Development of an Electrothermal Propulsion System to Deorbit Small Satellites. Robert Spina

Transcription

1 Development of an Electrothermal Propulsion System to Deorbit Small Satellites by Robert Spina A thesis submitted in conformity with the requirements for the degree of Master of Applied Science University of Toronto Institute for Aerospace Studies University of Toronto c Copyright 2017 by Robert Spina

2 Abstract Development of an Electrothermal Propulsion System to Deorbit Small Satellites Robert Spina Master of Applied Science University of Toronto Institute for Aerospace Studies University of Toronto 2017 The DAUNTLESS bus propulsion system, using carbon dioxide as the propellant, provides the capability to deorbit a satellite at the end of its mission. Large satellites often contain propulsion systems but the size, mass, and power limitations of small satellites makes commercial propulsion units generally inaccessible, not to mention very expensive relative to microspace budgets. Using off-the-shelf components a complete propulsion system was designed and successfully integrated onto a Space Flight Laboratory satellite that will be launching in late The development and testing of an electrothermal thruster was performed laying the groundwork for the design of the propulsion system. To accommodate the requirements of the DAUNTLESS bus a feed system and fill module were designed and assembled to deliver propellant to the thruster. Thorough testing not only validated individual components but also the system as a whole, proving its suitability as an option for a small satellite deorbiting mechanism. ii

3 Acknowledgements I would like to thank everyone who has supported me throughout the past two years. Specifically I would like to thank Dr. Robert Zee who has given me the opportunity to be part of the incredible team at the Space Flight Laboratory and has been a remarkable role model for developing my own skills in the space industry. Thank you as well to Ben Risi and Vince who took me under their wings and provided valuable mentorship throughout the entire propulsion development program. Without your help and guidance none of this would have been possible, I owe many of the achievements of my work to the two of you. And to the many other colleagues and students who have been part of this journey and have helped along the way, inside and out of the lab. I would finally like to thank my friends and family, especially my parents, for the support they ve shown me, not only throughout this Masters program but throughout my entire life. I would not be where I am today without their constant encouragement for me to follow my dreams. iii

4 Contents Acknowledgements Table of Contents List of Tables List of Figures Acronyms iii iv vii viii xi 1 Introduction Motivation Space Debris Mitigation Propulsion Systems State of Work Scope I Electrothermal Thruster Development 8 2 Micropropulsion Background Thruster Overview Key Requirements Trade-off Analyses Propellant Selection Catalyst Selection Thruster Design Chemical Propulsion Design Theory Performance Nozzle Design Thrust Chamber Design Propellant Mass Thruster Sizing Resistojet Monopropellant iv

5 3.2.3 Expected Performance Thruster Testing Risk Mitigation Heater Lifetime Testing Catalyst Lifetime Analysis Catalyst Deactivation Investigation Catalyst Restart Test Catalyst Lifetime and Comparison Test Performance Characterization Qualification Testing Vibration Cold Preheat Cold Propellant Operation Thruster Evaluation and Conclusions II DAUNTLESS Bus Propulsion System Development 53 5 Preliminary Analysis Requirements Thruster Re-Evaluation Propellant Selection Thruster Sizing Modifications Propulsion System Design System Architecture Flight System Configuration Feed System Fill/Drain Module Tank Selection Flight Assembly Testing Solenoid Valve Testing Lifetime Test Thermal Test Acceptance Additional Valve Testing Performance Characterization System Qualification Tank Hydrostatic Test System-level Vibration Test Flight Assembly Leak Testing Subassembly Nitrogen Leak Testing v

6 7.4.2 Helium System-level Leak Testing Conclusions 96 Bibliography 98 vi

7 List of Tables 2.1 Resistojet and Monopropellant Thruster (RMT) Key Requirements Summary Propellant Trade-off Summary RMT Design Summary Heater Lifetime Test Summary Catalyst Lifetime Test Results Summary Sintering Temperatures for Rhodium and Iridium Catalysts Phases of Alumina Results from Energy Dispersive X-ray Analysis Catalyst Restart Test Results Performance Test Results Summary Accelerations from the DNEPR Launch Vehicle Sinusoidal Sweep Test Levels Random Vibration Test Levels DAUNTLESS Bus Key Propulsion Requirements Summary Updated Resistojet Sizing and Performance Nominal Regulator and Thrust Valve Operational Parameters Test Parameters for Regulation and Thrust Valves Valve Lifetime Test Results Summary Valve Acceptance Test Parameters Hydrostatic Pressure Test Results Flight CRT Performance Parameters and Results QM Propulsion System Vibration Test Levels QM Propulsion System First Natural Frequency (FNF) Nitrogen Subassembly Leak Test Results Helium Leak Testing Results vii

8 List of Figures 1.1 The Micro-Propulsion System butane cold-gas thruster [1] The Orbital Propulsion Center hydrazine system [2] ECAPS green monopropellant thruster [3] An overview of the CNAPS module and a view of CNAPS within CanX Storage pressure vs temperature of nitrous oxide for various storage densities Percentage of specific impulse of an infinite expansion nozzle at different expansion ratios for different propellants Specific impulse as a function of chamber temperature for different propellants Schematic of a simple thruster showing the nozzle, thrust chamber, and standoffs that interface with the mounting structure (from right to left) Specific impulse as a function of decomposition percentage for nitrous oxide at a chamber temperature of 800 C Resistojet implementation of the RMT. Radiation shield not shown Final monopropellant implementation of the RMT (Nitro-100) An example of the test results for the heater lifetime test showing the temperature cycles Time needed for exhaust temperature to reach 750 C once the flow had been started, throughout the catalyst s life Thermal sintering of catalyst: diffusion of surface atoms (A) and bulk atom movement (B) Fresh catalyst sample at 10000x magnification (not used) Catalyst sample from Nitro-100 performance test at 10000x magnification ( 1 hour usage) Catalyst sample from protoflight performance test at 10000x magnification ( 1 hour usage) Catalyst sample from Nitro-100 restart test at 10000x magnification (about 30 hours usage) Catalyst sample from prototype lifetime test at 10000x magnification (greater than 50 hours usage) Catalyst sample from Nitro-100 restart test at 25000x magnification showing the needlelike features RMT performance characterization test setup Vibration test fixture with thruster mounted and coordinate system definition Placement of nine accelerometers plus control on the nozzle, housing, and interface plate Monopropellant thruster mounted on the vibration slip table (x and y axes) Monopropellant thruster mounted on the vibration head expander (z axis) Sinusoidal burst profile showing amplitude vs frequency viii

9 g shock test waveform Accelerometer response for the 0.5 g low-level sine test in the x-axis Accelerometer response for the 0.5 g low-level sine test in the y-axis Accelerometer response for the 0.5 g low-level sine test in the z-axis Sinusoidal burst accelerometer response in the y-axis Sinusoidal sweep accelerometer response in the y-axis Random vibration accelerometer response in the y-axis Negative 50 g shock accelerometer response in the y-axis Test setup for the cold preheat thermal vacuum test Temperatures during cold preheat test Temperatures during baseline test at 23 C Temperatures during cold propellant test at -10 C The updated Carbon dioxide Resistojet Thruster (CRT), radiation shield not shown DAUNTLESS bus propulsion system block diagram showing the fill/drain module on the left side of the tank and the feed system on the right Propulsion system mounted on the DAUNTLESS bus internal panel showing the tank, feed system, and fill/drain module Feed system design showing the configuration of all components Prototype feed system showing the initial plenum configuration. It was used for proof of concept and some risk mitigation testing A schematic of the solenoid valve describing the various internal components Fill module design Over-pressure system. The secondary rupture disk is mounted to the external side panel of the satellite Custom linerless composite tank Final DAUNTLESS bus assembly with internals visible (upper) and with the side panels mounted (lower) Modified prototype feed system schematic with updated plenum to accommodate leak testing of both valves Valve lifetime test flow control setup. Blue arrows indicate the direction of propellant flow The regulation frequency as a function of the selected regulated pressure deadband, determined experimentally Pressure over time in the secondary volume for all leak tests Pressure over time in the tertiary volume for all leak tests. Values were corrected to account for the uncalibrated pressure sensor Due to continuous operation over several hours, frost began to form on the tube downstream of the regulator valve, where most of the expansion was occurring Pressure over time in the secondary volume for all leak tests after dry actuations Pressure over time in the tertiary volume for all leak tests after dry actuations. Values were corrected to account for the uncalibrated pressure sensor Valve acceptance test setup inside a thermal chamber ix

10 7.10 Unit-level valve vibration test setup CRT performance test setup in the vacuum chamber A preliminary pressure sensor calibration that identifies the non-linear section below 90 psi An example of the pressure regulation algorithm controlling the pressure of the plenum at a setpoint of 125 psi with two different deadbands Typical CRT performance test heater and exhaust temperature profiles outlining the preheat period (0-600 s) and the thrust period ( s) Typical CRT performance test thrust output QM propulsion system mounted to the vibration table in all three axes: y-axis (upper left), x-axis (upper right), z-axis (lower) Tank bracket screws that experienced a loss of pre-load. This occurred on both sides of the tank Degradation of the tank bracket against the vibration fixture plate Degradation of the end bosses of the tank and the shims against the tank bracket Random vibration profile: control input, no notching Random vibration profile: control input, with notching Nitrogen subassembly leak test setup Outside-in helium leak testing setup Inside-out helium leak testing setup. The test article was placed inside the vacuum chamber and pressurized with helium x

11 Acronyms Al 2 O 3 Aluminum Oxide (Alumina) CO 2 Carbon Dioxide I sp I ssp Specific Impulse System-Specific Impulse NO 2 Nitrogen Dioxide N 2 Nitrogen N 2 O Nitrous Oxide SF 6 Sulfur Hexafluoride ADN Ammonium DiNitrimide CanX Canadian Advanced Nanospace experiment CHT Cylindrical Hall Thruster CNAPS Canadian Nanosatellite Advanced Propulsion System CoM Center of Mass COTS Commercial Off-the-Shelf CRT Carbon dioxide Resistojet Thruster CSA Canadian Space Agency DAQ Data Acquisition DI De-Ionized DUT Device Under Test EDM Electric Discharge Machining EDX Energy-Dispersive X-ray EoL End of Life FNF First Natural Frequency xi

12 GEVS General Environmental Verification Specification HAN Hydroxylammonium Nitrate HTP High-Test Peroxide IADC Inter-Agency Space Debris Coordination Committee ID Inner Diameter IPA Iso-Propyl Alcohol LEO Low Earth Orbit LFFT Long Form Functional Test LLS Low-Level Sinusoidal LPI Liquid Penetrant Inspection MEOP Maximum Expected Operating Pressure NANOPS NANOsatellite Propulsion System NASA National Aeronautics and Space Administration OD Outer Diameter POS Pressure Over Seat PSD Power Spectral Density PSLV Polar Satellite Launch Vehicle PUS Pressure Under Seat QM Qualification Model RMT Resistojet and Monopropellant Thruster RTV Room-Temperature-Vulcanization SEM Scanning Electron Microscopy SFFT Short Form Functional Test SFL Space Flight Laboratory SNF Second Natural Frequency SSTL Surrey Satellite Technology Ltd. TRL Technology Readiness Level TVAC Thermal VACuum XRD X-Ray Powder Diffraction xii

