MaCH-SR1:
Improving the Thrust-to-Weight Ratio of the MaCH-SR1
Hybrid Rocket Engine
Kevin Corcoran, Laurren Kanner, Alejandro Reiman-Moreno, and Joshua A. Stamps
University of Colorado, Boulder, CO, 80309
The MaCH-SR1 program is a project that aspires to produce a hybrid launch vehicle for
the University of Colorado. It is an ongoing project that is developed further each year by a
team of senior aerospace engineers in conjunction with the senior project program at the
University of Colorado in Boulder. This year, the team is working to improve the thrust-toweight ratio of a proven hybrid rocket engine design by decreasing the engine mass by more
than 30%. In order to isolate this objective, the proven 5,000lb f peak thrust engine has been
downsized to a 300lbf lab-scale engine. In total, four engines were designed, fabricated, and
tested including two baseline engines and two mass reduced engines. In result a mass
reduction of over 60% was achieved with no cost to thrust production by re-designing both
the nozzle and injector housings and by utilizing more appropriate materials for the nozzle,
chamber, component fittings, and injector plate. To compliment this improvement in the
overall thrust to weight ratio and to allow for future teams to optimize the fuel grain design,
direct regression measurements and fuel surface was also obtained during the combustion
process.
Nomenclature
T/W
O/F
TC
PC
=
=
=
=
Thrust-to-Weight ratio of the engine
Oxidizer to Fuel ratio
Chamber Temperature
Chamber Pressure
I.
Introduction
T
HE MaCH-SR1 project has been developed every year by a team of senior aerospace students at the University
of Colorado since the 2001-2002 academic school year. This report begins with a summary of the overall vision
of the program, the work that has already been done by previous teams, and finally the significance of the
contributions by this years’ team.
The focus of this paper, however, is to present the work done by the 2004-2005 team of students. The primary
objective was to improve the thrust-to-weight ratio of the proven engine design so as to allow future teams to
embark on the design of the rocket’s flight components. What this years’ team inherited was a heavy stainless steel
engine that could reliably produce an average of 2000lb f thrust and a maximum of 5000lbf. Before considering this
engine capable of flight, it’s thrust-to-weight ratio must first be increased from 7.8, to a value greater than 10. A
dynamic approach was taken in improving this performance metric as the proven design has been optimized at a
systems, component, and interface level in order to improve the thrust-to-weight ratio by more than 30%. This
improvement was achieved in the design and fabrication of two alternate engines. At a systems level, an aluminum
engine and a composite fiber engine have replaced the heavy stainless steel engine. At a component level, the
graphite nozzle has been replaced by a castable ceramic nozzle, the steel injector plate has been re-designed as a
titanium plate, and both phenolic and cera-fiber insulation have been utilized to allow for the use of aluminum and
composite fiber materials. The interfaces have also been optimized as flange interfaces have been replaced with
threaded adapters.
Colorado Space Grant Consortium
1
Error! Not a valid link.
As a secondary objective, this years’ team also created a foundation for future teams to continue reducing the
mass of the engine by acquiring combustion characteristics of the fuel for the first time such as fuel regression data
and fuel surface temperature data. With this data, future teams can optimize the fuel grain design further by
optimizing both its port geometry and its overall size.
II.
MaCH SR1 Project Overview
A. MaCH SR1 Vision
The MaCH-SR1 project was begun by a group of students at the University of Colorado at Boulder in 2000. This
team setup an ambitious goal: to give the University of Colorado, and its associated research labs, its own suborbital launch vehicle. To accomplish this ambitious task, a 5-stage method of approach was established, to guide
future teams to this ultimate goal:
Phase 1: Feasibility & Technology Base Investigations
Phase 2: Low-Powered Hybrid Rocket Applications (to 1,000 lb)
Phase 3: Mid-Powered Hybrid Rocket Applications (to 5,000 lb)
Phase 4: Flight Testing of Low to Mid-Powered Hybrid Rocket Applications (300 to 5,000 lb)
Phase 5: Flight Testing of High-Powered Hybrid Rocket Application, MaCH-SR2 (13,000 lb)
The goal of the project was to obtain sub-orbital flight (110 km) by 2010: only ten years to a completely studentdeveloped launch vehicle for exclusive use by the University and its associates.
B. MaCH SR1 Past
Currently, the project is on the cusp of breaking into Phase 4. In the 2000-2001 year, rocket feasibility studies
were done, and initial design was created, and the hybrid rocket fuel was developed. The original fuel was a mix of
HTPB and IPDI, with a Liquid Oxygen (LOX) oxidizer, and was used in small 100 to 300 lb lab-scale engines. The
next year’s team (2001-2002) proceeded to develop the rocket into the 1000 lb engine, thus completing Phase 2 of
the project. The 2nd team switched the oxidizer to high-purity Nitrous Oxide (N20), due to the complexity and
expense of LOX systems. The 2002-2003 team undertook the immense effort of developing the 5,000 lb engine.
With this, the University of Colorado’s MaCH-SR1 project became the largest undergraduate-developed rocket
engine in the world! Phase 3 was successfully completed with the static testing of the 5,000 lb engine in the spring
of 2004.
C. MaCH SR1 Future
Looking forward, the next step in the project is flight-testing. The development of a flight-ready vehicle, which
is safe yet effective, is a monumental task for any team: let alone a group of undergraduates. Phase 4 is perhaps the
longest phase of the project, as it encompasses flight development of multiple sizes of engines, and a huge jump in
technology. The rocket engines of Phase 4 must be completely self-contained in power, monitoring, control, and
communication. The flight vehicles must also utilize new and more advanced technologies than previous attempts in
order to be economically feasible: which means safety becomes a huge issue. For the flight vehicles, the designs can
no longer contain large factors of safety, which decrease the rocket’s efficiency. Streamlining the designs of the past
will be one of the greatest challenges to moving forward.
