GT2PropulsionSystemSubmissionDocument

Propulsion System
Annie Lin, Edmond Ngo, Edwin Romero, Darren Charrier, Diana Alsindy, Feennette Navarro,
Kenneth Benedictos
1. Abstract
The Students for the Exploration and Development of Space at the University of
California, San Diego (SEDS@UCSD) Chapter is researching additively manufactured
propulsion thrusters. Presented in this report is the first monopropellant engine, named
Callan, designed by SEDS@UCSD to be printed. This engine includes an additively
manufactured diffuser section, reaction chamber, and nozzle, which are printed in
separate pieces to be bolted together. The catalyst pack is not additively manufactured
and is assembled through traditional manufacturing processes. The design process
and emphasis on safety analysis for each section is explained throughout this
document, including calculations and considerations. With plans for future ground
testing and improved design iterations, the ultimate goal of this engine is to propel the
Triteia cube satellite into lunar orbit.
(a) (b)
Figure 1-1: A computer generated model of the monopropellant engine, Callan:
(a) Exterior view, (b) Interior cross-sectional view.

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Table of Contents
1. Introduction 4
1.1 Nomenclature 3
2. Design Requirements 5
2.1 Volume 5
2.2 Safety 6
2.3 Material 7
2.4 Electrical Power 7
2.5 Thermal 7
2.6 Storage and Handling 7
3. Hardware Designs and Analysis 8
3.1 Engine Design 8
3.1.1 Diffuser Plate Design 8
3.1.2 Chamber Design 15
3.1.3 Nozzle Design 17
3.1.4 Catalyst Bed Design 22
3.1.5 Propellant 29
3.2 CubeSat propellant Feed System Design 31
3.2.1 Delta-V Budget 32
3.2.2 Tanks 33
3.2.3 Valves 40
3.2.3.1 Quick Disconnects 40
3.2.3.2 Solenoid and Solenoid Latch Valves 40
3.2.3.3 Pressure Relief Valve 42
3.2.4 Thermocouples 43
3.2.5 Pressure Transducers 43
4. Static Fire System Design 44
5. Minimum Risk and Hazard Reduction by Design 46
6. Hazard Report Development (SLS Plan 217) 47
6.1 Summary of Derivation of System MDPs 47
6.2 Potential Hazards and Hazard Category 49
6.3 Preliminary Table 50

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6.4 Proposed Verification Approach for Controls to Ensure Pressure Integrity 52
6.5 Proposed Verification Approach Including Controls to Prevent Leakage 52
7. Material Compatibility, Toxicity, Flammability, and Toxic Off Gassing 52
7.1 Engine, C-Ring, Anti-Channel Baffles: Inconel 718 52
7.2 Nickel Screens 53
7.3 Silver Screens 53
7.4 O-Ring 54
7.5 Bolts 55
7.6 Tanks 55
8. Risk Matrix Plan 56
9. References 60
1.1 Nomenclature
Q= Volumetric Flow Rate
ṁ= Mass Flow Rate
= average molecular weightm
⍴= Density
Re= Reynold’s Number
P= Pressure
P0n= Stagnation pressure at nozzle throat
μ= Viscosity
f= Friction Factor
l= Length of Straight Pipe
g= Acceleration Due to Gravity
k= Specific Heat Ratio
= ThicknessxΔ
D= Inner Diameter of Pipe
ᵰ= Coefficient of Loss
A= Area
N= Number of orifices
t= Thickness
= residence timetΔ
σ= Stress
SF= Safety Factor
F= Force
ᵰ= Torque
Isp= Specific Impulse
Dcp= catalyst pack diameter
Lcp= length of catalyst pack

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1. Introduction
Building upon newly advanced research in Hydrogen Peroxide (H2O2)
propellants and engine design, Triteia is designed to utilize a 1 lbf thruster using 90%
H2O2 fuel through a blowdown system enclosed in a dual tank configuration. In
addition to reducing weight and reaction time, the thruster is expected to deliver a
more powerful specific impulse of 156.27 seconds in supplying a delta-v value of 464
m/s relative to its ion thruster counterpart. Made of Inconel 718, for its prolonged
endurance at high temperatures, the thruster is additively manufactured using Direct
Metal Laser Sintering (DMLS) techniques in three sections: the diffuser plate, reaction
chamber, and nozzle for assembly purposes. Additionally, the catalyst package will be
assembled separately by compressing multiple silver and nickel screens together along
with additively manufactured inconel anti-channel baffles. The engine stands at a full
length from inlet to nozzle of 8.79cm (3.46in) with a max width of 4.17cm (1.64in) at the
connection points, flanges, and a width of 0.81cm (0.32in) along the chamber.
Given the inherent advantages in the design of Triteia’s H2O2 monopropellant
thruster gained through its additive manufacturing, its fabrication and testing stages
are at a juncture in revolutionizing lunar exploration from the remainder of GT-2
through GT-4. By the 29th of February 2016, Triteia’s test engine, complete with an
already 3D printed truncated nozzle, is to be analyzed for its thrust capabilities and
firing performance. Subsequently, the engine will be tested with a sea-level nozzle
configuration on a test-stand that is currently being fabricated. After obtaining
data/error analyses on specific load, pressure, and temperature values, calculations on
corresponding critical locations throughout the engine will be obtained and optimized.
Constant and continual testing will be devised on an engineering model of the CubeSat
feed system for maximum efficiency and minimal environmental changes during the
actual mission. While the engine being tested is truncated, the flight-ready model will
have the full length of the nozzle for complete expansion.
The goal of this document is to explain how the propulsion system was
designed to meet all of the following criteria:

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2. Design Requirements
2.1 Volume
The CubeSat was allotted 3300 of volume within the 10cm x 20 cm x 30 cmcm3
structure, which the propellant feed system must be able to be constrained within the
given volume.
[SPS.SPL.003] Payloads shall not exceed a combined dispenser/payload (including
any thermal protection and vibration isolation) mass of 60 lbs (27.22 kg) for either a 6U
or 12U configuration.”
[SPS.SPL.004] Dispenser/Payload Center of Gravity. Payloads shall maintain a
combined dispenser/payload CG within the 6U or 12U enveloped in Table 3-1 and
depicted in Figure 3-4.
The system meets the center of gravity and maximum payload weight design
challenges, because the system has a symmetrical two-tank configuration and valve
assembly instead of one tank or an other arbitrary number of tanks. The dual tank
configuration allows the pressure vessels to hold less pressure than a one tank
configuration, thereby optimizing the thickness and weight of the tanks. This in turn
helps reduce tank length and allows the feed system to meet the volumetric
constraints of the CubeSat . Any tank configuration above two was eliminated because
the number of valves needed and the complexity of operation would outweigh (both
literally and figuratively) the benefits of smaller sized tanks.
[SPS.SPL.005]Dispenser/Payload Cleanliness. Payloads shall comply with the
GSDO-RQMT-1080, Cross-Program Contamination Control Requirements document
for visibly clean standard level.
The system meets the payload cleanliness design challenge because all fittings,
valves, tubes, and tanks that make up the system will be subject to oxidizer cleaning
by AstroPak. The process will include something similar to cleaning aluminum
components with sodium hydroxide, passivating the system with nitric acid, cleaning
the system with deionized water, and reviewing hydrogen peroxide properties for
compatibility36
. This also helps with contamination given the volatility of hydrogen
peroxide. The parts will be shipped for cleaning and will be later received by
SEDS@UCSD, and delivered to NASA under positive pressure with seals covering any
openings to prevent contamination.

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[SPS.SPL.006]Payload Storage: Payloads shall be storable up to 6 months under
conditions listed in Table 3-2.
The system meets the storage design challenge because the materials chosen
for the feed system are highly compatible with hydrogen peroxide. Aluminum 7X11-T6
was chosen for the tanks because it has an AOL of 0.33% and is considered a class
one material with rocket grade hydrogen peroxide . Aluminum 5254-H34 was chosen1
for the tubing and fittings, which also has a class one material compatibility with rocket
grade hydrogen peroxide . The temperatures during integration, rollout, and on the2
launch pad are all ideal for hydrogen peroxide since they are around room
temperature. The humidity will not be a matter of concern up until the system is filled
since the feed system will be sealed off and isolated before being handed off for
storage.
2.2 Safety
[IDRD.3.4.4.3] Pressurized systems with lines and fittings less than 1.5 inches
diameter (outside diameter (OD)) must have a Factor of Safety (FOS) for Pressure of
2.0x MDP for proof and 4.0x MDP for ultimate.
[IDRD.3.4.4.3] Pressurized systems with reservoirs / pressure vessels must have a
FOS of 1.5 x MDP for proof and a 2.0x MDP for ultimate.
[IDRD.3.4.4.3] Pressurized systems for other components and their internal parts
which are exposed to system pressure must have a FOS of 1.5x MDP for proof and
2.5x MDP for ultimate.
[IDRD 3.4.8.4.5.1] For sealed or vented containers:
1. Secondary payload sealed containers shall be designed to withstand the
maximum pressure differential created by SLS ascent. (15.2 psia for items exposed to
directly to vacuum).
2. Vented containers shall size vent flow areas such that structural integrity is
maintained with a minimum FoS of 1.4 for a depress rate of 0.15 psi/sec (9 psi/min).
1
Ventura, Mark. "Long Term Storability of Hydrogen Peroxide - 41st AIAA/ASME/SAE/ASEE Joint
Propulsion Conference & Exhibit (AIAA)."
2
Materials of Construction For Equipment in Use with Hydrogen Peroxide. Vol. 104. Philadelphia: FMC
Corporation, 1966. Print.

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2.3 Material
All material and components in contact with H2O2, must not have a severe chemical
effect in a reaction with H2O2 unless otherwise intended.
- Material used to store the propellant for durations exceeding one hour must
have a good or fair rating in terms of chemical effects.
[IDRD 3.4.8.5] Materials and processes shall be in accordance with NASA-STD-6016.
For materials that create potential hazardous situations as described in the paragraphs
below and for which no prior NASA test data or rating exists, the payload developer
will present other test results for SLS Program review or request assistance from the
MSFC in conducting applicable tests.
2.4 Electrical Power
The electrical power usage of the components of the valves and data acquisition
instruments of the propellant feed system were to minimize the electrical power
consumption to less than 8 watts due to battery and solar power constraints. The
safety / redundancy valves can not draw extra power.
2.5 Thermal
[Secondary Payload User’s Guide] The payload must be able to endure surface
temperatures ranging from 200F <TBR-001> with direct Sun on one side to -143F
<TBR-001> with deep space on the other side.
The secondary payload integrated with the deployer inside the MSA is not expected to
radiate heat or contribute to the thermal loading for the SLS vehicle.
2.6 Storage and Handling
[Secondary Payload User’s Guide] Any propulsion system flown aboard SLS must be
reported to the SLS Program in accordance with Section 3.11 of NPR 8715.3C, NASA
General Safety Program Requirements.
Secondary payload design must be compatible with storage of up to six months under
launch site environments while awaiting integration into the vehicle. Storage
temperatures can range from 65-85F. Other environmental conditions are discussed in
the following sections.
Secondary payload design must also be compatible with operations that place the
payload in horizontal as well as vertical attitudes during ground handling and
integration. Access to the secondary payload will not be allowed following integration
into the MSA.

