1 EXECUTIVE SUMMARY

1.1 Design Overview

1.2 System Performance

1.3 Design Development

2 MANAGEMENT SUMMARY

2.1 Organization

2.2 Schedule and Planning

3 CONCEPTUAL DESIGN

3.1 Mission Requirements

3.2 Rocket Design Tool

3.3 Rocket Design Concepts

3.4 Configuration Selection

4 PRELIMINARY DESIGN (Rocket 1)

4.1 Design Methodology

4.2 Propulsion System

4.3 Aerodynamics

4.4 Avionics System

5 DETAIL DESIGN (Rocket 1)

5.1 Rocket Dimensional Parameters

5.2 Sub-System Design, Selection, Integration, and Architecture

5.3 Weight and Balance

5.4 Rated Rocket Cost

5.5 Rocket and Mission Performance

5.6 Drawing Package

6 MANUFACTURING PLAN AND PROCESSES

6.1 Manufacturing Figures of Merit

6.2 Construction Method Selection

7 TESTING PLAN

7.1 Sub-System Tests and Objectives

7.2 Flight Testing

8 DESIGN AND TEST ITERATION

8.1 Rocket 3

8.2 Rocket 3 Flight Testing

9 PERFORMANCE RESULTS

9.1 Sub-System Evaluation

9.2 Demonstrated Rocket Performance

10 REFERENCES

1. EXECUTIVE SUMMARY (Julian)

This report describes the design process used by the Purdue University’s Team Galactic to develop a rocket that

could launch anytime within FAA regulation. The goal the design was to be capable of safe launch and recovery

procedures while also capable of recording altitude data during flight.

1.1 Design Overview

The rocket is restricted by FAA regulations to have no more than 4.4 ounce in propellant weight and no more than 3.3

pounds in total weight. In order to meet the FAA requirements, the design takes advantages of the 3D printer,

MakerBot, to produce lightweight and structural sufficient parts. The quick manufacturing process of the MakerBot

meant the rocket could be rapidly iterated between tests. This significantly reduced lag between tests.

The design consist of 6 main subsystems: nosecone, avionics bay, parachute bay, motor bay, motor cap, and fins.

The nosecone is meant to reduce overall aerodynamic. The avionics bay is meant to house the avionics system of

the rocket. The parachute bay is meant to contain the recovery system. The motor bay is meant to house the rocket

motor. The motor cap is meant to contain the rocket motor within the motor bay. Finally, the fins meant to increase

overall stability of the rocket. All of these components were manufactured by the MakerBot.

1.2 System Performance

The focus on the design was to allow the rocket to have a successful launch and recovery operation, while accurately

recording flight data using onboard avionics.

The rocket will launch from a guiding rail to ensure the rocket will fly towards the targeted destination. After 10 secs of

flight time, the ejection charge from the rocket motor will ignite, causing the rocket to separate into two sections in

order to deploy the parachute system. Once the parachute is deployed, the rocket will safely descend back to the

ground while keeping the entire rocket intact. In addition, the onboard avionics will record altitude data during the

entire duration of the flight.

System performance analysis of the system will be done by comparing the onboard altitude data with mathematical

models in order to examine what was the experimental results vs the theoretical results.

1.3 Design Development

The design philosophy was to create a rocket that could be rapidly iterated upon. This design wanted to challenge the

notion that rockets could be created using the Agile development with aspects of the standard Waterfall development

approach. The Waterfall approach focuses on sequence development towards a final product, while the Agile

approach focuses continuous improvement and flexibility towards modification.

In order to follow both the Agile and Waterfall philosophy approach, the design became very subsystem oriented.

Each subsystem was examined through comparative analysis to see which selection would provide the best results

within the design’s performance requirement. However if through testing and a subsystem was found not to be

performing up to expectations, it could also be quickly redesigned with minimal impact on the other subsystems.

2. MANAGEMENT SUMMARY (Julian)

The Purdue University’s Team Galactic is consisted of 5 team members, all of whom are seniors within the

Aeronautical/Astronautical school of engineering. The team was organized into 4 division leads and one program

manager. Due to the nature of having a small team, many aspects of the design were not done individually, but rather

cooperatively between division leads. Therefore a team member’s actual function is not be constrained by their

defined function.

The following figure describes the defined function of Team Galactic’s members.

2.1 Organization

The aerodynamics division is responsible for estimating and measuring the aerodynamic forces necessary on the

system. The avionics division is responsible for creating the avionics system that will be onboard the rocket. The

propulsion division is responsible for rocket motor selection. The systems integration division will be responsible for

sizing and modeling the rocket. The program manager’s role is to aggravate all of the results produced by the other

divisions and to lead the direction of the design.