D. MaCH SR1 Present
At the official end of Phase 3, the project was handed down to the current team. The next step was to obtain
flight data for the MaCH-SR1 rockets, but using the heritage designs was out of the question. They incorporated
heavy steel casings, carried excess fuel, used commercial off-the-shelf (COTS) products with limited design
efficiency, had no consistent design for factors of safety (FOS), and contained no consideration of aerodynamic
design. It was obvious that before Phase 4 could begin, a design that considered efficiency was necessary. The
current team began investigations of which of these problems was really holding back the onset of Phase 4. While
the self-containment was the most obvious item to tackle before anything could be launched, the real issue was the
weight. The previous design had a thrust-to-weight ratio (T/W) of 7.8. There is virtually no known database of mass
estimates for hybrid rockets, and so it is not well known what kind of T/W ratios can actually be achieved.
However, the average payload rocket has an overall T/W ratio of 1.5 (Humble, 18) including its engine, feed system,
aerodynamic shell, control, ignition, telemetry and of course it’s payload. Before embarking on the design of each
of these commodities it is felt that the engine mass must be further reduced. Thus, the goal of this year’s design thus
became to increase the T/W by way of decreasing mass. This allowed for the fuel configuration to stay the same, so
that the progress was measurable against the baseline of the previous designs.
Colorado Space Grant Consortium
2
Error! Not a valid link.
III.
Thrust-to-Weight Ratio Reduction
The 2004-2005 MaCH-SR1 team took an ambitious approach for minimizing the T/W ratio of the inherited
hybrid rocket engine design. This idea was to reduce the overall weight of the engine without any cost to its thrust
performance. The proven engine design that was inherited from past teams is a stainless steel structure that is
grossly over-designed in terms of structural support. The objective of past teams was to develop a fuel grain and
oxidizer combination that could reliably produce thrust and then increase its size to prove scalability. For this
reason, little attention was paid to developing a structure that is optimized for low mass, and hence fly-ability.
The inherited engine is designed to produce 300lb f of thrust for 15 seconds with HTPB fuel and Nitrous Oxide as
its combustants. The combustion process produces a chamber temperature in excess of 5000ºF and pressure of
500psi. So that it can be verified that any mass reductions obtained by this years team did not come at the expense
of performance, these same metrics have driven the design of all components for this years’ engines well.
The inherited engine will be presented followed by the new engines designed by this years’ team. Finally, the
reductions of mass that result in over a 60% mass reduction will be analyzed.
A. Inherited Engine vs. New Engines
1.
Inherited Design
This engine inherited from past MaCH-SR1 teams is illustrated in figure 1, and will be referred to from here on
out as the baseline engine. Only the top-level design of this baseline engine will be presented as the focus of the
paper is to present work done by this years’ team. There are six main components of the baseline engine: the
oxidizer adapter, injector, pre-combustor, fuel grain, post-combustor, and the nozzle. The purpose of the oxidizer
adapter (labeled oxidizer interface in the figure) is to connect with a feed system that supplies liquid Nitrous Oxide
(oxidizer), gaseous Oxygen (Ignition Source), and gaseous Nitrogen (Extinguishing Source) to the engine. These
gases pass though the injector plate, which is a conflate flange with 14 holes in it sized appropriately for the liquid
oxidizer to expand and gasify as it enters the chamber. The injector is not labeled in the figure but is the flange on
which the oxidizer adapter is welded.
After passing through the injector, the oxidizer first reaches the pre-combustor of the
chamber where the fluid is fully gasified and pre-heated before entering the fuel port of the
fuel grain. The fuel grain has a central port that is shaped so as to allow for a constant burn
area as the fuel combusts with the oxidizer. If the fuel port were circular, the burning
surface would increase as the fuel surface regresses outward. In an ideal world, the oxidizer
to fuel ratio (O/F) should remain constant throughout the burn in order to achieve a constant
thrust. What this means is that either the fuel port would have to be shaped so that its
exposed surface area remains constant as the fuel surface regresses, or that the oxidizer flow
is controlled so as to keep up with the changing surface area of the exposed fuel. However
the later approach is much more complicated as it a difficult task to monitor the real-time
Figure
1:
surface area of the fuel and vary the oxidizer feed rate simultaneously. In this engine the
Baseline engine
fuel port is shaped as a cross, the geometry of which will be expanded on in the propulsion
section view.
section of this report.
After the oxidizer enters the fuel grain and begins combusting with the solid fuel, the resulting mixture then
enters the post-combustor. The post-combustor is necessary due to the existence of competing boundary layers that
exist in the fuel port. As a result of these boundary layers, concentrations vary perpendicularly to the fuel surface.
The post-combustor is a mixing area for the gases to complete the combustion process before exiting the nozzle.
The mixed exhaust then exits the nozzle to provide thrust to the engine. The purpose of the nozzle is to regulate the
flow of the combusted gases in order to provide maximum thrust. A converging-diverging nozzle effectively
compresses the exhaust gasses to exactly sonic speeds at the throat of the nozzle and then increases the speed further
to supersonic speeds with the diverging part of the nozzle.
2. New Engine Design
In reducing the mass of the baseline engine, several optimizations were made. The most obvious way to reduce
mass was to find a more optimal material to base the structure on rather than stainless steel. Several materials were
considered before converging on the idea of using aluminum. Aluminum was found to be the most ideal material
considering its low weight, high strength, and low cost. Composite fiber was also selected for similar reasons.
However, composite fiber is less dependable than aluminum in regards to its fabrication. The physical properties of
a resulting carbon fiber structure are highly variable depending on the quality of the fabrication process. In result
two engine designs were pursued. The first would be completely made out of aluminum and the second would have
Colorado Space Grant Consortium
3
Error! Not a valid link.
a composite chamber with the rest of its components being made from aluminum. Both engine designs accomplish
the project objective of a 30% reduction in mass. However, the composite engine presents a more promising
foundation for further development, as fabrication processes can be perfected such that its high strength, low weight
properties are fully taken advantage of. Each of these new engines are shown in figure 2 along side the baseline
engine for comparison.
