Clay Blanchard characterized the performance of the green monopropellant AF-M315E in a 1N thruster. A test stand was designed and built to test the thruster under varied conditions inside a vacuum chamber. The thruster assembly utilized an iridium catalyst bed and injector to decompose the propellant. A series of heat exchangers were used to cool the hot exhaust gases before entering the vacuum pump. Initial testing verified that the catalyst bed heater functioned properly to preheat the catalyst to its operating temperature of 340°C. The overall goal was to collect performance data on the thruster and observe degradation of the catalyst over repeated firings.

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Characterization of AF-M315E in a 1N
Thruster
Clay Blanchard
Marshall Space Flight Center
August 8, 2014
Reviewed by NASA Mentor
Kevin Pedersen
ER – 23

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Characterization of AF-M315E in a 1N Thruster
Clay Blanchard1
Louisiana State University, Baton Rouge, LA 70820
NASA Marshall Space Flight Center, Huntsville, Alabama, 35812
AF-M315E is a “green” monopropellant developed by the USAF. This propellant offers advantages over
industry standard solutions such as hydrazine because of the simplified launch processing associated with propellant
toxicity. The motivation for the development of “green” monopropellants is to find alternative solutions to
Hydrazine. Hydrazine is a widely used propellant for many RCS thrusterapplications and is highly toxic, expensive,
and hard to store. The propellant is widely used because of its high reliability on orbit. The scope of this project is to
characterize the propellant AF-M315E. This propellant is an ionic liquid, which is decomposed and combusted by
catalysis in the thrusterhead end. The green monopropellant to be studied and characterized will be the Air Force
developed AF-M315E (a Hydroxyl Ammonium Nitrate based propellant). Analyses will be performed to predict the
performance of the thrusters,prepare and test the ground test system, prepare and lead design and test reviews, and
test and analyze the thrustersystems.
I. Introduction
The use of hydrazine has been dominant in monopropellant propulsion systems for many years. Problems
with this propellant include expensive storage, handling, disposal procedures, and most importantly propellant
toxicity. Traditional green propellants are used for low specific impulse and thrust applications. The purpose of AF-
M315E is to provide low toxicity, while having specific impulse comparable to that of Hydrazine. The 1N thruster is
an ionic liquid monopropellant thruster that, in this project, utilizes AF-M315E, an Air Force developed propellant.
AF-M315E reacts with an Iridium catalyst bed that is contained inside the 1N thruster. The decomposition of AF-
M315E requires the catalyst bed to be preheated. The monopropellant decomposes into different gases, which
provide thrust from expansion through the nozzle. This decomposition is hypothesized to result in higher thruster
performance than hydrazine. The propellant is pressurized at the top of the Propellant Tank using gaseous nitrogen.
The main goal of this project is to test and observe the performance of AF-M315E. In conjunction to the
aforementioned goal, the catalyst performance and life cycle will be sought after.
II. Design
A. Test Stand
The test stand described in this section was built to accomplish the goal of characterizing AF-M315E. This
stand will be used as a means of providing a secure structure, and a low pressure. All of the supporting structures for
the test stand were manufactured in MSFC building 4205 labs 104, 110, 111 and 106. This included fabricating a
test stand and designing and constructing plumbing for the fluid system. The fluid systemincludes the high and low-
pressure systems along with the cooling and propellant feed system. Precautions were taken to avoid over
pressurization of the tubing that can arise from misuse of the system. These precautions included following proper
operating procedures and safety measurements. To assure proper pressure rating in the tubing, hydrostatic testing
was done with all test stand plumbing. The pressure was set to twice the operating pressures. The test stand and
fluids system were designed such that steady state firing and pulse firing could be tested. A series of
instrumentation were integrated into the system in order to monitor the pressure and temperature of the fluid system
and certain thruster performance parameters, such as chamber pressure and temperature, mass flow rate, and thrust.
Once all of the test stand components were fabricated, every wetted article including valves, fittings, tubing, and
instrumentation were deemed compatible with AF-M315E according to safety regulations.
1 Research Assistant, Space Propulsion Systems, Marshall Space Flight Center

