The Design and organization aproach to a student build paraffin nitrous oxide hybrid sounding rocket

66th International Astronautical Congress, Jerusalem, Israel. Copyright ©2015 by the International Astronautical Federation. All rights reserved.

IAC-15,E2,3-YPVF.4,7,x30369 Page 1 of 11

IAC-15,E2,3-YPVF.4,7,x30369

THE DESIGN AND ORGANIZATIONAL APPROACH TO A STUDENT-BUILT PARAFFIN-NITROUS

OXIDE HYBRID SOUNDING ROCKET

Ashis Ghosh1, Adam De Biasi2, Jeremy Chan-Hao Wang3, Thomas Siu-Hong Leung4, Oleg Petelin5, Eric Jing-Bo

Yang6, Carl Pigeon7, Adrian Typa8, Mari Timmusk9

This paper presents the final design, testing methods and results, and organizational approach of Eos III (also

called Helios I), the University of Toronto Aerospace Team (UTAT) Rocketry Division’s third-generation sounding

rocket. Eos III was developed over a period of 10 months with the goal of delivering a 1.33kg 1U CubeSat scientific

payload to 3km above ground level, as part of the 2015 Intercollegiate Rocket Engineering Competition (IREC).

Because a standard carbon-fibre airframe was used, the design discussion focuses on propulsion, avionics, and payload

instead. The organizational approach is briefly discussed in terms of team structure and community impact.

Eos III was powered by the 8 100Ns (tested so far, with a target of 10 000-Ns) 'Bia III' hybrid rocket engine,

which used a mixture of paraffin-carbon black as fuel and nitrous oxide as the oxidizer. Fuel cartridges and shoulder-

bolted assemblies promoted ease of assembly and enabled multiple consecutive static test fires. Modular avionics

enabled independent system development, simplicity of design, and reparability. The payload contained an inertial

measurement unit, atmospheric sampling and weather-sensing units, and parachute recovery system, all arranged inside

a standard 1U CubeSat. Many of the components, including structures, internal flows and aerodynamics, engine, and

flight performance were simulated through in-house or commercial software. Ground tests validated these predictions.

The Rocketry Division was headed by two individuals (Lead and Chief Designer) and organized into five

subdivisions (Propulsion, Fluids, Avionics, Payload, and Structures). A team of four high school students was also

selected to develop the scientific payload under the mentorship of undergraduate and graduate students. Tandem with

the development of the rocket itself, the Rocketry Division was heavily involved in educating, inspiring, or simply

reaching out to members of the general public, high school students, and aerospace professionals.

Ultimately, careful simulation, strategic resource allocation, efficient organizational structure, and

collaboration with high school students led to a promising with valuable community impact. A limited launch window,

however, prevented completion of the launch procedures at IREC and a reattempt is scheduled for the future.

I. INTRODUCTION

Since the launch of the first man made satellite by

the Soviet Union, satellites have revolutionized our

civilization. For a large portion of human’s history in

space, the focus has been on developing large

multifunctional satellite missions. In recent years, the

global interest in nano (<10 kg) and microsatellites

(<100 kg) has increased. Beginning in 1999, California

Polytechnic State University and Stanford University

developed a standardized approach to nanosatellites

form factor referred to as a CubeSat [1].

Fig. 1: Distribution of Orbital Satellite Mass: 2000-

2009 for 0-10 kg Satellite Class [2]

As seen in Fig. 1, the 1 kg Cubesat is the most

popular due to its cost and short delivery time. Trends

show that small satellite development will continue to

grow and the need for launches will increase [2]. For

the trend to continue, developing services such as

launch services that support the microsatellite

community are needed.

Currently, small satellites rely heavily on piggy-

rides on medium to large launch vehicles such as

India’s PSLV and Russian rockets. The disadvantage

of being a secondary or tertiary payload on a large

rocket is that the availability, scheduling and orbit

parameters depend on the primary payload. Meaning

that preferred orbital locations cannot always be

achieved and compromises have to be made to obtain

a launch opportunity. New launch options dedicated

for small satellite payload would be a valuable service.