13 Chapter 1 Introduction 1.1 Motivation Micro and nanosatellites have become increasingly capable of achieving objectives typical of much larger satellites, while being significantly less expensive and having much quicker turnaround times. The increasing popularity of small satellites means that more spacecraft are being launched, especially into Low Earth Orbit (LEO), but there is an increasing concern for orbital debris colliding with operational satellites [4] Space Debris Mitigation A theory proposed by a National Aeronautics and Space Administration (NASA) scientist, Donald Kessler, states that there is a limit to the amount of space debris in Earth s orbit that when exceeded will result in runaway collisions eventually creating a dense belt of orbital debris that would make it extremely risky to launch satellites into orbit. This would heavily impact life on Earth as there would no longer be cell phone reception, wireless internet, weather forecasts, and many other things provided by satellites that people take for granted in their daily lives [5]. The Inter-Agency Space Debris Coordination Committee (IADC) is an international effort to mitigate the risks of collisions and has developed standards for the proper disposal of satellites after the completion of their operations [6]. Specifically for Low Earth Orbit (LEO) satellites this includes the requirement to deorbit within 25 years after the mission concludes, reducing the amount of man-made debris in orbit and likewise the risk of collisions [4]. There are several technologies available to meet these deorbiting standards and they fall into one of two general categories: passive drag devices and active systems. The passive systems include ribbons, balloons, sails, or electrodynamic tethers and are an attractive option for small satellites since they often do not have the capacity for larger and more complex active systems. They all rely on the deployment of lightweight structures at the end of their mission to increase their effective surface area thereby causing more interactions with atmospheric molecules. The momentum imparted by the molecules acts to decrease the satellite s orbital altitude. The Space Flight Laboratory (SFL) has experience with deployable drag sail technology having demonstrated it on-orbit on the Canadian Advanced Nanospace experiment (CanX-7) spacecraft. Alternatively, electrodynamic tethers have the added bonus of creating electrodynamic drag as well as aerodynamic drag. A current is induced in the deployed conductive tether as it 1

14 Chapter 1. Introduction 2 moves within the Earth s magnetic field and the interaction between the induced electric field and the geomagnetic field exploits the Lorentz force to cause the satellite to decelerate [7]. Active systems can be found in the form of steered drag devices, which resemble the passive systems although with active attitude control directing the drag structures into the optimal orientation for molecular collisions. Propulsion systems are also effective deorbiting devices and are especially desirable when the primary satellite mission already requires such a system. By thrusting in the direction opposite to the orbital velocity, the satellite loses energy and descends in altitude much like it would with passive techniques. This adds complexity but is often required for orbital altitudes where atmospheric particles are negligible Propulsion Systems Spacecraft propulsion methods can be classified based on whether or not they consume a propellant. Reaction engines expel mass (i.e., propellant) through a nozzle in accordance with Newton s third law to generate an equal and opposite reaction force thereby propelling the satellite. On the other hand zero-propellant systems rely on interactions with gravitation fields, magnetic fields, electromagnetic waves, and solar radiation to impart a propelling force on the spacecraft. Reaction engines are by far more common and large satellites often have one or more of these systems on-board to achieve such tasks as: orbit change maneuvers, interplanetary transfers, formation flying operations, station-keeping, collision avoidance, attitude control, and deorbiting. Reaction engines come in many different forms and can be divided into a few categories based on how the propellant is used. The simplest form is the cold-gas thruster that ejects compressed fluid out of a nozzle. These are simple, reliable, inexpensive systems that are optimal for low thrust applications yet they typically have poor efficiency. Resistojet (hot-gas) thrusters are similar to their cold-gas counterparts except they use heat exchangers to increase the temperature of the propellant prior to ejection. This increases the energy of the exhaust and provides much better performance and efficiency at the expense of a small energy input into the heaters. Taking the resistojet concept a step further, additional energy can be released if a decomposable propellant is used in the presence of a preheated catalyst, providing a boost in overall efficiency. This type of system is known as a monopropellant thruster, of which hydrazine systems have been used on many satellites for decades with great success. However alternate non-toxic propellants are beginning to replace hydrazine, especially in the small satellite market. Bipropellant and solid propellant rockets use the chemical energy generated from combustion of the fuel to generate high temperature exhaust, although they can be dangerous to store and test if proper precautions are not taken. Another type of thruster generates thrust by accelerating ions or plasma using electrostatic or electromagnetic fields. They are very efficient in terms of propellant usage, often having times higher efficiency than chemical thrusters, but have the disadvantages of a very low thrust output and extremely high power consumption making them undesirable for many small satellite missions. Although it is quite common for larger satellites to have propulsion systems, their use on the micro and nanosatellite scales is not as popular due to size, mass, and power limitations. However with smaller satellites now being used for many different types of missions there is a necessity to incorporate propul-

15 Chapter 1. Introduction 3 sive capabilities. A number of companies are developing micropropulsion technology with some actually being operated on-orbit. Surrey Satellite Technology Ltd. (SSTL) had developed the Micro-Propulsion System (Figure 1.1) which was implemented on the SNAP-1 nanosatellite launched in MPS was a cold-gas system using butane as the propellant having a thrust of 46 mn and a specific impulse (I sp ) a measure of the thruster s efficiency of 43 s [1]. SSTL has also expanded their design into a butane resistojet system that has flown on several satellites over the last 12 years including Alsat-1, Bilsat, UK-DMC-1/2, and exactview-1 to name a few. It has exhibited a thrust of 50 mn and an I sp of about 80 s at 300 C, proving that the resistojet implementation has a performance advantage over cold-gas thrusters [8],[9]. Figure 1.1: The Micro-Propulsion System butane cold-gas thruster [1]. The Orbital Propulsion Center division of ArianeGroup has had great success with hydrazine monopropellant thrusters which have been used on many different satellites since 1997 and are still being used today. They offer a number of thrust levels however their 1N version (Figure 1.2) could fulfill small satellite requirements, having a thrust of 1 N and a specific impulse of 220 s [2]. Despite hydrazine exhibiting incredible performance as a monopropellant, it is quite toxic, which makes it inappropriate for small satellite developers due to the difficulty in safely handling, integrating, and testing the propulsion system. Thus there has been a recent push in the industry to develop monopropellant systems that use green propellants to address the safety and environmental concerns of hydrazine. ECAPS, a Swedish space company, has developed a proprietary liquid propellant blend known as LMP-103S Storable Monopropellant, which combines Ammonium DiNitrimide (ADN), water, and various fuels like alcohols, acetone, and ammonia. It has a combustion temperature of 1600 C requiring high temperature catalysts and materials; however this allows it to achieve much higher performance. The 1 N thrust level version (Figure 1.3) has a specific impulse of about 235 s and was successfully operated on-board the PRISMA satellite in 2010 [3],[10]. These micropropulsion systems are available to be purchased, albeit typically at very high costs; thus a desirable alternative is to develop a system in-house. The challenge however lies in adhering to the aforementioned development philosophy, by making use of more readily available Commercial Off-the-Shelf (COTS) components and relying more on rigorous testing to achieve confidence in the design. SFL has

16 Chapter 1. Introduction 4 Figure 1.2: The Orbital Propulsion Center hydrazine system [2]. been developing propulsion systems since 2008 with the launch of the CanX-2 mission that contained NANOPS (NANOsatellite Propulsion System), a highly successful cold-gas propulsion demonstration. The technology proven with CanX-2 was then carried over to achieve the formation flying objectives of the CanX-4/CanX-5 mission in Using a more complex propulsion system that was capable of providing a constant thrust throughout its lifetime, known as CNAPS (Canadian Nanosatellite Advanced Propulsion System) (Figure 1.4 [11]), precise formation between the two identical spacecraft was maintained autonomously. The successful demonstration of formation flying provides opportunities for future missions involving high resolution Earth observation and interferometric imaging. Further expanding the capabilities of SFL satellites, a more advanced propulsion system was desired primarily to meet station-keeping and deorbiting requirements [12]. The investigation into more efficient and higher thrust options was undertaken by simultaneously developing both an electric propulsion system, in the form of a Cylindrical Hall Thruster (CHT), and an electrothermal thruster, in the form of a resistojet with expandability into a monopropellant system. Owing to the high power requirements of the CHT, which can often require more complex systems in small satellites, the electrothermal thruster is currently the more appropriate option for this bus size. 1.2 State of Work At the time the author joined SFL the preliminary trade-off analyses of the electrothermal thruster were completed, including propellant and catalyst selection; as well as the initial thruster sizing required to

17 Chapter 1. Introduction 5 Figure 1.3: ECAPS green monopropellant thruster [3]. Figure 1.4: An overview of the CNAPS module and a view of CNAPS within CanX-4. meet the desired performance parameters. In addition, the resistojet configuration was proven successful through a variety of vacuum tests in a representative space environment. The expansion to a full monopropellant implementation was being investigated with the preliminary testing of a prototype model. Some issues with the design were encountered after observing lower than expected performance, which prompted a redesign.

