1. FDM Hybrid Rocket Grains
Fused Layer ABS Rocketry Experiment (F.L.A.R.E.)
Critical Design Review
EML 4501/EAS 4700-Mechanical/Aerospace Engineering Design
April 20, 2015
Point of Contact:
Amy Besio
Team Members:
Jonathan Benson, Richard Horta, Joshua Rou, John Seligson
Faculty Advisor:
Justin Karl, Ph.D.

2. i
Ethics Statement
As engineer, we will uphold ourselves to the code of ethics set forth by the American
Society of Mechanical Engineers [1].
We, as engineers will uphold and advance the integrity, honor and dignity of the engineering
profession by:
I. using our knowledge and skill for the enhancement of human welfare;
II. being honest and impartial, and serving with fidelity their clients and the public;
III. striving to increase the competence and prestige of the engineering profession.
By signing this document, we agree to abide by these fundamental principles. We acknowledge
that this work is our original work and will provide credit when paraphrasing work that is not our
own.
Signatures Date
Amy Besio ________________________________________ _________________
Jonathan Benson ________________________________________ _________________
Richard Horta ________________________________________ _________________
Joshua Rou ________________________________________ _________________
John Seligson ________________________________________ _________________

3. ii
Abstract
This project explores utilizing fused deposition modeling (FDM) for optimization of hybrid
rocket fuel grains. FDM will allow for custom tailoring of fuel grain geometries, in order to target
desirable performance characteristics unobtainable through traditional manufacturing. The solid
propellant will be composed of acrylonitrile butadiene styrene (ABS), a common additive
manufacturing material. When exposed to an oxidizer, ABS performs comparably to commercially
available hydroxyl-terminated polybutadiene (HTPB) fuel grains. The liquid propellant will be
composed of nitrous oxide (N2O) and will provide the oxygen content to the fuel.
The scope of this project includes design, manufacturing, testing, and data review of the fuel
grains. Development of the grains entails forming appropriate mathematical models for solid and
liquid propellant characterization. Manufacturing encompasses fabrication of the ABS grains
using FDM and assembly of test bed components, which includes the test stand, thrust chamber,
and data acquisition and processing. Testing will consist of a baseline run, followed by subsequent
test fires. Data review includes the testing analysis and a comparison with computational
prediction.

4. iii
Table of Contents
Ethics Statement......................................................................................................Error! Bookmark not defined.
Abstract ……………………………………………………………………………………………………………….
iError! Bookmark not defined.
List of Figures.......................................................................................................................................................iv
List of Tables .........................................................................................................................................................v
Nomenclature........................................................................................................................................................vi
1.0 Project Overview..............................................................................................................................................1
2.0 Design Parameters............................................................................................................................................2
3.0 Parametric Design.............................................................................................................................................5
4.0 Detail Design.....................................................................................................Error! Bookmark not defined.
5.0 Bill of Materials………………..…………………………………………………………………………………27
References…………………………………………………………………………………………………………….29

5. iv
List of Figures
Figure 1.1: Classical Hybrid Configuration…………………………………………………………….....2
Figure 2.1: Fuel Grain Geometries By Casting…………………………………………………………....3
Figure 3.1: Example of Fuel Grain CAD Model…………………………………………………….….…6
Figure 3.2: I_sp vs. O/F Ratio……………………………………………………………………….….…8
Figure 3.3: Nitrous Oxide Phase Diagram…………………………………………………………..…….8
Figure 3.4: Commercial Rocket Configuration (Not Drawn to Scale)…………………………………...10
Figure 4.1: Combustion Chamber………………………………………………………………………..16
Figure 4.2: Forward Bulkhead (Front and Rear View) …………………………………………….……16
Figure 4.3: Aft Bulkhead (Front and Rear View)………………………………………………………..17
Figure 4.4: Thrust Chamber FEA-Displacement…………………………………………………...……17
Figure 4.5: Thrust Chamber FEA-Von-Mises Stress……………………………………………………18
Figure 4.6: Nozzle Design……………….………………………………………………………………18
Figure 4.7: Oxidizer Feed System………………………………………………………………….……19
Figure 4.8: Rail Frame Design…………………………………………………………………………...20
Figure 4.9: Platform Design……………………………………………………………………………...20
Figure 4.10: Platform and Wheels- Cross Section………………………………………………………20
Figure 4.11: Superstrut Bed with Clamps………………………………………………………………..21
Figure 4.12: Assembled Test Stand- Isometric View…………………………………………………....21
Figure 4.13: Assembled Test Stand- Top View………………………………………………………….22
Figure 4.14: Finite Element Analysis of Rail Frame- Displacement…………………………………….22
Figure 4.15: Measurement Apparatus……………………………………………………………………23
Figure 4.15: Pressure vs Burn Times………………………………………………………………….…25
Figure 4.16: Tiva™ C Series TM4C123G LaunchPad [Texas Instruments] ……………………………26
Figure 4.17: Control System Diagram…………………………………………………………………...27
Figure 6.1: Conduction Coefficient Formulation………………………………………………………..29
Figure 6.2: Temperature Equation Array………………………………………………………………..30
Figure 6.3: Temperature Distribution Per Node (Temp in K) ………………………………………….30

6. v
List of Tables
Table 4.1: Enthalpy of Formation………………………………………………………………………15
Table 5.1: Bill of Materials…………………………………………………………………………......27

7. vi
Nomenclature
𝑨 𝟐 Nozzle Exit Area
𝑭 Thrust
𝑮 Free Stream Velocity
𝒈 Acceleration Due to Gravity
𝑰 𝒔𝒑 Specific Impulse
𝒎̇ Propellant Mass Flow Rate
𝒑 𝟐 Nozzle Exit Pressure
𝒑 𝟑 Atmospheric Pressure
𝒓̇ Fuel Grain Regression Rate
𝒗 𝟐 Exhaust Velocity
𝒙 Axial Location for Combustion Port
𝜷 Non-Dimensioned Fuel Mass Flux from Fuel Vapor
𝝁 Combustion Gas Viscosity
𝝆 𝒇 Solid Phase Fuel Density

8. 1
1.0 Project Overview
1.1 Background
Bipropellant liquid and solid rocket motors that combine volatile and energetic propellants
have been the industry standard for the last fifty years [2]. Liquid engines have a high specific
impulse, Isp, and thrust-to-weight ratio necessary for launching large payloads. Typical liquid
engines use a liquid oxidizer and liquid fuel that are combined and burned in a combustion chamber
[2]. If the complex turbomachinery that mixes the two propellants fails, the combustion becomes
unstable and causes a loss of thrust and eventual loss of the vehicle. Solid rocket motors use a solid
fuel grain that consists of solid fuel and oxidizer particles that are mixed together with a binding
agent. For optimum solid motor performance, the oxidizer and fuel must be mixed to a specific
ratio as defined by the desired total impulse. The regression of the solid fuel grain depends greatly
on combustion chamber pressure. If the pressure increases too quickly, the motor is likely to
explode. Structural imperfections in the fuel grain can also cause an over pressurization of the
chamber as a result of an increase in local burning rate [3]. Both conventional launch systems have
a high susceptibility to failure of 8% according Claude Lafleur’s Spacecraft Encyclopedia. As
space exploration shifts from government to civilian space, a new market arises that is confronted
with cost, performance, and safety challenges that will not be satisfied by the aforementioned
launch systems [2].
1.2 Hybrid Rockets
Hybrid rocket motors have the ability to fulfil the above-mentioned flight requirements, as
they provide a cost effective and safer alternative to liquid and solid systems. A classical hybrid
configuration consists of a liquid oxidizer and solid fuel grain that are housed separately as seen
in Figure 1.1. They employ non-toxic and non-explosive propellants which makes them inherently
safer and reduces the cost of development, handing, and transportation. When compared to liquid
engines, hybrid motors exhibit mechanical simplicity, reduced fire and explosion hazards, and
higher fuel density.
Figure 1.1: Classical Hybrid Configuration

