝑑(20)
Where 𝑥𝑐𝑝 is the center of pressure and 𝑥𝑐𝑔 is the center of gravity. The center of pressure can be
measured from the nose cone tip using:
𝑥𝑐𝑝 =(𝐶𝑁)𝑛𝑋𝑛 + (𝐶𝑁)𝑓𝑋𝑓
(𝐶𝑁)𝑛 + (𝐶𝑁)𝑓
(21)
Where (𝐶𝑁)𝑛 is the ogive nose cone center of pressure, 𝑋𝑁 is calculated using:
𝑋𝑛 = 0.466𝐿𝑛 (22)
Where 𝐿𝑛 is the nose cone length. (𝐶𝑁)𝑓 is the fin center of pressure coefficient and is computed
using:
(𝐶𝑁)𝑓 = (1 +𝑅
𝑆 + 𝑅)
4𝑛 (𝑆𝑑
)2
1 + √1 + (2𝑙
𝑎 + 𝑏)
2(23)
Where S is the radius of the body between the fins, S is the fin semi-span, and 𝑛 is the number of
fins, 𝑙 is the length of the fin mid-chord line, 𝑎 is the fin root chord length, and b is the fin tip
chords length. And finally, 𝑋𝑓 is calculated using:
𝑋𝑓 = 𝑥𝑓 +𝑚(𝑎 + 2𝑏)
3(𝑎 + 𝑏)+
1
6(𝑎 + 𝑏 −
𝑎𝑏
𝑎 + 𝑏) (24)
Where 𝑥𝑓 is the distance from the nose cone tip to the front edge of the fin root, 𝑚 is the distance
between the fin root and fin tip. The center of gravity is computed using:
𝑥𝑐𝑔 =𝑑𝑛𝑤𝑛 + 𝑑𝑟𝑤𝑟 + 𝑑𝑏𝑤𝑏 + 𝑑𝑒𝑤𝑒 + 𝑑𝑓𝑤𝑓
𝑊(25)
Where 𝑊 is the sum of the whole rocket weight, 𝑑 is the distance between a specific center of
gravity (nose cone, recovery, body, engine, and fins, respectively) and the the aft end of the
rocket.
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The peak altitude is found through the sum of the boost phase while the motor is burning and the
coast phase from the motor burn-out to peak altitude.
𝑃𝐴 = 𝑦𝑏 + 𝑦𝑐 (26)
To find the altitude of the boost phase, 𝑦𝑏, the average mass is first calculated using:
𝑚𝑎 = 𝑚𝑠 + 𝑚𝑒 −𝑚𝑝
2(27)
Where 𝑚𝑠 is the structural mass of the rocket, 𝑚𝑒 is the motor mass, and 𝑚𝑝 is the propellant
mass. The aerodynamic drag coefficient, 𝑘, is
𝑘 =1
2𝜌𝐶𝑑𝐴 (28)
where 𝜌 is the air density, 𝐶𝑑 is the drag coefficient, and 𝐴 is the rocket cross-sectional area.
Knowing the average mass, 𝑚, and the aerodynamic drag coefficient, 𝑘, the burnout velocity
coefficient, 𝑞𝑏, is then computed using:
𝑞𝑏 = √𝑇 − 𝑚𝑎𝑔
𝑘(29)
Where T is the motor thrust, and 𝑔 is the gravitational constant. Equations 2, 3 and 4 are then
used to compute the burnout velocity delay coefficient, 𝑥𝑏, using:
𝑥𝑏 =2𝑘𝑞
𝑚𝑎
(30)
Then the motor burn time, 𝑡, is calculated using:
𝑡 =𝐼
𝑇(31)
Where 𝐼 is the motor impulse, and 𝑇 is the motor thrust. Equations 4,5, and 6 are then used to
calculate the burnout velocity, 𝑣𝑏, using:
𝑣𝑏 = 𝑞𝑏
1 − 𝑒−𝑥𝑏𝑡
1 + 𝑒−𝑒𝑏𝑡(32)
The altitude at burnout, 𝑦𝑏, can finally be computed using:
𝑦𝑏 =−𝑚𝑎
2𝑘ln (
𝑇 − 𝑚𝑎𝑔 − 𝑘𝑣𝑏2
𝑇 − 𝑚𝑎𝑔) (33)
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After the boost phase altitude is calculated, the coast phase altitude can be determined. At
burnout, the new mass during the coast phase is
𝑚𝑐 = 𝑚𝑠 + 𝑚𝑒 − 𝑚𝑝 (34)
With the new coasting mass comes the coasting velocity coefficient, 𝑞𝑐, and the coasting velocity
delay coefficient, 𝑥𝑐
𝑞𝑐 = √𝑇 − 𝑚𝑐𝑔
𝑘(35)
𝑥𝑐 =2𝑘𝑞𝑐
𝑚𝑐
(36)
With equations 10 and 11, the coasting velocity, 𝑣𝑐, can be computed using:
𝑣𝑐 = 𝑞𝑐
1 − 𝑒−𝑥𝑐𝑡
1 + 𝑒−𝑥𝑐𝑡(37)
The coasting phase altitude can then be calculated using:
𝑦𝑐 =𝑚𝑐
2𝑘𝑙𝑛 (
𝑚𝑐𝑔 + 𝑘𝑣𝑐2
𝑚𝑐𝑔) (38)
With the burnout altitude and the coasting altitude, the peak altitude can be determined using
equation 7 above.
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5.7.3 Flight Simulation
An OpenRocket simulation was conducted on the full-scale launch vehicle, which can be seen in
the figure below.
Figure 5.39: Graph of Launch Vehicle Motion vs. Time with 0 mph Wind Conditions
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The motor thrust curve for an AeroTech L1390G motor from OpenRocket can be seen below in
Figure 5.40. The motor burnout occurs at 2.87 seconds.
Figure 5.40: Motor thrust curve for an AeroTech L1390G motor
Open Rocket was used to simulate launches at various wind speeds from 0 to 20 mph in
increments of 5, using an AeroTech L1390G motor. A sample simulation can be seen above in
Figure 5.39, these simulations were used to predict apogee at various wind speeds, with a rocket
mass of 38.125 lbs. with a loaded motor.
Table 5.10: Projected Apogee at Different Wind Speeds Using OpenRocket
Wind Speed (mph) Projected Apogee (ft.)
0 5354
5 5304
10 5221
15 5151
20 5066
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5.7.4 Material Robustness Calculations
The AeroTech L1390G motor produces a maximum thrust of 1650 N. With an outer diameter of
6.17 inches and inner diameter of 6.00 inches of the full-scale launch vehicle, the compressive
stress in the airframe during maximum thrust can be calculated using the compressive stress
equation below:
𝜎 =𝐹𝑇
𝐴(39)
Where 𝜎 is the compressive stress through the airframe, 𝐹𝑇 is the maximum thrust produces by
the motor, and A is the cross-sectional area of the tube.
𝜎 =370.93 𝑙𝑏
𝜋(3.0852 − 32)= 228.28 𝑝𝑠𝑖
The maximum average compressive stress is 228.28 psi.
𝑀𝑂𝑆 =5000 𝑝𝑠𝑖
228.28 × 3 𝑝𝑠𝑖− 1 = 6.30
The maximum compressive strength of fiberglass is approximately 5000 psi [], which give a
margin of safety of 6.30, when using the factor of safety of 3. Therefore, the G12 fiberglass
airframe is capable withstanding loads due to the thrust from the motor.
5.7.5 Center of Gravity, Center of Pressure, and Static Stability Margin
Figure 5.41: OpenRocket side view of launch vehicle
The stability of the launch vehicle was projected by OpenRocket to be 2.5 calibers with a center
of gravity of 65.258 inches measured from the nose cone and a center of pressure 80.675 inches
from the nose cone as shown in Figure 5.x above. OpenRocket was cross referenced for
redundancy using the following equations.
𝑆𝑡𝑎𝑡𝑖𝑐 𝑆𝑡𝑎𝑏𝑖𝑙𝑖𝑡𝑦 =𝐶𝑃 − 𝐶𝐺
𝐷(40)
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Where S is the stability in calibers, CP is the length from the top of the nose cone to the center of
pressure, CG is the length from the top of the nose cone to the center of gravity, and D is the
diameter of the airframe.
Calculation of launch vehicle stability
80.675 𝑖𝑛𝑐ℎ𝑒𝑠 − 65.258 𝑖𝑛𝑐ℎ𝑒𝑠
6.17 𝑖𝑛𝑐ℎ𝑒𝑠= 2.5 𝑐𝑎𝑙𝑖𝑏𝑒𝑟𝑠
To verify the accuracy of the OpenRocket simulations of the mission, LBR has completed the
following hand calculations. The stability of the rocket yielded in the OpenRocket simulation
will be compared to the values calculated below and any discrepancies between the solution may
be a cause for concern. All given measurements for these calculations were taken from the full-
scale launch vehicle’s design.
Table 5.11: Definition of Hand Calculation Symbols
Symbol Definition Units
𝑆 Static stability cal
𝑥𝑐𝑝 Center of pressure in
𝑥𝑐𝑔 Center of gravity in
𝐷 Diameter of launch vehicle in
(𝐶𝑁)𝑛/(𝐶𝑁)𝑓 Individual center of pressure (nose cone/ fins) in
𝑋𝑛/𝑋𝑓 Distance of the component from the aft of the rocket in
𝐿𝑛 Nose cone length in
𝑛 Number of fins in
𝑟 Radius of the body of the rocket in
𝑠 Radius of the body between the fins in
𝑙 Length of the fin mid-chord line in
𝑎 Fin root chord length in
𝑏 Fin tip chord length in
94 Long Beach Rocketry | CDR 2017 - 2018
Figure 5.42: Hand calculations of the static stability
The stability shown on OpenRocket was 2.5 calibers, which is close to the hand calculated value
of 2.463; therefore, LBR can trust that the simulations generated by OpenRocket yield accurate
predictions of the launch vehicle’s stability during the flight duration. The 0.037 difference in the
two values is most likely a result of rounding in the hand calculation.
This stability of our rocket is the result of careful consideration in weight distribution and fin
design. LBR went for a much stable rocket to prevent the mishaps of last year’s rocket launching
off course due to low stability. This year, with a higher stability, that should no longer be an
issue. The stability is also sufficient for NASA’s minimum stability requirement of 2.00. LBR’s
stability is perfect in maintaining a steady launch while not becoming over stable where weather-
cocking becomes a greater issue. Weather-cocking is the result of a too high stability leaving the
launch vehicle more susceptible to the movements of wind. This is most commonly seen with a
stability over 3.00 calibers which LBR is safely under.
5.7.6 Kinetic Energy
When calculating parachute size, LBR ensured that the parachutes were large enough to have a
maximum kinetic energy of 75 ft-lb. The LBR launch vehicle has three sections: nose
cone/payload, recovery/AV bay, and propulsion bay, all of which are tethered together. The
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parachute that will be used as the drogue will be the Fruity Chutes 20-inch TARC Low and Mid
Power Chute and will be located between the Payload Bay aft end and Avionics Bay forward
end. The TARC Low and Mid Power Chute is compact and designed with a spill hole for high
stability; it is a six-gore elliptical shape chute that offers a drag coefficient of 1.5.
The descent velocity can be calculated for the main and drogue parachutes. The equation is as
follows:
𝑉 = √8𝑚𝑔
𝜌𝐷2𝜋𝐶𝑑 (41)
The following calculation uses the 20-inch TARC Low and Mid Power Chute specifications to
determine the approximate descent velocity for the drogue chute during the descent phase.
𝑉 = √8 × (0.9969) × 32.2
(0.0023769) × (1.67 )2 × 𝜋 × (1.50)= 90.67 𝑓𝑡/𝑠
This gives the launch vehicle a descent velocity of approximately 90.67 ft/s after the first
deployment (at apogee).
The parachute chosen as the main will be the Fruity Chutes 84-inch Iris Ultra Standard Parachute
and will be located between the Avionics Bay aft end and the Propulsion Bay forward end; the
parachute is annular (Standard Nylon Toroidal) and provides a coefficient of drag of 2.20.
The following calculation uses the same descent velocity equation to determine the descent
velocity of the launch vehicle once the main parachute deploys.
𝑉 = √8 × (0.9969) × 32.2
(0.0023769) × (7 𝑓𝑡)2 × 𝜋 × (2.20)= 17.86 𝑓𝑡/𝑠
This gives a descent velocity of 17.86 ft/s during the secondary deployment.
Table 5.12: Full-Scale Descent Velocity Calculations
Approx. Full-
Scale Launch
Vehicle Weight
= 32.1 lbs.
Type Size/Type Drag
Coefficient Descent Velocity (fps)
Drogue 20" FC TARC Chute 1.50 90.67
Main 84" FC Iris Ultra
Parachute 2.20 17.86
Throughout descent, the launch vehicle will separate at various stages and will have a main and
drogue parachute to enable a safe descent. Preliminary calculations were performed to ensure the
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launch vehicle sections will not exceed the maximum allowed kinetic energy. The following
equation is used to calculate kinetic energy:
𝐾. 𝐸. = 1
2 𝑚𝑣2
m = mass of independent section
v = descent velocity upon landing
Kinetic Energy Calculation for the Payload Bay (Top Section):
𝐾. 𝐸. = 1
2 (0.413)(17.86 )2 = 70.83 𝑓𝑡 − 𝑙𝑏
Kinetic Energy Calculation for the Avionics Bay (Middle Section):
𝐾. 𝐸. = 1
2(0.148)(17.86)2 = 23.63 𝑓𝑡 − 𝑙𝑏
Kinetic Energy Calculation for the Propulsion Bay (Bottom Section):
𝐾. 𝐸. =1
2 (0.4053)(17.86 )2 = 64.64 𝑓𝑡 − 𝑙𝑏
Table 5.13: Kinetic Energy Calculations for Each Independent Section
Kinetic Energy for Each Independent Upon Landing
Section Weight
(lb.)
Mass
(slugs) Descent Velocity (ft/s)
Kinetic Energy (ft-
lb)
Payload Bay
(Top Section) 14.30 0.44 17.86 70.83
Avionics Bay
(Middle
Section)
4.77 0.15 17.86 23.63
Propulsion Bay
(Bottom
Section)
13.05 0.41 17.86 64.64
Seen in the table above, the section with the highest kinetic energy is in the top section which is
the Payload Bay. Although it has the highest kinetic energy relative to the other sections of the
launch vehicle, it remains within the required 75 ft-lb limits. The highest kinetic energy upon
landing is 70.89 ft-lb and remains under that condition set by competition guidelines.
Furthermore, the other two sections being the Avionics Bay and Propulsion bay have a kinetic
energy of 23.65 ft-lb and 64.69 ft-lb respectively; these kinetic energies on their respective
sections also fall under the required maximum limit of 75 ft-lb.
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5.7.7 Drift Simulations and Calculations
Drift Calculation
All launch vehicle sections should land within the launch field; thus, LBR designed the launch
vehicle to drift no more than 2500 feet. The drift distance is calculated for five different wind
conditions: 0 mph wind, 5 mph wind, 10 mph wind, 15 mph wind, and 20 mph wind. The
calculations shown below accept that the launch vehicle is straight above the launch rail during
apogee and that the drift speed of the rocket is the same as the wind speed; with these conditions
established, the following equation is used to calculate drift distance:
𝐷𝑟𝑖𝑓𝑡 𝐷𝑖𝑠𝑡𝑎𝑛𝑐𝑒 = (𝑡𝑚 × 𝑆) + (𝑡𝑚 × 𝑆) (42)
td: descent time under drogue
tm: descent time under main
S: wind speed
The drogue deployment altitude will be 5,280 ft with the main deployment altitude being
approximately 500 ft. When the launch vehicle lands, it will have an altitude of 0 ft; using the
drogue parachute descent velocity of 90.67 ft/s and main parachute descent velocity of 17.86 ft/s,
this gives a drogue descent time of 52.61 seconds and a main descent time of 28.0 seconds. This
information allows for the calculation of different drift distances for each wind speed.
Total Drift Distance Calculation for 0 mph wind speeds:
𝐷𝑟𝑖𝑓𝑡 𝐷𝑖𝑠𝑡𝑎𝑛𝑐𝑒 = (28 × 0) + (52.61 × 0) = 0 𝑓𝑡
Total Drift Distance Calculation for 5 mph wind speeds:
𝐷𝑟𝑖𝑓𝑡 𝐷𝑖𝑠𝑡𝑎𝑛𝑐𝑒 = (28 × 7.33) + (52.61 × 7.33) = 591.14 𝑓𝑡
Total Drift Distance Calculation for 10 mph wind speeds:
𝐷𝑟𝑖𝑓𝑡 𝐷𝑖𝑠𝑡𝑎𝑛𝑐𝑒 = (28 × 14.67) + (52.61 × 14.67) = 1182.28 𝑓𝑡
Total Drift Distance Calculation for 15 mph wind speeds:
𝐷𝑟𝑖𝑓𝑡 𝐷𝑖𝑠𝑡𝑎𝑛𝑐𝑒 = (28 × 22) + (52.61 × 22) = 1773.42 𝑓𝑡
Total Drift Distance Calculation for 20 mph wind speeds:
𝐷𝑟𝑖𝑓𝑡 𝐷𝑖𝑠𝑡𝑎𝑛𝑐𝑒 = (28 × 29.33) + (52.61 × 29.33) = 2364.56 𝑓𝑡
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Table 5.14: Drift Distance Calculations
Wind Speed
(mph)
Wind Speed
(fps) Drogue Drift (ft) Main Drift (ft) Total Drift (ft)
0 0 0 0 0
5 7.33 385.84 205.30 591.14
10 14.67 771.68 410.60 1182.28
15 22 1157.51 615.90 1773.42
20 29.33 1543.35 821.20 2364.56
Using the 84-inch main chute and the 20-inch drogue chute for the launch vehicle recovery
parachutes allows feasible drift distances if wind speeds were to exceed 15 mph. In extreme
conditions of 20 mph wind speeds, the maximum drift distance of the drogue will be
approximately 1,543.35 ft and the maximum drift distance of the main will be 821.20 ft. This
gives a maximum total drift distance of 2,364.56 ft which is about 130 ft lower than the allotted
recovery radius of 2,500 ft. At reasonable 0 mph through 5 mph wind speeds, the maximum total
drift distance ranges from 0 ft to 591.14 ft and in extreme 20 mph wind speeds, the maximum
drift will remain within competition guidelines for launch vehicle recovery.
Drift Simulations
All launch vehicle sections should land within the launch field. To simulate the horizontal
distance the full-scale launch vehicle would drift with the given parachute sizes, LBR modeled
the full-scale launch vehicle using OpenRocket. In the OpenRocket simulation, the values of the
mass, parachute size, and apogee are equivalent to that of the actual launch vehicle. This
simulation allows the team to estimate the drift of all independent sections.
Figure 5.43 OpenRocket Simulation of Drift in 5 mph Winds
The drift calculation at 5 mph wind is cross referenced in Figure 5.43 above, which shows a drift
distance of approximately 900 ft. This contrasts to our previously calculated drift distance of 582
99 Long Beach Rocketry | CDR 2017 - 2018
ft, leaving a 318 ft difference. This a result of the previous assumptions that each section of the
rocket will descend at the same velocity and that the rocket’s flight path will go straight up
unaffected by the wind, according to the hand calculations.
5.8 Team-Derived Requirements
Table 5.15: Team-Derived Requirements – Launch Vehicle
Requirement Verification Method Verification Plan Status
The launch vehicle
will reach a altitude
approximate to the
target apogee of
5,280±50 feet.
Test The launch vehicle
will be tested for its
maximum altitude
during the full-scale
test launch.
Adjustments in the
launch vehicle design
will be made to reach
an apogee closer to
the projected altitude.
Full-scale launch
vehicle will be tested
on February 24, 2018.
Launch vehicle will
be able to maintain
above 2.0 caliber
static stability during
the flight duration.
Test/ Observation The launch vehicle
will be tested for
stability during the
full-scale test launch
Will be verified
during full-scale test
launch preparations.
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The launch vehicle
will recover safely
and be made reusable
by descending at a
maximum allotted
rate of 20-ft/s with a
maximum allotted
kinetic energy of 71
ft-lb.
Test/Analysis Preliminary
calculations on
parachute sizing for
the current weight
and intended velocity
and kinetic energies
of the launch vehicle
will be calculated and
utilized to choose a
main and drogue
parachute that will
ensure a safe descent.
The calculations will
provide analysis and
will help LBR choose
the best parachute
sizes; testing will be
done through the
upcoming full-scale
launch.
Will be verified
during full-scale test
launch preparations.
The shock cord must
successfully deploy
the launch vehicle’s
drogue and main
parachute at apogee
and at 500-ft
respectively.
Test/Observation The shock cord
reliability will be
tested during the full-
scale pop test. The
shock cords will be
assembled and will
undergo a pop test to
ensure the parachutes
deploy successfully
with the shock cords.
Observations will be
made during the pop
test for successful
deployment. Testing
will be done during
the full-scale launch
where altimeters will
be used to deploy the
shock cords and
parachutes at their
respective altitudes.
Will be verified
during full-scale test
launch preparations.
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The main and drogue
parachutes will not
have any holes or
tears that will hinder
its ability to safely
recover the launch
vehicle.
Inspection The main and drogue
parachutes will
undergo inspection
prior to the full-scale
pop test. The
parachutes will be
inspected for any
holes or tears that
would affect the
performance of the
parachutes.
Will be verified
during full-scale test
launch preparations.
All shock cord and
parachute tetherings
will remain untangled
throughout the launch
and recovery phases.
Test The wrapping of the
shock cords and
parachutes will be
tested during the full-
scale pop test. The
cords will be handled
meticulously and
folded in a manner
that will prevent any
possible tangling due
to the folding of the
shock cords or
parachutes.
Will be verified
during full-scale test
launch preparations.
All shock cords will
have a strength
capable of handling
the approximate 123
lbf from the drogue
and 97 lbf from the
main.