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3. Hardware Designs and Analysis
3.1. Engine Design
3.1.1. Diffuser Plate Design
The monopropellant engine features the use of two diffuser plates: a front and
an aft diffuser. The plates play an important role in a monopropellant propulsion
system by dispersing the propellant to achieve maximum spread and distribution such
that there is a uniform catalyst pack loading. This maximizes the contact of hydrogen
peroxide with the catalytic silver screens. The technology readiness of the diffuser
plate is 9 due to its common use in monopropellant engines. Successful launch
vehicles that use this concept include the Titan I, Titan II, and several other space
vehicles. The front diffuser contains a propellant inlet that is located on the central axis
of the plate. In order to achieve the uniform spread that is required for maximum
engine performance, the orifices on the front diffuser plate vary in size with no central
orifice. The smallest of the orifices begin closest to the center of the plate and increase
in diameter radially outwards. The orifice diameters used are 0.038cm (0.015in),
0.051cm (0.020in), and 0.064cm (0.025in). The quantity of each orifice size is 32, 18,
and 12, respectively.
(a) (b)
Figure 3-1: Front Diffuser Plate Module: (a) Internal geometry with view of central
propellant inlet, (b) Diffuser face with view of 0.038cm, 0.051cm, and 0.064cm diameter
orifices.

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According to previous research , the propellant should be fed through an3
injection or diffuser plate with 20% to 25% open area. As a result, the exit areas for
each orifice diameter were equated to be equal to one-third of the total exit area
required for 21% open area of the front diffuser plate. Through multiple iterations and
simulations using computational fluid dynamics, it was determined that this
arrangement of orifices is the best in creating a more uniform distribution of propellant
across the catalyst pack.
Figure 3-2: CFD simulation to determine the optimum placement of orifices to supply
maximum distribution of H2O2 over the catalyst pack.
The design calculations for the front diffuser plate are illustrated in the following.
The total area of the front diffuser plate is known to be equal to the cross-sectional
area of the catalyst pack. The diameter of the catalyst pack/chamber module was
found via Rocket Propulsion Analysis (RPA) to be 0.81cm (0.32in). Through a simple
calculation of the area of a circle, it is possible to find the total area of the diffuser
plate:
(E3.1)
3
Davis, Noah S., Jr., and James C. McCormick. "Design of Catalyst Packs For The Decomposition of
Hydrogen Peroxide." N.p., n.d. Web. 17 Jan. 2016.

9

The total area of the front diffuser plate is evaluated and yields 0.516cm2
(0.08in2
). Using this information, it is possible to determine the total open area required
for 21% open area:
(E3.2)
Evaluating Equation 2, the open area (Aopen) becomes 0.108cm2
(0.0168in2
). It
was determined that the front diffuser plate would utilize orifices with three different
diameters. The diameters were determined to be 0.038cm (0.015in), 0.051cm
(0.020in), and 0.064cm (0.025in) as stated previously. Their respective areas can be
found via an equation similar to Eq. 1. Their values are reported below:
It is important that all orifice sizes receive the same mass flow through them. To
ensure that all three different size orifices receive that same mass flow, it was
determined that the total open areas for all three orifice sizes were equal at 33% of the
open area. This can be done by using the following equation:
(E3.3)
Evaluating Eq. 3, yields the equated areas to be 0.516cm2
(0.0056in2
). Finally,
the required number of each orifice size can be found by dividing the equated area by
the area of each individual orifice:
(E3.4)
The required number of each orifice size are:
The aft diffuser plate exhibits a larger open area of 30%. The reason for this
larger open area is because a larger exit area is required to minimize any residual
propellant that may pool up in the catalyst pack. The aft diffuser is located at the
bottom of the chamber and is a permanent attachment to the chamber module.

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Furthermore, the aft diffuser contains 77 orifices, each sized at 0.051cm (0.020in)
diameter. They are arranged in a circular pattern around the central axis. In this case, it
is allowable to have a central orifice as the aft diffuser is not required to redirect
propellant flow away from the central axis.
Figure 3-3: Aft diffuser orifice placement located on the Nozzle Module. There are a
total of 77 orifices, each sized at a diameter of 0.051cm (0.020in) in diameter.
The design calculations for the aft diffuser plate are illustrated in the following:
Much like the front diffuser, the total area of the aft diffuser plate is known to be equal
to the cross-sectional area of the catalyst pack and front diffuser. As a result, the aft
diffuser plate has a total area (Aap) of 0.518cm2
(0.0804in2
). The total open area can be
easily found using:
(E3.5)
Therefore, the open area required for the aft diffuser plate is 0.155cm2
(0.024in2
). Since
the aft diffuser plate orifices are uniform in diameter, the number of orifices required to
meet the required open area can be calculated after determining the area of each
orifice. The orifice diameter for the aft diffuser was arbitrarily chosen to be
0.051c(0.020in). Using an equation similar to Equation 1, it is possible to determine that
the area of one orifice (Aorif) on the aft diffuser is 0.002cm2
(0.000314in2
). Then, the
number of orifices can be determined:

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(E3.6)
The number of orifices (N) required to achieve a 30% open area is 77.
The diffuser plate must withstand thermal loading from the decomposition of the
hydrogen peroxide as well as be stiff enough to minimize plate deflection. In order to
determine the minimum plate thickness for the front and aft diffuser plates, the
problem was simplified to a uniformly loaded circular plate with clamped edges. See
Figure 3 for sketch of simplified problem.
Figure 3-4.Uniformly loaded circular plate with clamped edges
Using the equation below , the minimum thickness of the front and aft diffuser4
plates could be evaluated using
(E3.7)
The designed chamber pressure (p) is 861.845kPa (125 psi). In addition, the
diffuser plates will compress the catalyst pack to a maximum 17.24MPa (2500 psia).
The material properties of Inconel 718 show that it has a yield stress between 634.32 ±
48 MPa (92 ± 7 ksi). Using the minimum yield stress of 586 MPa (85 ksi), a factor of
safety (FS) of 3, and a plate radius of 0.406cm (0.16in). The use of Equation 7 yields a
resulting minimum plate thickness of 0.107cm (0.0422in). In addition to numerical
calculations, a simulation was performed on the diffuser plate under the conditions that
were expected in the chamber during a typical burn. The diffuser plate was constrained
at the sidewalls such that there would be no expansion at that point. This constraint is
unrealistic, but was done to simplify the simulation. The actual results would not be as
extreme as the simulation results due to the constraint that was imposed on the
4
"Loaded Flat Plates." Loaded Flat Plates. Web. 31 Jan. 2016.
<http://www.roymech.co.uk/Useful_Tables/Mechanics/Plates.html>.

12

sidewalls. The plate would be allowed to expand during a typical burn. Ramping the
temperature and pressure from the decomposition of hydrogen peroxide yielded
results claiming a maximum deflection of 0.003cm (0.0011in). Images of the simulation
can be found below:
Figure 3-5: Simulation model of the diffuser plate
The thicknesses for the conical sections including the diffuser cone, and the
converging and diverging nozzle, were calculated using the following equations :5
(E3.8)
(E3.9)
The diameters used for the diffuser plate were 0.33cm (0.13in) for the small
diameter, Ds, and 0.81cm (0.32in) for the large diameter, D, which is also the diameter
of the chamber. The length of the diffuser conical section, L, is 0.37cm (0.147in). Since
the engine will be additively manufactured, it was assumed that the seam efficiency, E,
is 1.0. Based on the contents of Inconel 718 material being 50% nickel and 17%
chromium, both materials being reactive with H2O2, the corrosion factor, c, was
estimated to be 0.075 based on hydrogen peroxides decomposition with similar
5
"Shell Thickness Calculation." RedBag. N.p., n.d. Web. 17 Jan. 2016.

13

metals. With these parameters, the minimum thickness of the diffuser at a 45 deg
angle was calculated to be 0.06cm (0.024in).
Figure 3-6: Section view of the diffuser module illustrating dimensions used to
calculate minimum wall thickness.
The monopropellant engine is printed out of Inconel 718. Inconel 718 was
chosen for the engine material because of its availability for additive manufacturing and
its proven compatibility with rocket grade hydrogen peroxide. This relatively new
refractory material is the result of a need for space propulsion systems with high
performance requirements, lightweight materials capable of withstanding high
pressures at elevated temperatures, and a low ductile-to-brittle transition temperature
for high frequency vibrations at cryogenic temperatures (Special Metals).
Some design challenges that were addressed included the need for a good seal
between the interface of the different engine modules, and the available printable
geometries as a result of using additive manufacturing as a method of manufacture.
The engine is designed to be modular with three total modules: the front diffuser
module, chamber module, and nozzle module. This modular design requires the
interfaces between modules to be completely sealed. In order to achieve this, Parker
O-rings and C-rings are used between modules. The grooves required to seat the
O-rings and C-rings will be machined out after printing is completed so that they will
be as smooth as possible to maintain a perfect seal.
Furthermore, the monopropellant propulsion system is additively manufactured.
This presented potential challenges as certain geometries cannot be printed without
support material. Support material are temporary structures that are included in the
printing process to build geometries that would otherwise be impossible to build. A
common reason for the use of support material is when a design calls for an

14

unsupported ceiling structure. This is impossible to build without support material
because a printer cannot build over empty space. For our purposes in minimizing cost
and part production time, it is beneficial for a diffuser plate design that requires as little
support material as possible. The front diffuser plate is designed to be support
material-free. That is, the front diffuser module can be printed with no support material
as the structure is self-supporting. This is possible due to the restriction that required
the maximum angle with respect to the horizontal to be no more than 45 degrees.
3.1.2. Chamber Design
The role of the chamber is to house the catalyst pack where the reaction of
hydrogen peroxide and silver takes place.
Figure 3-7: Computer generated model of a cross-section of the decomposition
chamber with the pressure transducer port NPT fitting shown.
The chamber’s minimum thickness, Tp, was calculated using Barlow’s Formula
of pressure vessel design, taking into consideration corrosion allowance :6
(E3.10)
The inner diameter, Di, is 0.81cm (0.32in). With the same parameters as the
diffuser plate. The thickness of the chamber was calculated to be 0.022cm (0.009in).
To ensure uniformity and prevent chamber failure, the thickness was finalized to be
0.1016cm (0.040in).
6
"Shell Thickness Calculation." RedBag. N.p., n.d. Web. 17 Jan. 2016.