9. 2
When referenced to solid rocket motors, hybrids are chemically simpler, have throttling
and command shutdown capability, and higher Isp. Unlike the regression of solid fuel grains, the
regression rate of hybrid rocket solid fuel grains are most dependent on the oxidizer mass flow
rate. This property increases the tolerance to grain flaws and allows for geometry changes over the
length of the fuel grain.
Although advantageous, these motors are confronted with several technical and non-
technical challenges. Hybrid propulsion systems lack technological maturity and will have to
compete with systems currently implemented in industry. Integration of hybrids may prove
difficult as they typically suffer from lower performance characteristics and low regression rates
when compared to liquid and solid motors respectively.
1.3 Research Goals
The objective of this research is to design, fabricate, and test 3D-printed fuel grains that
optimize performance characteristics. This project seeks to enhance solid fuel grain burn rate by
increasing the surface area exposed in hybrid rocket fuel grains. The fuel grains will be
manufactured though Fused Deposition Modeling (FDM) as opposed to traditional manufacturing
methods. As the burn rate increases, the performance of the rocket improves significantly.
Individual goals for this project are four-fold and listed as follows:
1) Verify the feasibility of replacing HTPB with 3D-printed ABS plastic.
2) Characterize the performance of 3D-printed ABS/Nitrous Oxide fuel/oxidizer
combination.
3) Develop appropriate mathematical models for AB/Nitrous Oxide regression.
4) Apply the findings of (3) in the modeling, printing, and test firing of various ABS
fuel geometries.
2.0 Design Parameters
2.1 Problem Formulation
2.1.1 Design Problem Formulation
A major disadvantage of hybrid rocket systems is insufficient regression rate. Regression
rate is dependent on oxidizer contact with the fuel grain and is limited to the inner surface area. To
control resulting thrust curves, the geometry of the fuel grain surface area can be tailored for
increasing, decreasing, and constant surface area during combustion resulting in progressive,
regressive, or neutral burning respectively. The thrust curves and their respective fuel grain
geometries manufactured by casting shown in Figure 2.1 demonstrate the effect of surface area
design. Traditional methods of casting, tapping holes, and introducing air pockets in the fuel grain
have been used in an attempt to improve regression rate. These methods introduce complications.
Casting limits fuel grain geometry to constant cross-sectional area. Tapping holes and introducing

10. 3
air pockets can compromise the structural integrity of the fuel grain and present unreliable
performance characteristics.
Figure 2.1: Fuel Grain Geometries By Casting [Braeunig]
The most favorable method for fuel grain manufacturing would be capable of tailoring any possible
geometry with structural stability and consistent results. Such a method, FDM utilizes layer-by-
layer deposition, allowing for unlimited geometry tailoring. The results of FDM vary only in
proportion to layer resolution; high resolution layering leads to consistent results. FDM is currently
unavailable for commercially used fuel, HTPB. As an alternative, FDM is available for ABS, a
viable hybrid fuel source as seen in a study by Utah State. In the study, hybrid fuel grains of 82.6
cm diameter HTPB produced a mean thrust level of 755 N. ABS fuel grains of the same size
produced a mean thrust level of 717.8 N [4].
This project will first compare the performance of ABS by FDM to tubular casted ABS
and HTPB fuel grains. Second, FDM is expected to optimize control of hybrid fuel grain thrust
curves. Third, tailoring for optimal fuel grain geometry will attempt to increase hybrid rocket
motor performance up to low end solid rocket motor performance. Finally, improving the
performance of hybrid fuel grains necessitates increased safety requirements. The combustion
chamber and test apparatus must be designed with sufficient factors of safety to prevent test failure
and ensure operation personnel safety.
2.1.2 Design Variables
Hybrid rocket motor design variables are thrust, propellant mass flow rate, chamber
pressure and temperature. For total thrust 𝐹 𝑇
𝐹 𝑇 = 𝑚̇ 𝑣2
+ (𝑝2 − 𝑝3)𝐴2 [Equation 1]
the above equation shows the first product as the momentum of the motor, propellant mass flow
rate, m, and the exhaust velocity, v2. The second product that affects total thrust is the difference
of atmospheric p3, and nozzle pressure, p2 to the area A2, at the nozzle exit. The desired output is
to have exhaust pressure equal to or slightly higher than the ambient fluid pressure. As the exhaust
pressure reaches atmospheric pressure, the right side of the equation becomes negligible, and thrust
relies on mass flow rate of the propellant with the corresponding exhaust velocity. Known as
optimal expansion ratio.
Chamber pressure and propellant mass flow rate are related by the same equation. Chamber
pressure is calculated in 𝑀𝑃𝑎 and will be measured using a hardline pressure transducer while
propellant mass flow rate is measured in
𝑘𝑔
𝑠
. The chamber pressure is related to propellant mass
flow rate by,
𝑝1 =
𝑚̇ 𝑐∗
𝑔 𝐴𝑙
[Equation 2]

11. 4
where the equation shows the product of the propellant mass flow rate m, and characteristic
velocity c*, divided by the product of gravity g, and initial nozzle throat area At. As the propellant
mass flow rate increases, the internal combustion chamber pressure increases due to increased fuel
consumption.
In the design process of a hybrid rocket, temperature is related to the regression rate through heat
transfer rate. Temperature is related to heat transfer by,
𝑄̇ 𝑠 = 𝐹 𝑇
𝜕𝑇
𝜕𝑦
| 𝑦=0 [Equation 3]
in which the heat transfer rate per unit area of the active combustion zone to the fuel surface is
equal to that conducted [5]. The heat transfer rate 𝑄𝑠
̇ , is equal to the product of the change in
temperature dT, per change in area 𝑑𝑦, to the conductivity of gas kg. The temperature per area of
fuel is proportional to the heat transfer.
Hybrid rocket motor design performance variables are regression rate, specific impulse and fuel-
to-oxidizer ratios. For regression rate,
𝑟̇ = .36
𝐺.8
𝜌 𝑓
(
𝜇
𝑥
)
.2
𝛽.23
[Equation 4]
the free stream propellant mass velocity, 𝐺, is a main factor in the relation to the regression rate 𝑟̇.
As the axial location x increases, the free stream propellant mass velocity increases which increases
regression rate. As solid phase fuel density decreases, there is an increase in the mass velocity
which increases regression rate. From equation 4, as the fuel density decreases, the regression rate
increases due to increased mass velocity. Blowing coefficient, 𝛽, is an aerodynamic and
thermochemical parameter that describes the enthalpy relationship between fuel surface and the
flame zone, as well as regression rate [6].
A characteristic operating feature of hybrids is that the fuel regression rate is typically less
than one-third of composite solid rocket propellants [7]. Hybrid rocket motors have lower
regression rates because of their combustion process and combustion port to oxidizer mass flow
rate. During the combustion process the heat transfer rate is decreased by the vaporized fuel leaving
the fuel surface during combustion. This decrease in heat transfer rate causes a decrease in
regression rate.
Specific impulse of the rocket is the efficiency of a rocket. In equation,
𝐼𝑠𝑝 =
𝐹 𝑇
𝑚𝑔̇
[Equation 5]
the total thrust FT, is divided by the mass flow rate m , and gravity g. The specific impulse of the
rocket is how much thrust the rocket generates for how much fuel is used. This relates how much
force is being produced per propellant quantity being burned.
Fuel-to-oxidizer ratio, given by
𝑟 =
𝑚̇ 𝑜
𝑚̇ 𝑓
=
𝑂
𝐹
[Equation 6]
is the mass of the solid fuel grain to the mass of oxidizer in the mixture as the propellant burns
during combustion. The mixing ratio continually changes during combustion because surface area
of the fuel grain ports and oxidizer in the fuel grain are changing. The constant change in the
mixture ratio causes the specific impulse to vary with burn time [7]. This causes the overall
performance of the rocket to be less efficient.