The payload and launch vehicle designed by the

University of Toronto focuses on demonstrating a

technology that could offer solutions to the scarcity of

dedicated small satellite launches. Moreover, a hybrid

engine was chosen due its safety and increasing

promise for sounding rocket applications [3].



Fig. 2: A solid model of the final Eos III rocket design.

II. ROCKET DESIGN

The overall rocket design is discussed in terms of

the propulsion subsystem, avionics subsystem, and

payload subsystem. The mission profile was to launch

a single-stage hybrid rocket to 3km and deploy a

CubeSat 1U at apogee. Fig. 2 shows a solid model of

the final design for Eos III.

II.I Propulsion Subsystem

The propulsion subsystem consisted of all

mechanical and chemical design considerations

associated with engine performance. Discussed here

are the MATLAB engine performance suite, oxidizer

tank, plumbing bay, injector assembly, engine

chamber, and nozzle assembly.

II.I.I Simulation Tools: MATLAB and Excel

To estimate engine performance of proposed

designs, a MATLAB engine performance suite was

developed in tandem with a steady-state Excel

spreadsheet. Whereas the MATLAB program

provided a transient prediction of key performance

parameters like thrust and apogee via a rudimentary

aerodynamic model, the spreadsheet was a rapid

design tool in which promising designs could later

be inputted into the MATLAB simulation.

Key assumptions in both the MATLAB

simulation and Excel spreadsheet were: (1) the

oxidizer maintained liquid state until it entered the

engine; (2) the fuel core regression could be simply

modelled with the following equation [4]:

�̇� = 𝑎𝐺𝑜𝑥𝑛

(3) chemical combustion followed assumptions in

NASA Chemical Equilibrium with Applications

[5]; (4) isentropic expansion across the nozzle. The

nitrous oxide thermophysics was modelled

according to work by Fernandez [6]. Difficulties in

estimating Reynolds numbers led to the adoption of

industry practices in constraints for quenching or

blowout, borrowed from Humble [7]. In addition,

experimental or simulated values were used for the

airframe drag and injector discharge coefficient.

The overall analytical modelling and flow of

calculations for the MATLAB engine followed that

found in work by Genevieve [8], where instead of

invoking NASA CEA, a multivariable polynomial

regression equation was used to speed up

calculations. The rocket trajectory model was

simplified to a dynamics problem with a constant

but conservative drag coefficient.

The final design parameters were in

agreement between the MATLAB results and

spreadsheet outputs, and are covered in detail in the

remaining subsections. Given that rocket weight

(80 lbm) could not be further lowered before

IREC—mainly due to the technical interest in the

propulsion system despite its inherent weight— the

best performance predicted by MATLAB is shown

in Fig. 3 and features an apogee of 8000 ft (still

qualifying for IREC). The predictions agree with

existing work demonstrating that optimal

performance is reached near an oxidizer-fuel ratio

of around 7 [9]. The initial peak in thrust is due to

an initial high “guess” required for engine pressure,

based on the thermochemical solution for rocket

inputs used by NASA CEA. The MATLAB tool

was validated by comparison with existing data

from the Rocketry Division, and with data from the

University of Washington up to 15% error [9].

Fig. 3: Predicted performance by MATLAB

Program, including thrust and apogee.



II.I.II Oxidizer Tank

The custom oxidizer tank was designed to

the following specifications: (1) 9L capacity; (2)

maximum pressure rating of 2,000 psi with an

additional safety factor of 2; (3) maximum outer

diameter 5.4” to fit past the airframe connectors.

The tank was made from 6061-T6 aluminum alloy

with ellipsoid caps welded to the main right

cylindrical body. The preliminary dimensions of

the minimum weight version were calculated

using thin-walled pressure vessel theory, using a

set inner diameter of 4.5 in. to generate the

required wall thicknesses for the caps and the

main body. Due to the ellipsoid geometry of the

caps and the difference in thicknesses between the

caps and the main body of the tank, the minimum

weight design was further refined through finite

element analysis simulations to eliminate stress

concentrators while maintaining minimum weight.

