University of Pittsburgh

Swanson School of Engineering

3700 O'Hara St,

Pittsburgh, PA 15213

Room 636

Flight Readiness Report

NASA Student Launch 2018-2019

March 4th, 2019

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1 Summary 2

1.1 Team Summary 3

1.2 Launch Vehicle Summary 3

1.2.1 Vehicle Body Design Review 3

1.3 Payload Summary 10

2 Changes Since CDR 10

3 Vehicle Criteria 12

3.1 Changes Since CDR 12

3.1.1 Mechanical 12

3.1.2 Avionics 13

3.2 Design and Construction of Vehicle 17

3.2.1 Vehicle Safety Features 17

3.2.2 Flight Reliability Confidence 24

3.2.3 Vehicle Construction Procedure 26

3.3 Recovery Subsystem 41

3.3.1 Dual Deploy Recovery System & Safety Features 41

3.3.3 Electronic Recovery Component Robustness and Failure Modes 48

3.4 Avionics Systems 53

3.4.1 Mission-Critical Avionics Components and Processes 53

Recovery 54

GPS 54

Ground Control 54

3.4.2 Structural Avionics Components 55

3.4.3 Rocket Locating System (GPS) 58

3.4.4 GPS Failure Modes 60

3.5 Mission Performance Predictions 63

3.6 Vehicle Demonstration Flight 66

4 Safety 67

4.1 Safety Mission Statement 67

4.2 Safety Officer Identified-Responsibilities Defined 67

4.2.1 Safety Officer Identified-Responsibilities 67

4.3 Safety and Environment (Vehicle and Payload) 70

4.3.1 Approach to Analysis of Failure Modes 70

4.3.2 Personnel Hazards Analysis 72

4.3.3 Failure Modes and Effects Analysis 79

4.3.4 Environmental Hazards Analysis 85

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4.4 Launch Operations Procedures 89

4.4.1 Pre-launch day Altimeter Preparation 89

4.4.2 Recovery Preparation Checklist 91

4.4.3 Motor Preparation Checklist 94

4.4.4 Setup on Launch Pad Checklist 94

4.4.5 Igniter Installation Checklist 95

4.4.6 Launch Procedure Checklist 96

4.4.7 Troubleshooting 96

4.4.8 Post-Flight Inspection 97

5 Payload Criteria 98

5.1 Changes Since CDR 98

5.2 Payload Features 99

5.2.1 Design Overview 99

5.2.2 Mechanical Elements 101

5.2.2.1 Rover 101

5.2.2.2 Retention and Deployment System 104

5.2.3 Electrical Elements 107

5.3 Schematics 111

5.3.1 As-Built Mechanical Schematics 111

5.3.2 As-Built Electronic Schematic and Block Diagram 112

5.4 Construction 114

5.5 Flight Reliability Confidence 117

5.6 Payload Demonstration Flight 117

6 Project Plan 118

6.1 Testing 118

6.1.1 Mechanical Tests 118

6.1.2 Avionics Tests 120

6.1.3 Payload Tests 142

6.2 Rules Derived Requirements 147

6.3 Team Derived Requirements 162

6.4 Budgeting and Timeline 164

6.4.1 Line Item Budget 164

6.4.2 Funding Plan 168

1 Summary

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1.1 Team Summary

The University of Pittsburgh’s Rocketry Team, or the Pitt Rocketry Team (PRT) consists of

approximately 25 members that contribute to one of four sub-teams: systems, mechanical,

avionics, and payload. Team members either build and design the rocket or rover, control the on-

board electronics, or manage the team’s finances, logistics, and community presence. To aid

students the team is working with Tripoli Pittsburgh, TRA Prefecture #001, and Pittsburgh Space

Command, NAR Chapter #473. The Pitt Rocketry Team can be contacted at 3700 O'Hara St,

Pittsburgh, PA 15213.

Figure 1.1a: Team organization.

Table 1.1a: Important contact information.

Name Team Role Contact Information Additional Information

Professor Matthew Barry Team Advisor mmb49@pitt.edu (412) 624-9031

Fernando Marill-Tabares Student Team Leader fjm17@pitt.edu (412) 313-1133

Thomas Sullivan Harrington Safety Officer tsh25@pitt.edu (267) 421-1399

Duane Wilkey NAR Mentor duane@velocity.net NAR Level 3 #6342

1.2 Launch Vehicle Summary

1.2.1 Vehicle Body Design Review

Vehicle Size 95 in

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Vehicle Mass 26.9 lbm (with maximum ballast)

Launch Day Motor CTI PR75 K555-2G-W

Target Altitude 4750 ft

Rail Size 1515

Motor Specifications:

Motor CTI, Pro-X

Diameter 75mm

Burn Time 4.33 seconds

Average Thrust 555.1 N (124.79 lb-s)

Max Thrust 646.7 N (145.38 lbs)

Total Impulse 2406.2 Ns (540.94 lb-s)

Motor Type Reload

1.2.2 Recovery System Overview

Our launch vehicle will be using a dual-deploy recovery system. The dual-deploy recovery

system will require the launch vehicle to release a drogue parachute at apogee to prevent

excessive acceleration upon main deployment. The main parachute will be deployed at a lower

altitude to ensure the minimization of kinetic energy upon landing. Two independent sections

allowing for separation via two ejection charges will hold the drogue parachute and the main

parachutes. It is critical that the nose cone air frame and the Avionics bay booster section

separate when the two ejection charges go off. The first ejection charge will separate the nose

cone and the payload section releasing the drogue parachute at apogee which will be at an

approximate height of 4,750 ft. The second ejection charge will separate the motor mount from

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the avionics bay releasing the main parachute at our estimated height of 550 ft. Since the CDR,

we have finalized the mass of our rocket as well as finalized the parachute selections, the amount

of black powder charge needed, and the descent path calculations.

1.2.3 Milestone Flysheet

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1.3 Payload Summary

Title: WALL-E

Experiment: Autonomously travel at least 10 feet away from the launch vehicle after landing and

collect 10 mL of soil.

2 Changes Since CDR Table 2a: Mechanical changes.

Change Reason

Avionics Bay sled made of acrylic instead

of NylonX.

Final sled design was easier to manufacture without 3D

printing, and 3D printing is weaker than manufacturing

from solid sheet material.

Our ballast design was updated to be

placed aft of the avionics bay on the

threaded rods

The avionics bay is a place anchored in the rocket, and

the length of threaded rods was easily increased to allow

room for the attachment of ballast.

Spin tabs will not be used on the fins. Adding spin tabs to our fins would decrease the altitude

of our rocket and is unnecessary now that our rocket is

projected to get to the target apogee with the new mass.

Table 2b: Avionics changes.

Change Reason

Selection of 433MHz ISM Band Dipole

Antenna for onboard antenna.

The selection of this antenna followed from the analysis

of the testing data outlined in section 6.1.

Addition of a 3D printed antenna mount to

avionics sled.

For safety, an antenna mount should be used to secure the

antenna during flight.

Addition of indicator LEDs on the the

PCB to communicate the status of various

essential steps in the initialization of the

the avionics.

These indicators will allow us to quickly assess and

debug problems if they arise.

Addition of a linear voltage regulator to

the PCB and removal of switching voltage

regulator from the sled.

We chose to add a linear regulator to the PCB in order to

reduce the number of discrete components on the sled.

Using two ground control transceivers

instead of one.

Separating the payload and vehicle ground control

systems has simplified the development process for each

system and its software.

The ground control will utilize a Teensy

3.6 microcontroller instead of an Arduino

We chose to switch to a Teensy 3.6 because we were able

to solder it directly onto a perf-board allowing for a more

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Uno. robust system.

Using MATLAB instead of LabView for

the ground control GUI.

By using MATLAB we were able to achieve the desired

performance capabilities while also being able to design

and create it within the necessary time constraints.

Various components relocated on avionics

sled

Relocation of the altimeters and their batteries on the

recovery side of the sled was necessitated by the

manufacturing process. Because the avionics sled was

manufactured out of acrylic, we were unable to use

threaded brass inserts.

Table 2c: Recovery changes.

Change Reason

Use of posts instead of friction locking

disconnects to connect altimeters to black

powder charges through the bulkheads.

After considering the use of friction locking disconnects,

we chose to go with the posts described in Section 3.2.3

instead. The posts are a more robust solution that

minimizes the chances of the wires being pulled out of

the altimeter or loosing connection.

Use of a rotary switch instead of a

modified snap action switch.

We considered modifying such an important safety

component to be ill advised. We used rotary switches as

our safety switches. This decision was driven by the

priority of ensuring safety before, during, and after flight.

Parachute shape changed to increase

coefficient of drag to 1.85 from 1.6.

The weight of our rocket increased slightly, so to reflect

that increase we have increased the coefficient of drag of

the parachute.

Table 2a: Payload changes.

Change Reason

All electronic components are now

mounted on a PCB

Using a PCB significantly reduced the wiring

complexity compared to a perfboard and lowered the

risk of shorts or failed connections.

Minor changes to CAD models for

manufacturability

The bulk of the rover’s mechanical elements are 3D

printed, and to optimize the models for ease of printing

and reliability, small modifications had to be made to

the frame and the actuating arm.

3D printed bulkhead will no longer be

used to support the weight of the vehicle

To address concerns about the strength of 3D printed

parts, the eyebolt that connects the main parachute

with the forward section is now attached directly to the

bulkhead on the avionics bay.

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Motor Driver changed from L298N to

MX1508

Due to a lack of space in the rover, the MX1508 driver

is much easier to integrate with the PCB compared to

the larger L298N board

3 Vehicle Criteria

3.1 Changes Since CDR

3.1.1 Mechanical

This iteration of the PRT-01 includes altered fin shapes, an avionics bay sled made of acrylic,

and a new design for ballast storage. Our subscale rocket was designed with a narrow root cord

due to its relatively small physical size. To account of the altered dimensions and weight

distribution of our full scale, we have maintained the clipped delta design of the fin but have

greatly increasing solely the root cord. The resulting fin is much wider but was proven in

simulations to position our Center of Pressure properly with respect to the new Center of Mass.

We are also not using spin tabs on our rocket’s fins because our rocket reaches the correct

apogee without the use of the spin tabs. An increase in weight of the overall rocket from the

ballast retention design, manufacturing of the rocket, and other miscellaneous weights make up

the discrepancy we were previously unable to account for.

Our sled is made of sheet acrylic instead of being printed from NylonX because the NylonX

material requires an enclosed 3D printer, which we do not have access to. Additionally, we do

not have access to a printer large enough to print the sled in one continuous piece. Printing the

sled in pieces and connecting them would decrease the overall strength of the sled so it was

determined that we would use a sheet material to make the sled. Acrylic is a strong material and

it was clear which was helpful when lining up the electronics on our sled before attaching them

on either side. This way we could confirm bolts from either side would not be interfering with

each other and electronics. Acrylic sheet is more dense than NylonX so weight was added to the

avionics bay, but our motor was chosen with the thought that weight would be added during the

manufacturing process, so this extra weight did not affect the ability of our rocket to reach our

target apogee.

Our updated ballast design focuses on the adjustable addition of ballast mass depending on

weather conditions. Ballast discs are made of 0.06” thick stainless steel at either 100 or 50 gram

mass. Clearance holes along the discs’ diameter allow them to be slid onto the threaded rods and

the secured with washers and nuts. Additionally, a large center hole is included to fit the aft

avionics bay eye bolt. In total, we are able to add up to 2.2 lbs to our rocket just aft of the

avionics bay. Open Rocket simulations have confirmed that this variation of mass still allows for

a proper stability margin. Ballast were cut using a laser cutter to ensure accuracy and a uniform

design.

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Figure 3.1.1a: Ballast secured to avionics bay.

3.1.2 Avionics

Since the completion of CDR several changes have been made to the design of the avionics

system. Most of these changes were necessitated by manufacturing limitations or other

unforeseen circumstances that did not allow us to execute the implementation of our design as

expected. The changes to each avionics subsystem are outlined below.

Table 3.1.2a: GPS subsystem.

Change Reason

Selection of 433MHz ISM Band Dipole

Antenna for onboard antenna.

The antenna was selected as it performed the best

in our range tests.

Addition of a 3D printed antenna mount. It was determined that an antenna mount should be

used to secure the antenna during flight since the

largest antenna option was chosen. The antenna

mount will prevent shaking of the antenna which

could put stress on the PCB and transceiver,

potentially resulting in damage to the system.

Change in dimensions of the PCB. The PCB dimensions were changed from 160 x 33

mm to 120.65 x 36.82 mm. These design changes

were made in order to allow both the PCB and the

main battery to be placed along the same edge of

the sled instead of diagonal from each other as

previously proposed.

Replacement of switching voltage

regulator with linear voltage regulator on

the PCB.

We chose to add a linear regulator to the PCB in

order to reduce the number of discrete components

that had to be mounted to the sled and to increase

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space. The switching regulator more effectively

filters voltage ripple and provide more resistance

to temperature increases as compared to the linear

regulator. However, testing with a linear regulator

proved that all electrical components opperated

correctly with no effect on performance.

Additionaly we have determined that because the

difference between input and output voltage of the

linear regulator is only 2.4 volts, temperature

increase is of minimal concern. This was

determination was confirmed during the battery

longevity testing as described in 6.1.

Addition of indicator LEDs to the PCB. LEDs were added to the PCB in order to indicate

the status of various essential steps in the operation

of the avionics systems. These steps include

initializing the transcievers, setting the transceiver

frequency, and sending and receiving data. These

indicators will allow us to quickly assess the status

of the system and help us debug problems if they

arise.

Table 3.1.2b: Recovery subsystem.

Change Reason

Use of ring terminal posts instead of

friction locking disconnects to connect

altimeters to black powder charges

through the bulkheads.

The two designs for securely routing charge leads

through the bulkheads were friction locking

disconnects and ring terminal posts (which was

suggested by out mentor). After considering the

use of friction locking disconnects, we chose to go

with the posts described in section 3.2.3 instead.

We made this change for two reasons. The posts

are a more robust solution that minimizes the

chances of the wires being pulled out of the

altimeter or losing connection. Despite the friction

lock disconnects being easier to manufacture and

use, preliminary tests showed low resistance to

seperation. The decision to uses the posts also

made manufacturing significantly easier by

avoiding having to manufacture a non-uniform slot

in the bulkheads for the disconnects.

Use of a rotary switch instead of a

modified snap action switch.

We considered modifying such an important safety

component to be ill advised. We used rotary

switches that must be turned with a flat head

screwdriver as our safety switches. This decision

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was driven by the priority of ensuring safety

before, during, and after flight.

In addition to changes in the avionics design, there were changes to the layout of the avionics

sled. These changes resulted from the design changes outlined above and/or from unexpected

manufacturing challenges and limitations. These changes are outlined below in Table 3.1.2c.

Table 3.1.2c: Ground Control.

Change Reason

Using two transceivers for ground control

instead of one.

We will be using two LoRa transceivers for the

ground control system, one for the gps subsystem

and one for the rover. The reason for this was to

simplify the development process each system and

its software was able to be developed

independently of the other. We will still be using a

single transceiver on the avionics sled which will

both send the GPS data and receive the rover

release command. This information has been

reflected in the University of Pittsburgh’s 2019

Transmitter Data Sheet.

The vehicle ground control and the

payload ground control will each utilize a

Teensy 3.6 microcontroller instead of an

Arduino Uno.

We chose to switch to the Teensy 3.6 because we

were able to solder it directly onto a perf-board

allowing for a more robust system. We are also

familiar with using the Teensy 3.6 and have extras

on hand since we are using it in the on-board

avionics system.

Using MATLAB instead of LabView for

the vehicle and payload ground control

GUI.

MATLAB was chosen to replace LabView as the

software used to create our GUI due to the team’s

familiarity with it. By using MATLAB we were

able to achieve the desired performance

capabilities while also being able to design and

create it within the necessary time constraints.

Changes to the GUI functionality. Various functional improvements were made to

the GUI. The GUI now shows a live map of the

airframe’s position while also tracking the altitude,

distance from the launchpad, max distance, max

altitude, and total flight time.

Added software to prevent duplicate

transmissions of rover release and rover

This software functionality was added in order to

prevent duplicate commands which could interfere

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deploy commands. with the rovers operation. This prevents the rover

release sequence from being interrupted by a new

release command.

Changes to GUI appearance and

integration of rover control into the same

interface.

The ground control GUI visuals (see Figure

3.4.3b) are different due to the improvements

made to its functionality and the fact that the rover

control has been integrated into the GPS GUI.

Having them together will allow for ease of use

and provide ground control with all pertinent

information before, during, and after launch to

ensure a successful launch and recovery of the

rocket and successful deployment of the rover.

Table 3.1.2d: Changes in avionics sled layout.

Change Reason

The PCB was relocated on the sled. Because the design of the PCB changed, it was

necessary for us to modify the location of the PCB

to allow room for the antenna to protrude from the

board without causing interference with the

avionics bay tube when removing and inserting the

sled.

The altimeters were relocated on the sled. Relocation of the altimeters on the recovery side of

the sled was necessitated by the manufacturing

process. Because the avionics sled was

manufactured out of acrylic, instead of Nylon X as

anticipated, we were unable to use threaded brass

inserts. Because of this the screws securing the

altimeters protrude from the other side of the sled.

Thus, had to relocate the altimeters so that they

would not interfere with components such as the

battery on the opposite side of the sled.

Relocation of recovery batteries. This was required because of the relocation of the

altimeters.

Addition of the antenna mount. This was added to the layout of the avionics bay

after it was deemed necessary as discussed in

Table 3.1.2a.

Removal of the 5V regulator The external 5V switching regulator was removed

from the sled. The reasons for which are discussed

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in the table 6.1.2k.

Placement of PCB on standoffs. In order to mount the PCB on the opposite side of

the avionics sled relative to the altimeters, we had

to use 8 millimeter-tall standoffs. This allowed the

Teensy to be mounted on the bottom of the PCB to

be safe from interference with the sled or the

screws attaching the altimeters to the sled. The

PCB standoffs are discussed in further detail in

section 3.4.2.

3.2 Design and Construction of Vehicle

3.2.1 Vehicle Safety Features

Body Tubes

Our rocket body will experience forces throughout the launch, flight, and landing. We

determined that the rocket will undergo maximum stress in one of the following stages: takeoff,

flight, parachute deployment, or landing. Our rocket body, a Wildman Darkstar Extreme 4, is

made out of spiral-wound G12 fiberglass with a 4” diameter. Although the material properties

for G12 fiberglass are not available, we were able to obtain the flexural, compressive, and tensile

stress of G9, G10, and G11 fiberglass, as shown below.

Table 3.2.1a: Material properties of fiberglass types.

Material Flexural Strength

(psi)

Compressive

Strength (psi)

Tensile Strength

(psi)

G9 Fiberglass 50,200-60,400 45,000-70,000 39,000

G10 Fiberglass 45,000-55,000 35,000-68,000 45,000-55,000

G11 Fiberglass 59,600-76,700 32,900-63,000 59,600-76,700

With these properties, we can make a safe estimate for the material properties of G12 fiberglass.

We can determine that the flexural strength is most likely between 45,000 and 76,700 psi; the

compressive strength is likely between 32,900 and 70,000; and the tensile strength is likely

between 39,000 and 76,700 psi.

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Additionally, the hoop, or cylindrical strength of G12 fiberglass is ~50,000 psi, according to the

manufacturer of the rocket body. Using the lower limit of these material properties for our

simulations and calculations, we can accurately determine if the G12 fiberglass body is strong

enough to withstand any force that it could experience during the run.

During takeoff, the airframe, or body, will experience a compressive force from gravity and the

motor thrust. It will also experience a negligible drag force. Our motor, a Cesaroni K555, has a

maximum thrust of 646.7 Newtons, or 145.38 lbf, according to the manufacturer. The

gravitational force on the rocket is just the mass times the acceleration, which comes out to the

weight of the rocket -- 26.9 lbs. Since these two forces are working in opposite direction, the

maximum stress the rocket will undergo during takeoff will be the sum of the max thrust and

weight, which totals 172.28 lbf.

To determine the stress on the rocket, we divide the force by the cross sectional area using the

following equation:

𝜎 =𝐹

𝐴

Our rocket has an outer diameter of 10.16 cm (r = 5.08 cm) and inner diameter of 10.01 cm (r =

5.00507). Using the following equation, we can determine the cross-sectional area of the body.

𝐴 = 𝜋(router)2-𝜋(rinner )2

The cross-sectional area is 2.376 cm2 (0.368 in2), which means that the maximum stress endured

during takeoff is 3227780.63 Pa or 468.15 psi. This is significantly less than the estimated

minimum compressive stress of G12, 32,900 psi, so the rocket will easily be able to withstand

takeoff.

During flight, the rocket will experience a drag force due to the air around it. This creates a

compressive force on the airframe. According to our simulations, the maximum drag force,

which was calculated with the maximum wind speed that we are supposed to test with, is 393.99

lbf. While this is a sizable number, the stress (F/A) only comes out to 1070 psi, still well below

the lowest estimate for compressive strength of G12, 32,900 psi.

To show that the rocket will be able to withstand drag forces during takeoff as well, we will

assume that the maximum drag force will occur when the motor is providing maximum thrust.

Although this is not necessarily true, this situation will give us a stress that is higher than any

possible stress during takeoff. The drag force and motor thrust together create a compressive

force on the body. This force will total 566.27 lbf (172.28+393.99). Divided by the cross

sectional area, the stress is 1538.78 psi, again well below the estimate of 32,900 psi strength.

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Since the rocket will be able to withstand max drag force and max motor thrust at the same time,

it will be able to stand those forces occurring at different times.

When the parachutes deploy, a tensile stress is placed on the body. The force that is created is

from the difference in drag force from the rocket and main parachute. According to our

simulations, the maximum deceleration occurs when the main parachute deploys. This

deceleration is 176.2 m/s2 (578.08 ft/s2). Assuming this deceleration acts on the entire mass of

the rocket, it comes out to a force of 15,550.35 lbf. Divided by the cross-sectional area, this

yields a stress of 42,256.39 psi. Although this is slightly above the lowest estimate for tensile

strength of G12 at 39,000 psi, it is still well below the lower limit for tensile strengths of both

G10 and G11 fiberglass. Additionally, the tensile force only acts on the rocket for an instant

when the parachute deploys, so we can confidently state that the rocket will be able to withstand

all forces exerted by both parachutes.

During landing, the rocket will experience a compressive force from the motor as well as from

impact with the ground. We have calculated, with our rocket, that the maximum landing force

will be 57.7 lbf. However, for the purpose of maximizing stress we will assume that the rocket

landing will be at the maximum permitted with respect to the regulations, which is 75 lbf. To

analyze if the airframe can handle the stress during landing, we will perform calculations for two

different scenarios; the first of which is if the rocket lands vertically, landing on the entire cross-

sectional area and creating an axial compressive force. From the manufacturer of the body, we

know that the axial compressive strength of the tube is 20,000 psi. To calculate the stress on

impact, we divide the force by the cross sectional area: 75 lbf/0.368 in2. This yields a stress of

203.804 psi, well below the limit of 20,000 psi. For the second scenario, we will assume that the

rocket does not land upright and instead lands at an angle, therefore putting more stress on a

smaller section of the cross-section. Although we can’t find the peak stress, we can find the

smallest area that the rocket can land on within the lowest estimated compressive strength of the

body, 32,900 psi. To find the area, we divide the force by the strength: 75 lbf/32,900 psi, which

comes out to 0.00228 in2. This is an area that is approximately 1.47 mm2. The rocket coming

down on an area that small is virtually impossible, so it is extremely unlikely that the rocket will

fail on landing. Overall, the rocket will be able to withstand any forces exerted on it during

takeoff, flight, parachute deployments, and landing.

In addition to in-flight forces we also conducted a drop test to ensure the body tubes could

withstand an impact up to 75 ft-lbs of kinetic energy. All sections of the body were undamaged

during the test and were deemed sufficient for our use.

For additional protection, we added ¼” foam pipe tubing to areas of the shock cord that when

taught would be in contact with the body tubes. This foam tubing alleviates the pressure applied

on the body tube by the shock cord as the parachute is released by increasing the contact surface

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area. With thinner shock cords, there is a potential that deployment will either cause the shock

cord to cut into the body tube or to be cut by the tube’s thin lip. Both scenarios are avoided by

implementing this foam tubing thus improving safety and decreasing ware on the rocket each

launch.

Bulkheads

We are using stepped CNC cut fiberglass avionics bay bulkheads purchased from our Wildman

Darkstar Extreme kit and modified it to include holes for our eyebolt and posts for charge leads.

By using posts for our charge leads instead of pulling wires through drilled holes, we were able

to keep the holes smaller to reduce the stress on the bulkheads and protect the electronics in the

avionics bay from the blast of the black powder. We chose to use the stepped bulkheads so that

the bulkheads fit perfectly into the avionics bay and provides a seal from the black powder blast.

Nose Cone

Our airframe is topped with a 4” 4-1 G12 fiberglass wound aluminum tipped Von Karman nose

cone. The shape of the nose cone deceases overall drag by inducing laminar flow along the body

of the rocket. Von Karman nose cones have very low coefficients of drag for subsonic flights,

when compared to other popular designs, which leads to improved stability throughout flight. To

further improve flight stability and safety, we thoroughly sanded the entirety of the nose cone

following painting and before the addition of clear coat to eliminate excess drag forces caused by

surface roughness. Reducing surface abnormalities on the nose cone leads to improved flight

stability and consequently improved safety.

Fins

The design of the fins will remain the same as they were in the CDR report. We have opted for

three clipped delta shaped fins, with their dimensions listed below. They are made from G10

fiberglass, because this a strong, lightweight material. Each fin will be 120 degrees apart from

each other, so they are evenly spaced around the rocket.

Table 3.2.1b: Fin dimensions.

Dimension Measurement

Root Chord 15.494 cm

Tip Chord 7.010 cm

Height 10.008 cm

Sweep Length 8.484 cm

Sweep Angle 40.4°

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Thickness 0.318 cm

With this fin shape, our rocket will have a center of pressure located at 5.8 ft from the forward-

most point. The center of mass is at 4.9 ft from the same location. This will produce a static

margin of 2.5, which will ensure that the rocket maintains stability during the flight.

After analyzing the simulations of our as built rocket, we noticed that our apogee was already

very close to the our target apogee. Because of this, we decided not to implement the angled fin

design on this launch vehicle.

Fin Performance

To mitigate the risk that fins may break during flight due to aerodynamic forces, the fin flutter

equations can be used. In order to use the fin flutter equation, the aspect ratio of the fins must be

calculated first. This is done with the following equation:

𝐴𝑅 = 𝑏2/𝐴

where b is the height of the fin and A is the planar area of the fin. To calculate the planar area,

we will use the following equation:

𝐴 = (𝐶t + Cr)/2 x h

Where Ct and Cr are the root and tip chords respectively and h is the fin height. When we

evaluate this equation, we find that the planar area for the fin is 112.580 cm2. Substituting that

into the previous equation, we find that the aspect ratio of the fin is 0.89. This is close to 1,

which means that the fins produce only slightly more drag than lift. To minimize this effect, we

will also bevel the top edge of the fins to make the surface more aerodynamic. In addition,

proper sanding will also mitigate the drag caused by the fins.

Now that the aspect ratio has been calculated, the maximum flutter felt by the fins can be

determined.