Analysis/Test Research on the
current selected shock
cords give a
numerical input of
1,400-lbs. The shock
current Nylon
Webbed shock cords
chosen should be
enough to withstand
the force produced by
the main and drogue
parachutes
Will be verified
during full-scale test
lau ch preparations.
All avionics
electronic wiring will
remain in contact
throughout the launch
phase and descent
phase until physical
Test The wiring in the
avionics section will
be tested during the
full-scale pop test as
well prior to the full-
scale test launch.
Will be verified
during full-scale test
launch preparations.
102 Long Beach Rocketry | CDR 2017 - 2018
recovery of the
launch vehicle is
conducted.
Electrical wiring will
be fastened with their
terminal blocks and a
pull test will be
performed to ensure
all connections are
secure and at a low
risk for failure.
The GPS tracking
device will remain
operable throughout
launch and will be
able to read distances
of up to at least 2,500
feet.
Test The GPS device will
be tested during the
manufacturing and
assembly phase of the
full-scale launch
vehicle. The battery
will be fully charge
and then timed as it is
left on until the power
completely runs out.
Furthermore, the
transmitter will be
taken to a distance of
at least 2,500 feet
with the receiver
recording data to
verify its reliability.
Will be verified
during full-scale test
launch preparations.
The GPS tracking
device must be
secured in the Nose
Cone compartment of
the Launch Vehicle
and will remain in
contact throughout
launch until physical
recovery of launch
vehicle is conducted
Test The GPS housing
apparatus will be
tested during the
manufacturing and
assembly phase of the
full-scale launch
vehicle. It will be
fastened with its
respective nuts and
washers. The wiring
will be connected and
secured on the
apparatus as to not
cause any separation
throughout launch.
Will be verified
during full-scale test
launch preparations.
The launch vehicle
ejection charges will
Test/Analysis Preliminary analysis
on the amount of
Will be verified
during full-scale test
103 Long Beach Rocketry | CDR 2017 - 2018
house enough black
powder to produce a
pressure differential
inside the airframe of
approximately 15 psi.
black powder to be
used for the ejection
charge is calculated
for and mentioned in
the Recovery Section
of the Critical Design
Report; testing will be
conducted during the
pop test of the full-
scale launch vehicle
which will allow LBR
to gauge the amount
of black powder
required to produce a
pressure differential
within the internal
airframe of
approximately 15 psi.
launch preparations.
The thruster plate
must successfully
secure the motor for
the duration of the
flight.
Analysis The thruster plate will
undergo stress
analysis to ensure the
plate can withstand
the force experienced
from the motor equal
to 371 lbf.
Completed using
SolidWorks
Simulation along with
hand calculations.
The fins will be
secured 90֯ apart
from each other.
Inspection The fins will be
inspected by our team
for any
misalignments.
Will be verified
during full-scale test
launch preparations.
The recovery bay,
propulsion bay, and
payload bay
bulkheads shall
withstand the shock
cords.
Analysis/Calculation/
Test
Force experienced by
the bulkheads will be
analyzed and
calculated to ensure
bulkheads can
withstand pull from
the shock cords.
Will be verified
during full-scale test
launch preparations.
The 4-40 steel screws
shall be strong
enough to withstand
the shear stress
exerted on the
Analysis/Test Stress analyses will
be done to ensure the
steel screws are
strong enough to
secure the centering
Will be verified
during full-scale test
launch preparations.
104 Long Beach Rocketry | CDR 2017 - 2018
centering rings and
bulkhead.
rings and bulkheads.
The launch vehicle
shall be able to
separate using black
powder charge with
424 lbf without
damaging the
payload.
Test Perform ground
ejection tests and
ensure the aluminum
bulkhead behind the
RDM is tightly
secured.
Will be verified
during full-scale test
launch preparations.
RDM, DORITOS,
and Avionics Bay
shall land without
damage.
Test The safe landing of
the RDM, DORITOS,
and Avionics Bay
will be tested during
full-scale launch.
Will be verified
during full-scale test
launch preparations.
Fin epoxy must be
able to withstand the
forces experienced
during flight to keep
the fins in their
proper position.
Analysis AeroFinSim was used
to confirm that the
stress experienced by
the fin epoxy does not
exceed the maximum
allowable stress at
any point during
flight.
Completed
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Section 6: Payload Criteria – Rover Deployment Mechanism (RDM)
6.1 System Overview
6.1.1 RDM Objective
The Rover Deployment Mechanism objective is to separate the nose cone from the airframe so
that the rover may deploy from the internal structure of the launch vehicle.
6.1.2 RDM Mission Requirements
1. The LBR team will deploy a customized rover from the interior of the launch vehicle
following a safe landing.
2. The deployment process will be remotely triggered.
3. The RDM system must be able to activate in any landing orientation.
6.1.3 RDM System Summary
Figure 6.1: CAD of full integrated RDM design.
The Rover Deployment Mechanism (RDM) will deploy the rover from the internal structure of
the launch vehicle. It must comply with the constraints of the rover design to efficiently utilize
space while remaining within the weight restrictions. Since the rocket will be landing in an
arbitrary position, the deployment mechanism must deploy regardless of directions and be
powerful enough to move the rover payload and separate rocket components apart regardless of
the terrain or obstacles. This system must also be capable of securing the rocket frames together
106 Long Beach Rocketry | CDR 2017 - 2018
during flight and rugged enough to withstand the impact of the rocket landing. To satisfy these
functional requirements, LBR has decided to employ an electro-mechanical system of
disconnecting rocket components to expose and deploy the rover from the rocket’s internal
structure.
RDM will be comprised of a system of rods, both threaded and unthreaded, motor bulkhead,
nose cone bulkhead, a nut-embedded rover-pushing-plate, a motor, and a control electronics
inside the electronic bay. Since the rover will be housed in the upper portion of the launch
vehicle, RDM team chose the nose cone and payload interface as an area to deploy the rover by
disconnecting the nose cone from the rest of the airframe. The rotation of the threaded rod on the
embedded nuts placed at the center of the nose cone bulkhead and rover-pushing-plate will cause
the nose cone and the rover to translate simultaneously. This will create an opening for the rover
to be pushed off the rods which completes the deployment process.
6.1.4 RDM Changes Since PDR
After the subscale launch, LBR noticed that the motor used, 32 RPM Planetary Gear Motor,
slowed down greatly when trying to break 3 shear pins. The LBR team feared that using this
motor during the full scale would not have enough strength to break the shear pins due to a nose
cone weight increase. As a result, rather than using a 32 RPM Planetary Motor for the full scale,
the RDM team decided to use a 118 RPM HD Premium Planetary Gear Motor w/ Encoder. This
motor’s stall torque is 958.2 oz-in which is significantly greater than the 472.1 oz-in of the 32
RPM Planetary Motor [6.1]. This is further reflected in the gear ratios of the motors, which are
71.165:1 and 369.595:1 respectively. Additionally, the 118 RPM HD Premium Planetary Gear
Motor will spin much faster than the 32 RPM motor, so the rover will be deployed faster than
during the subscale launch.
Other differences in the motors include differences in weight, unloaded current draw, stall
current draw, and recommended voltage supply. The respective specifications for the 118 RPM
HD Premium Planetary Gear Motor are 12.70 oz, 0.53 A, 20 A, and 6V- 12V. The specifications
for the 32 RPM Planetary Motor are 4.15 oz, 0.21 A, 4.9 A, and 3V-12V. The 118 RPM HD
Premium Planetary Gear Motor requires more voltage and draws more current, but provides the
torque and speed needed to break successfully deploy the rover.
To ensure that the rover would not get stuck at the end of the rod, rather than make the rover
contain threaded nuts, LBR decided to implement a threaded plate that pushes the rover, “rover-
pushing-plate”, in which the rover would rest on top of. Rather than have the threaded rod longer
than the un-threaded rods for the subscale launch, seen in Figure 6.2.1, RDM team decided to
make the un-threaded rods longer than the threaded rods for full scale, Figure 6.2.2, so that the
rover-pushing-plate would stay attached to the rods; clearing any protruding rods once it reached
the end of translation and the rover is pushed off without getting caught which prohibits the
navigation. The reason the preliminary design almost failed was because the nuts and bolts on
the nose cone bulkhead were protruding and the bear threaded rod interfacing the hex nut in the
rover increased its chances of getting stuck.
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Figure 6.2.1: Half section of RDM subscale
Figure 6.2.2: Half section of RDM full scale
6.2 RDM System Design
6.2.1 RDM Design Selection
To deploy the rover from the internal structure of the launch vehicle, LBR team considered two
design options to separate the nose cone from the airframe. For Option 1, the team considered
separating the rocket into three components: airframe, coupler, and nose cone. Three threaded
rods in a triangular configuration, each driven by a motor, will run through all three components
and an un-threaded rod will run through the center. The rover will have a hole in the center, and
it will sit in the coupler with the un-threaded rod running through the center. Aluminum
bulkheads with threaded nuts built in will be secured to the coupler and nose cone so that when
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the threaded rods rotate, these components can move linearly along the rods. Two bulkheads
within the airframe will create a compartment for three high-torque motors. The motors will spin
and first move the nose cone and coupler away from the airframe. The rover will follow the
coupler since it is contained within. The nose cone will fall off the threaded rods so that it is
completely separated from the launch vehicle. The coupler will approach the end of the threaded
rod, and as it does, the rover will fall off the center un-threaded rod. The back end of the rover
will catch the tip of the un-threaded rod, and as the coupler is retracted, the rover will be left
behind on the ground which will allow it to be deployed.
On the other hand, Option 2 is a simplified approach to Option 1, with the use of a single
threaded rod driven by a single motor, running through the center of the airframe, the rover, and
the nose cone bulkhead. Additionally, two un-threaded rods will also be secured parallel to the
threaded rod, but their function will be to stabilize the translating components and to prevent the
rover from spinning. A threaded nut will be attached to the nose cone bulkhead, as well as the
rover. When the middle threaded rod rotates, it will move the nose cone bulkhead, the rover-
pushing-plate, and the rover which will allow the rover to be deployed. LBR plans on
successfully deploying the rover from the airframe and ensuring that the motor has enough
torque to detach the nose cone from the airframe to break the shear pins after the launch.
Table 6.1: Comparison of features: subscale design (Option 1 vs. Option 2)
Features Option 1 Option 2
Efficient Use of Space 1 5
Weight 3 4
Complexity 2 3
Coding 2 4
Torque Capabilities 5 3
Price 2 5
Total 15 24
Table 6.1 shows a careful weighted evaluation of each rover deployment mechanism design.
Based off the comparison between the two designs, LBR found Option 2 to be the best design.
This design utilizes a single motor with one threaded rod and two un-threaded rods, yielding less
points of failure and adhering to weight restrictions than Option 1 while still supplying sufficient
torque to successfully deploy the rover from the internal structure of the launch vehicle.
6.2.2 Mechanical Hardware
To deploy the rover from the internal structure of the rocket, LBR decided to use a system of
rods, bulkheads, a rover-pushing plate, and a motor to separate the nose cone from the airframe
of the launch vehicle. The interface between the nose cone and the airframe, in which the RDM
creates a gap, will allow the rover to escape the launch vehicle and drive away autonomously. In
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addition, the triangular shape of the rover complements the RDM design so that regardless of the
configuration of the rocket upon landing, the rover always deploys and lands on wheels.
The electronics of the subsystem will be housed in a 3D printed electronics bay to be discussed
in Section 6.2.3: “Electronics Bay”. Two 6061 aluminum ¼-14.5 in. un-threaded rods will
connect to the motor bulkhead and electronics bay by built in printed couplers which will be 4.3
in. apart. Perpendicular screws will secure the rods into their connecting cylinders. The purpose
of these rods will be to secure the rover into the airframe and to prevent the rover-pushing-plate
that will push the rover out of the rocket from rotating while translating the rover along the rod.
The electronics bay will also house the system’s motor. A Grade B7 Medium Strength Steel ¼”-
20 threaded rod that is 14 in. long will connect to the motor, using a ¼ in.- 6mm CNC Motor
Shaft Coupler D20 L25 Flexible Coupling, and it will run through the center of the rocket. Figure
6.3 shows the three rods assembled into the system. The threaded rods is coupled to the motor,
while the un-threaded rods are coupled to the motor bulkhead which secures to the face of the
electronics bay.
Figure 6.3: Figure of rod assembly of rods in full scale
The motor bulkhead with the un-threaded rods will attach directly to the electronics bay. Next,
the rover-pushing plate with a nut in the center and holes for the un-threaded rods will sit on the
rods next to the motor bulkhead. The threaded 5.539 in. diameter pushing-plate consists of two
halves, with a 0.236 in. threaded hole for the rod, which entraps a 0.445 in. hex nut, seen in
Figure 6.4.1. LBR decided that entrapping a nut, rather than epoxying a nut onto the face of the
plate, was a safer option. Entrapping the nut stabilizes the plate and rover during the unthreading
process and reduces the likelihood of failure in case the nut loosened in the epoxy. Four M3-15
mm bolts and nut pairs will screw into four 0.126 in. holes. In addition, located near the outer
surface of the plate are three 0.276 in. holes, which are 120 degrees apart from each other, which
are where the two ¼ -14.5 in. un-threaded rods will bypass. Having the un-threaded rods go
through the rover-pushing plate stabilizes the translation, which pushes the rover straight out.
This reduces the chances of the plate tilting and wobbling, which prevents the rover from being
pushed out unevenly and getting stuck onto the un-threaded rods towards the end of deployment.
The first plate is 0.236 in. thick, while the second is 0.118 in. so the hex nut can rest on the first
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plate without bending. Figures 6.4.2 and 6.4.3 show the dimensions of the mirror haves of the
pushing-plate that will be used to entrap the nut.
Figure 6.4.1: Assembly of rover-pushing-plate (Exploded on left)
Figure 6.4.2: Drawing of front half of rover-pushing-plate
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Figure 6.4.3: Drawing of back half of rover-pushing-plate.
Next the rover with a hole in the center will sit on the threaded rod in between the un-threaded
rods. As the motor spins the threaded rod, the pushing-plate will translate linearly along the rods.
Since the rover sits in front of the pushing-plate, the pushing-plate will cause the rover to linearly
translate as well. Figure 6.5 shows the location of the pushing-plate and rover in the RDM upon
initial assembly in the full scale RDM assembly.
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Figure 6.5: Complete full scale RDM assembly with rover
The final bulkhead is secured into the nose cone coupler. This bulkhead’s diameter is 5.775 in.
so it fits inside the nose cone coupler. The two pieces are bound together by four M3-20mm bolts
and nuts pair as shown in Figure 6.6. The assembled piece is then pressure fitted to the nose cone
coupler. The outer diameter of the nose cone coupler is 5.998 in. so that it fits within the nose
cone, and the inner diameter is 5.775 in. so that it can accommodate the bulkhead. This nose
cone design is the same as the subscale prototype nose cone, but rather than 3.998 in. in
diameter, the diameter is 5.998 in., seen in Figure 6.6. The dimensions for the full scale nose
cone bulkhead can be seen in Figures 6.7.1 and 6.7.2.
Figure 6.6: Assembly of full scale nose cone bulkhead (Exploded on left)
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Figure 6.7.1: Drawing of the front half of nose cone bulkhead.
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Figure 6.7.2: Drawing of back half of nose cone bulkhead.
Figure 6.8 shows the nose cone assembly. The coupler couples the nose cone to the airframe of
the launch vehicle. The holes in the nose cone match the rods, and the center hole has an
entrapped nut so that the nose cone can translate as the threaded rod spins. Three 4-40 Nylon
shear pins are screwed through the airframe and coupler so that the nose cone is secured to the
launch vehicle during flight. Securing of the nose cone, the RDM system is ready to deploy the
rover. Figure 6.1 shows the completely assembled RDM system.
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Figure 6.8: Assembly of full scale nose cone and nose cone bulkhead
LBR will remotely trigger the deployment of the rover by flipping a switch on a controller that
team members will hold. When the deployment commences, the motor and the threaded rod will
begin to spin. The rover-pushing-plate pushes the rover simultaneously with the bulkhead-nose
cone assembly both translating along the rods. The nose cone will reach the end of the rods and
stop translating. When the rover catches up and touches the nose cone bulkhead, it will push the
nose cone off the rods and out of the way. Finally, the rover-pushing-plate will push the rover
completely off the rods so that it will be able to autonomously drive and fulfill the project
requirements. The un-threaded and threaded rods will be different lengths to ensure that the rover
fully deploys; the un-threaded rods will be longer than the threaded rod, and the difference in
length will be 0.354 in., or the width of the rover-pushing-plate. Table 6.2 displays the estimated
mass of all mechanical RDM components.
Table 6.2: Estimated mass of mechanical components: full scale
Component Mass (oz)
Rod/Coupler Pair 6.24
Un-threaded Rod (x2) 2.89 (5.78)
Motor Bulkhead 2.33
Nose cone Bulkhead 4.10
Rover-Pushing-Plate 3.41
Electronics Bay 29.76 (1.86 lbs)
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6.2.2.1 Mechanical Design Validation
To validate the chosen design, RDM team verified using Autodesk Inventor Power Calculator
variables related to loading on screw emphasizing material design and screw diameter design.
For the specifics of the design, RDM team chose ¼-20 in. threaded rod-nut pair made of Grade
B7 Steel (also known as Steel SAE 4140). The team also anticipated that the peak load for the
RDM system will come from breaking the shear pins, specifically the planned application of
three 4-40 shear pins which rated at an average of 71 lbf to break [6.2]. By choosing the material
properties of Steel SAE 4140, ¼ in. diameter rod with 1/20 in. pitch, 0.236 in. nut length, and
maximum axial loading of 77.3 lbf, the result yielded a Factor of Safety (FOS) of 2 which means
the design complies twice what it was intended for. Table 6.3 shows the figures and results of the
calculations taken directly from the report that came out of the power screw calculator.
Figure 6.9: Screw and nut diagram with parameter designation
Table 6.3: Power screw calculator parameters and results
Guide Loads
Type of
Transmission
Rotation --> Shift Maximum Axial Force F 77.300 lbf
Type of Strength
Calculation
Material and Screw
Diameter Design
Maximum Torque T 4.947 lbf- ft
Load Input Force and Torque Thread Friction Factor f1 0.150 ul
Material Screw
Material Steel SAE 4140 Thread Diameter D 0.2500 in
Allowable
Thread Pressure
pA 10000.000 psi Pitch P 0.0500 in
Modulus of
Elasticity
E 30000000 psi Mean Screw Diameter ds 0.2188 in
Yield Strength Sy 238000 psi Min. Screw Diameter dmin 0.1812 in
Safety Factor ks 2.000 ul Nut Height H 0.236 in
Factor for End Conditions N 1.000 ul
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Max. Length L 12.000 in
Results Summary of Messages
Reduced Length Lred 12.000 in 10:05:51 PM : Calculation indicates design
compliance! Efficiency η 0.323 ul
Slenderness
Ratio
λ 219.429 ul
Pressure Stress σt 2995.943 psi
Torsional Stress τk 30792.047 psi
Reduced Stress σred 53417.471 psi
Rankin Critical
Stress
σR 5994.540 psi
Euler Critical
Stress
σE 6149.427 psi
Johnson Critical
Stress
σJ 999999.000
psi
Calculated
Thread Pressure
pc 1016.783 psi
Calculated
Factor of Safety
kv 2.001 ul
Helix Angle α 3.64 deg
Check Calculation Positive
The design also complies with the motor specifications that RDM chose during the subscale
launch and tests (at 2.349 lbf-ft = 450 oz-in). The new motor with the maximum torque of 4.947
lbf-ft (~ 950 oz-in) exceeds almost twice the subscale motor, thus, implies that the new motor
will be able to drive the RDM system.
6.2.3 Electronics Bay
The LBR team has designed an electronics bay integrated around the DC motor to house all the
electrical components efficiently. The electronics bay has been designed to have a secure
position for each of the components to prevent clutter and keep the circuit in place throughout
the mission. A clamping mechanism has been designed to secure the DC motor into the center of
the electronics bay and is held down with a bulkhead which will have the two un-threaded rods
used to assist in rover deployment.
The bulkhead consists of 4 parts: an electronics bay, detachable motor bracket, motor bracket
cover and the motor bulkhead cover. The electronics bay, 6 in. diameter and 4.42 in. in length,
will house a 118 RPM HD Premium Planetary motor, two 11.1 V Li-Po Batteries, Arduino
Nano, H-Bridge, and a 2.4 GHz Digital Receiver. Four M3 bolts will fasten the components
together. When assembled, the bulkhead utilizes 5.24 in. of the airframe, including the two
118 Long Beach Rocketry | CDR 2017 - 2018
cylindrical 3D printed couplers for the un-threaded rods. Figure 6.10.1 shows the electronics bay,
detachable motor bracket, motor bracket cover, and motor bulkhead. Figure 6.10.2 shows their
dimensions. The motor bracket (second component to the left) clamps on the motor. The 1.590
in. ×1.700 in. motor bracket has a 1.4 in. hole in the center for the motor, and when snapped in,
the cover secures the motor bracket in the slot within the electronics bay. Finally, the motor
bulkhead seals the components together.
Figure 6.10.1: Assembly and exploded view of full scale electronics bay
Figure 6.10.2: Drawing of full scale electronics bay components.
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The electronics bay portion of the motor bulkhead contains 1.4 in. diameter hole to entrap the
body of the motor and 1.763 in.×0.760 in. slots for the batteries on opposite sides of the motor.
On the bottom of the bulkhead, in between the two 3.125 in. long extruded components for the
motors, LBR created a cap with a 1 in.×0.5 in. slot to enclose the end of the motor and then
extruded a 2.673 in. long rectangular filament to screw in the H-Bridge, Arduino Nano, and
Receiver. LBR decided that creating compartments for the motor and batteries, but screwing in
the rest of the components was the optimal design for the electronics bay. Creating customized
compartments for all the components would result in an unnecessarily longer motor bulkhead to
prevent dismantling when wiring all the components together.
