Long Beach Rocketry | CDR 2017 - 2018...Long Beach Rocketry | CDR 2017 - 2018 4 Figure 5.43...

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


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


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


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 7.13: Rover gearbox system

Figure 7.14: Rover motor mount



Figure 7.17: Solar panel

Figure 7.18: Deployment mechanism before and after deployment side view


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


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 5.10: 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


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


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


AV- Avionic


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 [email protected]

(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 [email protected]

(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.


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


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

[email protected]

562-985-1518

Team Mentor

David Alexander Roy

Visual Artist, High Power Rocketry enthusiast

Otis College of Art and Design

[email protected]

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.


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


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.


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.


• 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.


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.


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.


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


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


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.


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.


C - Occasional Expected to occur several times or

occasionally within time.


D - Remote Unlikely to occur, but can be reasonably

expected to occur at some point within

time.


E - Improbable Very unlikely to occur and an occurrence

is not expected to be experienced within

time.

Probability < 5%


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


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


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


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



proper handling of

material will be


the MSDS




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




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


the chemicals MSDS

2E


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


students will be


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



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


students will be


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


will be supervised by

university personnel

that are aware of the

shop safety practices

such as the proper PPE

for the shop

2E


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


students will be


follow proper

operations when

working with shop

tools

2. Per section 4.4






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






such as the proper PPE

for the shop

3E


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


RAC


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



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


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


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


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 the motor

will done by

NAR/TRA team

mentor

2. Procedure available

in section 4.13.2

“Motor Preparation”

1E


Table 4.8: Failure Mode Hazard Risk Assessment – Recovery

Recovery Risk Analysis


RAC


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


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


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


“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



ensures that the

avionics system

altimeter beeps are

checked and give the

correct results before

the rocket is launched

1E


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


Table 4.9: Failure Mode Hazard Risk Assessment – Rover

Rover Risk Analysis


RAC


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


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


Table 4.10: Failure Mode Hazard Risk Assessment – Rover Deployment (RDM)

Rover Deployment (RDM) Risk Analysis


RAC


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




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


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




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


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


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


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


RAC


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



proper storage of

material will be


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


the procedure will

ensure that launch site

is check for trees

before the rocket is

launched

2E


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


4.13.3 “Setup on

Launcher” wind speed

will be measured

before the launch

occurs.

2. Section 4.13.3


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



proper storage of

material will be


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


procedure states to

check the sky before

launch to look for

clouds and abort

launch if conditions

are excessively cloudy

3E


Table 4.12: Environmental Risk Assessment – Rocket Impact on Environment

Rocket Impact on Environment


RAC


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


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


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


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



proper handling of

material will be


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


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


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.


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. _______________________________


Grease

Cesaroni L1350 CS-P Motor

Forward Extended Plugged Tapped Closure

Motor Casing

Aft Closure

54mm Retainer

End Cap


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.


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. _______________________________


Pen or pencil

Level 2 Certification card.

Rocket

Featherweight Screw Switch (Arming Switch) (2x)


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.


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. _______________________________


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. _______________________________


Closed toe shoes


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.


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.


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.


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


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


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.


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.


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.


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)


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


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.


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


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


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.


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.


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.


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.


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


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)


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,


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


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


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.


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


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.


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.


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.


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


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.


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.


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


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


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.


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.


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


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.


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.


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.


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)


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.



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)


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.


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)


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.


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


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


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)


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


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


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


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.


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


𝐷𝑟𝑖𝑓𝑡 𝐷𝑖𝑠𝑡𝑎𝑛𝑐𝑒 = (28 × 7.33) + (52.61 × 7.33) = 591.14 𝑓𝑡




𝐷𝑟𝑖𝑓𝑡 𝐷𝑖𝑠𝑡𝑎𝑛𝑐𝑒 = (28 × 22) + (52.61 × 22) = 1773.42 𝑓𝑡




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


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.


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



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




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



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



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


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




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



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



The launch vehicle

ejection charges will

Test/Analysis Preliminary analysis

on the amount of

Will be verified



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.


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



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



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




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



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



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


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


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.


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


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


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


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


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.


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)


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 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.


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)


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


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


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.


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


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


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.


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.


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.


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.


Figure 6.16.1: Assembly of top of subscale motor bulkhead

Figure 6.16.2: Exploded view of top of subscale motor bulkhead


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


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

Table 6.5: Mass of subscale electronic components

Component Mass (oz)

Motor 4.48


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


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.


Table 6.8: Team derived requirements verification status and plan (RDM)


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


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.


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


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.


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


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


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


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.



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.



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.


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.


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.


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


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


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.


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


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.


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


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.


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


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.


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.


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.


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


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


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.


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].


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)


(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.


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.


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


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


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.


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


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.


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.


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








individuals.




falling.



10) Insert the motor into the motor tube and secure using the motor retainer


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.



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.






individuals.




falling.




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


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.


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


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.


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


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


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.


Section 9: Project Plan

9.1 Requirement Verification

Table 9.1: General Requirements

Minimum Success


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

Completed


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

Completed


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

In Progress


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

PDF

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


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

Completed


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.


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.

In progress


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


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


main vehicle using its

own parachute

2.8. The launch

vehicle will be limited

to a single stage.


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


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


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


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


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


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.

In Progress


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.


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

Completed


adhere to this

criteria.

2.17. The launch

vehicle will accelerate

to a minimum velocity

of 52 fps at rail exit.


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


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


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-

In Progress


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.


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


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


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


test flight.

In progress


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

Completed


forward firing motors.

2.21.3. The launch


motors that expel

titanium sponges

(Sparky, Skidmark,

MetalStorm, etc.)

2.21.4. The launch


hybrid motors. 2.21.5.

The launch vehicle

will not utilize a

cluster of motors.

2.21.6. The launch


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


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


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.

Completed


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


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


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


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


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..


Table 9.4: Deployable Rover Requirements

Minimum

Success

Requirement

Verification

Method


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.


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,


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


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


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.


5.3.4. Assist in the writing and

development of the team’s hazard

analyses, failure modes analyses,

and procedures.


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.


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.


9.2 Timeline

Figure 9.1: Gantt Chart for 2017-2018 Competition


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


Brainstorm solutions for rover deployment 8/28/17 9/10/17

Design concept and initial layout plan 9/10/17 9/20/17


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


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


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


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


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


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


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


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


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


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.


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


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.


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


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].

https://www.mcmaster.com/#92942a716/=1b28l7i


[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

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