Each week the program manager will create deliverables for each division that will further progress of the design.

Each division will need the goals of the deliverables by the end of the week. At the end of each week, the division

leads will meet with the program manager in order to discuss weekly results. The program manager will take input

from all of the division heads and then evaluate the direction of the project by handing out new deliverables.

Throughout the week, it is also the role of the program manager to ensure each division will be capable of meeting

the goals of the deliverable.

2.2 Schedule and Planning

The overall schedule from the beginning of September 2014 to December 2014 was developed by the program

manager in conjunction with the division leads. The phases of the design process were depicted by a Gantt Chart.

The following figure shows the Gantt Chart that was used to visual the overall design schedule:

3. CONCEPTUAL DESIGN (Saphal)

3.1 Mission Requirements

The initial motivation of the project was to take part in a competition organized by Experimental Sounding Rocket Association (ESRA) which requires a team to build a rocket that can carry a 10 lb payload up to as close as 10,000 ft. Any altitude above or below 10,000 ft would accrue negative points. Hence, the scoring system was mainly based on the accuracy of the design Due to the limited knowledge on the rocket systems and poor experience with manufacturing process, the team planned to adjust the mission. The revised mission was to design a scaled down prototype. This will allow us to learn about different rocket sub-system and analysis techniques before jumping into the actual competition. The team would be able to test the propulsion system, recovery system and avionics system. Most importantly, the team would be able to gain some experience in manufacturing. Also, the scaled down prototype will require less powerful motor which would not require any kind of rocket certification.

The primary requirement of the revised mission was to verify the MATLAB codes with the actual flight data and optimize the accuracy of the MATLAB codes. Other requirements include successful recovery of the rocket and data acquisition.

3.2 Rocket Design/Analysis Tool

3.2.1 Vehicle Sketch Pad (VSP)

VSP is a geometry modeling tool for conceptual aircraft design. The software rapidly models

aircraft configurations without expending the expertise required for traditional Computer Aided Design (CAD)

packages. The software can also be used to design rockets. Initial designs for concept generation were

produced by using VSP.

3.2.2 SolidWorks

SolidWorks is solid modeling CAD (computer-aided design) software that utilizes a parametric

feature-based approach to create models and assemblies. The software was used to create detailed CAD

designs for manufacturing purposes.

3.2.3 Openrocket

OpenRocket is an open source program that allows the user to design a rocket, evaluate its

performance parameters, and produce a flight history of the rocket. With the tools provided by OpenRocket,

preliminary rocket designs were created. The user can design a rocket by adding nose cones, body tubes,

fins, couplers, rocket motors, etc. The user can also rearrange the configuration and sizes of the

components added to their rocket. Once a rocket motor is added to the model, OpenRocket can also

calculate the estimated apogee altitude that the rocket will reach.

Furthermore, OpenRocket can also provide the CD value of each of the components of the rocket,

as

well as the CD value of the entire rocket. OpenRocket can also evaluate the stability of the rocket by

calculating the mass of each rocket component to locate the center of gravity, calculate center of pressure,

and provide the caliber of stability.

3.2.4 MATLAB Model

The MATLAB model is the computer code used to calculate altitude of the apogee with the input

value of rocket configuration, launch angle and launch site altitude. It used Newtons 2rd Law of motion as

the fundamental mathematical model to calculate acceleration. It use compressibility model to measure the

change in drag coefficient and standard atmosphere code to calculate drag. It assumes an average thrust

curve profile and use the average thrust calculated by the total impulse over burn time. It use ODE45 as a

solver for differential equation. The details of mathematical model will be attached in the appendix.

3.3 Rocket Design Concepts

3.3.1 Nose Cone Design

The desire to develop a comprehensive aerodynamic nose cone is to reduce altitude loss from drag and

weight. Not all nose cones are created. Certain nose cones perform better under different Mach regimes.

Figure 4.1 - Comparison of Nosecone Performance over Several Mach Regimes

(1 is the best performance and 4 is the worst performance)

From Figure 4.1 the optimal nosecone will vary on depending on the flight history of a mission. Since the planned mission is mostly under transonic speeds, the LV-HAACK,, Power Series, or Parabolic nose cone would be the best.

Next, OpenRocket was utilized to do initial aerodynamic calculations.

Figure 4.2 – OpenRocket Analysis of CD Values for Haack Series Nosecone

Figure 4.3 – OpenRocket Analysis of CD Values for Parabolic Series Nosecone

Figure 4.4 – OpenRocket Analysis of CD Values for Power Series (x.5) Nosecone

From the initial OpenRocket nosecone analysis, it is not absolutely clear which nosecone is the best design. To obtain further data for nose cone selection, CFD analysis was done.