15

The length of the chamber took into consideration the length of the catalyst bed
with the catalyst package that was calculated using equation (E3.21) in the catalyst
bed design section below.
A pressure transducer port was added to the bottom of the chamber to measure
the pressure. It was assumed that the reaction of hydrogen peroxide will be complete
towards the end of the chamber with only oxygen and water vapor emissions. With this
implication, the gas in between the pressure transducer and decomposition chamber,
as seen in figure below, will give an accurate chamber pressure reading in order to
measure the varying thrust of the engine due to the blowdown system.
Figure 3-8: Cross-section of Chamber with view of the pressure transducer port
leading to chamber access towards the aft diffuser plate and nozzle module. The
pressure transducer port NPT fitting is not shown.
Flanges were placed at the beginning and end of the chamber, as well as on the
diffuser and nozzle, due to the assembly process of the engine. The engine is additively
manufactured in 3 separate pieces, the diffuser section, chamber, and nozzle. The
flanges have 5 holes that are 0.305cm (0.12in) in diameter for 0.284cm (0.112in) bolts
with a tolerance of 0.02cm (0.008in). Using stress equation ( ), the force was σ w = A
F
calculated to be 134.15N (30.159lbf) with a working pressure, , 2.58MPa (375 psi). σ
w
For 5 bolts, the force per bolt, F, was calculated to be 26.83N (6.0318lbf). The
minimum torque of each bolt was calculated using the below equation:
(E3.11)

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The friction coefficient, , is assumed to be 0.2 in the most general case. Theμ
diameter of the bolt diameter, , is 0.284cm (0.112in). The minimum torque wasDb
calculated to be 15.25 Nmm (0.135lbf-in) for each bolt to ensure a proper seal between
the sections of the engine. The maximum torque was calculated using the max force,
considering the minimum tensile strength of the bolt, , to be 1.034GPa (150 kpsi). σ min
Using the stress equation, the max force, , is 6.573 kN (1477.81lbf) and using F max
equation (E3.11), the max torque to tighten each bolt is calculated to be 37.28 Nmm
(33 lbf-in).
3.1.3. Nozzle Design
(a) (b)
Figure 3-9: (a) an illustration of a 3D model nozzle
The nozzle design was determined using the computing software, Rocket
Propulsion Analysis2
. A chamber pressure of 861.845kPa (125 psia), a thrust of 4.45N
(1 lbf), and an exit pressure of 3.04kPa (0.03 atm) were set as initial parameters. These
parameters were determined based off of an iterative approach to optimize the
efficiency of the engine, both in terms of volume and specific impulse. The combination
of initial parameters produced the highest specific impulse (156.27 sec) while still fitting
comfortably amongst other components within the cube satellite. Assumptions made
during the design were:
● The thrust was set as 4.45N (1 lbf) since a 4.45N (1 lbf) misfire was thought to
have less attitude control risk than a 22.24N (5 lbf) or 44.48N (10 lbf) misfire.
● A contraction area ratio above 3 was selected so that the inlet velocity could be
comparatively small to the exhaust velocity and thus be neglected as it is
assumed, theoretically, for larger thrust chambers. This ratio was also designed
in order to minimize energy losses due to the increase in pressure drops that

17

smaller thrust chambers produce because of smaller contraction area ratios, as
noted by Sutton .7
● The expansion area ratio was designed to perform optimally at 25,908m (85,000
ft) where the ambient pressure is 3.04kPa (0.03 atm) instead of at sea level
because it was found to increase the vacuum specific impulse from 149.77
seconds to 156.27 seconds at the set pressure and thrust. This was crucial
considering the relatively low specific impulse of hydrogen peroxide. An altitude
of 25,908m (85,000 ft) was found acceptable given the relatively small change
in pressure at 25,908m (85,000 ft) and pressure in space.
● The bell nozzle was designed in RPA, assuming the length to be 80% the length
of a reference conical nozzle with a 15 degree half-angle and the same
expansion area ratio.
● Multiphase flow and species ionization effects were assumed for nozzle flow
effects.
An overall engine efficiency of 94% was found using RPA. A divergence efficiency of
99.15% was produced in addition to a drag efficiency of 96.31% and a thrust
coefficient of 1.639 in a vacuum.
Challenges in the design include the relative length of the parabolic nozzle to
that of a cone-shaped nozzle. Rocket Propulsion Analysis estimates the length of a
parabolic nozzle to be 80% the length of a reference conical nozzle. An increase in
chamber pressure in order to achieve a higher specific impulse, however, forced a
parabolic nozzle length of 103% with respect to the reference conical nozzle. The
increased length resulted in a suboptimal engine weight for the designated altitude but
because the extra nozzle length affected only engine weight and not exit pressure, a
cost benefit analysis between the two determined it was an acceptable design.
Another challenge was deciding the optimal thrust for the engine. Engine size
increased with increasing thrust which limited the amount of space remaining for other
components. Below are images comparing the 44.48N (10 lbf) thruster that was
originally considered for the CubeSat to the 4.45N (1 lbf) thruster that was chosen for
the final design:
7
Sutton, George Paul., and Oscar Biblarz. Rocket Propulsion Elements. New York: John Wiley & Sons,
2001. Web.

18

(a) 1 lbf thrust (b) 10 lbf thrust
Figure 3-10. 1 lbf vs 10 lbf engine size comparison inside the CubeSat
For the purposes of testing the engine in the Mojave Desert (1500m (4,921ft)
above sea level), the following analysis explains how the nozzle was truncated in order
to avoid flow separation:
Assuming isentropic flow, the following relationship between area and pressure
for supersonic nozzles was used :8
(E3.12)
The desired pressure was constrained as:
(E3.13)
8
Sutton, George Paul., and Oscar Biblarz. Rocket Propulsion Elements. New York: John Wiley & Sons,
2001. Web.

19

This pressure was acceptable given the small change in pressure between sea level
pressure and pressure at 1500m (4,921ft) above sea level.
Parameters that were pre-set as constraints:
(E3.14)
Parameters determined using RPA:
(E3.15)
Tables with relevant data from RPA:
Table 3-1: Values of specific heat ratio, k, at various locations within the engine.
Parameter Injector Nozzle Inlet Nozzle Throat Nozzle Exit
k 1.2656 1.2656 1.2771 1.3476
Table 3-2: Various effective exhaust velocity values at different pressures
Parameter Sea Level Sea Level Flow
Separation
Optimum
Expansion
Vacuum Units
Effective Exhaust
Velocity
-591.69 533.78 1468.72 1532.44 m/s
Rearranging to solve for , the diameter was found for both the specific heat ratio atAx
the nozzle throat and at the nozzle exit afterwards:
(E3.16)

20

Table 3-3:Values calculated to determine distance of truncation.
Parameter Lowest Value Highest Value
k 1.2771 1.3476
px 101325 Pa 101325 Pa
p1 861845 Pa 861845 Pa
At 3.173 x 10-6
m2
3.173 x 10-6
m2
Ax 6.0776 x 10-6
m2
5.7885 x 10-6
m2
Dx 0.002781 m 0.002714 m
x 1.0025 mm 1.0025 mm
(a) (b)
Figure 3-11. An illustration of different nozzle lengths: (a) Full length nozzle, fully
expanded designed for use in a vacuum, (b) Truncated nozzle, designed for engine
testing at sea level.
Like the other modules that make up the engine, the converging diverging
section of the engine is also printed out of Inconel 718. This material was a factor in
nozzle thickness. The thickness of the nozzle was calculated using the same equations
for the diffuser plate design, namely equations (8) and (9). The thicknesses of the
converging and diverging sections of the nozzle were calculated to be 0.0965cm
(0.038in) and 0.03cm (0.012in) respectively. To prevent nozzle failure and ensure
integrity, the corrosion of the metal was accounted for, so the thicknesses was
finalized to be 0.102cm (0.040in).

21

3.1.4. Catalyst Bed Design
The design calculations for the catalyst bed length are illustrated in the following.
Several catalyst bed design parameters for 1lb thruster include the mass flow rate
determined at the engine throat of 0.003 kg/s (0.007 lb/s), chamber pressure of
861.845 kPa (125 psia), catalyst bed diameter of 0.813cm (0.32 inches):
(E3.17)
Under frozen flow assumptions, there is no chemical exchange during expansion at the
point of time, no rate processes, and molecules preserves their identity. Hence a
constant specific heat ratio, γ. The assumptions also include isentropic flow so friction
and heat losses are ignored. The average molecular weight of 90% hydrogen peroxide
decomposition product is 22.1g/mol with 0.7076 mol fraction of H2O and 0.2924 mol
fraction of O2
Common monopropellant characteristic velocity c* is 0.90, hence the chamber
temperature due to combustion chamber performance is :9
T0= Tad(ηC*)2
= 1029.5K x 0.92
= 834K
The stagnation pressure at nozzle throat was found using:
(E3.18)
The velocity at the nozzle throat :
(E3.19)
9
Othman, Norazila, Subramaniam Krishnan, Wan Khairuddin Bin Wan Ali, and Mohammad Nazri
Mohd Jaafar. Design and Testing of a 50N Hydrogen Peroxide Monopropellant Rocket Thruster.
Jurnal Mekanikal,December 2011,Vols. No 33, 7081.Web.

22

From RPA, the velocity of gas at the nozzle exit ,V2, is perceived to be 1,559.84 m/s
and the pressure P2 is 3.048kPa (0.4408psi). Average specific heat ratio, k, is assumed
to be constant with a value of 1.2967. Gas constant R is 367.1 J/kg*K. The
decomposition temperature at the catalyst chamber, T1, is 1,019.23 K and the pressure
P1 is 861.845 kPa (125 psia). With the numerical values, the velocity at the nozzle throat
V1 is found to be 246.698 m/s. It is important to note that the density of superheated
steam and oxygen decomposition product have an ideal standard gas density of
roughly 2.21 kg/m3
at temperature 1,039 K and pressure 861.845 kPa (125 psia)
(E3.20)
With the velocity found using equation 3 and the average density of superheated steam
and oxygen found using equation 4, stagnation pressure becomes:
= 861845pa +0.5(2.21kg/m3
)(246.7m/s)2
= 929096.28 PaP0n (E3.21)
Evaluating equation 1 with determined parameters:
Lcp= = 0.02948m= 1.16075in8314.3 0.00302kg/s 834K 0.0015 s* * *
(22.1 g/mol)(π (0.008128 m) /4)(929096.28 pa )*
2
The catalyst bed length above consists of only silver screens, each of which will
have a 0.036cm (0.014in) thickness. For the silver length of 2.95cm (1.16in), there are
going to be 83 silver screens and hence 83 nickel screens that are 0.036cm (0.014in)
thick to support the silver screens. However, due to the CubeSat physical size
constraint, the amount of the silver screens will need to be reduced. According to the
NASA Technical Note 1808 includes multiple silver screen catalyst bed designs for10
90% hydrogen peroxide, which are shown to use 60 pre-treated silver screens and
roughly 20 support screens. This previous silver screen catalyst design led to the
decision to cut the number of silver screens from 83 down to 70 screens and the
amount of nickel screens from 83 to 42.
According to McCormick , in small diameter silver catalyst bed loading factor value of11
a 5 to 10 is often employed for low temperature –starts.
10
Runckel, Jack F. Willis, Conrad M. Salters, Leland B. Jr. Investigation of Catalyst Beds for
98PercentConcentration Hydrogen Peroxide. Washington: National Aeronautics and Space Administration,
1963, Print.
11
Hydrogen Peroxide."