The ends of the tank's main body were modified to

facilitate the welding of the caps. The strength of

the welds were determined by the strength of the

filler metal used in the weld. The filler metal and

welding technique used were decided by the

welder according to the design specifications

provided. The tank was successfully hydro-tested

to 2,000 psi without issue. The tank solid model is

shown below in Fig. 4

Fig. 4: Oxidizer tank solid model, featuring top-view

with main and secondary outlets (right) and the

overall tank (left).

II.I.III Plumbing Bay

In between the oxidizer tank and engine was a

set of plumbing designed to enable diagnostics and

control the flow of oxidizer. There were two

systems running in parallel: (1) main feed system;

(2) safety & monitoring system.

The main feed system used a servo-actuated

1/4” ball valve, mounted on the valve with a 3D

printed actuation mount. ½” Swagelok

compression straights were used as the main feed

line for ease of assembly and minimization of

pressure losses. The stem of the valve and the

servo’s driving shaft were connected by a 3D

printed adapter, which used set screws to hold the

2 parts in place. The ½” straights were adapted to

a ¼” ball valve, which constrained oxidizer flow

but was implemented due to the higher torque and

size limitations with a larger ½” valve.

Safety, monitoring, and venting were achieved

with plumbing connected to a second, smaller

outlet beside the main ½” line. This second outlet

was connected to a coaxial vent line, pressure

transducer, thermocouple, 1100psi pressure relief

valve, and a 1/8” ball valve open to the

environment were included in this secondary

system, with ¼” Swagelok compression fittings for

space management and ease of assembly. The

oxidizer tank was filled by connecting a quick

disconnect fitting to the main nitrous tank, and

closing the 1/8” ball valve when liquid nitrous

entered the coaxial vent lines.

II.I.IV Injector Assembly

Oxidizer injection was designed with the

following major factors in mind: (1) mass flow

rate; (2) pressure drop; (3) oxidizer dispersion. In

the Bia III engine, the injector assembly consisted

of an injector manifold and an injector plate. The

plate diameter was 3.2” as compared to the

minimum 0.25” diameter of the main feeding

system. Sealing was achieved with O-rings placed

at various locations. A parabolic profile was chosen

for the injector manifold because recirculation

region would have been prominent in the case of a

rectangular profile. As well, the parabolic profile of

the manifold allowed radial bolts to be used for

securing the assembly to the engine chamber itself.

A cross section of the assembly is shown below.

Fig. 5: Section view of the injector assembly



Table 1: CFD and cold-flow tests for 3 injectors.

To balance the need for high �̇�𝑜𝑥 (~1kg/s

according to MATLAB) and sufficient pressure

drop (at least 30% against 𝑃𝑐𝑐 = 250psia, to buffer

against instabilities), fine tuning was required for

the size, number and angle of these holes. Initially,

three distinct arrangements were designed and

tested through computational fluid dynamics

(CFD) and cold flow tests using carbon dioxide.

The three designs considered were straight,

impinging and swirl (Table 1).

Results from the cold flow test, although

affected by �̇�𝐶𝑂2 due to limitations of valves on

industrial carbon dioxide cylinder, verified the flow

pattern generated using CFD simulations,

suggesting that the results obtained were fair

approximation of the actual performance. Based on

these flow patterns and engine testing data, a final

injector was designed to incorporate the dispersive

and azimuthal flow properties of the first and

second designs (Fig. 6).

Fig. 6: Final injector plate design.

II.I.V Engine Chamber

The engine chamber had ¼” aluminum walls

with an inner diameter of 4.5”. It was divided into

a 1” long pre-combustor, 12” long fuel core with

2.5” single circular port, and a 12” polyurethane-

fiberglass ablative postcombustor liner with ½”

thickness—a more cost-effective alternative to

industry-grade liners. Ignition was achieved

through three smaller B-size solid rocket motors

impregnated in the top of the fuel core.

Description Pattern Cold Flow Test CFD Simulation

31 holes, φ=3/32’’

Straight.

23 holes, φ=1/8’

Inner and outer ring counter-

clockwise, at 14°. Middle

ring clockwise, also at 14°.

24 holes, φ=1/8’

All holes are at 14° pointing

toward the center.

Not conducted due to

limited resources.

(red-blue:high-low pressure)



A standard 1” pre- combustor was selected

based on good practices by Sutton [10].