CAD model of the Haack Series was built using SolidWorks. Then using ANSYS Fluent, CFD data was

provided the of the LV - Haack analysis.

CFD analysis was done on the LV-Haack nose cone to compare results with OpenRocket. The assumptions made during CFD analysis were that operating conditions were sea level conditions. With these assumptions two analysis for two velocities were made. First analysis was made using 100 m/s which is about 0.3 Mach. Figure below shows that Cd value averages on 0.27 which greatly varies from OpenRocket.

Figure 4.6 - CFD analysis of LV-Haack nose cone (0.3 Mach)

Second analysis was run for 134 m/s which is 0.4 Mach. Figure below shows that Cd average changes significantly and goes up to 0.45.

Figure 4.7 - CFD analysis of LV-Haack nose cone (0.4 Mach)

3.3.2 Fin Design

Fins are essential for stability of the rocket. There are many types of fins differing in sizes and shapes.

Figure 1 shows some of the types of fins. In order to design a fin, many different types of analysis has to be

performed in order to choose the best one. Fins are used to shift center of pressure towards where fins are located,

and damp the roll during the flight. However, fins contribute to the overall drag coefficient of the rocket. Because of

huge number of selection in design, it is best to choose particular type of fin. In this case, delta wing would be the

best choice because of the large total area of the wing, and available resources on performance analysis data.

Figure 1. Types of fins

Fin performance depends on the count of fins, and the geometry; therefore, in order to understand the

performance of the fins comparison between numbers of fins attached is made, and geometry analysis is performed.

Number of fins attached influences roll damping and drag of the rocket. According to figure 2 higher count

of fins reduces the damping in roll coefficient. However, looking at figure 3, although, quantitative conclusion cannot

be made, it is obvious that higher number of fins create higher drag.

Figure 2. Variation with Mach number of damping in roll with delta wing plan forms

Sweep angle of the fins determines the roll damping and drag. From figure 2, it can be seen that higher

sweep angles have lower damping in roll coefficient.

Figure 3. The variation of drag coefficient with Mach number for different fin arrangements

For various Mach numbers of various sweep angles drag coefficient behaves differently. As it can be seen

in figure 4, for low subsonic flight of four fins with sweep angle of 45 degrees has less drag than 60 and 70 degrees.

Assumption that it behaves same for three fins configuration can be made. Starting from transonic speed, both for

three fin and four fin configuration, drag increases as angle decreases.

Figure 4. Variation of total drag coefficient at zero lift with Mach number

3.3.3 Motor Sizing (LEO)

The maximum size of a rocket motor can be installed on a model rocket is Class G, which has a range of total

impulse of 80Ns ~ 160Ns. As we are aiming on the Experimental Sounding Rocket Competition, which has a target

altitude of 10,000 ft. We want to choose the G class motor to maximum our apogee altitude. We compared multiple

Class G motor available in the local market as well as online retailer, the following table will provide details

information:

Table 3.3.3.1 Motor selection

Type Total Impulse

Maximum Thrust

Average Thrust Dimension Price with Shipping

Estes G80-10T 133 Ns 99.4N 77.9N 29*128 mm $24.99+tax (Local)

Estes G40-7W 99 Ns 74.3N 46.3N 29*128 mm $24.99+tax (Local)

Cesaroni G46-11W 127Ns 127.3N 45.4N 29*127 mm $21.99 + $ 7.99

Aerotech G40-7W 99 Ns 74.3N 46.3N 29*124 mm $23.75 + $ 7.99

As shown on the table above, within the constraints of a Class G motor, the motors provide similar performance

specification. Total impulse varies in a range of 99~133 Ns and all of the motors share a same diameter. The local

store provide Estes G80-10T and G40-7W with a reasonable price. Online retailer provided motors with cheaper

price. However, considering about the fact that it is impossible to have a one day shipping with a comparable price

and it is unsafe to store multiple rocket motor at the same time. Choosing the stock of the local store becomes the

best strategy.

To compare the 2 motors supplied by the local store. The G80-10T motor is more powerful than the G40-7W in total

impulse and maximum thrust. In our MATLAB mathematical model, we didn’t include the factor of variable thrust. By

taking comparision of two motor:

Table 3.3.3.2 Motor comparision

Type Maximum Thrust Average Thrust Percentage Difference

Estes G80-10T 99.4N 77.9N 27.60%

Estes G40-7W 74.3N 46.3N 60.48%

As shown on the figure 3.3.3.1, G80-10T motor provides neutral thrust profile which is steady and the maximum

thrust is closer to the average thrust than G40-7W.. Meanwhile, G40-7W has a constantly decrease in the thrust

which indicates it has a regressive thrust profile. This means a G80-10T motor can provide a much closer match of

MATLAB code altitude prediction with real flight condition than the G40-7W motor..