22

In the following equations, a is the speed of sound, P is pressure, G is the shear modulus of

fiberglass, E is the elastic modulus of fiberglass, v is Poisson’s Ratio, AR is aspect ratio, 𝜆is the

ratio of the tip chord/root chord, t is the fin thickness, and c is the root chord.

at T = 70℉

a = √1.4 × 1716.59 × (𝑇 + 460) = 1128.59 ft/s

𝑃 = 2116/144 × (𝑇 + 459.7/518.6)^5.256 = 16.43 lb/in^2

𝐺 = 𝐸/(2 × (1 + 𝑣))

𝐺 = 2.7 × 10^6/(2 × (1 + 0.12))= 1,205,000 psi

𝑉 = 𝑎 × √(𝐺/(1.337 × 𝐴𝑅^3 𝑃(𝜆 + 1))/(2(𝐴𝑅 + 2)(𝑡/𝑐)^3

𝑉 = 1128.59 × √1205000 ÷ (1.337 × 0.87^3 16.43(0.452 + 1))/(2(0.87 + 2)(0.125/6.1)^3

= 1899.56 ft/s = 1295.15 mph

As of our simulations, the maximum velocity is 373.568 mph which is significantly under this

velocity. This confirms that the fins will stay attached to the launch vehicle during the flight,

keeping the rocket stable during its ascent.

Attachment Hardware

The yield strength of the nuts securing the avionics bay is 35,000 psi, which is more than strong

enough to support the maximum force (286 lbf) felt on the avionics bay bulkheads that would

transfer to the nuts. The threaded rods in the avionics bay have a tensile strength of 120,000 psi

which again is significantly stronger than what is required to support 286 lbf. This calculation of

286 lbf as the maximum force felt on the avionics bay comes from multiplying the acceleration

from the

Thrust Plate

In order to alleviate the expected load produced from the motor of our rocket, we decided it

would be best to install a thrust plate in our rocket. The thrust plate we chose is composed of

Aluminum 6061 and manufactured by SC Precision. It has an outer diameter of 4” and fits on the

75mm motor mount tube of the rocket. It attaches to the tube itself via three countersunk screws.

The flanged motor retainer in turn is attached to the thrust plate via twelve screws as well.

As previously stated, its purpose is to reduce the load on any components attached on the inside

of the tube, such as centering rings and bulkheads. It allows for the entire thrust load from the

motor to go entirely on the tube itself, which eliminates many chances for parts to fail via

shearing.

23

Figure 3.2.1a: CAD model of the thrust plate.

To make sure the thrust plate can adequately withstand the expected loads, a simulation was

performed. These results can be seen below in Figure 3.2.1b. Our motor has a maximum thrust

force of 646.7N, therefore when under this load, the thrust plate should not fail. Figure 3.2.1b

shows the expected deformation under this load, whereas Figure 3.2.1c shows the von Mises

stress. The yield stress of Aluminum 6061 is 275000 kN/m2. Since the highest stress the thrust

plate will experience under the thrust of the motor is 4950 kN/m2, this part will not fail according

to the von Mises stress theory. Note, deformation in Figure 3.2.1b is not to scale.

Figure 3.2.1b: Simulation results of the thrust plate showing deformation.

24

Figure 3.2.1c: Simulation results of the thrust plate showing von Mises stress.

3.2.2 Flight Reliability Confidence

As stated in the section 3.1.2 Success Criteria in the Critical Design Review previously

submitted, our rocket and launch will be successful if we adhere to the guidelines and rules set

forth from the NASA SL 2018-2019 handbook. These rules and how we achieved compliance

can be viewed below in Table 3.2.2a:

Table 3.2.2a: Validation of guidelines.

Guideline How it was met Validation in Report

Launch vehicle (PRT-01)

will be ready for launch and

able to stay on the launch

pad for two hours and

remain functional.

The recovery and GPS

systems were tested to

ensure that they could

successfully operate after

holding on the launch pad

for two hours. Although the

rover deployment was not

successful, the altimeters

functioned as expected and

the results ensure that the

rocket will be able to safely

fly after a two hour delay.

The testing outlined in Table

6.1.2l demonstrates that the

launch vehicle will be ready

for launch and be able to stay

on the launch pad for two

hours and remain functional.

25

A direct current of 12V

without any outside

circuitry will be used to

ignite the motor when ready

to launch.

No external circuitry was

added in order to facilitate

motor ignition. The motor

will be ignited by 12V

direct current.

The absence, in this report, of

any components violating

this guideline validates the

meeting of this requirement.

The motor used will be

commercially available and

use ammonium perchlorate

composite propellant

(APCP).

Our motor, CTI PR75-2G-

W K555 Reusable Motor,

was readily available for

commercial uses. Likewise,

it indeed uses APCP.

The motor specifications

show it uses APCP and since

it was purchased from a

company it is commercially

available

The impulse will not exceed

5120Ns.

From our simulations and

the official specifications

of our motor provided by

the manufacturer, it was

determined that the impulse

will not exceed the limit of

5120Ns.

The launch vehicle summary

section of the report shows

the specifications of the

motor which state that the

impulse is below 5120Ns.

Our launch vehicle will

have a stability margin

greater than two upon rail

exit and its acceleration will

exceed 52fps.

From our calculations, it

was determined that the

stability margin of our

rocket is indeed greater

than two upon rail exit.

Similarly, the acceleration

was determined to be 53

fps which exceeds 52fps.

The mission performance

predictions section of our

report outlines the stability

margin and the rail exit

velocity.

PRT-01 will fly to its

apogee of 4750ft using a

single stage.

From both our simulations

and actual launch data, it

was determined that the

apogee our rocket will

reach is indeed 4750 ft.

We have designed the

From the OpenRocket

diagram of the fully

assembled rocket, it is clear

that our rocket has only a

single motor so it can only

have a single stage.

26

rocket to only have a single

stage.

PRT-01 will be both

recoverable and reusable.

The recovery system of our

rocket was designed such

that when the rocket fully

descends and touches

ground, we will be able to

recover and reuse it.

Likewise, the motor

selected is classified as a

‘reusable motor’ which will

allow us to relaunch it as

many times as desired.

From the strength ratings of

the materials in our rocket

and the projected forces

applied, it is clear that the

rocket will be recovered

intact and will be reusable.

PRT-01 will include no

more than four separate

sections with couplers and

shoulders at least one body

diameter in length.

PRT-01 was purchased and

built as three sections: Payload

Tube, Nose Cone, and Motor

Tube. The nose cone coupler is

greater than four inches in

length, and the Avionics bay

which acts as the coupler

between the payload and motor

tubes is greater than 4 inches in

length.

From the as-built schematics

section in this report, it is clear

that there is only 3 sections.

3.2.3 Vehicle Construction Procedure

Vehicle construction first began with the creation of a full SolidWorks model of the rocket.

Doing so improved spatial understanding of the rocket and eliminated the potential for future

conflicting designs. Moving from the CAD model we began physical construction of the rocket.

As our body tubes came straight from the factory, we thoroughly washed each section to remove

fiberglass particles and other potentially hazardous materials. The following sections outline our

manufacturing process.

27

Nose Cone

The nose cone and coupler were lightly sanded and cleaned before being epoxied together using

G5000 Rocketpoxy. Once the epoxy dried and the couper was secured, we used the fin slots in

the booster section to mark lines 120 degrees apart that would later be used for shear pin

installation. After marking positions for the shear pins, 1.5” up from the aft end of the coupler,

we epoxied in the nose cone bulkhead. A ¼” eyebolt was installed in the center of the bulkhead

to serve as an anchor point for the shock cords and main parachute. An epoxy fillet was added

around the edge of the bulkhead to add additional strength.

Figure 3.2.3a: Nose cone epoxy fillet.

Airframe

The payload section of the body is attached with shear pins to the nose cone and rivets to the

avionics bay. Housed within in it are the payload, payload bulkhead, actuating threaded rod, and

a shock cord guide. Due to alignment and removal issues with subscale shear pins, we altered

and improved our process for full scale Following a suggestion of our mentor, the below diagram

shows a traditional bolted joint versus our new method.

28

Figure 3.2.3b: Traditional bolted joint.

Figure 3.2.3c: Shear pin attachment method used.

A traditional bolted joint creates a pressure cone around the joint due to the downward force of

the bolt head and the upward force of the threads. When our shear pins sheared during ejection in

subscale, we were left with small portion of the shear pins threaded into the inside tube. These

pieces were difficult to remove and added additional, unnecessary time to our launch process.

For this application, we are not really trying to created a bolt joint, rather are looking for the

shear pins to align and support the interior tube in the exterior tube. Our method eliminates this

issues, as pins are simply able to fall out of the interior tube after shearing. The threads in the

exterior tubes hold the shear pins in place but allow pins to simply be unscrewed after being

sheared. Additionally, alignment issues are resolved as shear pins can more easily be engaged

into their respective threads.

The following process outlines the creation of shear pin holes between the nose cone and

payload section and the booster and avionics sections. First the respective tubes are lined up,

using the formerly marked 120 degree apart lines, and the taped together to avoid rotation during

drilling. Then a #50 drill bit is drilled through both sections. The two sections are then separated

and a #43 clearance hole is drilled into the interior tube. Finally, we hand threaded the exterior

tube using a 2-56 thread held into a tap vice.

29

Figure 3.2.3d: Through hole through avionics bay and booster section.

Figure 3.2.3e: Hand threading into the booster section.

After the process was completed for the first hole in a section, the sections are put back together

and the first shear pin is inserted. he inserted shear pin ensure alignment of each successive hole.

The process is then repeated until all shear pins have been installed.

Rivets between the avionics bay and payload tube provide a fixed connection point throughout

flight. We use four 5/32” rivets, spaced 90 degrees from each other. Rivets are lined up in such a

way that they are not collinear with avionics vent holes, to avoid potentially faulty altimeter

30

readings. A 3/16” hole is drilled at each location and a rivet is inserted before moving to the next

location to ensure proper alignment.

The payload bulkhead and shock cord guide are secured into the body tube using stainless steel

M4 bolts and nuts. Their position was determined during the creation of our CAD model and

were placed as such. They are bolted directly to the payload section with the rounded head of the

bolt on the exterior of the rocket. The shock cord guide, pictured below is made of 3-D printed

nylon and upon deployment keeps the shock cord from hitting our rover. The strength of the

system was confirmed during our ejection charge test, during which the rover remained

untouched.

Figure 3.2.3f: Shock cord guide.

Motor Mount

The motor mount system includes the motor tube, three centering rings, and our kevlar harness.

Using our CAD model for reference, the three centering ring locations were marked based upon

the location of the fins and rail buttons. The centering rings and motor were all sanded lightly to

increase epoxy effectiveness. We filed two slots into the forward two centering rings, wide

enough to allow the shock cord to fit underneath it.

The kevlar harness is made of a single length that is approximately two times the length of the

body tube. The cord is folded in half and at the top a loop is formed that will serve as a shock

cord anchor point. The two ends of the cord then are slid into the centering ring slots. The two

together are then epoxied to the motor mount with epoxy being applied both below and on top of

the shock cord leads. Epoxy fillets are then added to both sides of each centering ring for added

31

strength. The aft centering ring is not installed until after the fins are epoxied in. Finally, the

motor mount is epoxied into the booster section, with the addition of epoxy fillets to all

attachment points. The additional strength added at each step ultimately increases the overall

safety and likelihood of success of our rocket.

Fins

Our fins were constructed using ¼” fiberglass sheets. Fin design was finalized in our Open

Rocket simulations and then were cut on a vertical band saw. To ensure safety of teammates, all

fin edges were smoothed using a band sander Shown below are our full scale fins before being

epoxied into the frame. Their root chord is slightly enlarged to account for the distance from the

motor mount to the exterior of the booster tube.

Figure 3.2.3g: Full scale fins.

After seeing the effect of our previous failed subscale launch, we decided to attach fins directly

to the motor mount. Epoxy was first added to the fin edge that contacts the motor mount. The

fins were then slid into the premade slots in the booster section and were allowed to dry. Laser

cut fin alignment jigs, pictured below, were slid onto the rocket to ensure fins were precisely

oriented. Doing so increases flight stability and overall safety of the rocket.

32

Figure 3.2.3h: Fin alignment jigs.

Once the fin epoxy had dried, interior and exterior epoxy fillets were added to the booster tube

and motor mount. Tape on the fins was added before to avoid excess epoxy altering the fins. An

epoxy fillet on the middle centering is also visible here. These fillets greatly increase the strength

of the fin mounts and ensure that fins remain stable during flight. Additional epoxy around the

tube was removed before it was allowed to dry.

Figure 3.2.3i: Internal epoxy fillets.

33

Motor Retention and Thrust Plate

The thrust plate is attached to the aft centering ring of the motor mount via three, #10

countersunk screws. These screws provide more than enough strength to support the weight of

the thrust plate, motor, and motor retainer. The thrust plate was not epoxied to the tube itself to

ensure easy removal if necessary. The thrust plate helps distribute the force exerted on the motor

around the lip of the booster section. Doing this makes the system safer and reduces wear on the

booster section. The motor retainer, it is is attached to the thrust plate itself via twelve #4 screws.

Figure 3.2.3j: Thrust plate and motor retainer.

Avionics Bay

Fabricating the sled was the first step that we did in construction. We tried to print the sled using

NyolnX, but the 3D printer broke, prompting us to change the sled’s material to acrylic, a

material suggested by the manufacturing center at our. After the sled was fabricated, we tested

that the threaded rods fit through the holes that were printed to fit them in the sled.

Using the sled, we then outlined the positions of all the avionics onto the sled with expo marker

base on the CAD model that we had. An example of our initial outline is shown below. We soon

realized though that the components could not fit onto the sled in the CAD model’s layout

because of the screws that stick out of the other side of the bay when securing the altimeters. The

layout was redone as a result, but it raised another issue involving securing the PCB onto the

sled. We had to ensure that the threaded rods did not hit any components, which made it difficult

with the altimeter screws poking out the other side. Threaded holes were made in the sled to

34

remedy this issue. Every time the layout was changed, we placed the fiberglass tube over the sled

with the components taped on to check all components’ clearance.

Figure 3.2.3k: Initial component outlines.

RFI spray was applied on the main avionics side after all the holes for mounting components

were drilled based off the outline. Two coats of nickel spray were applied to the sled inside a

fume hood while wearing rubber safety gloves and safety glasses. After each coat, the sled was

left to dry for a minimum of 45 minutes and afterwards was tested for resistance to make sure the

coat was conductive.

The issue of wiring outside of the bulkhead was then addressed. To reduce stress on the

altimeters from the charge leads moving frequently, the posts were constructed. Initially only 8

posts were made, connection tested with a multimeter after each time. More posts were made

after when needed. Half of the posts made had a screw inserted so the posts could connect

through the bulkhead. To make an enhanced post, a 6.35mm minor diameter nut was tightened

onto the screw, and then the screw was inserted into the terminal ring of a post. After, another

6.35mm nut was tightened onto the screw until flush against the bulkhead. Resistance from the

screw is not a concern because testing revealed it to be only about 3 ohms.

Figure 3.2.3l: Post with screw.

At the same time, we stood the sled vertically on the bulkhead and marked places to drill for the

threaded rods and four holes on one bulkhead for the posts and eight holes on the other bulkhead

for the posts and the stepper motor driver leads. Each bulkhead was marked with sharpie,

identifying its purpose along with the location in which the posts/wires should be connected

through. The bulkhead on the main parachute side is labeled main and has four holes for the

stepper motor driver on it in addition to the four holes for the altimeter posts.

35

Figure 3.2.3m: Drilled bulkhead.

The altimeter presets were then set so we could easily change the settings between reasonable

altitudes while on the field. The presets are now as follows:

Figure 3.2.3n: Altimeter presets.

The altimeter switches then had 5” wire leads soldered into them:

36

Figure 3.2.3o: Wires soldered to arming switch.

The altimeters were placed onto the screw holes that were drilled for them and were screwed in

with 6.35mm nuts tightened on the opposite side to secure altimeters. The 9V batteries for the

altimeters were secured to the sled using two strips of Dual-Lock Velcro for each. The 7.4V

battery was secured in the same manner, instead using four strips of Dual-Lock (two strips for

each side). 9V adapters were then attached to the altimeter blocks. On the opposite side of the

sled, the smallest size PCB standoffs were screwed into their respective holes. The PCB was

then screwed onto the standoffs.

The subsequent connections attached to the altimeter blocks were the post leads. When

completing this step, we realized that after stripping it, the 16-gauge wire was too big to fit inside

the altimeter blocks. Since the posts are stranded core wire, a very small number of strands were

repeatedly cut until they fit into the altimeter blocks. The sled looked like this after connecting

all the posts into the altimeter blocks.

Figure 3.2.3p: Posts secured to altimeters.

The threaded rods were inserted through the holes in the sled. Rubber washers then 14.15mm

diameter nuts were tightened on one side so they were flush with the sled. The drogue bulkhead

was placed onto the threaded rods and secured with a regular, metal washer and 14.14mm nut on

37

the other side of it. The posts with screws were inserted through their properly identified holes

on the drogue bulkhead. On each post screw, a 6.35mm nut was tightened onto the screw,

followed by posts without screws and another 6.35mm nut to make the system secure.

Figure 3.2.3q: An example of some posts secured into a bulkhead.

Figure 3.2.3r: Fully connected bulkhead.

At this point in assembly we realized an issue with the threaded rods interfering with the 9V

adapters. We did not foresee, even though planning, how drastically the proximity of the

altimeters to the threaded rods would affect the wires coming out of it. The altimeters were

positioned so that we would have to wire the blocks closest to the threaded rods side before the

threaded rods would be inserted. We tested sliding the threaded rods in and out of the sled

multiple times with the 9V adapters already in the altimeter blocks. After these tests, the

insulation on the wires was starting to fade. Because of this issue, we decided to move the

altimeter over 6mm that had interference with the threaded rod later in assembly.

38

Figure 3.2.3s: Bulkhead attached to sled (before altimeter position was changed).

The arming switches were mounted into the fiberglass tube by first drilling into the fiberglass

then inserting the arming switches. The arming switches were wired to the altimeter blocks and

the sled was inserted into the fiberglass tube to test the clearance of all the components

considering the amount the arming switches stick into the bay. After taking the sled out of the

tube, one of the wires previously soldered onto the arming switch broke off. We decided to tape

the arming switch wires to the fiberglass tube and tape wires from the same origin together to

reduce the risk of tangling wires and having unnecessary stress on them. The arming switch was

re-soldered afterwards.

Figure 3.2.3t: Arming switches mounted to fiberglass tube.

The wires to connect the stepper motor driver to the stepper motor were soldered onto the motor

driver board next. After, the altimeter posts were connected through the main bulkhead as they

were for the drogue bulkhead. To secure the motor driver wires, the smaller gauge wires motor

driver wires were soldered onto post leads. These post leads were connected through the main

bulkhead. Rubber washers, then 14.15mm diameter nuts, were tightened on one side so they

were flush with the sled. The drogue bulkhead was placed onto the threaded rods and secured

with a regular, metal washer and 14.14mm nut on the other side of it.

39

At this point, the avionics bay was successfully constructed, but we felt that we could complete

this process faster. We practiced the parts of assembly numerous times, mainly securing the

bulkheads to the threaded rods and attaching the posts through the bulkheads. By launch day,

these assembly procedures should take about 10 minutes.

Figure 2.3.2u: Completed avionics bay.

3.2.4 As-Built Rocket Schematics

Figure 3.2.4a: OpenRocket model of launch vehicle.

The as built launch vehicle is a total of 25.7 lbs with full ballast. The rocket is 7.6 ft long, with a

diameter of 4’’. The center of pressure is calculated to be at 5.8 ft from the nose cone. The center

of gravity is 4.9 ft from the nose cone, giving our launch vehicle a stability margin of 2.5.

40

Figure 3.2.4b Avionics Bay Figure 3.2.4c Forward Section

Figure 3.2.4d Nose Cone Section

As-Built Schematic. As-Built Schematic. As-Built Schematic.

41

Figure 3.2.4e Booster Section

As-Built Schematic.

Figure 3.2.4.f Avionics Bay As-Built Schematic.

3.3 Recovery Subsystem

3.3.1 Dual Deploy Recovery System & Safety Features

42

In order to increase the success rate of our launch vehicle, our team decided to use a dual-deploy

recovery system. The dual-deploy recovery system will require the launch vehicle to release a

drogue parachute at apogee to prevent excessive acceleration upon main deployment. The main

parachute will be deployed at a lower altitude to ensure the minimization of kinetic energy upon

landing. Two independent sections allowing for separation via two ejection charges will hold the

drogue parachute and the main parachutes. It is critical that the nose cone air frame and the

Avionics bay booster section separate when the two ejection charges go off. The first ejection

charge will separate the motor mount from the avionics bay releasing the drogue parachute at

apogee which will be at an approximate height of 4,750 ft. The second ejection charge will

separate the nose cone and the payload section releasing the main parachute at our estimated

height of 550 ft. Since the CDR, we have finalized the mass of our rocket as well as finalized the

parachute selections, the amount of black powder charge needed, and the descent path

calculations.

The nose cone and air frame are tethered together by 1500# Kevlar shock cords to ensure

strength in attachment. In order to maximize securement, we will use a shock cord that is 3 times

the length of the entire rocket, approximately 30 ft. Both parachutes have 12x12 Nomex

parachute protectors to allow for protection against the hot gasses created by the ejection charge.

The 72” main parachute will be held in a deployment bag to ensure proper parachute release and

minimize the chance of entanglement. The main parachute has a barrel swivel attached to the end

of the suspension lines at the bridle. The swivel will be secured through a connection between a

quick link and the eyebolt on the bulkhead and help reduce the risk of entanglement. The 18”

drogue chute will have a similar connection system involving the three components listed above

as well. As a first year team, we will be purchasing Kevlar shock cords, parachutes, and recovery

system components (swivels, quick links, etc.) from outside vendors due to the difficulty of

fabricating the essential components. The sizing of our components as well as our parachute

sizes have not changed since CDR.

Using the maximum acceleration due to the main parachute and the combined mass of the

booster and payload sections of the fully loaded rocket, the max force on the shock cord and eye

bolt was calculated to be 286 lbf. The working load limit of the eyebolts is 430 lbf according to

the manufacturer, the carrying capacity of the quick link is 880 lbf, and the shock cord is rated to

be 1500 lbf, so all of the recovery attachment hardware is more than strong enough to withstand

the maximum forces it will experience during recovery.

Black Powder Ejection Charge

In order to precisely deploy the main and drogue parachutes at their given altitudes, we will use

ejection equipment consisting of black powder, shear pins, and wiring (including electronic

matches). The black powder will be enclosed in ejection canisters attached at the ends of the

avionics bay and payload section. In order to determine the mass of the black powder needed to

43

break the shear pins, we must look at four main components: pressure, volume, gas constant, and

combustion temperature. With an airframe diameter of 3.9 inches, we determined the

approximate shear force required to break the shear pins. With 4 shear pins, each pin shears at 44

lbf needing a total minimum force of 176 lbf. The additional 3.2 lbf acts as a buffer to ensure the

shear pins will in fact break off. Listed below are the Black Powder Ejection calculations using

the Ideal Gas Law. The calculations are necessary to find the correct amount of black powder

needed to deploy the parachutes.

Table 3.3.1a: Black powder ejection charge information.

𝐷𝑖𝑎𝑚𝑒𝑡𝑒𝑟 𝑜𝑓 𝑅𝑜𝑐𝑘𝑒𝑡 (𝐷) = 3.9 𝑖𝑛 𝐹𝑜𝑟𝑐𝑒 = 179.2 𝑙𝑏𝑓

𝐿𝑒𝑛𝑔𝑡ℎ𝑀𝑎𝑖𝑛 𝑃𝑎𝑟𝑎𝑐ℎ𝑢𝑡𝑒 𝐶𝑜𝑚𝑝𝑎𝑟𝑡𝑚𝑒𝑛𝑡 (𝑙)

= 18.75 𝑖𝑛

𝐵𝑙𝑎𝑐𝑘 𝑃𝑜𝑤𝑑𝑒𝑟 𝐺𝑎𝑠 𝐶𝑜𝑛𝑠𝑡𝑎𝑛𝑡 (𝑅)

= 266 𝑖𝑛 ∗ 𝑙𝑏𝑓

𝑙𝑏𝑚

𝐿𝑒𝑛𝑔𝑡ℎ𝐷𝑟𝑜𝑔𝑢𝑒 𝑃𝑎𝑟𝑎𝑐ℎ𝑢𝑡𝑒 𝐶𝑜𝑚𝑝𝑎𝑟𝑡𝑚𝑒𝑛𝑡(𝑙)

= 5.51 𝑖𝑛

𝐶𝑜𝑚𝑏𝑢𝑠𝑡𝑖𝑜𝑛 𝑇𝑒𝑚𝑝𝑒𝑟𝑎𝑡𝑢𝑟𝑒 (𝑇) = 3307∘ 𝑅

𝐴𝑟𝑒𝑎 =𝜋𝐷2

4=

𝜋 ∗ 3.92

4= 11.9 𝑖𝑛2

𝑃𝑟𝑒𝑠𝑠𝑢𝑟𝑒 (𝑝) =𝐹

𝐴=

179.2 𝑙𝑏𝑓

11.9 𝑖𝑛2= 15 𝑝𝑠𝑖

𝑉𝑜𝑙𝑢𝑚𝑒 (𝑉) = 𝐴𝑟𝑒𝑎 (𝐴) ∗ 𝐿𝑒𝑛𝑔𝑡ℎ (𝐿) =𝜋𝐷2

4∗ 𝑙

𝑝𝑉 = 𝑀𝑅𝑇

𝑀 =𝑝𝑉

𝑅𝑇

Table 3.3.1b: Minimum charge masses to separate the sections.

Length (in) Volume (𝑖𝑛3) Mass (lbs) Mass (g) per

altimeter

Main Charge 18.75 𝑖𝑛 223.125 𝑖𝑛3 0.003805 𝑙𝑏𝑠 1.73 𝑔

Drogue Charge 5.51 𝑖𝑛 65.569 𝑖𝑛3 0.001118 𝑙𝑏𝑠 0.507 𝑔

These charge masses are the minimum possible to separate the sections.

44

Figure 3.3.1a: Plan for descent (courtesy West Rocketry).

3.3.2 Parachute Choice

Parachute Location

The main parachute will be placed between the nose cone and the payload section. Between the

avionics bay and motor mount will be the location of drogue parachute. We have chosen this set

up as we expect the release of the drogue parachute to cause less stress on the payload bulkhead

than the release of the main parachute and because our mentor advised us to switch the

placement in order to minimize the chance of entanglement of the parachute cords. This will

ensure the safe deployment of both parachutes.

Main Parachute Selection

Since the CDR, we have stuck with our main parachute selection. We have chosen the

Rocketman elliptical 72” diameter parachute. After thorough research into main parachute

options, the Rocketman elliptical parachute was deemed most suitable for our purposes. The

Rocketman parachute is a multi-gore chute with a coefficient of drag of 1.85. The parachute has

45

attached suspension lines, a heavy bridle, and barrel swivel to ensure maximum strength and

security. It is brightly colored so that it will be easily visible as it falls, making it easier to track.

Drogue Parachute Selection

Following the CDR, we have finalized our decision to use the Apogee Fruity Drogue Parachute

(18” diameter model). Without a drogue parachute, there would be significant drift away from

the point of launch. In order to slow down our launch vehicle, we must have a drogue parachute

deployed at apogee so the main parachute can be deployed at the lowest elevation point possible.