The fuel core dimensions were the result of

simulations from the MATLAB performance suite,

discussed above. A small amount of carbon black

(2.5g/kg of fuel) was added to opacify the fuel and

protect the engine from radiation heating—the

effect on combustion performance was deemed

negligible. Fuel cores were poured into ½” thick

cardboard tubes, with a PVC inner mandrill for the

port diameter. These cartridges could be loaded and

removed from the engine to expedite testing and

transportation.

The length of the postcombustor was chosen to

allow enough time for mixing and combustion, as

approximated with characteristic combustor length

[11]. The postcombustor thickness was determined

through informal small scale testing using a 2”

inner diameter steel pipe as the stand-in

combustion chamber and observing the remaining

thickness of postcombustor after various tests.

II.I.VI Nozzle Assembly

The nozzle assembly was also held in with

radial shoulder-bolts and utilized a graphite conical

nozzle due to cost-effectiveness and machining

simplicity. The area ratio was 4.2 with a half-angle

of 12 degrees. Due to the high thermal conductivity

of graphite, a steel and not aluminum backing plate

was used to hold the nozzle in place. One-

dimensional thermal calculations assuming a

worst-case scenario of stagnation temperature

(3000K) at the nozzle inner wall showed that the

steel would not exceed melting temperatures. The

nozzle assembly is shown in (Fig. 7).

Fig. 7: The nozzle assembly, demonstrating the

shoulder-bolted design with graphite nozzle.

II.II Avionics Subsystem

The purpose of the avionics system was three-fold:

(1) to provide sensor (telemetry) data about the

rocket’s internal state during both flight and static

ground testing; (2) to communicate with a remote

ground station and initiate the launch sequence by

actuating and igniting the flow of oxidizer; (3) to

deploy the payload and drogue parachute when apogee

is reached and to deploy the main parachute when the

rocket is descending and is almost at ground level.

This section describes parts of the avionics

hardware and software system that were necessary to

safely launch and recover the rocket. At a high level

the avionics hardware was split between the ground

station and the avionics bay. The remote ground station

received telemetry data from the rocket via a wireless

link and sent ignition commands to the rocket via the

wireless and hard physical links. The avionics bay

contained all the sensors, power supplies and

processing devices necessary for static fire testing and

for launching the rocket. A half-page systems diagram

is depicted in Fig. 9 on the subsequent page.

II.II.I Power Distribution and Motherboard

The avionics bay had two power sources: (1)

a 12V 2000mAh battery pack providing 12V, 5V

and 3.3V to the motherboard where 5V and 3.3V

are derived from the 12V through switching buck

converters (150kHz) on the power board (Fig. 9);

(2) a 9V battery provided power to the redundant

parachute deployment system—the Raven 3

commercial altimeter (which was a redundant

system for parachute and payload deployment).

As shown in Fig. 8, the power board supplied

12V, 5V and 3.3V to the motherboard which

distributed these voltages to all the daughterboards.

Note that the 12V, 5V and 3.3V rails had their own

dedicated ground—this helped mitigate ground-

bounce noise seen by devices when high current

was being drawn through another supply rail (i.e.

when oxidizer valve motor draws 5A on the 12V

rail the “ground” reference jumped more

significantly on the 12V rail than the other rails).

Fig. 8 also shows the I2C bus which provided a

communication interface between the

daughterboards. The I2C bus was used in a single-

master multiple-slave configuration with the arbiter

daughterboard acting as the master.

Fig. 8: Power board-motherboard interface.



Fig. 9: Avionics systems diagram.

II.II.II Common Daughterboard Hardware

All daughterboards in the avionics system had

a common interface to the motherboard. As shown

in Fig. 10, each daughterboard connected to the

motherboard via a standard header. The header

supplied 12V, 5V and 3.3V, and connected the

daughterboard to the shared I2C bus.

A USB-to-UART (RS232) bridge allowed the

Atmel Atmega328p microcontroller to connect to a

PC and send/receive data via serial. Programming

the microcontroller could be done in one of two

ways: (1) through the USB connection if a

bootloader was present on the microcontroller; (2)

through the AVRISP2 six-pin header (standard

Atmel programming interface) or through the SPI

connection if the microcontroller was blank and did

not have a bootloader.