Figure 3.3.3.1 Thrust Curve of Different Motors

G80 was the best choice for the flight. However, during the propulsion subsystem test, the large acceleration

powered by the G80-10T motor exceed the structure limit of the prototype and caused structure failure. As the

maximum thrust of G40-7W motor is 25.3% less than that of G80-10T, and to increase the safety factor as well as the

efficiency of progress, A compromise of the accuracy had been made, and the G40-7W became the final motor

design.

3.3.4 Avionics

In order to obtain the data of the apogee, velocity, and trajectory of the flight, and the location of the rocket

after flight rocket has to have avionics installed on board. There are several essential components that avionics

should include in order to obtain the desired data. Following components are essential for a rocket avionics bay:

1. Arduino

Figure 4.19 Ardupilot chip

On board smart system(computer) which could be used to do altitude control, primary and secondary parachute deployment.

2. GPS

Figure 4.20 on-board GPS

sensor which measures altitude, velocity, acceleration for the rocket. Its feedback data to Ardupilot can be used for altitude control.

3. Altimeter

Figure 4.21 Altimeter

Pressure based sensor that can measure altitude on timely manner. It can be used for altitude control as well.

4. Battery pack

Figure 4.22 3V battery

Arduino, and other components like GPS and altimeter require power supply of 3.3V or 5V. Therefore, avionics bay needs to have batteries that would provide sufficient power to the components.

5. Memory Storage

Figure 4.23 SD card

Memory cards are essential for storing data obtained from parts that transmit data.

4. PRELIMINARY DESIGN (Rocket 1 and 2)

4.1 Design Methodology (Saphal)

The flowchart presented in the fig 4.1.1 shows the design methodology followed by our team. The team

established the mission requirements and performed some design research to get familiarized with different sub-

systems involved in a sounding rocket. Furthermore, the team also researched on multiple ways of manufacturing the

parts for the rocket. After the initial design research, the team was divided to perform different tasks focusing on

different sub-systems. The propulsion system, aerodynamics selection of nose cone and fins, avionics selection and

manufacturing techniques heavily influenced the preliminary sizing of the rocket. Furthermore, the motor selection

and the structural integration predicted the weight of the rocket which is one of the main driving factor for the mission

performance. The mission performance will be compared to mission requirements and we will identify any design

deficiencies and iterate the design with upgraded structural integration and motor selection.

4.2 Propulsion System (Leo)

Based on the altitude output during the concept generation stage, we choose Pro Series G40-7W as our rocket motor

for our design. Here G40 is the type of the motor, and ‘7W’ stand for an ejection charge will fire after 7 second delay.

G40-7W is a very common motor and it has no restriction on purchasing it. It is supplied in local store with ample

amount. The figure below shows the geometry of the motor.

Figure 4.3.1 G40-7 Motor (Amain)

The specification of this motor is presented by the following table.

Table 4.3.1 G40-7W Motor Specification

Name G40-7W

Initial Mass 123 g

Propellent Mass 53.8 g

Final Mass 69.2 g

Total Impulse 99 Ns

Maximum Thrust 74.3 N

Average Thrust 46.3 N

Time of Burn 2.13 s

Dimension* 29mm(D) * 128mm

It has a circular cross section

And a detail thrust profile is provided in the figure 4.3.1

Figure 4.3.1 G40-7W Trust Profile

G40-7W has a peak thrust of 74.3N initially, which can help us to increase the initial acceleration therefore to

increase velocity after launch rail. The ejection charge fires 7s after motor burned out, which gives us 9.13 second to

reach apogee.

4.3 Aerodynamics

Using G - 40 motor, rocket speed would stay at subsonic level. Therefore, looking at figure 4, it can be seen

that for subsonic case, sweep angle of 45 degrees provide least amount of drag.

Figure A. 45 degree delta wing fin

Furthermore, according to figure B, having straight tip reduces tip vortices and simplifies manufacturing

process. Top view of the fin is shown in the figure C.

Figure B. Tip vortices

Figure C. Top view of fin

Also, number of fins was decided to be 3, because at 45 degrees according to figure 3, it creates less drag.

Furthermore, three fins makes it faster to manufacture, and reduces overall weight.

4.4 Avionics System

Due to the small available space in avionics bay it was decided to use altimeter only incorporated with

arduino mini pro. Altimeter is capable of recording accurate data independently from axis. It had a capability to record

altitude versus time data which was essential for verifying calculations. SD card required another platform to record a

data which didn’t fit in avionics bay. Therefore, it was decided to use EEPROM memory of the arduino mini pro which

had 512 bits of memory, and could store 512 integers. Feet is three times more accurate than meters; therefore, code

was recording altitude in feet.