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1. Fluid System
The system is broken down into 5 sub-systems: High Pressure, Low Pressure, Propellant Feed, Vacuum and
Cooling systems.The high-pressure systemis used to pressurize the propellant tank and also provide a gas feed for a
GN2 purge to the vacuum chamber. The low-pressure system is used to provide inlet pressure to the electronic
controller that controls the pneumatic regulator on the high-pressure system. The propellant feed systemis used to
deliver the propellant into the thruster. This system starts at the propellant tank and ends with the thruster. The
propellant tank is pressurized by the high-pressure system, which causes the propellant to flow to the thruster. Along
the way to the thruster, a pressure transducer and flow meter show the properties of the propellant respectively. The
last sub-system is the coolant system. This system was created to cool the vacuumchamber, load cells, and exhaust
gas to the vacuumpump. Figure 1 shows the schematic of the entire fluid system.
A more in-depth description of the 5 sub-systems are shown below:
1a. High Pressure System
The propellant system is pressurized and purged with nitrogen supplied by a nitrogen storage bottle (K bottle).
The gas storage bottle supplies nitrogen at 4000 psig to two pressure regulators. These regulators are connected in
parallel with isolation valves (hand operated) to provide the systemwith pressure. The purpose of the hand-operated
regulator is for the system propellant purge, and the pneumatic regulator (remote operated) will be used during
testing to assure safety. These regulators flow nitrogen through either a solenoid valve or hand valve, connected in
parallel. Then through a series of two check valves. Then finally flows into the top of the propellant tank. A third
pressure-reducing regulator is tapped off of the K-bottle to provide low pressure purging of the vacuum chamber.
This is part of the cooling system as well. Relief valves and a burst disk are installed for safety precautions incase of
a systemover-pressurization. Pressure gauges and pressure transducers are located within the systemto observe the
regulated pressure in the systemat various locations.
1b. Low Pressure System
Low-pressure nitrogen is supplied from a facility line regulator, located inside the test cell (building 4205; 104).
The nitrogen supplies pressure to a current driven pressure transducer. The transducer will sense the pressure, which
will help provide feedback on regulating pressure to the pneumatic regulator on the high-pressure system. Another
important function of this system is to provide pressure to the pneumatic normally open valve (NOV). In case of
over-pressurization in the system, the NOV will vent the gas to the atmosphere.
1c. Propellant Feed System
The propellant feed systemstarts with the propellant tank and ends with the thruster. The propellant tank consists
of a short SST tubing, 6” long, with an outer diameter of ¾”. The system includes a filter, turbine flowmeter,
solenoid shutoff valve, hand operated valve, and the thruster. The propellant tank will be filled with propellant by
removing the cap off the top of the tube and distributing propellant into the top of the tube using a syring e or pipet.
The propellant tank will be pressurized up to 400 psig during test operations in order to flow propellant through the
system After the test, the remaining propellant will be drained out through the hand-operated valve or thruster. The
system will then be flushed with distilled water both through the thruster and out the valve. A camera will monitor
the propellant systemfor leaks, and also the pressure gauges during the testing operations.
1d. Vacuum System
The importance of the vacuum system is so that the thruster can be tested at low pressures. Testing at low
pressures allows more accurate results. The 1-Newton thruster that is being tested is located inside a small vacuum
chamber, with a volume of 2.62𝑓𝑡3
. Hot gas from the vacuum chamber flows through a shut off valve (hand valve)
outside the chamber, then through a cross-flow water-jacketed heat exchanger, through the rotary vacuum pump,
and finally discharges into the top of a drum located outside the walls of building 4205. The heat exchanger also
functions to cool the exhaust gas to prevent damage to the vacuum pump. The SST drum is used to collect and
condense fluids instead of exhausting directly to the atmosphere. A vacuum gauge (in bunker) and vacuum