Figure 2: Overview of engines from left to right: the
Aluminum Engine, Composite Engine, and Baseline
Engine
As can be seen in the figure, along with optimizing the base material of the engine structure, each of the
components was modified as well. The bulky steel flanges of the baseline engine were replaced with threaded
aluminum housings. The injector plate itself was changed from being the center of a steel flange to a much lighter
titanium plate. The nozzle, which was oversized in the baseline engine so as to fit snuggly inside the chamber, has
been replaced with a ceramic nozzle that is not only lighter but also much easier to manufacture. The oxidizer
adapter was replaced with an oxidizer cap. This is because aluminum cannot be as reliably welded as can steel and
so the oxidizer adapter had to be increased in size so as to be bolted to the injector assembly.
Each of these components will be expanded on in the following sections.
B. Chamber Materials
An extensive trade study was done on potential materials for the chamber of a reduced mass engine. Several
types of steel were examined including cold drawn and hot rolled 1018 steel, cold drawn 1019 steel, and 347
stainless steel. Also considered were 6061, 7475, 2011, and 5052 aluminum alloys as well as Grade 1 Titanium and
3k x 3k carbon fiber composites. The requirement for these materials is that the resulting chamber could withstand
pressures of up to 500psi, which is the expected and empirically verified chamber pressure during combustion of
HTPB and nitrous oxide. In result both aluminum and composite fiber materials were selected for alternate engine
designs. All materials provided adequate strength. In comparison all grades of steel were found to be much too
heavy, and titanium too expensive. Aluminum 6061 was the aluminum alloy of choice due to its low cost and
reliable availability. Carbon fiber composites were found to be the most appealing of all materials in question as it
carried the highest tensile strength (545ksi), and the lowest density (0.064lb m/in3). However, the manufacturability
of composite carbon fiber technology left enough doubt such that the aluminum chamber would be the prime focus
of this years’ team and a composite chamber would simply be pursued as a means to demonstrate technology for
future teams to develop further. In conclusion, it was determined that two chambers would be created: one
completely made from aluminum and the other a hybrid of aluminum components and a composite fiber chamber.
The main design obstacle in using aluminum and composite fibers is insulating these materials from the harsh
temperature conditions found in the combustion zone of the engine. The chamber temperature of the 300lb f baseline
engine has been seen to increase beyond 5000ºF. Meanwhile aluminum 6061 melts between 1000-1200ºF and
composite fibers melt at similar temperatures that are highly variable and not well documented. In order to
eliminate the risk of these materials melting a tube of 1/8th inch thick phenolic has been used to insulate them. This
is convenient for the fabrication of the composite chamber as it allows for a surface around which to wrap sheets of
carbon fiber. For both chambers, however, it increases their inner diameter by a quarter inch with respect to the
3.75in inner diameters of the stainless steel baseline engines. Each of these chambers are shown alongside the
baseline engine chamber below in figure 3.
Colorado Space Grant Consortium
4
Error! Not a valid link.
Figure 3: Chamber Comparisons
The design requirements for each chamber were similar to those driving the baseline engines. The chambers
must be able to withstand the combustion pressure of 500psi for a 300lb f rocket engine. They must have an inner
diameter of 4in, which is driven by the outer diameter of insulation required to isolate the chamber walls from the
high temperature combustion zone. They must also be 19.8in long, which accommodates the 17.8in long fuel grain
as well as a 2in post-combustion length. The baseline engine had to also account for the height of the nozzle and so
the baseline chambers were 24 inch long. In order to determine the required thickness of each of the new chambers,
thin-walled pressure vessel equations were used. A factor of safety of 2 was followed for the design of the
aluminum engine and 3 for the composite engine. The higher factor of safety for the composite chamber is due to
the uncertainties associated with fabricating carbon fiber composites. For the aluminum chamber, this meant a wall
thickness of 0.063in was required to withstand 1000psi and for the composite chamber a wall thickness of 0.054in
was necessary to withstand 1500psi. The thickness of the aluminum chamber was increased to 1/8 th of an inch due
to the difficulty in machining aluminum any thinner. As a result the actual factor of safety of the aluminum engines
increased to 2.7, indicating the chambers would rupture at 1350psi. The fabrication of a composite fiber chamber
did not present this problem as a 0.054in thickness simply corresponds to six wraps of a fiber sheet. However, in
adding the epoxy-resin to these wraps, the thickness of the composite system increased by about 25%.
C. Interfaces
In order to reduce the weight of the rocket it was easiest to first focus on the fitting mechanisms used to attach
the injector and nozzle fittings. This area was the heaviest and therefore had the most weight that could be lost. In
the base line engines these fittings were SS 304 conflate flanges. These fittings are very heavy and bulky as they
weigh about 8.5 lbs per pair and protrude over an inch outside of the combustion chamber. When compared to the
new design this is very heavy, the new injector fitting weighs only 1.7 lbs and the nozzle fitting only weighs 1.4 lbs
and only protrudes 0.2 inches past the outside of the combustion chamber. Taking a step back to the drivers of the
design and attempting to find solutions to those original problems rather than simply trying to improve the existing
design allowed for this years’ team to achieve such drastic changes in volume and mass.
When designing the new fittings it was decided that they should protrude as little as possible outside of the
combustion chamber in order to reduce the aerodynamic profile of the chamber. This excluded the use of an
external bolted flange. An additional constraint on the fitting was that the new design had to provide access to the
full diameter of the inside of the combustion chamber. This was necessary in order to slide the glass phenolic
insulation into the chamber and to ease fuel casting. This led to the option of either using a threaded fitting with oring seals or an epoxy bonded fitting that provides its own sealing. At first, based on simplicity and machinability, it
was decided that the fitting could be bonded to the ends of the chamber after the phenolic and fuel had been inserted
Colorado Space Grant Consortium
5
Error! Not a valid link.
and cast into the chamber. Not having to machine threads and o-ring glands made this design very attractive. It was
later decided that inserting the igniter into the fuel port would be too difficult if the fittings were not removable, this
lead to the use of the threaded fittings using o-ring seals.