Figure E. Arduino mini with altimeter

Code required a manual input of the altitude at the test site to zero the data, and would start recording the data after it detects that rocket is 20 feet above the ground. EEPROM was recording 2 altitude data points per second. After the launch it was found that 2 data points per second wasn’t enough for high speed. Therefore, avionics needed to record more data points per second, and EEPROM wasn’t appropriate for it.

Because of the limited data, next design of the rocket required more space for avionics to be able to fit the

SD card. On the second design of the avionics bay, it was possible to incorporate SD card to the avionics, and also

GPS receiver. During the design, it was found out that GPS was preventing the code to run properly; therefore, in

order to speed up the results, GPS was removed. Another problem that was encountered, was that altimeter wasn’t

recording proper data when using old batteries because it requires at least 2 V of power to operate. After changing

old batteries to new ones, altimeter started recording the data. Because of the large capacity of the SD card, code

was made to record data for 10 minutes instead of starting after 20 feet. However, code had to be changed after 3rd

test, when it was found out that in the cold weather batteries cannot supply enough power to the avionics, which lead

to data being corrupted. Therefore, next code was opening file in the SD card, transferring the data, and closing it

every time before the next iteration.

5. DETAIL DESIGN (Rocket 1 and 2)

5.1 Rocket Dimensional Parameters (Chris)

The first design was modeled in Openrocket.. This design has the largest motor that can launch without

clearance. The specifications for the rocket are listed in the table 5.1.1

Table 5.1.1: Rocket Design

Length 62.5 cm

Diameter 3.8 cm

Launch Weight 352 g (0.776 lbs)

Total Impulse (G-40) 99 N-sec

Delay Time 7 sec

Avionics Pressure Based Altimeter

Delay time is time between the motor burnout and deployment of the ejection charge. This is built into the

motor. Figure 5.1.1 is the rocket sim model of the configuration of the rocket.

Figure 5.1.1: Flight configuration of the rocket.

From figure 5.1.1 the avionics are nearest the the nose conse. The parachute and the shock cord are

directly below that. During parachute deployment the rocket separates between the avionics and the parachute. The

chord connect the avionics and nose cone to the rest of the body and motor.

5.2 Sub-System Design, Selection, Integration, and Architecture (Saphal)

5.2.1 Coupler

Integrated couplers were used as an interface between two sections. The end of one section would

have the outer walls carved out serving as a male attachment and the end of the receiving section would

have the inner walls carved out serving as a female attachment. The coupler would maintain a tight fit

between the sections. The figure below shows the example of coupler between avionics bay and nose

parachute bay.

5.2.2 Avionics Bay

The avionics bay would be comprised of two semi cylindrical structures; one resting on top of

another. The parts will have rails which can slide over each other. This design will provide us an ability to

work with electronics more freely. It will also enable us to make modification on the electronics without

disintegrating the rocket. The figure below shows the pictorial definition of the sliding rails.

5.2.3 Parachute bay

The parachute bay consists of two section. The first section will serve as a placement for the

parachute and the second section will serve as buffer region for the ejection charge and the flame coming

out of the ejection charge.

5.2.4 Motor Hub

The motor hub serves as a housing for the rocket motor. It consists of motor mount which takes the

structural load from the rocket motor and distributes on the entire body of the rocket. It also consists the

slots for the fins to slide in.

5.2.6 Launch Log and Rail

Standard launch log and the launch rail used in the model rocket was used for launching the rocket.

5.3 Weight and Balance (Chris)

The center of gravity of the rocket is located 40.4 cm from the nose of the rocket. This was generated in

Openrocket. All the dimensions were of the rocket were carefully modeled to ensure the accuracy. The total weight in

Openrocket was 352 gram (0.776lbs). The launch weight of the rocket was measured at 351g (0.774lbs). This check

helped verify Openrocket.

The center of pressure was also calculated in Openrocket as well. The center of pressure was located 49.8

cm. This value It is the equivalent of the aerodynamic center in an aircraft. This is the point where the pitching

moment remains constant independent of angle of attack. This value could not be easily verified by other means.

Static margin is measured as the distance between the center of gravity, Cg and the center of pressure Cp.

It is used a way to measure stability. The Cg is placed in front of the the center of pressure to keep the rocket stable.

For sounding rockets it is common practice to use calibers (number of diameter) to normalize the stability. For this

rocket the static margin is 9.4 cm, or 2.45 calibers. The sounding rocket competition requires 1-2 calibers of static

margin. A slightly higher static margin value was chosen to increase stability and reduce risk.