12. 5
2.1.3 Solution Evaluation Parameters
Manufacturing methods for hybrid motor fuel grain cannot produce thrust outputs that
compare to solid rocket motors. Casting is a manufacturing method for solid fuel grains. Casting
takes a cylindrical mold of the fuel grain with a removable casting tube down the middle of the
casted fuel grain. The oxidizer port is made once the casting tube is removed. Casted hybrid rocket
fuel grains are limited to simplistic geometries which exhibit low regression rate. Optimizing the
solid fuel grain geometry will increase regression rate and overall rocket performance. 3D printed
fuel grain geometries can be modified to increase thrust characteristics that casting cannot achieve.
Designing interstitial vacancies in the fuel grain will increase surface area and increase regression
rate.
2.1.4 Testing
A test stand and combustion chamber designed for specified factors of safety will be
manufactured to run multiple tests for the characterization of fuel grain performance. The
combustion chamber will include a removable rear bulkhead for installation of fuel grains. To
allow for multiple test runs, a phenolic lining will separate the combustion chamber from the fuel
grain, allowing for removal of a post-burn fuel grains. As a safety system, a pressure release valve
will be included in the forward bulkhead to prevent explosions. The test stand will incorporate
superstrut channels for modularity and a concrete foundation for stability. Another safety system
will incorporate a scatter shield, covering the test apparatus. The test stand will be oriented
horizontally, incorporating several measurement systems. Systems measuring thrust, combustion
chamber temperature, combustion chamber pressure, and propellant mass flow rate will be
incorporated on and around the test stand. Thrust will be measured by a load cell on the front end
of the test stand. Combustion chamber temperature will be related to measured plume temperature
by a multiwave IR meter. Combustion chamber pressure will be measured by a hardline inserted
through the forward bulkhead leading to a pressure transducer mounted on the test stand.
Propellant mass flow rate will be averaged by the addition of measured before and after weight of
the combustion chamber and the oxidizer supply. These measurement systems will be connected
to a data acquisition device, relaying data to a computer for measurement processing.
3.0 Parametric Design
3.1 Fuel Grain
3.1.1 Improving Regression Rate
Various solutions have been demonstrated to mitigate low regression rates in hybrid
rockets. The addition of oxidizer particles to the fuel grain matrix can increase burn rate due to
greater heat transfer via added surface reactions [8]. Oxidizing agents have included ammonium
perchlorate and iron oxide. Combined propellants pose similar safety risks to solid rocket systems
and diminish start-stop-restart capabilities. Microscopic particle additives, consisting primarily of
metals, can increase regression rate by promoting radiative heat flux from the grain surface [9].
These additives have include aluminum, lithium, and boron. Limitations in this method are due to
specific particle sizing in each application and combustion chamber pressure dependencies [8].

13. 6
Another method of boosting the regression rate is to increase the exposed surface area of
fuel grains, enabling more oxidizer to react at the surface [9]. Fuel grains with multiport designs
offer a way of increasing surface area without extending the length of the grains. Multiport grains
can lead to unburned portions of solid propellant and a weakened grain structure. The complex
design of multiport grains can be difficult to fabricate using traditional manufacturing methods
and may require additional support framework. Further drawbacks to multiport designs have
included requiring an injector for each combustion port and the need of a large pre-combustion
chamber [8].
3.1.2 Proposal of a Novel Method
The proposed solution is to optimize the exposed surface area of hybrid rocket fuel grains
through the use of FDM, commonly referred to as 3D printing. Components of a 3D printer
mainly consist of an extruder, print bed, filament spool, and a control system [10]. The process
of 3D printing requires converting a computer aided design (CAD) model to a standard
tessellation language (STL) file format [11].
Figure 3.1: Example of Fuel Grain CAD Model
This involves splitting the three-dimensional CAD model into successive layers of variably
numbered and spaced triangular geometries [12]. Printing can commence once the STL file is
interpreted by the 3D printer control system. 3D printing involves dispensing filament into a
heating element and extruding the resulting semi-liquid material through the extruder. Numerically
controlled motors actuate the extruder and print bed, precisely forming the designed geometries of
each layer.
3D printing offers advantages unobtainable through traditional casting methods. Support
material can be printed under ABS layers and dissolved in a lye bath to show the final geometry
[13]. This process allows flexibility in tailoring complex fuel grain geometries [14]. The precision
of 3D printing provides greater uniformity in fuel grain structure, while streamlining the
production process [4]. The material chosen to compose the fuel grain is ABS. It is a widely used
3D printing material and burns intensely when ignited in the presence of an oxidizer. Research at
Utah State University has demonstrated the viability of using ABS as hybrid rocket fuel grains and
has shown ABS to perform comparably to industry standard HTPB grains [4]. Additional benefits
of using ABS include being readily available and relatively inexpensive [14]. Fuel grains will be

14. 7
modeled to load into a 54 mm diameter commercially available hybrid rocket, with a maximum
anticipated thrust range of 500-1000 N.
3.1.3 Evaluation
ABS
The manufacturing precision of ABS, through the use of 3D printing, leads to consistent
performance. ABS is not a regulated substance and is relatively inexpensive. ABS is supplied in
spools of filament that are heated and extruded. Other chemicals are not required to be mixed with
ABS, and both ABS and the manufacturing method are readily available. Filament can be
purchased from multiple suppliers in a variety of lengths. It is limited in mixture composition but
is unlimited in sizing capabilities because of 3D printing. ABS burns intensely when ignited in the
presence of an oxidizer and is comparable to HTPB in terms of 𝐼𝑠𝑝 and regression rate. Precision
of manufacturing process promotes increased consistency between fuel grains. Unlimited fuel
grain geometries greatly increases tailorability. Fabrication time ranges from two to five hours
depending on size of the object, printing resolution, and infill. 3D printing allows the geometry of
the exposed fuel grain surface area to be tailored to meet desired performance characteristics.
Underlying layers of ABS can be structured to enhance regression rate and promote a more
uniform and complete burn.
HTPB
HTPB is a relatively safe material by itself. It burns intensely in the presence of an oxidizer.
HTPB grains are a combination of HTPB resin and a plasticizing agent such as PAPI 94 curative.
These chemicals are purchased separately which may increase shipping costs. Preliminary research
shows that the casting process requires a vacuum pump to remove interstitial air pockets and a
curing oven to set the rubber. HTPB is available from suppliers such as Aerocon or RCS Rocket
Motor Components. It is not as widely available as ABS. The performance of HTPB has been well
established by research and industry. HTPB is a very effective solid fuel grain for hybrid rockets
and is an industry standard, but is affected by low regression rates attributed to hybrid rockets.
Chemicals are mixed and poured into a mold for casting. Set time varies depending on the size of
the grain. Molds must be fabricated, and the portion of the mold that forms the core of the grain
must be removed. It is possible to construct the core mold out of styrofoam and dissolve it with
acetone. The casting process limits the tailorability of HTPB to single or multi-cored grains. HTPB
is difficult to manufacture complex internal geometries and requires considerable cure time.
3.2 Oxidizer Feed Design
3.2.1 Oxidizer
Nitrous oxide (N2O) will be the oxidizer used in this experiment and was chosen for being
non-toxic, self-pressurizing, and readily available. Its viability as a liquid propellant in hybrid
motors has been well established by research [3] and is relatively benign compared to other liquid
propellants [13]. N2O must be maintained at optimal pressure and temperature within the holding