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transducer (on chamber) will be used to monitor the vacuum chamber pressure at all times. An Infrared camera will
be used to monitor the internals of the vacuum chamber during the test.
1e. Coolant System
The purpose of the coolant system is to do just that; cool the hot exhaust from the thruster before it enters the
vacuum pump. It is also used to cool the load cells (within the chamber) with the nitrogen purge. In order to
effectively cool the thruster exhaust, three water-cooled heat exchangers are added in series. A flat plate heat
exchanger is located below the thruster, a copper coil below the thruster (on top of the flat plate), and a cross -flow
plate heat exchanger (outside the chamber). The hot exhaust gas will fire directly into the coil and hit the flat plate.
After this, the gas will get sucked through the vacuum tube and through the cross -flow heat exchanger outside the
chamber. The cross-flow heat exchanger will cool the gas even more before it enters the pump to prevent pump
damage. Water is supplied from a potable water source inside the test bunker. Water is flowed through these three
heat exchangers and ejected into a sink drain within the lab.
Figure 1. Fluid System Schematic
2. Part Servicing
The components of the fluids system were ultrasonically cleaned and passivated. The cleaning process involved
degreasing the components with a 15-minute soak in the ultrasonic cleaner. An employee at NASA completed all of
the passivation process. After these processes, all of the components were sealed in clean bags to ensure the parts
remained uncontaminated. When assembling the fluid system, all of the parts were handled with gloves, to prevent
contamination.

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3. Fabrication
Prior to beginning construction of the test stand, a CAD model was created. The vacuumchamber and test stand
were modeled in PRO-E. The height of the stand was determined based on the constraint of the viewport of the
vacuum chamber. The thruster needed to be seen through the viewport in the center of the chamber during t esting.
The width of the stand was already determined based off the physical limitations of the chamber. The depth was
built such that a coil and cold plate heat exchanger could fit under the stand. The stand was composed of four T-
Slotted aluminum legs and a flat SST plate on top, which is where the thruster assembly hangs. Many of the actual
thruster assembly parts were water-jet cut out of SST plates. The heat shield on the assembly was created by hand
pressing a ball bearing into a thin SST shim. The outcome was a half spherical shape with holes around the
perimeter to mount to a flange. Other parts were machined at the component development area (CDA), by lab
technicians. A pulley system was manufactured for a one point in-vacuum calibration of the load cells. The support
for the pulleys were water-jet cut out of SST and spot welded to the top of the thrust stand.
Figure 2 shows the thrust stand mounted inside the vacuum chamber. Note that the thruster will be fired directly
down into the copper coil heat exchanger shown in the picture.
Figure 2. Thrust Stand
The plumbing for the fluid systems was manufactured out of 0.035” and 0.049” thick SST tubing. The tubes
were hand bent and flared using a tube-flaring machine. The entire fluid systemwas mounted to an aluminum panel
for organization. Figure 3 shows the fluid systemincluding HOR, gauges,pressure transducers,and solenoid valves.

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Figure 3. Pressure System
B. Thruster
1. Injector
The injector was custom built by Plasma Processes. The injector consists of a 1/8” SST tube for the propellant
inlet and a flat face with 100micron diameters holes on the injector face. The injector is threaded so that it can
connect to the nozzle. On the perimeter of the injector is a flange. This flange is used to attach the thruster to the
thrust stand. Figure 4 shows the thruster, catalyst, injector, and load cell mount starting from the left respectively.
Figure 4. Thruster Assembly
The load cell mount was manufactured with a water-jet cutter in-house. The mount is made of 1/8” SST plate.
The triangular shape was designed in order to shed weight of the part. Figure 5 shows the load cells, load cell mount,
and flange spacer from top to bottomrespectively.