5.4 Rated Rocket Cost

Rocket

Table 5.5.3 Cost of Rocket

Nose Cone * (Free)

Avionics Bay * (Free)

Parachute Bay * (Free)

Motor Hub/ Fin * (Free)

SparkFun Altitude/Pressure Sensor Breakout -

MPL3115A2

$35.35

Wire Leader $3.49

Pro Series G40 Rocket Motor $26.74

24’’ Diameter Parachute $12.83

2 * CR2032 Battery $2.49

$2.49

Total $83.41

* The 4 parts are entirely made by 3D printer offered in KNOY Hall of Technology Boiler Make Lab which is free up to

200g of printing material each time for student project.

5.5 Estimated Performance

We estimated the rocket performance by using the open source code OpenRocket Simulator, and compare it with our

MATLAB code. These two analysis tools can give us the result based on our input which includes initial total mass,

propellent mass, thrust profile, total impulse, launch altitude. In OpenRocket, we can also get recovery rate from

setting the parachute size and parachute deploy time.

The detail of two mathematical tool comparison is shown in the following plot:

Figure 5.5.1 Predicted Altitude

As it shown, two mathematical tools have very similar curve on the predicted altitude plot. however, there is

difference in between. Which is possibly caused by the difference between thrust profile input.

Figure 5.5.1 First 0.7 second of the altitude prediction

The difference is caused by the uniform thrust profile assumed by the matlab model, but the G40-7W has a

regressive thrust profile. This explained why, matlab code has a lower prediction after 0.3 second. Since it takes

time(even though it is only 0.1 second) for the motor to reach the maximum thrust and then decreases with time

going, MATLAB code predicted a higher altitude in the first 0.3 second.

5.6 Drawing Package (Saphal)

6. MANUFACTURING PLAN AND PROCESSES (Julian)

To manufacture the rocket with the best result and efficiency, it was necessary to select the most appropriate

construction method. The appropriate construction method for the rocket was evaluated with the following

manufacturing figures of merit.

6.1 Manufacturing Figures of Merit

● Mass - The mass of the finished design

● Strength - The durability of the finished design

● Manufacturability - The ease and quality of manufacturing the finished design

6.2 Construction Method Selection

Several construction methods were examined before manufacturing. Machining the rocket parts using metal pieces

was automatically eliminated because of FAA restrictions towards metal pieces in rockets. Therefore the follow

manufacturing methods analyzed were:

● Fiberglass - Using a mold and laying up fiberglass around it

● 3D Printer - Using the MakerBot to print rocket parts with ABS filament as the construction material

● Plywood - Using milling tools to machine parts from plywood

The selection was based on which construction method scored the best based on the figures of merit developed

above. Therefore a weighted decision matrix was created to score which was the best option. However a scoring

rubric had to be created first. The following figure describes how the scoring was done:

Using the scoring rubric above, the following table details the weighted decision matrix used to score the construction

method:

The weighing was skewed towards manufacturability. The reason is because no matter how good something looks on

paper, if it can’t be manufactured then it is not an applicable manufacturing method.

● Fiberglass - The fiberglass provided many benefits. It was both lightweight and extremely strong, so the

mass and strength were rated as strong. However in order to actually manufacture the rocket using

fiberglass was incredible arduous. In order to lay up the fiberglass, the mold of the rocket component was

necessary. Therefore in order to have high precision results with the fiberglass lay up, the mold had to be

created with high precision. In addition, when wetting the fiberglass in preparation for the lay up, the epoxy

has to be uniformly spread out, otherwise uneven distribution in epoxy will cause loss in precision. Finally,

the epoxy had to be harden for at least 12 hours for the best results. The entire process for building just one

component could take up to day. Also, it was incredibly difficult to do fiberglass lay up around geometries

that have a pointed end, such as the nose cone. Therefore the manufacturability for fiberglass was rated as

poor.

● 3D Printer - The major benefit of 3D printing is the ease of manufacturability. The MakerBots were incredibly

easy to work with. All one had to do was create a CAD drawing of the desired part and then export to the

MakerBot to print. The MakerBot was also able to produce parts up to 0.1 mm of precision, reasonable

enough for the scope of the design. Also the MakerBot only took around 4 - 5 hours for each component and

while it is printing, the printer doesn’t require constant supervision. Therefore the manufacturability of the

MakerBot was rated as strong. Finally the MakerBot uses ABS filament for the construction. The density and

strength of ABS filament is sufficient for the scope of the design but it is poor in comparison to the other

design. Therefore the strength and mass was rated as poor.