15. 8
tank. Safety guidelines for the handling, use, and disposal of N2O have been outlined by Scaled
Composites and will be adhered to in the overall design [14].
The nitrogen within N2O serves the purpose of cooling the graphite nozzle [Newlands],
which aids in nozzle reuse for multiple test runs. The large mass flow content of nitrogen promotes
erosion rate of the fuel, and in turn, increases regression rate by exposing more oxygen to the fuel.
An oxidizer to fuel ratio by mass of 7:1 is common for burns in hybrid rockets. Oxygen content is
not greatly affected by changes in 𝑂/𝐹 ratio over the stoichiometric range, meaning it concedes
higher 𝐼𝑠𝑝 for a wider effective range in 𝑂/𝐹 ratio.
Figure 3.2: 𝐼𝑠𝑝 vs. O/F Ratio [Newlands]
N2O is subcritical at room temperature, which means the liquid and vapor phase exist
simultaneously within the tank. The liquid to vapor ratio of N2O varies with change in temperature.
N2O becomes supercritical at 309 K, so special care must be taken into consideration for launches
in high heat. A high temperature environment would necessitate a special injector.
Figure 3.3: Nitrous Oxide Phase Diagram [15]

16. 9
When N2O is subcritical, small pressure drops within the tank will produce an increase in
gas content of the mixture. This additional gas content then increases the pressure, effectively
returning steady pressure within the tank. Pressure drops at the outlet of the thrust chamber injector
indicate a change from liquid to a vapor phase will occur, so a maximum injector outlet diameter
of 1.5 mm must be used. The critical point indicates where the liquid and vapor saturation lines of
N2O incurs the largest density and pressure change. A drop from room temperature will cause a
loss of pressure in the run tank, meaning a reduction in thrust. Nitrous oxide entering the
combustion chamber through the injector is usually a lower density liquid. The low density liquid
content of N2O will be shot into the injector at high pressure due to the high vapor pressure of N2O
at room temperature of around 800 psi. This added benefit allows the thrust chamber to operate at
high pressure while still maintaining the pressure gradient. Aspirespace hybrid rockets were
designed to operate at higher combustion chamber pressures of 507 psi, which produced a higher
specific impulse. Nitrous requires a high temperature to break molecular bonds for release of
oxygen content, making it relatively benign.
3.2.2 Delivery System
A feed system will provide oxidizer to the injector of the thrust chamber. The system will
be comprised of a holding tank, valves, regulators, and routing lines [4]. The holding tank contains
the oxidizer at a prescribed temperature and pressure. A valving system will be implemented for
manual and remote cut-off of the flow. Regulators can be integrated into the system to control the
mass flow rate of the oxidizer and will cut off oxidizer flow in the event of backflow from the
thrust chamber. The oxidizer regulation system may be connected to a computer or microcontroller
for remotely controlling sequencing. Routing lines will connect all components of the feed system
and must withstand the self-pressurizing oxidizer [16].
3.3 Thrust Chamber Design
3.3.1 Design Considerations
Maximizing grain performance and obtaining accurate data, while minimizing cost and risk
of failure are integral to the design of this project. A two part test device will be developed in order
to adequately test the grains. The first subassembly, or thrust chamber, includes the combustion
chamber, an injector, and a nozzle. The combustion chamber will house the grains during
combustion with the oxidizer and is comprised of a cylindrical vessel sealed by a pair of bulkheads.
The injector distributes the oxidizer into the combustion chamber. After ignition, the vaporized
propellants will build pressure and are expelled out of the aft end of the chamber through a nozzle
that is optimized to increase motor performance.
The heat and stresses of combustion will subject components of the thrust chamber to
conditions that may cause failure if not properly accounted for. The melting point of materials
must be taken into consideration and relevant heat transfer analysis should be performed.
Materials and coatings need to be selected to minimize the oxidation of metals and erosion
of components [18]. Exposure to high temperatures and an oxygen rich environment may corrode
components, while high speed of combustion gases will deform the thrust chamber over multiple
uses.
The thrust chamber components must be designed to contain the combusting propellant
and allow pressure to build, resulting in an increase of exhaust velocity. The oxidizer entering the

17. 10
combustion chamber must exceed the pressure of combustion in order to prevent catastrophic
failure from hot gases flowing back into the oxidizer reserve. It is ideal to integrate a pressure relief
system to vent the gases, should combustion pressure exceed tolerable limits. The capability of
performing multiple tests is an important requirement of this experiment. Solid fuel grains, nozzles,
and supply of oxidizer must be replaced with minimal downtime between runs. A phenolic liner
will insulate the combustion chamber and prevent grain adhesion to the inner surfaces, facilitating
quick, repeated testing. The thrust chamber will be statically fired and must be mounted to a
restraint apparatus, which will subject the assembly to rapid accelerations and vibration.
3.3.2 Conceptual Solutions
Three solutions are available to develop the thrust chamber. The first method entails
performing testing on a pre-fabricated, commercially available hybrid rocket motor with minimal
modifications to the design. This method simplifies the design process and enables testing to be
performed on a reliable platform that integrates the thrust chamber and oxidizer reserve. A
commercial motor includes the oxidizer delivery system, nozzle, combustion chamber, bulkheads,
o-rings, and features the ability to be reloaded for multiple uses. This solution lacks a method to
relieve combustion pressure and poses a safety risk by storing the oxidizer supply near the
combustion chamber. The structural integrity of the rocket body could be compromised due to the
use of weaker, flight-weight materials, as a commercial flight rocket is not designed for repeated
static testing. This solution would not be cost effective to replace as it may employ proprietary
internal components that cannot be purchased as stand-alone parts. Attempting to implement
custom components could damage the motor as a whole. The budget may not allow for the
purchase of a replacement motor.
Figure 3.4: Commercial Rocket Configuration (Not Drawn to Scale)
The second solution attempts to mitigate some of these concerns by highly modifying a
commercial motor. A separate oxidizer Delivery system may be implemented to relocate the
reserve tank away from the thrust chamber, thus minimizing the risk of both components
simultaneously failing. A relief valve may be integrated with the combustion chamber and would
be activated if the pressure approaches a critical threshold. Though heavily modifying the
commercial rocket motor can give more freedom in the routing of oxidizer, this method still carries
some of the drawbacks that ruled out the slightly modified method. Dismantling the flight motor