18 Chapter 1. Introduction 6 At this point, the author was tasked with evaluating the second iteration thruster (known as the protoflight model) by conducting performance tests in both ambient and vacuum conditions, as well as lifetime tests on various thruster components. A third iteration on the design, intended to improve catalyst lifetime, was necessary and the author greatly contributed to achieving Technology Readiness Level (TRL) 6 of this thruster, known as the Nitro-100 system. A new satellite bus, DAUNTLESS, required propulsive capabilities for a new mission and the results from the electrothermal thruster testing were used to decide on an appropriate system for meeting the mission requirements. The author was then tasked with the development and testing of this propulsion system, and finally the integration of the system in the flight spacecraft. 1.3 Scope The main objective of this thesis is to demonstrate the development of a micropropulsion system for use on small satellites, by developing first an electrothermal thruster and second the associated modules required for use of the thruster in a satellite including the propellant feed system and the fill/drain module. Specifically it outlines the author s contributions to this objective including: Performance characterization testing of a nitrous oxide monopropellant thruster; Setup, execution, and analysis of risk mitigation testing for the thruster, including heater and catalyst lifetime tests; Setup, execution, and analysis for the catalyst deactivation investigation and the associated testing; Setup, execution, and analysis of qualification testing for the thruster, including unit level vibration, cold preheat, and cold propellant testing; Design and prototyping of a viable feed system for the DAUNTLESS bus propulsion system; System level assembly and integration of the propulsion system; Performance characterization incorporating both the feed system and thruster; Setup, execution, and analysis of risk mitigation tests, including valve testing and system-level leak testing; and Setup, execution, and analysis of qualification vibration testing for the propulsion system. The following sections describe the above contributions in detail while providing the relevant background information necessary for the reader s understanding. This thesis can be divided into two associated parts that ultimately describe the successful development of a micropropulsion system for a small satellite. The first deals with the design, development, and test campaign for an electrothermal thruster, specifically a resistojet thruster with expandability to a monopropellant configuration. Chapter 2 begins by outlining the development approach that would be followed to adapt the existing SFL cold-gas thrusters into higher performance, higher efficiency electrothermal thrusters. This chapter also discusses the requirements that will drive the design of the thruster and the trade-off analyses performed to select the propellant and catalyst. Chapter 3 discusses chemical propulsion design theory and how that defined the sizing of the thruster. Chapter 4 describes the testing required to prove the design of the thruster

19 Chapter 1. Introduction 7 successful. The second part of this thesis deals with the design, development, assembly, and testing of a complete carbon dioxide propulsion system that includes the previously developed thruster as well as the fill and feed systems necessary for operation in the DAUNTLESS bus. Chapter 5 outlines the key requirements for the specific mission and the subsequent modifications that needed to be made to the thruster and propellant choice. An overview of the system architecture and the configuration of the flight system is outlined in Chapter 6. Chapter 7 discusses the required testing, both on the unit-level and system-level, proving that the DAUNTLESS bus propulsion system will be a successful option for small satellites.

20 Part I Electrothermal Thruster Development 8

21 Chapter 2 Micropropulsion Background 2.1 Thruster Overview Using the microspace approach, a simple and inexpensive version of a propulsion system that would typically be found on large spacecraft could be achieved. The results from the development of this system including SFL s electrothermal thruster, known as the Resistojet and Monopropellant Thruster (RMT), would prove to be beneficial for future missions that require small propulsion systems with thrust levels below 1 N. The development approach began with lessons learned from SFL s highly successful cold-gas propulsion system CNAPS. This system used Sulfur Hexafluoride (SF 6 ) as propellant and was extensively operated on-orbit to achieve formation flying of the two satellites. By scaling up the CNAPS design to meet higher thrust level requirements as well as replacing SF 6 with Nitrous Oxide (N 2 O), a substantial increase in efficiency was expected. The first step was to operate the thruster in the resistojet configuration, where a heater core inside the thrust chamber was used to heat the propellant to about 500 C. At the expense of some power to operate the heaters a significant increase in performance over cold-gas systems would be possible. Upon verifying the performance of the resistojet thruster, the next step was to determine the viability of achieving monopropellant operation, which would replace the heater core with a catalyst bed. This would help initiate the decomposition of the nitrous oxide once the catalyst bed was pre-heated. It was expected that this would cause a performance increase as well as reduce the input power required to heat the propellant, as the decomposition would be self-sustaining once initiated. The performance in both configurations would be evaluated against the key requirements derived for a reference spacecraft mission. 2.2 Key Requirements There are several important requirements that define the performance capabilities of the RMT. They were used to perform trade-off analyses (as described in Section 2.3) as well as determine the appropriate sizing of the thruster (Section 3.2) and are summarized in Table

22 Chapter 2. Micropropulsion Background 10 Table 2.1: Resistojet and Monopropellant Thruster (RMT) Key Requirements Summary Number PROP-01 PROP-02 PROP-08 PROP-14 PROP-15 PROP-18 PROP-37 PROP-38 PROP-62 Requirement Description Minimum V of 100 m/s to the reference spacecraft Reference spacecraft dry mass of 150 kg Minimum impulse bit of 50 mns or less Less than 5 minute warmup time Maximum continuous power of 20 W Minimum life of 7 years Operate within specification over temperature range of 30 C to 60 C Survival temperature range of 40 C to 80 C Propellant shall not be toxic, flammable, corrosive, or damaging to the environment The minimum velocity change ( V) combined with the spacecraft reference mass derive the desired performance characteristics of the system, including I sp and thrust. The minimum impulse bit requirement allows for a trade-off between the desired thrust level and on-time of the thruster heaters, which translates to power consumption. Driving the design and feasible operating temperature are the warm-up time specification and the power limitations, which define the efficiency that heat must be transferred to the thrust chamber to obtain the specified temperature. The lifetime requirement defines the survivability of the electronics and the overall system leak rate, the latter of which must be evaluated over the entire mission duration and accounted for in total propellant calculations. The temperature limitations control the storage parameters of the propellant. For most self-pressurizing propellants at low temperatures nominal thrust levels are difficult to maintain due to the vapour pressure dropping so low. At high temperatures the vapour pressure increases significantly resulting in the necessity to use either larger tanks with decreased storage density or use higher pressure valves to distribute the propellant, which can often be difficult to source. The relationship between temperature and pressure of nitrous oxide at various storage densities is shown in Figure 2.1. It is evident that if one is trying to remain within a certain pressure range there is a trade-off between the volume (and thus the physical size) of the tank, which defines the storage density, and the allowable temperature fluctuations that can be accommodated. Often it is those pressure limitations that have a contribution to the derivation of the temperature range requirements. 2.3 Trade-off Analyses The main areas of variability in determining the operational and performance characteristics of an electrothermal thruster are the choice of propellant and, in the case of a monopropellant thruster, the catalyst material Propellant Selection A number of propellant options were examined to determine which one would be the most appropriate for the RMT. They were evaluated based on several criteria including: specific impulse, which describes the

23 Chapter 2. Micropropulsion Background 11 3,000 Pressure (psia) 2,500 2,000 1,500 1,000 1 L/kg 1.3 L/kg 2 L/kg 2.2 L/kg 3 L/kg 6 L/kg 15 L/kg 33 L/kg 60 L/kg Supercritical Temperature ( C) Figure 2.1: Storage pressure vs temperature of nitrous oxide for various storage densities efficiency of the thruster; system-specific impulse (I ssp ), which takes into account the overall mass of the system and provides a more appropriate comparison for different propellants; vapour pressure, defining the storage characteristics; space heritage, since proven technology can often be more reliable; and handling safety, since SFL team members will frequently be working with and around the propellant during testing. In addition, the viability of each propellant to be operated in the monopropellant configuration (i.e., the ability for the propellant to be decomposed) to allow for a more seamless transition from resistojet to monopropellant systems. The results from the trade-off analysis are summarized in Table 2.2. Table 2.2: Propellant Trade-off Summary Propellant System Type Isp (s) Issp (Ns/kg) Vapour Pressure (psi) Space Heritage Handling Safety SF6 Cold-Gas Yes Safe R-236fa Cold-Gas No Safe R-134a Hot-Gas No Safe Butane Hot-Gas Yes Moderate HAN Monoprop No Safe LMP-103S Monoprop low Yes Safe Hydrogen Peroxide Monoprop 150 N/A 0 Yes Dangerous Hydrazine Monoprop Yes Very Dangerous Although SF 6 has been used extensively at SFL, is relatively easy to obtain, safe to handle, and shows good performance in a cold-gas configuration, it has very low efficiency even when pre-heated making

24 Chapter 2. Micropropulsion Background 12 it unsuitable for use in the RMT. Freon R-236fa and R-134a are typically used as refrigerants and can be difficult to obtain since licensing is often required. Despite being safe to handle they have no space heritage and also exhibit low performance compared to other options. As a warm gas (resistojet) option, butane seems reasonable as it has space heritage, however its moderate performance and high flammability and toxicity make it less desirable. In terms of options that are capable of monopropellant operation, hydrazine has long been the most popular propellant. It has great performance and very low vapour pressure however it is extremely hazardous, which has driven researchers to develop more green replacements that aim to maintain or improve the performance over hydrazine while being much safer to handle. Two such options were examined, Hydroxylammonium Nitrate (HAN) and LMP-103S, which both have exceptional performance. The latter has even been used operationally on the PRISMA spacecraft. The downside to these propellants is that they are proprietary and difficult to obtain. Hydrogen peroxide, or High-Test Peroxide (HTP) as it is commonly known for propulsion applications, has good performance but is dangerous to handle, store, and transport and its instability results in auto-decomposition over time causing large increases in pressure. The final propellant examined and the one that was ultimately selected for the RMT was nitrous oxide (N 2 O). It was chosen due to its good performance, ease of procurement, and handling safety; the only disadvantages being the high vapour pressure, which may make selecting valves more difficult and the potential creation of the noxious gas, Nitrogen Dioxide (NO 2 ), that is an intermediate product in the decomposition. Although the creation of NO 2 can be easily alleviated by ensuring full decomposition as well as providing adequate venting during testing. Overall these factors and the ability for N 2 O to be used in both resistojet and monopropellant configurations made it the most desirable candidate for the RMT Catalyst Selection To initiate the decomposition of nitrous oxide it must be heated to above 520 C, which can be achieved using a heat exchanger inside of the thrust chamber. A constant power input is then required to maintain decomposition. By introducing a catalyst into the chamber the decomposition temperature decreases and self-sustained decomposition can be achieved without providing any further power, which is a very attractive quality for spacecraft where power consumption is typically limited. A desirable catalyst must minimize the decomposition temperature significantly, be resilient through several lifetimes of operation, be safe to handle, and easy to procure. Catalysts come in many forms including: pure metals, pure oxides, mixed oxides, supported systems, zeolites, and although not practical some gases can even act as catalysts. A lot of research has been done on metals like platinum, palladium, rhodium, and iridium as well as oxides of transition metals like rhodium, cobalt, iron, and nickel and their ability to decompose N 2 O [13]. It has been found that they can reduce the activation energy required for decomposition from about 250 kj/mol to between 80 and 170 kj/mol. From a practical standpoint, supported systems such as those using aluminum oxide (alumina), silicon oxide (silica), or zirconium oxide (zirconia) as the carrier for the above stated catalyst metals are more appropriate due to their higher dispersion over the larger specific surface area of the support [13]. Two catalyst types, one rhodium-based and one iridium-based were considered as viable options for the design with both consisting of an aluminum oxide substrate as a support for the catalyst material.

25 Chapter 2. Micropropulsion Background 13 Iridium catalysts have been used extensively for many years in a number of monopropellant systems. They have a higher sublimation temperature than rhodium, but they decompose nitrous oxide at about 450 C and are more expensive to acquire [14]. Rhodium on the other hand is much more inexpensive and has a lower activation temperature of about 250 C, which makes it a desirable alternative to iridium. However, it lacks in survivability over multiple restarts, a characteristic that was examined thoroughly through testing and the results are described in Section Rhodium was selected for initial testing but if determined insufficient, iridium can be investigated as an alternative.