23

With the catalyst pack cross sectional area of 0.080 in2
and 0.00302kg/s (0.0067 lb/s)
mass flow, the loading factor of the catalyst pack is 5.
Silver’s coefficient of expansion is greater than Inconel 718, as a result, the
screens will be crushed on the side (where it comes in contact with Inconel) upon high
temperature decomposition and a decrease in screen diameter upon cooling. A gap
will form between the screens and wall, allowing H2O2 to bypass the catalyst bed and
create a wet start. Anti-channel baffles are necessary to be included in the catalyst
pack to seal any potential gap between the chamber wall and the pack. The drawback
of anti-channel baffles is that they also decrease contact between HTP and the catalyst
pack therefore if sized incorrectly, the efficiency of the catalyst pack will decrease.
The 2 anti-channel baffles in the catalyst pack are 0.036cm (0.014in) thick 3D
printed Inconel 718 rings with an outer radius 0.406cm (0.16in) and an inner radius of
0.381cm (0.15in). With this design, the area of the anti-channel baffle will be 12.1% of
the screen area. The anti-channel baffle will be placed at the beginning and the middle
of the catalyst pack. Placing it further down of the pack will reduce its useful life due to
high decomposition gas temperature. A single silver screen will sit in in the middle of
anti-channel baffle ring to prevent voids in the compressed catalyst pack.
Figure 3-12: Screen configuration in catalyst bed
Starting from the left to the right, the first layer is in contact with the front
diffuser plate and consists of 10 silver screens. The reason this layer is in between the
front diffusor plate and the first anti-channel baffle is to maximize the exposure of silver
to aid decomposition. The number 2 in the figure, labels the first anti-channel baffle.
After the first anti-channel baffle at layer number 3 is 20 silver screens sandwiched
between 2 sets of a single piece nickel screen. This bundle of silver screens is also

24

where the majority of the decomposition process will take place. Next layer at number
4, which is also prior to the second anti-channel baffle is a set of alternating 1 silver
screen and 1 nickel screen up to total of 10 screens of silver and 10 screens of nickel.
Number 5 labels the second anti-channel baffle located at the middle of the catalyst
pack. The silver-nickel alternating configuration of layer 4 is also used at layer 6 after
the second anti-channel baffle up to 20 silver screens and 20 nickel screens. While
towards the end of the catalyst pack will be a higher temperature environment, the
purpose of nickel screens as silver screens support becomes increasingly important.
Hence, the silver-nickel alternating configuration ensures the support the catalyst pack
requires. It is important to orient each screen 45 degrees from the previous to
maximize the surface area contacting the propellant. Any empty spaces in the chamber
after compressing a total of 112 silver and nickel screens with the arrangement stated
above during the catalyst bed assembly will be filled in with additional nickel screens to
prevent void and to provide extra support. The additional nickel screens will be located
after layer 6. The 70 silver screens of 0.036cm (0.014in) thickness, 42 nickel screens of
0.036cm (0.014in) and 2 anti-channel baffles of 0.036cm (0.014in) thickness totals the
length of the catalyst bed to be 4.054 cm (1.596in).
To make sure the diffuser plates can handle the stress load caused by
compressed catalyst bed expansion due to high thermal temperature, it is important to
calculate thermal stress for silver screens and nickel screens that makes up the
catalyst bed. In the following equation, E represents Young’s modulus of the material in
N/m2
, α, is the coefficient of expansion of the material in o
F-1
and dt is temperature
change in o
F which is estimated as the difference of the decomposition temperature
1013.13K (1364 o
F) and room temperature 294.2K (70 o
F).
(E3.21)
Table 3-4: Table of material properties of materials used in the catalyst package
Young’s modulus12
Coefficient of Thermal Expansion 8
Silver 72.4 x 109
N/m2
1.29 x10-6
/
K (11 x 10-6
/o
F)
Nickel 200 (99.6% wrought nickel) 170 x 109
N/m2
9.96 x10-7
/
K (8.5 x 10-6
/o
F)
12
"Modulus of Elasticity or Young's Modulus and Tensile Modulus for Some Common Materials." The
Engineering ToolBox. N.p., n.d. Web. 2 Feb. 2016.

25

Stainless Steel 316 190 x 109
N/m2
1.87 x10-6
/
K (16 x 10-6
/o
F)
Evaluating equation 5 and multiplying it by the screen area to 1 m2
ratio, it is
found that the thermal stress for each 0.32 inches diameter silver screen is 53,443.45
N/m2
and nickel screen is 97,018.31 N/m2
. The total catalyst bed thermal stress with 70
silver screens and 42 nickel screens will be 7.815 x 106
N/m2
. The additional 2 smaller
0.30 inches diameter silver screens that fit in the 2 anti-channel baffles have a total
thermal stress of 93,943.57 N/m2
, which totals the catalyst bed thermal stress to be
7.909 x 106
N/m2
. Through ANSYS simulation, it is concluded that the Inconel 718
diffuser plate is thick and strong enough to encounter 7.909 x 106
N/m2
of thermal
stress.
Shear of the catalyst pack, on the other hand in comparison to the thermal
stress, is not a big concern potentially contributing to catalyst pack inefficiency. Shear
stresses, which may occur at the interface between catalytic screens may cause
catalyst contact surface area to be inadequate in the decomposition process of
hydrogen peroxide resulting in unreacted propellant. However, such an assumption
can be considered to be negligible because of the extensive length of the pack. It can
be assumed that any remaining hydrogen peroxide that has not reacted will continue
and finish its decomposition by the end of the catalyst pack. In addition, with
increasing shear at higher temperature, the latter part of the catalyst pack near the aft
diffuser plate will experience more shear than the initial portion. This is not problematic
because by the end of the catalyst pack most of the hydrogen peroxide propellant will
already be decomposed into superheated steam and oxygen.
The largest challenge in creating a design for a silver screen catalyst for a
hydrogen peroxide monopropellant thruster was the dearth of information available for
such a design. As rocketry technology grew more sophisticated after the 1960’s
hydrogen peroxide was eventually replaced by hydrazine and other higher specific
impulse propellants, leading to a halt in hydrogen peroxide propulsion systems
research and development. It wasn’t until the 1990’s that interest in hydrogen peroxide
renewed due to growing environmental concerns over the impact of propellants such
as hydrazine and research began toward its use in micro-propulsion units for its ease
of handling, good performance, low cost, and environmental friendliness. As its
reintroduction to the space industry is relatively new and cube satellite production
began in 1999 without a launch until 2003 the majority of the information compiled in
this document was largely empirical and lacked guiding design parameters easily
applicable to the design above.

26

The second obstacle was the sizing constraint placed on the engine by the size
of the 6U cube satellite and its components. For this reason, it was a challenge to
create an extremely high performing pack with accurate output values as opposed to
some of the examples in papers where designs and decisions were not optimized.
While determining the length of the decomposition chamber, it was found that there
was a discrepancy between using the diameter of the screen wire as the thickness of
the catalyst screen and using the actual thickness of the screen. The diameters of
silver wire and nickel wire screens are 0.014 inches but the actual thickness of the
screens is 0.028 inches. The thickness of the anti-channel baffle is to also increase
from the previous 0.014 inches to 0.028 inches so that exactly one screen can be
placed in the center. The increase in screen thickness will result in the predicted
catalyst chamber length to double and may cause problems with the placement of the
engine and other CubeSat components. The catalyst pack’s total displacement after
applying 2500 psi of compressive load normal to the pack screen is unclear. It is for
this reason that the team decided to continue with 70 silver screens in the catalyst
pack design. The number of silver screens and nickel screens and the length of the
decomposition chamber will be adjusted based on actual displacement. The remaining
catalyst bed design information will continue to be relevant for the catalyst bed design
with 70 silver screens.
In addition, the catalyst pack requires utmost cleanliness and care. When
inserting the catalyst pack into the chamber, careful handling precautions must
followed so the screens do not develop nicks, wrinkles, or deformations. Such damage
would create voids or cavities in the catalyst bed, which in turn creates performance
and life degradation. During assembly the pack must not be contaminated with oils,
dust, dirt or other foreign object debris that will coat the catalyst’s surface causing it to
become less porous and inactive. Other oxides or sulfur are also harmful to the pack.
Teflon or polyethylene from gaskets or packaging in the feed line or engine will
decompose with 90% hydrogen peroxide decomposition gases and cause the screens
to melt and fuse . Lastly, any inadequate or insufficient purging of the engine post-test13
with water pumped nitrogen gas will leave residual H2O2 in the pack and continue the
oxidation reaction reducing the pack’s life and efficiency. Normal use and on-off pulse
operation, which imposes shock loads, will also degrade the pack over time.
As a common catalyst material for hydrogen peroxide decomposition, silver
screens work excellently as catalyst pack in a hydrogen peroxide powered engine. It
satisfies many requirements of a sufficient catalyst bed in addition to having the ability
13
Hydrogen Peroxide."

27

to decompose hydrogen peroxide to create desired thrust. The silver screen catalyst
pack provides a starting transient below 100 milliseconds and a rapid decay rate after
the hydrogen peroxide fuel is shut off. The low starting transient means we receive
rapid responses to the controls because the engine meets the desired thrust shortly
after the propellant is injected. Rapid decay rate, or the rapid decomposition rate of
hydrogen peroxide in the silver catalyst pack, prevents the propellant from continuing
to exert thrust after being deactivated. The drawback of using silver as a catalyst is
that it has a relatively low melting point of approximately 1088.7K to 1144.2K (1500 o
F
to 1600 o
F) in comparison to other hydrogen peroxide catalyst materials. As a result, it
is undesirable to use with a hydrogen peroxide concentration greater than 90% due to
concentrations higher than 90% resulting in higher decomposition temperatures that
are close to silver’s melting point. However, the melting point of the silver catalyst pack
is sufficiently high enough to sustain a 90% hydrogen peroxide decomposition
temperature of approximately 1033.15K (1400 o
F). 11
In “Method of Making a Catalyst Coated with Samarium Oxide”, McCormick
presented a method to treat pure silver catalyst pack screens with samarium nitrate
Sm(NO3)3. Treating the silver screens with samarium nitrate allows a discontinuous film
of samarium oxide to form on the screen surface . According to McCormick, the14
samarium oxide film prevents large gas bubbles generated during the decomposition
process to build up into pockets surrounding the catalyst pack and blocking the
hydrogen peroxide from accessing catalyst screens. However, due to lack of technical
data proving the result of samarium nitrate treatment having a significant impact on the
catalyst pack performance, a testing of untreated silver catalyst pack will be first
conducted to verify whether the treatment is necessary.
Nickel screens in the catalyst bed act as supporting material for the silver
screens. It has a higher material strength, and it is more resilient to unique load and
stress caused by high temperature than silver. Furthermore, nickel screens were
chosen to be the supporting material as opposed to the commonly used stainless steel
support because nickel also catalyzes hydrogen peroxide. Using nickel in the catalyst
pack aids the decomposition process, whereas stainless steel would exist as inert
weight. The Technology Readiness Index value of silver and nickel woven screen as
14
McCormick, James C. Method of Making a Catalyst Coated with Samarium Oxide. Fmc Corp,
assignee. Patent US3560407 A. 2 Feb. 1971. Print.

28

hydrogen peroxide catalyst pack is 6. It has been extensively tested in an engine by
NASA .15
Similar to the engine’s body, Inconel 718 was chosen as the anti-channel baffle
material because of its proven material compatibility with rocket grade hydrogen
peroxide and its availability at the company this engine is to be printed from. This
relatively new refractory material is the result of space propulsion systems with the
requirement of high performance, lightweight materials capable of withstanding high
pressures at elevated temperatures in addition to a low ductile-to-brittle transition
temperature for high frequency vibrations at cryogenic temperatures .16
3.1.5. Propellant
The propulsion system was designed for the use of hydrogen peroxide of 90%
concentration. At this concentration, H2O2 has a density of 1400 kg/cm3
and may exist
as a liquid at a temperature of 273K (32.1°F) . Due to decomposition from heat17
exposure, the temperature of vaporization of H2O2 has been estimated around 394 K
(250°F) before rapid acceleration of decomposition occurs. The technology readiness
of 90% hydrogen peroxide is 9 due to its common use as a propellant for
monopropellant engines. This propellant was used in space vehicles like the Mercury
space capsule. There are several factors that were considered when this concentration
was chosen:
- Monopropellant propulsion systems significantly reduce the added weight of
bipropellant systems and helped meet the volumetric size constraints
pre-determined by the 6U CubeSat requirements.
- Information regarding safe and proper handling of hydrogen peroxide is more
readily available than that of other monopropellants such as variations of
hydrazine (MMH and UDMH). According to a published AIAA document , high18
concentrations of hydrazine may be regarded as a carcinogen and as a
flammable substance when exposed to very high temperatures. Additionally, it is
not an asphyxiant nor a mutagen such as liquid CO2 or hydrazine of 100%
concentration. Moreover, since high concentrations may be properly handled by
15
Runckel, Jack F. Willis, Conrad M. Salters, Leland B. Jr. Investigation of Catalyst Beds for
98PercentConcentration Hydrogen Peroxide. Washington :National Aeronautics and Space Administration,
1963, Print.
16
Physical Constants and Thermal Properties (n.d.): n. pag. Specialmetals.com. Special Metals
Corporation. Web.
17
Corporation, 1966. Print..
18
Melof, Brian M., and Mark C. Grubelich. "Investigation of Hypergolic Fuels with Hydrogen Peroxide 37th
Joint Propulsion Conference and Exhibit (AIAA)." Aiaa.org. American Institute of Aeronautics and
Astronautics, n.d. Web. 31 Jan. 2016.