18. 11
to better suit static testing may decrease its structural integrity and damage components that are
expensive to replace. Purchasing a prefabricated motor is costly, especially in the event of a failure.
The third solution addresses the deficiencies of the previous methods directly by
fabricating the thrust chamber from the ground up with cost-effective and easy to replace
components.
3.3.3 Our Solution
It is possible to fabricate a thrust chamber that is comparable with commercial performance
and meets the requirements of a repeated static experiment. The thrust chamber will be compatible
with 54mm commercially available HTPB and 3D printed ABS grains. Baseline performance of
the fabricated motor will be determined by testing HTPB and forming a comparison to similarly
sized commercial motors. The thrust chamber will be over-designed to withstand pressures of 3000
kPa, delivering a minimum factor of safety of two. Excessive combustion pressure is to be vented
out of the chamber by via the actuation of a relief valve. Finite element analysis will assist in
determining an appropriate set pressure to actuate the relief valve. The diverted propellant should
be directed away from equipment and personnel, and in a manner that will not cause a torque about
the testing apparatus.
The prototype rocket motor will be manufactured by machining a metal housing with
bulkheads. A reload system will be implemented on the aft end of the motor to streamline the grain
and nozzle loading process. It is preferred to fabricate the test bed from materials used on a live-
fire flight rocket, but this is not a primary concern due to the static nature of this experiment. It is
more important to account for the heat transfer from the hot gasses to the housing, as the
propellants of hybrid motors can reach temperatures up to 3500 K [17]. It is also feasible to apply
extra insulation to the combustion chamber by increasing the phenolic liner thickness to mitigate
these concerns.
The nozzle will be machined from graphite stock. A precedent for graphite nozzles has
been well established because of its low density and high temperature resistance [18]. Large
graphite stock machines well, but is abrasive and can quickly dull out tooling bits [18]. Graphite
is prone to brittle failure and requires high technical proficiency to prevent fracturing. Graphite
stock can be purchased from various industrial suppliers.
The oxidizer injector will be fabricated to provide an optimum amount of liquid propellant
for various grain geometries
3.3.5 Nozzle Evaluation
Purchase
Defined nozzle dimensions would simplify the grain making process and promote
consistency between tests. It is limited in terms of suppliers and sizing of length and diameters.
Purchasing a nozzle with known dimensions increases performance consistency between runs,
assuming the material does not deteriorate. If the material does deteriorate, the performance and
safety of the nozzle will be affected and may necessitate replacement. Purchasing the nozzle
reduces the fabrication process but may require modification for compatibility with the motor
casing.
Fabricate
A graphite rod is much more cost effective than buying a nozzle, but machining costs are
present. G raphite is very abrasive to tooling bits and may dull them out quickly. Rod stocks may

19. 12
be cut to length to fabricate multiple nozzles but may require a significant amount of time
to fabricate. Multiple sizes of graphite are readily available. Performance of graphite fabrication
is based on machining precision, which is determined by tolerances in the machining process and
the skill of the machinist. The material has a high temperature threshold leading to durability and
possible repeated use.
3.4 Test Stand Design
3.4.1 Design Considerations
The second subassembly of the testing apparatus is the test stand, which is used to restrain
rocket thrust chamber during live-fire testing for fuel grain characterization purposes. The test
stand includes the frame, foundation, and rocket motor restraint system [4]. A transparent scatter
shield may also be incorporated to contain debris and allow observations to be made safely. The
frame members must withstand thrust and transfer those forces to a foundation secured to the
ground. Materials used to construct the test stand are selected based on maintaining structural
integrity through numerous test runs. The minimum factor of safety for the test stand will be 5
times the maximum expected thrust output of 1000 N.
3.4.2 Conceptual Solutions
Three rocket test stand orientations are commonly used in industry and research. Two
vertical orientations include upward or downward firing exhausts, while the third orientation is a
horizontal exhaust configuration [17]. Both vertical test stand orientations require a high strength
structure affixed to the ground, in order to withstand large bending moments from the additional
height. The downward exhaust design enables a relatively simple oxidizer routing system at the
forward bulkhead and emulates test flight conditions. Upward exhaust orientation allows for the
test stand to be positioned in an excavated hole, providing added safety in the event of a failure.
This orientation requires an oxidizer Delivery system that can be difficult to route. In terms of test
stand rigidity and economic viability, both vertical orientations require additional material for
structural support, meaning an increase in overall cost.
3.4.3 Idealized Solution
In contrast to the vertical orientations, the horizontal structure is closer to the ground,
decreasing the bending moment and stresses on the members and in effect, reducing the required
members and the cost of materials. It also provides a simple oxidizer routing system via the
forward bulkhead. The thrust chamber will be mounted and secured with clamps to a linear rail
system, which will constrain the motor laterally in the thrust direction. The stand will be
compatible with thrust chambers of varying lengths and diameters by changing the clamps and
distance between them. The frame will be fastened to a solid foundation comprised of concrete, to
prevent any test stand movement. Transportation of the stand will be possible by unbolting the
frame from the foundation. High strength, corrosion resistant materials will be used for the frame

20. 13
members to combat high stress levels, vibrations, high temperatures, and possible sources of
corrosion. These measures will ensure overall test stand robustness for numerous uses.
Modifying Existing Prototype
The existing test stand prototype is a functioning proof of concept, but requires significant
modifications to the rail system and to its supports to be a viable testing apparatus. The platform
requires additional wheels to keep it from sliding off the rails laterally and moving vertically.
Components must be purchased and cut to size and existing welds must be grinded off for new
members to be welded on. The structure is very heavy and rigid, yet collapsible and portable. The
prototype's main structural elements are mostly fabricated which decreases labor and material
costs. The platform needs heavy modifications as well as the foundation. This test stand must also
be compatible with measurement devices and mounting brackets. Overall, this method has less
control over the design parameters than a ground up build.
Full Fabrication
The integration of measurement devices, restraint system, and foundation can be addressed
directly by fabricating the test stand from the ground up. This method allows the most control of
design parameters and the parts from the existing test stand can be repurposed to develop a new
stand. New components of the test stand would be higher quality than the existing prototype parts.
A linear rail system will constrain the platform and the rolling supports can be developed to fix all
motion in the lateral and vertical direction. Building from the ground up also allows for a more
rigid foundation to be made. All of the frame members and a full linear rail system will need to be
purchased. The components of the test stand are readily available in a variety of sizes and
configurations.
3.5 Measurements
Obtaining accurate measurements plays a vital role in characterizing grain performance,
and all measurement devices used for the experiment must be properly calibrated.
3.5.1 Thrust
Thrust will be measured by a load cell attached to the test stand. Important qualities of a
load cell are its effective loading range, measurement sensitivity, and cost. The anticipated thrust
range is 500 N to 1000 N and the ideal sensitivity is less than or equal to 1N. It is important to
account for oscillations of high-impact members and apply appropriate damping.
Thrust data may be obtained in two different ways. The first and least cost effective method
is to purchase a load cell that operates within the desired thrust range. The second option is to
create a force transducer using strain gauges on a metal block but may require more time to
fabricate and calibrate.
Fabricate Load Cell
The load cell will have the total thrust bearing down on it. The load cell will be mounted
onto the test stand using brackets. By fabricating the load cell, the manner by which it is mounted
can be addressed directly in the design. Cost of fabrication would include the block of metal and
machining tools, strain gauges, developing a wheatstone bridge out of wires and resistors, and the
necessary wires and connectors to the DAQ. These components are not very expensive and easily