26 Chapter 3 Thruster Design 3.1 Chemical Propulsion Design Theory Performance Resistojet and monopropellant thrusters operate under the same basic principles as any other chemical rocket, where a force is exerted on the vehicle that the rocket is part of by ejecting mass at a high speed. The force (F ), known as thrust, is given by: F = I sp ṁg 0 (3.1) where ṁ is the mass flow rate of the propellant, g 0 is the acceleration due to gravity at the Earth s surface, and I sp is the specific impulse, which is the most important performance parameter in sizing the thruster since increasing specific impulse means a higher thrust is achieved for a given mass flow rate. It is a function of nozzle geometry, propellant temperature, and chemical properties and describes the number of seconds that a unit of propellant can generate a unit of thrust and is often representative of the system s efficiency. It can be calculated using the following set of equations. I sp = C g 0 γ ( ) ( ) 1+γ ( 2 2 γ 1 1 p e γ γ p c γ 1 ) γ (3.2) where pe p c is the pressure ratio between the nozzle exit and the thrust chamber, γ is the ratio of specific heats of the propellant, and C is the characteristic exhaust velocity, given by: C = ( γ ( a ) 1+γ 2γ γ ) (3.3) where a is the speed of sound of the propellant in the thrust chamber, determined by: a = γrtc m M (3.4) where R is the universal gas constant, T c is the temperature of the propellant in the thrust chamber, and m M is the molecular mass of the propellant [15]. 14

27 Chapter 3. Thruster Design 15 The pressure ratio from equation 3.2 is a function of the nozzle expansion ratio through the exit Mach number (M e ) and the ratio of specific heats. It is also noteworthy that the ratio is independent of the chamber pressure since a change to chamber pressure causes a proportional change to the exit pressure assuming choked flow. It is defined as: ( p e = p c 1 + γ 1 Me 2 2 ) γ γ 1 (3.5) The Mach number can be found by iteratively solving the following equation for the nozzle expansion ratio (ɛ) [15]: ɛ = 1 ( ( (γ 1) M 2 )) γ+1 2γ 2 e M e γ (3.6) Nozzle Design The main parameter that will determine both the resulting performance characteristics of the nozzle as well as the geometry is the nozzle expansion ratio. Ideally the propellant exiting the nozzle expands to match the vacuum of space meaning the pressure ratio is zero and the expansion ratio is infinite. From a practical perspective it is not possible to achieve this expansion ratio due to size and mass limitations however in many cases it is possible to approach the performance of an ideal nozzle with a reasonable expansion ratio. This is dependent on the specific heat ratio of the propellant, where the higher it is the closer to the ideal specific impulse one can get as is seen in Figure 3.1. Xenon (γ = 1.666) can achieve 99% of the specific impulse of an infinite expansion nozzle with a ratio of 100, whereas SF 6 (γ = 1.056) can only achieve 54% of the ideal specific impulse with the same nozzle. Once a sufficient expansion ratio is selected and the I sp determined, a specific nozzle half-angle (α n ) can be selected initially. With this, the nozzle efficiency (η) can be calculated according to: η = 1 + cosα n 2 (3.7) After selecting the desired thrust level through compromise with thrust duration to meet the minimum impulse bit requirement the mass flow rate that is required to meet performance levels can be found by rearranging equation 3.1 while implementing the nozzle efficiency term: ṁ = F g 0 I sp η (3.8) The actual geometry of the nozzle is important when manufacturing one or purchasing a pre-existing one. The geometry can be defined by the nozzle throat area, exit diameter, and total length. The nozzle throat area can be found by rearranging an equation for characteristic exhaust velocity as follows: A t = ṁc p c (3.9) where p c is the chamber pressure. Having already determined a nozzle expansion ratio, the exit area

28 Chapter 3. Thruster Design % of Infinite Expansion Specific Impulse SF 6 Butane Xenon N 2 O Hot N 2 O Mono CO Expansion Ratio, ɛ (A e /A t ) Figure 3.1: Percentage of specific impulse of an infinite expansion nozzle at different expansion ratios for different propellants and subsequently the exit diameter, can be easily found by: A e = ɛa t (3.10) Although bell-shaped nozzles are better from a performance standpoint, conical nozzles are much simpler to manufacture and offer almost as good performance. Thus the nozzle length (L n ), which is measured from the throat to the exit, can be calculated using the previously determined geometry: L n = (d e d t ) 2 tan α n (3.11) where d e and d t are the exit and throat diameters respectively. The half-angle can then be iterated on using the above equations to establish a nozzle geometry that provides the performance necessary while also being practical to manufacture and implement into a spacecraft Thrust Chamber Design The thrust chamber is the section where the propellant is heated and/or decomposed to the appropriate temperature which is necessary to achieve the desired performance characteristics. In resistojet thrusters this takes the form of a heat exchanger, while in monopropellant systems it contains the catalyst bed as well as the heater that preheats the catalyst.

29 Chapter 3. Thruster Design Heat Exchanger It is evident that when nozzle geometry and chemical properties are fixed increasing specific impulse is only achieved by increasing the temperature of the propellant. This is done in electrothermal systems by adding a heat exchanger to cold gas thrusters. The increase in specific impulse that is caused by this increase in temperature for a variety of propellants is shown in Figure 3.2. Resistojet (and monopropellant) systems gain a performance increase for a small increase in complexity and are often ideal in situations where energy is more plentiful than mass Specific Impulse (s) ,000 1,100 1,200 1,300 1,400 1,500 1,600 Chamber Temperature ( C) SF 6 Butane Xenon N 2 O Ammonia CO 2 R-134a Figure 3.2: Specific impulse as a function of chamber temperature for different propellants The heat exchanger has the task of adding energy to the propellant in the form of heat, which subsequently increases the specific impulse (or efficiency) of the thruster. The geometry of the nozzle (i.e., expansion ratio) defines the exit Mach number, however the actual flow velocity increases with exhaust temperature. Thus, the more heat that is added to the propellant the better the performance. This is limited practically by material limitations and available power which requires a trade-off between performance and a design that is reasonable to build and operate. The amount of power ( q) that must be imparted to the fluid to raise its stagnation temperature from its initial temperature (T 0i ) to the desired final temperature (T 0f ) can be found using: q = c p ṁ(t 0f T 0i ) (3.12)

30 Chapter 3. Thruster Design 18 where c p is the isobaric specific heat capacity of the fluid. Equation 3.12 represents the amount of power that is needed to increase the energy of the fluid but it does not necessarily represent the actual amount of energy that the heater must output since there will be losses due to conduction and radiation. Figure 3.3: Schematic of a simple thruster showing the nozzle, thrust chamber, and standoffs that interface with the mounting structure (from right to left). Assuming steady-state conditions and a simple thruster design, as seen in Figure 3.3, conduction occurs through the standoffs that connect the thruster to the mounting point according to: q cond = ka l (T h T c ) (3.13) where k is the thermal conductivity of the standoff material, A is the effective cross-section of the standoffs, l is the length of the standoff, T h is the temperature of the hot body (i.e., thruster or heat exchanger), and T c is the temperature of the cold body (i.e., the mounting surface). It is apparent why the standoff material must be carefully selected to be able to handle the high temperatures of the thruster, while also prohibiting conduction as much as possible to prevent heat losses. In addition, there will be radiative losses from the body of the thruster to space. This is mitigated by implementing a radiation shield that almost completely encompasses, yet is thermally isolated from, the thruster to absorb some of the radiation. Despite this, both the exposed nozzle and the radiation shield will still lose some heat to space based on the following generic relation for radiation: q rad = σɛat 4 (3.14) where σ us the Stefan-Boltzmann constant representing the power per unit area that a black-body emits as a function of temperature and is equal to m 2 K. ɛ is the emissivity of the body 4 representing how close it is to acting like a black-body, A is the surface area of the body, and T is the body s temperature. By ensuring that the emissivity of a particular surface is low by using surface coatings or certain materials one can limit the amount of radiation that is lost. Equation 3.14 can be used to determine the amount of radiation lost through the exposed portion of the nozzle as well as the radiation shield, however the temperature of the radiation shield is still unknown. For that the following equation describing the radiation transfer between two objects (i.e., the outer surface of the W

31 Chapter 3. Thruster Design 19 thruster and the inner surface of the radiation shield) can be used: σ(t1 4 T2 4 ) q rad = 1 ɛ 1 A 1ɛ (3.15) A 1F ɛ2 A 2ɛ 2 where F 12 is the view factor from body 1 to body 2 (assumed to be equal to 1 for the thruster/radiation shield scenario, since they are treated as concentric cylinders). In addition to radiation heat transfer, there will also be some conductive heat transfer through the standoffs that isolate the radiation shield from the thruster, as described by equation By summing the two equations that describe how heat is transferred from the thruster to the radiation shield (conduction from 3.13 and radiation from 3.15) and then equating the total to equation 3.14, the temperature of the radiation shield (T 2 ) can be determined. This will then allow one to determine the total losses that will be experienced. L total = q condstandoff + q radnozzle + q radshield (3.16) Adding the total losses to the required power allows one to determine how much power the heater in the heat exchanger needs to provide to raise the propellant temperature to the desired amount Advantage of Decomposition If the propellant can be decomposed exothermically an additional performance increase can be achieved. Using heaters to preheat a catalyst bed, a small amount of energy can be input releasing the stored chemical energy in the propellant and starting the decomposition reaction. In the case of nitrous oxide the reaction is defined as: N 2 O N O 2 82 kj mol (3.17) where one mole of nitrous oxide decomposes forming one mole of nitrogen and half a mole of oxygen. The negative sign on the energy term indicates an exothermic reaction, which allows the reaction to be self-sustaining where subsequent propellant is decomposed with a reduced energy input. Another benefit is the fact that the products have a smaller molecular (molar) mass than the reactant, which results in a higher specific impulse. Nitrous oxide has a molar mass of g/mol, whereas the products, nitrogen and oxygen, have molar masses of g/mol and 32.0 g/mol respectively. Two-thirds of the produced molecules are nitrogen, while one-third is oxygen, resulting in the overall molar mass of the products to be g/mol. The same relationship applies to the specific heat ratio of the products that leave the rocket s nozzle, where the higher the specific heat ratio of the exhaust the higher the specific impulse. Since the overall effect of decomposition is to increase system performance over resistojet or cold gas systems it is necessary to achieve as close to complete decomposition as possible. The relationship between I sp and the percent decomposition for nitrous oxide at a constant temperature is seen in Figure 3.4. In the presence of a catalyst the decomposition can occur at much lower temperatures and the amount of the propellant that ends up getting decomposed is generally a result of the catalyst bed design. The catalyst comes in the form of small pellets about 1-2 mm in diameter and are contained in the thrust chamber in between the propellant input and the nozzle throat. The catalyst loading is a function of the length and diameter of the thrust chamber, the total mass of catalyst, and the size of each catalyst pellet,