29

diluting in large amounts of water, which is easily accessible, hydrogen peroxide
was chosen as the propellant.
Figure 3-13 : The density of hydrogen peroxide at temperature ranges from 0 - 100
deg-Celsius.
The main benefit of using Hydrogen Peroxide as a monopropellant are the
significant cost-savings associated with the drastic simplification of the health and
safety protection procedures necessary during propellant production, handling, and
storage. High-energy propellants such as green propellants depend on complex
organic particles and repay the vast sub-atomic weight of their decay items with high
operational temperatures of the exhaust gasses. Hydrogen peroxide does not
experience those disadvantages of these detriments and has consequently been
rethought as a promising green charge for low and medium thrust applications.
Hydrogen peroxide is a high density liquid having the characteristic of being able to
decompose exothermically into water (steam) and oxygen according to the reaction:
2H2O2 (l)→ 2H2O(g)+ O2 (g)
The physical properties of hydrogen peroxide are close to those of water, with
two imperative contrasts: H2O2 has a higher thickness and a much lower vapor
weight. It stays in the liquid state at enveloping weight in a broad assortment of
temperatures and is modestly easy to handle concerning other fundamental liquid

30

rocket fuel oxidizers like dinitrogen tetroxide, nitric destructive and liquid oxygen
(Ventura and Muellens3
). The propulsive execution of hydrogen peroxide
monopropellant rockets is around 20% lower than hydrazine, yet the volume particular
drive achievable with 90% H2O2 is higher than most diverse powers as a result of its
high thickness.
The most critical challenge for the acknowledgment of hydrogen peroxide
monopropellant thrusters is the improvement of viable, dependable, extensive reactant
beds, giving quick and repeatable execution, inconsiderateness to harming by the
stabilizers and pollutants contained in the fuel, fit for managing the expansive number
of warm cycles forced by run of the mill mission profiles and not requiring (if
conceivable) pre-warming for proficient operation.
3.2 CubeSat Propellant Feed System Design
Figure 3-14. The Final Propellant Feed System design in CubeSat
The feed system was designed to monitor the status of the propulsion unit by its
temperature and pressure readings throughout various sections in the system. The
system is a multiple tank system due to volume constraints and parameters within the
CubeSat, which were designed to connect via tubing to ensure the pressures would

31

equalize in both tanks. A tee fitting was designed along the connecting tube in addition
to a pressure transducer in order to monitor tank pressures. The tanks were each
designed to have a thermocouple in order to monitor the temperature of the hydrogen
peroxide, so as to provide the updated status on pressure dependent readings.
Directly under the propellant tanks, which is symmetric on both sides, is a tee
fitting that leads to a quick disconnect and the first latch valve. The filling process will
require a manual ball valve, which will detach after filling is complete. The quick
disconnect will reduce the weight of having 2 manual ball valves for filling, while
ensuring a proper seal. The first valve that will be used is a latch valve, which was
chosen for the reason of reducing power usage and which also acts as the first level of
redundancy. Unlike solenoid valves, latch valves do not require a continuous power
draw to remain open. Instead, latch valves use power to turn a mechanical part inside
the valve. This part keeps the valve open until power is supplied again to close the
valve. Pressure transducers will measure the pressure after the first latch valves to
determine if the valves opened successfully. Subsequently, the tubing from both tanks
will lead to a custom-made aluminum block that will converge to an additional latch
valve. This second latch valve serves as the second level of redundancy. Following this
latch valve is another pressure transducer to confirm the opening of this valve. Lastly,
the main solenoid valve is the third level of redundancy which will be the last valve
before the propellant is fed into the thruster. This is a solenoid valve, different from the
latch valve, functioning as the main control of operation for the thruster.
3.2.1 Delta-V Budget
Maneuvers will play a crucial role in transitioning the spacecraft between
different mission phases. Due to the constraints on the allowed mass and dimensions
of the spacecraft, it is imperative that these maneuvers be as efficient as possible in
order to minimize the overall mass of propellant required.
There will be two major maneuvers necessary from CubeSat deployment to
propel Triteia into Lunar orbit. The first is a trajectory correction to be performed as
soon as possible upon deployment from the ICPS, once detumbling is complete and a
communication link has been established. This maneuver ensures that the spacecraft
will be oriented to enter a polar orbit upon arrival at the moon. It was minimized under
the constraints that the spacecraft’s final orbit must remain stable for at least 365 days
and its inclination must remain between 85° and 95°. The second maneuver is
performed at periapsis in the retrograde velocity direction. This burn will take the
spacecraft off of its hyperbolic approach trajectory and onto an elliptic capture orbit. It

32

was minimized under the constraint that the CQC defines a lunar orbit as having a
perilune greater than 300 km and an apolune less than 10,000 km.
The following table summarizes the Delta-V allocation for each maneuver in the
mission, as well as the additional budget in case of misfires or other emergencies. All
velocities are in meters per second.
Table 3-5:Overview of Delta-V budget
Mission Phase Maneuver Name Delta-V
(m/s)
Comments
Cruise Trajectory Correction
Maneuver (TCM)
46
Orbit Insertion Lunar Orbit Injection
(LOI)
350
Error Margin 56.8 To be used for
additional correction
maneuvers, the
protection of historic
lunar sites, etc.
452.8 Grand total Delta-V for
the mission
3.2.2 Tanks
Table 3-6: Values calculated to design the propellant tanks.
Raw Value Marginal Value
Delta-V, VΔ 396.0 m/s 452.8 m/s
Dry mass, mf 6.482 kg 6.799 kg
Propellant mass, mp 1.910 kg 2.227 kg
Wet mass, mo 8.392 kg 8.709 kg
Burned off Mass, mburn,1 0.248kg 0.257kg
Burned off Mass, mburn,2 1.6623kg 1.726kg

33

First Burn Time, t1 82.143sec 85.246sec
Second Burn Time, t2 550.446sec 571.203sec
Mass Flow Rate, m˙ 0.00302 kg/s N/A
Specific Impulse, Isp 156.27 sec N/A
The tanks were designed to hold the propellant volume that would be needed
for mission maneuvers. An iterative approach was followed to acquire the mass values
of the dry, wet, and total masses. The very first constraint was to use the given delta-V
required for orbital trajectory and maneuvers. This value of 396 m/s was determined by
orbital dynamics as the total change in velocity needed. Subsequently, the propellant
mass and propellant volume were found using the equation below:
(E3.22)V g ln( ) Δ = Isp o mf
m
o
An estimated final dry mass, , of 6.482 kg was determined from the totalmf
mass of the other subcomponents within the CubeSat. To recall, RPA calculated the
specific impulse as 156.27 seconds. Rearranging and solving for the wet mass, ,mo
gave:
(E3.23)
The propellant mass was calculated and a 16.6% margin was added in order to satisfy
the requirements from the judge’s workbook :19
(E3.24)m m 1.91 kgmp = o − f =
(E3.25) (1 .166) 2.2272 kgmp, margin = mp + 0 =
The margin in propellant mass provided a new initial wet mass with a margin and a
new delta V:
(E3.26) m m .7092 kgmo,margin = p, margin + f = 8
(E3.27)V g ln( ) Δ = Isp o mf
m
o,margin
(E3.28)V 452.8 m/sΔ margin =
19
"Ground Tournament Submittal Requirements and Standardized Judging Criteria." National Aeronautics
and Space Administration, n.d. Web.

34

The Delta V budget for two burn times were calculated using the total value of
396m/s instead of 452.8 m/s. This was because the two delta V values calculated
by orbital dynamics, which were 46m/s and 350 m/s respectively, were desired
delta V values. Although these values were found taking the worst case scenario for
initial wet mass (14 kg), it was assumed that the mass of the satellite in comparison
to the celestial bodies would be so small, that it would not affect the desired delta V
values. Therefore, the burn times for these two delta V values were calculated using
the marginal value for initial wet mass but the raw value for delta V, indicating that
the extra propellant would be accounted for but never burned. Using the burn time
equation shown below :20
(E3.29) [1 ]tburn = m˙
mo,margin
− 1
exp( )ΔV
I gsp o
The first burn time with the preliminary values:
5.248 secondst1 = 8
After the first burn, the engine has expelled:
(E3.30) m t 0.257 kg of total fuelmburn,1 = ˙ * 1 =
Subtracting this from the initial wet mass, the new wet mass is:
(E3.31) mo,1 = m 8.452 kgmo,margin − burn,1 =
Updating the new total mass, and the second burn time:,m0,1 V ,Δ
71.609 seconds t2 = 5
Subtracting the amount of propellant expelled for the second burn, the final dry mass
of the cube satellite was found:
(E3.32)m m m 6.726 kgmf,new = o,margin − burn,1 − burn,2 =
20
Braeunig, Robert A. "Basics of Spaceflight: Rocket Propulsion." Basics of Spaceflight: Rocket
Propulsion. N.p., n.d. Web. 31 Jan. 2016. <http://www.braeunig.us/space/propuls.htm>.

35

The propellant mass including the margin was then the difference in the initial wet
mass and final dry mass:
(E3.33)m m .983 kgmp,new = o,margin − f,new = 1
The result of this value indicated that using and with the same raw valuem o m o,margin
for delta V produced nearly the same propellant mass. Taking the raw propellant mass
and dividing it by the new propellant mass:
(E3.34)/m 6.31%mp p,new = 9
This proved that the 16.6% margin changed the required propellant mass by 3.69%, a
small amount.
The propellant volume was found using the density of rocket grade hydrogen peroxide
and the propellant mass with the 16.6% margin in order to account for all of the21
propellant on the cube satellite:
(E3.35) 590.714 cmV cyl,tot = ρH O2 2
mp,margin
= 2.227 kg
.0014 kg/cm3 = 1 3
Since a dual tank configuration fits the volume constraints of a CubeSat better than a
single tank configuration, the propellant volume for each tank:
(E3.36) 795.357 cmV cyl = 2
V cyl,tot
= 3

The thickness, , of each tank was then found using the hoop stress equation belowxΔ
where it was assumed that the hoop stress was equal to the yield strength of22
Aluminum 7X11-T6, 5.674x108
Pa:σ
yield
= 23
(E3.37)σh,cyl = Δx
p ri
Rearranging the equation to solve for thickness, xΔ :
21
Corporation, 1966. Print..
22
MATHalino.com." Thinwalled Pressure Vessels. N.p., n.d. Web. 17 Jan. 2016.
23
Ventura, Mark. “Long Term Storability of Hydrogen Peroxide 41st AIAA.”