21. 14
obtained. The strain gauges, bridge, and DAQ connections are more difficult to acquire.
Fabricating the load cell gives a degree of control of the effective range and device sensitivity.
Developing the bridge, making the device compatible with the DAQ, machining and calibration
time, all play a factor in fabrication time.
Purchasing Load Cell
Externally purchased load cells have already been tested by the manufacturer to perform
within the specified load range. Pricing depends on supplier, load capacity, and configuration of
load cell. Suppliers and variety of load cells are more limited in the higher thrust range.
Performance is very high due to quality of materials. Fabrication time is minimal and consists of
mounting the load cell to the test stand.
3.5.2 Temperature
Temperatures during testing are expected to range from 293 K to 3500 K [15] and can be
measured by an infrared (IR) thermometer or IR camera. The cost of these solutions is directly
proportional to their respective measurable temperatures, with the IR thermometer being more cost
effective than the IR camera. The IR thermometer takes measurements at a point and would be
better suited to measure high temperatures further into the plume or chamber. An IR camera can
pinpoint critical temperature sites on the verge of failure and displays important information about
heat transfer. Using both devices in tandem is preferred, and it may be possible to acquire them
from outside sources, mitigating the high cost of these solutions
3.5.3 Propellant Mass Flow Rate
The mass flow rate of the propellant will be obtained by the addition of fuel mass flow rate
and oxidizer mass flow rate. A scale will be used to weigh the combustion chamber and oxidizer
supply before and after each test run. A mass loss calculation will give the difference in weight,
and dividing the difference by gravitational acceleration and the burn time will give the average
mass flow rate for each system. To acquire an accurate burn time, a digital image correlation of
the combustion chamber will indicate the transition of oxidizer blowdown to steady state
combustion, taken as the initial time. The end time will be indicated by the extinguishing of the
combustion. Once the initial flux of oxidizer is burned through the flame and plateaus, the oxidizer
mass flow rate can be calculated using a piecewise function. The two average mass flow rates are
summed, resulting in average propellant mass flow rate. This mass flow rate will be recorded and
used in data processing.
4.0 Detail Design
4.1 Thrust Chamber
4.1.2 Thermochemical Evaluation
Before testing the fuel grains, chemical equilibrium analysis was completed for
combustion of ABS and 𝑁2 𝑂. The thermodynamic properties of ABS were obtained from the
research done by Utah State University which analyzed HTPB and ABS fuel grains. The ABS
filament available

22. 15
for 3D printing consists of 50% butadiene, 43% acrylonitrile, and 7% styrene. Table 4.1 shows the
mole fractions, heats of polymerization Δ𝑄 𝑝𝑜𝑙𝑦, the corresponding polymer heat of formation Δ𝐻𝑓
0
,
and the enthalpy contributions of the individual monomers completed by Utah State [4].
Table 4.1: Heat of Formation of ABS
Monomer Δ𝐻𝑓
𝑜
, monomer
(kJ/mol)
Δ𝑄 𝑝𝑜𝑙𝑦
(kJ/mol)
Δ𝐻𝑓
𝑜
polymer
(kJ/mol)
Mole
Fraction
Enthalpy
Contribution
(kJ/mol)
Acrylonitrile 172.6218
74.3119
98031 0.43 42.27
Butadiene 104.1 72.1020
32.00 0.50 16.00
Stryene 146.9121
84.60 63.31 0.07 4.63
ABS Total 62.63
The stoichiometric reaction for combustion that corresponds to the reduced chemical
formula from the calculated value of Δ𝐻𝑓
𝑜
is seen below. Using the mole numbers found from
chemical equilibrium and the molecular weights, the oxidizer to fuel ratio
𝑂
𝐹
is found to be 8:1.
Using this formula, properties of the combustion products including mole fractions, density𝜌 𝑝, and
gas constants 𝑅 𝑝 are found for future reference.
𝐶15 𝐻17 𝑁 + 38.5𝑁2 𝑂 → 15𝐶𝑂2 + 8.5 𝐻2 𝑂 + 15.5 𝑁2 [Equation 7]
𝑂
𝐹
=
𝑚 𝑜𝑥̇
𝑚 𝑎𝑏𝑠̇
=
𝑀 𝑛2𝑜 𝑁 𝑛2𝑜
𝑀 𝑎𝑏𝑠 𝑁 𝑎𝑏𝑠
[Equation 8]
The adiabatic flame temperature, 𝑇𝐹, is the maximum temperature of combustion achieved
adiabatically, without heat entering or leaving the system. Assuming complete combustion, the
expected 𝑇𝐹 is 3500 K. The properties of the combustion products for the expected 𝑇𝐹 are
calculated using equations 9,10,11.
𝑐 𝑝̅ = 𝑎 + 𝑏𝑇 + 𝑐𝑇2
+ 𝑑𝑇3
[Equation 9]
𝐶𝑣 = 𝐶 𝑝 + 𝑅 𝑝 [Equation 10]
𝛾 =
𝐶 𝑝(𝑇)
𝐶 𝑣(𝑇)
[Equation 10]
4.1.2 Fuel Grain
The length of the fuel grain is directly dependent on the burn time. For this application, the
desired burn time is five seconds which leads to a fuel grain length of 180 mm. The fuel grain will
be inside a phenolic liner with a thickness of 2.03 mm. This limits the fuel grain to have a maximum
outer diameter of 51 mm.

23. 16
4.1.3 Combustion Chamber
To compare the performance to those used in industry, the combustion chamber will have
an inner diameter of 54mm. The combustion chamber has a length of 200 mm. The additional 20
mm will act as a post combustion chamber, ensuring complete combustion of the reactants.
Figure 4.1 Combustion Chamber
The combustion chamber will be sealed by two bulkheads constrained with threaded rods
and hex nuts. The forward bulkhead will house the injector and pressure hardline, whereas the aft
bulkhead will contain the nozzle. Each bulkhead will have an appropriately sized O-rings that will
aid in the sealing of the chamber.
Figure 4.2 Forward Bulkhead (Front and Rear View)

24. 17
Figure 4.3 Aft Bulkhead (Front and Rear View)
The combustion chamber and bulk heads will be manufactured from 6061-T6 round
aluminum stock. This material was chosen for its affordability, strength, and ease of
manufacturing. To comply with the specified factor of safety, the chamber was analyzed using
FEA as seen in figures 4.4 and 4.5.
Figure 4.4 Thrust Chamber FEA-Displacement

25. 18
Figure 4.5 Thrust Chamber FEA-Von-Mises Stress
4.1.4 Nozzle
Once combustion is complete, the nozzle converts the chemical energy produced into
kinetic energy. For performance efficiency, supersonic flow is achieved through the design of a
converging-diverging nozzle. The flow at the throat will reach Mach 1 and will be supersonic in
the diverging section.
The thermodynamic properties of the combustion produces as defined above, were used to
calculate the initial velocity, Mach number, and initial area of the nozzle. Assuming the nozzle is
ideally expanded, the expansion ratio 𝜖, is 2.3. This relates the nozzle exit area to the area at the
throat as seen in Figure 4.6.
Figure 4.6 Nozzle Design