32 Chapter 3. Thruster Design Specific Impulse (s) Percent Decomposition Figure 3.4: Specific impulse as a function of decomposition percentage for nitrous oxide at a chamber temperature of 800 C which all must be carefully selected to achieve the optimal temperature of propellant at the throat. Too little catalyst and the propellant will pass through the pack without getting completely decomposed. Too much catalyst and the temperature of the propellant will begin to decrease further down the chamber due to increased thermal mass. In both cases, the optimal temperature will not have been achieved at the end of the chamber. Through vigorous testing Zakirov determined that the optimal chamber length was between 40 and 50 mm [14]. In addition, how tightly the catalyst is packed will determine if a specific mass flow rate of propellant is able to pass through the catalyst pack and is represented by the loading factor (F L ): F L = ṁ A = ρu (3.18) where ṁ is the mass flow rate of the propellant through the catalyst bed and A is the cross-sectional area of the thrust chamber. Alternatively, ρ is the fluid density and u is the propellant velocity. Zakirov determined that loading factors between kg sm 2 were deemed to be reasonably sufficient [14] Propellant Mass Knowing the specific impulse as well as the V that the system must impart on the spacecraft allows for the total propellant mass (m p ) to be calculated. implemented: m p = m dry ( e V g 0 Isp 1 ) For this Tsiolkovsky s rocket equation can be (3.19) where m dry is the dry mass of the spacecraft (i.e., mass without any propellant) [15]. Due to the exponential nature of the equation it is evident that small changes in the specific impulse can have large effects on the propellant mass. Although overall the lower the I sp is, the greater the effect on propellant mass a small change makes. This is sometimes referred to as the tyranny of the rocket equation since a portion of the fuel carried is only used to move the rest of the fuel. If the fuel is inefficient (i.e., low specific impulse) than more fuel is needed for the vehicle mass, and even more is needed to move the extra fuel.

33 Chapter 3. Thruster Design 21 As noted in Section 2.2, the dimensions of the tank needed to store the propellant can be found using the maximum operating temperature and pressure to determine the minimum storage density (ρ min ). Simple geometry can then be used to calculate the dimensions of a spherical tank needed to contain the propellant according to: ( 6mp d tank = πρ min ) 1 3 (3.20) 3.2 Thruster Sizing Using the equations described in Section 3.1 the actual design of the electrothermal thruster can be determined, including the design iterations that were made based on the results from several tests that were performed (described in detail in Chapter 4) Resistojet The resistojet thruster, as mentioned earlier, is a higher performance extension of the simple cold-gas thrusters (like those in CNAPS). By heating the propellant in a heat exchanger prior to expulsion through the nozzle the efficiency of the thruster can be significantly increased. The design of the resistojet thruster, seen in Figure 3.5, is made up of a feed line to deliver the propellant to the thruster, the heat exchanger to increase the propellant temperature, the nozzle to accelerate the propellant, and the radiation shield to limit thermal losses from the heater. Also, thermocouples are included inside the thruster to provide temperature feedback at various points (mainly just before the throat). The thruster is mounted to a base that in turn can be mounted to the spacecraft in any orientation to achieve a desired thrust vector, while also providing thermal decoupling to the rest of the spacecraft. The feed line is a stainless steel tube that is welded into the base of the heat exchanger to provide a leak tight seal, while connecting to the rest of the feed system (described further in Part II) using a standard nut-and-ferrule compression fitting. This feed tube can also be easily bent into any orientation to accommodate various applications. The heat exchanger is made up of a stainless steel chamber with a central heater core running axially down the chamber. The core is a resistive heater element encapsulated in a ceramic powder all contained within a stainless steel sheath for protection and is often termed a cartridge heater. In addition the propellant is forced around a series of baffles to prolong the duration that it is in contact with the central heater core, maximizing heat transfer. Referring back to Figure 3.2 to maximize specific impulse, while staying within material and power limitations, a chamber temperature of 500 C is desired. Thus, the dimensions of the heat exchanger were derived from the heat flux required to increase the propellant temperature to this desired level; also taken into consideration were the dimensions that would be required for the eventual expansion to a monopropellant system. The nozzle has a converging-diverging conical shape machined out of stainless steel by way of electric discharge machining (EDM) and is welded to the end of the heat exchanger. Although for prototyping and developmental testing purposes, the initial design was to have the nozzle thread into the chamber and sealed with a high temperature, vacuum rated, non cementing thread sealant, allowing for internal

34 Chapter 3. Thruster Design 22 inspection in between tests. The nozzle was designed based on a nominal thrust level of 100 mn. The propellant is accelerated in the converging section until reaching Mach 1 (i.e., choked flow) at the throat, where it is then accelerated supersonically through the diverging section, generating the majority of the thrust. Finally, the above mentioned components are covered by an aluminum radiation shield to reduce the radiative losses that the hot components would emit to deep space. This can be further reduced by implementing a low-emissivity coating. Figure 3.5: Resistojet implementation of the RMT. Radiation shield not shown Monopropellant By replacing the heat exchanger baffles with a catalyst bed, the resistojet thruster can be easily converted into a monopropellant system, with all other components unchanged. However, due to the slightly different operating conditions that the monopropellant thruster would experience some fine-tuning of the geometry can be done to optimize the design. With a higher operating temperature of C, the thrust chamber remains the same, but if necessary the length can be modified to accommodate a specific amount of catalyst with the desired loading characteristics. As seen in Figure 3.1, for a monopropellant system using N 2 O an equivalent percentage of the infinite expansion specific impulse can be achieved at a much lower expansion ratio than for a resistojet (hot-gas) system. Thus a lower expansion ratio was selected and by maintaining the overall length of the nozzle the efficiency can be increased by reducing the half-angle. Through a variety of tests described thoroughly in Chapter 4, the thruster parameters were iterated on to optimize the design Mark I (Prototype) The first iteration of the design, known as the Prototype model, used the same dimensions as the resistojet thruster, albeit with the nozzle modifications as described above. It contained 7.09 g of

35 Chapter 3. Thruster Design 23 catalyst pellets, which were 0.5% rhodium (by weight) on 1 mm alumina (Al 2 O 3 ) spheres. However after performing a catalyst lifetime test they were determined insufficient and were replaced with 10% rhodium (by weight) on 1.5 mm alumina spheres. The cartridge heater used in the resistojet heat exchanger was replaced by a thin wire resistive heating element in a helical configuration that was embedded in the catalyst pack. Despite seeing good lifetime of the catalyst over several restarts, the performance was not sufficient and it was suspected that there was too much catalyst, causing the exhaust temperature at the throat to be lower than the optimal value (as explained in ). This was addressed in the next iteration Mark II (Protoflight) The protoflight model was the second iteration primarily aiming to optimize the length of the thrust chamber leading to a much shorter design. In addition, the cartridge heater was re-introduced and with these two changes the catalyst mass that was able to be packed into the chamber became 3.32 g. After performing vacuum tests on this new design it was determined that despite optimizing the length of the thrust chamber from a performance standpoint, the reduction in catalyst created a lifetime issue. After the first test, the thruster was not able to re-initiate decomposition under the same preheat conditions. This was thought to be an indication that too much of the catalyst was becoming inactive, either from temperature effects or physical damage Mark III (Nitro-100) The third iteration, the Nitro-100, was designed to address the catalyst lifetime issue and had an increased length over the protoflight model to increase the catalyst capacity and is shown in Figure 3.6. A second cartridge heater was added as redundancy (although only one would be operated at a time) and the temperature feedback from the heaters was improved by adding a thermocouple directly in contact with the heater sheaths. This would help to optimize the heat supplied to the catalyst bed. In addition a diffuser plate was added at the beginning of the chamber to help distribute the flow evenly throughout the catalyst pack in hopes of preventing localized deactivation of the catalyst along the nominal flow path. This redesign of the chamber resulted in a total catalyst pack mass of 6.58 g. Furthermore to reduce thermal degradation of the catalyst, the outer surface of the thruster was sandblasted to increase the emissivity and reduce the overall operating temperature to about 700 C. It was deemed necessary to accept the potential sacrifice in performance at this lower temperature to improve the lifetime of the catalyst Expected Performance Using the design parameters described above for both the resistojet thruster and the Nitro-100 monopropellant thruster, and the performance equations from Section 3.1.1, the expected specific impulse and mass flow rates can be determined. These are summarized in Table 3.1. It can be seen that the monopropellant implementation will have an increase in specific impulse of over 22 s (about 21%), and along with the reduced power input throughout operation it is obvious why this type of thruster is more desirable than the resistojet. This increase in I sp can be attributed to the fact that the decomposed exhaust products have a lower molar mass and a higher ratio of specific heat,

36 Chapter 3. Thruster Design 24 Figure 3.6: Final monopropellant implementation of the RMT (Nitro-100). Table 3.1: RMT Design Summary Resistojet Monopropellant (Nitro-100) System Performance Requirements Spacecraft Dry Mass (kg) 150 Total V (m/s) 100 Minimum Impulse Bit (mns) 50 System Design Parameters Thrust (mn) 100 Chamber Temperature ( C) Nozzle Expansion Ratio Nozzle Half-Angle ( ) Propellant Chemical Properties Specific Heat Ratio - N 2 O (@ 700 C) Specific Heat Ratio - N 2 (@ 700 C) Specific Heat Ratio - O 2 (@ 700 C) Molar Mass - N 2 O (g/mol) Molar Mass - N 2 (g/mol) Molar Mass - O 2 (g/mol) 32.0 Performance Outputs Expected Specific Impulse (s) Minimum Thrust Duration (s) Mass Flow Rate (mg/s) Total Propellant Mass (kg) further emphasizing the importance of complete decomposition of the propellant and adequate catalyst lifetime.