Figure 4: Fitting Comparisons. (A) Baseline flange interface of oxidizer adapter with
the chamber. (B) Baseline flange interface of nozzle with the chamber. (C) New
engine threaded injector interface. (D) New engine threaded nozzle interface
In order to reduce weight it was decided that the nozzle fittings would be conical in order to minimize the
inefficient use of material in the form of a flat plate. The conical shape absorbs the load of the internal pressure in
tension rather than in shear as in a flat plate. The design of flat fittings is driven by deformation rather than by
stress. This means that if the fitting where simply designed based on stress calculations it would most likely deform
so much that the seal to the chamber would be unreliable. For this reason the design would have to be deformation
driven in order to produce a viable part, which would require much more material and therefore mass. In the book
by Dennis Moss called Pressure Vessel Design Manual (Moss, 2003), Moss shows that modified hoop stress and
axial stress equations can be used to design the conical portion of the fitting. These equations can be seen in
equation 1 and can be compared to the standard hoop stress and axial stress equations in Equation 2 (Vable). This
allowed the structure to be optimized for strength instead of for deformation as is required for the conflate flange.
As an additional benefit the fitting structure reduced the amount of ceramic material required for the nozzle. Using
these cone equations it is possible to design a fitting that is only marginally thicker than the combustion case.
Equation 1
Colorado Space Grant Consortium
6
Error! Not a valid link.
Equation 2
Due to the required interface with the oxidizer adapter it was not possible to optimize the oxidizer fitting in the
same way as the nozzle fitting. At first the conical shaped fitting was attempted, and a smaller and thinner injector
plate was designed to drive the shape of the fitting. The full weight saving could not be achieved due to the extra
material required to attach the oxidizer cap using bolts. This approach still brought weight savings by reducing the
volume of material required for the injector plate which requires high strength, high temperature tolerant material.
This allowed the use of aluminum in place of the SS 304 of previous years, which is much less dense. In addition
since the injector plate was so much smaller it was possible to purchase titanium instead of stainless steel for the
injector plate, further reducing mass.
The drastic changes in mass mentioned previously would not have been possible had the design process not
taken a step back from the existing design and analyzed instead the drivers of the project.
D. Injector Plate
The purpose of the injector plate is to provide the correct mass flow rate of oxidizer to the engine and to maintain
the desired pressure gradient between the liquid oxidizer feed and the chamber. The baseline injector plate was part
of a flange that was welded to the baseline chamber and then bolted to another flange with the oxidizer adapter
welded to it. In replacing flanges with threaded adapters, the team was presented with
the opportunity to design the injector plate as a separate piece with a reduced size that is
made from an optimized material. To begin a trade study was conducted on materials
for an injector plate. After selecting an optimum material, the thickness of the plate
was determined so as to withstand the pressure gradient between 800psi liquid nitrous
and a 500psi combustion chamber. Finally the hole pattern was designed so as to
provide the required oxidizer flow rate of 1.06 lb/s.
Four materials were evaluated for the fabrication of an injector plate: 304 stainless
steel, 347 stainless steel, grade 5 Titanium, and C360 Brass. Titanium turned out to be the most ideal material as it
has the highest strength, the highest melting point, and the lowest density. Stainless steel was found to be far too
heavy while brass showed the lowest strength. To determine the required thickness of the plate, an FEA analysis
was performed via CosmosWORKS. It was determined that a 3in titanium plate would
have to be 1/4in thick to withstand the expected 300psi pressure gradient with a factor
Figure 5: Injector
of safety of 2.
Plate
The hole pattern in the injector plate was selected based on trade studies of past
MaCH-SR1 engines as well as other proven hybrid rocket engines such as that used for SpaceshipONE. At this
stage in hybrid rocket technology, the ideal hole configuration is highly debatable while the most commonly
preferred is a showerhead configuration. In order to obtain the desired flow rate, ten holes of 0.05in diameter are
required. The resulting injector plate is shown in figure 5.
E. Nozzle
In a sense, the nozzle is the single most important component to the performance of a rocket. The job of the
nozzle is to convert high thermal energy of chamber gases into kinetic energy and thrust. For this reason, the nozzle
is subjected to the highest velocities, heat fluxes, and pressure gradients found in the rocket engine. To begin this
presentation on the nozzle design, the baseline nozzle will first be discussed followed by the optimized nozzle
design developed by this years’ team.
3. Baseline Nozzle
The nozzle that has been inherited from former MaCH-SR1 teams is machined from ATJ Iso-molded Graphite.
Figure 6 illustrates this nozzle design. The nozzle is located at the end of the rocket engine, immediately downINLET
stream of the post-combustor where the combustion process is completed. The internal geometry
of the nozzle is
sized so as to convert the thermal energy of the combustion process into kinetic energy, which ultimately produces
thrust. This is accomplished by choking the flow of combusted gases to sonic speedsTHROAT
at the throat of the nozzle and
INLET
then by expanding the gases further to super-sonic speeds as they are ejected from the rocket
engine.
Colorado Space Grant Consortium
EXIT
INLET
7
Error! Not a valid link.
The converging section of the nozzle effectively compresses the sub-sonic combusted gases to exactly sonic
speeds. The throat is then considered ‘choked’. The diverging area of the nozzle is then designed so that the flow of
exhausted gases remains super sonic such that further expansion accelerates the super sonic gas flow. This ideal
case is achieved by expanding the gases far enough such that the exhaust
pressure is equal to the ambient pressures of the atmosphere into which
it is ejected. If the exhaust pressure of a choked flow is not equal to the
ambient pressure, the nozzle is said to have over or under-expanded the
flow, in which case the resulting flow will have shock waves, which will
act to dissipate the valuable thrust energy produced by the nozzle.
This nozzle has been developed and verified by former teams, and so
the internal geometry of the nozzle has been left alone by this years’
team. This years’ team has optimized this nozzle however as the
material selection, fabrication process, and exterior geometry has been
Figure 6: Baseline Nozzle Design
improved in the interest of mass reduction.