The final piece for the motor bulkhead is a flat 6 in. diameter cylinder, which is 1.50 in. thick,
contains 3D printed cylindrical couplers, with ¼ in. holes, for the two ¼- 14.5 in. un-threaded
rods. To tighten the un-threaded rods to the couplers, LBR will screw in two M3.5 set screws on
the side of the couplers. LBR designed the final piece as a cap to separate the electronics from
the rover and the rover-pushing plate. This prevents the electronics components from entangling
with other RDM components. Figure 6.11 shows the electronics bay full assembly. Table 6.4
shows the estimated mass of the electronics components.
Figure 6.11: Full scale electronics assembly with components
Table 6.4: Estimated mass of full scale electronics
Component Mass
118 RPM HD Premium Planetary Gear Motor
w/Encoder 12.70
Arduino Nano Microcontroller 0.21
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L298N H-Bridge 0.85
2.4 GHz Digital Receiver 0.63
11.1V Li-Po Battery (x2) 7.94 (15.88)
Limit Switches (x2) 0.30 (0.60)
6.2.3.1 Motor
The team originally intended to use a 32 RPM planetary geared motor with encoder for the rover
deployment mechanism. This motor was used and tested for the subscale model to break through
four shear pins and deploy the rover, but through further investigation the team has noticed a
potential point of failure that has been considered for design revision. The DC motor functions at
a low current when operating, but once the motor needs to break the shear pins the current draw
spikes up to about 1.65 Amps as shown in Figure 6.12. The H-Bridge the team will be using has
a maximum current output of 2 A but the motor has a stall current of 4.9 A and 472.1 oz-in. stall
torque. Since the full-scale launch vehicle is larger the motor will need to push a larger load
which may exceed the motor driver current output and potentially prevent the rover from
deploying.
Figure 6.12: Graph of current drawn to break three shear pins using a 32 RPM motor
To avoid this problem and speed up the rover deployment process, the team has decided on using
a larger planetary motor with encoder which can provide more torque. The new motor is 118
RPM which is faster than the previous motor but has a potential 20 A stall current at 958.2 oz-in
of torque, which is a higher current draw but around double the torque of the previously
considered motor which will guarantee the deployment of the rover. Further test must still be
121 Long Beach Rocketry | CDR 2017 - 2018
conducted to find the new current draw when attempting to break the shear pins, but if it exceeds
the current limitation of the motor driver the team will seek a replacement motor driver capable
of providing more current.
6.2.3.3 Battery
Due to the high power demand of the DC motor, the battery in the RDM must be capable of
providing power throughout the whole mission which includes the system idling while preparing
for launch and the deployment of the rover. On idle the team can approximate a current draw of
500mA for 1.5 hours to be safe, and according to Figure 6.12 the system will draw roughly 1.6 A
for 5 seconds but due to the full scale, a larger load, the current draw will be estimated at 3 A.
With the DC motor spinning at 118 RPM the rover will take roughly 4 minutes to deploy under 2
A of current draw. These calculations have been overestimated to allow for tolerance in both
power consumption and mission duration. Through using the same formulas as seen in the rover
payload section, we can approximate the battery necessary to complete the mission.
𝐵𝑎𝑡𝑡𝑒𝑟𝑦 𝐶𝑎𝑝𝑎𝑐𝑖𝑡𝑦 𝑖𝑛 𝑚𝐴ℎ =(𝐵𝑎𝑡𝑡𝑒𝑟𝑦 𝐿𝑖𝑓𝑒 (ℎ𝑟𝑠)) ∗ (𝐿𝑜𝑎𝑑 𝐶𝑢𝑟𝑟𝑒𝑛𝑡 𝑖𝑛 𝑀𝑖𝑙𝑙𝑖𝑎𝑚𝑝𝑠)
0.7(43)
(1.5 ℎ𝑜𝑢𝑟𝑠)∗(500 𝑚𝐴)
0.7 +
(0.002 ℎ𝑜𝑢𝑟𝑠)∗(3000 𝑚𝐴)
0.7 +
(0.07 ℎ𝑜𝑢𝑟𝑠)∗(2000 𝑚𝐴)
0.7 = 1280 mAh
The team has originally designed the electronics bay for the subscale to accommodate two 11.1V
1100mAh batteries which were connected in parallel to provide 2200mAh to endure the entirety
of the mission. Though the batteries had a high capacity, they had a low discharge rate which
may be a constraint with the new motor. To fix this issue, the team will be using a single 11.1V
1800mAh battery with higher discharge “C” value capable of powering the larger motor.
6.2.3.4 Control
After the launch vehicle lands, LBR will initiate the rover deployment using a 2.4 GHz
transmitter. The signal received by the 2.4 GHz digital receiver inside the RDM was processed
by the Arduino Nano which sends information to the H-Bridge to spin the motor
counterclockwise to translate the rover and nose cone forward. Using a rotary encoder, attached
to the 118 RPM Planetary Motor, LBR receives data regarding speed and distance of the
threaded rod. When the load is increased, the encoder sends a signal back to the Arduino Nano,
which tells the motor to reduce its speed to increase the torque and avoid motor burnout. Figure
6.13 shows the electric schematic of the electronic components in the RDM.
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Figure 6.13: Electric schematic of full scale RDM
6.2.4 Launch Vehicle Integration
The RDM sits inside the front tip of the airframe and extends into the nose cone. The base of the
electronics bay is 17 in. from the top of the 36 in. airframe. The 14.5 in. un-threaded rods and 14
in. threaded rod stick approximately 4 in. into the nose cone. Figure 6.14.1 shows the RDM
integrated into the airframe and nose cone. The RDM system components are arranged as shown
in Figure 6.14.2. The first half of the separating piece will include the nose cone, followed by the
nose cone coupler which also accommodates the nose cone bulkhead. The second half portion of
the RDM system is integrated in the payload airframe section. This includes the rods, rover-
pushing-plate, motor bulkhead, and electronic bay. The nose cone bulkhead will be pressure
fitted to the nose cone coupler. The motor bulkhead will be fastened by 4-40 bolts at 4 locations
around the perimeter from the outside of the airframe. The electronic bay will be fastened by 4
M3-20mm bolts and nuts pairs to the motor bulkhead. To secure the nose cone to the airframe,
LBR will use three 4-40 Nylon shear pins around the nose cone coupler and airframe
interconnect.
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Figure 6.14.1: Location of the RDM section relative to the rocket
Figure 6.14.2: Internal components of RDM system: Labelled
6.3 Subscale Prototype
For the RDM subscale system, LBR used one ¼-12 in. threaded rod and two ¼-8 in. 3-D printed
un-threaded rods to prevent the air frame from rotating as it separated from the nose cone. The
three rods were attached to a 3D printed bulkheads, 4 in. in diameter and 0.63 in. thick.
Additionally, the system contained a plain 3D bulkhead in the nose cone, 3.8 in. in diameter and
0.63 in. thick, with a trapped nut in the center. When activated, the motor spins the tip of the ¼-
12 in. threaded rod connected to the trapped nut translating the nose cone forward. Using ⅛ in.
screws, LBR externally secured both bulkheads. In addition, the subsystem contained a custom
3D printed electronics bay which housed the electronics operating the RDM subsystem, which
was attached to the bottom of the motor bulkhead.
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6.3.1 Nose Cone Bulkhead
The 3D printed nose cone bulkhead was flushed against and press-fitted at the bottom edge of the
nose cone. RDM chose to press-fit the nose cone bulkhead because in the final assembly, the
nose cone coupler was secured by 3 shear pins. Both the threaded and un-threaded rods will run
through the payload section and nose cone, i.e. starting from the motor bulkhead to nose cone
bulkhead. There are three 0.276 in. holes that are separated 120 degrees apart in which any two
of these holes will be where the un-threaded rods will pass through. The four 0.128 in. holes will
be used to fasten together the other half of the nose cone bulkhead to trap a flange hex nut inside
using four M3-20mm bolts and nuts. The bulkhead contained a trapped ¼ in. flange hex nut,
mating the threaded rod tip to the nut, which threaded out the nose cone. Figure 6.15 shows the
components of the nose cone bulkhead. When the motor spun counterclockwise, the nose cone
separation from the air frame created an opening for the rover to deploy.
Figure 6.15: Assembly of subscale nose cone bulkhead (Exploded on left)
6.3.2 Motor Bulkhead
The bulkhead housing the 32 RPM Planetary Gear Motor and the two ¼-8 in. un-threaded rods
externally screwed into the airframe using ⅛ in. screws. A customized square 1.57 in. detachable
motor bracket which securely clasps into the motor bulkhead. The detachable bracket contained
two halves, connected by tightening two pairs of M3 bolts and nuts 20 mm in length, located on
both sides of it to a hex nut and a washer. When tightly screwed in, the washer prevents cracking
by evenly distributing the weight. Additionally, LBR 3-D printed a motor cap, with a centered
1.02 in. hole for the aluminum helical flexible shaft coupler and the ¼-12 in. threaded rod
enclosing motor bracket for additional security. Figures 6.16.1 and 6.16.2 show the detachable
motor bracket and motor cap, and motor bulkhead. The two 0.71in. diameter printed cylindrical
coupler, part of the motor bulkhead, with ¼ in. holes allow the RDM team to secure the two 8 in.
un-threaded rods. LBR considered 3D printing the motor bulkhead fully rather than creating
detachable parts, but agreed that separating the components was safer in case the length needed
to be altered for future redesigns.
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Figure 6.16.1: Assembly of top of subscale motor bulkhead
Figure 6.16.2: Exploded view of top of subscale motor bulkhead
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6.3.3 Electronics
The RDM electronics bay, shown in Figures 6.17.1 and 6.17.2, housed the following electrical
components: a L298N H-Bridge, Arduino Nano Microcontroller, two 11.1 V Li-Po batteries, and
2.4 GHz Digital Receiver. The electronics bay lies right below the bulkhead containing the motor
and the two ¼-8 in. un-threaded rods. Since the end of the 32 RPM Planetary Gear Motor
protrudes out of the bulkhead, LBR created a pocket on the electronics bay to enclose the end of
the motor. The electronics bay was 3.898 in. in diameter and was fastened by four ⅛ in. screws
from the outside of the airframe. In addition, in order for RDM system electronics to be
activated, a rotary main switch was used and secured by drilling a ½ in. hole in the 4 in. motor
bulkhead. This allowed the main switch to be easily accessible to turn on the system prior to
rocket assembly, ensuring the battery life of the RDM system electronics would not die before
launch. In order to reduce the size of the subsystem, LBR decided that the best electronics bay
design would have to be a cylindrical in shape which would allow for electronics to be mounted
in various positions and levels to maximize space. Through creating a custom electronics bay the
team was able to connect and securely mount the H-Bridge, Receiver and Arduino with 4-40
screws as well as two Li-Po batteries as shown in Figures 6.18.1 and 6.18.2. To ensure that the
wires wouldn’t get tangled and disconnect, thin strips of electrical tape were used to hold the
wires in place. Table 6.5 shows the mass of the electronic components used in the subscale
RDM.
Figure 6.17.1: Bottom view of subscale RDM electronics bay
127 Long Beach Rocketry | CDR 2017 - 2018
Figure 6.17.2: Top view of subscale RDM electronics bay
Figure 6.18.1: Top view of subscale electronics bay with components
128 Long Beach Rocketry | CDR 2017 - 2018
Figure 6.18.2: Side views of subscale electronics bay with components
Table 6.5: Mass of subscale electronic components
Component Mass (oz)
Motor 4.48
Arduino Nano Microcontroller 0.21
L298N H-Bridge 0.85
2.4 GHz Digital Receiver 0.63
11.1 V Li-Po Battery 7.94
Rod/ Coupler Pair 6.24
Threaded Rod 2.89
Un-threaded Rods (x2) 2.89
3 Bulkheads 13.86
Total 39.99 (2.5lb)
6.6 Team Derived Requirements – RDM
As described in the NASA Student Launch Handbook in Experiment Requirements section 4,
each team can choose one of three design experiments to incorporate into their launch vehicle.
LBR has chosen Option 2, to build a deployable rover, so the team will follow the requirements
129 Long Beach Rocketry | CDR 2017 - 2018
in section 4.5. The first of these requirements is to design a custom rover that will deploy from
the internal structure of the launch vehicle. LBR has separated this design objective into the
RDM subsystem, where it must also follow the second requirement of deploying the rover
through a remotely activated trigger by the team upon landing the launch vehicle. Once the rover
is deployed, it must autonomously traverse at least 5 feet in any direction away from the launch
vehicle. Once the rover achieves this objective and comes to its destination, it must deploy a set
of solar panels to simulate the charging of the vehicle. The rover payload and deployment
mechanism have been designed to accomplish the NASA experimental objectives to ensure a
functional design capable of completing the mission.
Table 6.7: NASA Experiment Requirements
Requirement Verification
Method Verification Plan Status
Teams will design a
custom rover that
will deploy from the
internal structure of
the launch
vehicle.
Test
The team will complete a full-
scale rover that must be capable
of autonomously traversing
rough terrain and deploy solar
cells. Rover will be placed in
launch vehicle and deployed
through the RDM.
Subscale prototype
rover deployment
was successful. Full-
scale rover test will
be verified by
February.
At landing, the team
will remotely activate
a trigger to deploy
the rover from the
rocket.
Test
Team will flip a switch on a
remote to activate deployments
mechanism. Rover must be
completely deployed and
capable of driving away even
under non-ideal conditions.
Subscale RDM
deployed rover after
launch. Full scale
RDM will be
verified by
February.
After deployment,
the rover will
autonomously move
at least 5 ft. (in any
direction) from the
launch vehicle.
Test
Rover must autonomously
choose best direction to go and
maintain a straight heading until
the distance surpasses 5 feet.
Will be verified by
February.
Once the rover has
reached its
destination, it will
deploy a set of
foldable solar cell
panels.
Test
Rover must autonomously know
when to deploy solar panels and
completely open them up
regardless of the position the
rover is in.
Will be verified by
February.
130 Long Beach Rocketry | CDR 2017 - 2018
Table 6.8: Team derived requirements verification status and plan (RDM)
Requirement Verification Method Verification Plan Status
RDM components
must be removable
from the launch
vehicle.
Demonstration
The RDM electronics bay
was designed to easily
access the electronic
components if necessary.
The components will be
secured using screws. Once
the assembly is completed,
the RDM team will perform
this demonstration.
Will be tested
by mid-
February after
the assembly of
the electrical
components.
The RDM system,
can respond to
signal, without error.
or delay, from at
most a mile away.
Test/Inspection
Following the assembly of
the electronics the RDM
team members team will
stand a mile with the 2.4
GHz Radio Transmitter, and
observe the RDM system
for errors or delays.
Will be tested in
early February
2018 after the
coding of the
electronics bay.
Electronics must be
properly wired. Inspection
Following the design of the
motor bulkhead, the RDM
team will assemble the
electronic components into
the electronics bay
components of the
bulkhead. The RDM will
carefully inspect the wiring
to ensure that the wires are
properly soldered and
connected.
Will be tested in
mid-February
2018 after the
assembly of the
entire
electronics bay.
The 2.4 GHz Radio
Transmitter and
Receiver must be
able to retain charge
for 1.5 hours.
Test
Leave the transmitter
operating on idle and test to
determine if the RDM
system can operate after at
least 1.5 hours, then
deploying the rover.
Will be verified
by the end of
January 2018.
The rover can fully
disconnect from the
rods and rover-
pushing plate.
Test
After assembling the RDM
mechanical hardware, place
a simulated rover on top of
the rover-pushing plate and
activate the system. When
Will be verified
after the
complete
assembly of the
RDM
131 Long Beach Rocketry | CDR 2017 - 2018
the rover-pushing plate
reaches the end of the
threaded rod and deploys
from the airframe, this will
verify this requirement.
mechanical
hardware and
electronics by
the end of
February 2018.
After dropping the
RDM system, the
nose cone remains
attached to the
airframe.
Test
To verify this, RDM will
perform a drop test on the
completed RDM system
when the electronics and
mechanical hardware are
secured in the airframe.
Will be verified
after the
assembly of the
electronics bay
in mid-
February 2018.
Identification of
designated channel
values when the 2.4
GHz Radio
Transmitter-Receiver
is turned on.
Test/Inspection
RDM will utilize color
coding and tagging for easy
identification after the
channels are verified
through serial monitor
output.
Will be tested in
early February
2018 after the
coding of the
electronics bay.
Sufficient space
during airframe and
nose cone separation
for rover
deployment.
Test
RDM team will perform
several deployment tests to
ensure a smooth release of
the rover to the field
Will be verified
in late February
2018 after the
complete
assembly of the
RDM system.
The 118 RPM HD
Planetary Gear
Motor must be
capable of breaking
three shear pins.
Test/Analysis
The test has already been
made for the 32 RPM motor
breaking four 4-40 shear
pins. RDM team will repeat
the same test for the new
motor.
Will be verified
in late February
2018 after the
complete
assembly of the
RDM system.
Electronics remain
intact throughout
launch.
Test/Inspection
LBR will perform a
vibration test on the
connected airframe and
rover. After the test, the
RDM team will disassemble
the payload and inspect the
electronics bay. Following
the inspection, the RDM
team will retest the system
to ensure that the electronics
survived the vibration test.
Will be tested in
mid-February
2018.
132 Long Beach Rocketry | CDR 2017 - 2018
Threaded and
unthreaded rods
remain linear
throughout entire
mission.
Test/ Inspection
One of the main concerns
was that the rover’s weight
would bend the end of the
rod. Following multiple
tests, the team will inspect
the rods to verify the
requirement.
Will be verified
in late February
2018 after the
complete
assembly of the
RDM system.
Threaded rover-
pushing plate pushes
rover out evenly.
Test
Following the electronics
assembly, the RDM team
will test to ensure that the
rover is deployed evenly
and doesn’t get caught on
the un-threaded rods.
Will be verified
in late February
2018 after the
complete
assembly of the
RDM system.
A custom rover will
deploy from the
internal structure of
the airframe.
Test
In order to verify this
requirement, RDM will
assemble the electronics bay
and couple the 14 in.
threaded rod to the 118
RPM motor to ensure that
motor has enough torque to
separate the nose cone and
airframe.
Will be verified
in late February
2018 after the
complete
assembly of the
RDM system.
The 118 RPM HD
Planetary Motor
must supply at least
50 lbf of force and
0.2083 lbf-ft of
torque to shear a
minimum of two 4-
40 shear pins.
Verify/Test
RDM team will put to test
this motor loaded with two
to four shear pins through an
existing setup from the
subscale launch. The team
will only verify the result
since a 32 RPM (~ 450 oz-
in) half the torque of 118
RPM HD motor
successfully break four 4-40
shear pins
Will be tested in
mid-February
2018 after the
assembly of the
entire
electronics bay.
Current drawn by
motor required to
break shears pins
must not exceed the
maximum current
output of the H-
Bridge.
Test
Connect the motor to an
Ammeter and measure the
current of RDM breaking
three shear pins. Using this
data, compare the current to
the maximum current of the
H-Bridge.
Completed for
32 RPM motor
using L298N H-
Bridge. Test
must be
repeated for 118
RPM motor on
the 1st of
February
133 Long Beach Rocketry | CDR 2017 - 2018
A single 11.1V
1800mAh Li-Po
batteries must be
able to supply
enough power to
power the 118 RPM
HD Planetary Motor.
Analysis/ Test
Due to the increase in torque
and RPM of the new motor,
the team will conduct up to
3 full deployments using a
single battery with a much
higher ‘C’ ratings than the
two batteries used in the
subscale.
Will be tested in
mid-February
2018 after the
assembly of the
entire
electronics bay.
Encoder must be able
to detect variation in
speed when different
amounts of forces are
felt in order to avoid
motor burnout.
Test/Monitor
RDM team will perform an
actual deployment tests
while the Arduino
microcontroller is plugged
in to the computer to
monitor and log the data
coming from the encoder
Will be tested in
early February
2018 after the
coding of the
electronics bay.
Arduino Nano must
be able to receive
information from the
receiver and rotates
the motor enabling
rover deployment.
Test
RDM team will perform
multiple tests after
assembling the electronics
to verify that the motors run
during each trial.
Will be tested in
mid-February
2018 after the
assembly of the
entire
electronics bay.
Installation of limit
switches: (1) to
prevent over
assembly of the load
to the system, (2)
ensure that the rover
is released to ground.
Test
RDM team will survey/test
installation on the
appropriate location within
the space between the motor
bulkhead and rover-
pushing-plate for the
location for the switches.
Will be tested in
first week of
February.
134 Long Beach Rocketry | CDR 2017 - 2018
Section 7: Payload Criteria – Dynamically Oriented Rocket
Integrated Triangular Object (D.O.R.I.T.O.)
7.1 System Overview
7.1.1 DORITO Objective
The payload of the Long Beach Rocketry launch vehicle will be an autonomous, deployable
rover that must be capable of traveling 5 feet from the rocket and then deploy solar panels. In the
Preliminary Design Review, LBR presented numerous design choices to be considered for the
design of the payload. The team chose on a final design that is light yet robust for impact,
capable of traversing rough terrain, and is easily and effectively implemented into the rover
deployment mechanism.
7.1.2 DORITO Mission Requirements
1. The rover will be a custom designed autonomous vehicle that will be stored within a
payload bay on the launch vehicle.
2. Once the launch vehicle has landed, the LBR team will remotely initiate an autonomously
deployment of the rover through the RDM which will traverse across a rugged terrain
until it is at least 5 feet away from the rocket.
3. When the rover has surpassed the required distance, it will stop and deploy a set of solar
panels which will be used to charge the power supply onboard the rover.
7.1.3 DORITO System Summary
During the PDR stage, team members considered two main designs for the rover located inside
of the rocket. The challenge for the rover design was to create a vehicle that is efficiently stored
within the payload bay of the rocket that maximizes the use of available space, so it must be
designed with the internal structure of the deployment system in mind. This vehicle must be
capable of driving in rough terrain regardless of the orientation the rocket lands due to the rover
being deployed in an arbitrary manner. It must also withstand forces of being pushed out of the
payload against the nosecone yet light enough that it doesn’t inhibit the performance of the
rocket. Through analysis of the two potential rover designs, the team has decided to go with the
triangular rover design due to its ability to perform well in the requirements mentioned above.