In the process of selecting the drogue parachute, the three factors we are concerned with are the

material, weight, and dimensions. The Apogee Fruity Drogue Parachute (18”) is not only bright

in color, but has an elliptical shape which is considered optimal for high drag and minimum

weight and material. The typical drag coefficient is 1.5-1.6 which fits the estimated drag

coefficients tested in our simulations. The strong nylon cloth material used for the parachute and

the suspension lines that attach to a nylon bridle and a barrel swivel ensure for maximum

strength and minimization of twisting. The 18” diameter is the proper size to minimize drift

between its deployment at apogee and the deployment of the main parachute. Additionally, it

will maintain a factor of safety in the force applied to the main parachute shroud lines and shock

cords.

Numerical Analysis of Drogue Parachute

By analyzing the surface area and coefficient of drag on the respective drogue parachute, we are

able to find the terminal velocity for the drogue chute:

𝑝𝛥 = 𝐹𝑛𝑒𝑡 ∙ 𝛥𝑡

0 = 𝐹𝑔𝑟𝑎𝑣𝑖𝑡𝑦 − 𝐹𝑑𝑟𝑎𝑔

𝐹𝑔𝑟𝑎𝑣𝑖𝑡𝑦 = 𝐹𝑑𝑟𝑎𝑔

So,

𝑔 ⋅ 𝑀𝑟𝑜𝑐𝑘𝑒𝑡 =𝐶𝑑 ∙ 𝜌 ∙ 𝑉2 ∙ 𝐴

2

𝑉𝑡𝑒𝑟𝑚𝑖𝑛𝑎𝑙 = √2𝑀𝑔

𝜌𝐶𝑑𝐴

𝑉𝑡𝑒𝑟𝑚𝑖𝑛𝑎𝑙 = √(2)(11.516 𝑘𝑔)(9.81 𝑚/𝑠2)

(1.225 𝑘𝑔/𝑚3)(1.6)(0.164𝑚2)

𝑉𝑡𝑒𝑟𝑚𝑖𝑛𝑎𝑙 = 26.5 𝑚/𝑠

Where Cd is the drag coefficient, 𝑀 is the mass of the system after motor burnout, 𝑔 is the

acceleration due to gravity, 𝜌 is the air density, 𝑉 is terminal velocity, and 𝐴 is the cross sectional

area of the parachute.

46

Numerical Analysis of Main Parachute:

To find the terminal velocity achievable by solely the contributions of the main chute, we use

similar calculations to those above:

𝑝𝛥 = 𝐹𝑛𝑒𝑡 ∙ 𝛥𝑡

0 = 𝐹𝑔𝑟𝑎𝑣𝑖𝑡𝑦 − 𝐹𝑑𝑟𝑎𝑔

𝐹𝑔𝑟𝑎𝑣𝑖𝑡𝑦 = 𝐹𝑑𝑟𝑎𝑔

So,

𝑔 ⋅ 𝑀𝑟𝑜𝑐𝑘𝑒𝑡 =𝐶𝑑 ∙ 𝜌 ∙ 𝑉2 ∙ 𝐴

2

𝑉𝑡𝑒𝑟𝑚𝑖𝑛𝑎𝑙 = √2𝑀𝑔

𝜌𝐶𝑑𝐴

𝑉𝑡𝑒𝑟𝑚𝑖𝑛𝑎𝑙 = √(2)(11.516 𝑘𝑔)(9.81 𝑚/𝑠2)

(1.225 𝑘𝑔/𝑚3)(1.85)(2.63𝑚2)

𝑉𝑡𝑒𝑟𝑚𝑖𝑛𝑎𝑙 = 6.157 𝑚/𝑠

Total Landing Kinetic Energy

To calculate the total kinetic energy at landing, the system is defined as the rocket and parachute.

Since the CDR, we have calculated the total mass of our rocket resulting burnout mass of 11.516

kg. We can now calculate the total kinetic energy at landing with the following equations:

Using the terminal velocity achievable with solely the drogue parachute:

𝐾𝐸 =1

2𝑀𝑉2

𝐾𝐸 =1

2(11.516 𝑘𝑔)(26.5 𝑚/𝑠)2

𝐾𝐸 = 4044 𝐽 = 2983 𝑓𝑡 − 𝑙𝑏𝑠

Using the terminal velocity achievable with the main parachute:

𝐾𝐸 =1

2𝑀𝑉2

𝐾𝐸 =1

2(11.516 𝑘𝑔)(6.157 𝑚/𝑠)2

𝐾𝐸 = 218.3 𝐽 = 161.0 𝑓𝑡 − 𝑙𝑏𝑠

Landing Kinetic Energy Limit

47

To ensure we meet the kinetic energy requirement set by NASA, we performed calculations on

the each subsection of our rocket to test that we meet the kinetic energy requirement of a

maximum value of 75 ft-lbs (101.68 J). The burnout mass of our rocket is 8.04 kg and has been

split to illustrate the mass of each section. Below, each subsection mass will be used to test that

we fulfill the 75 ft-lb requirement per each subsection.

Table 3.3.2a: Section mass values.

Booster Mass 3.899 kg

Payload Mass 4.114 kg

Nose Cone Mass 0.663 kg

Energy calculations using the drogue parachute terminal velocity:

Booster Mass

𝐾𝐸 =1

2𝑀𝑉2

𝐾𝐸 =1

2( 3.899)(26.5 𝑚/𝑠)2

𝐾𝐸 = 1369 𝐽 = 1010 𝑓𝑡 − 𝑙𝑏𝑠

Payload Mass

𝐾𝐸 =1

2𝑀𝑉2

𝐾𝐸 =1

2( 4.114 𝑘𝑔)(26.5 𝑚/𝑠)2

𝐾𝐸 = 1445 𝐽 = 1066 𝑓𝑡 − 𝑙𝑏𝑠

Nose Cone Mass

𝐾𝐸 =1

2𝑀𝑉2

𝐾𝐸 =1

2(0.663 𝑘𝑔)(26.5 𝑚/𝑠)2

𝐾𝐸 = 233 𝐽 = 172.8 𝑓𝑡 − 𝑙𝑏𝑠

Energy calculations using the main parachute terminal velocity:

Booster Mass

𝐾𝐸 =1

2𝑀𝑉2

𝐾𝐸 =1

2( 3.899 𝑘𝑔)( 6.157 𝑚/𝑠)2

𝐾𝐸 = 73.9 𝐽 = 50.5 𝑓𝑡 − 𝑙𝑏𝑠

48

Payload Mass

𝐾𝐸 =1

2𝑀𝑉2

𝐾𝐸 =1

2(4.114 𝑘𝑔)( 6.157 𝑚/𝑠)2

𝐾𝐸 = 78.0 𝐽 = 57.5 𝑓𝑡 − 𝑙𝑏𝑠

Nose Cone Mass

𝐾𝐸 =1

2𝑀𝑉2

𝐾𝐸 =1

2(0.663 𝑘𝑔)( 6.157 𝑚/𝑠)2

𝐾𝐸 = 12.6 𝐽 = 9.3 𝑓𝑡 − 𝑙𝑏𝑠

The above calculations confirm that with a successful main parachute deployment, each section

will land with a kinetic energy less than the upper limit of 75 ft-lbs.

3.3.3 Electronic Recovery Component Robustness and Failure Modes

The electronic components of the recovery system are determined to be highly unlikely to fail

based on testing and manufacturer failure ratings. When proper procedure is followed with

installing altimeters, securing wires in their block terminals, and replacing batteries between

flights, the failure rate is minimized with the greatest apparent risk being due to the electric

matches failing. The less than 1% chance of their failure is further reduced by the design of the

redundant recovery system, with distinct altimeter setups terminating their electric matches at the

same black powder canisters. According to the (rounded up for a margin of safety) estimate of

ignition failure for the electric matches, the failure rate of the electronic recovery system should

be less than 0.01%. The failure modes of each individual component are evaluated below. All

tests are further elaborated on in Section 6.1.

49

Figure 3.3.3a: Schematic of the recovery electronics.

Safety switch flipped off during flight

The rotary safety switches are rated for high power rocket flights. They must be toggled by a flat

head screwdriver. The switch used on the subscale was in its proper position even after the initial

subscale impact, demonstrating that this failure point is highly unlikely even in the absolute

worst case scenario. Redundant altimeter setup reduces the risk of this event.

Safety switch Disconnected from altimeters

The switches perform as expected when used in altimeter tests. Their wires were soldered on

after being looped through hole for security and then covered with heat shrink for additional

protection. The switches are securely connected to altimeter block terminal. Extra slack in the

wires prevents them from being tugged out of the block terminal. The switches are epoxied to

airframe. Redundant altimeter setup reduces the risk as well.

50

Figure 3.3.3b: Wires are soldered on looped like the white one. The blue one has since been

redone.

Batteries disconnect from altimeters

The 9V batteries are attached to sled with dual lock that remained attached during the subscale

crash. The adaptors are standard snap ons for 9V batteries. Their wires securely held in altimeter

block terminals. Redundancy reduces the risk as well, so that if one altimeter remains connected

the rocket can be safely recovered.

Figure 3.3.3c: Standard 9V battery snap on adaptors used for the altimeters.

51

Batteries fail to supply enough power to ignite electric matches

Fully charged 9V batteries were tested by igniting three times as many electric matches as

necessary during a test. The altimeters will reliably be able to fire the standard amount during

flights if batteries are replaced after every flight. Redundant routing of electric matches to black

powder charges reduces this risk as well.

Posts disconnect from altimeter block terminals or e-match or charge leads

The ring terminals crimped to the wires routed to the posts fit securely around the post. A

significant impact would have to occur to cause these to fail, which would only occur due to a

failed recovery. This event would occur after the posts serve their purpose.

High resistance from bulkhead posts

The posts have had their resistance tested after their construction. They have a negligible

resistance of less than three ohms each and will not impact the altimeter’s output.

Bulkhead posts short circuit:

The posts are the only portion of the wiring between the altimeters and the black powder charges

that have exposed wire. The nearest two posts are over an inch apart, making it highly unlikely

for them to short circuit and prevent an electric match from firing. The only event in which this

may occur is after the occurence of a severe structural failure deemed improbable during flight.

Charge lead connectivity failure

Leads routed to electric matches will be protected by heat shrink and secured to the inside of the

airframe with electrical tape. The end connections will be secured to the electric matches.

Stranded wire will be used to prevent any internal breakage of the leads.

Electric match ignition failure

The electric matches that our mentor supplies us with have a failure rate of less than one percent.

Redundancy in the recovery electronics minimizes this failure rate to at most one hundredth of a

percent of either parachute failing due to the electric matches.

Altimeter connection failures

The charge leads, battery, and safety switch for each altimeter can all be securely fastened in

block terminals with a flat head screwdriver. The other ends of these connections are all fixed

and will not tug on the altimeter to cause connection loss during flight.

52

Charge activation failure

The altimeters have been tested with simulated flights at minimum ten times each. The flight

logs recorded all demonstrate that they have activated the drogue output at apogee and the main

output at 550 feet.

Figure 3.3.3d: Screenshot of a test of the altimeter.

An oscilloscope was used to observe the presence of the output in addition to the logger’s record.

They have been marked as the altimeters that went through this minimum number of tests to be

our flight cleared altimeters. Any altimeters including spares brought to launches have

experienced this amount of testing. As no altimeters have ever failed to demonstrate this output,

they have been determined as highly unlikely to fail. The altimeters were also tested multiple

times with the transmitters running at maximum operating power on the other side of the sled

and successfully created appropriate flight logs. In addition to standard flight simulations

53

performed on the altimeters, a set of flight simulations with the altimeters attached to the

avionics sled as the PCB transmits data to ground control were performed. The LoRa

transmission activity did not affect the charge outputs or logging of the StratologgerCFs.

Figure 3.3.3e: Screenshot of an altimeter test performed next to the transceivers with

shielding.

Using the avionics bay partially sealed with tape as a pressure chamber makes it more difficult to

slowly release the pressure and allow for an ideal main deployment, but the logging on these

tests is steadily maintained over the course of the flight simulation.

3.4 Avionics Systems

3.4.1 Mission-Critical Avionics Components and Processes

54

Recovery

The recovery system is comprised of two Stratologger CF altimeters, two 9 Volt batteries, two

charge leads to the drogue parachute electric match and black powder canister, two charge leads

to the main parachute electric match and black powder canister, and two rotary safety switches.

In order to achieve mission success, the rocket must be launched and recovered safely. The

drogue parachute must deploy at apogee and the main parachute must deploy at 550 feet above

the ground after apogee. The recovery electronics were designed with safety as the primary

objective. The two altimeters ignite the same charges, therefore if one fails the other one is still

able to deploy parachutes and recover the airframe safely. Each altimeter has its own power

supply independent from the other altimeter’s power supply such that if one battery were to fail,

both altimeters will not lose power and the altimeter still being powered can send the ejection

charges.

The application of RF shielding material to one side of the avionics sled between the recovery

electronics and the rest of the electronics was used to ensure that RF interference would not

cause an altimeter to malfunction. This is particularly important since the GPS system includes

an on-board transceiver operating at 433 MHz. The use of RF shielding allows both systems to

be housed in the avionics bay without risking safety or mission success.

Another concern regarding the recovery system is the potential for the altimeters to fire the black

powder charges prematurely when team members are working on the rocket. The use of rotary

safety switches mitigates this risk by only enabling the altimeters when the switch has been

turned using a flat head screwdriver. This safety measure prevents a premature firing of the

altimeters and allows the rocket to be safely handled before launch.

GPS

The GPS plays an important role in the success of the mission. It consists of a 7.4 volt LiPo

battery, and a Teensy 3.6 microcontroller, an Adafruit Ultimate GPS Breakout, a 433MHz LoRa

transceiver, and an antenna integrated onto a custom designed PCB. The GPS is responsible for

transmitting NMEA-formatted GPS data back to the ground control in order to allow for a quick

and safe recovery of the rocket after landing. The GPS allows us to know the precise location of

the rocket when the rocket is within range (approximately 1500 feet of the ground system) and to

visualize its location using a live map integrated into the ground control GUI allowing the rocket

to be easily located.

Ground Control

The ground control portion of the GPS consists of an Adafruit 433 MHz RFM96W LoRa

breakout module, a moxon rectangle antenna, and a Teensy 3.6. It also has a GUI interface

created in MATLAB to allow ground control to monitor the state of the GPS and send rover

release commands to the vehicle as shown in the figure below. In addition to allowing us to see a

55

map of the vehicle during flight, the ground control GUI provides further assurance during flight

in that the last known location of the vehicle is displayed so even when the GPS transceivers

disconnect prior to apogee, we will know the latest location of the vehicle and its general

direction relative to the launch site. Because the mapping functionality of the GUI uses the

Google Maps API, it requires an internet connection. This will be provided by a mobile wireless

hotspot. Th 4G LTE data coverage map for AT&T was checked and it was confirmed that

cellular data will be available at the launch site in Huntsville.

3.4.2 Structural Avionics Components

Each component in the avionics bay is secured to the sled in a way that maximizes flight

survivability and safety. There are three methods with which electronic components are secured

directly to the sled. The first method, used to secure the altimeters, is a through hole with 3/4

inch 4-40 screws and nuts as shown in Figure 3.4.2a. The nuts are secured using Loctite

Threadlocker to ensure that the altimeters do not come loose during flight or on landing due to

vibration or flight forces.

Figure 3.4.2a: Altimeters are secured to the sled.

The second way in which components are secured to the sled is with 8 millimeter M3 nylon

standoffs. This technique is used solely for securing the PCB to the sled. The standoffs were

attached to the sled by screwing them into a pre-drilled and tapped hole. The use of standoffs

was necessitated by the fact that the PCB has components through-hole soldered onto both sides,

therefore the board had to be elevated above the surface of the sled using standoffs as shown

below to prevent the Teensy from touching the sled. By properly securing the PCB using

standoffs we eliminated potential structural failure of the GPS, which could result in inability to

recover the vehicle.

56

Figure 3.4.2b: The PCB mounted on standoffs.

The final method of securing components to the sled is using Dual Lock. Dual Lock has been

used to secure the main battery, recovery batteries, and the newly added antenna mount. Dual

Lock was used to attach the batteries to allow us to easily remove and replace or charge them.

Dual Lock also allowed us to more easily attach the antenna to the avionics sled. We were able

to mount the PCB first then mount the antenna separately and secure its mount using Dual Lock

instead of having to fit the antenna into the mount at the same time as we mount the PCB to the

standoffs. In this way not only does Dual Lock promote a safe flight and recovery, it allows for

ease of access to the electrical components it is used to secure.

Figure 3.4.2c: Recovery battery (left) and antenna mount (right) mounted using Dual Lock.

Another factor considered in the construction of the avionics sled was how to secure the wires to

minimize the probability of them disconnecting during flight. Since the charge canisters

connected to the altimeters could move during flight, it is important to prevent this movement

from pulling on the wires. Any pulling on the wires poses a significant risk to flight success and

safety as it could cause them to disconnect from the altimeters, resulting in a failure to deploy

one or both of the parachutes.

In order to mitigate this risk, we have designed and manufactured posts that have been attached

to the bulkheads. The posts are made out of 4-40 stainless steel screws, ring terminals, and nuts

secured with Loctite Threadlocker. The ring terminal is crimped onto the end of the 16 AWG

57

charge leads on both sides of the bulkhead, one connecting to the charge canister and one

connecting to the altimeter. A nut is screwed up to the head of the screw and then the screw is

put through the ring terminal. Another nut is then screwed on to securely clamp the ring terminal.

The screw is then inserted through a hole in the bulkhead, and secured with a nut on the other

side. The second ring terminal is attached and the final nut is screwed on to secure the post to the

bulkhead and Threadlocker is applied to prevent the nut from coming loose during flight. This

setup exists for the total of eight wires connecting the charges to the two altimeters. By using

this design we have prevented tugging on the altimeters that could cause the system to fail.

Instead, the force resulting from any tugging on the wires during flight will be delivered to the

more robust posts, thus greatly increasing the probability that the connections between the

charges and the altimeters will remain intact for the duration of the flight.

Figure 3.4.2d: Posts with ring terminals and nuts connected on each side of the bulkhead

(left) and individual post (right).

In addition to preventing the charge leads from becoming disconnected, we also had to consider

the possibility of one or both of the safety switches becoming disconnected from the altimeters.

Because the rotary switches are in the “on” position during flight, if one were to disconnect it

would cause the altimeter to be disabled. To prevent this from happening we have secured the

rotary switches to the avionics bay tube with super glue, and a retaining device is screwed on

front the inside to ensure that the switch cannot be pulled through in either direction.

Additionally, it has been ensured that all electrical connections between the switches and the

altimeters are robust and structurally sound.

58

Figure 3.4.2e: Switch with leads soldered on (left) and switch secured to the avionics bay

tube (right).

3.4.3 Rocket Locating System (GPS)

The GPS for PRT-1 consists of a vehicle portion and a ground control portion. The vehicle GPS

communicates coordinates to the team through the ground control portion of the GPS. The

transceivers on the vehicle and ground control transmit at 433 MHz with a peak wattage of 100

mW. The system has been tested at a range of 400 meters and the anticipated range of the system

is approximately 600 meters, just short of the 762 meter (2500 foot) range goal the team set out

to build. Since the transceivers are able to successfully reconnect after going out of range, as

long as the vehicle stays within approximately 2000 feet of the launch site, we will be able to

successfully locate the vehicle after landing with this system.

The vehicle GPS is located on the top side of the avionics sled in the avionics bay and consists of

an Adafruit Ultimate GPS breakout module, an Adafruit 433 MHz RFM96W LoRa breakout

module, a 433 MHz ISM band dipole antenna, and a Teensy 3.6. Each component was through-

hole soldered on to a custom-designed printed circuit board (PCB) to minimize continuity and

wiring issues during manufacturing, the design of which is shown below.

Figure 3.4.3a: Final PCB design

59

The PCB is powered by a Zippy 2200 mAh 2S 25C lithium-polymer battery. The PCB also

contains a linear regulator circuit to step the 7.4 volt output from the lithium-polymer battery

down to 5 volts. While all components could operate off 3.3 volts (and the Teensy 3.6 natively

operates at 3.3 volts), through testing we discovered that the system’s range is best when the

transceiver is powered with 5 Volts, therefore the entire PCB (with the exception of the motor

driver which is also on the PCB but not part of the GPS) operates on 5 volts. On average, the

GPS module takes 30 seconds to a minute to obtain a fix and begin outputting NMEA data.

During manufacturing, we realized that the antenna sticking off the PCB was only secured by the

SMA adapter which connects it to the transceiver. To provide extra stability to the antenna, a

small antenna mount was 3D printed from PLA to secure the antenna to the avionics sled. The

antenna mount was connected to the avionics sled with Dual Lock so the mount could be

removed to easily detach the antenna from the system.

To ensure that the GPS can transmit coordinates to ground control without interference, several

measures were taken. First, the LoRa transceivers being used are hardcoded to only be able to

communicate with other identical LoRa transceivers. As an additional measure, we added 4-byte

headers to all transmissions between the vehicle and ground control. The vehicle transmissions

have a 4-byte header which ground control knows so ground control only processes

transmissions containing that header. The remaining 82 bytes in transmissions coming from the

vehicle contain NMEA data from the GPS followed by a null terminator to make parsing

transmissions easier. Similarly, ground control transmissions have a different 4-byte header

which the vehicle knows so the vehicle only processes transmissions containing that header. The

remaining 16 bytes in transmissions coming from ground control contain nothing unless a rover

release command is issued from ground control at which point these bytes contain a command

followed by a null terminator, again to make parsing transmissions easier.

The ground control portion of the GPS consists of an Adafruit 433 MHz RFM96W LoRa

breakout module, a moxon rectangle antenna, and a Teensy 3.6. It also has a GUI interface

created in MATLAB to allow ground control to monitor the state of the GPS and send rover

release commands to the vehicle as shown in the figure below. In addition to allowing us to see a

map of the vehicle during flight, the ground control GUI provides further assurance during flight

in that the last known location of the vehicle is displayed so even when the GPS transceivers

disconnect prior to apogee, we will know the latest location of the vehicle and its general

direction relative to the launch site. This alleviates concerns over not reaching the intended 2500

foot range of the GPS.

60

Figure 3.4.3b: Ground control GUI

Several tests were performed on the GPS to ensure that it performs well enough to locate the

rocket after landing. First, a simple test was performed on the GPS module to test the time it

takes for the module to get a fix. We found that it took approximately 36 seconds for the GPS

module to get a fix when obstructed by fiberglass on a moderately cloudy day. As a result of this

test, we are confident that on the day of a launch, the GPS will be able to obtain a fix within a

minute of being powered on, which is sufficient.

Additionally, since CDR, we performed tests on the various antenna options for the vehicle

transceiver. After performing 350-meter range tests with identical setups, changing only the

antenna used for the vehicle transceiver, we determined that the 433 MHz ISM band dipole

antenna offers the strongest performance which is why this is the antenna we used in our final

flight configuration. These test results can be found in section 6.1.

3.4.4 GPS Failure Modes

Any circumstance that prevents the ground control GUI from displaying the final location of the

airframe will be considered a GPS failure. A failure of this type would jeopardize mission

success and the successful and safe recovery of the airframe after landing. Because of the

complexity of the GPS system, there are a variety of failure points, which have the potential to

prevent a successful flight and recovery. Each of these failure points have been identified,

analyzed, and addressed individually in order to ensure the highest probability of the GPS

subsystem’s success and overall mission success. The potential failure points that we have

identified are listed below, along with the steps that have been taken mitigate them.

61

Loss of Power

Loss of power could be the result of a drained battery or a mechanical failure resulting in the

battery becoming disconnected from the PCB. Either scenario would result in the flight computer

being unable to transmit the final location of the airframe.

The potential for loss of power due to a drained battery will be mitigated by following the proper

pre-flight procedures outlined in Section 4.4.2. Following those procedures will ensure that the

battery is fully charged prior to the GPS system being powered and prepared for launch. In

addition to being fully charged before flight, our system power tests outlined in Table 6.1.2l have

proven that, when fully charged, our battery has sufficient charge to power the GPS and

transceiver for extended periods of time.

Loss of GPS fix

The loss of our GPS fix would most likely occur if the GPS lands with its patch antenna oriented

directly downward. Without a GPS fix upon landing, we will have no way to determine the final

position of the airframe and thus a GPS failure will occur.

Although we are unable to control the orientation of the GPS antenna upon landing, our testing

has shown that a loss of GPS fix is unlikely once a fix has been acquired. The only condition

under which our GPS has been observed to have lost its fix once acquired is the loss of power.

The GPS has also shown in the test outlined in Table 6.1.2a that it is capable of rapidly acquiring

a fix when powered on. We are therefore confident that the GPS will be able to acquire and

maintain a fix for the duration of flight and upon landing.

Failure to initialize transceiver

A failure to initialize either the on-board or ground transceiver would result in the inability to

transmit the location of the airframe. For this to occur we have determined that the transceiver

must be either malfunctioning or wired incorrectly.

We have mitigated this risk by performing rigorous testing on each of the transceivers that will

be used and by maintaining proper documentation of both ground control and flight computer

wiring diagrams. Furthermore, the use of a PCB will prevent inadvertent errors in manufacturing

which could cause improper connections resulting in transceiver failure.

In the unlikely scenario that one or more of the transceivers do fail to initialize, we will be able

to quickly identify the problem through the use of indicator LEDs. We will then take corrective

action before launch to ensure success of the GPS system.

Failure to set transceiver frequency

62

A failure to set the transceiver frequency would result in the ground control transceiver being

unable to communicate with the flight transceiver.

We deem this event to be highly unlikely as it has never occurred across all of our testing with

the transceivers. In the unlikely event that it does occur, we will be able to identify it by using

the indicator lights on the PCB and the ground control system.

Failure of transceivers to reconnect

If the transceivers are temporarily unable to communicate it is essential that they are able to

reconnect. If they cannot reconnect then the ground control will not be able to receive the final

location of the airframe. Because the possibility that the transceivers will disconnect near

apogee, it was important that we confirm the ability of the transceivers to reconnect.

The test results outlined in Table 6.1.2c show that the transceivers are capable of reconnecting.

We are confident that as long as the airframe lands within range of ground control, they will be

able to reconnect.

Airframe goes beyond transceiver range

If the airframe lands beyond the range of our transceivers, we will be unable to identify the final

location of the airframe. Our mechanical team has determined that the probability of the airframe

landing beyond this range is unlikely as outlined in Section 3.5.

Flight computer software failure

An unidentified bug in the fight software could result in a GPS system failure. To minimize this

possibility, the flight software to be used has undergone significant testing and review. The flight

software has been used to perform the transceiver and GPS tests. In addition to being thoroughly

tested, the flight computer software underwent a rigorous code review process wherein it was

critiqued and refined by members of Pitt Rocketry’s avionics and mechanical subteams. Because

of this testing and review we deem a failure of the flight software to be highly unlikely.

Ground control software failure

A failure of the ground control software would result in the inability to receive the GPS

coordinates of the airframe. Like the flight computer software, the ground control software has

been used to perform the tests outlined in Section 6.1.2. It too underwent a rigorous code review

process where it was optimized and potential bugs were resolved. Because of this we deem a

failure of the ground control software resulting in a GPS system failure to be highly unlikely.