26. 19
4.2 Oxidizer Delivery System
An aluminum N2O tank with a 10lb. capacity will comprise the holding tank and features
a high-flow valve with a siphon tube to aid in liquid N2O removal [19]. Stainless steel braided
hoses of 1/4in. diameter will be used to form the routing lines connecting the components [20].
Prevention of flowback in the feed line will be accomplished with a brass one way male to male
check valve of 1/4in. diameter, which is capable of withstanding up to 3000psi [21]. A 2-way
normally closed Gem solenoid valve with a 900psi rating will be used to remotely actuate oxidizer
release and shut-off [22]. Compatibility between fitting sizes will be accomplished through the use
of adapters [23]. T-junctions will function to connect the the solenoid valve to the routing lines
[24].
Figure 4.7 Oxidizer Feed System
4.3 Test Stand
4.3.1 Fabrication Process
The process of manufacturing the test stand will involve bolting channeled members and
welding frame members. Before and after each part is assembled, measurements will be made to
ensure accurate geometry. The test stand factor of safety of 5 contributes to the tolerance for error
in geometry. The test stand is divided into three subassemblies: the rail frame, the platform, and
the superstrut bed. The rail frame is the outer structure of the stand and is fabricated completely
out of 1-¼ x 1-¼ square steel tubing welded together. Two vertical rectangles members will
support the horizontal rails on which the platform will roll along. The rails are oriented at a 45°
angle to allow two wheels to constrain the platform at each corner opposed to three in previous
designs. The rail frame will be bolted to pre-formed concrete slabs Provisions for load cell
integration will be implemented onto the rail frame.

27. 20
Figure 4.8: Rail Frame Design
The platform consists of 1-¼ x 1-¼ square steel tubing welded to form a rectangle that fits
within the rail frame. 2” sections of tubing will be welded at a 45° angle at each of the corners.
Two holes will be welded into each of the 2” sections and bolts will be welded into the holes. The
upper set of wheels will be fastened to the bolts, then the stand will be placed onto the rail frame.
The second set of wheels can then be fastened to constrain the platform to the rail frame. A
provision for the hardline pressure transducer will be made onto the platform to enable the system
to move with the thrust chamber.
Figure 4.9: Platform Design
Figure 4.10: Platform and Wheels- Cross Section

28. 21
The superstrut bed consists of two 1-⅞ x 1-⅝ lengths of superstrut welded onto the axial
members of the platform and clamps to restrain the thrust chamber. Two lateral lengths of
superstrut will be bolted onto the axial superstrut lengths and the clamps to restrain the thrust
chamber. This restraint system will be compatible with thrust chambers of varying lengths and
diameters by changing out the clamps and distance between them.
Figure 4.11: Superstrut Bed with Clamps
Figure 4.12: Assembled Test Stand- Isometric View

29. 22
Figure 4.13: Assembled Test Stand- Top View
Figure 4.14: Finite Element Analysis of Rail Frame- Displacement
4.3.2 Scatter Shield
For the safety of operating personnel and testing apparatus, a scatter shield will be
manufactured to surround the test stand, capable of withstanding shrapnel and debris. In order to
observe the test stand during testing, the material of the scatter shield will include acrylic sheets
and metal corner braces covering a 48” x 24” x 18” space over the test stand. Acrylic will be clear
and 0.22” in thickness, covering the forward, sides, and top areas. 5” corner braces will be used to

30. 23
assemble the acrylic sheets and will ensure rigidity. The forward acrylic sheet will be cut for the
entrance of the oxidizer routing and measurement cables.
4.4 Data Acquisition
The signals from the load cell, pressure transducer, and IR meter will simultaneously feed
into a DAQ that will relay the data via USB cable to a computer running LabVIEW software. The
DAQ may be housed on the test stand and wiring must be able to withstand high accelerations and
intense vibrations to prevent loss of contact due to testing. The DAQ will also provide voltage
excitation to power the load cell and pressure transducer. LabVIEW software will be used to
compile the data and an appropriate sampling rate must be selected to minimize aliasing. A virtual
instrumentation subroutine, or VI, will collect the raw voltage outputs from the sensors and record
them for post processing. The sensors must be calibrated before testing by relating the voltage
outputs to known quantities. Calibration data will be processed on Microsoft Excel to perform
statistical analysis, formulate visual representations, and develop models for the interpolation and
extrapolation of data. The data can then be exported for post processing and used to make future
adjustments in the grain geometry. The figure below displays the measurement apparatus.
Figure 4.15: Measurement Apparatus
4.4.1 Load Cell
Manufacturing a load cell in house is advantageous because of design intended for this
project’s specific application. Drawbacks arise with the application of a strain gage to the load
cell’s surface. Strain gages are isolated to measurement in one strain direction while deflection in
any other direction is not observed by the strain gage. This requires the strain gage to be applied
to a manufactured load cell under critical tolerances. Loading applied in a direction non-
perpendicular to the load cell’s surface will also not be measured by the strain gage. To ensure the
uniformity of the load cell’s geometry and the accuracy of its strain gage, an externally purchased
load cell from Omegadyne, Inc. will be included in the measurement apparatus. The LCCD-500
S-beam load cell has a maximum capacity of 2224.11N with accuracy of ±0.25%. A five point
calibration procedure will be performed before operation. In a study at Utah State, the same load

31. 24
cell was used in a similar application testing tubular casted ABS fuel grains. The thrust levels
increased from zero to over 800N within 1s, providing validity to the response time of the load
cell [26]. A hysteresis of ±0.02% also contributes to the load cell’s sufficiency for dynamic
loading. The ultimate overload of 300% maximum capacity extends beyond the factor of safety of
five for the test stand design, ensuring safety and reliability of the measurement system. For
integration to the test stand, a 1/2"-20 male thread bolt will be used to join the load cell to the
superstrut cross member at the front end of the test stand. The 24 AWG jacketed cable containing
excitation and output wiring will be connected to the DAQ for 10V excitation and 3V output [27].
4.4.2 IR Meter
Externally purchased IR meters within this project’s budget are not capable for measuring
maximum combustion chamber temperatures. To compromise, measurements of a single point of
the plume will be related to the combustion chamber’s internal temperature via the following
equation:
𝑇𝑜2 = 𝑇2(1 +
𝛾 𝑝−1
2
𝑀𝑒
2
) [Equation 10]
in which 𝑇𝑜2 is the adiabatic internal combustion temperature, 𝑇2 is the exit temperature, 𝛾𝑝 is the
specific heat of the mixed propellant, and 𝑀𝑒 is the exit Mach number. The measurements will be
collected through a DAQ and sent to a computer for data processing, incorporating the equation
above.
An Ircon UX-40P will be loaned for the purposes of this project as a sufficient IR meter
for plume temperatures well above 473 K would be very expensive. The meter covers a spectral
region of 8 mm to 13 mm and is capable of measuring a maximum temperature of 1273 K. The
meter is able to be placed at a safe distance from the plume. An optical resolution of D/40 and
focus distance of 700 mm or larger. The accuracy of the reading is ±1.0% or ±2 K, whichever is
greater. The resolution of the meter is 1 K. The response time of 1.0 sec will be considered during
the burn, as the measurements will be delayed. The analog output of 1.0 V will be connected to
the DAQ for data acquisition. The meter is powered by 4 type AA batteries requiring no excitation
voltage [27].
4.4.3 Pressure Transducer
The expected internal combustion chamber temperature is above the maximum operating
temperature of a pressure transducer within this project’s budget. To decrease the temperature of
the combustion gasses, sufficient heat transfer through the hardline is required for cooling to 398
K. A two dimensional heat transfer solution was implemented to determine the length of the
hardline necessary for this cooling. Considering radial and axial heat transfer, a nodal analysis was
performed with nine nodes with an initial temperature at the first node. Each node incorporated
radial convection with ambient temperatures and axial conduction through the stagnant
combustion gasses present in the hardline. Using the mass percentage of the combustion products,
a conduction coefficient was formulated for the total homogeneous combustion gasses as a
function of temperature. The function of temperature was interpolated from known conduction
coefficient values of the individual combustion products from Fundamentals of Heat and Mass
Transfer [27]. Creating an array of heat transfer equations for each node, a Newton Raphson solver
was used to solve for the axial temperature distribution for the hardline. This solution allowed for
specification of hardline diameter and length, sufficient for the desired heat dissipation. See the
appendix for calculations. The result of the solution specified a 0.12 m hardline of 3.175 mm