36

(E3.38)x Δ =
p ri
σh,cyl
Since the height of the CubeSat constrained the largest radius of the tank, the radius
was rewritten in terms of defined variables, yield strength and height:
Figure 3-15: Top view 2D and 3D representation of the tanks
(E3.39) 2r 2Δx 2(r x) 2(r ) 2r(1 )h = + = + Δ = + σh
p ri
= +
pi
σh
(E3.40) r = h
2(1+ )
pi
σh
(E3.41) tcyl = σ h
p ri
= σh
pi
h
2(1+
pi
σ )h
=
p hi
2(σ + p )h i
Here, it was found that the maximum design pressure, , for the tank,σh
accounting for the headloss, was 135 psia. The factor of safety (FOS) from table 3-9 in
Interface Definition Requirements Document (IDRD) is x1.5 for proof and x2.0 for
ultimate. Therefore, internal pressure was rated for 270 psia. The tanks were designed
to operate up to a pressure of 600 psi however, since previous iterations had shown
that tank pressures below 600 psia produced a very thin shell design. At 270 psia, the
tank thickness was found to be 0.0174 cm as opposed to a tank thickness of 0.0385
cm for 600 psia. From a practical approach, it was thought that reducing the tank
thickness to 0.0174 cm would present issues during fabrication and during handling.

37

Therefore, the tanks have a factor of safety (FOS) of 4.44. For a thickness of 0.0385
cm, the radius value was then found to be:
(E3.42) .286 cmr = pi
t σcyl h
= 5
As for the ends, 2:1 elliptical heads were chosen due to their cheaper cost and easier
manufacturability than hemispherical heads. Given the internal radius found above, the
head parameters were found using the image below :24
Table 3-7: Listed dimensions of the tank end caps.
Parameter Value
Internal Diameter 10.562 cm
Knuckle Radius 1.825 cm
Internal Crown Radius 9.553 cm
Internal Depth of Dish 2.643 cm
Length of Straight Flange 1.406cm
Material Thickness 0.038cm (same as cylinder: conservative)
24
"Elliptical Head Blank Dimensions." Titan Metal Fabricators Inc., n.d. Web.
<http://www.titanmf.com/wp-content/uploads/docs/TITAN-Elliptical-Head-Dimensions.pdf>.

38

Overall Head Height 4.087cm
According to the tanks’ intended use, the tank material is, by NASA standards
(NASA-STD-6012), in a class 4 environment because it will be exposed to a potentially
corrosive chemical (90% hydrogen peroxide). NASA-STD-6012 states that only the
1000, 3000, 5000, and 6000 series aluminum alloys are corrosion resistant. However,
Aluminum 7X11-T6 was chosen for the tank material because of its high compatibility
with rocket grade hydrogen peroxide . A material usage agreement must be requested25
for this aluminum alloy given its lack of availability within MAPTIS.
The assumption that the propellant volume was equal to the tank volume can be
interpreted as the cylindrical portion of the tank holding all the propellant. Rather, the
elliptical heads do not hold any of the propellant. The total ullage volume is then the
total volume of the heads. From SolidWorks, the total tank volume below was used to
find the ullage volume in each tank:
(E3.43)
Challenges in the design and use of the tanks were the volumetric constraints of
the CubeSat. The two tank design is from understanding that one tank would not fit
within the volumetric constraints. As a result however, the two tanks design allows
each tank to hold less pressure, be thinner, and weigh less. Tank configurations above
two were eliminated because the number of valves required and the complexity of
operation would outweigh (both literally and figuratively), the two tank configuration. In
addition to volumetric constraints, how the tanks would release the same amount of
fluid at the same time was another issue that presented itself. This issue was resolved
by connecting the two tanks above with the same tubing, allowing the two tanks to
reach equilibrium pressure.
The overall technology readiness level of the tanks is 5 due to the unique
methods of manufacturing and testing. Propellant tanks utilizing bladders to expel
propellants were used in the Syncom II, Syncom III, and other satellites. The tanks will
be tested on the ground by Ground Tournament 3.
25

39

3.2.3 Valves
3.2.3.1 Quick Disconnects
Manual ball valves were originally considered for filling but were replaced soon
afterwards. Manual ball valves use a ball with a hole carved through it to start and stop
flow. This presented a possibility that hydrogen peroxide would get trapped and begin
decomposing inside the ball, potentially causing a critical hazard. The manual ball
valves were strategically replaced with quick disconnects that hold back pressure in
order to reduce the weight and easily detach the connection after filling. These quick
disconnects were considered over check valves because they allow the propellant to
be taken out of the system without having it run through the engine.
3.2.3.2 Solenoid and Solenoid Latch Valves
The CubeSat propellant feed system consists of four valves, three of which are
solenoid latch valves and the last being a solenoid valve. All four valves are provided
by Moog Space and Defense Group. Figures 3-17 and 3-18 are drawings of the latch
solenoid and solenoid valves that will be utilized in the propellant feed system. Both
the latch solenoid and solenoid valves are exactly the same, except that the latch
solenoid valves have the ability to “latch” open or closed allowing the system to save
on power consumption.

40

Figure 3-16: Specification sheet of Moog’s (latch) solenoid valves.

41

Figure 3-17: Drawing illustration of Moog’s (latch) solenoid valves.
Figure 3-18: Side profile drawing illustration of Moog’s (latch) solenoid valves.
3.2.3.3 Pressure Relief Valves
The pressure relief valve (PRV) at the top of the feed system was considered
given the intrinsic property of hydrogen peroxide to decompose over time. The
continuous decomposition of hydrogen peroxide within the tanks would cause a

42

pressure increase, thereby presenting both safety and mission concerns for the SLS
and the cube satellite. As a result of design concerns and specifications within the
Interface Definition Requirements Document (IDRD), a pressure relief valve was
included in the design of the feed system. The integration, however, also presented
concerns of its own, namely, the offgassing of potentially hazardous vapor inside the
Middle Stage Adapter (MSA). This was an issue that was also described in IDRD.
However, the design choice of a bladder and a cost benefit analysis of integrating a
pressure relief valve to account for decomposing hydrogen peroxide was thought to be
safer than not integrating a PRV at all in order to completely contain any off gassing. As
a result, the system includes a PRV but other issues such as its frequency of actuation
and time of actuation remain to be found.
3.2.4 Thermocouples
Thermocouples are utilized in the CubeSat propellant feed system to give
accurate data and updates on the status and performance of the engine. Three
thermocouples are placed on the monopropellant engine to provide accurate
temperature information. Two out of the three probes are placed inside the
decomposition chamber to provide catalyst pack thermal information. The final
thermocouple is placed on the outside wall of the monopropellant engine to monitor
wall temperature.
The acquired thermocouples and thermal probes will be required to operate in
extreme environments. Temperatures inside the decomposition chamber are expected
to reach 1033.15K (1400 o
F). Consequently, thermocouples and probes will need to
have operating temperatures that can survive such extreme temperatures.
Furthermore, material compatibility is also an important characteristic to address in the
selection of thermocouples as some may come into contact with undecomposed
hydrogen peroxide.
3.2.5 Pressure Transducers
Pressure transducers are utilized in the CubeSat propellant feed system as well.
The pressure transducers are placed at various locations along the lines of the feed
system to confirm flow and to take real-time tank and chamber pressure readings. Six
transducers are used to accomplish this.

43

4. Static Fire System Design
Figure 4-1: The above image is a Plumping and Instrumentation Diagram (PID) created
by SEDS@UCSD to test the engine before installing it on the CubeSat.
The static firing test stand is a pressure fed system built to test the
monopropellant CubeSat engine designed by the SEDS@UCSD team. The system will
measure the engine’s pressure, thrust, temperatures, vibration, and flow velocity to test
for efficiency. The purpose of testing the engine is not only to verify the design but also
to determine the amount of pressure loss
The static fire system is a pressure-fed liquid propellant system with a 2800 psi
nitrogen k-bottle. Unlike the blowdown propellant feed system on Triteia, the goal of
the static tests is to obtain the steady state thrust. Due to the variances in components
between the flight feed system and test stand, the data obtained will be significantly
different if the test system were to be a blow down. The goal of the static tests is to

44

obtain the engine efficiency and confirm the designed thrust of 1 lbf at ideal steady
state operation.
During the test, the nitrogen k-botte will be strapped vertically behind the beam.
Nitrogen flows through a filter to refine any impurities and particles in the gas. ¼” (size
of the regulator) stainless steel tubing was used because it can withstand nitrogen at
the pressure of the k-bottle (2800 psi). The gas splits and one end flows through an
inline regulator, which allows the users to determine the operating pressures of the test
stand. Through the other path the gas flows through a dome regulator with a pilot
pressure from the inline regulator. This is done so that the mass flow rate of nitrogen
will match the mass flow rate of the propellants. The flow of nitrogen then goes through
an electrically actuated ball valve that is normally closed. This will give us control over
when to pressurize the tanks. Then the gas will flow through a check valve. The check
valve is in place to prevent any possible backflow of propellants to protect the valve
and regulators upstream of the check. From there, the nitrogen flows to a manifold
aluminum block, which acts as a 5 way fitting. The 5 ports used are for: the flow into
the block, out of the block into the tanks, a pressure gauge and transducer, a pressure
relief, and a vent valve. The pressure gauge is a visual indication of the pressure at this
point in the system and the transducer will give the user an electronic sensor reading.
The pressure relief valve is a high pressure pop safety valve that will release pressure
at 150 psi to make sure the system is not over pressurized. The vent valves are spring
return actuator valves, which are normally open. These valves open whenever the
system is not being pressurized before the test and after the test when the system
needs to be released of pressure. The pressure then flows through the top of each tank
and pressurizes the propellants. Below the 90% Hydrogen Peroxide tank with a 1379
kPa (200 psi) pressure, is a tee fitting connected to manual ball valve that will allow
filling of the tank. To actuate the ball valve, a dual actuator will be connected to 3-way
solenoid, which would allow flow into either the force close or force open passage of
the dual actuator. To control the actuation of these dual actuators, electronically
actuated solenoid valves were used in order to distribute pressure into the respective
ports of the main and vent valves. In addition, a compressor is connected to a solenoid
valve in order to open the main valve due to it's a activation pressure at 827.37kPa
(120 psi).
The calculated pressure of 200 psi accounting for head loss will give the
theoretical pressure to pressurize the system. An iterative approach will be followed
through measuring the efficiencies and thrust of the thruster, which will acquire the test
data needed to determine the actual ideal pressure. Using this data, and accounting for
the varying headloss for this system, a better understanding of the thruster can be
gained for the time testing on engineering model hardware can be done.