Ground control GUI failure

A failure of the ground control GUI would be unlikely to prevent the determination of the final

location of the airframe. Because the GUI reads from the serial port, if it were to fail, the GPS

63

coordinates could be viewed from the serial monitor and then manually converted into the

appropriate form. This however would be detrimental to overall system performance and ease of

use. Thorough testing of the GUI makes us believe that it will perform as expected during the

launch and recovery of the airframe.

3.5 Mission Performance Predictions

Figure 3.5a: Open Rocket simulation predictions.

Our as built rocket with 500g of ballast is predicted to reach an apogee of 4776.9 feet at 0 mph

wind speed. This is perfectly in line with our target altitude of 4750 ft. The predicted altitudes at

5, 10, 15, and 20 mph winds are 4754 ft, 4701 ft, 4635 ft, and 4589 ft respectively. Depending on

the weather conditions of the given day, more ballast will be removed allowing our apogees to

stay within our target range. However, for all wind speeds up to 20 mph, the K-555 motor will be

sufficient to bring our vehicle to the target altitude.

The descent time of our rocket is 79 seconds from apogee to the ground. From this, the drift from

the launch pad can be calculated. For 0, 5, 10, 15, and 20 mph winds, the landing distances are

simulated to be 0, 579.3, 1158.7, 1738, and 2317.1 ft according to our simulations. Using the

calculation of (windspeed * descent time), we have found the landing distances to be 0, 579,

1158.6, 1738, and 2317.3 ft for 0, 5, 10, 15, and 20 mph winds respectively. These results match

the simulation results almost exactly. For all of the wind speeds up to 20 mph, our rocket should

land within 2500 ft of the launch site.

64

The final masses of the components of our rocket are shown in the table below.

Table 3.5a: A final breakdown of the masses of the different components of our rocket.

Component Mass (g)

Drogue Parachute 394

Main Parachute 1739

Nomex Parachute Protector (1) 119

Nomex Parachute Protector Large (2) 306

Quick Links (for all 4) 129.6

30 ft Kevlar Shock Cord (2) 120

Small Ballast (all 8) 397.9

Big Ballast (all 6) 577.2

Avionics Bay 378.6

Nose Cone 631.1

Payload Section 635.5

2 AV Bulkheads 179.6

Booster Section 2270

Misc. Fasteners (3ft kevlar for parachute, shear pins, rivets, payload

bracket and bolts) 229

AV Bay Sled 827

Rover 290

Black Powder 4

Rover Retention System 235

Motor 2759

The parachutes deploying will cause an acceleration on the rocket. When the main parachute

deploys, the maximum acceleration is 176.2 m/s2. The drogue deployment causes 1.75 m/s2 of

acceleration according to our simulations.

The center of gravity of the rocket as-built is 4.9 feet from the tip of the nose cone. The center of

pressure as-built is 5.8 feet from the tip of the nose cone. The stability margin of the rocket as-

built is 2.5. These relationships are shown in the figure below.

65

Figure 3.5b: View of rocket showing center of gravity (blue ) and center of pressure (red).

As shown in section 3.2.1, these are the calculated kinetic energy at landing for each independent

and tethered section of the rocket (namely, the booster, payload section, and nose cone).

Booster Mass

𝐾𝐸 =1

2𝑀𝑉2

𝐾𝐸 =1

2( 3.899 𝑘𝑔)( 6.157 𝑚/𝑠)2

𝐾𝐸 = 73.9 𝐽 = 50.5 𝑓𝑡 − 𝑙𝑏𝑠

Payload Mass

𝐾𝐸 =1

2𝑀𝑉2

𝐾𝐸 =1

2(4.114 𝑘𝑔)( 6.157 𝑚/𝑠)2

𝐾𝐸 = 78.0 𝐽 = 57.5 𝑓𝑡 − 𝑙𝑏𝑠

Nose Cone Mass

𝐾𝐸 =1

2𝑀𝑉2

𝐾𝐸 =1

2(0.663 𝑘𝑔)( 6.157 𝑚/𝑠)2

𝐾𝐸 = 12.6 𝐽 = 9.3 𝑓𝑡 − 𝑙𝑏𝑠

The simulated motor thrust curve is shown in the figure below.

66

Figure 3.5c: Simulated motor thrust curve.

In summation, all calculations and simulations verify that our vehicle is robust enough to

withstand the expect loads before, during, and after flight.

3.6 Vehicle Demonstration Flight

Due to inclement weather, we have been unable to launch our full scale on any of the dates listed

below in Table 3.6a. Despite being ready to fly since the February 17th, all launches within a

day’s trip have been cancelled for the past several weeks. It has been frustrating and unfortunate

for our team to have not yet see the fruit of our labor. Having completed all testing up to and

including ejection charge testing, we are ready to fly.

Table 3.6a: Attempted launches

Scheduled Launch

Date

Location Result Due To

February 17, 2019 Grove City, PA Canceled High Winds

March 2, 2019 South Charleston, OH Could Not Fly Low Cloud Ceiling

March 3, 2019 Grove City, PA Canceled Snow Accumulation

March 3, 2019 Churchill, MD Canceled Weather

67

Although we have not been able to fly, our testing has shown that we are ready. In order to show

that, we intend to launch at the next available opportunity. We have located several launches this

upcoming weekend which are outlined in Table 3.6b below. Despite the University of

Pittsburgh’s spring break beginning this Friday, our team is committed to doing whatever is

necessary to launch. With a variety of locations available, we hope one will provide suitable

weather. Our mentor has already agreed to join us at any of these locations. Should the weather

cooperate, we will provide NASA with all flight results within 48 hours of launch.

Table 3.6b: Upcoming launches

Scheduled Launch Date Organization Location

March 9, 2019 Tripoli Central Virginia Rapidan, VA

March 9, 2019 WVSOAR Rio Grande, OH

March 9, 2019 MORA Easton, MD

March 9, 2019 Tripoli Mid-Ohio South Charleston, OH

4 Safety

4.1 Safety Mission Statement

Safety is the primary responsibility for every member of the team; no single facet of the team’s

operation is of greater importance. It is each individual’s responsibility to speak up when

uncertain, to work safely, and to ensure the safety of others.

4.2 Safety Officer Identified-Responsibilities Defined

4.2.1 Safety Officer Identified-Responsibilities

The Safety Officer, Thomas Harrington, is the team’s primary resource for all safety related

issues. There is a set responsibility required of the safety officer in order to maintain the safest

environment possible for all stakeholders possibly affected by the actions or inactions of the

team. The Safety Officer Identified-Responsibilities are outlined in Table 4.2.1.

68

Table 4.2.1a: Safety Officer Identified Responsibilities

Requirement Design Feature Verification Status Req #

Each team will

use a launch and

safety checklist.

The team’s

safety officer

has created and

will

continuously

improve a launch

and safety

checklist for

launch day

operations.

This requirement

shall be verified

through

inspection.

Verified. 5.1

Each team must

identify a student

safety officer.

Thomas

Harrington is the

team’s student

safety officer.

This requirement

shall be verified

through

inspection.

Verified. 5.2

Roles and

responsibilities

of safety officer

shall be outlined.

Safety officer

and/or appointed

safety deputies

will attend all

team activities to

monitor and

emphasize safety

This requirement

shall be verified

through

inspection.

Verified. 5.3.1

Safety officer

shall monitor

design of vehicle

and payload.

Safety officer

and/or appointed

safety deputies

will periodically

review the

overall design

for safety.

This requirement

shall be verified

through

inspection.

Verified. 5.3.1.1

Safety officer

shall monitor

construction of

vehicle and

payload.

Safety officer

and/or appointed

safety deputies

will be present

during

construction and

fabrication

processes.

This requirement

shall be verified

through

inspection.

Verified. 5.3.1.2

Safety officer

shall monitor

Safety officer

and/or appointed

This requirement

shall be verified

Verified. 5.3.1.3

69

assembly of

vehicle and

payload.

safety deputies

will be present

during assembly

processes.

through

inspection.

Safety officer

shall monitor

ground testing of

vehicle and

payload.

Safety officer

and/or appointed

safety deputies

will be present

during ground

testing

processes.

This requirement

shall be verified

through

inspection.

Verified. 5.3.1.4

Safety officer

shall monitor

subscale launch

test(s).

Safety officer

and/or appointed

safety deputies

will be present

during subscale

launch.

This requirement

shall be verified

through

inspection.

Verified. 5.3.1.5

Safety officer

shall monitor full

scale launch

test(s).

Safety officer

and/or appointed

safety deputies

during full scale

launch.

This requirement

shall be verified

through

inspection.

Unverified. 5.3.1.6

Safety officer

shall monitor

team activities

on launch day.

Safety officer

and/or appointed

safety deputies

for launch day.

This requirement

shall be verified

through

inspection.

Unverified. 5.3.1.7

Safety officer

shall monitor

recovery

activities.

Safety officer

and/or appointed

safety deputies

will be present

during recovery

activities.

This requirement

shall be verified

through

inspection.

Unverified. 5.3.1.8

Safety officer

shall monitor

educational

engagement

activities.

Safety officer

and/or appointed

safety deputies

will be present

during

educational

This requirement

shall be verified

through

inspection.

Verified. 5.3.1.9

70

activities.

Safety officer

shall implement

standard work

procedures

developed by the

team for

construction,

assembly,

launch, and

recovery

activities.

Safety officer

and/or appointed

safety deputies

will strictly

implement

standard work

procedures for

construction,

assembly,

launch, and

recovery

activities.

This requirement

shall be verified

through

inspection.

Verified. 5.3.2

4.3 Safety and Environment (Vehicle and Payload)

4.3.1 Approach to Analysis of Failure Modes

After deliberation and further research of Failure Modes and Effects Analyses, the decision to

redact the previous risk assessment code (RAC) matrix outlined in the team’s CDR report was

made. This decision was made primarily for the creation of a “Risk Priority Number”(RPN) for

the team’s benefit to help them understand and plan for the most debilitating risks. The new

grading system for risks includes severity, occurrence probability, and detection. The severity

scale is outlined in The RPN is calculated as (Severity*Occurrence*Detection). The highest RPN

is our most debilitating risk. Table 4.3.1.1 outlines the severity of the risk, Table 4.3.1.2 outlines

the occurrence probability of the risk, and Table 4.3.1.3 outlines the detection probability of the

risk.

Table 4.3.1a: Severity Scale

Effect Criteria: Severity of Effect Rankin

g

Casualty Equipment

Damage

Project Plan Environmenta

l

Catastrophic Permanent

disability or

death

Irreparable

damage or loss

of system,

machinery, or

equipment

Heavy delays

or budget

overruns

resulting in

failure to

complete

project

Long-term or

irreversible

damage (>5

years)

5

71

Critical Severe injury

or illness

temporarily

preventing

normal

activities

Significant

damage to

system,

machinery, or

equipment

resulting in

prolonged non-

operation

Significant

delays of

budget

overruns that

severely limit

project

performance

Medium term

(1-5 years)

4

Marginal Injury or

illness without

effect on daily

activities

Damage to

system,

machinery, or

equipment

resulting in

temporary non-

operation

Delays or

budget

overruns that

impact non-

critical project

performance

factors

Short-term (<1

year)

3

Negligible Insignificant

injury treated

through basic

first aid

Minor damage

to system,

machinery, or

equipment

fixed by

immediate

repairs

Minor delays in

non-essential

operations

Minor damage

readily

repaired.

2

None No Effect 1

Table 4.3.1b: Occurrence Scale

Probability of

Failure

Qualitative

Definition

Quantitative Definition Rankin

g

Very High Likely to occur

repeatedly during the

life of the project

Probability >10-1 5

High Likely to occur

several times during

the life of the project

10-1 > Probability >10-2 4

Moderate Likely to occur

sometime within the

life of the project

10-2 > Probability >10-3 3

Low Not likely to occur

sometime within the

10-3 > Probability >10-6 2

72

life of the project

Very Low Occurrence is not

expected during the

life of the project

10-6 > Probability 1

Table 4.3.1c: Detection Scale

Detection Criteria: Likelihood the existence of a defect will be

detected by process controls

Rankin

g

Almost Impossible No known controls available to detect failure mode 5

Low Low likelihood current controls will detect failure mode 4

Moderate Moderate likelihood current controls will detect failure mode 3

High High likelihood current controls will detect failure mode 2

Almost Certain Almost certain current controls will detect failure mode 1

4.3.2 Personnel Hazards Analysis

The Personnel Hazards Analysis, outlined in Table 4.3.2.1, outlines possible hazards to any

individuals involved with the Pitt Rocketry team’s actions and/or inactions. Here, the “Risk

Priority Number” (RPN) allows the team to effectively and quickly identify the highest risk to

personnel present. According to the Personnel Hazards analysis the biggest risk to the launch

personnel is the failure of the parachutes to deploy. It has an RPN of 80/125.

Table 4.3.2a: Personnel Hazards Analysis

Potential

Failure Mode

Potential

Failure Effects

S

E

V

E

R

I

T

Y

(1 - 5)

Potential Causes O

C

C

U

R

R

E

N

C

E

(1 - 5)

Current

Controls

Verification

Plan

D

E

T

E

C

T

I

O

N

(1 - 5)

R

P

N

(1-125)

Rocket motor

explodes

Death or serious

injury, total loss

of launch

5 Manufacturing

flaws, careless

handling, improper

3 Follow launch

pad rules for

safe distance.

Buy motor

from reputable

dealer, ship

5 75

73

vehicle installation, directly to

NAR mentor,

have NAR

mentor oversee

installation of

motor.

Injury during

ground testing

Death or serious

injury

5 Improper design of

tests

3 Design tests to

remove human

factor as much

as possible.

Safety officer

signs off on

tests

2 30

Injury due to

negligence

Death or serious

injury

5 Failure of safety

officer to fulfill

responsibilities

3 Safety officer

will not be

allowed to

participate in

any of the pre-

launch

procedures. It

will be their sole

responsibility to

oversee each

member’s

actions and

monitor for

safety.

All pre-launch

procedures will

be assigned to

specific

individuals

before arrival

at launch site.

None of the

pre-launch

procedures will

be assigned to

the safety

officer.

1 15

Inhalation of

fiberglass dust,

epoxy fumes,

paint fumes, or

other

dangerous

fumes during

construction of

launch vehicle

Injury or illness

temporarily

impairing

individual’s

daily activities

4 Lack of proper

PPE

2 Team is supplied

with dusk masks

in order to

prevent this type

of injury

Safety officer

will oversee all

manufacturing

operations

1 8

Aluminum or

fiberglass

splinter from

handling

airframe

during

Insignificant

injury treated

through basic

first aid

1 Burrs on freshly

manufactured

materials

3 Manufacturing

team is

instructed on

proper

deburring

techniques.

Manufacturin

g lead

engineer will

inspect all

components to

make sure they

1 3

74

construction Team is also

supplied with

proper

deburring tools

are properly

deburred

Injury from

machinery

Bodily harm to

team member(s);

Damage to

machine and/or

rocket

5 Mishandling of

machines, fatigue,

failure to comply

with safety

guidelines, or

machine

malfunction.

3 Safety guidelines

and instructions

will be carefully

observed with all

machines used.

Certification

tests for SCPI

and the SSoE

Makerspace

must be passed

before using the

facilities.

Access to

machinery will

be limited to

those with

certification;

Powered

machinery will

be operated

only while

another

certified user

is present.

2 30

Chemical

Injury

Bodily harm to

team member(s);

Damage to

rocket or rocket

parts.

5 Mishandling of

chemicals and/or

failure to comply

with MSDS

guidelines and

warnings

3 Proper safety

equipment will

be worn by all

team members

working with

any chemical.

MSDS

guidelines will

be carefully

observed with all

chemicals used.

Chemicals will

only be

available to

members who

have read and

understand the

MSDS

guidelines;

Chemicals will

only be used

when there is

at least one

other member

present.

2 30

Injury from

erratic rocket

flight

Injury to team

members and/or

bystanders

resulting from

rocket impact.

5 Poor rocket

stability;

Rail system

malfunction.

3 Accurately

calculate the

center of

pressure and

center of mass;

Calculations

and

simulations

will be

performed

multiple times

and reviewed

2 30

75

Perform

simulations

before fight.

by team

advisor

Matthew

Barry; A

spotter will be

assigned to

watch each

launch and

alert others if

another rocket

is dangerous

Premature

rocket ignition

Injury,

including severe

burns, to team

members and/or

bystanders.

5 Ignition

malfunction;

Failure to follow

safety procedures.

1 Conduct

briefings at pre

launch

meetings;

Ensure

reliability of

ignition safety

switch.

Launch day

operations will

be supervised

by safety

officer Thomas

Harrington

and NAR

mentor.

5 25

Premature

black powder

ignition in

ejection

charges

Injury to team

members form

explosion and

resulting

shrapnel.

5 Recovery system

malfunction;

Faulty testing

procedures;

Failure to follow

safety procedures.

2 Perform black

powder tests

within a testing

enclosure;

Follow launch

day safety

procedures.

A safety officer

or mentor will

be present for

all black

powder tests;

Launch day

operations will

be supervised

by safety

officer Thomas

Harrington

and NAR

mentor.

2 20

Parachute

deployment

fails

Damage to

rocket and/or

injury to team

members on the

ground from

5 Recovery system

fails to deploy.

4 There will be

ground and

subscale testing

of the entire

recovery system

The recovery

system will be

declared as

functional

before test

4 80

76

free-falling

projectiles.

and its

components.

Test launches

will be carried

out on days with

optimal weather

conditions and

minimal clouds

below projected

apogee.

All persons

present during

launch will be

notified to

remain attentive.

launch. All

members at the

launch will be

reminded to

stay attentive

during the

entirety of the

flight, from

launch to

landing.

Separation of

sections does

not occur even

though the

charges go off.

Death or serious

injury from

ballistic launch

vehicle

components.

5 Improper launch

vehicle

construction

2 Pre-launch

procedure

carefully

outlines launch

vehicle build

Safety officer

will oversee

vehicle build

1 10

Premature

separation of

vehicle

sections occurs

during flight

Death or serious

injury from

stray launch

vehicle

components

5 Altimeter failure 1 Altimeter test

verify that it will

deploy the

ejection charges

at the right time.

Ejection

charge test

verifies that

the ejection

charges deploy

as planned

1 5

Soldering tools

burn team

members

Insignificant

injury treated

through basic

first aid

2 Team member not

trained on how to

use soldering tools

2 Avionics lead

teaches each

member of their

subteam how to

solder properly.

Safety officer

verifies that

soldering

training has

taken place.

2 8

Injury from

electric shock

while members

handle

electrical

Insignificant

injury treated

through basic

first aid

2 Team member not

trained on how to

properly handle

electronics

2 Avionics lead

teaches each

member of their

subteam how to

handle

Safety officer

verifies that

electronics

handling

training has

2 8

77

components electronics

components

properly.

taken place.

Injury from

finger

pinching from

electric motor

Insignificant

injury treated

through basic

first aid

2 Team member’s

hands not clear of

machinery when

motor operations

are initiated

2 Mechanical lead

and payload lead

teaches their

subteam

members how to

avoid pinching

from the motors

Safety officer

verifies that

this training

has taken

place

2 8

Battery fire or

explosion

Heat and/or

chemical burns

to team

members,

damage to

rocket and other

equipment.

5 Overcharge, over-

discharge,

overheating,

puncture, or

physical impact to

lithium cells

2 Care will be

taken when

charging,

discharging,

handling, and

storing lithium

batteries.

Batteries will not

be overcharged

or over-

discharged.

Batteries will

be stored away

from

flammable

materials.

Voltage

monitoring

will occur

before and

after charging,

and every

flight to ensure

nominal

function.

Personnel

working with

lithium

batteries will

check for

shorts prior to

powering

circuit.

Batteries will

be kept away

from sources

of heat such as

4 40

78

soldering

irons.

Batteries will

be routinely

checked for

physical

damage and

swelling.

Batteries will

only be

charged under

supervision

and with a

proper Lipo

charger.

Batteries in the

rocket will be

secured-down

and located

away from

potential

points of

impact.

PM2.5 emitted

during

production of

parts

PM2.5 can

penetrate deeply

into the lungs

and affect

respiratory

systems

3 Fumes of material

created during

laser cutting might

contain PM2.5

2 Turn on the

ventilator and

make sure

people does not

work alone

Use materials

that creates

less PM2.5

and make sure

only trained

members have

access to

manufacturing

equipment.

2 12

79

4.3.3 Failure Modes and Effects Analysis

The Failure Modes and Effects Analysis (FMEA) outlines possible hazards to the launch vehicle

and rover. Here, the “Risk Priority Number”(RPN) allows the team to effectively and quickly

identify the highest risk to the launch vehicle and rover present. According to the FMEA the

biggest risks to the launch vehicle and rover are the centering rings failure to secure the motor

casing and altimeter failure. They both have an RPN of 75/125.

Table 4.3.3a: Failure Modes and Effects Analysis (FMEA)

Process

Step/Input

Potential

Failure Mode

Potential

Failure

Effects

S

E

V

E

R

I

T

Y

(1 - 5)

Potential

Causes

O

C

C

U

R

R

E

N

C

E

(1 - 5)

Mitigation Verificatio

n Plan

D

E

T

E

C

T

I

O

N

(1 - 5)

R

P

N

(1-125)

Airframe Rail button

breaks during

launch.

Trajectory

alteration

causing

danger to

anything or

anyone on

the ground

level.

5 Improper

material

selection for

rail buttons;

insufficient

fastening

method for

rail buttons;

ambitious fin

angle.

2 Rail buttons

will be tested

for security

and integrity

before launch

Visual and

mechanical

inspection

will verify

the security

and

integrity of

the rail

buttons

5 50

Rocket motor

explodes

Vehicle is

destroyed

resulting in

mission

failure and

explosion

could cause

injury or

death.

5 Manufacturin

g anomaly;

damage upon

and/or during

delivery

and/or

transportation

.

2 Motor is

purchased

from reliable

vendor. Motor

and casing is

shipped Also

consult NAR

advisor Duane

Wilkey on the

reliability of

the rocket

Test the

viability for

use of the

motor.

1 10

80

motor.

Centering

rings fail to

secure the

motor casing

Motor

travels

through the

vehicle

instead of

propelling

the vehicle

upward

causing

destruction

of the

airframe,

avionics

bay, and

rover

5 The shear

stress from

motor ignition

is too great

for the

centering

rings or epoxy

joint failure,

causing

mechanical

failure on the

airframe

during

launch.

2 The centering

rings are

attached

securely to the

airframe using

G5000

Rocketpoxy

with

sufficiently

strong fillets.

Both

subscale

and full-

scale

launch

confirmed

that the

mounting

system

sufficiently

secures the

motor to

the

airframe

for launch

5 75

Shear pins fail

to shear

Parachutes

fail to

deploy,

launch

vehicle goes

ballistic.

5 Improper

installation of

shear pins,

ejection

charges fail to

separate

launch

vehicle

3 Continuity

tests in the

wires

supporting

ejection

charges.

Careful outline

of shear pin

installation

procedure.

Multiple

ejection

charge tests

with fully-

built rocket

1 15

Fin

Detachment

Fin falls

from vehicle

mid-flight

causing

potential

injury or

death; loss

of fin causes

vehicle to

lose

stability.

5 Structural

weakness at

fin

attachment

point.

2 Take special

precaution

when attaching

fins to

airframe to

ensure quality

of attachment.

Perform

structural

tests on fins

to ensure

proper

attachment

2 30

Recovery

Systems

Premature

parachute

Parachutes

deploy prior

4 Altimeters not

properly

2 Verify and test

altimeter

Test the

recovery

5 40

81

deployment to apogee

causing the

vehicle to

lose

stability,

rendering its

flight path

off-nominal

and

unpredictabl

e.

calibrated;

faulty

deployment

logic;

improper

packing.

calibration

procedure;

verify and test

deployment

logic.

system

during

subscale

launch.

Parachutes fail

to deploy

Launch

vehicle goes

ballistic

5 Altimeters not

properly

calibrated;

faulty

deployment

logic;

improper

packing.

2 Verify and test

altimeter

calibration

procedure;

verify and test

deployment

logic.

Test the

recovery

system

during

subscale

launch.

5 40

The recovery

parachutes are

burned by the

separation

charges.

Launch

vehicle goes

ballistic

4 Nomex

blankets fail

to protect

parachute

from thermal

damage

2 Carefully

install the

nomex

blankets to

properly shield

the parachutes.

Test the

recovery

system

during

subscale

launch and

with a

ground test

of full-scale

5 40

Kevlar shock

cord snaps

Vehicle

separates in

flight;

certain

components

have no

parachute

or safety

mechanism,

turning

them into

potentially

5 Excessive

force on

shock cord

during

parachute

deployment.

2 Appropriate

calculations

simulation,

and testing can

demonstrate

that our choice

of shock cord

significantly

reduces the

likelihood of

this event.

Develop

tests and

simulations

to ensure

that the risk

of this

occurring

is at least

extremely

unlikely,

but

preferably

3 30

82

deadly

projectiles

entirely

mitigated.

Ejection

charges do not

fire

Launch

vehicle goes

ballistic

5 Lack of

continuity in

wires to

ejection

charges

3 Use stranded

wire to connect

ejection

charges to

avionics .

Continuity

tests using

multimeter

1 15

Ejection

Charges fire

prematurely

Launch

vehicle

separates

before

Apogee,.

4 Improper

altimeter

programming.

Wire shorting

2 Use PCB to

reduce the

number of

loose wires.

Perform

ejection

charge tests

2 16

Drogue chute

does not deploy

at Apogee

Launch

vehicle

speed too

great when

main chute

deploys

4 Ejection

charge misfire

3 Careful

packing of

parachute and

setup of

ejection

charges.

Perform

ejection

charge tests

Avionics Altimeter

failure

Recovery

system is not

deployed

resulting in

an

uncontrolled

free fall

possibly

leading to

injury or

death.

5 Loss of power

to altimeter;

hardware

malfunction.

2 Perform

rigorous

testing of

recovery

system

altimeter;

purchase

reliable

altimeter.

Test the

recovery

system and

altimeter

during the

subscale

launch.

5 75

GPS failure Difficult or

impossible

to locate the

vehicle after

landing.

2 Loss of power

to GPS;

Satellite Fix

failure;

Transceiver

signal loss;

2 Perform

rigorous

testing of GPS

system;

purchase

reliable GPS.

Test the

GPS system

prior to and

during the

subscale

launch.

2 12

Transceiver

failure

Failure to

receive GPS

3 Transceiver

exceeds range

2 Test

transceivers at

Test

transceivers

2 18

83

coordinates

and

therefore

failure to

locate the

vehicle after

landing.

and is unable

to operate

upon

returning to

range;

multiple

ranges within

and exceeding

maximum

expected

range.

independen

t of system

and during

subscale

launch.

Battery

thermal

runaway

Battery

explodes,

catches on

fire,

damages

internals.

5 Any damage

to battery

including

dents,

punctures,

and heat/cold.

2 Rigorous

inspection of

batteries prior

to installation;

removal of

batteries until

final assembly.

A pre-flight

checklist

will include

the battery

installation

step.

Battery to

be tested

before

installation.

4 50

Power failure Launch

vehicle goes

ballistic

5 Battery

connection

issue.

3 Using dual-

lock to secure

batteries in

place and

proper

electrical

connectors.

Listen for

altimeter

beeps

before

leaving

launchpad

2 30

Payload Payload

Electrical

failure

Failure or

loss of

power to one

or more

onboard

electrical

component(s

).