Fig. 10: Common daughterboard hardware.



II.II.III Oxidizer Actuation

The oxidizer actuation system was composed of

an H-bridge (TLE5206) and the relevant pins used

to interface with the encoder and limit switches of

the oxidizer actuation housing. The encoder pin

was fed into the microcontroller’s timer/counter

input, which enabled counting of the square-wave

pulses of the encoder without the use of interrupts.

The switches were fed directly into GPIO pins of

the microcontroller.

II.II.IV Pyrotechnic Actuation

The ignition circuit was an implementation of a

‘firing circuit’, a custom modular-design solid-

state pyrotechnic actuation circuit (Fig. 11). The

circuit had two different MOSFETs that controlled

the current flow, labelled ‘arm’ and ‘fire’. When

the circuit was armed (the ‘arm’ FET switched into

saturation mode) the e-match attached in series was

energized to 9V, and when the ‘fire’ FET was

switched on also, sufficient current flowed in order

to activate the e-match. The circuit was designed

such that before the arming, there was no voltage

on the e-match, and therefore no risk of a short

circuit. The reason behind the solid-state

construction was that relays (typically used for

pyrotechnic actuation) are susceptible to launch

vibrations. An addition feature of the firing circuit

was continuity detection across the e-match, which

allowed for the detection of incorrectly attached e-

matches. The circuit was designed to be modular,

and could be implemented in such a way that

multiple channels of e-matches were armed

together and fired separately. Each channel could,

in turn, fire three e-matches.

II.II.V Sensors

The pressure transducers present were 1000 psi

models that fed their data output in the form of a

current between 4-20mA full scale. In order to

convert that to a voltage that the microcontroller’s

on-board ADC could read, a noninverting amplifier

op-amp circuit was used, ultimately amplifying the

signal to a range of 0.8~4V.

The thermocouple amplifier used was

MAX31855, which interfaced with the

microcontroller over the SPI protocol (Fig. 12).

This particular chip performed cold-junction

compensation, enabling the use of a thermocouple

without a constant temperature junction on the

other side. As well, since the thermocouple

amplifier also had an internal temperature sensor, it

could be used to sense the temperature of the

interior of the avionics bay.

Fig. 11: Pyrotechnic actuation circuit (i.e. firing

circuit for test or launch).

Fig. 12: Temperature signal amplification circuit.

II.III Payload

The objective was to demonstrate that it is

technically feasible to deploy a small satellite from a

small rocket. The system was made for small diameter

rockets capable of suborbital flight. It would be

feasible to adapt the system to medium sized rockets

capable of achieving orbital mission or air launch

missiles carried to high altitudes by aircraft before

being launched. The flexibility and versatility of this

payload aim to open conversation and possibilities of

commercial adaptation of the technology.

Key requirements of the CubeSat included: (1)

mass under 1.33kg; (2) conformation to the 1U

CubeSat form factor; (3) presence of a remove-before-

flight pin and inactivation of all electronics when

engaged; (4) non-transmissions and no power while



Fig. 13: Deployment system, cross-sectional view.

stowed in the deployment bay; (5) inclusion of a

parachute in the design, to achieve a descent rate of

3.5m/s to 4.5m/s

II.III.I CubeSat Contents

The final design included a stowed parachute as

shown in Fig. 14 and an electronics compartment

which enabled the scientific aspect of the mission.

An Arduino Uno was incorporated, with weather

sensors for the measurement of atmospheric

temperature, pressure, humidity, acceleration and

rotational rates. A static line and deployment bag

enabled the parachute to be deployed upon exiting

the deployment bay. The deployment bag would

remain with the rocket once the parachute was been

deployed. The recovery of the CubeSat would be

done by radio tracking with one of IREC’s radios.

Fig. 14: The completed CubeSat 1U.

II.III.II Deployment

The CubeSat is deployed perpendicular to the

axial direction of the rocket, via four springs

located on the top and bottom of the satellite (Fig.