● Plywood - The major benefits of using plywood is its availability and price. However this benefit is diminished

by the fact that the fiberglass and 3D printing are provided for free by Purdue University. Manufacturing

components from the wood composites will also require a large learning at first in order to learn how to

operate the mill machine. Machining parts will require constant supervision as well. However machining the

components can provide good precision. Therefore the manufacturability is rated at moderate. In addition,

the strength and density of plywood is moderate. Therefore the strength and mass is rated at moderate.

Therefore, based on the weight decision matrix, it was determined that 3D printing was the best option to proceed

with manufacturing.

7. TESTING PLAN

7.1 Sub-System Tests and Objectives

Test 1: The first test is a sub systems test. In the event of something catastrophic happening to avionics bay,

with avionics on board, the avionics would need to be reorder and reprogrammed. There was not time in the

semester to reorder this hardware. If the rocket without the avionics fails, the design could be improved and easily

reprinted. The objective of this test is work resolve any flight and recovery issues the rocket has before flying the

instrumentation.

The flight test test took place 12/07/2014. Battle Ground Middle School was selected as the location for the launches.

This location was surrounded by cornfields so the rocket had plenty of area to land safely. The wind was 12 mph

ESE. The fins were separate components from the motor hub. They were tight fit into slots and reinforced with tape.

The larger G-80 motor was used as well. The launch pad was an unmodified commercial launch rail. Figure 7.2.1

shows the rocket on on launch rail.

Figure 7.2.1: Flight test 1 Rocket on the Launch Pad.

The tape attaching the fins can clearly be seen in the figure. Also the cement was used to help keep the launch rail as

vertical as possible.The rocket had an estimated apogee a little over 800m. No exact data was retrieved for this flight

altitude, however rocket, didn’t reach anywhere close to that. Figure 7.2.2 shows the rocket in flight.

Figure 7.2.2: Flight 1 Trajectory

Two of the fins can be seen in the image on the left in figure 7.2.2. The one is right below the smoke trail.

The other is to above the smoke trail almost straight across from the rocket. In the left picture also shows the rocket

turned 90 degrees to the smoke trail. The image on left show wild flight path that ensued from the fin loss.

This disastrous flight test was not a total failure. There some very important lessons learned. The first is the

fins needed to be secured better. The second was to use a smaller motor, in case the rocket would become unstable.

7.2 Flight Testing

This was the flight test with avionics. The flight conditions for this flight were a little bit better than the first.

The wind was only 8 mph. Also the new motor hub that was 3-d printed with the fins as one part was installed. The

motor hub had the new G-40 motor which had less power than the the first tests G-80.

The second flight had greatly improved trajectory. Figure 7.2.3 shows the the second flight test.

Figure 7.2.3 Flight 2 Trajectory

The relatively straight trail of smoke shows the path of the rocket on the second flight in figure 7.2.3. This

flight path was greatly improved when compared to the first flight. The assent part of the flight looked really great

however the recovery did not go as well. The upper rocket (avions bay and nose cone) and lower rocket (parachute

tube and motor hub with fins) were found in different locations. Figure 7.2.4 shows the lower section as it landed.

Figure 7.2.4 Lower Section Landing

There is clearly no parachute attached attached to the lower section and from the the tub dug in it appears

to have been in free fall the entire descent as seen in figure 7.2.4. The avionics bay also appeared to be in free fall

without the parachute. The avionics bay was completely shattered, but luckily the avions were safely recovered.

This flight was clearly better that the first, however failure of the recovery system was still a big issue. The

inside of the tube was dark and chard. The shock cord that was used was elastic and only taped at the top of the

tube. The shock cord should be make from a fire resistant material and attached to the bodies more securely..

8. DESIGN AND TEST ITERATION (All)

8.1 Rocket 3 Design

As the original design of rocket was destroyed during the test. An improved design was made to cover all

the drawbacks we found during the test. Neither of the subsystem had changed. However, some details improvement

had been made to improve the overall performance.

Shock cord had been placed by the steel leader to improve the strength and fire resistance.

Figure 8.1.1 Steel leader

Two attachment points were built on the end of avionics bay and motor hub to connect with the shock cord. They

were used to replace the method which glued shock cord with rocket body by epoxy. This new method largely

improved the efficiency and strength of the attachment.

Figure 8.1.2 Attachment point on Avionics Bay

Figure 8.1.2 Attachment point on Motor Hub

8.2 Rocket 3 Flight Testing

This was the flight test with the iterated design. Overall flight condition is good. Wind speed is about 6 mph to the

east. A 10 degree launch angle was used during the launch to compensate the wind drifting effect. The figure below

shows the launch pad set for this flight test:

Figure 8.2.2 Launch pad setup of the test

The third flight had greatly improved trajectory. Figure 7.2.3 shows the the second flight test.