The team has implemented a subscale model of the triangular rover with limited functionality in
order to test the design concept during the subscale launch vehicle test launch. Through building
a functional scaled down prototype the team was able to perform a systems test and an attempt to
make the rover traverse through harsh terrain, where the team was able to collect valuable data
on design flaws and potential design improvements which will be corrected for to increase the
likelihood of a successful mission.
The team will use a microcontroller in the rover which will be used to control the motors for the
drive mechanism and a servo which will be used to deploy the solar cells. The rover will know
its position and distance with respect to the launch vehicle by using an inertial measurement unit
to know the heading and orientation, as well as a wireless transceiver module which will be used
135 Long Beach Rocketry | CDR 2017 - 2018
to communicate back to the rocket to provide a real-time location system. A control system
including yaw suppression is used to keep the rover in the correct heading in case it strays off
course. This will ensure the rover surpasses the minimum 5 feet of distance it needs to travel
from the rocket.
7.2 Rover Design
Based on design analysis and the NASA interview’s positive opinions of the design, the team has
decided to use the triangular rover design. Various tests are performed on the rover to achieve
the best design based on wheels, sensors, and propulsion combinations.
Figure 7.1: LBR rover design
7.2.1 Implementation into RDM
The rover will be deployed by the RDM with a threaded rod that will run through its
body. Originally, LBR intended to use threaded hex couplers that would allow the rover to be
moved from the rocket as the threaded rod rotates. Through testing of the subscale rover, the
team was able to find an issue during deployment from the launch vehicle. The rover was unable
to be completely deployed from the rocket due to part of the hex coupler remained on the rod. In
the new RDM design, the rover will be deployed by a circular panel that will push the rover out.
The circular panel will be threaded and will move along the threaded rod of the deployment
mechanism which must still go through the entirety of the rover. It is therefore unnecessary for
the center driveshaft to be threaded, and LBR will use a hollow unthreaded shaft instead to drive
the external gearbox from the internal motors.
7.2.2 Center Driveshaft
LBR considered two design choices for the center drive shaft of the rover, an aluminum hex
coupler and a hollow aluminum tube. Long Beach Rocketry has decided to use the hex coupler
136 Long Beach Rocketry | CDR 2017 - 2018
design due to its higher tensile stress, making it more favorable than the aluminum shaft. The
payload will be seated vertically into the rocket, and will be under intense compressive forces
along the shaft during the launch sequence. In order to ensure the safety and survivability of the
rover for post landing deployment, LBR must make high tensile strength a priority. Finding a
proper hex coupler that fits the requirements of the other components of the rover could prove
challenging; however, LBR has access to the tools necessary to machine this part. This item will
be hollowed out and cut to be used as the center driveshaft.
Figure 7.2: Rover center driveshaft
The center driveshaft will have an outer diameter of 0.5 in and an inner diameter of 0.3126
in. The length will be 4 in, 3 in of which will be used to hold the wheels. Two of these
components will be used for the driveshaft: one for each side of the payload; additionally, this
material will be used on each of the three wheels on both sides of the payload. These six
components are shafts that hold the wheels in place.
The hex shaped driveshaft will transfer power to the rover gearbox from the motors more
efficiently as compared to a circular shape which would not allow for enough grip for the gears
to clamp and "lock" onto the driveshaft. A threaded driveshaft will also put unnecessary stress on
the drivetrain since the drivetrain will counter the RDM's unthreading using the nature of the
high gear ratios of the drivetrain. Also, clamping hubs to hold the shaft in place will face
unnecessary perpendicular force to keep the shaft in place. The hex coupler is important for the
functionality of the bogie system of the payload. In order for this system to work, it will need a
pivot point. The center driveshaft will move along with the body of the rover which it is attached
to and therefore can act as the pivot point for the bogie system.
It is important to consider whether the shaft will be able to handle the torsion stresses caused by
connecting the main body to the gearbox. The general equation for max shear stress due to
torsion for a hexagon cross section beam is:
𝜏𝑚𝑎𝑥 =𝑇
𝑠3(44)
𝜏 is shear stress, T is torque, and s is the diameter of the cross section. The output torque of the
motors is 159.7 oz in [7.11], and the distance, s, is 0.5 in. Solving for𝜏𝑚𝑎𝑥:
𝜏𝑚𝑎𝑥= 159.7
0.53= 1277.6 oz/in2
137 Long Beach Rocketry | CDR 2017 - 2018
This is within the allowable shear stress of the cross sectional beam.
7.2.3 Drivetrain
LBR made the decision to implement a gearbox on the rover to transfer the power from the
motors to the wheels. With the gearbox, there will be many possible combinations of gears and
different orientations for the motors. Since the rover will traverse through rough terrain, a setup
with the most torque will be ideal for the rover since speed is not a priority. The desired torque
setup is a small center gear and large planet gears, but large planet gears will impede ground
clearance. Therefore, LBR has decided to use similar sized gears all around as a compromise to
torque and ground clearance. The center gear has 35 teeth while its surrounding planet gears
have 25 teeth.
Figure 7.3: Rover gearbox
All gears within the gearbox are 20 pitch hardened steel gears, allowing for a resilient drive
system. 20 pitch gears were chosen because due to design, it is required that the gears lock on a
0.5in wide center hex driveshaft, and the only type of gear with a 0.5in wide center hex bore
were 20 pitch gears. A 0.5in hex center bore to allow the gears to lock onto its driveshaft.
Gear strength can be calculated with the Lewis Factor Equation for Gear Tooth Calculations:
𝜎 = 𝑃𝑊𝑡
𝐹𝑌(45)
𝑊𝑡 is tangential tooth load, F is the face width of a tooth, P is diametral pitch, Y is the Lewis
Factor, and is bending stress is a gear tooth. The large 72 tooth gear transferring power from the
motor to the gearbox can handle a load of 52.41 pounds. This level of strength is well above any
amount of load that will be necessary in this mission. The smallest 25 tooth gear can handle a
104.85 pound load. Strength shows that the gear are more than strong enough to transmit power
to the wheels.
138 Long Beach Rocketry | CDR 2017 - 2018
Figure 7.4: Rover exterior body
The power is transmitted through the center hex drive shaft through the center of the triangular
gearbox into the interior of the rover, where is it connected by a 32 pitch 72 teeth gear. 32 pitch
gears were used because the 16 tooth pinion gear on the motor has a 4mm shaft that can only
accommodate a 32 pitch center gear that was also compatible with a hex bore clamping hub with
a 0.770" pattern, which clamps onto the 0.5 in wide center hex driveshaft. Torque will be
transferred to the wheels from the motor. Power goes along a 16 tooth pinion gear to a 32 tooth
center gear (both of which are 32 pitch gears), then to a 35 tooth center to 25 tooth planetary
gears (both of which are 20 pitch gears).
A gearbox will also cover the gears and protect them from dust and debris. It will also allow the
set of gears to pivot around a central axis to allow for the bogie system to work.
7.2.4 Bogie System
Based on results seen in the subscale launch, it would be preferable if the rover had additional
traction so LBR has decided to add a bogie system. The bogie system will help the rover contact
the ground acting as suspension. Even with rough terrain, all four out of six wheels will make
contact with the ground compared to having the wheels fixed to the body without suspension.
The bogie system will allow for additional traction and ability to traverse harsh terrain which
would increase likelihood of mission success. While the bogie system has it advantages, it would
add complexity to the rover. With the increased complexity, there is also a higher chance of parts
failing with the bogie system. The bogie system also takes away interior space from the rover
electronics. Though the bogie system will take additional space, the system has been designed
with the other systems in mind so there is sufficient space for all the other rover components.
139 Long Beach Rocketry | CDR 2017 - 2018
Figure 7.5: Rover bogie system
7.2.4.1 Individual Components
Titanium Rod: This rod will be used to connect to a pivot at its midpoint and it will be
connecting the two gear boxes in both ends of the rover. The length is 6in and the diameter is
3/16in. Titanium is the material the team chose to use due its similar strength to steel while being
lighter than steel. The density of steel is 7.8g/𝑐𝑚3[7.1]; titanium is 56% that of steel [7.2]
therefore titanium is considered a better combination of low weight and high strength ratio when
comparing to steel. With titanium, the team can use less material to yield the same amount of
strength therefore saving weight.
Figure 7.6: Bogie system rod
Ball Joint Link: The ball joint link will be the connecting part for the attachment to the
gearboxes as well as the titanium main rod attached to the pivot point. Since it is a ball joint
connection, when the gearbox moves the ball joint would allow some movement on the ball joint
link. The ball joint link allows up to maximum 45 degrees of ball swivel. Since the ball joint link
is a smaller part of the bogie system as a whole, the team has decided to have the link made out
of plastic as it is more weight and cost efficient.
140 Long Beach Rocketry | CDR 2017 - 2018
Figure 7.7: Bogie system ball joint
3d Printed Parts: Along with the ball joint connections and the titanium rod, the team designed
custom parts for the bogie system. Some of the parts include the pivot point used to attach the
titanium rod to the triangular shell and also the connecting part for the gearbox and the bogie rod.
The team has made the decision to use the 3D printer material ABS as it is a strong, sturdy
material. ABS is stronger compared to PLA, which is more brittle with a lower melting point.
141 Long Beach Rocketry | CDR 2017 - 2018
Figure 7.8: Integrated bogie system and components
7.2.5 Wheel and Tire Design
One of the defining features of any rover is the design and material choice of the wheels. The
rover must have tires that can grip the terrain very well and be light. In order to grip the terrain
well it must have deep treads and be soft. 3D printed wheels and RC rock crawler tires have been
chosen as an optimal setup for the triangular rover because it provides stability while providing
ground clearance and grip over rough terrain. 3D printed wheels are also more cost and weight
efficient as compared to aluminum beadlock wheels.
Based off the results of the subscale launch, LBR has decided to use expandable wheels in lieu of
the non-expandable alternative. In addition to be more volume efficient within the RDM, these
wheels are better suited to traverse the required terrain. After the payload was deployed during
the subscale testing, the rover subsystem team found that non-compressible wheels were
ineffective in propelling the craft forward. The wheels would simply dig the rover into the
ground over soft, sandy terrain. Compressible tires will be able to mold to the shape of the
ground and have stronger grip. Also, utilizing an off road wheel and tire setup, the wheels will be
thinner than the tire to prevent the tire from de-beading off the wheel and prevent the wheel from
getting caught on obstacles. Since buying RC beadlock wheels were costly and heavy, the team
opted to design and manufacture 3D printed beadlock wheels. Each side of the triangular rover
will have 3 wheels and tires and are attached to each vertex of the triangular rover. All the
wheels on the rover will be powered by the gearbox.
142 Long Beach Rocketry | CDR 2017 - 2018
Figure 7.9: Wheel design
The wheels will be 3D printed and have a beadlock design. They will be 1in in diameter and
0.57in wide, which is narrower than the width of the inner rim of the tire to prevent the tire for
de-beading off the wheel. It will also help prevent the wheel from getting caught on obstacles. A
beadlock wheel design will allow for tire reusability and the ability to change to different tire
designs. Also, a beadlock wheel will allow for a more secure attachment between tire and wheel.
The wheel will have a 0.5in wide hex bore to allow it to attach to its respective axle on the
gearbox to roll and carry the weight and transport the rover. A 0.5in wide hex shaft was used to
keep consistency throughout the rover, since the center driveshaft will also be made from the
0.5in wide hex rods.
Figure 7.10: Tire selection
The tire has a maximum width of 0.87in. Its outer diameter is 2.43in and the inner diameter is
1.0in [7.3]. It is a standardized tire for a 1/24th scale micro crawler remote controlled truck. The
tire is most optimal for max ground clearance while being able to fit within the payload bay. The
143 Long Beach Rocketry | CDR 2017 - 2018
tire is made out of a material marketed by RC4WD as an Advanced X2SS Compound, which
will allow for the tire to have a soft and sticky characteristic. It will allow the tire to flex over
rough terrain for the most grip. The tire is modelled after its real life counterpart of a rock
crawler tire for the most effective tread pattern for grip. The tire will be semi-filled with air,
meaning that while air will not be pumped into the tire, it will retain the air it has inside the tire
since it the air inside the tire will be sealed off from the rest of the environment. The semi-
pneumatic tire is filled with atmospheric pressure air to reduce weight since filling the tire with
foam will increase the weight of the entire wheel and tire assembly. The tire weighs 0.42oz.
Figure 7.11: Wheels and tires on rover
When the shaft on the gearbox turns, it will turn the wheel which will in turn the tire, which will
make the rover move forward. Each of the triplets of wheels on each side of the rover will move
in unison.
7.2.6 External Frame
LBR will use ABS plastic for the rover shell. The plastic is readily available to the team and is
easy to work with; however, this material is not as strong as the alternatives of aluminum and
carbon fiber which have a low tensile strength of 42.5 - 44.8 MPa [7.4]. During the subscale
launch, the subscale rover was made with an ABS plastic shell to test whether this material
would be able to survive the launch of the rocket. The results showed that this material is
144 Long Beach Rocketry | CDR 2017 - 2018
sufficient because the rover remained intact and was not deformed after deployment, though
additional test will be performed to confirm its strength. ABS is favorable over carbon fiber and
aluminum due to its cost efficiency, easy accessibility, and simple application and machining.
The entire shell of the payload will be made out of ABS plastic.
Figure 7.12: Rover exterior body
The total length of the shell is 5.8 in with a height of 3.25 in. The general thickness is 0.07 in.
and 5.8 in for the length of the rover shell which allows enough room to fit all the necessary
components into the rover and gives the rover 2 in of tolerance in the RDM.
Along with the shell of the rover, the gearbox covers will also be made with the same ABS
plastic
145 Long Beach Rocketry | CDR 2017 - 2018
Figure 7.13: Rover gearbox system
LBR will use two gear box covers: one for each side of the rover. The total width of the box is 1
in with a height of 3.7 in with a thickness of 0.07 in.
Plastics tend to generate high amounts of friction when they come into contact with one
another. LBR will use metal bearings on the gearbox to ensure that the tires will be able to rotate
smoothly along the gear shaft. The payload shell and gearbox will not come into contact with the
inner shaft of the rocket, so there will be no contact between these two surfaces.
7.2.7 Motor Mount
The motor mount will allow for easy assembly and a means to secure the internal components.
The motors, battery, and electronics will be mounted to the motor mount allowing them to be
installed into the rover as a unit. The battery will sit between the motors and will allow for
clearance for a center rod to go through the middle of the rover above the battery. A shelf will
provide clearance, allowing for electronics to be mounted above the center rod from RDM to go
through the middle of the rover.
Figure 7.14: Rover motor mount
The motor mount which serves as a shelf for the electronics has been constructed efficiently
store the components within the rover, as it cannot interfere with our deployment rod and the
146 Long Beach Rocketry | CDR 2017 - 2018
triangular external frame. The motor mount houses the motor and battery as well as housing the
servo, Arduino Nano, and the motor driver. The motor mount shelf extension has been designed
and placed to not obstruct the deployment rod as well as appropriately holding the electronics.
The motor mount design is also efficient at minimizing the empty space inside the rover, which
helps reach the goal of ultimately maximizing the limited space. It has been constructed to exist
around the rod while holding together the electronics in place. These components are what
makes up the rover internally. The motor mount acts as a structural branch within the rover to
withhold its internal components, as these components are of importance. Once all the
components are secured onto the motor mount assembly they can be inserted into the rover as a
single unit.
Figure 7.15: Rover motor mount
The team has chosen to use ABS 3D printed material as it is light and simple to manufacture.
ABS is also a stronger material than PLA. The rover is held inside the rocket therefore the
internally placed motor mount must be light as it has a contribution to the weight of the rover.
The motor mount will be retaining components with high mass such as the motors which,
therefore the material holding the inputs must be capable of withstanding impact with the
additional inertia of the components. Through testing the team will confirm that ABS will be
capable of withstanding impacts and securely hold all the components in place during the
mission.
The thickness of where the motor attaches to the mount is designed so that the smaller gears
make contact with the bigger gears. The thickness of the two posts are slightly thicker than the
initial design so it can withstand a greater amount of stress. The thickness of the Arduino mount
and all the mounting holes are also taken into consideration so that the Arduino fits directly onto
the motor mount. All of the components attached to the motor mount are placed so that all of the
147 Long Beach Rocketry | CDR 2017 - 2018
parts fit into the rover shell even with the limited space. The thickness of the motor shield
mounting point has also been carefully considered so that the motor shield would be fully
supported under the stress generated by the rocket during launch.
Figure 7.16: Rover bogie system
7.2.8 Solar Panel Mechanism
In order to complete the mission, the rover must deploy a set of foldable solar panels. These solar
cell will be mounted to two of the inside walls of the rover and will be deployed along a hinge to
be exposed to the sunlight. This method of solar cell deployment will protect the cells while the
rover is traversing the terrain and will increase the surface area of the rover to allow for more
light to be captured. LBR will be using two 0.15W 5V solar panels [7.5] that are 2.08 in long and
1.18 in wide which. The solar panels will be functional but will not be wired for the mission due
to additional required components and complexity which would add to potential failure.
Figure 7.17: Solar panel
To open the solar panels a HS7950TH servo motor will be mounted to the motor mount with a
servo horn attached. Two linkage arms will be connected to the servo horn, one 2 inch arm and
one 3 inch arm that attach to a hinge on the solar panel. As the servo turns it will push the solar
panel out, exposing it to sunlight.
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Figure 7.18: Deployment mechanism before and after deployment side view
Figure 7.19: Deployment mechanism before and after deployment
The servo motor LBR is using for the solar panel deployment mechanism shall be mounted on
the motor mount such that it does not interfere with any electronics or other motors as shown in
the cutaway view above. The servo shall turn approximately 90 degrees counterclockwise,
allowing for full extension of the linkage arm without coming into contact with any electronics
or other motors.
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Figure 7.20: Rover side panels which house solar cells
An additional feature of the solar panel deployment mechanism is that it will always cause the
rover to correct itself to an upright position under the circumstance that the rover has landed in a
different position. The servo has up to 486 oz-in of torque [7.6] to cause the rover to right itself
even when it is deployed onto one of its sides ensuring that the solar panels will be opened right
side up. There will be a slight loss in torque at the servo linkage arm due to leverage but it is
negligible due to the high torque of the servo.
Figure 7.21: Rover with side panel assembly
For the material, the primary linkage arm will be “2 inches long by .2 inches wide by .1 inches
thick”. The secondary linkage arm will be 3” inches long by .2 inches wide by .1 inches thick”
[7.7]. These dimensions were determined to have the least effect on the surrounding structure
while still being sturdy enough to support the torque required to flip the rover in the event of the
150 Long Beach Rocketry | CDR 2017 - 2018
rover landing on one of its sides. LBR has elected to purchase these parts from a vendor who
sells the linkage arms of the same length and width but were .1 inches too thick, so LBR shall
shave them down for application on the rover.
7.3 Electronics
7.3.1 Design and Components
The electrical system on board the rover will consist of 2 propulsion devices, a microcontroller to
serve as the primary controller, an inertial measurement unit (IMU) to detect roll and yaw, and
multiple sensors to detect objects which need to be avoided. In addition, the rover will have a
servo that will deploy a solar panel once the rover reaches the necessary distance. The image
below shows the wiring diagram of the rover’s electrical circuit.
Figure 7.22: Rover Circuit Diagram
Arduino Nano
The microprocessor board chosen by LBR is the Arduino Nano because of its compact
dimensions while containing the necessary amount of general purpose input/output pins needed
for the electronics. The Nano contains 22 digital pins and 8 analog pins. The microprocessor can
151 Long Beach Rocketry | CDR 2017 - 2018
be powered by a battery between 7 and 12 volts and can output 5 volts at maximum of around
40mA [7.8]. It has a low power consumption of 19mA which is preferable while the rover is on
idle. This board will be used as a master-writer and master-reader, reading data from the sensors
and sending commands to the motor shield. In addition, LBR has previous experience with
Arduino based microprocessor boards and concluded that knowledge from previous projects can
be applied.
Figure 7.23: Arduino Nano
Full-Bridge Motor Driver Dual - L289N There are clear advantages to having a motor shield versus attaching the microcontroller directly
to the motors. The first is it simplifies directional and magnitude control over each of the motors.
Second, the shield allows for a maximum of 12V and 2A [7.9] output instead of the Arduino
Nano wired directly to the motors, which would result in a maximum of 5V and 40mA which
would be insufficient to power the motors. As shown by figure 6.21, the shield contains 4 pins
for input signals, 2 pins for power input and ground, 4 pins for outputting power to 2 motors, and
1 pin for outputting 5V of power that will be used to power the Arduino Nano.
The 4 input pins are grouped into pairs as 8 and 9 and 10 and 11 (figure 7.24) that intake a digital
input. One of the pins of the pair will act as a power while the other as a ground. Direction of the
motor can be changed by swapping the values of the 2 pins in the pair. In addition pins 7 (paired
with 8 and 9) and 12 (paired with 10 and 11) are designed to intake analogous input and set the
speed of the corresponding motor based on its paired pins. The 2 power input pins (labeled 4 and
5 on figure 7.24) are used to supply power to the driver through a direct connection to a battery
of a maximum voltage of 12 volts. The 4 power output pins (figure 7.24) are labeled (in pairs
respectively) as 1 and 2 with the other pair labeled as 13 and 14. For each of these pairs, one wire
will act as a positive node while the other acts as a ground. In order to reverse the motor, the
roles of a pair will be flipped and the motor will rotate the opposite direction.
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Figure 7.24: L298n Motor Driver
Inertial Measurement Unit (IMU) - MPU6050
LBR has chosen the MPU6050 for onboard heading information. This sensor provides readings
in 6 degrees of freedom from the 3-axis gyroscope and 3-axis accelerometer [7.10]. The utility of
this sensor is to detect when the rover is veering off-course, the IMU will detect and signal the
Arduino to modify the output to the motors to steer it back on course. The second utility of this is
to detect roll so in the event of the rover folding along the z axis, the rover will understand the
current path is inaccessible and search for an alternate route.