2 Insufficient

battery

voltage;

voltage ripple

from 5v

switching

voltage-

regulator

4 Fully charge

all batteries

prior to

launch;

Choose

batteries with

capacities

suited for

devices’ power

draw and

expected run

time; Choose a

switching

voltage-

regulator

designed for

Measure

power draw

of all

components

during

typical use.

Endurance

test

switching

regulator

with

correspondi

ng

electronic

devices.

5 40

84

use with

sensitive

electronics.

Rover

mechanical

failure once

deployed

Rover

unable to

complete

mission

1 Design flaws,

damage from

flight, damage

from

deployment

2 Design rover to

withstand

forces from

flight and

deployment.

Perform

ejection

charge tests

and rover

operations

tests.

1 2

Payload power

failure

Rover

unable to

complete

mission

1 Battery

connection

issue.

3 Using dual-

lock to secure

batteries in

place and

proper

electrical

connectors.

Perform

rover tests

1 3

Payload

Integration

Premature

rover release.

Recovery

does not

work; rover

is not able to

perform as

expected.

3 Failing

avionics;

launch

failure;

separation

failure.

1 Ensure that

rover release

mechanism

adheres to

rules outlined

in Student

Launch

Handbook.

Test the

rover

release

mechanism

and the

altimeters

before

flight.

5 15

Coupler tube

does not

separate at the

correct time.

Parachutes

to do not

deploy,

rocket hits

ground at

terminal

velocity.

4 Improper

fastening

techniques

(cross

threading,

under-

torqued, etc);

improper

shear-pin-

selection.

3 The rocket will

be assembled

and the shear

pins will be

tested and

sheared by

hand to

validate their

strength

Assembly

checklists

will include

proper

fastening

techniques

for coupler

tube

5 60

Coupler tube

separates

during flight

(not at the

Flight does

not reach

target

altitude;

5 Improper

fastening

techniques

(cross

1 The rocket will

be assembled

and the shear

pins will be

Assembly

checklists

will include

proper

5 25

85

correct time). endangers

spectators if

the rocket is

still in

powered

flight.

threading,

under-

torqued, etc);

improper

shear pin

selection.

tested and

sheared by

hand to

validate their

strength

fastening

techniques

for coupler

tube

Launch

Support

Equipment

Launch rail

breaks.

Trajectory

alteration

causing

danger to

anything on

the ground

level.

5 Previous

damage to

launch rail;

launch rail

improperly

assembled.

1 Consult NAR

advisor Duane

Wilkey on the

reliability of

the launch rail.

Consult

NAR

advisor

Duane

Wilkey.

5 25

Launch

Operations

Igniter fails to

ignite.

Launch

failure

1 Manufacturin

g anomaly;

damage.

2 Careful

inspection and

handling of

igniter.

Consult

NAR

advisor

Duane

Wilkey.

1 2

Propellant fails

to ignite.

Launch

failure

1 Manufacturin

g anomaly;

damage.

2 Careful

inspection and

handling of

motor.

Consult

NAR

advisor

Duane

Wilkey.

1 2

Non-uniform

propellant

burn.

Launch

failure;

target

altitude

missed;

ground

danger.

5 Manufacturin

g anomaly;

damage.

1 Careful

inspection and

handling of

motor.

Consult

NAR

advisor

Duane

Wilkey.

5 5

4.3.4 Environmental Hazards Analysis

The Hazards to Environment Analysis outlines possible hazards to the environment due to the

Pitt Rocketry Team’s actions and/or inactions. Here, the “Risk Priority Number”(RPN) allows

the team to effectively and quickly identify the highest risk to the launch to the environment

present. According to the Hazards to Environment Analysis, the biggest risks to the environment

is fin detachment. It has an RPN of 75/125.

86

Table 4.3.4.1: Hazards to Environment Analysis

Potential Failure

Mode

Potential Failure

Effects

S

E

V

E

R

I

T

Y

(1 - 5)

Potential

Causes

O

C

C

U

R

R

E

N

C

E

(1 - 5)

Current

Controls

Verification

Plan

D

E

T

E

C

T

I

O

N

(1 - 5)

R

P

N

(1-125)

Water pollution with

perchlorate

Perchlorate

inhibits NIS-

Sodium Iodide

symporters in

thyroid, which

NIS is essential

for Iodine

transport, which is

needed for

synthesizing

T3(thyroxine) and

T4

(triiodothyronine)

hormones

4 Perchlorate

from the

ammonium

perchlorate

composite

propellants

release to air

2 Dispose of all

spent motors

properly;

Safety officer

and team

mentor will

confirm the

proper

handling of

all spent

motors.

5 40

Parts of rover break off

from system

Pieces of rover are

littered into the

environment.

2 Structural

weakness in

rover due to

poor design

choices or

fabrication

errors.

2 Verify quality

of rover

assembly and

minimize

number of

separate, small

(potentially

detachable)

components on

rover.

Inspect rover

before and

after each

flight to

ensure no

components

were lost.

2 8

Parts from rocket

become detached and

become projectiles

Debris might be

left in the

environment.

5 Parts

experience

impact force

from moment

of vehicle

acceleration

3 Pay close

attention to

potential

detachment

points during

fabrication.

Thoroughly

test potential

detachment

points

4 75

87

until vehicle

is stable after

landing.

Battery rupture that

spreads hazardous

chemicals

Hazardous

chemicals are

spread into the

environment.

5 Battery is

punctured by

a hard crash

or by general

mishandling.

3 Protect the

batteries and

locate them

away from

potential points

of impact on

launch vehicle.

Properly

recycle any

non-functional

or damaged

batteries

Test battery

enclosure to

ensure it is

sufficient to

protect

batteries

from a hard

crash. Team

mentor and

Safety

Officer will

oversee

proper

disposal of

damaged

batteries.

2 30

High temperature

exhaust damages

grounds around

launching area

Ground of the

launch site could

be damaged or

even catch fire.

3 The motor

expels high

temperature

exhaust at the

ground

during

launching.

4 Ensure the

ground around

the launch pad

has no

flammable

material

surrounding it.

Examine the

ground

condition

before and

after every

launch.

3 36

Accidental black

powder ignition before

or after launch.

Brushfire, or

property fire

3 Sparks

created near

black powder

1 Fire

extinguisher to

be present at

every launch

Safety officer

will verify the

presence of

fire

extinguisher

1 3

Team members leave

trash at the launch site

after launch, such as

cardboard, plastic,

spent motors, etc.

Pollution of the

environment

contaminates

ecosystem

3 Carelessness 5 Culture of

respect fostered

among the

team. Plenty of

trash bags

packed for

launch day

Safety officer

will check for

left-behind

garbage after

team has

packed up

after launch

1 15

88

Launch vehicle

impacts the ground

with high kinetic

energy.

Hole needs to be

dug in ground in

order to recover

rocket.

Destruction of

property,

environment, or

individuals.

5 Rocket goes

ballistic

5 Continuity tests

on the ejection

charge wires.

Ejection

charge test of

every launch

vehicle setup

2 50

Additionally, the Hazards from Environment Analysis outlines possible hazards from the

environment to the Pitt Rocketry Team’s project. Here, the “Risk Priority Number” (RPN)

allows the team to effectively and quickly identify the highest risk to the launch from the

environment present. According to the Hazards from the Environment Analysis, the biggest risk

from the environment is Flight becomes unstable due to high-speed wind. It has an RPN of

30/125.

Table 4.3.4.2: Hazards from Environment Analysis

Potential Failure Mode Potential Failure

Effects

S

E

V

E

R

I

T

Y

(1 - 5)

Potential

Causes

O

C

C

U

R

R

E

N

C

E

(1 - 5)

Current

Controls

Verification

Plan

D

E

T

E

C

T

I

O

N

(1 -

5)

R

P

N

(1-125)

Flight becomes

unstable due to high-

speed wind.

Rocket fails to

reach apogee

and parachutes

do not deploy.

5 Unpredicted

high-speed

winds occur

during flight.

2 Fins mitigate

instability.

Fly the

rocket.

3 30

Launch vehicle

becomes damaged by

landing in water.

Components of

the vehicle

(particularly

electronics)

experience

permanent water

damage.

4 Launch site has

water elements

present within

2500 feet of

launch site.

2 Silicone

waterproofing

spray was

applied to the

most

vulnerable

components of

the rocket

(everything in

Perform

waterproofin

g test on

avionics bay

and verify

avionics

functionality

after test.

1 8

89

the avionics

bay).

High winds affect the

altimeters’ pressure

readings.

Altimeters deploy

the parachutes at

the wrong time

(too early or too

late) based on

erroneous

pressure

readings.

5 Unpredicted

high-speed

winds occur

during flight.

1 Minimize size

of altimeter

vent holes to

prevent excess

wind from

entering the

avionics bay.

Exert

maximum

expected

wind on vent

holes and

verify that

altimeter

pressure

reading are

accurate.

3 15

Dirt or other debris are

caught in the rover

(outside intended soil

collection window).

The rover is

unable to

operate.

2 Large clumps of

soil obstructing

rover exit

location.

3 Motor driver

pushing motor

out of rocket

will attempt to

push the rover

past obstacles.

Perform

rover

deployment

with

obstructions

over rover

egress point.

2 12

4.4 Launch Operations Procedures

4.4.1 Pre-launch day Altimeter Preparation

Before launch day, the altimeters and unused 9V batteries will be mounted on the sled with leads

due towards the posts installed in block terminals. The switches and their leads will be set in the

rocket. Snap on attachments for batteries will also be present.

90

Figure 4.4.1a: altimeter wiring diagram

The first step to prepare the altimeters for launch is to use a flat head screwdriver to set the safety

switches to the on position marked “220V” and their leads are to be set into the block terminals.

Next, the batteries will be plugged in. If the altimeters fail to begin sounding their starting

sequence of beeps, the first method for troubleshooting is to replace the battery. If an altimeter

then fails to begin sounding, remove the switch and perform a connection test across the

terminals when it is in the on position. If this fails, replace the wires. If this passes, retry

plugging the battery and switch wires into the block terminals. If the altimeter does not begin

sounding, replace the altimeter with a spare. Once the altimeters begin sounding their starting

sequences, attach the charge leads to the electric matches. Check that at the end of the starting

sequence the altimeters begin a repeated sequence of three beeps. If the altimeter stops making

noise after the full starting sequence, check that the leads are properly wired. If the altimeter

makes a repeating set of two beeps, one of the leads is not properly wired. Perform a connection

test to find the improperly wired lead. When both altimeters begin sounding a repeated set of

three beeps, turn the safety switches off until the rocket is mounted on the pad and ready to

launch.

91

4.4.2 Recovery Preparation Checklist

Parachute Setup Procedure

Required PPE: Safety Glasses

❏ Designated Mechanical member will carry out parachute setup procedure

❏ Attach nomex parachute protector to 30’ shock cord

❏ Attach parachute to 3’ shock cord using a quick link

❏ Fold main parachute

❏ Hold parachute from top center of canopy

❏ Ensure all lines are straight and loop is at top of parachute

❏ Fold parachute in half

❏ Fold both sides into the middle tightly

❏ Fold parachute in half again, zig-zagging the lines back and forth

❏ Fold in half one more time

❏ Fold Drogue parachute

❏ Hold parachute from top center of canopy and straighten out lines

❏ Place parachute on the ground flat

❏ Fold in half to make a semicircle

❏ Fold in half again to make a wedge shape

❏ Fold shroud lines up into parachute and then back down again

❏ Fold parachute in half again

❏ Fold parachute in a z-formation

❏ Have certified NAR/TRA member approve parachute folding

❏ Ensure nomex parachute protector has plugged the airframe so that black powder blasts

will not harm the parachute

❏ Have certified NAR/TRA member approve parachute folding

❏ Insert parachute into airframe

Black Powder Charge Setup Procedure

Required PPE: Safety Glasses, Rubber Gloves

❏ Prior to launch day, the mechanical team will calculate the required amount of black

powder charges for each canister

❏ Prior to the preparation of the black powder charge canisters, the Safety Officer shall

ensure no power is connected to the recovery avionics and that the safety switches are in

the off state. All team members will put on safety glasses at this point, and keep glasses

on for the duration of the pre-launch set up procedure.

92

❏ Team mentor with NAR/TRA certifications shall prepare black powder charges

❏ The team mentor shall be the only person handling the black powder, and all other

team members will be alerted to avoid hazardous activities that could set off the

black powder

❏ Avionics member will wear gloves and secure the ends of the electric matches to the post

Airframe Separation Test

Required PPE: Safety Glasses

❏ Team mentor will perform the airframe separation tests

❏ The entire rocket will be assembled, including all fasteners and shear pins.

❏ The black powder canisters will be filled

❏ The leads of the canisters will be connected to a battery in the avionics bay

through a post in the bulkheads and will come out through one of the vent holes.

❏ With the leads of the charge exposed and the rocket fully assembled, the tips of

the cables will be made to touch the terminals of a 9V battery.

❏ The test is successful if the shear pins are broken and sections separate.

❏ Follow parachute setup and black powder charge setup procedure again to prepare for

launch

Avionics Setup Procedure

Required PPE: Safety Glasses

* Avionics and Payload Setup Sections may occur concurrently*

❏ Avionics team will prepare recovery electronics

❏ Prior to launch day, the altimeters will be set to deploy charges at the appropriate

altitudes

❏ Connect charges to appropriate terminals on altimeters

❏ Connect safety switches to appropriate altimeter terminals

❏ Plug 9V battery adaptors into batteries

❏ Temporarily activate safety switches to ensure proper setup. Both altimeters

should repeat a pair of beeps to indicate that both charges are connected and the

altimeters are prepared for launch

❏ Avionics team will prepare GPS and ground control electronics

❏ Plug LiPo into PCB/replacement

❏ Plug ground system board into designated laptop with Matlab GUI set up

❏ Make sure Matlab GUI is connected to wifi/hotspot to enable mapping

functionality

❏ Verify successful communications between ground station and PCB/replacement

by inspecting Matlab GUI

93

❏ Wait for GPS module to get a fix (red light on board will stop blinking and GPS

coordinates will start to be sent to ground system)

❏ Verify accuracy of fix by inspecting Matlab GUI

❏ Screw in PCB/replacement to sled

❏ Mount PCB/replacement LiPo battery to sled using velcro

❏ Avionics representative will ensure electronics are properly working

❏ Mechanical team member will secure the sled setup in the avionics bay tube

(coupler) by sliding dampening washers onto the threaded rods, sliding the tube

and bulkheads on, and then putting another dampening washer, a regular washer,

and a nut on each end to secure the avionics sled in the tube

Payload Setup Procedure

Required PPE: Safety Glasses

❏ This procedure will be carried out by 2 members of the payload team

❏ Ensure that the rover power switch is off during the preparation and assembly of the

payload and retention system.

❏ The rover is mated to the retention and deployment system before the whole assembly is

put inside the rocket

❏ Attach threaded rod designated for rover to the stepper motor

❏ Line rover up with threaded and smooth rods

❏ Run stepper motor in reverse to thread rover onto threaded rod

❏ Run charge leads through payload bulkhead

❏ Run shock cord from the main parachute through the shock cord guide on the forward

section airframe and through the designated hole in the payload bulkhead

❏ Activate rover electronics by turning on the power switch

❏ Attach payload assembly to the forward section with screws and nuts

❏ Attach all wires to connection posts on the top bulkhead of the avionics bay

❏ After rover is turned on, ensure reliable connectivity with ground control by sending a

test signal and verifying the receipt of a response from the rover

Airframe Setup Procedure

Required PPE: Safety Glasses

❏ Mechanical and Avionics representative will ensure avionics bay and payload bay are

correctly positioned in forward airframe

❏ Mechanical team member will secure avionics bay into forward airframe using rivets

❏ Safety officer and mentor will ensure the black powder charges are secure in airframe

94

❏ A Nomex cloth will be placed to cover the rover to protect from black powder charges by

a member of the payload team

❏ Parachutes will be placed in the airframe by a mechanical team member

❏ Main parachute placed between forward and nose cone sections

❏ Drogue parachute placed between coupler and booster sections

❏ Wind speed will be checked, and ballast will be added accordingly by a mechanical team

member

❏ Using a scale, ballast rings will be added to extended threaded rod on the booster

section side of the avionics bay until correct amount of ballast is added

❏ A nut will be placed on each threaded rod between the avionics bay bulkhead and

the ballast rings, and at the end of the ballast rings

❏ Black powder charge leads will be threaded through the holes in the ballast rings

and reconnected to the black powder charge canisters

❏ Shear pins will be added to connect the forward and nose cone sections, as well as the

forward and booster sections with a screwdriver

4.4.3 Motor Preparation Checklist

Motor Setup Procedure

Required PPE: Safety Glasses

❏ Motor shall be handled exclusively by team mentor with correct NAR/TRA certifications

❏ Motor will be placed in casing by mentor, steps are outlined below:

❏ Inspecting the motor casing to ensure there is no damage

❏ Gently twisting the delay/ejection module into the forward end of plastic liner

❏ Remove nozzle from top of liner

❏ If grains are not already assembled inside the motor, slide the grains inside the

motor liner, making sure it is place beveled end first.

❏ Insert the nozzle into the rear end of the plastic liner and remove nozzle cap

❏ Insert liner assembly , including the delay/ejection module, grains, and the nozzle,

into the rear end of the motor

❏ Casing will be placed in motor mount tube

❏ Motor retainer will be attached to end of rocket

❏ An overall check will be completed by mentor and safety officer

4.4.4 Setup on Launch Pad Checklist

Launchpad Procedure

Required PPE: Safety Glasses

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❏ Rocket will be carried to the launch pad by at least 2 team members, with 2 others

walking nearby to alert of possible tripping hazards

❏ Inspect launch pad for debris, damage

❏ Inspect rocket for damage

❏ Count that there are 3 shear pins connecting forward and nose cone sections, as well as 3

shear pins connecting the avionics bay and booster section

❏ Count that there are 4 rivets holding avionics bay in place

❏ Count that there are 3 Rivets holding payload bay in place

❏ Lower rail to horizontal position to easily attach rocket

❏ Slide the rocket onto 1515 rail using 1515 size rail buttons on rocket

❏ Turn on safety switches for avionics equipment

❏ Listen to ensure altimeters are ready for launch (repeated pair of beeps)

❏ Rotate rail to correct angle and lock in place

❏ Use an inclinometer application to check the tilt of the rail in the pitch and yaw axes

❏ Ensure the rail is not tilted towards any onlookers

❏ Ensure team members are at least 300 ft from the launch site other than safety officer and

mentor

4.4.5 Igniter Installation Checklist

Igniter Installation Procedure

Required PPE: Safety Glasses

❏ Safety officer will complete igniter procedure with the supervision of the TRA/NAR

certified mentor

❏ Inspect igniter for damage

❏ Remove motor retainer and motor casing cap

❏ Turning away from the sun so there is a shadow, run the 2 clips of the launch leads

together to check for a spark

❏ If there is a spark, have a team member notify the RSO immediately to prevent an

accidental igniter firing

❏ If no spark appears, feed igniter into motor cavity until reaching a hard stop against the

propellant grain

❏ Mark where the bottom of the motor touches the igniter

❏ Remove the igniter and make a loop at the marked spot

❏ Reinsert igniter into rocket

❏ Put cap on motor casing and retainer

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4.4.6 Launch Procedure Checklist

Launch Procedure

Required PPE: Safety Glasses

❏ Ensure there is no one within 300 feet of the rocket

❏ Alert all people in the vicinity of the impending launch

❏ A designated team member will launch the rocket

❏ After a safe landing and the approval from the RSO, send signal to trigger rover release

❏ Safety officer will allow retrieval of the rocket once the rocket is safely landed

❏ Team members will retrieve the rocket using ground control

❏ Rocket will be inspected for damage and possible danger such as sharp shards before

being touched

❏ If rocket is not safe to touch, proper protective equipment will be used to ensure safety

when transporting it back to the launch set up table

❏ Once rocket is deemed safe to touch, it will be brought back to the launch set up table for

post-launch inspection

4.4.7 Troubleshooting

Required PPE: Safety Glasses

Altimeters are not responding to arming command when on launch pad:

1. Turn off altimeters (avionics lead).

2. Return Avionics Bay to workbench (avionics lead).

3. Check for continuity (avionics lead).

4. If continuity is not the problem, replace wiring and electronic matches (avionics lead).

Required PPE: Safety Glasses

Misfire:

1. Immediately turn off power to the launch area (safety lead).

2. Wait for five minutes after the last misfire before approaching the rocket or wait until the

all clear is given (safety lead).

3. Remove the igniter (mechanical lead).

4. Check if igniter has discharged (mechanical lead).

5. Perform continuity test on ignitor (mechanical lead).

6. Replace ignitor with a new one if it is defective (mechanical lead).

7. If ignitor was not defective, examine cause of the misfire at the workbench (mechanical

lead).

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Required PPE: Safety Glasses

Cracks in the rocket’s body or broken:

1. Make quick repairs to the rocket if problem isn’t too serious

2. · If repairs cannot be made postpone the launch to another day

Required PPE: Safety Glasses

GPS fails to send coordinates to ground systems:

1. Replace/Charge batteries

2. Check to see if Antenna

Required PPE: Safety Glasses

Electronics fail to turn on:

1. Replace/Charge batteries

2. · Check wiring for any disconnections.

Required PPE: Safety Glasses

Motor does not ignite:

1. Check leads to see if they are connected to the igniter ends

2. · Feed the igniter into motor cavity until it stops against propellant grain

Required PPE: Safety Glasses

Black powder charges did not ignite during test:

1. Check and see if the leads are connected to the terminals

2. · Open charge canister and determine if the black powder needs to be packed tighter.

4.4.8 Post-Flight Inspection

Required PPE: Safety Glasses

1. Use GPS information to locate the launch vehicle if needed (avionics lead).

2. Ensure only qualified personnel approach the launch vehicle.

3. Inspect launch vehicle for charges that have not gone off. If any explosives are still intact,

immediately disarm them (safety lead).

4. Ensure all sections of the rocket are present.

5. Take photos of the recovery sight and launch vehicle.

6. Have the safety officer check for any possible hazards that could result in injury (sharp

edges, electronic/physical damage).

7. Bring launch vehicle back to determine the official altimeter reading (safety lead).

8. Remove motor and ejection charges after cooling (mechanical lead).

9. Clean residue off of the motor (mechanical lead).

10. Clean all parts of the motor casing (mechanical lead), preferably with wet wipes or

acetone.

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11. Have the mechanical/manufacturing lead check:

a. Main body, motor section, nose cone for cracks or anomalies

b. Drogue/Main parachute status and performance

c. Any other possible defects or damage that may have occurred

12. Report any anomalies to the safety officer.

5 Payload Criteria

5.1 Changes Since CDR

The following changes have been made to the payload design since the CDR. All changes were

made on the basis of improving the safety and reliability of the payload and retention system.

Printed Circuit Board

All electronic components of the rover are now mounted on a custom designed printed circuit

board (PCB). This significantly reduces the complexity of wiring and optimizes the use of space

compared to a perfboard or a free-floating “rat’s nest” wiring of components. The reduced

number of soldered connections and exposed wires improves safety due to the lower probability

of detachment or entanglement of wires.

Figure 5.1: The assembled PCB.

New Motor Driver

The L298N motor driver had been selected for CDR due to its popularity and ease of use.

However, the driver would have been too large to fit in the rover’s frame with the other

components on the PCB. A new motor driver based on the MX 1508 dual H-bridge chip was

selected. It has a much smaller footprint but can still provide ample power to the two geared DC

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motors while also having a much lower idle current. The driver provides full speed control in

forward and reverse directions through the use of PWM pins on the Arduino Nano

microcontroller.

Changes to CAD Model

Minor changes had to be made to the design of the 3D printed mechanical components to ensure

reliable printing performance while retaining strength. Most fillets were replaced by chamfers to

reduce the overhang angles, and the rear wheel mounting mechanism was made into a separate

part that would be attached to the frame after being manufactured. The actuating arm was

modified to include a bend in the middle of the arms and a notch on one side to accommodate the

length of the internally threaded T-nut used to deploy the rover.

Eyebolt Location for the Main Parachute

To address concerns about the strength of 3D printed parts, the payload retention bulkhead no

longer supports the weight of the vehicle under the main parachute. Instead, the eyebolt that

attaches the main parachute to the forward section of the rocket is mounted directly on the

avionics bay bulkhead.

5.2 Payload Features

5.2.1 Design Overview

Rover:

Figure 5.2.1a: Rendering of the final rover CAD model.

The rover’s design has remained largely unchanged since CDR. It’s a three wheel taildragger

vehicle with two main wheels driven by geared DC motors and a trailing wheel for stability. The

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drive wheels also perform the function of soil sample collection by turning in reverse after the

rover has travelled the required ten feet from the landing site. The frame is constructed with 3D

printed nylon and has attachment points for all other components. An actuating arm, also 3D

printed with nylon, serves as the mechanism to flip the rover in case it is deployed upside down.

The arm is also deployed during soil collection since it can provide extra drag as the wheels turn

in reverse.

Figure 5.2.1b: Top view of the rover showing the locations of the components.

The rover is powered by a 7.4V LiPo battery connected to the PCB through a one-way, single

throw safety switch. There are three servo motors, two of which are used to actuate the arm and

the third turns the ultrasonic distance sensor side to side to detect obstacles.

Retention System:

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Figure 5.2.1c: CAD model of the payload retention and deployment system assembly.

The payload is located in the forward section of the launch vehicle, above the avionics bay. A

bulkhead constructed with 3D printed nylon is used to secure the assembly to the airframe. A

slotted hole in the bulkhead allows the shock cord for the main parachute to be routed to the

eyebolt which is located on the avionics bay bulkhead. The nylon bulkhead is only used to

support the payload hardware, not the weight of the launch vehicle after parachute deployment.

A stepper motor is mounted on the bulkhead which is connected to the threaded rod via a flexible

shaft coupling. The threaded aluminum rod engages with the internally threaded t-nuts on the

rover and is the primary payload retention mechanism. A smooth nylon rod is used to stabilize

and guide the rover during flight and deployment. A pusher that also has a threaded insert is

deployed along with the rover and serves to push the rover all the way out of the airframe. A

shock cord guide with a slotted hole allows the shock cord to be routed around the rover to

protect it from the forces during and after parachute deployment.

The deployment sequence starts with ground control verifying that the vehicle has landed safely.

A radio signal is sent to the flight computer to initiate rover release, which commands the stepper

motor to turn the threaded rod counterclockwise for a predetermined number of revolutions.

After the rover has fully exited the vehicle, ground control then sends a startup command to the

rover’s transceiver after which the rover begins driving.

5.2.2 Mechanical Elements

5.2.2.1 Rover

The rover is composed of the following primary mechanical subsystems, discussed in detail

below: the frame, the drive system, the actuating arm, and the soil sample collection mechanism.

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Frame:

The frame holds all components together and provides protection to the battery and electronic

components during flight and after deployment. It is 3D printed with Taulman Nylon Alloy 910,

which is a nylon blend designed specifically for 3D printing and offers an as-printed tensile

strength of 8100 psi. Nylon also offers favorable toughness and impact resistance compared to

most other plastics. Figure 5.2.2.1a shows the as-printed CAD model of the frame.

Figure 5.2.2.1a: As-printed CAD model of the single-part frame.

Drive System:

The rover is powered by a pair of geared DC motors with encoders for recording the distance

covered. The motors operate at a nominal voltage of 6V to 12V and have a stall current of 2A.