4. New Nozzle Design
The most straightforward method of reducing mass in the nozzle is to reduce the excess mass surrounding the
nozzle. This excess mass is included only so that the outer diameter of the nozzle matches the internal diameter of
the chamber that houses it. The problem encountered in simply reducing the outer diameter of the nozzle is that the
new engines utilize aluminum to house the nozzle. As mentioned, the nozzle is subjected to extremely high
temperatures and aluminum has a much lower melting temperature than does the stainless steel used with the
baseline engines. Furthermore, the baseline nozzle is made from graphite, which has a relatively high thermal
conductivity. For this reason, new materials were considered for the nozzle that would conduct heat less readily
than graphite to the surrounding aluminum. A trade study was conducted in which graphite was compared to copper
and two types of castable ceramics. The two ceramics consist of zirconium oxide (ZrO2) and silica oxide (SiO2).
The former was considered for its high melting point and the later for its resistance to thermal shock.
Table 1 illustrates the trade study conducted on each of the four materials. The properties considered were
density, melting point, thermal conductivity, availability, manufacturability, and cost.
Table 1: Nozzle Material Trade Study
Density (lbm/in3)
Melting Point (ºF)
Thermal
Conductivity
(BTU-in)/(hr-ft2-ºF)
Availability
Manufacturability
Cost ($/lbm)
ATJ
ISO-molded
Graphite
0.064
800
0.32
1560
ZrO2
Cast
Ceramic
0.14
4000
SiO2
Cast
Ceramic
0.064
2700
804.8
403.1
6.5
4.0
High
Med
$17.87
High
High
$29.60
High
Med
$9.90
High
Med
$7.50
Copper
As is clearly illustrated in the table, silica oxide is more desirable than the other materials investigated in all
categories except melting point. However, all materials in question have melting points below the expected
chamber temperature, which is in excess of 5000ºF. In fact, the lowest melting point is graphite, which melts at
800ºF. The adverse affects of utilizing a material that melts at temperatures lower than the expected temperature it
will be exposed to has been encountered by former MaCH-SR1 teams in their use of graphite nozzles. Due to the
short burn time of 15 seconds the resulting erosion of the nozzles hasn’t been seen to cause adverse affects in the
nozzles ability to produce thrust. It is expected that the silica oxide will form glassy surfaces as its surface melts
during the firing of the engines. To minimize this occurrence, an ablative is applied to the internal surface of the
nozzle, which will absorb excess heat before it reaches the ceramic nozzle. This theory has been tested and verified
to hold true, as will be discussed later in the testing section of this report.
Graphite is preferable to the other materials in the trade study only because of its low density. Silica oxide has
the same low density as does graphite and is less conductive while carrying a smaller price tag. Furthermore, since
silica oxide is castable, there is considerably less waste of material in its fabrication than in machining graphite.
Although they are both indicated as having medium difficulty in manufacturability, the only complexities in
Colorado Space Grant Consortium
8
Error! Not a valid link.
fabricating silica oxide lie in the fabrication of a mold. However, many nozzles are needed in order to put the
engine through various tests and so this one time cost of fabricating a nozzle mold is much more desirable than
repeatedly machining graphite each time a nozzle is needed. In result, silica oxide was found to be a superior
material for the engines nozzle than graphite.
The ceramic nozzle is fabricated by first creating a nozzle insert that matches the internal geometry of the nozzle,
placing it inside the nozzle housing and then pouring the liquid ceramic material directly between the nozzle insert
and the nozzle fitting. The process is shown in figure 7.
Ceramic
Nozzle
Nozzle
Housing
Figure 7: (A) Rubber insert molds; (B) Casted insert molds; (C) Assembled insert
mold; (D) Silica oxide casting process in nozzle housing; (E) Finished nozzle in
nozzle housing; (F) SolidWORKS depiction of nozzle in housing
The nozzle insert molds used to create the nozzle insert are shown in 7.A. A liquid rubber is poured into these
molds and then allowed to cure. The rubber pieces that are produced from these molds are shown in 7.B. As shown
in 7.C, the nozzle insert is then assembled by connecting these two pieces with a dowel pin. The assembled nozzle
insert is then placed inside the nozzle housing where the liquid ceramic mixture is poured between then nozzle insert
and the nozzle housing (7.D). The cured ceramic nozzle is shown in 7.E, with a solid model depiction of the nozzle
assembly illustrated in 7.F.
F. Mass Analysis
The result of these systems level and component level mass reductions are shown below in figure 8:
Colorado Space Grant Consortium
9
Error! Not a valid link.
Mass Comparison
Weight (lbs)
40.00 lbs
35.00 lbs
30.00 lbs
25.00 lbs
20.00 lbs
Baseline
15.00 lbs
10.00 lbs
Aluminum
Composite
5.00 lbs
0.00 lbs
Injector Chamber
Fuel
Nozzle
Total
Component
Figure 8: Mass comparison of each engine.
As can be seen in the figure, this years’ team obtained significant reductions in the mass of the MaCH-SR1
hybrid rocket engine. An overall mass reduction of 49.8% was obtained with the aluminum engine and a 61.1%
reduction was achieved with the composite engine. In the composite engine, only the chamber was fabricated from
composite fibers. While the aluminum engine is a great success, the composite engine shows even greater promise
for continued development. Furthermore, as this years’ team obtains fuel regression data, mass reductions can also
be achieved with the fuel grain that will ultimately lead to further reductions of mass with each of the other
components, as their sizes are ultimately driven by the size of the fuel-grain. In summary there is still more room
for improvement, as will be expanded on in following discussions of each component.