32. 25
diameter for heat dissipation reducing the combustion gas temperature at the pressure transducer
to 374.526 K.
A MLH01KPSB06A sealed gauge male, 3.175mm 27 NPT port pressure transducer made by
Honeywell manufacturing will be purchased from Digi-Key Electronics. The operating pressure
of the pressure transducer is 0 to 6.894 MPa with a standard output of .5 to 4.5Vdc ratiometric at
.5% accuracy at Full Scale Span (FSS). The FSS is the algebraic difference between the output
signal measured at the maximum pressure and minimum pressure limits of the pressure range [30].
Excitation voltage is 5Vdc with a response time of less than 2ms. The error of the pressure
transducer is the Total Error Band of ± 2% FSS. This pressure transducer is rated for 10 Hertz to
2000 Hertz of vibration which is more than what is expected during testing. With accordance to
MIL-STD-810C, Figure 214.2-5 curve AK, this pressure transducer can withstand a minimum of
20.7 G-rms [29]. The operating and compensated temperature range of this pressure transducer is
233 to 398 K. A pressure transducer like this has been used in ABS hybrid rocket testing by Utah
State University [4]. Using a pressure transducer like Utah State, chamber pressure plot is expected
to be similar to Figure 4.15.
Figure 4.15 Pressure vs Burn Times
Hardline Tube
A ⅛ outer diameter stainless steel hardline tube will connect to the ⅛ diameter pressure
transducer port to dissipate the combustion chamber heat. The tube will be inserted into the forward
bulkhead of the combustion chamber and placed in the middle of the combustion chamber. The tip
of the tube in the combustion chamber will be angled 90° to acquire the static pressure of the
chamber. The stainless steel tube will be coated with LOCTITE Mil Spec Silver Grade Anti-Seize
to protect the tubing from high heat temperatures. This coating will increase the resistance of the
stainless steel by 1143 K. This insulator coat will prevent the stainless steel from melting and
warping. Any failure in the stainless steel tubing will show error in the pressure readings. During
combustion, pressure will increase in the tube and will be read by the pressure transducer. The
stainless steel tube has a melting range of 1673-1723 K which would hold for the five second burn
time of the combustion chamber. After each run, the tube will be inspected for any signs of failure.
Extra hardline tubes will be bought to replace any failing hardline tubes.

33. 26
4.4.4 Combustion Chamber/Oxidizer Supply Scale
Initial and final weight measurements will be conducted for the combustion chamber and
the oxidizer supply. Both the combustion chamber and oxidizer supply will be removed from the
test apparatus and weighed on an externally purchased digital scale. The accuracy of each
measurement will be ±0.1 lbs, relating to ±0.045kg [28]. Initial weight will be subtracted from
final weight and divided by the burn time to give propellant mass flow rate.
4.4.5 DAQ
Once the oxidizer is ignited and the fuel grain burns, the pressure from the combustion
reaction will be recorded through the hardline and to the pressure transducer. To acquire data, the
pressure transducer and load cell will be connected to a NI USB-6008 12 bit, Low-Cost
Multifunction DAQ. The DAQ has 8 analog inputs and 2 analog outputs with a max sampling rate
of 10 kS/s. A single ended input and working range of ±10 Volts. The DAQ’s working range of
10V would power the pressure transducer’s operating voltage of 5Vdc and the load cell’s operating
voltage range of 10 - 15V. The measurements from the pressure transducer and load cell would
be sent to the DAQ and then sent to the computer in LABVIEW.
4.5 Control System
The sequencing of the experiment will be controlled by a Tiva C Series TM4C123G digital
signal controller. The Tiva is simple to code and is compatible with a vast library of open source
Arduino programs.
Figure 4.16: Tiva™ C Series TM4C123G LaunchPad [Texas Instruments]
The Tiva will time the burn, signal a solenoid valve to open and close the valves, and prime
the ignition source. An emergency stop function will be added to the code to close the solenoid
valve in case of failure. A check valve will be implemented to permit oxidizer flow to the thrust
chamber and inhibit backflow into the holding tank. Pressure relief valves on the thrust chamber
and holding tank will vent gasses that reach excessive pressures.

34. 27
Figure 4.17: Control System Diagram
5.0 Bill of Materials
Thrust Chamber
Component Unit Price ($) Quantity Cost
Injector $60.00 1 $60.00
4" Round Aluminum Stock $116.00 1 $116.00
Threaded Rod $9.50 2 $19.00
Hex Nuts $0.20 22 $4.40
Medium Extruded Graphite Rod $27.00 1 $27.00
O-rings $5.00 12 $60.00
Phenolic Liners $8.00 6 $48.00
HTPB $90.00 1 $90.00
Total $424.40
Oxidizer Delivery System
10-pound Aluminum Nitrous Bottle $230.00 1 $230.00
Nitrous oxide 65lb. tank $135.00 1 $135.00
Stainless Steel Braided Hose with -4AN Blue Fittings $43.16 1 $43.16
Inline NPT Check Valve $37.81 1 $37.81
Gem Solenoid Valve 900-1000psi $120.00 1 $120.00
Blue Anodized Aluminum -4AN to 1/8" NPT Straight Flare To Pipe Fitting $6.83 1 $6.83
AN Male to AN Female Swivel on Side T-Fitting $18.99 1 $18.99
Tiva™ C Series TM4C123G LaunchPad $12.99 1 $12.99
$604.78

35. 27
Test Stand
Radial Bearings 0.3750 x 0.8750 x 0.2812 in. $0.99 8 $7.92
Plain Steel Square Tube, 1-¼ in. x 1-¼ in. 1/16 Thickness $0.32 200 $64.00
Concrete Pad, 16 x 16 x 4 in. $4.37 4 $17.48
Total $89.40
Measurements
MLH01KPSB06A Pressure Transducer $110.00 1 $110.00
Stainless Steel Hardline Tube ⅛" Diameter x 6' $31.71 1 $31.71
LOCTITE Mil Spec Silver Grade Anti-Seize, 8oz brushtop, $19.99 1 $19.99
Digital Scale $39.99 1 $39.99
NI DAQ USB-6008 $199.00 1 $199.00
Total $400.69

36. 29
Appendix
Figure 6.1: Conduction Coefficient Formulation

37. 30
Figure 6.2: Temperature Equation Array
Figure 6.2: Temperature Distribution Per Node (Temp in K)

38. 31
References
[1] ASME Publications Committee, 1995, “Ethical Obligations of Authors,” from
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