37 Chapter 4 Thruster Testing A number of tests were required to ensure that the thruster was operating at the necessary performance levels and would survive throughout its entire mission duration. In addition, the system must be able to withstand the intense loads of launch and the space environment, including vacuum conditions and a wide range of temperatures. To ensure that the thruster met all of these objectives a series of tests were performed. On a component level, lifetime testing was necessary to retire any risks of using those components since in most cases they were not necessarily designed to be used in space. Performance characterization tests in representative conditions were then preformed to verify the expected results. To qualify the thruster for use in real missions, rigorous vibration and thermal tests were performed to ensure survivability. The following sections outline the goals, methodology, and results for each of the tests. 4.1 Risk Mitigation Given that certain components have never been used in previous designs it is often desirable to perform extensive testing on the individual units to rule out any perceived risks that could cause problems in the later stages of the thruster development. This is especially necessary for components that do not have space heritage and more importantly were not designed specifically for use in space. For the RMT, both the cartridge heater and the catalyst present a concern for whether or not they will survive the cycling that would be involved with a lifetime s worth of operation Heater Lifetime Testing The custom cartridge heaters used in both the resistojet and monopropellant implementations will experience very wide temperature extremes, potentially getting as cold as 3 K (-270 C) when inactive and as hot as 1100 C during sustained operation. These large, and often abrupt, changes in temperature can put a lot of strain on the materials within the heater. Based on the system performance requirements such as the dry mass and total V values (outlined in Table 3.1), the total number of cycles that the heater would experience (i.e., total number of thrusts required to achieve the V) was calculated to be 1784 cycles. This corresponds to one lifetime of the cartridge heater for the purposes of this test. 25

38 Chapter 4. Thruster Testing Test Setup Three cartridge heaters with rated operating temperature limits of 760 C were tested in this experiment, each with a K-type thermocouple secured to the outer sheath using a ceramic adhesive to ensure sufficient contact even under extreme thermal loads. They were all placed in a thermal chamber with the temperature set to -73 C (thermal chamber s lowest possible setting), which was necessary to cool the heaters down in a reasonable amount of time to facilitate more frequent cycling. The thermocouples were connected to a data acquisition (DAQ) device for logging and were ultimately used as the feedback to determine when the heaters would be turned on or off. Each of the heater s power leads were connected to their own dedicated power supply, which were in communication with custom Matlab and Python scripts that would control the supplies on and off at the necessary voltage levels Methodology The following procedure was implemented to examine the cycle lifetime of each of the three heaters: 1. Initially the heaters were allowed to cool to the chamber set point of -73 C. 2. The previously developed Matlab/Python algorithm cycled all three heaters on and off simultaneously between -30 C and 600 C at 28V and 1.2A using a bang-bang control scheme. 3. This was repeated until failure of each heater. An example of the test methodology can be seen in Figure Temperature ( C) Time (s) Figure 4.1: An example of the test results for the heater lifetime test showing the temperature cycles Results The results from the test confirmed that all heaters successfully completed at least two lifetimes worth of cycles. Table 4.1 summarizes these results. Heater 3 had a premature failure after its thermocouple detached. The heater no longer had temperature feedback to tell the script to turn it off and so remained to be powered until failure. Despite this, it still achieved over two lifetimes of cycles and the lowest natural failure occurred beyond four lifetimes worth of cycles. Heater 2 even lasted well past six lifetimes, thus

39 Chapter 4. Thruster Testing 27 it is safe to say that these custom cartridge heaters will likely be able to withstand the required cycles necessary for the reference mission. Table 4.1: Heater Lifetime Test Summary Heater Achieved Cycles Number of Lifetimes Notes Heater burned out Premature failure due to test apparatus fault Heater burned out Catalyst Lifetime Analysis Proof of concept tests proved that the catalyst would perform well for short periods of time; however there was some uncertainty in its long duration performance. Zakirov obtained similar results and indicated that identifying a suitable catalyst is often the most difficult aspect of monopropellant design [14]. Thus, four main tests were completed to better understand how the selected catalyst would perform under the expected conditions. First, using the prototype model thruster a dedicated lifetime test was performed. Despite observing promising results, future lifetime issues with the Mark II and III models of the thruster provoked an additional three tests that aimed to identify possible sources of catalyst deactivation and techniques that could be employed to mitigate them Test Setup The prototype thruster (with the internal helical resistance wire heater) was filled with a fresh batch of catalyst and the nozzle was threaded and sealed onto the chamber. Five external K-type thermocouples were attached at various positions along the exterior of the thrust chamber. A tank of medical grade nitrous oxide was used as the propellant supply with a pressure regulator and gauge being used for pressure control. Just upstream of the thruster was a flow meter which was used as feedback to adjust the regulator to the desired flow rate. All of the thermocouples were connected to the DAQ to log temperature data. Based on the spacecraft dry mass, V requirement, and the thruster performance (I sp of s and mass flow rate of 65 mg/s) the operational lifetime of the thruster was determined to be 43.1 hours with 10.7 kg of N 2 O propellant being used. These are the parameters that the results were compared against Methodology 1. Using the internal resistance wire heater and the internal thermocouple as feedback, the catalyst pellets and chamber were preheated to between 300 C and 400 C. 2. The pressure regulator was opened and adjusted to achieve the desired mass flow rate of 65 mg/s into the thrust chamber. 3. Once decomposition was initiated, power to the heater was shut off. 4. The system was allowed to reach steady-state and then was run continually for 8-hour periods.

40 Chapter 4. Thruster Testing The test was paused overnight by shutting off the flow to the thruster and the system was allowed to cool back to room temperature. 6. Testing was repeated daily until failure or at least two lifetimes were achieved Results The lifetime test was composed of 8 sub-tests and the important results are summarized in Table 4.2. Over the course of the test campaign it was evident that the time needed to reach steady-state temperature was increasing with catalyst age as shown in Figure 4.2. Test Day Preheat Temperature ( C) Table 4.2: Catalyst Lifetime Test Results Summary Time to Peak Time to Reach Reach Temperature Peak 750 C (s) ( C) Temperature (s) Steady-State Temperature ( C) Hot Flow Duration (s) Mean Total Time Tested (hours) 50.4 Time to Reach 750 C (s) Cumulative Test Time (s) Figure 4.2: Time needed for exhaust temperature to reach 750 C once the flow had been started, throughout the catalyst s life. The thruster was operated for a total of 50.3 hours and the results show that the catalyst sustains decomposition beyond the required lifetime of 43.1 hours. There is however evidence of degradation over time starting at around 20 hours, which is displayed as an increased amount of time to achieve 750 C. In the final test (8) it was shown that the increased heat up time can be mitigated by increasing the preheat temperature from 350 C to 500 C, which is why that data point seems to be an outlier. Based

41 Chapter 4. Thruster Testing 29 on these results it seems that the selected catalyst is sufficient for at least one lifetime, however after iterating the thruster design to Mark II and Mark III, performance testing in vacuum conditions brought to light further issues with continual restarts. Thus, additional analysis and testing was required, which will be discussed in detail in Sections 4.1.3, 4.1.4, and Catalyst Deactivation Investigation To better understand what could be done to improve the lifetime of the catalyst it was first necessary to investigate how decomposition works on a molecular level and what mechanisms could be causing the deactivation. Then using Scanning Electron Microscopy (SEM) techniques and Energy-Dispersive X-ray (EDX) analyses the actual rhodium content on catalyst samples of varying degrees of wear were examined. Simultaneously a series of tests were performed with the Nitro-100 model thruster to see if changing the preheat conditions would initiate decomposition consistently throughout the catalysts life. As a last resort, investigations into alternate catalyst materials would be evaluated Decomposition of Nitrous Oxide Nitrous oxide undergoes decomposition when it is heated above 520 C resulting in products that are akin to air (N 2 and O 2 ). Homogeneous (thermal) decomposition occurs at appreciable rates in the temperature range of about C and is a unimolecular process involving fission of the weaker N-O bond as described in Section Homogeneous decomposition requires a constant heat input to maintain the temperatures necessary for decomposition, while on the other hand, heterogeneous (catalytic) decomposition can occur with significantly less heat input as long as it is in the presence of a suitable catalyst. In general heterogeneous decomposition is initiated when a nitrous oxide molecule reacts with an active site on the catalyst surface resulting in a nitrogen molecule and an adsorbed oxygen atom, as in: N 2 O + () N 2 + (O) (4.1) where () denotes an active site on the catalyst. From here there are two mechanisms at work: the Langmuir-Hinshelwood mechanism and the Eley-Rideal mechanism [16] and [17]. The former occurs when atomic oxygen migrates over the catalyst surface, recombines, and desorbs to form an oxygen molecule and two active sites as in: 2 (O) O () (4.2) The reversibility of this reaction allows for the possibility of inhibitor molecules in the gas feed such as oxygen (O 2 ), nitrogen monoxide (NO), and water (H 2 O) to occupy the active sites and prevent the nitrous oxide molecules from using them for decomposition. The second mechanism occurs when a nitrous oxide molecule comes into contact with an adsorbed oxygen on the catalyst surface forming nitrogen and oxygen molecules according to: N 2 O + (O) N 2 + O 2 + () (4.3) The nitrogen and oxygen escape as products while the active site provides another location for decomposition of additional nitrous oxide molecules.

42 Chapter 4. Thruster Testing Mechanisms of Deactivation The mechanisms for deactivation can be grouped into three categories: chemical, mechanical, and thermal. Each has a possibility for occurring in the monopropellant thruster due to potentially impure propellant supplies, high loads, and high temperatures. From a chemical standpoint deactivation could be caused by poisoning, vapour-solid reactions, or solidsolid reactions. Poisoning is the strong chemisorption of reactants, products, or impurities on sites otherwise available for catalysis. This includes the inhibitors mentioned above and depends upon the species adsorption strength relative to the other species it is competing against for the active sites [18]. Vapour-solid reactions can occur between the vapour phase of the fluid with materials on the catalyst producing inactive bulk or surface phases, including the formation of oxides of the catalyst material like rhodium oxide. Solid-solid reactions can also occur between multiple catalyst materials and/or support materials such as the formation of rhodium aluminate (RhAl 2 O 4 ) when rhodium is used on alumina supports. Mechanical deactivation mechanisms like fouling, crushing, or erosion can also occur. Fouling is similar to poisoning in that it is a physical deposition of a foreign species from the fluid phase onto the catalyst surface. The difference is that the particles are not chemically absorbed onto an active site, they are simply blocking the propellant from interacting with the catalyst material. Crushing under loads and erosion can also cause deactivation by destroying the catalyst pellet and reducing the overall size available for decomposition. These mechanisms, although possible, are unlikely under the nominal operating conditions of the monopropellant thruster. Thermal sintering is suspected to be the most likely mechanism for deactivation, involving the loss of catalytic surface area due to crystallite growth and/or loss of support surface area due to support collapse. Sintering of the catalyst generally occurs at higher temperatures (>500 C) and increases exponentially with temperature. When the Huttig temperature (30% of the melting point (T mp )) is reached less strongly bound surface atoms at defect sites like edges or corners dissociate and diffuse readily over the surface. At the Tamman temperature (50% of T mp ) atoms in bulk become mobile and move across the surface (Figure 4.3). Table 4.3 shows the Huttig and Tamman temperatures for rhodium and iridium. Table 4.3: Sintering Temperatures for Rhodium and Iridium Catalysts Metal Melting Huttig Tamman Temperature (K) Temperature (K) Temperature (K) Rhodium Iridium Sintering of the carrier can also occur, where single-phase oxide carriers (like alumina) can undergo surface diffusion, solid-state diffusion, evaporation of volatiles, grain boundary diffusion, or phase transformation. Alumina has a particularly rich phase behaviour as a function of temperature and preparation, where certain phases have different surface areas for holding catalyst materials as seen in Table 4.4 [18].