1. Chamber
The greatest total mass reduction was achieved in the redesign of the engine chamber. This is not a
straightforward comparison, however. In the redesign process the pre and post combustors were relocated from the
chamber to the injector and nozzle housings, respectively, while these regions were part of the chamber itself in the
baseline design. Also the flanges that were welded to each end of the baseline chamber were eliminated completely
and were replaced with threaded fittings. Also the weight of insulation was actually increased in the new engine
designs in comparison to the baseline engine design. This is due to the more sensitive thermal properties of
aluminum and composite fiber, which required that phenolic insulation line the entire chamber of each engine while
it was used only in pre and post-combustor areas of the steel baseline chamber.
In result the aluminum chamber, with full phenolic insulation, weighs in at 8.29lbs – a 52.9% mass reduction
from the 17.61lb baseline chamber. The composite chamber was even more impressive as the total weight of two
aluminum thread rings, chamber length insulation, and composite wrap is a total of 3.97lbs. This is a 77.5% mass
reduction from the baseline chamber. As the fabrication process of composite engine components is better
developed during the MaCH-SR1 program, this mass reduction can potentially be achieved with the injector and
nozzle housings as well as with the chamber.
2. Injector
The baseline injector consisted only of a steel feed system adapter and a steel flange. In the new engines, the
injector consists of a titanium plate, aluminum housing, pre-combustor area, cera-fiber insulation, and an aluminum
oxidizer cap. Even with these added parts, an overall mass reduction of 55.6% was achieved over the baseline
design as the mass was reduced from 8.26lb to 3.27lb in each of the new engines.
More reductions can be made with this component by future MaCH-SR1 teams in three significant ways. One
way of doing this would be by utilizing composite fiber as the primary structural material. A second possibility for
mass reduction is to consider the necessity of a pre-combustor area. It is widely debated whether a pre-combustor is
necessary for single-port hybrid rocket engines. However, at this point in the overall MaCH-SR1 design process the
injector plate is still being developed, and the existence of a pre-combustor area is necessary to ensure that the
Colorado Space Grant Consortium
10
Error! Not a valid link.
oxidizer is fully gasified before entering the fuel port. The third way to further reduce mass with the injector
component comes with modifying the oxidizer cap. Currently its outer diameter is the same size as the outer
diameter of the injector housing so that it can be bolted down into the thick walls of the injector housing. The
oxidizer cap does not need to be this large as its only function is to interface with the feed system.
3. Nozzle
Perhaps the most successful way that mass was reduced came with the re-design of the nozzle. In the baseline
engine, the nozzle assembly consisted of a flange and a graphite nozzle. In the new engines, the size of the nozzle
was greatly reduced by replacing the flange with a nozzle housing and the material of the nozzle itself was replaced
with a castable ceramic. In result the nozzle assembly was reduced by 71.7%, as the nozzle assembly in the new
engines weighs 2.06 lbs compared to the 7.29lb baseline nozzle assembly. Considering the use of composite fibers
with the nozzle housing can further reduce the mass of this assembly.
4. Fuel
The fuel was the only component that was not optimized for mass. The reason for this is because of the lack of
empirical data available regarding the actual fuel regression rate achieved with the developed fuel port geometry.
As a result, a maximum fuel regression rate has been utilized to drive the required thickness of the fuel grain. For
this reason, it is much thicker than necessary and much of the fuel remains unburned after each engine firing. This
means that the chamber diameter and mass is greater than it actually needs to be. Fuel regression data acquired by
this years’ team will allow for reliable regression data that will allow for future teams to decrease the size of the fuel
grain, ultimately allowing for a decrease in the chamber diameter and chamber mass.
IV.
Fuel Regression Data
As mentioned previously, the main obstacle faced by this years’ team in optimizing the engine mass came with
uncertainties in the combustion characteristics of the fuel. The fuel was intentionally oversized by previous teams in
order to ensure that there was sufficient fuel left over after the burn to act as insulation between the combustion zone
and the steel chamber walls. This years’ team has solved this problem by designing and verifying the performance
of reliable insulation materials. Furthermore, no previous team has been able to empirically determine the
regression rate of the fuel during the combustion process. The estimated regression rate used to size the fuel comes
from historical data of the regression rate of a circular port. However, the fuel port developed by previous teams so
as to achieve a constant burn rate, and utilized by this years’ team is a cross-shaped port. In acquiring actual
regression data of this cross port geometry, the ability to reduce the size of the fuel grain will be achieved leading to
reductions in the size of the engine chamber, nozzle, nozzle housing, and injector housing.
In order to acquire this data, Heather Chluda was approached. Heather is a graduate student at the University of
Colorado, and former MaCH-SR1 team member, who is researching the regression rate of HTPB fuel in hybrid
rocket engines that utilize nitrous oxide as an oxidizer. This years’ team was able to work with her on designing and
acquiring a sensor suite that could be embedded into the fuel grain in order to learn about how the combustion
process actually evolves during a burn. In result, a collection of eroding thermocouples and regular K-Type
thermocouples were placed into the baseline fuel grains such that the local temperature of the fuel port and fuel
surface at various locations could be attained as well as local measurements of the regression rate of the fuel at
various locations. Together this data would provide an unprecedented understanding of how the combustion process
develops in the engine, allowing future teams to use this data to optimize the size and shape of the fuel grain.
To obtain the fuel port and fuel surface temperatures, eroding thermocouples developed by Nanmac (figure 9)
were acquired.
Figure 9: Nanmac Eroding Thermocouple
These thermocouples work by shielding the two thermocouple leads with a stainless steel cover. The cover is
sized such that the thermocouple leads erode at a specific temperature, and as they do so, the thermocouple weld at
the tip of the sensor is continuously renewed. For our application two types of eroding thermocouples were utilized.
The first being a G-Type thermocouple which was designed to withstand the temperature of the combusting fuel for
Colorado Space Grant Consortium
11
Error! Not a valid link.
the entire duration of the burn, ultimately providing instantaneous temperature measurements at their tip for the
duration of the combustion process. The second type is a C-type thermocouple, which was designed to burn with
the HTPB fuel, providing instantaneous measurements of the burning surface. Due to cost restriction, a total of five
eroding thermocouples were obtained. Two of these were the G-type and the other three were the C-type
thermocouples. They were distributed between each of the two baseline engines.