Figure 7.2.3 Flight 3 Trajectory

The trajectory is very clear and steady. The parachute successfully deployed and a nearly soft landing was

performed. All of the structure and subsystem were recovered,and there was a minor damage on the one fin structure

which was totally fixable.

Figure 8.2.4 Nearly Soft Landing

It is clear to see all the structure survived after flight, which demonstrated how well our parachute system worked.

However, the data stored in the SD card was not readable. It is caused by the cold weather which shrinked the

battery life and the program could not finish with the shrinked battery life. It caused a power cut off before the

program closed the file writer. This gave us a lesson to insulate the battery from the environment to ensure the

battery life, and close the file writer every iteration inside the program to ensure the safety of data.

9. PERFORMANCE RESULTS (All)

9.1 Sub-System Evaluation

9.1.1 Coupler

Integrated couplers performed exactly as intended. Separation of the rocket components after

ignition of the ejection charge performed successfully throughout all three launch tests. No changes were

needed between iterations.

9.1.2 Avionics Bay

The avionics bay housed and protected the avionics system as intended. Even when the avionics

bay shattered due to impact during the second launch, the avionics system was still intact operation. Also,

the holes drilled into the avionics bay also provide the altimeter sufficient change in pressure in order to

calculate the altitude. No changes needed between iterations.

9.1.3 Parachute bay

The parachute bay initially caused a lot of issues with parachute deployment. The parachute would

be deployed, but would only be attached to the avionics bay and nose, leaving the motor bay to free fall.

After the 2nd rocket iteration, all major issues regarding the parachute bay were solved. Sturdier shock cord

attachments were created to ensure the parachute would deploy with both sections of the rocket still

attached. In addition, a larger parachute bay was designed so the parachute would install in the rocket more

easily.

9.1.4 Motor Hub

The motor hub performed as expected throughout all the launches, with exception to the first test

launch where the fins fell off. Only changed that needed to made was printing the fins and the motor bay as

one part. The motor bay housed the rocket motor in place without any issues and the fins kept the entire

rocket stable throughout all launch test.

5.2.6 Launch Log and Rail

The launch log and rail was successful in guiding the rocket to an expected trajectory throughout all

test launches.

9.2 Demonstrated Rocket Performance

Table 9.2.1 Comparison of Altitude Prediction and Real Flight

Method Altitude to Apogee (m)/(ft) Time to Apogee (s) Error

Real Flight 816.6/ 2679.0 9.5 N/A

MATLAB Code 854.0 / 2801.8 9.67 4.58%

OpenRocket 867.5 / 2846.1 9.9 6.23%

Figure 9.2.1 The Altitude History and Comparison with prediction

From the figure 9.2.1 and table 9.2.1, it is clear that there is a close match from the mathematical model and the test

flight condition. There was only a 4.58 % error from the MATLAB Code and a 6.23% error from the OpenRocket.

However from the plot of test flight, at 8 second after launch, there was a huge transition of trajectory indicates that

the ejection of parachute. This is 1 second earlier than the prediction. If we count the fact of the early ejection of

parachute, the altitude prediction at the time of ejection would be more useful than that atapogee.

Table 9.2.2 Altitude at ejection

Method Altitude (m)/(ft) Time when charge fired Error

Real Flight 809.6/ 2656.2 8 N/A

MATLAB Code 817.1./ 2680.8 8 0.93%

OpenRocket 827.6 / 2715.2 8 2.22%

As both mathematical model provide accurate result (less than 5%), and the MATLAB Code has even less than 1%

error. We can conclude that our mathematical models are correct and the mission requirement has been satisfied.

10. REFERENCES

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Cai,, Z., & Ross, R. (n.d.). Mechanical Properties of Wood-Based Composite Materials. Retrieved

December 21, 2014, from http://www.ncsu.edu/wmtrp/publications/15IW

Mandell, G., & Caporaso, G. (1973). Topics in advanced model rocketry. Cambridge, Mass.: MIT Press.

Mastrocola, N. (1947). Effect of Number of Fins on the Drag of a Pointed Body of Revolution at Low

Supersonic Velocities. Langley Field, VA: Langley Memorial Aeronautical Laboratory.

Sanders, E. Claude. (1952). Damping in roll of models with 45, 60, and 70 degree delta wings determined

at high subsonic, transonic, and supersonic speeds with rocket -powered models. Washington, D.C.:

National Advisory Committee for Aeronautics.

Figure 4.3.1 G40-7W (Amain) http://www.amain.com/images/medium/est/est9776.jpg