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Figure 5. Load Cell Mount
Figure 6 shows the scale of the full thruster assembly. This assembly is mounted above the copper coil as
described above.
Figure 6. Thruster on stand

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2. Catalyst Bed
For monopropellant thrust applications of AF-M315E, a catalyst must be present to start the decomposition
reaction. Plasma Processes also manufactured the catalyst that this systemused and therefore is proprietary. Figure 7
shows the monolithic catalyst bed at various magnifications. The catalyst is composed of Iridium. In order for the
reaction to occur, the catalyst bed must be preheated to about 340°C. In the testing section below, catalyst heater
verification was done. The results are shown in Fig. 9.
Figure 7. Iridium Catalyst
3. Nozzle
Plasma Processes manufactured the nozzle used in this system as well. The manufacturing of the nozzle
consisted of electroforming, with the substrate being Rhenium and the coating being Iridium. The Iridium
coating is used to protect the part fromoxidation. The nozzle can be seen in Fig. 4 and 6 above.
C. Heat Exchanger
Since the thrusterwill be expending very hot gasses during the test,a series of heat exchangers will be used to
absorb heat from the gas. The decision was made to have the thrusterfire directly down onto a plate heat exchanger.
The copper coil was added in order to cool and contain the surrounding gas. Both the plate and coil heat exchanger
were located in the vacuum chamber. The cross flow heat exchanger was mounted outside of the chamber to further
cool the gases before they enter the vacuum pump. If the pump were to expel these hot gasses,it would quickly burn
up. This is why the heat exchangers are very important. The copper coil heat exchanger was created by hand
bending 3/8” coppertubing around a solid aluminum cylinder of the diameter needed. The result was 10 loops of
coil. Wire was used in between the loops of the copper in order to tightly pack the coppercoil. The cross-flow and
plate heat exchangers were purchased. The three heat exchangers were assembled in series. The inlet of the heat
exchanger assembly was hooked to a water tap located in the lab. The exit was attached to plastic tubing and
discharged into a sink. The three heat exchangers can be seen in Fig. 8 below.

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Figure 8. Heat Exchanger
III. System Testing and Safety
A. Testing
1. Objectives
The primary goal of this development was to characterize the performance of the monopropellant AF-M315E. A
secondary goal for this project was to observe the degrading of the catalyst bed over time.
Performance of a 1N AF-M315E thruster will be assessed by:
• Flowing AF-M315E propellant through the thrusterwhile measuring flow rate, thrusterinlet pressure,
and thrust
• Perform tests at different inlet pressures
• Measure thruster’s chamber pressure and combustion temperature
• Measure the thruster’s outersurface temperature using an IR camera
• Vary test firing duration and time between firings
• Repeat a baseline firing periodically to determine if performance has changed
• Measure temperature at thrustervalve interface
• Measure temperature near injector tube supply O-ring
As mentioned above, the catalyst bed must be preheated in order to combust the monopropellant. The bed must
be heated to 340°C. The test performed was to make sure the heater functioned properly. The heater was only heated
to around 130°C to test functionality. Figure 9 shows the thrust assembly with the preheater turned on.

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Figure 9. Catalyst heater with Infrared Camera
D. Test Matrix
The different testing plans that are to be completed are listed in the following table:
Test Description
Trim/SS tests Dial in test pressures and developed steady state response data
The following test matrix will be performed:
Test Description Inlet Pressure [psig] Duration [sec]
Trim Facility checkout 300 5
SS SS 400 10
SS SS 350 10
SS SS 300 10
SS SS 250 10
SS SS 200 10
SS SS 150 10
Table 1: AF-M315E Monopropellant Test Matrix
E. Safety and Hazards Analysis
The following safety analysis was performed by this project’s team leader, Gustavo Martinez. Since this project
used high pressure and a class 1.3 explosive, several measures of safety had to be taken into considerations in ord er
to run the planned tests safely. The high pressure comes form a 5000psi K-Bottle. In accordance with NASA safety
standards, a Hazards Analysis had to be addressed in order to complete the Test Readiness Review. The Hazards
Analysis or Assessment was performed to enhance safety of the AF-M315 1N Thruster testing and to ensure
compliance with NASA, OSHA, and national consensus standards. The assessment also serves as a summary of
identified hazards to be used by decision makers when examining the risk level for the test.
Each safety hazard has been assessed to identify severity and probability of the hazard and assigned a Risk
Assessment Code (RAC). Definitions used for Severity and Probability in this assessment are as follows:
Initial RAC: The RAC is expressed as a combination of the severity and the probability classifications with the
severity classification listed first such as IA, IIA, IID, etc.
Residual RAC: Indicates the residual hazard probability and severity after all the identified controls are
incorporated.