In one engine, two C-types and one G-type thermocouple was embedded into the fuel. The G-type was placed
such that its tip is located directly in the center of the fuel port, allowing for direct temperature measurements in the
center of the combustion zone. The two C-type thermocouples were then placed adjacent to each other, one in a
peak and the other in a trough of the fuel grain. This allows for an understanding of how the peak of the fuel grain
erodes in comparison to the trough at these locations in the fuel grain.
In the second engine, one C-type and one G-type thermocouple were placed into the fuel grain. This time, the Gtype thermocouple was placed such that its tip is located at the initial fuel surface in the same axial location of the
fuel grain that the G-type thermocouple in the first engine is placed. This allows for data that will show how the
boundary layer develops near the surface of the fuel. In the first engine, the G-type thermocouple remains
completely outside of the boundary layer. In the second engine, the G-type thermocouple is initially at the base of
the boundary layer but will eventually protrude outside of it as the fuel surface regresses. The C-type thermocouple
in the second engine is placed in the same relative location as the two in the first engine. This allows for verification
that the data acquired from both engines is repeatable, and thus reliable.
Along with the eroding thermocouples, regular k-type thermocouples are used to directly measure the local fuel
regression rate at the same location as the eroding thermocouples. These thermocouples are placed in pairs, aligned
radially inside the fuel grain. The welded tips of each thermocouple pair are separated by 0.2 inches. They are
intended to fail at temperatures well below the combustion temperature. Therefore, the local regression rate can
easily be determined by dividing the separation of 0.2inches between each thermocouple by the time delay between
their respective failures.
The locations of each sensor in each engine are shown below in figure 10. In this figure both the axial and radial
location of each sensor is identified.
Figure 10: Locations of eroding and regular thermocouples in each baseline engine.
The engine with two eroding thermocouples is labeled as the CG engine, while the baseline with three eroding
thermocouples is labeled the CGC engine. The two images that correspond to each engine show the axial and radial
locations of each sensor. This number indicator shown in the axial images of the fuel grain provide the depth (given
in inches) from the inlet side of the fuel grain. The radial location of each sensor can then be identified by referring
to the same number identifier along with the type of thermocouple found at that depth.
Colorado Space Grant Consortium
12
Error! Not a valid link.
V.
Testing
A. Component Tests
1. Chamber Pressure Tests
In order to verify the pressure that the chambers where designed to withstand it was necessary to reach a pressure
of 750 psi at room temperature. This value includes a FOS of 1.5 added to what the chamber will be exposed during
real firing (500 psi). The theoretical analysis done in Solid Works to the new chambers (Aluminum and Composite)
gave a critical value of 1000 psi, which must not be surpassed during testing. It is for this reason that the right
equipment was very important to carry out the test. The external equipment used for the test was a manual hydraulic
jack, a hydraulic gage with the correct resolution and hydraulic fluid. The chambers where in full configuration
excluding the following: injector plate, fuel, nozzle, insulation and sensors. The parameter measured during this test
was the maximum pressure reached with out any leakage or fracture of the chambers up to a maximum of 750 psi.
The test where successful for all the chambers, they all withstood 750 psi as it was expected.
2. Oxidizer Cap Pressure Tests
This pressure test concentrated on the oxidizer-cap, injector-plate and fitting assembly, which is to withstand a
maximum pressure of 1600 psi with a FOS of 1.4. The reason why this is so much higher than the pressure seen in
the chamber is because the pressure of the N2O could go up to 1400 psi in order to maintain the desired mass flow
rate provided by the injector plate. This test has the same external equipment as the Chamber Pressure test. The
only difference is that in this assembly the injector plate is placed inside the fitting. In order to prevent the fluid
from going through the holes of the plate an EPDM rubber sheet was place on top. Pressure was the main
parameter in this test, and as expected the assembly was able to withstand 1600 psi with out sign of leakage or
fractures in the structure.
3. Insulation Flame Torch Tests
The mass flow rate provided by the injector plate is a very important factor affecting the thrust performance of
the rocket. For this reason a flow rate test was necessary to determine the mass flow rate provided by the injector
plate. The flow rate provided by the plate is strictly dependent on the pressure at which the fluid is being applied
through the orifices of the plate; therefore the purpose of the test was to measure the flow rate with respect to
pressure. During the actual fire test, the mass flow rate should be close to 0.48 kg/s (1.06 Lb/s) at a pressure of 300
psi. However, being able to conduct a test at such high pressure with N 2O is out of our capability for testing at the
component level. This lead to the following decision, the test would be conducted at lower pressures using water as
the main fluid. The results are then interpolated up to the hot fire pressure of 300 psi. The reason why water was
chosen as the fluid was due to the fact that it has a density close to liquid N 2O and it can be pressurized very easily
and safely with a pump. Since water was used volume flow rate was easier to measure than mass flow rate, the
experiment consisted of simply timing how long it takes for the injector plate to expel water measured in volumetric
units. Therefore the desired volume flow rate is 0.48 l/s for tap water.
4. Nozzle and Insulation Flame Torch Tests
The nozzle and the other materials used as insulation (glass phenolic and ceramic fiber) are there to protect the
chambers wall from being exposed extreme temperatures. The temperature of combustion can reach up to 3000 K
while the chamber is designed to withstand a maximum of 500 K. It is for this reason that a flame torch test was
required in order to ensure the thermal performance of the nozzle and the insulation. The flame torch test consisted
of applying direct flame on the nozzle and the insulation for at least 30 seconds, twice as long as the actual hot fire
test, while measuring the temperature on the cold side of the sample. The test results where not satisfactory at first,
but with the usage of thin layer of ablative material the results where satisfactory.