The motor has a no-load speed of 160 rpm at 6V and a rated torque of 0.2 kg.cm.

Figure 5.2.2.1b: CAD drawing of the geared DC motor.

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The two motors are powered by a dual H-bridge motor driver board based on the MX 1508 chip.

The driver operates at the battery voltage of 7.4V nominal, is rated for 2 amps to each motor and

includes a built-in thermal protection circuit.

Actuating Arm:

The actuating arm is attached to two servo motors and serves two functions: (1) to flip the rover

in case it is deployed upside down and (2) to provide drag to the wheels to aid in soil sample

collection. The arm is 3D printed with the same nylon filament as the frame.

Figure 5.2.2.1c: CAD model of the as-printed actuating arm

The arm has cavities on either side for attaching servo horns. One side has a notch for providing

clearance for the threaded inserts on the frame.

Soil Sample Collection Mechanism:

The wheels have scoops and valves that collect soil samples after the rover has driven the

required 10 feet away from the landing site.

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Figure 5.2.2.1d: The wheel assembly.

The wheel assembly consists of a base, six valves, and a lid. The base and valves are printed with

PLA and the lid is printed with translucent PETG. As the wheel turns in reverse, the scoops pick

up soil and the valves allow the soil to fall into the internal cavity while keeping the existing soil

sample from falling out. The valves are designed such that they turn freely on the posts and are

constrained by the inner wall of the wheel. The wheel’s outer diameter including the scoops is 70

mm, and the inner circle is 49 mm in diameter. Each wheel has an internal free volume of

approximately 18 cubic centimeters.

The wheels are connected to the motor with an M3 screw that threads into the shaft of the motor

and holds the assembly together. The soil sample in the wheel is accessible by removing the

screw and disassembling the wheel assembly.

5.2.2.2 Retention and Deployment System

The Retention and Deployment System consists of the bulkhead, stepper motor, shaft coupling,

pusher, and the threaded and smooth rods.

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Figure 5.2.2.2a: The payload assembly showing the retention system components.

The bulkhead is 3D printed from nylon and is attached to the airframe with M4 stainless steel

screws and nuts. The stepper motor is attached to the bulkhead with four M3 screws. The pusher

(shown in black) is a part 3D printed with PLA and is threaded into the retention system before

the rover. The ejection charge for the main parachute is located in the space between the

bulkhead and the pusher. The bulkhead has a hole for the shock cord to pass through and attach

to the eyebolt underneath, and additional holes for the ejection charge leads. The threaded rods

are made of aluminum with a single thread and a standard thread pitch of 20 threads per inch.

The rod engages with the rover at two points in the frame, where internally threaded t-nuts are

located.

Having two points of contact ensures that the rover will not want to twist away from the rods

during deployment and this was confirmed by our deployment test. The flexible shaft coupling

ensures effective transmission of torque even when the aluminum rod is slightly bent from the

axis of the motor’s shaft due to weight. The smooth nylon rod goes through holes located at the

bottom of the rover’s frame and provides stability during flight and deployment. The deployment

test also confirmed that there is no binding or other issues related to over-constrainment of any

part when the rover is deployed.

Rover retention is achieved by the self-locking nature of the threaded rod and nuts. A power

screw is said to be self-locking if there is a positive torque required to lower the weight of the

load on the nut. The formula for calculating the torque to lower a load on a power screw is given

by:

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T𝐿 =𝐹∗𝑑𝑚

2 (

𝑙 + 𝜋𝜇𝑑𝑚

𝜋𝑑𝑚−𝜇𝑙)

Where F is the load applied, dm is the mean diameter of the screw, l is the lead, and μ is the

coefficient of friction between the screw and the nut. The lead angle is based on the pitch and the

mean diameter, and in our case we can simplify the above equation to state that the threads are

self-locking if the coefficient of friction is greater than the tangent of the lead angle. This is

equivalent to saying that the part of the equation in parentheses is positive. In our case, the mean

diameter is 0.25 in and the pitch is 20 threads per inch. This gives us a lead angle of 3.64°, and

the tangent of it is

tan(3.64°) = 0.0636

The condition for self-locking is met if this value is lower than the coefficient of friction between

the aluminum rod and the steel nuts. Commonly cited values of μ between clean and dry steel

and aluminum surfaces are between 0.4 and 0.6, which means our lead angle is well within the

requirement for self-locking.

The fact that an external torque would be required to move the rover on the rod proves that the

tensile or compressive forces during flight will not be able to unscrew the rover out of its

retention assembly or otherwise make it unsafe. The shaft coupling that attaches the threaded rod

to the motor shaft is a single-piece aluminum alloy spring coupling that can absorb

misalignments and tension in the rod through elastic deformation. In the highly unlikely case that

the self-locking condition fails and the rod experiences a torque due to tension, the holding

torque from the stepper motor will be able to stop it from turning and releasing the rover

prematurely.

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Figure 5.2.2.2b: Image from the payload deployment test.

5.2.3 Electrical Elements

The rover is composed of the following primary electrical subsystems, discussed in detail below:

the PCB, power, control, radio communication, and navigation.

PCB:

The PCB was custom-designed to house the four discrete electronic components that comprise

the electronic system of the rover: the Arduino Nano microcontroller, the MPU 6050

Accelerometer Breakout Board, the Adafruit RFM96W LoRa Transceiver, and the MX 1508

Dual H-Bridge Motor Driver.

Figure 5.2.3a: Rendering of the PCB showing the electronic components.

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The PCB has through-hole pads for soldering the Arduino Nano microcontroller, the MPU 6050

Accelerometer Breakout Board, the Adafruit RFM96W LoRa Transceiver, and the MX 1508

Dual H-Bridge Motor Driver. Pads for power input and connection to the external sensors and

motors are located around the edges of the board. The board is attached to the rover via nylon

standoffs on the frame with screws at the mechanical attachment holes shown.

Figure 5.2.3b: PCB layout showing the traces and pad locations.

The circuit board schematic and layout were made in the online Electronic Design Automation

editor called easyEDA and the design files were sent to a prototype PCB fabrication house to be

manufactured.

Power:

Power is supplied by a 7.4V 450 mAh lithium polymer battery. The battery is wrapped in an

orange colored tape for visibility and is attached to the frame of the rover using Dual Lock which

has been shown to be strong enough to hold a 9V battery even during the high impact landing of

our first subscale flight.

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Figure 5.2.3c: View of the battery, safety switch, and switch holder.

The safety switch is a simple single pole, single throw switch which is accessible from the top of

the rover and can be used to cut off the power supply from the battery to the PCB at any point.

Control:

Control is provided by an Arduino Nano microcontroller which performs all functions pertaining

to driving, orientation correction, obstacle avoidance, travel distance measurement and radio

communication with ground control. The microcontroller polls the accelerometer when it is first

activated on the ground, and periodically during driving to get orientation data. If it is

determined that the rover is upside down, the actuating arm is deployed and retracted to flip the

rover. To keep track of the distance covered, the microcontroller takes interrupt signals from the

hall-effect encoder on the left driver motor. After the rover has travelled the required distance, it

stops driving and begins the soil collection sequence. This involves deploying the arm, driving in

reverse, driving forward for a few seconds and repeating the sequence a number of times. After

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soil collection is complete, the rover sends an end-of-mission signal to ground control and stops

all other operations.

Radio Communication:

The Adafruit RFM96W LoRa module is used by the rover to communicate with ground control.

The transceiver is placed on the PCB and communicates with the Arduino Nano through the SPI

protocol. The transceiver communicates with ground control at 434.0 MHz. The rover has an

omnidirectional helical antenna and the ground control transceiver uses a high gain, directional

Moxon rectangle antenna to ensure a reliable wireless connection.

Figure 5.2.3d The transceiver module seen in the middle with a helical antenna.

When the rover is first powered on, the transceiver is in listening mode and does not transmit

anything until a message is received from ground control. When the ground control sends a test

signal that contains a unique header, the rover replies with a character array to indicate a

successful connection and goes back to listening. All powered mechanical systems of the rover,

i.e the DC motors and servos are inactive when the rover is in listening-only mode. After landing

and deployment, the ground control transceiver then sends an activation signal after which the

rover begins driving. After the rover has completed its mission, it sends a message to ground

control and stops all other operations until the ground crew can retrieve it.

Navigation:

To avoid obstacles while driving, the rover uses an ultrasonic sensor mounted on a servo motor.

The servo motor sweeps side to side to cover a total angle of 40° while the sensor takes distance

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measurements at each corner. The driving algorithm is written such that the servo completes one

sweep cycle every second, and the distance information is stored in variables which are used to

determine whether to drive straight, turn right, turn left, or to make a sharp turn with one wheel

stationary if an obstacle is detected on both sides. The rover does not reverse while navigating to

avoid collecting any soil before it travels the full distance.

Figure 5.2.3e: The ultrasonic sensor mounted to the servo at the forward facing side of the

rover.

5.3 Schematics

5.3.1 As-Built Mechanical Schematics

The completed schematic drawings of the rover and the retention system hardware are shown in

figures 5.3.1a and 5.3.1b below.

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Figure 5.3.1a: As-built CAD drawing of the rover.

Figure 5.3.1b: CAD drawing of the retention system showing the relevant dimensions.

The fully assembled mass of the rover is 290 grams and the retention and deployment system

weighs 235 grams, not including the forward section airframe where it is located. Due to the fact

that most mechanical components of the payload were 3D printed, the final dimensions were

nearly identical to the dimensions of the CAD models and assemblies.

5.3.2 As-Built Electronic Schematic and Block Diagram

The figure below shows the final electronic schematic which was used to design the PCB layout.

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Figure 5.3.2a: Final electronic schematic of the rover on which the PCB is based on.

Figure 5.3.2b shows the block diagram that represents the operation of the rover. The

microcontroller maintains two-way communication with ground control through the LoRa

transceiver through which it can receive test and activation signals. It takes inputs from the

accelerometer, ultrasonic sensor and the hall-effect encoder, and sends outputs to control the

three servo motors and the two DC motors.

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Figure 5.3.2b: The block diagram for the rover.

5.4 Construction

The rover and the retention system were constructed in iterative steps while building and testing

the fit and tolerances of all mechanical elements and going back to the CAD models and making

changes wherever necessary. The use of 3D printing allowed us to rapidly prototype and test

designs and solve any issues that showed up during testing.

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Figure 5.4a: Early prototype of the frame being used to decide on the circuit board layout.

Multiple iterations of the frame were printed in PLA due to its ease of printing and low material

cost. The PCB had to be designed to conform to the limited space in the rover while still

accommodating all of the necessary components. After ironing out issues related to assembly and

fit, the parts were printed with nylon and the final assembly was started.

Figure 5.4b: Nylon standoffs screwed into the brass heat-set inserts.

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The DC motors and servos are attached to the frame with stainless steel screws, and the PCB is

mounted on the nylon standoffs shown above using M3 nylon screws. The wheel well is super-

glued to the bottom of the frame.

Figure 5.4c: Parts of the rover before assembly and testing.

The software for the rover was written concurrently with the construction of the hardware which

allowed some testing to be conducted before the final iteration of the hardware was ready.

Testing of the transceiver connection, encoder function, and orientation correction were carried

out in conjunction with ongoing hardware development.

Figure 5.4d: Inverted view of the rover prior to being flipped by the actuating arm.

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5.5 Flight Reliability Confidence

Each element of the rover was selected for its predicted and measured reliability and safety.

Through theoretical demonstration and informal testing, we are highly confident that the

retention and deployment system will successfully retain the rover prior to rover deployment and

that the rover will be able to successfully deploy from the vehicle upon landing.

Additionally, the rover was tested many times to ensure that it is able to orient itself and navigate

around obstacles at the level of the rover’s ultrasonic sensor following landing. Should the rover

exit the vehicle in anything other than nominal orientation, it will be able to detect this and use

the actuating arm to correct itself to nominal orientation. Additionally, should the rover

encounter obstacles during its 10 foot path away from the vehicle, we are confident that the

ultrasonic sensor will detect it and that the rover’s obstacle avoidance algorithm will be able to

successfully avert all obstacles.

To take extra precaution and improve the flight reliability confidence of the rover, the rover’s

battery is wrapped in high visibility tape, the rover has a functional safety switch, and the rover’s

battery is fastened with the flight-tested Dual Lock which successfully retained the batteries on

our first subscale launch.

There was initial concern about having the payload so close to the black powder charges when

they go off. It was thought that the payload might be damaged by either the heat or force of the

ejection charges. Though the charge ejection test on the full scale vehicle, however, the rover

sustained very minimal cosmetic damage to one wheel. As a result of this test, we are confident

that the rover will be able to survive the heat and force of the ejection charges, further

demonstrating the flight reliability confidence of the rover.

Lastly, though there were initial concerns about the strength of 3D printed materials, the strength

of 3D printed nylon on the rover has been tested and sufficiently verified such that we are

confident that the frame rover and all 3D printed elements will be able to withstand any stresses

during flight.

5.6 Payload Demonstration Flight

The Payload Demonstration Flight has not been performed yet and will be performed in the

future to verify the success of the payload operations and of the retention and deployment

system.

We plan to fly the rocket on March 9th along with the payload. The flight will be deemed

successful if the rover is safely retained in the vehicle throughout the duration of the flight and is

fully deployed after landing.

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6 Project Plan

6.1 Testing

6.1.1 Mechanical Tests

The following test were designed and performed to improve team safety and efficiency.

Performing the airframe drop test validated the strength of the fiberglass body tubes. The black

powder sectioning test was conducted to ensure that the deployment charges provided enough

force to shear the connecting pins and separate the body tubes. Successful separation is required

for nominal parachute deployment and safe recovery of the rocket. The initial runthrough test

was conducted to ensure team understanding of launch preparations. Fully understanding the pre-

launch procedure helps to eliminate missed steps on launch day and ultimately leads to a safer

team.

Table 6.1.1a

Test Initial Runthrough

Description This test proves whether members of the team can adequately set up our

rocket without any delays or interruptions.

Rationale During the launch we will perform at the competition in April, it is

imperative that our team will be able to adequately set up and ready our

rocket when it is our turn to launch. In order to guarantee the highest

efficiency possible in this process, we must perform this test.

Environmental

Requirements

Be in a setting with ample space where our team will be able to fully

assemble our rocket.

Procedure 1. Record list of equipment and applications needed for launch

2. Assemble members who are willing to be part of the group that will

be putting the rocket together.

3. Practice the assembly procedure and record what was done, what

needs improved, what was exceptional, and who did what.

4. Repeat step 3 until the team is able to assemble the rocket with the

highest possible efficiency.

Required

resources

Rocket (and all of its components)

Any equipment needed throughout the assembly (screwdrivers, computer,

etc.)

Success Criteria This test will be considered a success if we can properly set-up the rocket

in a functional manner without any interruptions or missing components

needed.

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Proof of Success Successfully set up rocket for launch on launch field.

Table 6.1.1b

Test Airframe Drop Test

Description Drop a section of the airframe from a height to allow it to reach 75ft-lbs of

kinetic energy

Rationale To ensure the airframe can withstand the forces of landing within the

NASA requirements

Environmental

Requirements

Be in a setting with enough vertical space to drop the body tubes from a

height that will cause the airframe to land with 75 ft-lbs of kinetic energy.

Ensure landing and surrounding area is clear of persons before dropping.

Drop rocket outside to best approximate the conditions of a landing.

Procedure Use conservation of energy to calculate the height needed to drop a section

of the rocket to reach 75 ft-lbs upon landing.

Required

resources

Section of the rocket, access to specified height outside

Success Criteria Inspection of the dropped airframe finds that it is not damaged by the drop.

Proof of Success Dropped airframe was not damaged and was able to be used in the launch

vehicle cleared for flight.

Table 6.1.1c

Test Black Powder Sectioning Ground Test

Description Test the ejection charges to ensure the calculated black powder is enough

to separate the sections of the rocket

Rationale The separation of the rocket is integral to a safe recovery and landing of the

rocket. The amount of black powder in the ejection charge canister must be

tested to ensure it is adequate to separate the rocket while not being so

much that unnecessary force is applied to the shock cords. This test also

serves to check that the parachutes and electronics are adequately protected

from the black powder charge.

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Environmental

Requirements

Test should be carried out in an open outdoor environment with adequate

space for the rocket to be separated by a distance the length of the shock

cord. All personnel should be a safe distance away from the rocket

especially in the direction the rocket will be sectioning.

Procedure 1. Calculate the required amount of black powder charges for each

canister

2. Safety Officer shall ensure no power is connected to the recovery

avionics and that the safety switches are in the off state.

3. All team members will put on safety glasses

4. The team mentor shall be the only person handling the black

powder

5. Avionics member will wear gloves and secure the ends of the

electric matches to the post

6. The entire rocket will be assembled, including all fasteners and

shear pins

7. The black powder canisters will be filled

8. The leads of the canisters will be connected to a battery in the

avionics bay through a post in the bulkheads and will come out

through one of the vent holes.

9. With the leads of the charge exposed and the rocket fully

assembled, the tips of the cables will be made to touch the terminals

of a 9V battery

Required

resources

Black powder, fully set up rocket,

Success Criteria The test is successful if the shear pins are broken and sections separate

Proof of Success Rocket sectioned properly

6.1.2 Avionics Tests

After carefully designing and constructing the avionics systems, we performed thorough testing

on each aspect of the system. This testing was performed in order to ensure proper function of

the systems as well as to identify any weaknesses or flaws in our design that were not revealed

during the initial design phase. The following tables outline each of the tests that were performed

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by the avionics team. These tests encompass all aspects of the avionics system including, GPS,

recovery, ground control, and rover release.

Table 6.1.2a

Test Name Obstructed GPS Fix Test (GPS Test 1)

Description This test verifies that the GPS module can obtain a fix and that the GPS is

sufficiently accurate. The main avionics system will be powered on and the

GPS coordinates from the GPS module will be read until it obtains a fix.

Rationale In order to be usable, the GPS module must be able to obtain a GPS fix

within a reasonable amount of time (approximately 1 minute).

Additionally, the GPS coordinates obtained must be reasonably accurate

(within 10 feet) in order to be used to locate the vehicle after landing. If the

GPS cannot obtain a fix or if the GPS is not sufficiently accurate, it will not

be able to help us locate the vehicle.

Environmental

Requirements ❏ System must be located in a field with no tall buildings within at

least 50 feet of the system. Environment should be as similar to

launch field as possible.

❏ Must not be raining.

❏ Ideally performed on a day with minimal cloud coverage (clear

skies).

Procedure 1. Connect teensy to laptop using micro-USB.

2. Power on system and start timer.

3. Open Arduino IDE Serial Monitor with 9600 baud rate and verify

that GPS does not have a fix but that the GPS module’s red LED is

blinking. Serial Monitor output should look like this (there’ll be a

bunch of commas separating nothing):

4. Continue to monitor the Serial Monitor until GPS has a fix (will

look similar to this):

5. Once GPS acquires a fix, stop the timer and power off the GPS

system record the time shown on the timer and copy the data from

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the serial monitor into a separate file.

6. Plug most recent NMEA sentence into converter (like this one:

http://www.gonmad.co.uk/nmea.php) to convert from

degrees/minutes/seconds format to decimal degrees.

7. Plug resulting decimal degrees latitude/longitude into a tool (like

this one: https://www.maps.ie/coordinates.html) to make sure GPS

fix is close to actual position (test location).

8. Get current GPS location using a reliable site (like this one: ) and

plug both module-calculated latitude/longitude and the

latitude/longitude from reliable site into a distance calculator (like

this one: http://boulter.com/gps/distance/). Record distance between

two points.

Required

Resources ❏ Timer (smartphone timer is sufficient)

❏ Laptop with Arduino IDE installed

❏ Micro-USB cable for connecting teensy to laptop

Success Criteria ● GPS is able to get a fix within 1 minute (ideally within 45 seconds

but 1 minute is acceptable).

● Coordinates obtained from GPS fix are within 10 feet of actual

coordinates (spec sheet says 6 feet but 10 feet is acceptable).

Proof of Success ● Time recorded from timer

● Serial monitor output from time powered on to time powered off

Results Testing showed that it takes around 37 seconds for the GPS to get a fix

through G12 fiberglass.

Analysis These results were in-line with what we expected to see and show that the

gps system will be able to easily acquire a fix between the time the avionics

bay is assembled and powered and the launch.

Table 6.1.2b

Test Name 2500 Foot Obstructed Range Test (GPS Test 2)

Description This test verifies the transceiver range meets the system requirements. The

ground and main avionics systems will be powered on and separated until

they are 2500 feet apart.

Rationale In order to be able to recover the vehicle after landing, the system must be

able to connect at the largest acceptable range from the launch site -- 2500

feet. The system must be tested to ensure that this range can be met. This

test was consolidated from (and therefore encompasses) all range tests

below 2500 feet (e.g. 100 foot range test, 500 foot range test, etc.).

Environmental ❏ System must be located in a field (or close to a field) with no tall

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Requirements buildings within at least 50 feet of the system. Environment should

be as similar to launch field as possible.

❏ Must not be raining.

❏ Ideally performed on a day with minimal cloud coverage (clear

skies).

❏ Must be at least 2500 feet of flat, traversable ground.

Procedure 1. Determine location of stationary ground system and record GPS

coordinates of this location. Use Google Maps’ distance measuring

system to determine a suitable stationary ground system location

and the final location of the vehicle system (2500 feet from the

ground system).

2. Position ground system and vehicle system at location of stationary

ground system. Make sure the ground system antenna is pointed in

the direction the vehicle system will move and do not move the

ground system or antenna once positioned. Try to keep the ground

system at roughly the same height as the vehicle system (this means

holding it a few feet off the ground).

3. Connect ground system microcontroller to one laptop and vehicle

teensy to the other laptop.

4. Power on ground system and vehicle system.

5. Open the Arduino IDE Serial Monitor with 9600 baud rate on each

computer.

6. Wait for GPS to get a fix and transceivers to successfully pair (as

indicated on Serial Monitor). GPS fix will look like this:

Successful transceiver pair will look like this:

7. Once transceivers are paired and GPS has a fix, record the serial

port the ground system is connected to. This is needed to initiate the

logging script. The name of the serial port can be found by opening

the Arduino IDE and navigating to Tools then Port. The

checked/listed port is the serial port the microcontroller is

connected to.

8. Close the Arduino IDE on the ground computer. In order to run the

logging script, nothing should be interacting with the serial port

(otherwise the logging script will fail to start). Keep the Arduino

IDE open on the vehicle computer (the Serial Monitor will be used

to tell those carrying the flight system whether or not the

transceivers have lost connection which would be an indication for

the test to stop).

9. On the ground computer, open a bash shell (Terminal on

macOS/Ubuntu) or Command Prompt (if using Windows).

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10. On the ground computer navigate to the root directory of the GPS

repository (“cd [path to repository]”).

11. On the ground computer, start the logging script by running “python

logging_script.py”.

12. When prompted, enter the USB port (identified in step 7), a

descriptive prefix for the log which the current date and time will be

appended to, the number of the test being performed,, and a

description of any relevant information about the test (e.g. weather

that day, location of test, etc.). This information will be used to set

up the logging environment, name the log file, and record details

about the test at the top of the log file.

13. After entering the test description, a line saying “cat /dev/[USB port

name] > [path info]/logs/[log file name] which indicates that the

ground system has started logging data to the file. The terminal

output should look something like this:

Check the log file to ensure that data is being recorded properly.

Log file should look something like this:

14. Once its verified that data is being logged to the ground computer,

those carrying the flight system walk away from the ground system,

avoiding blocking the vehicle system with anything but the

materials that naturally encompass the flight system.

15. If the transceivers disconnect (may happen if RSSI values reach

below -120), terminate the logging script and return the vehicle to

the stationary ground system location. The test has failed.

16. Once the full range (2500 feet) has been reached by the vehicle

system, walk the vehicle system back to the ground station in the

same manner. The ground station should not move until the flight

system is back with the ground system then terminate the logging

script.

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a. It’s important to not terminate the logging script at 2500 feet

from the ground system. By walking the vehicle system

back to the ground system, we are doubling the number of

data points this test will collect, thus improving the accuracy

of the distance vs. RSSI plot and other further analysis of

the logging script.

17. Whether or not the test failed, once the vehicle system is back at the

ground station, open the log file and visually verify successful data

collection.

18. Run the data processing script on the resulting log file. Do this by

running “python test_processing.py [name of log file]”. Only enter

the name of the log file found in /logs -- the script searches through

this directory since the logging script places all log files here. The

processing script will create a processed version of the log file and a

png plot of the distance vs. RSSI data. The processed version of the

log file will be placed in /logs/processed and the file name will

match that of the log file with “_processed” appended to the end of

the filename. The generated plots will be placed in /logs/plots and

will be named with “_plot” appended as a png. After running the

test processing script, verify that the processed and plots were

created.

19. Power off the vehicle and ground stations.

20. Push the resulting raw log file, processed log file, and png plots to

Github and specify in the commit message that the files pushed are

from a system test.

Required

Resources ❏ Two micro-USB cables for connecting teensy to laptop

❏ Latest version of GPS repository

(https://github.com/PittRocketryTeam/GPS)

❏ Data logging/collection script

(https://github.com/PittRocketryTeam/GPS/blob/master/log

ging_script.py)

❏ Test processing script

(https://github.com/PittRocketryTeam/GPS/blob/master/test

_processing.py)

❏ Two laptops each with the Arduino IDE, Python 2.7, and Python

3.6 installed

Success Criteria ● Transceivers are able to maintain connection up to 2500 feet.

Proof of Success ● Processed test data and graph of distance vs. RSSI.

Results Considerable difficulty was had attempting to find an area large and flat

enough for us to perform this complete range test. Ultimately we were able

to test to about 1300 ft (400 m).

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Analysis The resulting plot of distance vs. RSSI shown in the results section

indicates that the RSSI began to settle into a range of between about -85

and -110. Based upon this pattern we expect that we would be able to

achieve distances significantly longer than 1300 feet before the RSSI

dropped to an unacceptable level.

Lessons

Learned

During the planning and execution of this test we learned some important

lessons. We learned that finding a location with the necessary range

required is much more difficult than expected. In the future we will find a

location that meets the requirements and then make a trip there to perform

our testing. We also learned that the process as described above can be

rather cumbersome to perform in the field. It is for this reason that we

began to use the ground control GUI to perform subsequent range tests. In

the future, the ground control GUI will be used for all GPS range tests.

Table 6.1.2c

Test Name Transceiver Reconnection Test (GPS Test 3)

Description This test verifies that the transceivers can reconnect if they lose range. To

perform this test, the transceivers will be connected then forced to

disconnect through distance or obstruction. The distance or obstruction will

then be removed to ensure that the transceivers can reconnect.

Rationale While the system has/will be range-tested to 2500 feet, the vehicle’s target

apogee is 4750 feet. Because the system will not be range tested to 4750

feet, a test needs to be performed to ensure that if and when the transceivers

disconnect they will be able to reconnect.

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Environmental

Requirements ❏ System must be located in a field (or close to a field) with no tall

buildings within at least 50 feet of the system. Environment should

be as similar to launch field as possible.

❏ Must not be raining.

❏ Ideally performed on a day with minimal cloud coverage (clear

skies).

❏ Must contain adequate distance/obstructions to temporarily interfere

with transceiver signals.

Procedure 1. Follow steps 1-13 of GPS Test 3.

2. Obstruct the vehicle and ground systems in any way possible until

the transceivers lose connection, as shown on the Arduino IDE

Serial Monitor (but be sure to record loosely how the transceivers

were obstructed).