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Status: Indicates whether the identified hazard is “Open” or “Closed.”
Open - Corrective action to eliminate or control the hazard has NOT been completed. The hazard will remain
open pending completion of corrective action and verification.
Closed - Corrective action to eliminate or control the hazard has been implemented. Closure of a hazard will
be based on Industrial Safety Office acceptance and approval.
Risk Assessment Classification (RAC): The RAC indicates the risk associated with each individual hazard. It
is derived by considering both the severity and probability of a hazard.
Hazard Severity: An assessment of the worst-case potential injury or system damage if an identified hazard
were to result in an accident defines the hazard severity. Severity categories are listed below.
Hazard Probability: The likelihood that an identified hazard will result in a mishap defines the hazard
probability.
Figure 10. RAC Definitions [1]

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Figure 11. Severity and Probability Definitions [1]

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The following table provides a summary of the identified hazards for the AF-M315E 1N Thruster test.
Conce
rn No.
Concern
Initial
RAC
Residua
l RAC
Residu
al Risk
Status
1 Rupture of High Pressure Tubing System 3C 3E Low Closed
2 Rupture of Low Pressure (<150 psig
MEOP) Tubing System
3C 3E Low Closed
3 Rupture of K-Bottle due to over-
pressurization results in damage to test facility
and/or injury to personnel
3C 3E Low Closed
4 Rupture of water coolant system due to
over pressurization of water
3C 3E Low Closed
5 Electrical shock due to exposure to water 1C 1E Low Closed
6 Loss of power causes unsafe conditions 3C 3E Low Closed
7 Unplanned ignition or detonation of
propellant
2C 2E Low Closed
8 Asphyxiation due to venting of GN2 in an
enclosed room results in death or injury to
personnel
1C 1E Low Closed
9 Slips, trips, and falls due to water spillage 3C 3E Low Closed
10 Injury to test personnel or the environment
due to exposure to propellant
3C 3E Low Closed
Table 2. Identified Hazards
IV. Conclusions and Future work
In conclusion of the project, a sophisticated propellant testing system was designed and built. Even
though the thrust stand and all necessary systems were complete, the system could not be operated for total
project completion. Due to time constraints of this internship, a hot test fire was not conducted for this
thruster. Some problems that came up during this project included the manufacturing of the radiation
shield, and safety requirements. Using a thinner shim to properly deform the spherical shape solved the
radiation shield problem. The safety requirements inhibited the process of the project. The teamwas unable
to test the completed testing system because the NASA safety team had to inspect and improve the lab
setup via a TRR. This issue was not resolved because of the internship time constraint.
In order to complete this work, the thrust stand will be tested and the propellant characterized. The
future work for this project will consist of the testing article listed in the section above. The first test will
consist of a steady state thrust test. After this test, catalyst will be inspected and results written. Once the
steady state data is acquired, the team will move on to pulse testing. Once testing data is acquired, the
propellant AF-M315E can be characterized.

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V. Acknowledgments
The author would like to thank the following people for their invaluable assistance with our project:
Chris Burnside
Kevin Pedersen
Tyrone
Daniel
Charles Pierce
Tony Yarbrough
John Wiley
Darrel Gaddy
Boyd McNutt
Brad Bullard
Erik Kraus
Jeremy Kenny
Luke Scharber
Logan Kennedy
Jonathon Jones
Texas Space Grant
Louisiana Space Grant
Ohio Space Grant
Alabama Space Grant
VI. References
1. MSFC NASA Safety Code, MWI 8715.15 Rev E