Motor
Choosing the correct motor for the craft is crucial. The motor must be able to provide ample
power and torque to allow the vehicle to traverse the terrain, and have a reasonable volume and
weight in order to be implemented into the rover effectively. Since speed in not a requirement of
the rover, LBR will prioritize high torque. The motor will be attached to the motor mount which
the batteries and electronics will also be mounted to; consequently, the motor must have a small
enough volume to fit into the mount with the other components. The motors will be placed on
two of the three ends of the rover body and will have a shaft coupler to connect to the gears to
spin the wheels. Each motor controls a side of the rover and is connected to a gearbox which
rotates three wheels the same direction, using differential steering to change heading. In order to
make the rover move across the terrain, the team is using two high-torque DC motors. The
motors have a gear ratio of 104:1 resulting in 159.7 oz-in of torque and a speed of 116 RPM. The
input voltage range is 3V-12V, and due to the battery selection will be operating at roughly
11.1V. Under no load the motors draw 0.2A each, but the team estimates a current draw of 0.4A
per motor under normal load conditions [7.11].
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Figure 7.25:116 RPM planetary gear motor
Servo
The solar panels mechanism will be deployed using a Hitec HS-7950TH servo. Servos consist of
a DC motor, gear train, potentiometer, and integrated circuit which provide accurate feedback of
motor position and make it simple to control deployment angle. The selected servo is capable of
403 oz./in. of torque at 6 volts which provides adequate power to lift and flip the rover into the
correct position for its solar panels to face the sun. Under no load at 6V the servo draws 300mA
of current, but if stalled it can pull up to 4.8A which would quickly drain the battery [7.6]. Since
the solar panels do not deploy until the end of the mission, the servo will remain on standby
where it will only draw a no-load current for the majority of the mission, and will only drain
slightly more of the battery to deploy the solar panels in the end.
Battery Choice
The battery chosen has sufficient power to drive the rover and enough capacity to power the
rover throughout the mission. The rover must be turned on prior to entering the launch vehicle
and remain powered the full duration of the rocket launching procedure up until the rocket
landing and deploying the rover, then proceed with its mission. Battery selection has been made
to ensure there is enough capacity to last throughout the mission. It also fits within the required
dimensions of the rover in between the motors and under the center rod.
The team found the required battery capacity through the formula of [7.12]:
Battery Life (hours) = 0.7∗Battery Capacity in Milliamps per hour
Load Current in Milliamps (46)
Through subscale testing and experience from previous years, the rocket is left to idle for up to
an hour prior to launch during launch procedures, and needs roughly an additional half an hour
until the rocket lands and the team can locate it to prepare for rover deployment. The team has
considered the idle time of approximately 1.5 hours as well as the rover run time of
approximately 0.2 hours, consuming 0.5 amps and 2 amps respectively. With information on the
run time and current usage the battery capacity can be found using the formula:
Battery Capacity in mAh = (𝐵𝑎𝑡𝑡𝑒𝑟𝑦 𝐿𝑖𝑓𝑒(ℎ𝑟𝑠))∗(𝐿𝑜𝑎𝑑 𝑐𝑢𝑟𝑟𝑒𝑛𝑡 𝑖𝑛 𝑀𝑖𝑙𝑙𝑖𝑎𝑚𝑝𝑠)
0.7(47)
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(1.5 ℎ𝑜𝑢𝑟𝑠) ∗ (500 𝑚𝐴)
0.7+
(0.2 ℎ𝑜𝑢𝑟𝑠) ∗ (2000 𝑚𝐴)
0.7= 1642 𝑚𝐴ℎ
The power source selected for the rover is an 11.1V battery with 1800mAh of capacity to ensure
power throughout the mission.
7.4 Control System
7.4.1 Proportional Integral Derivative (PID) Controller
A PID controller is a control mechanism that calculates the error based on the difference between
the expected input and the measured input. The controller provides an output based on the
proportional, integral, and derivative terms to calculate response to the error signal. The
proportional variable is calculated from the current error, the integral variable is calculated by
the past error, and the derivative variable is calculated from the expected or future error. These 3
variables are then analyzed for overshoot and steady-state error. Overshoot is when the system is
overdamped and the response to an input overcorrects and the angle exceeds the error in the
opposite direction, and causing another overshoot to then correct that response leading to an
unstable system. Steady-state error occurs when the systems settles at a point that is not within
the given range. The integral term corrects this, but at the cost of causing an increasingly
unstable system. The derivative term is used to increased stability and can therefore overcome
the instability induced by the integral term.
7.4.2 Discrete closed-loop response
The system will be designed around a discrete closed loop response. A closed loop system was
picked over an open loop system because data will always be received from the sensors and the
decision of action chosen is determined by the system. Closed loop systems are designed to have
a desired goal to achieve, which is an error of 0. The system will use motors in an attempt to
achieve this goal by alternating power between motors. The responses to achieve this are
determined to be discrete because the rover will have an established goal that it must reach. The
goal will be to move over 5 feet, which the rover will promptly exit the loop and then enter a
second loop to fully deploy the solar panel.
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Figure 7.26: System Response Flowchart
7.5 Subscale Prototype
For the rover subscale prototype, LBR created a scaled down version of the system with limited
functionality to test out the general design concept. This subscale model allowed the team to
pinpoint any design flaws and get a general understanding of system weaknesses and where to
improve on. One result of building the subscale rover is that the team learned that designing a
rover that was easy to assemble is important to the process. The system was built primarily using
3D printed material to test the rigidity during launch and rocket impact. The frame, wheels, and
motor mounts were all made using ABS 3D printer filament with the hexagonal nuts for the
RDM threaded rod press fit and epoxied into place. The rover was driven by two continuous
rotation servos which were connected to the Arduino Nano microcontroller board, powered by a
compact 7.4V LiPo battery.
The team was able to test the prototype rover during the subscale test of the launch vehicle to see
how the design can perform during a mission. Once the launch vehicle landed and he team found
the rocket, the rover deployment mechanism was activated remotely and the rover was released.
As anticipated, the deployment mechanism was able to function even with a mound of dirt in
front of the nose cone but the team found a flaw with the RDM since the rover was not able to
fully disengage from the threaded rod, a problem which will be fixed by using a longer rod. With
some assistance the rover was able to make contact with the surface with four wheel and the
team gave it power to attempt to traverse the terrain. Due to it being a subscale model the team
had to make the rover one-third smaller than the final design and only two wheel drive so it
lacked ground clearance and proper traction so the rover was not able to make it very far on the
loose sandy surface of the desert. Through this test the team learned the importance of ground
clearance, tire traction, and the necessity to create a gearbox so all wheels can receive power.
The team was also able to conclude that the 3D printed ABS material was able to withstand the
rocket launch and landing and that the symmetrical design is practical.
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Figure 7.27: Rover subscale model
7.6 Mass Budget
Table 7.1: Mass of DORITO
Estimated Mass of Components (oz)
Motors (2) 6.46
Arduino Nano Microcontroller 0.25
L298N Motor Driver 0.92
MPU6050 IMU 0.54
11.1 V LiPo Battery 4
GY-530 Rangefinder 0.16
Wiring 0.5
Dorito body 3.83
4x Gearbox covers 0.201
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18x 1/2" Hex Bore Bearings 14.4
Scrambler Off-road 1.0" Scale Tires 2.52
Wheels 0.949
6x 25t planetary gears 10.214
2x 35t big center gear 9.22
2x pinion gears 0.5
2x interior gears (aluminum hub gears) 0.8
2x ½” wide center hex shafts 2.56
Hardware 1.7
Servo (Hitec HS-7950TH) 2.29
4x Solar panels 4
Bogie 0.05
2x Solar Panel Deployment Linkage Arms 0.64
Micro Servo Horn 0.32
Total 67.024 oz
7.7 Team Derived Requirements – DORITO
In addition to the NASA requirements listed in the previous section, LBR has derived additional
requirements for the current rover design. Through testing and analysis, the team will fulfill all
the requirements in order to achieve a fully operational and reliable system.
Table 7.2: Team Experiment Requirements
Requirement Verification
Method
Verification Plan Status
Rover must be
capable of translating
5 feet after deploying
in any orientation.
Analysis/Test Design rover symmetrically
to always land on its wheels
regardless of launch vehicle
landing orientations.
Will be verified by
February 2018 after
complete
construction of
mechanical
hardware and
electronics for
rover.
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Rover must be
capable of traversing
through rough terrain,
such as rocky
surfaces, dirt mounds,
and tall grass fields.
Analysis/Test Rover designed with
reasonable ground clearance
and obstacle avoidance
control.
Will be verified by
February 2018 after
completion of the
rover.
Rover must be
capable of handling
various impact speeds
and still perform.
Test Rover will be subjected to
indirect impact of up to 30
ft/s to simulate rocket
landing and remain operable
Will be verified by
February 2018 after
completion of the
rover.
During launch the
rover will be subject
to vibrations which it
must withstand.
Test The team will create a
testing apparatus that will
vibrate the rover for an hour
of which after the rover
must remain operable
Subscale launch
verified. Full-scale
will be verified by
the end of February
2018.
Rover must traverse a
variety of
environmental
conditions.
Test Rover will be tested in a
variety of ground conditions
including large obstacles to
test object avoidance
system.
Will be verified by
February 2018 after
completion of the
rover.
Rover must be able to
transverse up an
incline of up to 45°.
Test Rover will be tested on
rough terrain with 45° slope.
If this angle is exceeded it
must use obstacle avoidance
to find an alternate route.
Will be verified by
February 2018 after
completion of the
rover.
Towards the end of
deployment, rover
must not bend or be
constrained on end of
threaded rod.
Redesign/Test Rover must completely
disengage from RDM and
traverse away from the
launch vehicle.
Subscale test failed.
Full-scale will be
verified by the end
of February 2018.
Solar panel must be
capable of flipping
rover upright after
rover deployment.
Test Solar panels open one
direction, so if the side is
facing the ground the servo
must be capable of using the
flaps to flip the rover
upright.
Will be verified by
February 2018
upon completion of
the solar panel
deployment code.
After deployment,
rover must traverse
the terrain and
maintains a desired
Analysis/Test Through a control system
and IMU readings, the rover
must return to a heading
even when encountering
Will be verified by
the end of February
2018 after
completion of the
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heading. disturbances. rover.
Electronics must be
properly wired for
full circuit
functionality.
Inspection/Test All connections will be
soldered securely and tested
through the assembly
process.
Will be verified by
February 2018
upon completion of
the electronics.
Throughout the
launch, the 11.1V
1800mAh battery
must have sufficient
power to complete the
mission.
Analysis/Test Calculations performed to
find minimum required
battery capacity. A full
mission test must be
performed to pass.
Will be verified by
February 2018
upon completion of
the electronics.
L298n motor driver
must provide
additional 5V supply
to microcontroller
board and sensors.
Analysis/Test Built in 5V step-down
voltage regulator on motor
driver must be capable of
providing power to
additional rover electronics.
Will be verified by
February 2018
upon completion of
the rover
electronics.
118 RPM motor must
be controllable in
both speed and
direction of rotation
through H-bridge.
Test The L298n motor driver
shall provide control and
sufficient power to drive the
geared DC motors and move
the rover.
Will be verified by
February 2018
upon completion of
the rover
electronics.
Hitec HS7950TH
must receive power
and control signals to
deploy solar panel.
Analysis/Test The servo shall receive 5V
through the motor driver
and a control signal from the
Arduino. The servo will be
powered throughout the
mission but deploy when
prompted by the
microcontroller signal.
Will be verified by
February 2018
upon completion of
the electronics.
Rangefinder sensor
must be able to
identify obstacles
higher than 2” and
maneuver according.
Analysis/Test If the rover does not have
sufficient ground clearance
the VL6180 rangefinder
sensor will be used to detect
oncoming obstacles. The
control system must be
capable of reading the input
and mitigate the obstacle.
Will be verified by
February 2018
upon completion of
the electronics.
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Section 8: Testing Plan
8.1 Approach to Testing
LBR will conduct tests on all mission critical components that will be on board the full-scale
launch vehicle to ensure that they will function properly during launch. Successful completion of
the following tests will allow LBR to improve upon each subsystem’s design in order to ensure
that all subsystems are functioning properly.
8.2 Launch Vehicle Testing Plan
8.2.1 Launch Vehicle Major Test
The launch vehicle test consists of several sub tests which will prove the integrity of the design.
These tests will verify the requirements that pertain to the performance of the full-scale launch
vehicle.
Subscale Vehicle Ground Ejection Test
Test Objective: This test will demonstrate the system’s ability to separate sections of the launch
vehicle to allow the recovery equipment to deploy during flight.
Success Criteria: Payload bay successfully separates, forcing drogue parachute and shock cord
to be removed from the launch vehicle. The propulsion bay successfully separates, forcing the
main parachute and its shock cord to be removed from the launch vehicle.
Procedure:
1) Put on safety glasses and operate in a space that is away from any individual that is not
participating in the separation test.
2) Have all other subsystems fully assembled into the launch vehicle for the most optimal
and realistic simulation.
3) Fold the main and drogue parachute and set aside for later assembly.
4) Acquire electronic matches to connect the wire connectors to the battery (DO NOT
CONNECT ANYTHING TO THE BATTERY)
5) Connect the electronic matches to the wire connectors (with no external connections on
the other side) and place the electronic matches into their respective ejection charge
containers.
6) The wire connectors should allow the performer to still connect a battery to it while at a
distance away from the launch vehicle. The performer should never be in close proximity
of the launch vehicle while ejection charges are loaded.
7) Measure out black powder on a scale into a small container for the main and drogue
parachute ejection charges. Set aside that container for later in an area away from any
individuals.
8) Acquire cellulose insulation and masking tape.
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9) Pour the black powder container into the main and drogue ejection charge holders and fill
the charge holders with the cellulose insulation. Tape the insulation to prevent it from
falling.
10) Connect the parachutes and parachute shock cords together onto the Avionics Bay and
assemble to launch vehicle completely.
11) Carry the launch vehicle to an area that is safe for detonation. Remember, the wire
connectors must be available for connection to the performer while the performer is
safely away from the launch vehicle.
12) Perform a countdown to assure all participants of the event and once ready, connect a
battery to the wire connectors.
13) Once the first detonation occurs (main or drogue), safely transition the launch vehicle in
an orientation that is ready for the next separation test for either the main or drogue
depending on which section separated first.
Results: Pass
The drogue and main deployment ejection charge successfully separated the launch vehicle into
3 sections.
Subscale Vehicle Flight Test, December 2nd, 2017
Test Objective: This test will demonstrate the flight characteristics, recovery, and structural
integrity of the launch vehicle.
Success Criteria: The subscale launch vehicle is considered success if its meet all NASA
requirements. The recovery of the subscale vehicle must successfully deploy and land safely.
Procedure:
1) Manufacture the subscale launch vehicle by the day of subscale test day
2) Program the two altimeters to ignite an ejection charge at apogee and at 700 feet
3) Fold the main and drogue parachute and set aside for later assembly
4) Acquire electronic matches to connect the wire connectors to the battery (DO NOT
CONNECT ANYTHING TO THE BATTERY)
5) The wire connectors should allow the performer to still connect a battery to it while at a
distance away from the launch vehicle. The performer should never be in close proximity
of the launch vehicle while ejection charges are loaded.
6) Measure out black powder on a scale into a small container for the main and drogue
parachute ejection charges. Set aside that container for later in an area away from any
individuals.
7) Acquire cellulose insulation and masking tape.
8) Pour the black powder container into the main and drogue ejection charge holders and fill
the charge holders with the cellulose insulation. Tape the insulation to prevent it from
falling.
9) Connect the parachutes and parachute shock cords together onto the Avionics Bay and
assemble to launch vehicle completely.
10) Insert the motor into the motor tube and secure using the motor retainer
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11) Check for the center of gravity to make sure the subscale launch vehicle has the same
static stability as the one in the simulation
12) Attach the subscale launch vehicle to the launch pad and prepare to launch
13) Make sure that all spectators are at least 100 feet away from the launch pad
14) Arm each altimeter by turning each rotary switch to the on position, which are accessible
from the outside of the airframe
15) Only Mentor/or certified personnel is allowed to insert the igniter into the motor,
ensuring that the igniter tip is inserted far enough into the motor
16) Check continuity between the igniter and the launch system
17) Launch the subscale vehicle
Results: Pass
The subscale launch and recovery system met the success criteria of the test. The details of the
flight can ben seen in Section 5.5
Full Scale Launch Vehicle Ground Ejection Test
Test Objective: This test will validate that black powder charges are sufficient to break the shear
pins and separate all 3 sections of the launch vehicle.
Success Criteria: Payload bay, AV bay, and propulsion bay are successfully separate into 3
sections, forcing the drogue and the main parachute and shock cord to be removed from launch
vehicle.
Procedure:
1) Put on safety glasses and operate in a space that is away from any individual that is not
participating in the separation test.
2) Have all other subsystems fully assembled into the launch vehicle for the most optimal
and realistic simulation.
3) Fold the main and drogue parachute and set aside for later assembly.
4) Acquire electronic matches to connect the wire connectors to the battery (DO NOT
CONNECT ANYTHING TO THE BATTERY)
5) Connect the electronic matches to the wire connectors (with no external connections on
the other side) and place the electronic matches into their respective ejection charge
containers.
6) The wire connectors should allow the performer to still connect a battery to it while at a
distance away from the launch vehicle. The performer should never be in close proximity
of the launch vehicle while ejection charges are loaded.
7) Measure out black powder on a scale into a small container for the main and drogue
parachute ejection charges. Set aside that container for later in an area away from any
individuals.
8) Acquire cellulose insulation and masking tape.
9) Pour the black powder container into the main and drogue ejection charge holders and fill
the charge holders with the cellulose insulation. Tape the insulation to prevent it from
falling.
163 Long Beach Rocketry | CDR 2017 - 2018
10) Connect the parachutes and parachute shock cords together onto the Avionics Bay and
assemble to launch vehicle completely.
11) Carry the launch vehicle to an area that is safe for detonation. Remember, the wire
connectors must be available for connection to the performer while the performer is
safely away from the launch vehicle.
12) Perform a countdown to assure all participants of the event and once ready, connect a
battery to the wire connectors.
13) Once the first detonation occurs (main or drogue), safely transition the launch vehicle in
an orientation that is ready for the next separation test for either the main or drogue
depending on which section separated first.
Results: To be determine in February 2018
Full Scale Vehicle Flight Test, February 24th, 2018
Test Objective: This test will validate the flight characteristics, recovery and structural integrity
of the full-scale launch vehicle. It also demonstrates that the shear pins keep the launch vehicle
together during flight, verify that the black powder charges are sufficient, and the payload is
fully protected.
Success Criteria: All launch vehicle sections separate, and the drogue and the main parachute
deploy at the correct altitude.
Procedure: See subscale launch procedure and Section 4.13
Results: To be determine in February 2018
8.2.2 Launch Vehicle Minor Test
8.2.2.1 AV Bay Switch Test
Test Objective: Test that the altimeter can turn on/off from the outside of the launch vehicle
Success Criteria: The altimeter is successfully turn on/off from the outside
Procedure: Install the rotary switch to the AV bay. Make sure to connect the wire from the
rotary switch to the altimeter.
Results: Pass. The team was successfully able to arm the altimeter during the subscale launch
8.2.2.2 Shock Cord Strength Test
Test Objective: Test the strength of the shock cord
Success Criteria: The shock cord must withstand at least 1000 lb and can support the weight of
the payload bay and the propulsion bay
Procedure: Attach the shock cord into the material load frame and let the machine pull stretch it
until it breaks
Results: Pass. The shock cord was snapped at 1400 lb
8.2.2.3 Launch Vehicle Structural Integrity Test
164 Long Beach Rocketry | CDR 2017 - 2018
Test Objective: Test the airframe to prove that it can fully protects the internal components of
the launch vehicle by performing the drop test
Success Criteria: All components are not damaged after drop test
Procedure: Assemble the launch vehicle and drop it 5 feet above the ground into a similar
terrain at the competition launch site
Results: Pass. None of the internal components was damaged and the dummy rover was able to
successfully deploy
8.2.2.4 Vibration Test on the Payload bay and the AV bay
Test Objective: Ensure that all the electrical components are remain connected and secured
during the flight vibration
Success Criteria: All electronics component and wire remain connected
Procedure: Fully assemble the AV bay and tighten it using the hex nuts. Put the AV bay into the
vibration test stance. The assembled bay then is vibrated up to the level which is expected during
the launch.
Results: Pass. All components were undamaged, and all electrical connections were able to
function normally
8.3 RDM Testing Plan
Table 8.x: RDM full scale testing comparison chart
Testing Plan
Test Objective
Success Criteria
Communication
Range Test
Ensure that the system
can operate from at least
a mile away.
The RDM system can respond to signal,
without error or delay, from a distance of at
least a mile.
Full Rover
Deployment Test
Have the rover deploy
from the airframe.
Ensure that the rover can fully disconnect
from the rods and rover-pushing plate.
Rover
Deployment
Drop Test
Ensure that the system
can operate after impact.
After dropping the RDM system, the nose
cone remains attached to the airframe and can
successfully activate to deploy the rover.
Rotary Encoder
Test
Determine the real-time
rotation speed of the
motor and the change in
rotation speed.
After benchmarking the real-time speed of the
motor in normal operation, detection of
change in speed especially during loading such
as shear pin breaking, the motor’s speed must
also change as programmed accordingly.
165 Long Beach Rocketry | CDR 2017 - 2018
Radio
Transmitter-
Receiver Channel
Value Test
Ensure that the correct
output values for the
designated channels are
determined correctly for
programming the
control.
When the 2.4 GHz Radio Transmitter-
Receiver is turned, the designated channels’
values can be identified when the switches are
moved from lower extreme to higher extreme.