3. Remove the obstruction so that the transceivers are in the initial

position (where they were connected prior to introducing the

obstruction).

4. Follow steps 17-20 of GPS Test 3.

Required

Resources ❏ Two laptops each with Arduino IDE installed

❏ Two micro-USB cables to connect laptop to teensy

Success Criteria ● Transceivers successfully reconnect and transmit data following

disconnection.

Proof of Success ● Log file showing transceiver connection followed by transceiver

disconnection followed by transceiver reconnection.

Results The results of this test indicate that the transceivers will be able to

reconnect after being disconnected. Furthermore, this has been observed

throughout other tests requiring the use of the transceivers. Occasionally a

packet would be dropped due to unexpected interference or other

environmental interference. The transceivers were observed to be able to

reconnect after such instances.

Analysis After completion of this test we are confident that the transceivers will be

able to reconnect if they are to temporarily lose connection. In the event

that the airframe goes beyond the range of the transceiver, as is likely at

apogee, the transceivers will be able to reconnect and allow us to observe

the final location of the airframe upon landing.

Table 6.1.2d

Test Name Transceiver Interference Test (GPS Test 4)

Description This test verifies that other transceivers on the 433 MHz band will not

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interfere with our transceivers.

Rationale The 433 MHz band is a common transmission frequency. If other users

transmitting at 433 MHz causes interference, it could result in corrupted

data. This is a risk to mission success as it could prevent the location of the

airframe upon landing.

Environmental

Requirements ❏ GPS system must be able to successfully transmit over adequate

range so that any interference would not be negligible.

Procedure 1. Set up GPS system so that it is operating as it would be during a

flight.

2. Separate the ground and flight systems by at least 20 ft and observe

RSSI

3. Set up a seperate set of transceivers transmitting at 433 MHz.

4. Using the Arduino IDE serial monitor, observe the RSSI values and

record any changes in RSSI after interference is introduced.

Required

Resources ❏ Two laptops each with Arduino IDE installed

❏ Two micro-USB cables to connect laptop to teensy

❏ Two 433 MHz transceivers to create interference.

Success Criteria ● The RSSI does not change with the introduction of another pair of

transceivers transmitting at 433 MHz.

Proof of Success ● Observation of the Arduino IDE serial monitor.

Results The addition of another pair of transceivers operating at 433 MHz had no

effect on the RSSI of the GPS system transceivers.

Analysis Based on the results of this test, the presence of other devices operating on

the 433 MHz band will not affect the performance of our transceivers.

Therefore we are confident that we will be able to recover the airframe

even while other teams operate their systems.

Table 6.1.2e

Test GPS Fix Test Oriented Downward (GPS test 5)

Description This test determined whether or not the GPS would be able to acquire a fix

while in the worst possible orientation, upside down.

Rationale Because we cannot control the orientation in which the airframe lands, we

need to be sure that the GPS will still be able to get a fix no matter how it is

oriented.

Environmental ❏ This test must be performed outside

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Requirements ❏ An open area with a limited number of tall buildings around is

optimal.

Procedure 1. Locate an open space outside sufficiently away from lage buildings

2. Orient the Adafruit GPS patch antenna towards the ground.

3. Power the GPS system and start a timer.

4. Observe the GPS fix light. It will blink once every second when it is

attempting to acquire a fix and once every 15 seconds once it has

acquired a fix.

5. Stop the timer once it is observed that the GPS has acquired a fix.

6. Power off the GPS system.

Required

resources ❏ GPS system

❏ Stopwatch

Success Criteria ● The GPS acquires a fix in under two minutes, and acceptable

amount of time to wait until receiving the final location of the

rocket.

Proof of Success ● The recorded time it took to achieve a fix.

Results The GPS took 76 seconds to acquire a fix when its patch antenna was

oriented downward.

Analysis The results of this test show that the GPS will be able to acquire a fix even

in the worst case scenario where the GPS is oriented downward upon

landing. This reinforces our design decision to not affix an external antenna

on the GPS and shows that we will be able to locate the rocket upon

landing no matter the orientation.

Table 6.1.2f

Test Name Basic altimeter function test (Recovery test 1)

Description This test demonstrates that the Stratologger CFs will be able to measure the

altitude that the rocket will meet and retrieve that data post flight

Rationale Every component in the recovery system must be thoroughly tested.

Verifying the function of the altimeters is important before launching

Environmental

Requirements

No environmental requirements

Procedure 1. Connect resistors to main and drogue terminals

2. Connect oscilloscope probes across the resistors

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3. Power the altimeter

4. Place altimeter and battery in container

5. Depressurize container to simulate flight

6. Observe the oscilloscope readings

7. Download flight data produced by the simulation

Required

Resources ❏ Two high power 2.5 ohm resistors

❏ StratologgerCF

❏ Two new 9V batteries

❏ Oscilloscope

❏ Altimeter data cable

❏ Long wire leads

❏ Plastic container large enough to hold altimeter and batteries with a

conduit for leads and another for depressurization

❏ Vacuum for depressurization

Success Criteria ● The altimeters fire both drogue and main charges at the appropriate

altitudes

Proof of Success ● Multiple altimeter flight logs showing expected performance.

Results Multiple iterations of this test consistently produced the expected results of

the drogue and main charges being fired at apogee and the preset altitude

respectively. The image below shows the logged results of the 17th test

performed. As can be seen, the drogue was fired at apogee and the main

was fired at an altitude of 550 ft on descent.

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Analysis The results of this test have verified that the altimeters reliably function as

expected. We are confident after performing these tests that the altimeters

will function correctly.

Lessons

Learned

While performing this test, we discovered the importance of slowly

pressurizing the container after simulated apogee was reached. This more

closely simulates a real flight. Re-pressurizing the container too quickly

could cause the altimeters to appear to fire too late because of of the high

rate of pressurization as compared to the sampling rate of the altimeters.

Table 6.1.2g

Test Name Electric match test (Recovery test 2)

Description This test demonstrates that our choice in batteries for the altimeters will

sufficiently activate the electric matches in order to excite lift off

Rationale Every component in the recovery system must be thoroughly tested.

Demonstrating that the battery can ignite three times the amount of matches

that it would normally have to will satisfy this rigor of testing by a

desirable margin

Environmental

Requirements ❏ Test must be performed outdoors on a non-flammable surface

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Procedure 1. Put on appropriate PPE including safety glasses before beginning

test

2. Acquire a new 9V battery and use multimeter to confirm its voltage

is at least 9 volts.

3. Make two groups of three e-matches.

4. Twist the leads of each group together, separating the halves to lead

into the positive and negative terminals.

5. Wire normally into the altimeter as shown in Figure 3.3.3a, one

group of three to the main, one group of three to the drogue.

6. Depressurize the altimeter to simulate a flight

7. Check to see if all three matches fire normally. At least two need to

fire per terminal block.

8. If there are any matches that fail to fire, test them using a new

battery. Do not wire to the altimeter, just touch to the battery’s

terminals. If the e-match fires, this indicates that a poor battery

choice was made. If the e-match fails to fire, it is defective.

Required

Resources ❏ Six electric matches

❏ StratologgerCF

❏ Two new 9V batteries

❏ Plastic container large enough to hold altimeter and batteries with a

conduit for matches and another for depressurization

Success Criteria ● All six electric matches wired to the altimeter fire during the

simulated flight

Proof of Success ● Image proof of all six matches igniting within one test flight

Results During this test the altimeter was able to fire all six electric matches. A

picture of the matches firing is below.

Analysis Because the altimiters were able to fire three electric matches we are able

to conclude that they will be able to fire the two that we are using for each

black powder charge.

Table 6.1.2h

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Test Name Altimeter test with electromagnetic interference (Recovery test 3)

Description This test verifies that the altimeters work as designed while the transceiver

operates its flight procedure

Rationale While initial tests of the conductive coating demonstrate that it shields an

area from the electromagnetic emissions from the transceiver, the recovery

system must be tested with the GPS running to ensure that it has been

properly shielded.

Environmental

Requirements

This test has no environmental requirements

Procedure 1. Set up avionics bay as if for flight

2. Insert LEDs into altimeters instead of charge leads

3. Activate GPS and check ground control to ensure that transmissions

are being received

4. Tape over all but one vent hole on avionics bay

5. Apply a vacuum to the uncovered hole

6. Inspect both altimeter logs

Required

Resources ❏ Fully assembled avionics bay

❏ Four LEDs

❏ Altimeter data cable

❏ vacuum

Success Criteria ● Flight simulation is nominal, both charges fire/LEDs burn out

Proof of Success ● Altimeter test flight data and plot showing charges fired

Results The altimeters functioned as expected during the simulated flight while the

GPS system was turned on and transmitting. The results and test setup

(removed from avionics bay) are shown below.

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Analysis The results of this test shows that the operation of the GPS system will not

135

prevent the altimeters from functioning properly. Therefore, we are

confident that we will be able to operate the GPS without compromising

mission safety.

Table 6.1.2i

Test RF Shielding Tests (Recovery test 4)

Description Our goal for this test was to see how the RF shielding affected the RSSI of

the transceivers.

Rationale Our reason for conducting these tests is to see that the RSSI dropped when

RF shielding is applied. This will ensure that the RF coating provides

protection to the recovery side of the avionics sled.

Environmental

Requirements

Space with accommodation of roughly 30 ft or more

Procedure 1. Create Matlab Script

2. Move transceivers 30 feet apart

3. Get baseline RSSI data including 20 samples without any obstructions

4. Get 20 RSSI data samples with flight transceiver placed on RF coated

sled and the sled between the two transceivers

5. process data and compare using script

Required

Resources ❏ MATLAB

❏ Transceivers

❏ RF Shield avionics sled

Success Criteria ● RSSI is reduced by the presence of RF coating

Proof of Success ● MATLAB processing script and produced plot showing a drop in

RSSI when the RF shielding is used.

Results The test data was processed using the matlab script below and the the

following plot was produced.

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Analysis In the plot above it can be seen that the RSSI is higher for the unshielded

test (blue line) than it is for the shielded test (red line). This indicates that

eh RF coating reduces the strength of the signal. It is important to not that

because RSSI is a non-linear measurement, the difference is more

significant than it appears to be.

137

Table 6.1.2j

Test Avionics Bay PCB Testing Procedure (System test 1)

Description This provides a detailed plan to test the functionality of the PCB and all its

components for use in the avionics bay of the launch vehicle.

Rationale This test is required to ensure the PCB safely and effectively integrates the

teensy, gps, transceiver, and motor driver into one compact board.

Environmental

Requirements

This test has no environmental requirements.

Procedure 1. Put on appropriate PPE including safety glasses.

2. Solder all components onto the board according to the PCB

schematic, checking connectivity with a multimeter as each

component is added to the board.

3. Ensure the 5V regulator circuit is functioning properly.

a. Using a multimeter, measure the voltage difference between

the input of the linear voltage regulator and ground. It

should be about 7.4 Volts depending on the charge of the

LiPo battery.

b. Again using a multimeter, measure the voltage difference

between the output of the linear regulator and ground. It

should read 5 Volts.

c. Measure the voltage between ground and Vin of each

component. The multimeter should read 5 Volts for the

LoRa, GPS, and Teensy and about 7.4 Volts for the motor

driver.

4. Upload a simple blink LED sketch to the teensy to confirm that it is

functioning properly.

5. Once it is confirmed that the teensy is operating as expected, upload

the flight code.

6. Using the ground system, confirm that the flight computer is

transmitting GPS data as expected.

7. Finally, use the ground control GUI to test the rover release to

ensure that the motor driver is working properly.

8. Once these steps are taken and operation is confirmed, the PCB is

ready to be mounted onto the avionics sled.

138

Required

Resources ❏ Eagle Schematic of PCB

❏ PCB

❏ Adafruit Ultimate GPS

❏ Adafruit LoRa

❏ Pololu DRV8880 Stepper Motor driver

❏ Female lipo connector

❏ 5V Linear Regulator

❏ 10 μF Capacitor

❏ 100 μF Capacitor

❏ 0.1M Capacitor

❏ Multimeter/Voltmeter

Success Criteria ● All components power on

● LoRa initializes

● GPS is sending and receiving data

● Motor driver is functioning

Proof of Success ● Images of completed PCB with voltage measurement

● Use of PCB for further testing

Results The PCB was built and tested successfully, all voltage measurements were

consistent with the expected values. Images of the completed PCB and the

voltage reading from the output of the 5 Volt linear regulator are below.

The PCB was used to perform further testing including the rover release

test outlined in Table 6.1.2m.

Analysis After completing construction and testing of the PCB we have determined

that it is functioning as expected. By using the PCB we have streamlined

all of our non-recovery systems into one compact board. This will increases

the amount of free space in the avionics bay and makes it easier to

troubleshoot any potential problems with the avionics system. The design

and construction of the PCB also allowed us to streamline other testing

procedures such as the rover release test.

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Table 6.1.2k

Test Name Waterproofing test (system test 2)

Description This test verifies that the team derived requirement of waterproofing the

avionics bay is met

Rationale To increase the chances of a successful recovery, the electronics on our

rocket should be safe from water to allow reuse

Environmental

Requirements ❏ This test must be performed in an area where water can be drained

safely away from any electronics not involved in the test

Procedure 1. Generously apply a layer of silicone conformal coating to the top

side of a spare altimeter and a spare copy of the PCB. Avoid

interfering with the silver barometer and serial data port on the

altimeter

2. Allow the conformal coating to set

3. Thoroughly wet both components under a faucet

4. Wipe down the block terminals and bottom side of the altimeter to

avoid short circuiting the battery

5. Perform a simple flight simulation test on the altimeter using LEDs

as substitutes for charges

6. Perform a full PCB function test

Required

Resources ❏ Spare altimeter

❏ Spare PCB

❏ Silicone conformal coating

❏ Two LEDs

❏ Altimeter data cable

Success Criteria ● Both charges fire at appropriate times during simulated flight

● PCB functions properly

Proof of Success ● Tests confirming the proper function of both the PCB and

altimeters after the waterproofing test has been performed.

Results The altimeters performed as expected and were able to fire the charges at

the appropriate time during a simulated flight. The PCB however was

unable to function properly after the waterproofing test was performed.

Analysis Successfully tested altimeters have received the silicone coating and been

re-tested to ensure that they still function. Each of them passed the retest.

The reason for the PCB failure after testing is unknown however due to this

failure we will not be waterproofing the PCB.

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Table 6.1.2l

Test Launchpad delay test (system test 3)

Description This test determines whether our system can function properly after a 2

hour maximum launchpad delay

Rationale There is potential for there to be a significant delay between setting up the

rocket at launch and actually launching it. If the rocket is unable to function

after this delay, recovery will become very difficult or even impossible.

Environmental

Requirements

This test has no environmental requirements.

Procedure 1. Prepare avionics systems for launch

2. Power the systems as if they were set up on the launch pad but not

cleared to launch yet.

3. Turn the safety switches to engage arm the altimeters.

4. Begin GPS transmission with the ground control system.

5. Measure voltage to the stepper motor driver and from the 5 Volt

regulator.

6. Measure the voltage of both altimeter batteries.

7. After two hours perform a payload release test to ensure that the

battery has not been drained so much as to prevent the payload from

being released.

8. Perform basic altimeter testing procedure to ensure the altimeters

are still able to fire the black powder charges.

9. Repeat steps 5 and 6 and record results.

Required

resources ❏ GPS system

❏ Two altimeters

❏ Two 9 Volt batteries

❏ Avionics bay

❏ vacuum

❏ payload

❏ payload release system

❏ 7.4 Volt LiPo

❏ ground control system

❏ ground control GUI for rover release

❏ multimeter

Success Criteria ● Teensy, LoRa, and GPS are each receiving 5 Volts.

● Rover release is successful

● Altimeters successfully complete simulated flight

Proof of Success ● Voltage measurements taken before and after and visual

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confirmation of rover release.

● Confirmation of successful simulated flight based on audio from

altimeters.

Results The LiPo voltage dropped from 8.28 V to 5.65 V and the rover release was

unsuccessful. The 9 Volt recovery batteries started with a voltage of 9.4 V

and dropped to 9.28 V. The altimeters successfully fired the main and

drogue charges during its simulated flight.

Analysis It was determined that the stepper motor was drawing power for the

entirety of the test due to its idle torque. This drained the battery and

prevented a successful release of the payload. To address this problem we

will be purchasing the a stepper motor that does not draw power when it is

not in operation.

Although the stepper motor failed to release the payload, the GPS system

was still receiving 5 volts and still transmitting which would have allowed

us to retrieve the rocket. Most importantly the altimeters performed as

expected, ensuring that the airframe would have been safely recovered after

launch, even with a two hour launch delay.

Table 6.1.2m

Test Name Payload release test (system test 4)

Description This test verifies that the payload is able to be released from the rocket

using the ground control GUI.a

Rationale In order for the payload to perform its designated task, it must first be

released from the rocket. To do this a command must be sent to the flight

computer to engage the payload release stepper motor. This complicated

process is being tested to ensure that the rover can be released without any

unexpected complications.

Environmental

Requirements ❏ This test must be performed on the ground to prevent the rover from

falling after being deployed from the airframe.

Procedure 1. Power on the flight and ground control systems.

2. Open the ground control GUI on a laptop and connect the ground

control system.

3. Begin transmitting between the flight system and the ground

system.

4. Toggle the rover release safety button on the GUI.

5. Press release button on the GUI.

6. Ensure that the rover fully exits the payload section of the airframe.

Required ❏ Flight PCB

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Resources ❏ Ground control

❏ Ground control GUI

❏ Payload release system

❏ Payload

❏ Laptop

Success Criteria ● Rover is successfully released without any damage to the rover or

the airframe.

Proof of Success ● Images of test and visual observation

Results The rover was successfully released from the payload section of the rocket.

A photo of the rocket leaving the tube is below.

Analysis The rover release was performed as expected with the rover leaving the

payload section without damage occurring. We expect the system to

perform the same on launch day.

6.1.3 Payload Tests

The following tests were conducted to ensure proper operation of the various subsystems of the

rover and the retention system, and to reveal any flaws in the design that would need to be

corrected before finalizing the design.

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Table 6.1.3a

Test Orientation Detection and Correction Test

Description Test the ability of the rover to sense orientation and flip itself if it is upside

down.

Rationale There is a possibility that the rover is deployed upside down after landing.

The rover needs to be able to reorient itself in order to drive and collect

soil.

Environmental

Requirements

There are no special environmental requirements for this test. A flat surface

on which the rover can be placed is adequate

Procedure 1. Attach the microcontroller and the accelerometer to the frame of the

rover.

2. Ensure that the actuating arm is securely attached to the servos.

3. Upload the program to detect orientation and send information

through serial to the computer

4. Turn the rover upside down and verify that the accelerometer

readings are reasonable

5. Let the rover flip itself

Required

resources

Rover frame with servos and actuating arm, microcontroller, accelerometer

module, computer with the Arduino IDE, USB cable

Success Criteria Test is successful if the rover can correct its orientation by using its

actuating arm

Proof of Success The rover was able to flip successfully every time the test was performed.

Table 6.1.3b

Test Obstacle Avoidance Test

Description Test the ability of the rover to detect and avoid obstacles while driving

Rationale The rover is expected to encounter obstacles while driving and needs to be

able to navigate without getting stuck

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Environmental

Requirements

An open area on which the rover can drive on that has simulated obstacles

Procedure 1. Upload the final program to the rover

2. Attach the battery connection and turn on safety switch

3. Ensure the rover starts up properly

4. Send test signal to rover from ground control to ensure reliable

wireless connection

5. Prepare driving area by placing simulated obstacles such as boxes

or chairs

6. Place rover on the driving area

7. Send activation signal to the rover so it begins driving

8. Observe driving behavior

Required

resources

Fully assembled rover, ground control microcontroller and computer,

driving course with obstacles

Success Criteria Test is successful if the rover can drive while avoiding obstacles

Proof of Success The rover was able to avoid most obstacles but sometimes struggled with

smaller obstacles

Notes We expect the rover to still navigate the launch field successfully since it

can drive over grass and most hills

Table 6.1.3c

Test Transceiver Communication Test

Description Test the ability of the rover and ground control to maintain a reliable radio

communication link

Rationale The rover needs to be activated remotely after deployment and it also needs

to send the end of mission signal back to ground control

Environmental

Requirements

There are no special environmental requirements for this test.

Procedure 1. Upload the test program to the rover and the ground control

microcontrollers.

2. Ensure that the transceiver is wired properly on both ends

3. Start monitoring the serial monitor on both computers

4. Send test signal to the rover

5. Verify that a reply is received

6. Verify that the serial monitor is still working and the rover is still

able to receive and send messages

Required Rover with the microcontroller and transceiver, ground control

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resources microcontroller with transceiver, computer with Arduino IDE

Success Criteria Test is successful if the rover and ground control can maintain a

communication link without hanging up or resetting unintendedly.

Proof of Success Verification of successful, continued communication through the serial

monitor on both the rover and ground control

Table 6.1.3d

Test Ejection Charge Endurance Test

Description Test the ability of the rover to survive the main parachute ejection charge

going off in close proximity

Rationale The ejection charge is stored in the same compartment as the payload so

the payload is exposed to the hot gases of the flash from the black powder

going off. A successful deployment and operation depends on the rover’s

ability to stay intact after the charge is ignited.

Environmental

Requirements

The test should be performed in an outdoor environment away from groups

of people.

Procedure 1. Assemble the payload and retention system

2. Run shock cord form the main parachute through the guide and

bulkhead hole to the eyebolt

3. Attach ejection charge to wires leading out of vent holes from the

avionics bay

4. Fold parachute and attach the forward section and nose cone

5. Insert shear pins

6. Set up the assembly in a safe spot

7. Set off charges by connecting leads to 9V batteries

8. Disassemble payload assembly from airframe and assess the

condition of the rover

Required

resources

Payload assembly, Forward section and nose cone, main parachute,

Ejection charge, 9V battery

Success Criteria Test is successful if the rover is fully intact after the ejection charges are

fired.

Proof of Success The rover was unharmed and functioned normally after the test.

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Table 6.1.3e

Test Soil Collection Test

Description Test the ability of the rover’s wheels to collect soil

Rationale The rover’s mission after deployment is to travel 10 feet and collect soil.

Environmental

Requirements

The test is to be performed outdoors where the ground resembles typical

soil conditions at the launch site.

Procedure 1. Upload test code to the rover

2. Place rover on the ground where it is free to drive in either direction

3. Turn on safety switch to power on the rover

4. Wait for the rover to complete collecting soil

5. Examine the contents of the wheel

Required

resources

Fully assembled rover, computer with Arduino IDE, USB cable to upload

program

Success Criteria Test is successful if the rover can collect at least 10 ml of soil across its two

wheels

Proof of Success The rover was able to collect dry and soft soil. Wet mud and larger clumps

proved more difficult and required more digging time to collect any

significant amount

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Notes Test was partially successful and success was highly dependent on the soil

condition at the site.

6.2 Rules Derived Requirements

Table 6.2a

Requirement Verification

General requirements

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).

Demonstration was used to verify that

students on the team did 100% of the project

by recording all members involved any given

task.

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

assignments, STEM engagement events, and

risks and mitigations.

This was demonstrated with the Gantt charts

provided in previous reports, the recorded

work of the systems team, the personnel

hazard analysis, and the failure modes and

effect analysis.

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 certain

activities.

All Foreign National Team Members have

filled out the appropriate paperwork and

have been identified in the PDR

The team must identify all team members

attending launch week activities by the Critical

Design Review (CDR). Team members will

include:

● Students actively engaged in the

project throughout the entire year.

● One mentor

● No more than two adult educators.

All PRT members actively engaged in team

activities starting September 2018 through

the CDR, our mentor Duane Wilkey, our

advisor Matthew Barry were recorded as

being a part of the Pitt Rocketry Team from

2018-2019 by the CDR.

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The team will engage a minimum of 200

participants in educational, hands-on science,

technology, engineering, and mathematics

(STEM) activities, as defined in the STEM

Engagement Activity Report, by FRR. To

satisfy this requirement, all events must occur

between project acceptance and the FRR due

date and the STEM Engagement Activity

Report must be submitted via email within

two weeks of the completion of the event.

As described in our proposal, through

collaboration with Pitt’s Society of Physics

Students, PRT presented to 1 school about

the technical information regarding rocketry

as well as opportunities in STEM fields. At an

event in collaboration with the Society of

Women Engineers, PRT introduced

elementary school students to rocketry and

other topics in STEM. Through these 2 events,

PRT presented to 250 students.

The team will establish a social media

presence to inform the public about team

activities.

Our team’s Instagram account can be found

at

https://www.instagram.com/pittrocketryteam/

where we post about team activities.

Teams will email all deliverables to the NASA

project management team by the deadline

specified in the handbook for each milestone.

In the event that a deliverable is too large to

attach to an email, inclusion of a link to

download the file will be sufficient.

All team deliverables have been sent to NASA

by the deadline as requested.

All deliverables must be in PDF format. All deliverables are in PDF format

In every report, teams will provide a table of

contents including major sections and their

respective sub-sections.

Table of Contents provided as seen in

beginning of report

In every report, the team will include the page

number at the bottom of the page.

See bottom of page

The team will provide any computer

equipment necessary to perform a video

teleconference with the review panel. This

includes, but is not limited to, a computer

system, video camera, speaker telephone, and

a sufficient Internet connection. Cellular

phones should be used for speakerphone

capability only as a last resort.

PRT will reserve all necessary space and

computer equipment to teleconference with

the review panel prior to each design review.

All teams will be required to use the launch

pads provided by Student Launch’s launch

services provider. No custom pads will be

permitted on the launch field. Eight foot 1010

PRT has not and will not create a custom pad

for the launch. The launch vehicle is

compatible with the launch pad provided by

the Student Launch’s launch services

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rails and 12 foot 1515 rails will be provided.

The launch rails will be canted 5 to 10 degrees

away from the crowd on launch day. The exact

cant will depend on launch day wind

conditions.

provider.

Each team must identify a “mentor.” A mentor

is defined as an 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 teams he or she supports. The

stipend will only be provided if the team

passes FRR and the team and mentor attend

launch week in April.

The PDR demonstrates that Duane Wilkey, a

level 3 certified NAR member is our team’s

mentor. Duane possesses evidence of the

necessary requirements to assist our team as

the designated mentor.

Vehicle requirements

The vehicle will deliver the payload to an

apogee altitude between 4,000 and 5,500 feet

above ground level (AGL). Teams flying below

3,500 feet or above 6,000 feet on Launch Day

will be disqualified and receive zero altitude

points towards their overall project score.

The mechanical subteam has designed the

rocket to reach an apogee of 4,750 feet, and

the result will be demonstrated by the

readout of the altimeters during test flights

and on launch day.

Teams shall identify their target altitude goal

at the PDR milestone. The declared target

altitude will be used to determine the team’s

altitude score during Launch Week.

Our target altitude is 4,750 feet.

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The vehicle will carry one commercially

available, barometric altimeter for recording

the official altitude used in determining the

Altitude Award winner. The Altitude Award

will be given to the team with the smallest

difference between their measured apogee

and their official target altitude on launch day.

The avionics bay contains two Stratologger

Altimeters that will be used to record the

altitude.

Each altimeter will be armed by a dedicated

mechanical arming switch that is accessible

from the exterior of the rocket airframe when

the rocket is in the launch configuration on

the launch pad. Each altimeter will have a

dedicated power supply.