43 Chapter 4. Thruster Testing 31 Figure 4.3: Thermal sintering of catalyst: diffusion of surface atoms (A) and bulk atom movement (B). Table 4.4: Phases of Alumina Alumina Phase Crystal Structure Surface Area (m 2 /g) (BET Analysis) Transition Temperature ( C) Boehmite Orthorhombic γ-alumina Cubic Defective Spinel δ-alumina Orthorhombic θ-alumina Deformed Monoclinic Spinel α-alumina Hexagonal Close Pack (ABAB Stacking) The monopropellant thruster was originally designed to operate at temperatures between C, which is above the Tamman temperature of rhodium, it is very possible that there is bulk movement of rhodium atoms reducing the effective surface area available for decomposition. In addition, since the selected catalyst support was γ-alumina and the operating temperature is nearing the transition temperature of θ-alumina, there may be a reduction of 75% of the surface area. In some areas of the catalyst pack, especially close to the heater, temperatures could be even hotter possibly causing a transition to α-alumina and a reduction of 200 times the surface area. The α-alumina phase (also known as corundum) is the most thermodynamically stable phase of alumina [19], which could give further reason to believe that the crystal structure of the substrate had transitioned to this phase. It was this suspicion that prompted the modification of the outer surface emissivity for the Nitro-100 model in an attempt to reduce the operating temperature to 700 C. If achieved it would prevent phase transition of the γ-alumina and greatly reduce the thermal sintering of the rhodium Scanning Electron Microscopy Analysis Catalyst samples from various tests were examined against a batch of fresh catalyst pellets to see how they compared in rhodium content. Using a focused beam of high-energy electrons a variety of signals at the surface of the test specimen can be generated deriving from the electron-sample interactions. These signals reveal information about the specimen including texture, chemical composition, and crystalline structure [20] and were specifically used to image the catalyst at 10000x magnification in this test.

44 Chapter 4. Thruster Testing 32 The following five figures (4.4, 4.5, 4.6, 4.7, and 4.8) show images taken through the scanning electron microscope at 10000x magnification. They are presented in increasing levels of wear with the first one being an unused batch of catalyst, the second two considered having medium levels of wear (about 1 hour of operation), and the final two considered to have high levels of wear (greater than 30 hours operation). Figure 4.4: Fresh catalyst sample at 10000x magnification (not used). Figure 4.5: Catalyst sample from Nitro-100 performance test at 10000x magnification ( 1 hour usage).

45 Chapter 4. Thruster Testing 33 Figure 4.6: Catalyst sample from protoflight performance test at 10000x magnification ( 1 hour usage). Figure 4.7: Catalyst sample from Nitro-100 restart test at 10000x magnification (about 30 hours usage). Figure 4.4 shows a fresh batch of catalyst where the alumina substrate can be clearly identified as having large crevices (that in general signify the surface area); as well as many rhodium particles in the form of light-coloured, flaky specs. This is to be the visual baseline for which the other samples will be compared against. A catalyst sample from the performance testing of the Nitro-100 model was examined (Figure 4.5) and despite there being a moderate amount of rhodium particles remaining it appears that the number of crevices and their size has been severely reduced, which could indicate a phase transformation of the alumina. Even though the Nitro-100 thruster has a nominal operating temperature of 700 C it is possible that in certain locations within the catalyst bed there are hot spots that can exceed the transition

46 Chapter 4. Thruster Testing 34 Figure 4.8: Catalyst sample from prototype lifetime test at 10000x magnification (greater than 50 hours usage). temperature of about 850 C to form δ-alumina. At these temperatures sintering of the rhodium is likely to have occurred causing diffusion and coalescence of the surface rhodium, which would explain why it seems like there is less rhodium. In addition the fluid flow could also have ablated the rhodium particles away. The catalyst sample from the protoflight thruster performance test (Figure 4.6) has a more similar structure to that seen in the fresh sample although with smaller crevices, again indicating a likely phase transition, especially since the nominal operating temperature was about 900 C. It also appeared to contain much fewer rhodium flakes, both of these are likely causes of the catalyst deactivation observed after about 1 hour of operation. The two catalyst samples that underwent significant usage are seen in Figures 4.7 and 4.8 where both the crevices and rhodium particles are greatly reduced. Another feature worth noting are small needle-like formations about 1 µm long that were observed in these highly used samples and can be seen in Figure 4.9 (same as Figure 4.7 but at 25000x magnification). Although it is uncertain exactly what these needle-like structures are some preliminary research has indicated that it may be rhodium or rhodium oxide crystallized nanofibers [21]. Alternatively it could be an indication of mullite formation, which is a silicate mineral that is created by combining alumina with silica through sintering [22] Energy Dispersive X-ray Analysis In conjunction with the SEM imaging an EDX analysis was performed to quantify the percentage of rhodium relative to other elements. It does this by detecting x-rays emitted from the sample that has been bombarded by the electron beam [23]. Each of the above samples were examined using this technique and the results have been summarized in Table 4.5. The results from the EDX analysis confirm that rhodium content decreases with catalyst usage, even

47 Chapter 4. Thruster Testing 35 Figure 4.9: Catalyst sample from Nitro-100 restart test at 25000x magnification showing the needle-like features. Table 4.5: Results from Energy Dispersive X-ray Analysis Sample Description Usage (hours) % of Rhodium Fresh Catalyst Nitro-100 Performance Test Protoflight Performance Test Nitro-100 Restart Test Prototype Lifetime Test after just 1 hour of use the relative rhodium content had decreased by almost half. It is suspected that two mechanisms are at play here as mentioned earlier. Thermal sintering of the rhodium particles could cause a formerly even distribution over the substrate surface to clump into smaller areas. This decreases the effective surface area that is available for decomposition. Also, fluid moving past the particles could be ablating the rhodium particles and exhausting them through the nozzle all together. Although an analysis of the exhaust composition would have to be performed to confirm this hypothesis. In addition to the EDX testing, an X-Ray Powder Diffraction (XRD) analysis could also identify the phases of the various crystalline materials, which would help to verify if thermal sintering of the alumina is occurring. Overall, visual and chemical analyses indicate a reduction in potential decomposition sites leading to the observed deactivation of the catalyst Catalyst Restart Test The main objective of the Catalyst Restart test was to identify if a particular preheat operation could maintain consistent decomposition of the nitrous oxide regardless of the catalysts level of wear. This test aimed to discover if there was a certain preheat temperature that would initiate decomposition under any circumstance.

48 Chapter 4. Thruster Testing Test Setup The Nitro-100 thruster was filled with a fresh batch of catalyst, which would be used for the entirety of the restart test, and the nozzle was threaded and sealed onto the chamber and was then placed within fire bricks to limit convective heat loss. The interior heater and exhaust thermocouples were used for temperature feedback. A tank of medical grade nitrous oxide was used as the propellant supply with a pressure regulator and gauge being used for pressure control. Just upstream of the thruster was a flow meter which was used as feedback to adjust the regulator to the desired flow rate. All of the thermocouples were connected to the DAQ to log temperature data Methodology 1. The thruster was preheated to the nominal conditions (i.e., until the heater temperature reached 400 C). 2. Nitrous oxide was then flowed through the thruster at the desired flow rate of 76 mg/s and the heaters were turned off. 3. If decomposition was achieved the system was allowed to reach steady-state conditions (i.e., when the exhaust temperature increased by less than 1 C/min) 4. After 1 hour at steady-state temperatures the flow was stopped. 5. Steps 1 to 4 were repeated until decomposition was not able to be initiated or was unsustained for the entire 1 hour duration. 6. Once failure was observed at a particular preheat temperature, the preheat temperature was increased by 50 C. 7. At each preheat temperature the number of successful operations were recorded. 8. This was repeated until either: (a) Decomposition was not initiated or unsustained at any temperature. (b) A preheat temperature of 700 C was reached (due to thermal limitations on the heater elements it was not recommended to operate higher than 760 C, so 700 C was selected as a safe limit) Results The results from the catalyst restart test have been summarized in Table 4.6. They identify how many successful 1 hour thrust operations were achieved at a particular preheat temperature consecutively before decomposition was either not initiated or was unsustained. At the temperature range of C decomposition was able to be sustained consistently for many more times than at other temperatures. This could be an indication that the optimal preheat temperature is in this range. To confirm a dedicated lifetime test using these preheat conditions was performed.

49 Chapter 4. Thruster Testing 37 Table 4.6: Catalyst Restart Test Results Preheat Number of Successful Temperature ( C) Operations Before Failure Failure Method 400 (Nominal) 1 No decomposition No decomposition No decomposition Unsustained decomposition Unsustained decomposition Unsustained decomposition Unsustained decomposition Total Successful Operation 28.5 hrs Catalyst Lifetime and Comparison Test Seeing as the catalyst was still not performing as expected from a lifetime standpoint, a final test was performed to observe if a fresh batch of catalyst pellets operated at the seemingly optimal preheat temperature (based on the restart test) would provide better lifetime characteristics, to quantify this lifetime, and eventually compare alternate catalyst materials for longevity and overall performance. A dedicated lifetime test was performed previously with the prototype model thruster (Section 4.1.2), however it was necessary to perform the lifetime test again with the new Nitro-100 thruster to have a baseline for comparison against other catalyst materials. The other material obtained was iridium metal on an alumina support, as has been used in many other catalysis applications Methodology The test setup was the same as for the restart test, however the procedure was slightly different. 1. After the thruster was filled with fresh catalyst and the nozzle was threaded and sealed onto the chamber, the thruster was preheated to 650 C. This was selected since it was the highest temperature from the restart test that was able to initiate sustained decomposition repeatedly. 2. The propellant was then flowed through the thruster at 76 mg/s and the heaters were subsequently turned off. 3. Upon reaching steady-state conditions, the thruster was allowed to operate for 8 hour durations. The focus here was not on how the catalyst performed during restarts but how it performed over many hours of continuous operation. 4. This was repeated until decomposition was not initiated or unsustained. 5. The entire procedure was then repeated for alternate catalyst materials Results The lifetime of the rhodium on alumina catalyst was determined to be approximately 22 hours, which is close to what was observed in the restart test (28.5 hours), indicating that changing the nominal preheat temperature to 650 C (from 400 C) did not solve the issue of premature deactivation. Although the iridium catalyst was procured, due to reallocation of resources towards other projects, further investigations into whether it would have proved to have better lifetime were never performed. It is recommended

50 Chapter 4. Thruster Testing 38 that before any future modifications to the monopropellant thruster, testing the iridium catalyst should be the first step. 4.2 Performance Characterization A series of thrust tests were performed for each model of thruster to evaluate their performance against the expected results. Due to the limitations in manufacturing tolerances of the small nozzles it is important to characterize the thrust, specific impulse, and mass flow rate levels and ensure they are performing as designed Test Setup A vacuum chamber was used for all thrust testing since it can adequately reduce the pressure of the environment around the thruster to below Torr using a combination of roughing pump and turbopump to evacuate the majority of the air. Throughout thrusting, a copper plate that is mounted inside the chamber and known as the cold head since it is cooled to about -220 C condenses any remaining molecules to maintain high vacuum in the chamber. Inside the vacuum chamber the thruster is mounted to a thrust stand which in turn sits on a microbalance (Figure 4.10). Having the thruster thrust downwards onto the balance allows one to measure the effective mass imparted by a thrust and ultimately this can be converted to a force. All thruster models have been designed with one to two internal thermocouples to monitor exhaust and sometimes heater temperatures and for the purposes of performance testing, as well as being outfitted with an additional three to five thermocouples on the exterior for monitoring various locations of interest. All wiring and plumbing attached to the thruster base was secured to a standoff to prevent spring-like effects that would affect the thrust measurements. Temperature, mass flow rate, and force (mass) were logged using DAQ s, while the thruster heaters were connected to a power supply. Figure 4.10: RMT performance characterization test setup.