45

The static fire system used 0.635cm (0.25in) because of the sizing of the
regulator from the nitrogen k-bottle. An increase in tubing size from 0.635cm (0.25 in)
to 1.27cm (0.5in) took place to adapt the tubing to the same size as the pneumatic
cylinders. A reduction of tube size from 1.27cm to 0.3175cm (0.5in to 0.125in) was
used between the main valve dual pneumatic actuator and the engine to adapt to the
monopropellant thruster to achieve the desired flow rates of the cube satellite thruster.
5. Minimum Risk and Hazard Reduction by Design
The latch valves act as levels of redundancy during ascent to prevent the
thruster from firing prematurely. As soon as the CubeSat is deployed, there will be
pre-burn burp tests of the engine to warm up the chamber. In order to do this, the
sequence of valve openings will be as follows:
1. 1st Level of Redundancy, Latch Valve - Open
2. Confirm Pressure reading for successful test.
3. 2nd Level of Redundancy, Latch Valve - Open
4. Confirm Pressure reading for successful test.
5. 3rd Level of Redundancy, Solenoid Main Valve - Open for 500 ms.
6. 3rd Level of Redundancy, Solenoid Main Valve - Close.
7. Repeat steps 5-6 for 3 times.
After this sequence, the latch valves will remain open and the operational
controls for firing the thruster will become dependent on the solenoid main valve. Once
in space, there will be significantly less vibration from external sources that will prohibit
the main valve from actuating compared to the ascent. The main solenoid valve as well
as all of the latch valves will be thoroughly tested and certified to guarantee success
and prevent premature actuation.
There were a considerable number of design challenges associated with the
components. The system was designed to meet the triple redundancy system but
some of the parts included still need to be verified for their technology readiness level.
The system is designed to contain only AN fittings and not NPT fittings so as to avoid
using teflon for seals. Both the solenoid valve and the latch valves being considered for
the system still need to be verified as space rated valves.

46

6. Hazard Report Development (SLS Plan 217)
6.1. Summary of derivation of system MDPs
● MDP Derivation (Tanks): This derivation is presented above under the
tank section.
● MDP Derivation (Tubing):
To help determine the tank pressures, the major and minor head loss were calculated
to find the pressure drop across the propellant feed system. To start off, the volumetric
flow rate needed to be calculated by the given equation:
(E6.1) .16 m /sQ = ρ
m˙
= 1400 kg
m3
0.00302 s
kg
= 2 × 10−6 3
Once the volumetric flow rate was found, the velocity through the propellant line was
calculated.
(E6.2) .0589V line = D2
0.1273 Q
= (0.002159m)2
0.1273[2.16×10 (m /s)]6 3
= 0 s
m
To calculate the major head loss, the friction factor must be found. To find the friction
factor we must first find the Reynold’s number. Since the Reynold’s number is well
below 2040, making the flow laminar and allowing the use of the specific equation for
finding the friction factor.
(E6.3).145Re = μ
ρV D
= 1.23 kg
m∙s
(1400 )(0.0589 )(0.002159m)kg
m3 s
m
= 0
(E6.4) 42.08f = Re
64
= 64
0.145 = 4
Now it is possible to solve for the major head loss, which is given by the following
equation:
(E6.5) ( ) ( )hL,major = f l
D 2g
V 2
Using the headloss, the pressure drop can finally be calculated.
(E6.6)P h (ρg)Δ = L,major

47

This process is repeated for each straight length of pipe. Taking each length of pipe,
finding its pressure drop, and then adding them all together to find a total pressure
drop due to major head loss:
(E6.7)P PΔ total = ∑

Δ
After finishing everything for the major head loss, the minor head loss must be
calculated as well. Similarly to the major head loss, the minor head loss for every valve,
fitting, and bend must be calculated. The minor head loss is given by:
(E6.8)ξ( )hL,minor = 2g
V 2
Now that minor head loss has been calculated, we can calculate the pressure drop due
to the minor head loss. This can be achieved by the following equation:
= (E6.9)PΔ ρVξ2
1 2
The aforementioned equations showed that there will be about a 70 kPa (10.15 psia)
drop across the propellant feed system. The majority of the pressure drop resulted
from major head loss 66.094kPa (9.151 psia), whereas the minor head loss resulted in
less than 0.046kPa (0.005 psia) pressure drop, and can be regarded as negligible.
Since the tank pressures are 930792 Pa (135 psia), the inner radius of the tubing can
be found assuming the yield strength of AL 5254-H34 ( and using7000E6 Pa)σyield = 1
the outer diameter of 0.1375cm (⅛in):
σt = t
(p −p )ri o
uter diameter 0.1375cm r t o = = 2 + 2
where the hoop stress equation was rearranged to solve for thickness (t) and the result
was inserted into the second equation:
r(1 ) 1375 cm2 + σt
p ri
= .
The radius of the tubing was found:
1587 cmrtubing = .
From table 3-9 where the factor of safety is x1.5 for proof and x2.0 for ultimate, a
factor of safety of 2.0 was selected for the lines. Therefore, the maximum proof test
pressure was found as:
2.0(930792Pa) = 1861584Pa (269.99psi)
MDP Derivation (Latch Valve): A suitable latch valve could not be found

48

MDP Derivation (Solenoid Valve): 400 psig = 2.578x106
Pa (provided by Moog)
Table 6-1: Factors of safety for design of pressure systems
6.2. Potential Hazards (flammability, toxicity, explosion, corrosion):
According to “ Centers for Disease Control and Prevention”, hydrogen peroxide 2
decomposes on warming or under influence of light producing oxygen, which
increases fire hazard. The substance is a strong oxidant and reacts violently with
combustible and reducing materials causing fire and explosion hazard
particularly in the presence of metals. Attacks many organic substances, e.g.,
textile and paper. Poisonous gases are produced in fire. containers may explode
in fire. Hydrogen peroxide may ignite combustibles (wood, paper and oil).
Contamination of hydrogen peroxide can possibly yield a self-quickening
deterioration response, contingent upon the relative rate of warmth misfortune
and the rate of decay, which are affected by vessel size, ambient temperature,
insulation, initial concentration, amount of contamination, etc. Although high
concentration hydrogen peroxide solutions can be put away securely, with
special engineered safety precautions and tight control of systems and the
quick environment, the quickened response energy connected with higher focus
arrangements gives an expanded danger to the capacity of high fixation
hydrogen peroxide arrangements, along these lines Hazard and Risk
Assessment forms regularly lead most mechanical users to stay away from
capacity at fixations more prominent than half. Flammable material in the vicinity
of high fixation hydrogen peroxide might rapidly blast into fire. Ignition, with
lower fixation hydrogen peroxide concentrations, might be deferred after initial
contact.

49

At ambient temperature, it has been difficult to get a spreading explosion
in commercial hydrogen peroxide solutions. Organic material dissolved in
hydrogen peroxide, especially in stoichiometric amounts, may form an explosive
mixture. Under customary storage and taking care of conditions there is no
danger of a vapor phase explosion.
The more common danger is a "hazardous" pressure rupture due to a gas
generation rate exceeding the vent capacity of a container. Catalytic
decomposition results in the freedom of oxygen and heat. Vaporization of
solution water results in further centralization of the hydrogen peroxide solution.
Under ordinary circumstances, hydrogen peroxide at factories is put away at
half fixation. There is adequate water present to retain any warmth advanced
because of typical moderate (unaccelerated) peroxide disproportionation. There
is likewise limit in the "framework" to retain the warmth of a moderate level of
item tainting, and give adequate time to undertake actions that may securely
keep away from the framework going "critical" in a accelerating decomposition
reaction.
The decomposition procedure is normally moderate to begin with, and
might take days or weeks to end up 'critical'. In case of extreme contamination
in storage, the disintegration can reach hazardous extents rapidly. A
temperature increment of 1-2°C every hour, at 30-35°C, is indicative of a
decomposition event. Decomposition reactions in storage can be quenched by
external cooling or dilution.
Table 6-2: Preliminary Table
Item Yield
Strength
(Pa)
Ultimate
Strength
(Pa)
Max
Design
Pressure
(Pa)
Factor
of
Safety
(actual)
Proof
Test
Factor
Leak
Rate
Method
Common
Integrity
Required
Pressure
Relief
Valve
N/A N/A N/A N/A N/A N/A N/A
Cross
Fitting

50

Quick
Dis-
connect
AL
7X11-T6
Tanks
5.6743
E826
N/A 2.758 E6 4.44 4.44 N/A N/A
Tee
Fittings
Latch
Valves
AL 5254
Custom
Block
17000E627
35000E626
N/A N/A N/A N/A N/A
Solenoid
Valve
N/A N/A 2.758E628
2.0 2.0 N/A 1E-6
SSC/sec at
4.137E619
AL 5254
Tubing
17000E626
35000E626
930792 2.0 2.0 N/A N/A
Inconel
718
Engine
6.34 E829
(min; as
built)
9.8 E820
(min; as
built)
8.61845
E530
3
(min; all
parts)
3 N/A N/A
The factor of safety must be specified if it is different than what is required from the specified table.
The proof test factor is equivalent to the maximum proof test pressure divided by the maximum design
pressure.
The leak rate method is the method used for hazardous materials.
Common Integrity required is simply the maximum allowed leak rate.
26
27
"ALUMINUM 5254." Alloy Digest (1979). ASM International. Web.
<http://maptis.ndc.nasa.gov:2052/ac/dataFile.aspx/al157.pdf?dbKey=grantami_ac_datasheets&data=20053
5&record=88412>
28
"Valve, Propellant Dual Solenoid Installation Moog Model." Perkin Elmer, n.d. Web.
<http://www.moog.com/>.
29
Stirling, Robert. Material Safety Data Sheet. Fairfax, VA: Defense Mapping Agency, Safety & Health
Division, Human Resources Directorate, 1993. Electro Optical Systems. Web.
<http://gpiprototype.com/images/PDF/EOS_NickelAlloy_IN718_en.pdf>.
30
Ponomarenko, Alexander. Rocket Propulsion Analysis. Computer software. Rocket Propulsion Analysis.
Vers. V2.2. N.p., n.d. Web. 30 Jan. 2016. <http://www.propulsionanalysis.com/>.

51

If the proof test factor is less than 1.5 x MDP, provide an explanation.
6.4. Proposed Verification Approach for Controls to Ensure Pressure Integrity:
The first verification approach is the design of the tanks. The tanks were
designed with a higher factor of safety in order to account for any overpressurization
due to decomposition of rocket grade hydrogen peroxide and oxygen loss into the
MSA. The second verification approach are the pressure transducers located
throughout the system. The pressure transducer above the tanks was designed to
ensure that both tanks hold the same pressures. There are also pressure transducers
located after every type of valve in order to measure the pressure differential across
them and monitor when they open. The practical approach will be to construct the
propulsion system and run the system at higher pressures than the expected operating
pressures.
6.5. Proposed Verification Approach Including Controls to Prevent Leakage:
The proposed verification approach to prevent leakage will be to assemble the
propulsion system in such a way that minimizes leakage. Metal gaskets have been
considered to seal all AN fittings. After assembling all the components, every fitting in
the propulsion system will be torque striped to get a physical indication of any leaks
associated with loose fittings. Just like for the proposed verification approach to
ensure pressure integrity, the approach to ensure minimal leakage will also involve
running the system at higher pressures.
7. Material Compatibility, Toxicity, Flammability, and Toxic Off Gassing
7.1.Engine, C-Ring, Anti-Channels Baffles: Inconel 718
Inconel 718 was used to print the engine’s nozzle, combustion chamber, and
diffuser plates as well as the interface between the nozzle and the combustion
chamber and the catalyst pack’s anti-channel baffles. Data for this metal with 90%
H2O2 is not provided in the Materials and Processes Technical Information System
(MAPTIS) used by NASA, therefore the SEDS team will conduct experimental testing
and publish the results in a paper titled, “Testing the Compatibility of Hydrogen
Peroxide With Inconel 718”.
The goal of the paper is to establish whether compatibility increases or
decreases with increasing exposure times and temperatures through active oxygen
loss tests (AOL) and stability tests. The tests will focus on ensuring contact between
Inconel 718 and rocket grade hydrogen peroxide does not react to form explosive
by-products, or degrade the test metals in the form of: discoloration, weight change,
visual cracking, corrosion, oxidation, or mechanical/structural property loss.