B. Static Thrust Tests (Josh – 2 Pages)
At the publish date of this paper, the static thrust tests of each of the engines has not been completed. These tests
are to be conducted on April 1 and April 4, 2005. Results of the tests will be presented during the presentation of
this paper on April 8, 2005
1. Test Procedures
The purpose of the static thrust tests are to verify that the mass reductions achieved in the design of the
aluminum and composite engines do not come at the cost of thrust performance. Also, these tests will allow for the
fuel regression data to be obtained from firing the two baseline engines. Each of the engines require the same test
Colorado Space Grant Consortium
13
Error! Not a valid link.
setup and procedures. They will be conducted at Lockheed Martin in Littleton, Colorado, under the supervision of
Lockheed Martin engineers.
The procedure for each test is shown below in figure 11.
Figure 11: Static thrust test procedures.
The ignition of the combustion process is achieved by use of steel wool and an electric match. The steel wool is
located inside the engine chamber, and is wrapped around an electric match. The procedure begins by saturating the
engine with pure oxygen, so that the match can reliable ignite the steel wool. It takes 5 seconds for enough oxygen
to fill the chamber. As soon as the chamber is fully saturated, the electric match is triggered and the steel wool is
ignited. At this point the oxidizer flow is initiated and allowed to flow for 15 seconds. During this time, the
oxidizer flow is throttled by use of a control valve in the oxidizer feed line, so that an average of 300lb f thrust is
achieved. At the end of the 15-second burn, nitrogen is used to extinguish the burn completely before test
conductors are allowed to approach the test apparatus. During the process, Labview software will collect sensor
data including the regression sensor suite, the chamber pressure, the inlet oxidizer pressure, the load cell, a flow
meter, and various temperature and pressures sensors in the feed system so as to determine any leaks in the piping.
2. Feed System
This years’ MaCH SR1 team designed a top level feed system, defining the overall feed system requirements. Ty
Meusburger of Lockheed Martin then fine-tuned the design to account for its interfaces with Lockheed Martin
facilities. The oxygen supply is pressurized to 120psi, which will reliable saturate the engine with oxygen without
resulting in too strong of a flow such that the steel wool is blown out of the chamber before it is ignited. The
nitrogen supply is pressurized to 500psi so that it is strong enough to overcome the combustion pressure of the
engine. The nitrous oxide supply is pressurized to 1250psi in order for the desired flow rate to be achieved. This
presented the biggest design challenge in the feed system as nitrous is packaged at 750psi. An external helium
source was required to increase this pressure to 1250psi, which required modification of the bottle in which nitrous
is shipped. Nitrous bottles typically have only dip-tubes from which to access liquid nitrous. A pressure port had to
be added to the bottles so that the helium source could be interfaced with it.
VI.
Conclusion
This years’ MaCH-SR1 team aspired to bridge the gap between designing a hybrid rocket engine and designing a
full rocket assembly. The main obstacle in the overall MaCH-SR1 project was the low T/W ratio of the proven
rocket engine design. Therefore the goal of increasing this T/W ratio by more than 30% was undertook with
remarkable success. The team ended up doubling this goal by reducing the mass of the proven rocket engine by
more than 60%. Furthermore, the team created a foundation from which even greater mass reductions could be
attained by exploring composite technologies and by further characterizing the combustion process. During this
project many new technologies were explored and demonstrated such as the use of aluminum and composite fiber as
Colorado Space Grant Consortium
14
Error! Not a valid link.
structure materials, the use of the ceramic alumina oxide as a nozzle material, the dependability of phenolic and
cera-fiber insulation as applied to high temperature rocket engines, as well as the use of sensors inside the engine so
as to characterize the combustion process. Because of these findings future teams will have a difficult decision on
which technologies to further develop first in the pursuit of a launch vehicle for the University of Colorado.
Acknowledgments
A. MaCH SR1 Family
1. Current Team Members
Everyone on this years’ team has contributed to the work presented in this paper and so it represents their
collective efforts. In addition to the authors of this paper, there are six others on the team. These individuals
include Nick Anderson, Andrew Campbell, Kurt Etter, Chris Madsen, Nicole Ortmann, and Joel Wecker.
2. Past Team Members
The successes of all former MaCH-SR1 teams have laid the foundation from which this years’ team has built.
Several individuals from these former teams have even offered their advising and support directly throughout the
last year. These individuals include Heather Chluda, Ervin Krause, Otto Krause, and Kevin McWilliams.
3. Team advisors
As this project is part of the aerospace senior project program at the University of Colorado, a special thanks is
due to the project advisors who have repeatedly offered their advice and expertise. These advisors include Dr. Brian
Argrow, Dr. Donna Gerren, Dr. Jean Koster, Dr. Kurt Maute, Matt Rhode, and Dr. Jeffrey Thayer.
B. Industry Support
1. Lockheed Martin
Lockheed Martin was a very active supporter of this project. They allowed the MaCH-SR1 team use of their
facilities for firing each of the engines. They also provided support in the design, integration, and final assembly of
the feed system, control software, and test stand. In each of these systems they also donated hardware that the team
was otherwise unable to afford. The Lockheed Martin engineers who volunteered their time to this project include
Charles Box, Dean Lenz, Ty Meusburger, Nick Patzer, and Mike Webber.
2. Grants
Several organizations contributed funding to this project including the CU Undergraduate Research
Opportunities Fund, Orbital Science Corporation, Lockheed Martin, Dow Corning, and the CU Engineering
Excellence Fund, and the CU aerospace department.
References
1.
2.
3.
4.
5.
6.
7.
8.
MaCH SR1 2004-2005 Fall Report
MaCH SR1 2003-2004 Final Report
MaCH SR1 2002-2003 Final Report
Anderson, Campbell, and Kanner “Investigation of a Castable Ceramic Nozzle in a Mid-Powered Rocket Application”,
AIAA 2005
Humble, Henry, and Larson, Space Propulsion Analysis and Design. Space McGraw Hill Publishing, 1995
Moss, Dennis, Pressure Vessel Design Manual. Gulf Professional Publishing
Ortmann, “Advanced Processing of Hybrid Rocket Solid Propellant” AIAA 2005
Vable, Madhukar, Mechanics of Materials Book. Oxford University Press ©2002 New York
Colorado Space Grant Consortium
15