13). The springs were attached to an aluminum

back plate which the satellite is held up against in

compression when the door is in place. The door is

kept closed by two explosive bolts. Once apogee is

reached, the explosive bolts simultaneously

detonate thus releasing the stored energy of the

springs and ejecting the satellite.

Horizontal deployment was primarily done to

not interfere with the parachute bay stowed at the

nose cone. Additionally, the sideways deployment

meant that a number of these deployment systems

could be stacked onto each other in order to deploy

multiply CubeSats independently aboard a single

rocket.

II.III.III Structure

The structure of the payload bay was made from

6061 aluminum with a carbon-fibre shell for

aerodynamic streamlining. Rails machined into the

payload guide the CubeSat out of the bay. UTAT

female bay connector grooves were added for quick

integration with the rest of the rocket.

II.III.IV Camera Module

A camera module was an additional payload

carried on-board the rocket to serve as a means of

acquiring photographic evidence confirming the

deployment of the CubeSat at apogee. The camera

module used a GoPro camera system remotely

operated from the ground station. The data was

stored locally and could be viewed once the rocket



has been recovered. Since the camera module was

connected with a standard UTAT bay connector to

the deployment system payload bay, the camera

module was optional and could be reused in future

years as a stand-alone unit.

III. ORGANIZATIONAL APPROACH

The organizational approach to Eos III

development may be discussed in terms of: (1) team

leadership; (2) key projects and division of labour; (3)

timeline; (4) community involvement.

Rocketry Lead and Chief Designer had already

spent two years with the Rocketry Division. The

former handled administrative elements and overall

project management. The latter handled systems

integration and providing design advice—especially in

terms of general mechanics and aspects of manufacture

or assembly—to the subsystem leads.

By contrast, the subsystem leads were responsible

for projects relevant to their expertise. Propulsion was

chiefly responsible for the oxidizer tank, plumbing,

and engine mechanical design as well as combustion

performance. Avionics was responsible for the on-

board sensors, microcontrollers, and effectors (e.g.

firing circuit for the engine), as well as the ground

station. Structures was responsible for the airframe and

recovery system (not discussed here due to standard

approaches) and advising propulsion structures. Fluid

mechanics was responsible for designing the injector

plate, nosecone, and fins. Payload was responsible for

the deployment bay and actual scientific

instrumentation, namely the CubeSat 1U. The test

facility was a collaborative effort between all

subsystems. Following this matrix organizational

structure, the team could harness the technical

knowledge of its members to accomplish the projects.

Of the 10-month period allotted for development,

testing, and integration, the first three months were

spent setting high-level requirements, performing

conceptual design, and developing or becoming

familiar with simulation tools. At that point,

approximately 40 new members joined the team as part

of recruitment activities taking place at the start of the

new Academic Year. These new members were

integrated into the team through ‘beginner’ projects

such as simple mechanical designs or fabrication tasks.

New members were also encouraged to attend the

University of Toronto Aerospace Team’s general

aerospace seminar series which covered topics ranging

from solid modelling to aircraft and spacecraft

electronics to machining and fabrication. The

remaining seven months witnessed two 3-month

engine testing phases, the first phase involving minor

iterations on the initial design and the second phase

allowing for major changes relative to the initial

concept. The schedule primarily motivated by engine

development—the most challenging technical

element—the other subsystems planned accordingly to

aim for integration at T-2 months before the

competition. Delays in manufacture and further

required engine testing resulted in integration taking

place at T-1 month before competition. This prevented

the possibility of any pre-competition launch attempt,

during which if a severe failure occurred there might

not be enough time to repair systems for the

competition.

Major bottlenecks throughout this time were: (1)

manufacture limitations as the entire rocket was

designed, built, and tested by students (i.e. all

mechanical parts machined or wet-laid-up by

students); (2) design and debugging of avionics

systems, which was typically the last to begin of any

subsystem because it could be done only after the

mechanical parts they controlled/were housed

in/sensed from were specified.

Lastly, being a student-led initiative, a secondary

motivation of this project was to educate and inspire

members of the local and international community.

Throughout the design and testing of Eos III, the team

interacted with members of the public and high school

students at events such as Science Rendezvous, guest

lectures at local secondary schools, appearances at

conferences (e.g. International Space Development

Conference), on-campus design showcases, university

applicant events or open-houses, and talks at

organizations affiliated with students or the team at

large, such as the German Aerospace Center (DLR).