These extreme values will be used to program
the controls.
Maximum Load
Test
Test motor to determine
if there is enough torque
to detach the nose cone
from the airframe.
Using the nose cone, LBR, will repeatedly run
the motor to ensure that the nose cone and
airframe will separate and allow enough room
for the rover to separate from the internal
compartment.
Shear Pin
Breakage Test
To test the current
drawn by breaking 3 and
4 shear pins.
The current drawn from breaking three or four
shear pins, preferably four, is less than the
maximum current of 2 A which the H-Bridge
can provide.
8.3.1 RDM Test
8.3.1.1 Communication Range Test:
Testing Variable: Strength of the 2.4 GHz Radio Transmitter.
Validation: This test is necessary to ensure that after landing, the RDM system can successfully
activate and deploy the rover. A reliable connection between the transmitter and the receiver is
essential, without this, no signal would be received, which would prohibit the beginning of the
rover deployment. If the transmitter fails, LBR needs to consider utilizing a new set of
transmitter-receiver from another manufacturer. This may call for a minor redesign of the
electronics bay due to the size difference between the new and old receivers.
Procedure: To test this, two members of the RDM team will stand a mile with the 2.4 GHz
Radio Transmitter, while two other members will observe the RDM system for errors or delays.
8.3.1.2 Full Rover Deployment Test:
Testing Variable: Rover deployment from the airframe after launch.
Validation: An issue that occurred during the subscale launch was that the nuts and bolts on the
nose cone bulkhead protruded out interfering with the hex nut in the rover, which resulted in an
unsuccessful deployment of the rover. To prevent the same issue for full scale, the LBR team
redesigned the deployment process. Testing and observing for deployment issues will allow the
team to make necessary design changes before the competition. Failure of this deployment would
cause the RDM team to either redesign the threaded rover-pushing-plate or change the rod
lengths.
Procedure: To test the full rover deployment, the rover will sit on top of a threaded movable
rover-pushing-plate attached to a threaded rod in the center of the airframe, while a 2.4 GHz
Radio Transmitter will be activated to move the rover-pushing-plate along the threaded rod. The
166 Long Beach Rocketry | CDR 2017 - 2018
team will observe the un-threaded rover, be pushed along the rod until it reaches the end and
deploys.
8.3.1.3 Rover Deployment Drop Test:
Testing Variable: Functionality of electronics after impact.
Validation: Testing the electronics functionality after impact allows RDM team to determine
whether it would survive the rocket’s landing and be able to successfully deploy the rover.
Failure of this test would require RDM to either redesign the electronics bay to better secure the
components or consider applying reinforcements, such as encasing the electronics bay for
security. Severe damage to the external nose cone or airframe would require LBR members to
reconsider the external material used or decrease the load to reduce the amount of force the
airframe and nose cone feel during impact.
Procedure: The RDM team plans to assemble the electronics into the airframe, secure the nose
cone using shear pins, then drop it multiple times, at least 3, from a height of 3 to 5 ft above a
concrete surface. After the drop, the team will activate the transmitter to confirm that the
electronics were undamaged during the drop. After activating the transmitter, the members will
observe the nose cone and airframe for significant damages.
8.3.1.4 Rotary Encoder Test:
Testing Variable: Speed Variability of the 118 RPM HD Premium Planetary Gear Motor.
Validation: With the use of the motor’s built-in rotary encoder, this test should verify the rated
RPM (118) of this motor given the recommended voltage of 6V-12V without any loading on the
motor shaft. Varying motor speeds according to different loading forces prevents motor burnout.
When the motor is subjected to a loading, i.e. shear pin breakage, the change in speed will be
noted so that the RDM can be programmed accordingly.
Procedure: RDM team will use pulse-width-modulation (PWM) to control the speed of the
motor, and the rotary encoder to read the real-time speed, both via Arduino Nano
microcontroller. The Arduino Nano will be connected to the serial port of the PC to read the
output of the encoder. For the first test, the motor will not have any load on it and will be turned
on (no preferred direction) with 100% PWM. The team will note the resulting average value. The
second test will have loading in the motor, preferably breaking the three to four 4-40 shear pins.
The change in speed at the point of breaking the shear pins, peak load, will be noted and will be
use as the trigger in the program to slow down the speed of the motor during peak load.
8.3.1.5 Radio Transmitter-Receiver Channel Value Test:
Testing Variable: Value Integrity of the 2.4 GHz Radio Transmitter-Receiver.
Validation: The importance of this test is to ensure that the correct output streams acquired from
the receiver will be displayed for each transmitter channel input. If the transmitter-receiver pair
presents a huge fluctuation or error in the output values, the transmitter-receiver pair will not be
suitable for deploying the RDM due to the chances of inoperability.
Procedure: The test must be performed around buildings and areas prone to radio frequency
interference. To do the test, RDM team will read input received from the transmitter to the
receiver through the serial console monitor of either Arduino IDE or Microsoft Visual Studios.
167 Long Beach Rocketry | CDR 2017 - 2018
Each channel switch will be moved from lower extreme to higher extreme. The theoretical
extreme values range from 1000 uS to 2000 uS. This test will verify the actual output values
which should fall within the aforementioned theoretical values and will be used to construct the
validation statements when programming the RDM control.
8.3.1.6 Maximum Load Test:
Testing Variable: The Torque of the 118 RPM HD Planetary Gear Motor.
Validation: Before testing the shear pin breakage, RDM needs to test to ensure that the motor’s
torque is strong enough to separate the nose cone from the airframe. Failure of the motor to
separate the nose cone and airframe without screwing in 4-40 shear pins requires RDM to
reconsider the motor used. Since it is essential for the nose cone and airframe to separate for
rover deployment, the RDM team will have to consider using either experiment with a
completely different motor or use one with a lower RPM with a higher torque.
Procedure: To perform this test, RDM will need to connect the nose cone to the airframe
without shear pins. In the airframe, the 118 RPM HD Planetary Gear Motor will be coupled with
a 14 in threaded rod connected to a bulkhead in the nose cone, with a trapped hex nut. Activating
the 2.4 GHz Radio Transmitter, the RDM team will observe the nose cone to see if the motor is
strong enough to detach it from the airframe.
8.3.1.7 Shear Pin Breakage Test:
Testing Variable: Maximum current drawn to break three or four shear pins.
Validation: Testing the maximum current to break the shear pins is necessary to ensure that the
current required to break the shear pins does not exceed the maximum current that the H-Bridge
can handle. This is extremely important because exceeding the H-Bridge’s maximum current
would cause the whole system to stall and eventually fail. Failure of this would require RDM to
consider getting a new H-Bridge that can withstand a greater amount of current or if the issue is
that four shear pins draw too much current, but three shear pins fall require less than 2A, RDM
will choose to use three shear pins to secure the rocket.
Procedure: Testing shear pin breakage requires the assembly of the electronics by connecting
the 118 RPM HD Premium Planetary Gear Motor to the two 11.1V Li-Po batteries. To ensure
that the motor has enough torque to shear the three 4-40 shear pins, attach the nose cone to the
airframe. Connect the motor connection in series to an Ammeter and observe the current drawn.
Repeat using 4 shear pins. After the test, compare the collected data to 2 A to determine the best
choice.
8.3.2 Completed Test:
Shear Pin Test: To determine whether breaking three shear pins requires more current than the
maximum current the H-Bridge could take, LBR assembled the electronics portion of RDM and
connected an ammeter in series to the motor. Throughout the entire test, the RDM team
monitored the current and determined that the minimum current drawn was 0.15 A while the
maximum current was 1.65 Amps. Since the maximum current of the H-Bridge was 2 A, this test
was considered a success. A graph of the current drawn during this test can be seen in Figure
6.12. However, since this was tested using a 32 RPM Planetary Motor and the full scale RDM
168 Long Beach Rocketry | CDR 2017 - 2018
system will use a 118 RPM HD Planetary Gear Motor, the results demonstrate how dangerously
close the drawn current was to the maximum H-Bridge current. Exceeding this would end up
frying the H-Bridge resulting in a failure for the RDM system. RDM took note that a new H-
Bridge may be needed since the new motor will draw more current. However, future analysis and
testing are needed to decide whether a new H-Bridge with a larger maximum current is needed.
8.4 DORITO Testing Plan
8.4.1 DORITO Test
8.4.1.1 Impact Testing:
Testing Variable: Rover durability and functionality after impact.
Validation: The rover must be capable of handling an indirect impact of 20 ft/s with a maximum
speed of up to 30 ft/s. To be a success, after an impact, the rover must be able to have a fully
functional electrical system and have all the sensors and motors remain unharmed. In the event
of the failure of these systems upon impact, the rover would be inoperable an unable to complete
its mission. In the event that the LBR rover design proves to be unable to withstand the impact,
LBR must design further subsystems inside the rover that will mitigate the amount of force on
the systems of the rover, but with the side effect of increasing weight.
Procedure: The rover will be dropped onto dirt from various heights housed within a simulated
rocket airframe to simulate rough impact of the rocket and tested to ensure it can still function.
8.4.1.2 Vibration Testing:
Testing Variable: Functionality of mechanical components and electronics during rocket
launch.
Validation: During flight, the rover will be subjected to vibration. The vibration can interfere in
such ways as causing a breakage between soldered wires that results in the rover being
inoperable. After performing this test, LBR will search for any loose soldering points or
connection and will amend these to prevent future problems from occurring.
Procedure: To ensure the rover will be able to withstand flight, the rover will be put it in a
testing apparatus that will shake it for an hour continuously. Then the rover will be tested if it is
still operable.
8.4.1.3 Environmental Conditions:
Testing Variable: Capability of vehicle traversing over diverse terrain.
Validation: The rover will be subjected to various environmental conditions that could prove too
difficult to traverse. Examples of this may be terrain such as sand or a steep incline that would
cause the rover to flip. If the rover detects itself flipping due to steep terrain it must use its
obstacle avoidance system, otherwise the rover would continue along the same route and
repeatedly fail. The result would show in the wheels need to have a different design or the code
would need changing to overcome a particular environmental condition.
Procedure: For testing, LBR will let the rover run in various environments such as the sand on a
beach or a desert and test as to whether it can maintain a path and avoid becoming stuck. In
169 Long Beach Rocketry | CDR 2017 - 2018
addition, the rover will be placed in an environment with steep inclines and when the rover flips,
it must see that route as non-traversable and search for an alternative.
170 Long Beach Rocketry | CDR 2017 - 2018
Section 9: Project Plan
9.1 Requirement Verification
Table 9.1: General Requirements
Minimum Success
Requirement Verification Method Verification Plan Status
1.1. Students on the
team will do 100%
of the project,
including design,
construction,
written reports,
presentations, and
flight preparation
with the exception
of assembling the
motors and
handling black
powder or any
variant of ejection
charges, or
preparing and
installing electric
matches (to be done
by the team’s
mentor).
Verify all project design,
construction, written reports,
presentations, and flight
preparation are performed
100% by students. Verify that
assembling the motors and
handling black powder or any
variant of ejection charges, or
preparing and installing
electric matches is performed
by team mentor.
LBR is a 100%
student-run
organization. Team
leads will ensure
all written and
physical work is
performed by
student members
of LBR. Safety
Officer will ensure
assembling the
motors and
handling black
powder or any
variant of ejection
charges, or
preparing and
installing electric
matches is
performed by team
mentor.
In Progress
1.2. The team will
provide and
maintain a project
plan to include, but
not limited to the
following items:
project milestones,
budget and
community support,
checklists,
personnel assigned,
educational
engagement events,
and risks and
mitigations.
Verify the team will maintain
a project plan that will include
project milestones, budget
community support, assigned
personnel, educational
engagement and risk
mitigation.
LBR utilizes and
accessible team
calendar to track
assignments due
dates and event
dates. LBR has
also created an
online storage
filled with budget,
personnel assigned,
and education
engagement event.
The safety officer
ensures daily that
every LBR follows
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171 Long Beach Rocketry | CDR 2017 - 2018
the posted lab
safety rules.
1.3. Foreign
National (FN) team
members must be
identified by the
Preliminary Design
Review (PDR) and
may or may not
have access to
certain activities
during launch week
due to security
restrictions. In
addition, FN’s may
be separated from
their team during
these activities
Verify that all foreign
nationals are identified before
PDR by provided having all
team members provide proof
of US citizenship.
Team leads will
verify every team
member's
citizenship and
submit the
necessary
documents to the
NASA
representative.
Completed
1.4. The team must
identify all team
members attending
launch week
activities by the
Critical Design
Review (CDR).
Team members will
include: 1.4.1.
Students actively
engaged in the
project throughout
the entire year.
1.4.2. One mentor
(see requirement
1.14). 1.4.3. No
more than two adult
educators.
Students must verify to the
team advisor that they will be
able to attend launch before
CDR is submitted. One
mentor and adult educator
must also commit to the
launch week.
LBR will require
that every team
member that can
attend launch week
sign up before
CDR is submitted.
LBR will keep a
record of every
student who will be
attending launch.
Completed
1.5. The team will
engage a minimum
of 200 participants
in educational,
hands-on science,
technology,
engineering, and
Verify team will engage a
minimum of 200 participants
in educational STEM
activities. An accurate activity
report will be completed and
submitted two weeks after the
event. Verify for an
Team members are
required to attend
at least one
outreach event to
maintain team
engagement.
Outreach events
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172 Long Beach Rocketry | CDR 2017 - 2018
mathematics
(STEM) activities,
as defined in the
Educational
Engagement
Activity Report, by
FRR. An
educational
engagement activity
report will be
completed and
submitted within
two weeks after
completion of an
event. A sample of
the educational
engagement activity
report can be found
on page 31 of the
handbook. To
satisfy this
requirement, all
events must occur
between project
acceptance and the
FRR due date.
educational event to count it
will be completed between
project acceptance and FRR
due date.
are planned out
week by week by
the outreach chair
who will be in
charge of planning
each event and
submitting each
educational
engagement report.
1.6. The team will
develop and host a
Website for project
documentation.
LBR will host and update
their team website
http://longbeachrocketry.com/
LBR has specified
a specific
webmaster that
will handle all
social media and
website posts.
Team leads will
check for their
quality and inform
the webmaster
what information
they want to be
posted
Completed
1.7. Teams will
post, and make
available for
download, the
required
Verify LBR will make all
necessary documentation
available for download on the
team website by the due dates
of the project.
LBR will prepare
and finish all
documentation 2
weeks before the
due date to ensure
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173 Long Beach Rocketry | CDR 2017 - 2018
deliverables to the
team web site by
the due dates
specified in the
project timeline.
everything is
posted before their
due dates. All links
will be tested to
ensure that there
will be no errors on
the due date.
1.8. All
deliverables must
be in PDF format.
Verify that all deliverables are
in PDF format
Team leads will
use PDF reader to
compile all
documents to a
In Progress
1.9. In every report,
teams will provide
a table of contents
including major
sections and their
respective sub-
sections
Verify every report provides a
table of contents including
major sections along with
their respective sub-section.
Team leads will
structure each of
their sections to
follow NASA’s
report guidelines.
They will check
that any subsection
that they add to
their system is
added in the Table
of contents.
In Progress
1.10. In every
report, the team
will include the
page number at the
bottom of the page
Verify in every report the
page numbers are included at
the bottom of the page.
LBR’s PDF
compiler will add
the page numbers
to the bottom of
the documents.
Team members
will check and re
read every
document to ensure
no errors have
occurred.
In Progress
1.11. The team will
provide any
computer
equipment
necessary to
perform a video
teleconference with
the review panel.
LBR has access to the
CSULB college of
engineering rooms which
have access to all of the
computer equipment
necessary to perform a video
teleconference.
LBR will get in
contact with the
college of
engineering in
advance of the
teleconference to
reserve the rooms
In Progress
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This includes, but is
not limited to, a
computer system,
video camera,
speaker telephone,
and a broadband
Internet connection.
Cellular phones can
be used for
speakerphone
capability only as a
last resort.
necessary for the
conference.
1.12. All teams will
be required to use
the launch pads
provided by
Student Launch’s
launch service
provider. No
custom pads will be
permitted on the
launch field.
Launch services
will have 8 ft. 1010
rails, and 8 and 12
ft. 1515 rails
available for use.
LBR launches off standard
launch rails that will provided
at the FAR launch site. To
ensure that standard launch
rails are used.
LBR will confirm
on launch day with
the FAR advisor
that the launch rails
available are of
standard sizes.
Will be verified on
launch day.
1.13. Teams must
implement the
Architectural and
Transportation
Barriers
Compliance Board
Electronic and
Information
Technology (EIT)
Accessibility
Standards (36 CFR
Part 1194)
Verify that Architectural and
Transportation Barriers
Compliance Board Electronic
and Information Technology
are implemented
LBR will
thoroughly read
and acknowledge
to the Architectural
and Transportation
Barriers
Compliance Board
Electronic and
Information
Technology (EIT)
Accessibility
Standards.
Completed
1.14. Each team
must identify a
“mentor.” A mentor
is defined as an
LBR will verify a team
mentor who is in possession
of an NAR certification and in
good standing to handle all of
LBR has identified
David Roy as their
team mentor. He is
currently in
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175 Long Beach Rocketry | CDR 2017 - 2018
adult who is
included as a team
member, who will
be supporting the
team (or multiple
teams) throughout
the project year,
and may or may not
be affiliated with
the school,
institution, or
organization. The
mentor must
maintain a current
certification, and be
in good standing,
through the
National
Association of
Rocketry (NAR) or
Tripoli Rocketry
Association (TRA)
for the motor
impulse of the
launch vehicle and
must have flown
and successfully
recovered (using
electronic, staged
recovery) a
minimum of 2
flights in this or a
higher impulse
class, prior to PDR.
The mentor is
designated as the
individual owner of
the rocket for
liability purposes
and must travel
with the team to
launch week. One
travel stipend will
be provided per
mentor regardless
of the number of
the motors. Who also has a
minimum of 2 flights with
motor class or higher and
capable of traveling with the
team to launch week.
possession of an
NAR certification
and in good
standing. David
Roy is also
available to attend
every launch and
go to Huntsville,
Alabama with the
team.
176 Long Beach Rocketry | CDR 2017 - 2018
teams he or she
supports. The
stipend will only be
provided if the team
passes FRR and the
team and mentor
attends launch week
in April.
Table 9.2: Launch Vehicle Requirements
Minimum Success
Requirement
Verification Method Verification Plan Status
2.1. The vehicle will
deliver the payload to
an apogee altitude of
5,280 feet above
ground level (AGL)
Based on simulations
LBR will be reaching an
apogee of 5,400 feet
AGL
LBR will utilize an
air brake system to
ensure that the
launch vehicle will
reach an apogee as
close to 5,280 feet
as possible.
Will verify on
launch day.
2.2. The vehicle will
carry one
commercially
available, barometric
altimeter for recording
the official altitude
used in determining
the altitude award
winner. Teams will
receive the maximum
number of altitude
points (5,280) if the
official scoring
altimeter reads a value
of exactly 5280 feet
AGL. The team will
lose one point for
every foot above or
below the required
altitude.
Verify that the launch
vehicle will carry one
commercially available,
barometric altimeter to
record the altitude.
LBR will utilize
two Perfectflite
StratoLogger
Altimeter in order
to recorder the
launch vehicle's
altitude The second
recorder will be
used for redundancy
to verify the first
altimeters results.
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177 Long Beach Rocketry | CDR 2017 - 2018
2.3. Each altimeter
will be armed by a
dedicated arming
switch that is
accessible from the
exterior of the rocket
airframe when the
rocket is in the launch
configuration on the
launch pad
Verify that each altimeter
will be armed by a
dedicated arming switch
that is accessible from the
exterior of the rocket.
LBR will only be
able to arm every
altimeter with
arming switches
located outside of
the launch vehicle.
Will verify during
full-scale test
launch preparations.
2.4. Each altimeter
will have a dedicated
power supply.
Verify each altimeter will
have a dedicated power
supply.
LBR will use
separate lipo
batteries to power
each altimeter
separately.
In progress
2.5. Each arming
switch will be capable
of being locked in the
ON position for
launch (i.e. cannot be
disarmed due to flight
forces).
Verify each arming
switch is capable of being
locked in the ON position
for launch.
LBR only purchases
arming switches
from apogee rockets
that are capable of
being locked in the
ON position.
In progress
2.6. The launch
vehicle will be
designed to be
recoverable and
reusable. Reusable is
defined as being able
to launch again on the
same day without
repairs or
modifications.
Verify the launch vehicle
is designed recoverable
and reusable.
LBR has designed
the launch vehicle
to be recoverable
and reusable.
In progress
2.7. The launch
vehicle will have a
maximum of four (4)
independent sections.
An independent
section is defined as a
section that is either
tethered to the main
vehicle or is recovered
separately from the
Verify in the design stage
that the launch vehicle
will have a maximum of
4 independent sections.
LBR has designed
their launch vehicle
to be made of three
sections
independent
sections:
propulsion,
avionics, and
payload bay.
In progress
178 Long Beach Rocketry | CDR 2017 - 2018
main vehicle using its
own parachute
2.8. The launch
vehicle will be limited
to a single stage.
Verify the launch vehicle
will be limited to a single
stage.
LBR is a single
stage Aerotech
motor.
Completed
2.9. The launch
vehicle will be capable
of being prepared for
flight at the launch site
within 3 hours of the
time the Federal
Aviation
Administration flight
waiver opens.
Verify the the launch
vehicle is capable of
being prepared for flight
within 3 hours.
LBR will practice
preparing the
launch vehicle
before official
launch dates in
order to ensure the
launch vehicle is
capable of being
prepared within the
time frame.
In progress
2.10. The launch
vehicle will be capable
of remaining in
launch-ready
configuration at the
pad for a minimum of
1 hour without losing
the functionality of
any critical on-board
components.
All vehicle components
will be capable of
remaining on the launch
pad for 1 hour without
losing functionality,
LBR has
implemented lipo
batteries that have a
battery life of over
an hour so launch
vehicle can
maintain
functionality for
over an hour.