52

Additionally, it will assess each material’s contamination level after exposure to rocket
grade hydrogen peroxide and the catalyst bed.
7.2. Nickel Screens
Nickel 200 is used in the catalyst pack as a series of mesh screens located
primarily near the end of the package. Their purpose is to support the more malleable
silver screens and prevent deformation from occurring due to the high temperatures of
the chamber during the decomposition of hydrogen peroxide into oxygen and
superheated steam.
Information about this material’s corrosive properties under such an
environment were gathered from NASA’s Materials and Processes Technical
Information System (MAPTIS). Nickel 200’s corrosion resistance to water and humidity
is listed as being ‘Excellent’. MAPTIS also lists Nickel 200 as being non flammable,31
and states that all nickel compounds should be regarded as toxic and that some can
cause cancer and/or fetal abnormalities. Due to a lack of information on MAPTIS32
about this material’s toxic off gassing properties a material usage agreement will need
to be provided.
7.3. Silver Screens
The majority of the catalyst package is composed of 99% pure silver screens.
Silver decomposes hydrogen peroxide into oxygen and superheated steam when they
come into contact. While the silver meshes near the front diffuser plate will be mainly
exposed to the hydrogen peroxide, the silver meshes at the end of the catalyst pack
close to nozzle will be in a steam and oxygen rich environment as the majority of the
hydrogen peroxide has decomposed. Information about this material’s properties as
well as the following two corrosion data tables, one of silver in oxygen and the other
silver in steam, are provided by the Materials and Processes Technical Information
System (MAPTIS) and the ASM Alloy Center Database Corrosion Performance Data. It
lists silver’s corrosion resistance to water as excellent.33
31
  "Nickel, Commercial Purity, Grade 200, Spring Temper, Wire." Materials and Processes Technical
Information System (MAPTIS). Web.
<https://maptis.ndc.nasa.gov/mi/datasheet.aspx?record=367445&search=true&locate=false&dbkey=MI_Mat
erialUniversePolymer>
32
  "Nickel, Commercial Purity, Grade 200, Spring Temper, Wire." Materials and Processes Technical
erialUniversePolymer>
33
  "Silver, Commercial Purity, Fine, Soft (annealed)." Materials and Processes Technical Information
System (MAPTIS). Web.

53

Table 7-1: Corrosion data table for silver in oxygen34
Rate (mil/yr) Equivalent rate
(mil/yr)
Temperature
(°F)
Condition/Comment Source of data
≤2 ≤2 212 Pure, Attack
becomes
appreciable at 200 C
(390 F)
Smith and Zysk
1987e
Table 7-2: Corrosion data table for silver in steam35
Rate (mil/yr) Equivalent
rate (mil/yr)
Temperature
(°F)
Condition/Comment Source of data
≤2 ≤2 1110 Pure, Without
Pressure
Smith and Zysk
1987e
The total amount of time the engine stays in firing state is 32.5 min, hence the
corrosion rate of 2 mils penetration per year for silver in oxygen and 2 mils penetration
per year for silver in steam is negligible due to the relatively short amount of time silver
will be exposed to oxygen and steam. Silver is not toxic. Due to a lack of information36
on MAPTIS about this material’s toxic off gassing properties a material usage
agreement will need to be provided.
7.4. O-ring
The engine contains one Perfluoroelastomer (FFKM) o-ring used to create a seal
between the injector plate and the engine’s body to prevent hydrogen peroxide from
seeping out the engine. Perfluoroelastomer FFKM was considered as the material for
the O-rings in the front diffuser plate for its temperatures and material properties which
erialUniversePolymer>.
34
“Silver, Silver in Oxygen”, ASM Alloy Center DataBase, Materials and Processes Technical Information
<http://maptis.ndc.nasa.gov:2052/ac/index.aspx?profileKey=grantami_ac_corrosion>
35
“Silver, Siver in Steam”, ASM Alloy Center DataBase, Materials and Processes Technical Information
36
  "The Facts on Silver." Silver FAQ. Web. 04 Feb. 2016.
<http://www.dartmouth.edu/~toxmetal/toxicmetals/moremetals/silverfaq.html>.

54

were cross referenced through MAPTIS. FFKM received a satisfactory rating for its
compatibility with both 87% hydrogen peroxide and distilled water, which hydrogen
peroxide would most closely resemble after decomposition, as well as a non
flammable rating. Additionally, Viton near or above 392ºF might release small37
amounts of hydrogen fluoride which is toxic and may cause prolonged irritation to the
respiratory tract.38
7.5 Bolts
Zinc yellow-chromate plated steel is used for the bolts that will secure the three
different modules. The standard for high-strength cap screws, these are made from a
steel alloy and have a minimum tensile strength of 150,000 psi. They also meet ASME
B18.2.1 and SAE J429.v specifications. The zinc yellow-chromate plating makes the39
bolts rust resistant. Information on the specific type of steel alloy used in this bolt was
unavailable from the manufacturer, however a customer service representative assured
a certificate of chemical composition will be available after purchase. A material usage
agreement will need to be provided for these zinc yellow-chromate plated steel bolts
given its lack of availability within MAPTIS.
7.6 Tanks
The tanks are made of Aluminum 7X11-T6. A material usage agreement will
need to be provided for this tank material given its lack of availability within MAPTIS.
37
  "Perfluoro Elastomer (FFKM)." Materials and Processes Technical Information System (MAPTIS). Web.
<https://maptis.ndc.nasa.gov/mi/datasheet.aspx?dbKey=MI_MaterialUniversePolymer&record=369046&hist
ory=16467&locate=True>.
38
  "DuPont™ Viton® Handling Precautions for Viton® and Related Chemicals." (2010). Chemours. Web.
<https://www.chemours.com/Viton/en_US/assets/downloads/HandlingPrecautions_VitonandRelatedChe
micals.pdf>.
39
  "McMasterCarr." McMasterCarr. Web. <http://www.mcmaster.com/#92620a401/=10z3p6n>.

55

8. Corrosion, Prevention, and Implementation Plan
Over Pressurized Tank
The tanks have a safety factor of x3 in addition to the low tank operating
pressure of 135 psi. The pressure increases can result from active oxygen loss from
the decomposition of hydrogen peroxide and pressure difference in ascent to a
vacuum. The tanks can corrode over time, especially after sitting in storage for months;
however, this was accounted for in the design of the tanks. Since the tanks were
designed considering these scenarios and was over constrained, it is very unlikely to
occur. If the tanks were to over pressurize, the whole propulsion system will be
dysfunctional because of the lack of pressure having a significant impact on the
mission.
Flanges Not Properly Secured
The installation of the flanges is an intricate procedure; however, due to the
multiple tests, practice, and experience acquired in assembling the engine using a
torque wrench, it will not be likely that the flanges will leak. If they were to, however,
there will be a lot of wasted propellant and rapid pressure loss, which will produce a
much lower thrust. This will impact the mission moderately depending on the phase of
the mission this happens.
Pressure Transducer Failure
Pressure transducers are known to be robust especially when certifications for
space grade sensors are issued. It is unlikely that the pressure transducers will fail
because of this. If they were to, however, since these transducers monitor the tank
pressures, the initial valve actuations and reaction chamber pressure, there would be
no way for mission control to have eyes on the components they are sensing. The
chamber pressure is significantly impacted however, because this determines the

56

thrust of the engine and will give a reading of engine functionality. The other
transducers impact the mission moderately because there can be alternate ways to
measure the pressures in tanks and valves, which are not the most critical data
acquisition reading.
Engine Not Properly Pre-Heated
The engine must be pre-heated with short bursts before the first burn can be
initiated. If the engine is not pre-heated, it will have a minor effect on the mission
because extra fuel and pressure will be used to heat the engine up before efficient
thrust will be produced. This, however, is accounted for with extra marginal propellant
as well as pressure. This may be a possible occurrence due to an improper procedure
in space; however, there will be multiple ground tests to practice firing the engine at
max efficiency before it is flown.
Thermocouple Failure
Thermocouples and thermal probes are utilized in the propellant feed system to
monitor engine and tank temperatures. A failure in the performance of the
thermocouples that are placed on the catalyst pack would result in the inability to
monitor catalyst pack temperature. Such an occurrence would mean that the readiness
of the catalyst pack for an engine burn would be unknown. This could result in an
inefficient engine burn and an insufficient amount of propellant to complete the burn.
Such a risk can be considered very unlikely due to the simple nature of the
thermocouple and the impact of this risk would be negligible due to risk mitigations.
Modes of risk mitigation include extensive testing to determine the correct number of
pulses to adequately warm up the catalyst pack for a full burn.
Engine Not Properly Insulated
During CubeSat operation, it is possible for the engine to not be properly
insulated due to the insulating mylar shaken off by vibrations from SLS and the
possibility that the aluminum tape isn’t properly sealed. An improperly insulated engine
may cause satellite components to overheat and further creating a necessity for the
satellite engine to shut down to cool the system. Restarting the engine after shutdown
will waste propellant, power, and lower the catalyst pack life. Adhesive thermal
insulation on the engine will decrease the probability of the risk occurrence from
possible to very unlikely. However, the consequences of risk impact after mitigation will
remain the same.

57

Propellant Line Leak
The possible risk of a leak in the propellant lines may be caused by bad AN
fitting connections due to gasket failure, flare misalignment and/or broken flare. While
the risk is unlikely to occur, the impact of the risk is significant. If the propellant line
experiences a leak then the satellite will have an insufficient amount of propellant to
complete the mission. The probability of a propellant line leak can be minimized
through extensive testing and leak checks of the propellant feed system to ensure
proper line fittings. The impact of a propellant line leak will remain the same.
Valve Failure
It is possible for an electronic signal failure and the existence of a cryogenic
environment to cause the valves in the propellant feed system to fail to actuate during
operation. The electronic and mechanical failure of the valves may cause the engine to
miss its desired burn time which may result in the satellite entering into a solar orbit. To
reduce the risk, valve triple redundancy is designed to ensure that the possibility of
failure with one valve is mitigated by the existence two other valves. Although the
probability of valve failure to occur remains the same, triple redundancy lowers the
impact of failure to moderate.
Tank Bladder Burst
There are several issues in regards to the use of tank bladders, namely, internal
wrinkling during the expulsion cycle. If the bladder is elastomeric, internal wrinkling can
be avoided at ambient temperatures but not at cryogenic temperatures. At cryogenic
temperatures, the majority of elastomeric materials become extremely brittle and lack
the ability to withstand extensive wrinkling and folding without fracturing. Since in
space temperatures were given as -143 F to 200F by IDRD, it is possible for this to
occur. Other issues presented for bladders are their level of porosity. This is essential
in order to determine whether offgassing will occur. As a result, the ideal properties of
bladders are a high cryogenic cycle life, low values of porosity, minimum weight, and
barrier materials that remain flexible at cryogenic temperatures and resistant failures
caused by rolling three-corner folds. Accounting for these ideal properties as risk
mitigation lowers the probability to very unlikely.
Inconvenient Pressure Relief Valve Actuation
The pressure relief valve will only actuate if the propellant tank is over pressurized
which may be caused by excessive decomposition of hydrogen peroxide under high
propellant storage temperature. Although the risk is likely to happen, the
consequences are minor. To reduce the risk, a bladder has been implemented in order
to prevent hydrogen peroxide vapor from mixing with the pressurant gas. This can be

58

considered to be true as long as the bladder has a low porosity and remains
undamaged.
Table 7-3. Natural Environment during integration and assembly

59

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