These opportunities provided an easy way to engage

others requiring minimal preparation aside from

bringing items and multimedia for display and

explaining concepts well. It is estimated that the team

was able to directly interact (i.e. speak with) just under

1000 individuals across these settings.

IV. MAJOR TESTS AND RESULTS

This section discusses the end results of engine

testing, as well as qualitative challenges with avionics

and launch operations at IREC. Overall, the Bia III

engine could provide a maximum thrust of 280-lbf

and an average of 200-lbf over the course of the 9-

second burn. This is approximately 2/3 the thrust of,

and 4 seconds longer than, the MATLAB prediction.

The primary suspected reason is low oxidizer flow

rates due to choking at main valve in the plumbing.

IV.I Engine Testing

A number of static test fires were conducted at a

student-constructed, inverted-engine static test fire

facility at the University of Toronto Institute for

Aerospace Studies. A photo from the last test

conducted is shown in Fig. 15, along with engine thrust

and pressure data.



Fig. 15: Photo of last engine test on the Bia III engine

noting the perfectly expanded nozzle gases (top);

thrust and pressure versus time graph (bottom).

Smoke near the bottom of the engine (top picture)

was from e-match cables spit out of the engine.

It is evident from the thrust curve that there is a

rapid increase followed by a relatively steep and then

even and shallower decrease until the engine burn

concludes at 9 seconds. Compared with the MATLAB

predictions of Fig. 3, the experimental thrust curve

bears similar shape but with 1/3 less thrust. The

pressure, too, is consistently about 1/3 less than that

predicted by MATLAB.

The main suspected reason for this is that the ¼”

main valve, which was the largest valve that could be

accommodated, was likely choking the flow and

limiting the usefulness of the otherwise ½”

components in the feeding system. Furthermore, the

extended burn time the presence of more leftover fuel

than expected add to the pool of evidence supporting

this hypothesis. In total, this resulted in a total impulse

of only 8100Ns as opposed to the 10 000Ns originally

predicted, but it is strongly suspected that 10kNs if not

more is possible given future refinements and the

performance of engines of similar scale [9].

IV.II Avionics Challenges and Testing

Throughout prototyping, it was found that the

oxidizer actuation’s microcontroller was not very

consistent at reading the states of the switches. When

observed using an oscilloscope, it was found that the

switch lines had a considerable amount of noise. This

was attributed to the fact that the high frequency pulses

of the square-wave encoder were adjacent to the switch

lines. However, a solution was found in delaying the

switch polling code, perhaps minimizing interference

between the lines.

One potential concern with the pressure transducer

circuit was its temperature stability, as the resistors

used in this analog circuit were temperature dependent

and therefore the readings could change with changing

ambient temperatures. However, during multiple fire

tests in subzero and standard room temperatures, the

circuit continued to provide reliable pressure data once

re-calibration was performed.

Overall the modularity of the avionics system came

at the expense of complexity and whether the system

will provide its promised long-term value will depend

on future attempts to ‘evolve’ the system instead of

replacing it entirely.

IV.III IREC Performance

A limited launch window resulted in a delay of the

original launch of the Eos III due to the time required

for launch preparation and on-site debugging. The

rocket has not yet flown and in the future UTAT will

seek to make improvements to the design before

reattempting launch before July 2016. Pre-flight

ground testing did show however that the payload

deployment mechanism was successful, and that the

CubeSat parachute packaging was conducive to

opening up when dropped from a tall building.

V. CONCLUSION & NEXT STEPS

A promising design has been designed and ground

tested, pending future launch tests. The MATLAB/

Excel performance suite demonstrated accuracy as a

design and prediction tool. During ground tests, the Bia

III underperformed, providing only 8100kNs of the

target 10kNs, but this was largely attributed to

restrictive oxidizer plumbing that will be changed. The

avionics system was able to support rigorous test

campaigns, and was successfully implemented despite

its modularity leading to more complexity than

0

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traditional avionics systems. The payload deployment

was effective in ground tests but flight qualification is

needed along with integrated qualification of Eos III.