In progress
2.11. The launch
vehicle will be capable
of being launched by a
standard 12-volt direct
current firing system.
The firing system will
be provided by the
NASA-designated
Range Services
Provider
The launch vehicle is
capable of being
launched by a standard
12-volt direct current
firing system.
LBR has configured
the launch vehicle
to be capable of
being fired by a
standard 12-volt
firing system. The
same system will
also be used at
every test launch.
Completed
2.12. The launch
vehicle will require no
external circuitry or
special ground support
equipment to initiate
launch (other than
Verify the launch vehicle
will not require any
external circuitry or
group support system to
initiate launch.
LBR will not make
use of any external
group support
system to initiate
launch
Completed
179 Long Beach Rocketry | CDR 2017 - 2018
what is provided by
Range Services)
2.13. The launch
vehicle will use a
commercially
available solid motor
propulsion system
using ammonium
perchlorate composite
propellant (APCP)
which is approved and
certified by the
National Association
of Rocketry (NAR),
Tripoli Rocketry
Association (TRA),
and/or the Canadian
Association of
Rocketry (CAR).
2.13.1. Final motor
choices must be made
by the Critical Design
Review (CDR).
2.13.2. Any motor
changes after CDR
must be approved by
the NASA Range
Safety Officer (RSO),
and will only be
approved if the change
is for the sole purpose
of increasing the
safety margin
Verify that the launch
vehicle uses a
commercially available
solid motor propulsion
system using ammonium
perchlorate composite
propellant (APCP) that is
approved by the National
Association of Rocketry
(NAR) and the Tripoli
Rocketry
Association(TRA).
Verify all motor changes
will be made before CDR
and any changes will
only be approved by the
NASA representative.
LBR will use an
Aerotech motor that
has been approved
and certified by the
NAR and TRA.
LBR also
acknowledges that
the motor decision
has to be made by
the CDR due date.
But if changes were
made LBR will
reach out to their
NASA
representative.
Completed
2.14. Pressure vessels
on the vehicle will be
approved by the RSO
and will meet the
following criteria:
2.14.1. The minimum
factor of safety (Burst
or Ultimate pressure
versus Max Expected
Operating Pressure)
will be 4:1 with
Verify all pressure
vessels will be approved
by the RSO. Verify each
vessel will have a factor
of safety of 4:1, with a
relief valve the full
pedigree on display.
LRR will not be
using pressure
vessels on the
vehicle. But if that
were to change
LBR will follow all
of the following
criteria.
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180 Long Beach Rocketry | CDR 2017 - 2018
supporting design
documentation
included in all
milestone reviews.
2.14.2. Each pressure
vessel will include a
pressure relief valve
that sees the full
pressure of the valve
that is capable of
withstanding the
maximum pressure
and flow rate of the
tank. 2.14.3. Full
pedigree of the tank
will be described,
including the
application for which
the tank was designed,
and the history of the
tank, including the
number of pressure
cycles put on the tank,
by whom, and when
2.15. The total
impulse provided by a
College and/or
University launch
vehicle will not
exceed 5,120 Newton-
seconds (L-class).
Verify that the total
impulse of the vehicle
will not exceed 5,120
Newton-seconds.
LBR’s launch
vehicle will have a
total impulse of
3946 Newton-
seconds. If any
changes were to be
made, the team
leads will adhere to
this criterion.
Completed
2.16. The launch
vehicle will have a
minimum static
stability margin of 2.0
at the point of rail exit.
Rail exit is defined at
the point where the
forward rail button
loses contact with the
rail.
Verify the launch vehicle
will have a minimum
static stability margin of
2.0.
LBR’s will have a
static stability
margin of 2.6
according to
calculations. Team
leads have been
made aware of this
requirements and
any changes made
to the vehicle must
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181 Long Beach Rocketry | CDR 2017 - 2018
adhere to this
criteria.
2.17. The launch
vehicle will accelerate
to a minimum velocity
of 52 fps at rail exit.
Verify that the launch
vehicle will accelerate to
a minimum velocity of 52
fps at rail exit.
LBR’s launch
vehicle will have a
velocity of 75.9 ft/s
at rail exit based on
the OpenRocket
simulation. Team
leads have been
made aware of this
requirement and
any changes made
to the vehicle must
adhere to this
criteria.
Completed
2.18. All teams will
successfully launch
and recover a subscale
model of their rocket
prior to CDR.
Subscales are not
required to be high
power rockets. 2.18.1.
The subscale model
should resemble and
perform as similarly as
possible to the full-
scale model, however,
the full-scale will not
be used as the subscale
model. 2.18.2. The
subscale model will
carry an altimeter
capable of reporting
the model’s apogee
altitude.
Verify a subscale rocket
will be launched that
should resemble the full
rocket prior to CDR.
LBR launched a
subscale rocket on
November 18, 2017
well before the
CDR due date.
Completed
2.19. All teams will
successfully launch
and recover their full-
scale rocket prior to
FRR in its final flight
configuration. The
rocket flown at FRR
Verify LBR will launch a
full-scale rocket before
FRR in its final flight
configurations
successfully. Verify that
if the payload is not
flown mass simulators
LBR will work
towards having a
successful launch
two weeks prior to
FRR. To plan in the
event that there is a
failure in the full-
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182 Long Beach Rocketry | CDR 2017 - 2018
must be the same
rocket to be flown on
launch day. The
purpose of the full-
scale demonstration
flight is to
demonstrate the
launch vehicle’s
stability, structural
integrity, recovery
systems, and the
team’s ability to
prepare the launch
vehicle for flight. A
successful flight is
defined as a launch in
which all hardware is
functioning properly
(i.e. drogue chute at
apogee, main chute at
a lower altitude,
functioning tracking
devices, etc.). The
following criteria must
be met during the full-
scale demonstration
flight: 2.19.1. The
vehicle and recovery
system will have
functioned as
designed. 2.19.2. The
payload does not have
to be flown during the
full-scale test flight.
The following
requirements still
apply: 7 2.19.2.1. If
the payload is not
flown, mass
simulators will be
used to simulate the
payload mass.
2.19.2.1.1. The mass
simulators will be
located in the same
approximate location
will be used to simulate
mass in the approximate
location the payload will
be kept in.
scale launch. Also,
if the payload is not
ready by the time of
the full scale launch
LBR will use mass
simulators to
simulate the
payload.
183 Long Beach Rocketry | CDR 2017 - 2018
on the rocket as the
missing payload mass.
2.19.3. If the payload
changes the external
surfaces of the rocket
(such as with camera
housings or external
probes) or manages
the total energy of the
vehicle, those systems
will be active during
the full-scale
demonstration flight.
Verify if the payload
changes the external
surfaces of the rocket or
manages the total energy
those systems will be
active during flight of the
full-scale rocket.
LBR’s payload does
not changes the
external surface of
the rocket nor
manages the energy
but if any changes
were LBR will
follow these
criteria.
In Progress
2.19.4. The full-scale
motor does not have to
be flown during the
full-scale test flight.
However, it is
recommended that the
full-scale motor be
used to demonstrate
full flight readiness
and altitude
verification. If the full-
scale motor is not
flown during the full-
scale flight, it is
desired that the motor
simulates, as closely
as possible, the
predicted maximum
velocity and maximum
acceleration of the
launch day flight.
Verify that the full-scale
motor is not required
during the full-scale test
flight.
LBR plans to use
the full-scale motor
during the full-scale
flight test.
However, if any
changes were
necessary and the
full scale motor is
not used LBR will
plan to use a motor
as close to the
actual as possible.
In progress
2.19.5. The vehicle
must be flown in its
fully ballasted
configuration during
the full-scale test
flight. Fully ballasted
refers to the same
amount of ballast that
will be flown during
Verify that the launch
vehicle will be flown in
its fully ballasted
configuration during the
full-scale flight test.
LBR will be
launching their
launch vehicle in its
fully ballasted state
during the full-scale
test flight.
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184 Long Beach Rocketry | CDR 2017 - 2018
the launch day flight.
Additional ballast may
not be added without a
re-flight of the full-
scale launch vehicle.
2.19.6. After
successfully
completing the full-
scale demonstration
flight, the launch
vehicle or any of its
components will not
be modified without
the concurrence of the
NASA Range Safety
Officer (RSO).
Verify after the full-scale
flight no components will
be modified without the
concurrence of the
NASA RSO.
After the successful
full-scale
demonstration flight
any modifications
to any of the launch
vehicle's
components will not
be made without the
concurrence of the
NASA RSO.
In progress
2.19.7. Full scale
flights must be
completed by the start
of FRRs (March 6th,
2018). If the Student
Launch office
determines that a re-
flight is necessary,
then an extension to
March 28th, 2018 will
be granted. This
extension is only valid
for re-flights; not first-
time flights.
Verify all full-scale
flights must be completed
by the start of FRRs but
if necessary for a re flight
an extension will be
granted
LBR plans to have
all full-scale flights
completed by FRRs.
If a re-flight is
scheduled LBR has
made plans to
account for this in
their scheduling.
In progress
2.20. Any structural
protuberance on the
rocket will be located
aft of the burnout
center of gravity.
Verify any structural
protuberance of the
rocket will be located aft
if the burnout center of
gravity.
LBR’s airbrake
system will
protrude aft of the
burnout center of
gravity.
In progress
2.21. Vehicle
Prohibitions 2.21.1.
The launch vehicle
will not utilize
forward canards.
2.21.2. The launch
vehicle will not utilize
Verify that all the vehicle
prohibitions will be
followed by LBR.
LBR has followed
everyone the
vehicle prohibitions
when designing the
rocket. If any
changes are made
with the launch
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185 Long Beach Rocketry | CDR 2017 - 2018
forward firing motors.
2.21.3. The launch
vehicle will not utilize
motors that expel
titanium sponges
(Sparky, Skidmark,
MetalStorm, etc.)
2.21.4. The launch
vehicle will not utilize
hybrid motors. 2.21.5.
The launch vehicle
will not utilize a
cluster of motors.
2.21.6. The launch
vehicle will not utilize
friction fitting for
motors. 2.21.7. The
launch vehicle will not
exceed Mach 1 at any
point during flight.
2.21.8. Vehicle ballast
will not exceed 10%
of the total weight of
the rocket
vehicle LBR will be
careful to not
violate any of these
verifications.
Table 9.3: Recovery Requirements
Minimum Success
Requirement
Verification Method
Verification Plan Status
3.1. The launch
vehicle will stage the
deployment of its
recovery devices,
where a drogue
parachute is
deployed at apogee
and a main parachute
is deployed at a
lower altitude.
Tumble or streamer
recovery from
apogee to main
parachute
deployment is also
Verify that the launch
vehicle will stage the
deployment of its
recovery devices.
The launch vehicle
will utilize a dual
deployment recovery
system to execute a
successful deployment
of the drogue
parachute at apogee
(5,280-ft) and the
main parachute at 500-
ft.
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186 Long Beach Rocketry | CDR 2017 - 2018
permissible,
provided that kinetic
energy during
drogue-stage descent
is reasonable, as
deemed by the
RSO.
3.2. Each team must
perform a successful
ground ejection test
for both the drogue
and main parachutes.
This must be done
prior to the initial
subscale and full-
scale launches
Verify that each team
will perform a
successful ground
ejection test for both
the drogue and main
parachutes before
subscale and full-
scale launches.
LBR will utilize black
powder ground
ejection charge testing
(primary and backup)
to ensure the drogue
and main parachutes
successfully eject from
the drogue and main
airframes. Ground
ejection testing will be
conducted prior to
every subscale and
full-scale launch to
ensure maximum
reliability.
In progress
3.3. At landing, each
independent sections
of the launch vehicle
will have a
maximum kinetic
energy of 75 ft-lbf
Verify that at landing,
each independent
sections of the launch
vehicle will have a
maximum kinetic
energy of 75 ft-lbf
Using known
equations for descent
velocity, drift distance
and kinetic energy, the
recovery system will
utilize a 20” drogue
with a 96” main that
will create a kinetic
energy less than the
maximum allowed,
ensuring that each
section of the launch
vehicle does not
exceed a kinetic
energy of 75-ft-lbf.
In progress
3.4. The recovery
system electrical
circuits will be
completely
independent of any
Verify that the
recovery system
electrical circuits will
be completely
independent of any
The launch vehicle
utilizes its own Single
Avionics Bay that is
separate from the
Payload Bay,
separating both
In progress
187 Long Beach Rocketry | CDR 2017 - 2018
payload electrical
circuits.
payload electrical
circuits.
electronics and circuits
from each other
entirely.
3.5. All recovery
electronics will be
powered by
commercially
available batteries.
Verify that all
recovery electronics
will be powered by
commercially
available batteries.
The recovery system
electronics will utilize
a standard 9V
commercial battery as
its main power source.
In progress
3.6. The recovery
system will contain
redundant,
commercially
available altimeters.
The term
“altimeters” includes
both simple
altimeters and more
sophisticated flight
computers.
Verify that the
recovery system will
contain redundant,
commercially
available altimeters.
The launch vehicle
avionics bay will
house two PerfectFlite
StratoLoggerCF
altimeters to maintain
redundancy and
reduce the risk of
recovery failure.
In progress
3.7. Motor ejection is
not a permissible
form of primary or
secondary
deployment.
Verify that the motor
ejection is not a
permissible form of
primary or secondary
deployment.
The launch vehicle
motor will not be used
as a form of primary
or secondary
deployment.
Completed
3.8. Removable
shear pins will be
used for both the
main parachute
compartment and the
drogue parachute
compartment.
Verify that
removeable shear
pins will be used for
both the main
parachute and the
drogue parachute
compartments.
The launch vehicle’s
main and drogue
parachute airframes
will utilize four ⅛”
shear pins on both the
front and aft ends of
the Avionics Bay as
the nylon pins for the
separation events.
In progress
3.9. Recovery area
will be limited to a
2500 ft. radius from
the launch pads
Verify that the launch
vehicle will descend
and land within the
2,500-ft recovery
area.
Utilizing known
equations on drift
distance and descent
velocity and with the
20” drogue and 96”
main being used, the
launch vehicle will
recover under the
specified recovery
In progress
188 Long Beach Rocketry | CDR 2017 - 2018
radius of 2500-ft.
through the drogue
and main parachute
deployments.
3.10. An electronic
tracking device will
be installed in the
launch vehicle and
will transmit the
position of the
tethered vehicle or
any independent
section to a ground
receiver. 3.10.1. Any
rocket section, or
payload component,
which lands
untethered to the
launch vehicle, will
also carry an active
electronic tracking
device. 3.10.2. The
electronic tracking
device will be fully
functional during the
official flight on
launch day
Verify that an
electronic tracking
device will be
installed in the launch
vehicle and will
transmit the position
of the tethered
vehicle or any
independent section
to a ground receiver;
any separate section
that lands untethered
to the launch vehicle
will also carry an
electronic tracking
device and each
electronic tracking
device will be fully
functional during
flight.
The launch vehicle’s
Avionics Bay will
utilize the Big Red
Bee BRB900
electronic GPS
transmitter device and
receiver to transmit the
position of any
tethered vehicle or
independent section of
the launch vehicle.
The current design of
the launch vehicle
yields no untethered or
independent sections
during flight. The
payload will deploy
the rover after landing.
The BRB900
electronic GPS device
will undergo multiple
ground testing with
recorded data to
ensure reliability
during flight.
In progress
3.11. The recovery
system electronics
will not be adversely
affected by any other
on-board electronic
devices during flight
(from launch until
landing). 3.11.1. The
recovery system
altimeters will be
physically located in
a separate
compartment within
the vehicle from any
other radio frequency
Verify that recovery
system electronics
will not be adversely
affected by any other
on-board electronic
devices during flight;
verify the altimeters
will be separate from
any frequency
transmitting device
and that all recovery
system electronics are
shielded from any
transmitting
device or magnetic
The Avionics Bay will
have its own separate
section on the launch
vehicle with dedicated
electronics to ensure
that it is not affected
by other on-board
electronic devices
during flight. The
StratoLoggerCF
altimeters will be
located within the
avionics bay on the
front wooden housing
tray on the top side of
In progress
189 Long Beach Rocketry | CDR 2017 - 2018
transmitting device
and/or magnetic
wave producing
device. 3.11.2. The
recovery system
electronics will be
shielded from all
onboard transmitting
devices, to avoid
inadvertent
excitation of the
recovery system
electronics. 3.11.3.
The recovery system
electronics will be
shielded from all
onboard devices
which may generate
magnetic waves
(such as generators,
solenoid valves, and
Tesla coils) to avoid
inadvertent
excitation of the
recovery system.
3.11.4. The recovery
system electronics
will be shielded from
any other onboard
devices which may
adversely affect the
proper operation of
the recovery system
electronics.
wave generating
devices. Verify that
the recovery
electronics will not be
adversely affected by
other onboard
devices.
the 3D printed
avionics center casing,
ensuring separation
from any the BRB900
GPS system, which
will be attached on the
lower wooden housing
tray, separated by the
avionics center casing,
maintaining separation
between the altimeters
and frequency
transmitting devices.
Since the Avionics
Bay internal structure
is separated by a 3D
printed center housing
case, the recovery
electronics will not be
affected by any
onboard transmitting
devices. The avionics
bay will be its own
separate section and
has its own dedicated
recovery system
electronics in the
airframe, removing
any devices that may
generate magnetic
waves. Since the
avionics bay will be its
own separate section
any other onboard
devices which may
affect the proper
operation of the
recovery system
electronics will be
physically located in a
separate section of the
launch vehicle..
190 Long Beach Rocketry | CDR 2017 - 2018
Table 9.4: Deployable Rover Requirements
Minimum
Success
Requirement
Verification
Method
Verification Plan Status
4.5.1. Teams
will design and
manufacture a
custom rover
that will deploy
from the
internal
structure of the
launch vehicle.
Test The team will complete a
full-scale rover that must be
capable of autonomously
traversing rough terrain and
deploy solar cells. Rover
will be placed in launch
vehicle and deployed
through the RDM.
Subscale prototype rover
deployment was
successful. Full-scale
rover test will be verified
by February.
4.5.2. At the
launch vehicle’s
landing site, the
team will
remotely
activate a
trigger to
deploy the rover
from the rocket.
Test Team will flip a switch on a
remote to activate
deployments mechanism.
Rover must be completely
deployed and capable of
driving away even under
non-ideal conditions.
Subscale RDM deployed
rover after launch. Full
scale RDM will be
verified by February.
4.5.3. After
deployment, the
rover will
autonomously
move at least 5
ft. (in any
direction) from
the launch
vehicle.
Test Rover must autonomously
choose best direction to go
and maintain a straight
heading until the distance
surpasses 5 feet.
Will be verified by
February.
4.5.4. Once the
rover has
reached its final
destination, it
will deploy a set
of foldable solar
cell panels.
Test Rover must autonomously
know when to deploy solar
panels and completely open
them up regardless of the
position the rover is in.
Will be verified by
February.
191 Long Beach Rocketry | CDR 2017 - 2018
Table 9.5: Safety Requirements
Minimum Success Requirements Verification Method Verification Plan
5.1. Each team will use a launch
and safety checklist. The final
checklists will be included in the
FRR report and used during the
Launch Readiness Review (LRR)
and any launch day operations.
LBR will create safety
checklist that will be
included in FRR report
and Launch Readiness
Review.
LBR Safety Officer with
ensure that the safety
checklist are complete. They
will also verify that safety
checklist is used during the
launch day operations.
5.2. Each team must identify a
student safety officer who will be
responsible for all items in section
5.3.
Shawn Everts is the
official Safety Officer
for LBR.
He will assure that the team
adheres to all regulations
pertaining to the
construction, assembly,
testing, flight, and recovery
phases of the launch vehicle.
5.3. The role and responsibilities of
each safety officer will include, but
not limited to: 5.3.1. Monitor team
activities with an emphasis on
Safety during: 5.3.1.1. Design of
vehicle and payload 5.3.1.2.
Construction of vehicle and
payload 5.3.1.3. Assembly of
vehicle and payload 5.3.1.4.
Ground testing of vehicle and
payload 5.3.1.5. Sub-scale launch
test(s) 5.3.1.6. Full-scale launch
test(s) 5.3.1.7. Launch day 5.3.1.8.
Recovery activities 5.3.1.9.
Educational Engagement Activities
LBR Safety Officer will
be responsible for all
subsections listed.
LBR Safety Officer will be
present during the design,
construction, assembly,
ground testing, sub-scale
launch, full-scale launch,
launch day, recovery and
education engagement
activities. The Safety Officer
will be monitoring all these
activities for safety concerns.
5.3.2. Implement procedures
developed by the team for
construction, assembly, launch, and
recovery activities
All members of the Long
Beach Rocketry team
will be given a safety
briefing and are required
to sign a safety contract.
These briefings will cover
proper procedures for
construction, assembly,
launch and recovery
activities.
5.3.3. Manage and maintain current
revisions of the team’s hazard
analyses, failure modes analyses,
procedures, and MSDS/chemical
inventory data
LBR Safety Officer will
be responsible for
maintaining the team’s
hazard analysis, failure
modes analysis,
procedures, and
MSDS/chemical
inventory data.
LBR Safety Officer will be
briefed on all components of
the launch vehicle from each
subsystems lead to be able to
properly maintain hazard
analyses and failure mode
analyses.
192 Long Beach Rocketry | CDR 2017 - 2018
5.3.4. Assist in the writing and
development of the team’s hazard
analyses, failure modes analyses,
and procedures.
LBR Safety Officer will
be responsible for
writing the team’s
hazard analysis, failure
modes analysis,
procedures, and
MSDS/chemical
inventory data.
LBR Safety Officer will have
a understanding of all
components of the launch
vehicle to be able to properly
write the hazard analyses and
failure mode analyses.
5.4. During test flights, teams will
abide by the rules and guidance of
the local rocketry club’s RSO. The
allowance of certain vehicle
configurations and/or payloads at
the NASA Student Launch
Initiative does not give explicit or
implicit authority for teams to fly
those certain vehicle configurations
and/or payloads at other club
launches. Teams should
communicate their intentions to the
local club’s President or Prefect
and RSO before attending any
NAR or TRA launch
All members of the Long
Beach Rocketry Team
will follow the range
safety regulation.