Inspection has shown that our recovery

system has been designed including these

specifications.

Each arming switch will be capable of being

locked in the ON position for launch (i.e.

cannot be disarmed due to flight forces).

Our choice in arming switches adheres to this

guideline.

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.

There are no expendable components on the

vehicle. The vehicle is re-armable following a

launch.

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 main vehicle

using its own parachute.

The PRT-01 consists of three sections:

A nose cone, the body tube ( which contains

the avionics bay) , and the booster section

Coupler/airframe shoulders which are located

at in-flight separation points will be at least 1

body diameter in length.

The lengths of the airframe shoulders were

measured and compared to that of their

respective diameter and it was found that the

lengths are indeed at least their diameter.

These values can be seen in figures 3.2.4a-e.

Nosecone shoulders which are located at in-

flight separation points will be at least ½

body diameter in length.

The length and diameter of the nose cone

shoulder were measured and it was proved

that the length was at least ½ its body

diameter. These values can be seen in figure

3.2.4d.

The launch vehicle will be limited to a single

stage.

The only motor used is a single stage

refuelable motor.

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The launch vehicle will be capable of being

prepared for flight at the launch site within 2

hours of the time the Federal Aviation

Administration flight waiver opens.

The rocket is capable of going through

preflight preparations within two hours.

The launch vehicle will be capable of

remaining in launch-ready configuration on

the pad for a minimum of 2 hours without

losing the functionality of any critical on-

board components.

All sensitive (namely electronic) components

were left idle in flight configuration for a

minimum of two hours to ensure that the

launch vehicle is capable of remaining in

launch-ready configuration on the pad for a

minimum of 2 hours without losing the

functionality of any critical on-board

components.

The launch vehicle will be capable of being

launched by a standard 12-volt direct current

firing system. The firing system will be

provided by the NASA-designated launch

services provider

The motor in use is able to be launched with

a standard 12-volt DC firing system. This was

verified by inspection.

The launch vehicle will require no external

circuitry or special ground support equipment

to initiate launch (other than what is provided

by the launch services provider).

The motor in use requires no external circuitry

or special ground support to be launched.

This was verified by inspection.

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).

The motor used is a commercially available

APCP fueled motor. APCP purchased was

certified by the NAR or TRA.

Final motor choices will be declared by the

Critical Design Review (CDR) milestone. Any

motor change 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. A

penalty against the team’s overall score will

be incurred when a motor change is made

after the CDR milestone, regardless of the

reason.

Motor has already been chosen through

research, and has not changed since the CDR

milestone

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Pressure vessels on the vehicle will be

approved by the RSO and will meet the

following criteria:

● The minimum factor of safety (Burst or

Ultimate pressure versus Max

Expected Operating Pressure) will be

4:1 with supporting design

documentation included in all

milestone reviews.

● Each pressure vessel will include a

pressure relief valve that sees the full

pressure of the tank and is capable of

withstanding the maximum pressure

and flow rate of the tank.

● 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.

The final rocket design does not utilize a

pressure vessel.

The total impulse provided by a College or

University launch vehicle will not exceed 5,120

Newton-seconds (L-class).

The chosen motor has an impulse below

5,120 Ns

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.

Masses within the launch vehicle and fin

surface area were adjusted as necessary

throughout design process to ensure stability

margin is greater than 2.0

The launch vehicle will accelerate to a

minimum velocity of 52 fps at rail exit.

The motor was chosen to ensure that the

launch vehicle will accelerate to a velocity

greater than 52 fps at the point of rail exit.

All teams will successfully launch and recover

a subscale model of their rocket prior to CDR.

Subscales are not required to be high power

rockets.

A subscale rocket has been manufactured

and launched prior to the CDR deadline.

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.

The subscale rocket has same design as the

full-scale model but smaller to ensure the

subscale performs as similarly as possible to

the full scale rocket. The subscale rocket is

not full-scale size.

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The subscale model will carry an altimeter

capable of recording the model’s apogee

altitude.

Two Perfectflite StratoLoggerCF altimeters

were used on the subscale model.

The subscale rocket must be a newly

constructed rocket, designed and built

specifically for this year’s project.

Our team has completed construction of the

subscale rocket that launched on Saturday

January 5th .

Proof of a successful flight shall be supplied in

the CDR report. Altimeter data output may be

used to meet this requirement.

This altimeter output from the subscale flight

is included in the CDR.

An FRR Addendum will be required for any

team completing a Payload Demonstration

Flight or NASA required Vehicle

Demonstration Re-flight after the submission

of the FRR Report.

If PRT requires a NASA required Vehicle

Demonstration Re-Flight or a Payload

Demonstration flight, an FRR addendum will

be submitted to NASA after the FRR report.

Teams required to complete a Vehicle

Demonstration Re-Flight and failing to submit

the FRR Addendum by the deadline will not

be permitted to fly the vehicle at launch week.

PRT will complete a Vehicle Demonstration

Re-Flight if necessary, in a timely manner to

ensure FRR Addendum is submitted by the

correct deadline.

Teams who successfully complete a Vehicle

Demonstration Flight but fail to qualify the

payload by satisfactorily completing the

Payload Demonstration Flight requirement

will not be permitted to fly the payload at

launch week.

The Pitt Rocketry Team will complete all tasks

by their deadlines.

Teams who complete a Payload

Demonstration Flight which is not fully

successful may petition

the NASA RSO for permission to fly the

payload at launch week. Permission will not

be granted if the RSO or the Review Panel

have any safety concerns.

The Pitt Rocketry Team will complete all tasks

by their deadlines.

Any structural protuberance on the rocket will

be located aft of the burnout center of

gravity.

All structural protuberances such as fins are

located aft of the center of gravity after

burnout.

The team’s name and launch day contact

information shall be in or on the rocket

airframe as well as in or on any section of the

This information is listed on the fins of the

rocket, verifiable by inspection. Additionally, it

is listed on the bottom of the rover.

154

vehicle that separates during flight and is not

tethered to the main airframe. This

information shall be included in a manner that

allows the information to be retrieved without

the need to open or separate the vehicle.

Vehicle demonstration flight

All teams will successfully launch and recover

their full-scale rocket prior to FRR in its final

flight configuration. The rocket flown must be

the same rocket to be flown on launch day.

The purpose of the Vehicle Demonstration

Flight is to validate 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 the intended lower

altitude, functioning tracking devices, etc.).

Our final, full-scale design of the rocket was

not tested prior to FRR in its final flight

configuration due to numerous weather

cancellations.

The vehicle and recovery system will have

functioned as designed

The vehicle and recovery system has been

thoroughly tested and proven to operate as

designed

The full-scale rocket must be a newly

constructed rocket, designed and built

specifically for this year’s project.

The full scale rocket is a newly constructed

rocket built and designed by PRT for the

NASA 2019 Student Launch competition.

The payload does not have to be flown during

the full-scale Vehicle Demonstration Flight.

The following requirements still apply:

● If the payload is not flown, mass

simulators will be used to simulate the

payload mass.

● The mass simulators will be located in

the same approximate location on the

rocket as the missing payload mass.

Payload is ready to be flown for a vehicle

demonstration flight.

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

The PRT payload is not designed to change

the external surface of the rocket.

155

active during the full-scale Vehicle

Demonstration Flight.

Teams shall fly the launch day motor for the

Vehicle Demonstration Flight. The RSO may

approve use of an alternative motor if the

home launch field cannot support the full

impulse of the launch day motor or in other

extenuating circumstances.

The launch day motor will be used during the

Vehicle Demonstration Flight.

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 the launch

day flight. Additional ballast may not be

added without a re-flight of the full-scale

launch vehicle.

A check will be performed to verify that the

vehicle flown during the full-scale test flight is

in its fully ballasted configuration.

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).

Following the successful completion of the

full-scale demonstration flight, the launch

vehicle and its components will not be

modified without the concurrence of the

NASA Range Safety Officer.

Proof of a successful flight shall be supplied in

the FRR report. Altimeter data output is

required to meet this requirement.

The recovery altimeters will collect and store

flight data in their on-board loggers. The

flight data will be recovered following the

flight for use in the FRR report.

Vehicle Demonstration flights must be

completed by the FRR submission deadline. If

the Student Launch office determines that a

Vehicle Demonstration Re-flight is necessary,

then an extension may be granted. This

extension is only valid for re-flights, not first-

time flights. Teams completing a required re-

flight must submit an FRR Addendum by the

FRR Addendum deadline.

The Vehicle Demonstration Flight was not

completed by the deadline due to weather

cancellations in all driveable distance

launches for the entire month of February

and first weekend in March.

Payload Demonstration Flight

The payload must be fully retained

throughout the entirety of the flight, all

retention mechanisms must function as

The retention mechanism is designed such

that the payload is always supported and

rigid during flight. The deployment stepper

156

designed, and the retention mechanism must

not sustain damage requiring repair

motor will be powered and apply a holding

torque so any of the components don’t move

inadvertently.

The payload flown must be the final, active

version.

The final version of the rover is ready for the

payload demonstration flight.

If the above criteria is met during the original

Vehicle Demonstration Flight, occurring prior

to the FRR deadline and the information is

included in the FRR package, the additional

flight and FRR Addendum are not required.

The payload is ready for the payload

demonstration flight.

Payload Demonstration Flights must be

completed by the FRR Addendum deadline.

No extensions will be granted

The team will ensure that the Payload

Demonstration Flights are completed by the

FRR Addendum deadline.

Vehicle Prohibitions

The launch vehicle will not utilize forward

canards. Camera housings will be exempted,

provided the team can show that the

housing(s) causes minimal aerodynamic effect

on the rocket’s stability.

The vehicle was designed to not contain any

forward canards or camera housings. This was

verified by inspection.

The launch vehicle will not utilize forward

firing motors.

The vehicle design does not utilize forward

firing motors. This was verified by inspection.

The launch vehicle will not utilize motors that

expel titanium sponges (Sparky, Skidmark,

MetalStorm, etc.)

The vehicle design does not utilize that expel

titanium sponges. This is verifiable by

inspection.

The launch vehicle will not utilize hybrid

motors.

The vehicle design does not utilize hybrid

motors. This is verifiable by inspection.

The launch vehicle will not utilize a cluster of

motors.

The vehicle design does not utilize a cluster

of motors. This is verifiable by inspection.

The launch vehicle will not utilize friction

fitting for motors.

The vehicle design does not utilize friction

fitting for motors. This is verifiable by

inspection.

The launch vehicle will not exceed Mach 1 at

any point during flight.

The motor utilized and the overall final design

of our rocket is incapable of producing

enough thrust force to achieve Mach 1.

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Vehicle ballast will not exceed 10% of the

total unballasted weight of the rocket as it

would sit on the pad (i.e. a rocket with and

unballasted weight of 40 lbs. on the pad may

contain a maximum of 4 lbs. of ballast).

The ballasted weight of our rocket design will

be checked prior to launch and made sure

not to exceed 10% of the unballasted weight.

This will be done by calculating and summing

the individual weights of the parts, then

comparing it to the weight expected to be

used for unballasting purposes.

Transmissions from onboard transmitters will

not exceed 250 mW of power

The transceivers and antenna used are

incapable of transmitting a signal of 250 mW

of power.

Excessive and/or dense metal will not be

utilized in the construction of the vehicle. Use

of lightweight metal will be permitted but

limited to the amount necessary to ensure

structural integrity of the airframe under the

expected operating stresses.

Only the desired and appropriate amount of

metal needed for our design is used in our

design. Likewise, dense metal is not used.

Recovery System Requirements

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

permissible, provided that kinetic energy

during drogue-stage descent is reasonable, as

deemed by the RSO.

The recovery system is designed to stage the

deployment of the drogue and main

parachutes, with the main to be deployed at a

lower altitude.

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.

Tests for the recovery systems were properly

tested before the launches.

At landing, each independent section of the

launch vehicle will have a maximum kinetic

energy of 75 ft-lbf.

Appropriate parachute sizes to reduce the

kinetic energy of the rocket below 75 ft-lbf

have been calculated and are used in the

recovery system.

The recovery system electrical circuits will be

completely independent of any payload

electrical circuits.

The recovery system electrical circuits have

been designed to be completely independent

of all other electrical circuits on the vehicle.

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The recovery system was tested and verified

independent of the rest of the vehicle.

All recovery electronics will be powered by

commercially available batteries.

The recovery system design includes only

commercially available batteries.

The main parachute shall be deployed no

lower than 500 feet.

The altimeters were tested to confirm that

they can precisely deploy the main parachute

at a height greater than 500 feet. Analysis of

the flight logs will verify that this requirement

is satisfied.

The apogee event may contain a delay of no

more than 2 seconds.

The altimeters have been tested to confirm

that they can precisely deploy the drogue

within 2 seconds of reaching the apogee.

Analysis of the flight logs verify that this

requirement is satisfied.

The recovery system will contain redundant,

commercially available altimeters. The term

“altimeters” includes both simple altimeters

and more sophisticated flight computers.

The design of the recovery system includes

two Perfectflite StratologgerCF altimeters,

both able to activate the charges for the

parachutes.

Motor ejection is not a permissible form of

primary or secondary deployment.

The motor will not be ejected during flight.

Removable shear pins will be used for both

the main parachute compartment and the

drogue parachute compartment.

Removable shear pins are used in the design

of the parachute compartments.

Recovery area will be limited to a 2,500 ft.

radius from the launch pads.

The recovery area will be limited to a 2,500 ft.

radius from the launch pad based on the

subscale design and simulations, as well as

initial testing of the rocket.

Descent time will be limited to 90 seconds

(apogee to touch down).

The descent time is under 90 seconds based

on simulations, the subscale design, and

mathematical calculations.

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

All sections of the rocket are to be tethered

to each other, allowing the GPS system in the

avionics bay to satisfy this requirement. The

tethers were chosen to withstand the tensile

forces that may be imposed on them during

flight.

159

Any rocket section or payload component,

which lands untethered to the launch vehicle,

will contain an active electronic tracking

device.

The rocket and payload will be the only two

separated components.

The electronic tracking device(s) will be fully

functional during the official flight on launch

day.

The GPS units will be powered and sending

data at a constant frequency during launch

and recovery. The batteries chosen for the

rocket will have enough power to sustain this.

The recovery system electronics will not be

adversely affected by any other on-board

electronic devices during flight (from launch

until landing).

The recovery system electronics will be in a

separate compartment from all other on-

board electronics.

The recovery system altimeters will be

physically located in a separate compartment

within the vehicle from any other radio

frequency transmitting device and/or

magnetic wave producing device.

The recovery system electronics is in a

separate compartment from all other on-

board electronics.

The recovery system electronics will be

shielded from all onboard transmitting

devices to avoid inadvertent excitation of the

recovery system electronics.

The compartment housing the recovery

system electronics is protected with radio

frequency shielding.

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.

Any device that may create enough magnetic

waves to affect the recovery system will be

surrounded with a high permeability metal to

prevent the waves from reaching the recovery

electronics compartment. As of the current

design, it is highly unlikely that this would be

necessary, but proper magnetic protection

can be verified through appropriate testing.

The recovery system electronics will be

shielded from any other onboard devices

which may adversely affect the proper

operation of the recovery system electronics.

Proper testing has verified that the recovery

system will be unaffected by other onboard

devices.

Payload Experiment Requirements

Each team will choose one experiment option

from the following list.

● Option 1: Deployable Rover/Soil

The team has built a deployable rover that

will recover a soil sample after landing

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Sample Recovery

● Option 2: Deployable UAV/Beacon

Delivery

An additional experiment (limit of 1) is

allowed, and may be flown, but will not

contribute to scoring.

The team will not be flying any additional

experiments so there is no verification plan in

place.

If the team chooses to fly an additional

experiment, they will provide the appropriate

documentation in all design reports so the

experiment may be reviewed for flight safety.

The team will not be flying any additional

experiments so there is no verification plan in

place.

Deployable Rover / Soil Sample Recovery Requirements

Teams will design a custom rover that will

deploy from the internal structure of the

launch vehicle.

The team has developed a custom rover and

a deployment mechanism inside the launch

vehicle to deploy the rover upon landing and

meeting deployment conditions.

The rover will be retained within the vehicle

utilizing a fail-safe active retention system.

The retention system will be robust enough to

retain the rover if atypical flight forces are

experienced.

The rover will be threaded into an aluminum

rod and supported on both sides by solid

surfaces. For active retention, the deployment

stepper will be powered on during the flight

and set to hold torque to avoid any

movement.

At landing, and under the supervision of the

Remote Deployment Officer, the team will

remotely activate a trigger to deploy the rover

from the rocket.

The avionics bay will be able to receive a

remote signal to deploy the rover.

After deployment, the rover will

autonomously move at least 10 ft. (in any

direction) from the launch vehicle. Once the

rover has reached its final destination, it will

recover a soil sample.

The rover is designed to carry out these

actions which will be verifiable by inspection.

The soil sample will be a minimum of 10

milliliters (mL).

The rover soil sample containment device and

mechanism will be operated prior to launch

and the containment device will be inspected

following operation to ensure that a soil

sample of at least 10 milliliters was collected.

This test will be performed a statistically

significant number of times to ensure that the

161

rover reliably collects an adequate sample

size.

The soil sample will be contained in an

onboard container or compartment. The

container or compartment will be closed or

sealed to protect the sample after collection.

The rover soil sample containment device and

mechanism will be operated prior to launch

and the containment device will be inspected

following operation to ensure that the device

closed or sealed to protect the sample as

intended.

Teams will ensure the rover’s batteries are

sufficiently protected from impact with the

ground.

Frame of the rover is constructed with Nylon

which has a high impact tolerance, and the

batteries are secured to the frame using Dual

Lock to ensure that the batteries remain

secure during flight and upon landing.

The batteries powering the rover will be

brightly colored, clearly marked as a fire

hazard, and easily distinguishable from other

rover parts.

The battery on the rover are covered in a

bright non flammable material and marked

with warning signs indicating a fire hazard.

Safety Requirements

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.

The team has developped a launch and safety

checklist included in the FRR report and used

in the Launch Readiness Review and any

launch day operations.

Each team must identify a student safety

officer who will be responsible for the

following requirements:

● Monitor team activities with an

emphasis on Safety during: Design of

vehicle and payload, Construction of

vehicle and payload, Assembly of

vehicle and payload, Ground testing of

vehicle and payload, Subscale launch

test(s), Full-scale launch test(s), Launch

day, Recovery activities, STEM

Engagement Activities

● Implement procedures developed by

the team for construction, assembly,

launch, and recovery activities.

Thomas Sullivan Harrington has been

identified as the student safety officer and will

perform the listed requirements.

162

● Manage and maintain current

revisions of the team’s hazard

analyses, failure modes analysis,

procedures, and MSDS/chemical

inventory data.

● Assist in the writing and development

of the team’s hazard analyses, failure

modes analysis, and procedures.

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 does not give explicit or

implicit authority for teams to fly those

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.

The team will not launch the vehicle designed

for the NASA Student Launch at any NAR or

TRA launch unless allowed by the local

President or Prefect and RSO. If the team

wishes to launch the vehicle at any NAR or

TRA launch, a member will contact the

President or Prefect and RSO for permission.

Teams will abide by all rules set forth by the

FAA.

The team has read the rules set forth by the

FAA and ask any necessary questions to

ensure that the rules are fully understood.

6.3 Team Derived Requirements

Table 6.3a: Updated and Reviewed Team Derived Requirements

Requirement Verification Procedure Adherence and Results

Vehicle:

Avionics bay must be

easily accessible.

Bulkhead at access point will be

removable such that the avionics

bay can be removed from launch

vehicle.

The bulkheads are removable

which creates an easy access

point for the avionics bay.

Additionally, the team has

practiced removing the

avionics bay, and the process

averages 5 minutes.

Electronics in the avionics

bay must be protected

from water which could

Test the water resistance of a

spare altimeter and PCB treated

with silicone conformal coating

Altimeters were tested and

verified as being resistant to

water, however the spare PCB

163

permanently damage

flight-critical hardware.

by performing standard flight

tests on them post treatment.

that received the silicone

treatment failed to pass basic

functional testing. The reason

for this is under investigation.

Outside of airframe must

be smooth.

Airframe will be sanded and a

clear coat will be added on top of

sticker used to identify rocket.

The airframe was painted and

had two layers of clear coat

added over our team name

and sponsor stickers. The

airframe was sanded before

clear coat was added and is

sufficiently smooth.

Fins must be properly

spaced and attached.

Create and utilize a jig for fin

attachment.

A laser cut fin jig was made

and used during fin

application. The fins are

properly oriented and

securely attached.

Recovery:

Shear pins must break

during recovery stage of

flight at parachute

deployment.

Simulation and testing confirm

that the black powder is able to

appropriately break the shear

pins.

Due to inclimate weather, we

were unable to fly our full

scale. However, shear pins

performed properly during

ground ejection charge

testing.

Parachute must not get

tangled to ensure the

recovery system operates

successfully.

Parachute ejection tests and

simulations of parachute

placement will verify that the

parachute does not tangle during

recovery. Swivels will be used to

minimize parachute and shock

cord entanglement.

We have not yet been able to

fully test parachute

deployment. However,

swivels are prepared and we

have practiced parachute and

shock cord folding to

minimize the risk of

entanglement.

Payload:

Rover must be rigid when

held in its enclosure.

Extensive testing will confirm

that the rover stays secured under

various loads. A mechanical rig

will be built for this purpose.

Informal tests to confirm

rover security have been

performed but no formal tests

of rover security have been

performed.

Rover egress must not be

hindered by any launch

vehicle components such

Simulate landings and test rover

deployments. Adjust stepper

motor power and speed until

Rover egress is not hindered

by other components and has

successfully been deployed in

164

as bulkheads or shock

cords.

reliable deployment is confirmed.

testing.

Wheel scoops must work

in a variety of common

soil conditions.

Test and refine scoop design to

make it work in different soil

conditions.

Wheel scoops have been

tested in sand but no other

soil types.

Wheel scoop valves must

open to allow maximum

soil containment for a

given interval volume.

Build various prototypes with

different valves and hinge angles

to choose the best design.

The current hinge angle (the

first angle proposed) is the

optimal angle so it is the final

design.

Obstacle avoidance

system must work

reliably.

Test the rover with simulated

obstacles.

Informal tests confirm the

rover can avoid obstacles.

Team Performance:

Avionics tests are to be

performed by team

members not involved in

their creation to remove

bias from test results.

The avionics lead will assign tests

being conducted to the

appropriate team members to

meet this requirement.

Avionics tests were

performed in groups by

appropriate team members

picked by the avionics lead.

6.4 Budgeting and Timeline

6.4.1 Line Item Budget

Item Vendor Price Quantity Total Price

PR75-2G-W Wildman

Rocketry

$157.99 3 $473.97

StratoLoggerCF Perfect Flite $49.46 6 $296.76

Arduino Mega Arduino $33.00 1 $33.00

GPS Breakout Adafruit $39.95 5 $199.75

Antenna

SREI038-S9P

Mouser $10.11 2 $20.22

Antenna Adaptor Amazon $9.98 2 $19.96

165

HC-12

Communication

unit

Amazon $4.86 12 $58.32

Fire Starters Apogee Rockets $12.83 1 $12.83

Epoxy Apogee Rockets $60.04 1 $60.04

Shear Pins Apogee Rockets $3.22 8 $25.76

Darkstarr Jr Kit Darkstarr $125.99 2 $251.98

Main Chute (72”) Apogee Rockets $55.50 2 $111.00

Drogue Chute

(18”)

Apogee Rockets $59.33 1 $59.33

Ball Bearings for

Subscale

Apogee Rockets $4.86 2 $9.72

Swivels for

Subscale

Apogee Rockets $8.74 2 $17.48

Rivets for

Subscale

Apogee Rockets $3.35 2 $6.70

Fiberglass for

Subscale

Eplastics $69.43 1 $69.43

Shaft Collars McMaster-Carr $15.51 1 $15.51

Motor Adapter Apogee Rockets $31.01 1 $31.01

Ejection Canisters Apogee Rockets $18.00 3 $54.00

HC-SR04

Ultrasonic Sensor

Amazon $5.25 1 $5.25

Rail Button Apogee Rockets $3.35 2 $6.70

433Mhz SI4463

Wireless HC-12

Transceiver

HiLetgo $7.99 3 $23.97

RFM96W LoRa

Radio Trver

Breakout

Adafruit $19.95 4 $79.80

Edge-Launch

SMA Connector

for 1.6mm / 0x062

Adafruit $10.00 4 $40.00

166

uFL SMT

Antenna

Connector

Adafruit $3.00 4 $12.00

Motor Retainer AeroPack $55.56 1 $55.56

Thrust Plate $44.49 1 $44.49

NylonX Spool Matter Hackers $58.00 2 $58.00

Rocket Body Wildman Darkstar $339.00 1 $339.00

Hardened Steel

Nozzle for

Lulzbot TAZ 6

printer

Matter Hackers $22.50 2 $45.00

Kevlar Wildman

Rocketry

$1.25/ft 135 $168.75

Subscale Chute

Protectors

Wildman

Rocketry

$6.95 2 $13.90

Subscale Motor

Retainer

Wildman

Rocketry

$25.00 1 $25.00

Subscale Main

Chute (30”)

Apogee Rockets $14.75 2 $29.50

Subscale Drogue

Chute (18”)

Apogee Rockets $7.49 1 $7.49

HC-12 for

Subscale

HiLetgo $15.98 1 $15.98

Teensy Amazon $29.95 8 $239.60

TOTAL $3,036.76

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Business and Travel

Description Cost

Competition Travel $1850.00

Lodging $4186.00

Outreach $80.00

Emergency Fund $1000.00

Non-Competition Travel $500.00

TOTAL $7616.00

Payload Budget

Description Vendor Cost

DC Gear motor with encoder DFRobot $26.80

L298N motor driver Amazon $5.99

MX 1508 motor driver Amazon $8.65

Arduino Nano clone Amazon $13.86

Ultrasonic sensor Amazon $5.99

7.4V LiPo Battery Amazon $23.98

7.4V LiPo Battery Amazon $14.88

SG90 Servo Motor Amazon $10.99

NEMA 14 Stepper Motor Amazon $14.99

A4988 Stepper Driver Amazon $5.99

270 degree servo motors AliExpress $50.80

168

Flexible Coupling Amazon $6.96

M8 T-nuts Amazon $7.49

M8 Nylon threaded rod McMaster-Carr $17.57

¼” Aluminum threaded rod McMaster-Carr $27.70

¼” nylon smooth rod McMaster-Carr $18.94

⅜” extruded Nylon rod Gamut $8.29

Nylon filament spool Amazon $37.89

Wheels On-hand $0.00

TOTAL $307.73

Description Cost

Vehicle $3036.76

Payload $307.73

Business and Travel $7616.00

TOTAL $10,960.49

6.4.2 Funding Plan

Table 6.4.2a: Funding Sources

Source Amount Funds Allowed for

University of Pittsburgh

Mechanical Engineering

Department

$5000.00 All costs of project

University of Pittsburgh

Swanson School of

Engineering

$5000.00 All costs of project

University of Pittsburgh $5000.00 Competition travel and

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Student Government Board lodging

SolidWorks SolidWorks licenses for team CAD modelling

ANSYS ANSYS licenses for team Simulations

WorkZone Project Management Software

for team

Creating Gantt Charts,

assigning tasks, setting

deadlines

Any further required funds for travel during competition week will be acquired through team

fundraising.