In the future, the current valves and piping will be

replaced with a burst disk system for simplicity,

weight-saving, and increased oxidizer mass flow rates.

Flight telemetry may also be added to the avionics

system to supply additional data regarding the rocket

trajectory and engine status. High school students may

work with the team again for community impact, but

the Rocketry Division will likely partner with another

design team at the University of Toronto to develop a

more complex payload. Lastly, a flight will be

attempted either in Canada or the United States, not

only to demonstrate Eos III’s capabilities but to serve

as a simulated launch operations sequence for the

Rocketry Division. If a successful payload deployment

at 3km is achieved, then the Division will continue

increasing the target altitude and stepping up engine

performance with it.

VI. ACKNOWLEDGEMENTS

To the University of Toronto Division of

Engineering Science, Institute for Aerospace Studies,

Department of Electrical & Computer Engineering,

Department of Mechanical & Industrial Engineering,

and Engineering Society, for funding or otherwise

supporting the bulk of this project; to past and present

members of the University of Toronto Aerospace

Team’s Powered Flight, UAV, Space Systems, and

Outreach Divisions for their advice and moral support;

to Adam Paul Trumpour, for readily critiquing and

having assisted with cold-flow testing; finally, to the

numerous sponsors and partners who shall go unnamed

in writing but were instrumental in enabling the team’s

hybrid rocket programme.

VII. REFERENCES

[1] R. Nugent, R. Munakata, A. Chin, R. Coelho

and J. Puig-Sairi, "The CubeSat: The

Picosatellite Standard for Research and

Education," AIAA, San Diego CA, 2008.

[2] D. DePasquale and A. C. Charania, "Analysis

of the Earth-to-Orbit Launch Market for

Nano and Microsatellites," AIAA, Anaheim

CA, 2010.

[3] E. Doran, J. Dyer, K. Lohner, Z. Dunn, M.

Marzoña and E. Karlik, "Peregrine Sounding

Rocket," Stanford University, Stanford, CA,

2008.

[4] G. Ziliac and M. A. Karabeyoglu, "Hybrid

Rocket Fuel Regression Rate Data and

Modelling," AIAA, Sacramento CA, 2006.

[5] NASA, "Chemical Equililbrium with

Applications (CEA)," NASA, Cleveland,

OH, 2014.

[6] M. Fernandez, "Propellant tank

pressurization modelling for a hybrid

rocket," Rochester Institute of Technology,

Rochester NY, 2009.

[7] R. W. Humble, G. H. Henry and W. J.

Larson, "Space Propulsion Analysis and

design," McGraw-HIll, New York, NY,

1995.

[8] B. Genevieve, M. Brooks, P. Beaujardiere

and L. Roberts, "Performance Modeling of a

Paraffin Wax / Nitrous Oxide Hybrid

Motor," AIAA, Orlando, 2011.

[9] T. Edwards, V. Hansen, T. Slais, C. Chu, M.

Hughes, G. Li, G. Finnegan, T. Ip, A. Hatt

and B. Degang, "University of Washington

DAQ Destroyer Hybrid Rocket," Seattle,

2012.

[10] G. Sutton and O. Biblarz, Rocket Propulsion

Elements, New York City, NY: Wiley, 2010.

[11] B. T. C. Zandbergen, "Hybrid Rocket

Motors," Delft University of Technology,

Delft, 1999.

1 University of Toronto Mechanical & Industrial Engineering, Canada, [email protected] 2 University of Toronto Mechanical & Industrial Engineering, Canada, [email protected] 3 University of Toronto Engineering Science, Canada, [email protected] 4 University of Toronto Engineering Science, Canada, [email protected] 5 University of Toronto Electrical & Computer Engineering, [email protected] 6 University of Toronto Engineering Science, Canada, [email protected] 7 University of Toronto Institute for Aerospace Studies, Canada, [email protected] 8 University of Toronto Mechanical & Industrial Engineering, Canada, [email protected] 9 University of Toronto Mechanical & Industrial Engineering, Canada, [email protected]

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The Design and organization aproach to a student build paraffin nitrous oxide hybrid sounding rocket

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