LBR will follow the range
safety regulation as stated in
the Long Beach Rocketry
Team Safety Agreement that
all members have to read and
sign to be able to participate
in the project.
5.5. Teams will abide by all rules
set forth by the FAA.
LBR Safety Officer will
be responsible for
verifying that all rules
set forth by the FAA or
followed.
LBR Safety Officer has
created LBR Team Safety
Agreement which is sign by
all members saying that they
have read and understood all
rules set forth by the FAA.
193 Long Beach Rocketry | CDR 2017 - 2018
9.2 Timeline
Figure 9.1: Gantt Chart for 2017-2018 Competition
194 Long Beach Rocketry | CDR 2017 - 2018
Table 9.6: Expected Development Schedule
Competition Timeline
Task Expected Date of Completion
Proposal
PDR
Subscale Launch
CDR
Full-Scale Launch
FRR
Competition
9/20/17
11/3/17
11/4/17
1/12/18
2/3/18
3/5/18
4/4/18
Table 9.7: Expected Development on Each Team
Launch Vehicle
Task Start Date End Date
Brainstorm and design launch vehicle 8/28/17 9/20/17
Build subscale launch vehicle 9/24/17 11/4/17
Design full scale launch vehicle 11/6/17 12/17/17
Construction of full scale launch vehicle 12/17/17 2/17/18
Adjustment based on launch 2/3/18 2/3/18
Additional full-scale launch (if needed) 2/17/18 2/17/18
Rover Deployment Mechanism
Task Start Date End Date
Brainstorm solutions for rover deployment 8/28/17 9/10/17
Design concept and initial layout plan 9/10/17 9/20/17
195 Long Beach Rocketry | CDR 2017 - 2018
Build RDM prototype for the subscale 9/20/17 10/30/17
Integrate RDM with subscale launch vehicle 10/30/17 11/3/17
Subscale Launch (with prototype RDM on board) 11/4/17 11/4/17
Post-subscale-test flight analysis 11/4/17 11/19/17
Build RDM protected box for electronic components 11/19/17 12/17/17
Integrate RDM with full scale launch vehicle 12/17/17 2/17/18
Troubleshooting and adjustments 2/17/18 2/30/18
Autonomous Rover
Task Start Date End Date
Brainstorm solutions for the rover 8/28/17 9/20/17
Preliminary CAD 9/26/17 10/26/17
Preliminary analysis 10/20/17 11/1/17
Prototype construction 11/1/17 11/20/17
Prototype testing 11/20/17 12/17/17
Installing a foldable solar panel on the rover 12/17/17 1/5/18
Integrate the rover with RDM and launch vehicle 1/5/18 2/15/18
Test flight the experiment 2/3/18 2/3/18
Additional test flight (if needed) 2/17/18 2/17/18
Electrical & Software
Task Start Date End Date
Develop high level integration plans 9/1/17 10/4/17
Design, analyze, and refine all schematic designs 10/4/17 10/26/17
Build initial prototypes and order PCBs 10/26/17 11/15/17
Assemble and test individual PCBs 11/15/17 11/22/17
196 Long Beach Rocketry | CDR 2017 - 2018
Integrate assembled modules; revise if needed 11/22/17 11/29/16
Complete system level integration testing between all modules 11/29/17 1/17/18
Complete integration with the rover 1/17/18 2/1/18
Perform full-scale integration testing 2/1/18 2/3/18
Design final revision and create backups 2/28/18 3/8/18
9.3 Budget
LBR has a projected budget of approximately $14,200 which greatly exceeds their original
estimate of $12,000. This is due to the addition of airbrakes on the launch vehicle requiring a
larger motor and different motor casing to house it. These efforts have been made in order to
secure new sponsors for this year in order to account for this increase in budget detailed in table
9.8.
Table 9.8: Projected Expense
ROVER
MATERIALS
AndyMark
1/2" Hex Bore, Flanged, Heavy Duty Inner
Race Shielded Ball Bearing (FR8ZZ-
HexHD) (am-2986) $6.00 18 $108.00
RC4WD
SCRAMBLER OFFROAD 1.0" SCALE
TIRES $8.99 3 $26.97
AndyMark
25 Tooth 20 DP 0.500" Hex Bore, Steel
Gear (am-3535) $12.00 6 $72.00
ServoCity
Hex Bore, Face Tapped Clamping Hubs,
0.770" Pattern 545674 $7.99 2 $15.98
AndyMark
Collar Clamp, 1/2 Hex Bore, 2 Pc with Flats
(am-2871) $7.00 2 $14.00
McMaster-Carr
Turnbuckle-Style Connecting Rod 5/16 '' -
24 Internal Thread, 4 '' Overall Length $18.68 2 $37.36
ServoCity 116 RPM Premium Planetary Gear Motor $27.99 2 $55.98
AndyMark
35 Tooth 20 DP 0.500" Hex Bore, Steel
Gear (am-3486) $15.00 2 $30.00
197 Long Beach Rocketry | CDR 2017 - 2018
AMain Hobbies
Turnbuckle-Style Connecting Rod 5/16"-24
Internal Thread, 2" Overall Length $17.89 6 $107.34
Amazon black abs filament $23.99 1 $23.99
ServoCity 1.00" Bore 32 Pitch Aluminum Hub Gears $12.99 2 $25.98
ServoCity
4mm Bore 32 Pitch, 16T Shaft Mount
Pinion Gear $7.99 2 $15.98
Total $533.58
RDM
McMaster-Carr
Medium-Strength Grade B7 Steel Threaded
Rod $12.38 1 $12.38
McMaster-Carr
High-Parallel-Misalignment Flexible Shaft
Coupling (1/4 in) $14.46 1 $14.46
McMaster-Carr
High-Parallel-Misalignment Flexible Shaft
Coupling (4mm) $14.46 1 $14.46
McMaster-Carr
93 in.-lbs. Acetal Disc for 3/4" OD High-
Parallel-Misalignment Flexible Shaft
Coupling $2.81 1 $2.81
ServoCity
32 RPM Premium Planetary Gear Motor
w/Encoder $49.99 1 $49.99
McMaster-Carr
Stainless Steel High-Torque 12-Point Flange
Nuts $3.54 2 $7.08
McMaster-Carr 6061 Aluminum Rod with Certification $15.27 1 $15.27
McMaster-Carr Zinc-Plated Steel Coupling Nut $1.87 4 $7.48
McMaster-Carr
316 Stainless Steel Washer Oversized, 1/4"
Screw Size, 0.281" ID, 3" OD $2.88 1 $2.88
McMaster-Carr
Medium-Strength Grade B7 Steel Threaded
Rod $12.38 1 $12.38
McMaster-Carr
Durometer 98A Spider for 3/4" OD for
Clamping Vibrate-Damping Precision
Flexible Shaft Coupling $8.89 1 $8.89
McMaster-Carr
Vibrate-Damping Precision Flexible Shaft
Coupling Clamping Hub, 1-7/64" Overall
Length, 3/4" OD $13.22 1 $13.22
McMaster-Carr
Vibrate-Damping Precision Flexible Shaft
Coupling Clamping Hub, 1-7/64" Overall
Length, 3/4" OD $17.54 1 $17.54
Total $178.84
198 Long Beach Rocketry | CDR 2017 - 2018
AVIONICS
PerfectFlite Direct StratoLogger CF Altimeter $49.46 2 $98.92
PerfectFlite Direct
Altimeter Mounting Hardware (Standoff
Screw) $1.79 2 $3.58
Mouser 2-Pole Rotary Arming Switch $6.48 5 $32.40
BigRedBee
BRB900 TX/RX Base GPS BigRedBee (w/
LCD) $378.00 1 $378.00
Wooden Plate
Walmart 9 Volt Batteries $11.51 1 $11.51
Walmart Digital Gram Scale $8.23 1 $8.23
Walmart 9 Volt Battery Clip Connector (5 pcs) $5.99 1 $5.99
Total $538.63
RECOVERY
Fruity Chutes 24" Elliptical Parachute - 2.2lbs @ 20fps $60.00 1 $60.00
Fruity Chutes
Iris Ultra 72" Compact Parachute - 28lbs @
20fps $265.00 1 $265.00
Giant Leap Rocketry
Shock Cord Protector Sleeves (Length of
30") $9.92 2 $19.84
Giant Leap Rocketry
Parachute Protective Blanket (Up to 7.5"
Airframe) $12.45 2 $24.90
Fruity Chutes
Parachute Protective Blanket (Up to 7.5"
Airframe) $43.00 1 $43.00
Strap Works
3/8" Shockcord~Flat Nylon Webbing
(1400lbs TEST) $0.23 per foot $0.23 100 $23.00
McMaster-Carr Steel Eyebolt with Shoulder - for Lifting $3.20 2 $6.40
McMaster-Carr
3/8" Thickness, 1/2" Opening, Quicklink
(2,200 lbs test) $3.47 4 $13.88
McMaster-Carr
U-Bolt~Galvanized Steel with Mounting
Plate, 3/8"-16 (1,075 lbs Test) $2.23 4 $8.92
McMaster-Carr
Type 18-8 Stainless Steel Flat Washer 3/8"
(100 Per Pack) $5.01 1 $5.01
McMaster-Carr
Threaded Rod 5/16"-18 Thread, 2 Foot
Long (Grade 8 Steel) $8.20 2 $16.40
199 Long Beach Rocketry | CDR 2017 - 2018
McMaster-Carr
High-Strength Steel Hex Nut Grade 8, Zinc
Yellow-Chromate Plated, 3/8"-16 Thread
Size (pack of 100) $7.46 1 $7.46
The Thread Exchange
Kevlar - Size 46 (Tex 40) - Natural -
Bonded - Nominal 1/2 Oz Spool - 312 Yards
- Strength 14 Lbs $15.00 1 $15.00
McMaster-Carr
Black Nylon Pan Head Machine Screw
Phillips, 2-56 Thread, 1/2" Length (packs of
100) $5.67 1 $5.67
McMaster-Carr
Zinc-Plated Steel Pan Head Phillips
Machine Screw4-40 Thread, 1/2" Length
(100 per pack) $1.75 1 $1.75
McMaster-Carr
Low-Strength Steel Hex Nut Zinc-Plated, 4-
40 Thread Size (100 per pack) $0.87 1 $0.87
Total $517.10
LAUCH
VEHICLE
Madcow Rocketry 6'' Fiberglass G12 60'' length $288 2 $576
Madcow Rocketry 6'' Fiberglass G12 14'' length coupler $60 2 $120
Apogee Rockets CTI 4263-L1350-CS-P Motor $185 3 $555
Chris’ Rocket Supplies Cesaroni 75-3 Grain Hardware Set $289.95 1 $289.95
Madcow Rocketry 4" Fiberglass G12 12" length coupler $32 1 $32
Madcow Rocketry 4" Fiberglass G12 60" length $116 2 $232
Bay Area Rocketry AeroTech K1499 $134.99 2 $269.98
McMaster-Carr Aluminum sheet $120 1 $120
McMaster-Carr Carbon Fiber Sheet $100 1 $100
Total $2,249.93
TRAVEL
Competition Transportation $5,000 1 $5,000
Competition Lodging $2,000 1 $2,000
Sponsorship Materials $1,000 1 $1,000
Website Domain $70 1 $70
Shipping and Taxes $600 1 $600
Total $8,670
200 Long Beach Rocketry | CDR 2017 - 2018
Total Competition Expense $13,870.91
Table 9.9: Projected Income
College of Engineering $4,200
AIAA - CSULB $1,500
Fundraisers $1,500
ASI Travel Grant $7,000
Total Income $14,200
Figure 9.2: Budget and Present Expenditures
0
1000
2000
3000
4000
5000
6000
7000
8000
9000
DORITO Materials RDM Avionics Recovery Launch Vehicle Travel
Am
ou
nt
in D
olla
r($
)
Category
Total Amount Spent for each Category
Amount Spent Amount Budgeted
201 Long Beach Rocketry | CDR 2017 - 2018
The team tries to be as material efficient as possible and always considers cost in the decision-
making process. The team lead will continue to create budget projections and monitor the team’s
spending as the year progresses.
9.4 Funding Plan
Based on our initial proposed budget the initial funding for this project will be $12,140. To reach
this requirement LBR has reached out to California State University College of Engineering, and
CSULB AIAA chapter, ASI and various corporations for sponsorships, seen in Figure XX .
Figure 9.3: Fulfillment Percentages of Various Funding Methods
College of Engineering
AIAA’s team will be working with the College of Engineering to complete the IRA
Annual Funding Request form in order to secure funding. The goal of the Instructionally Related
Activities program is to provide student fee funding for out-of-class experiences for students
participating in an academic program, discipline, research or department where those
experiences are integrally related to one of its instructional courses. Such activities are deemed
essential for providing a quality educational program and constitute a vital and enhanced
instructional experience for students. This year LBR received $1,000 from the college of
Engineering.
Associated Student Body Travel Grant
The Associated Student Body (ASI) at CSULB offers student travel grants. This grant covers
airfare and accommodation for instruction related activities or students participating in
competitions representing the school. This grant will be applied for later in the project when
202 Long Beach Rocketry | CDR 2017 - 2018
travel arrangements are being made. LBR plans to receive $7,000 from ASI to completely cover
our travel costs.
AIAA
CSULB's AIAA chapter will be using funding obtained from fundraising and student
membership fees to fund this project. The goal of AIAA chapter is to expose students to the
Aerospace Industry and to show that the Aerospace industry is very diverse and filled with every
different major from the STEM field, meaning students can have more to look forward to in
joining the Aerospace Industry. This year LBR received $1,000 from AIAA which will exclusive
used for gas money and food which is not recorded in our budget. The fundraising total is
separate from this $1,000 and LBR plan to use to fulfill any gaps in the travel request from ASI
if they are unable to complete our travel request in full. The events LBR will host will be
transferred over from last year via Staff Pro and ice skating event. Both of which were highly
effective fundraising events last year and will be used to fill any gaps in our funding.
Associated Student Body Research Grant
The Associated Student Body (ASI) at CSULB offers student research grants throughout the
year. The grant will cover the cost of parts and materials needed for outreach. The team manager
will submit the grant application and create a presentation to be presented in front of the ASI
grant committee. This year LBR received $4,200 from the research grant which will fund
majority of the manufacturing of this project
Ice Skating Fundraising
LBR will host an ice skating fundraiser at the Lakewood ICE rink to raise additional funds for
competition travel. Each ticket will cost approximately $15 and LBR will receive 53% of the
profits. Last year, approximately 50 people attended and $400 was raised. This is the second time
LBR is hosting an ice skating event to fundraise and hopes that this will be as effective as last
year.
203 Long Beach Rocketry | CDR 2017 - 2018
Figure 9.4: Sponsorship tier system
Sponsorships and Donations
In order to cover the remaining cost of the competition, LBR has created a sponsorship proposal
to submit to companies offering advertising on the rocket, on social media platforms, on team
shirts, and displayed banners, which will be present at all team events during outreach and
competition. Additionally, LBR created a tier system to offer to donors, a sample of the tier
system has been provided below in Figure 9.4.
9.5 Educational Engagement
Outreach Objective
LBR aims to encourage and support students, children, professors, and organizations within the
community to pursue careers in STEM and apply their knowledge of engineering to practical and
rewarding applications.
Moving forward, LBR plans to revamp their outreach programs by strengthening relationships
with professor’s organizations and programs in order to engage a larger range of individuals
within the Long Beach campus community. Additionally, LBR hopes to broaden their influence
and reach out to members of the local community. Through events with other educational
204 Long Beach Rocketry | CDR 2017 - 2018
institutions, LBR members hope to inspire students and youth alike to pursue STEM careers by
showcasing what engineers in these fields have accomplished before.
In LBR’s first year they had exceeded expectations by reaching out to 900 students. This year
LBR is setting their expectations even higher and are aiming to surpass previous year’s results.
Past and Planned Outreach Events
Girls Day at the Beach
Last March, LBR participated in Girls Day at the Beach alongside the university’s Society of
Women Engineers (SWE). Elementary school girls visited the team’s laboratory for a tour of the
team’s projects and equipment. Additionally, the team constructed and programmed remote
control vehicles that the girls were able to test and drive.
As a result of previous success, LBR will participate in Girls Day at the Beach again. This time,
the team plans to continue to showcase past projects and equipment as well as explain current
activities. Girls from k-12 schools will see the LBR 2016-2017 launch vehicle, payload,
parachutes and motor casings. Additionally, the girls will observe demonstrations of the lab’s
new Virtual Reality room.
Aerospace Rocket Symposium
On September 7th, the LBR team, alongside AIAA and other CSULB rocket teams, held a booth
at the Aerospace Rocket Symposium that took place at UC Irvine. The team displayed past
projects to students attending the event. Through their presentation, they explained the design
process and requirements it takes to participate in a NASA competition. The LBR team hoped to
lead prospective aerospace students by example by encouraging them to get involved in projects
and research once in college.
Intro to Engineering Presentations
To support engineering students within the university, the team will be conducting presentations
in “Introduction to Engineering” courses. The goal is to get incoming freshman excited about
engineering and the things they can achieve when pursuing a STEM career. MAES Latinos in
Engineering Bottle Rocketry
LBR will cooperate with MAES to teach local middle schoolers about basic rocketry. The
presentation concludes with the students applying their newly found knowledge to build bottle
rockets and experience their first launch.
High School Engineering Presentation
LBR members will return to their high school’s STEM programs in order to give presentations
about rocketry concepts, engineering in college, and advice in order to get involved in hands on
projects like USLI in college. This has been a popular event in the past but has fallen out of
practice, so they plan to revive it this year returning to our alumni colleges. Previously it was a
highly successful event that many LBR members enjoyed participating in. They are expecting
the same result in the following year.
205 Long Beach Rocketry | CDR 2017 - 2018
Table 9.10: Educational Engagement Schedule
Event Date Estimated Attendees
Girls Day at the Beach (1) 3/2017 100
Aerospace Rocket Symposium 9/7/2017 200
Girls Day at the Beach (2) 9/2017 200
Introduction to Engineering Presentations 11/2017 100
MAES Latinos in Engineering Bottle Rocketry 4/2018 60
High School Engineering Presentation 12/2018 500
TOTAL 1160
206 Long Beach Rocketry | CDR 2017 - 2018
Section 10: Appendices
10.1 References
[6.1] Servo City 118 RPM HD Premium Planetary Gear Motor w/Encoder. (Date accessed:
2017-12-12). [Online] Available: https://www.servocity.com/118-rpm-hd-premium-
planetary-gear-motor-w-encoder
[6.2] McMaster-Carr Nylon Pan Headed Slotted Screws. (Date accessed: 2016-12-
14) [Online] Available: https://www.mcmaster.com/#92942a716/=1b28l7i
[7.1] density_accepted_values. [Online]. Available:
http://honorsph.startlogic.com/honorsphysicalscience/labs/density_accepted_values.htm.
[Accessed: 11-Jan-2018].
[7.2] “Titanium - Element information, properties and uses | Periodic Table,” Royal Society of
Chemistry - Advancing excellence in the chemical sciences. [Online]. Available:
http://www.rsc.org/periodic-table/element/22/titanium. [Accessed: 11-Jan-2018].
[7.3] Intashu, and Seth Stiles. “Scrambler Offroad 1.0.” RC4WD Store,
store.rc4wd.com/Scrambler-Offroad-10-Scale-Tires_p_5142.html.
[7.4] Edge, LLC. Engineers. “Lewis Factor Equation for Gear Tooth Calculations.” Engineers
Edge, www.engineersedge.com/gears/lewis-factor.htm.
[7.5] Banggood.com, “5V 30MA 0.15w 53 x 30 x 3mm Polycrystalline Mini Solar Panel
Photovoltaic Panel,” www.banggood.com. [Online]. Available:
https://www.banggood.com/5V-30MA-0_15w-53-x-30-x-3mm-Polycrystalline-Mini-
Solar-Panel-Photovoltaic-Panel-p-1020641.html?cur_warehouse=CN. [Accessed: 11-Jan-
2018].
[7.6] “HS-7950TH High Voltage, Ultra Torque, Titanium Gear, Coreless Ultra Premium
Servo,” HiTec. [Online]. Available: http://hitecrcd.com/products/servos/ultra-premium-
digital-servos/hs-7950th-ultra-torque-hv-coreless-titanium-gear-servo/product.
[Accessed: 11-Jan-2018].
[7.7] “Redcat Racing Steering Link: Toys & Games,” Amazon.com: Redcat Racing Steering
Link: Toys & Games. [Online]. Available: https://www.amazon.com/Redcat-Racing-
02074-Steering-Link/dp/B00E5L6LFS/ref=sr_1_2?s=toys-and-
games&ie=UTF8&qid=1515566361&sr=1-
2&keywords=Redcat%2BRacing%2BSteering%2BLink. [Accessed: 11-Jan-2018].
[7.8] Arduino Nano, store.arduino.cc/usa/arduino-nano.
[7.9] Fezder, “Full-Bridge Motor Driver Dual - L298N,” COM-09479 - SparkFun Electronics.
[Online]. Available: https://www.sparkfun.com/products/9479. [Accessed: 11-Jan-2018].
[7.10] “MPU-6050 Accelerometer Gyro.” Arduino Playground - MPU-6050,
playground.arduino.cc/Main/MPU-6050.
[7.11] “116 RPM Premium Planetary Gear Motor,” ServoCity.com. [Online]. Available:
https://www.servocity.com/116-rpm-premium-planetary-gear-motor. [Accessed: 11-Jan-
2018].
207 Long Beach Rocketry | CDR 2017 - 2018
[7.12] “Battery Life Calculator,” Battery Life Calculator | DigiKey Electronics. [Online].
Available: https://www.digikey.com/en/resources/conversion-calculators/conversion-
calculator-battery-life. [Accessed: 11-Jan-201