51 Chapter 4. Thruster Testing Methodology The parameters used for the performance tests of each thruster were slightly different (as outlined in Table 4.7), however they all shared a general procedure: 1. The vacuum chamber was pumped down to about Torr. 2. The thruster was then preheated to its nominal temperature. 3. Nitrous oxide was flowed into the thruster at the desired flow rate. 4. In the case of the monopropellant only, the heater was powered off since the decomposition is self-sustaining. 5. The system was allowed to reach steady-state conditions. 6. The flow of nitrous oxide was stopped after the desired thrust duration Test Parameters and Results The test parameters and the subsequent results for each of the tests have been summarized in Table 4.7 below. Table 4.7: Performance Test Results Summary Resistojet Monopropellant (Nitro-100) Preheat Temperature ( C) Mass Flow Rate (mg/s) Steady-State Temperature ( C) Thrust (mn) Specific Impulse (s) Although the resistojet implementation did not perform as well as expected, exhibiting slightly lower thrust and specific impulse, it was still considered to be successful seeing as the design of the heat exchanger could be modified easily to improve the efficiency. The monopropellant thruster on the other hand performed exceptionally showing slightly higher than expected specific impulse. Even though the Nitro-100 model was designed to operate at a much lower temperature than previous models (700 C compared to 900 C+), the expected performance decrease was not observed likely because the flow was more evenly heated and the chamber length adjusted for a more optimal exhaust temperature. The results of the performance characterization tests show that a highly-capable electrothermal thruster is achievable. 4.3 Qualification Testing A series of qualification testing was performed on the Nitro-100 model thruster to ensure the design would withstand the harsh environments of launch and space. Vibration tests in each axis aimed to

52 Chapter 4. Thruster Testing 40 ensure no loosening or damage will occur during launch. A cold preheat test ensured the heaters had enough authority to preheat the catalyst bed. Finally, it was necessary to examine how the thruster would respond if the incoming propellant was at a low temperature in the cold propellant test Vibration During launch severe vibrations and mechanical loading are experienced by the rocket and its payloads. Since each launch vehicle has slightly different vibration profiles it was necessary to test the thruster to a worst-case combination of the most common launch vehicles maximizing its utility over any mission or launch opportunity. The test profiles are designed to subject the system to representative vibration and shock loads and are multiplied by a safety factor for even higher confidence Test Setup The Nitro-100 model thruster was fully assembled and mounted to the vibration mount (Figure 4.11), with all bolts being torqued to the proper specifications and secured with a room-temperature-vulcanization (RTV) silicone to prevent loosening. Nine accelerometers were mounted to the thruster: one in each axis at the nozzle, housing, and interface plate (Figure 4.12). The thruster and mount were then secured to the vibration table (being rotated for each orientation) as in Figures 4.13 and A control accelerometer was also mounted to the table to accurately perform the specified profiles. Figure 4.11: Vibration test fixture with thruster mounted and coordinate system definition Methodology The vibration test can be split up into three major segments, each corresponding to one of the principal axes (x, y, and z). For each axis: sinusoidal burst, sinusoidal sweep, random vibration, and shock profiles were performed. Before and after each of the these profiles a low-level sinusoidal (LLS) test was performed initially to provide a baseline vibration response for the test article, with subsequent ones compared against the baseline. Any change in the response could indicate that loosening or damage may have occurred, prompting an inspection before continuing additional profiles. In addition, a functional test of the thruster electronic components (heaters and thermocouples) was performed before and after each

53 Chapter 4. Thruster Testing 41 Figure 4.12: Placement of nine accelerometers plus control on the nozzle, housing, and interface plate. Figure 4.13: Monopropellant thruster mounted on the vibration slip table (x and y axes). axis. many of the selected levels were based on either General Environmental Verification Specification (GEVS) since it is an accepted classification system, or the DNEPR or Polar Satellite Launch Vehicle (PSLV) launch vehicles since those are ones that have typically been used by SFL or are common for launching small satellites.

54 Chapter 4. Thruster Testing 42 Figure 4.14: Monopropellant thruster mounted on the vibration head expander (z axis). Low-Level Sinusoidal Performed as a standard test at SFL, the goal of the LLS test was to determine the natural frequencies of the test article before and after each major profile. If the structural integrity of the test article was compromised, the natural frequency would likely shift from the baseline allowing operators to identify issues before proceeding. The test used the following parameters: frequency range of Hz, sweeping through at a rate of 2 octaves/minute, at an amplitude of 0.5g (where g is the acceleration due to gravity of 9.81 m/s2 ). Sinusoidal Burst A quasi-static acceleration test was performed based on the levels expected on the DNEPR launch vehicle (Table 4.8). The sinusoidal burst test is essentially a strength test where the low frequencies and high displacements allow one to achieve as close to a static test as possible without exciting the natural frequencies of the system. Table 4.8: Accelerations from the DNEPR Launch Vehicle Acceleration Platform A Platform B Longitudinal g-load +10.7, , -1.3 Transverse g-load +1.0, , -3.4 Thus, using the worst-case results the test was performed at a vibration level of 10.7g over a frequency range of 9-10 Hz at a sweep rate of 0.5 Hz/second. The profile of which is given in Figure Sinusoidal Sweep Similar to the LLS test described above, the sinusoidal sweep is a test intended to identify where the natural frequencies and modes occur for the test article. The launch vehicle

55 Chapter 4. Thruster Testing 43 Figure 4.15: Sinusoidal burst profile showing amplitude vs frequency. provider requires this testing be performed to ensure the test article does not get excited from the nominal operating frequency of the rocket engines, which typical occurs at frequencies below 100 Hz. The test levels used correspond to those suggested for the Polar Satellite Launch Vehicle (PSLV) and are summarized in Table 4.9. Frequency (Hz) Table 4.9: Sinusoidal Sweep Test Levels Acceleration (g) Displacement (mm, 0-peak) Sweep Rate: 2 octaves/minute Random Vibration The random vibration test is typically the most vigorous test out of all the profiles and is the one that best represents the randomness of an actual launch. The selected levels were based on the GEVS classification for Device Under Test (DUT) under 22.7 kg. The frequency ranged from Hz and although the acceleration could spike up to some very high levels for extremely short durations, the average acceleration level was specified as 14.1g rms (where g rms is the root mean square of the acceleration across all tested frequencies). Also the test duration was specified to be 120 seconds. Table 4.10 summarizes these levels and the power spectral density (PSD) at various frequencies. Table 4.10: Random Vibration Test Levels Frequency (Hz) PSD (g 2 /Hz) g rms 14.1 Duration 120 sec

56 Chapter 4. Thruster Testing 44 Shock Test Since the PSLV is a likely potential option for SFL small satellites the levels experienced in past PSLV missions can be used as a baseline for the shock test. A level of 50g, in both the positive and negative directions, was used for the qualification of the thruster. This is the worst-case observed in past missions (40g) plus a 25% margin. It was conducted with a half-sine waveform having a duration of 10 ms, as shown in Figure Figure 4.16: 50 g shock test waveform Results The following plots show the results from the qualification vibration test of the thruster and display the accelerometer responses for the various vibration inputs. Figures 4.17, 4.18, and 4.19 show the LLS responses for the x, y, and z-axes respectively. These plots bring to light the first natural frequencies (FNF) of the test article in each axis. The lowest FNF observed for any axis was 470 Hz, which was well above the 90 Hz minimum requirement that most launch vehicles provide. It can also be seen that there was some coupling between the in-axis and off-axis accelerometers, likely a result of the three-point mounting configuration of the thruster to the interface plate. Owing to the high frequency at which it occurs, this coupling was not considered to be of concern. As examples, Figures 4.20, 4.21, 4.22, and 4.23 show the responses from the sinusoidal burst, sinusoidal sweep, random vibration, and shock tests respectively. These plots are just for the y-axis responses however similar results were seen for the other two axes as well. As can be seen in the y-axis results, the accelerometer responses for each test closely followed the input vibration profiles and this was also the case for the other two axes. In between profiles, visual inspection

57 Chapter 4. Thruster Testing 45 Figure 4.17: Accelerometer response for the 0.5 g low-level sine test in the x-axis. Figure 4.18: Accelerometer response for the 0.5 g low-level sine test in the y-axis. and LLS tests confirmed that no loosening or damage was done to the test article throughout the entirety of the test. Also, functional tests in between axes verified that the electronics did not suffer from the vibrations either. Thus, it can be concluded that the Nitro-100 monopropellant thruster would survive the launch environment.

58 Chapter 4. Thruster Testing 46 Figure 4.19: Accelerometer response for the 0.5 g low-level sine test in the z-axis. Figure 4.20: Sinusoidal burst accelerometer response in the y-axis Cold Preheat Preheating the catalyst bed is an important aspect of initiating decomposition of the propellant, where the decomposition temperature for rhodium catalyst is around 250 C. The goal of the cold preheat test was to ensure that the heaters in their current configuration within the thruster would have enough authority to heat the catalyst bed to a sufficient temperature in a reasonable amount of time. The

59 Chapter 4. Thruster Testing 47 Figure 4.21: Sinusoidal sweep accelerometer response in the y-axis. Figure 4.22: Random vibration accelerometer response in the y-axis. preheat duration had been designed to be less than 5 minutes. The worst-case scenario to be tested is when the thruster has reached equilibrium at its coldest operational temperature of -30 C Test Setup The Nitro-100 monopropellant thruster was fully assembled onto the Thermal VACuum (TVAC) mount with all bolts torqued to specifications. Temperature sensors were also mounted to the thruster housing,

60 Chapter 4. Thruster Testing 48 Figure 4.23: Negative 50 g shock accelerometer response in the y-axis. interface plate, and TVAC mount. The thruster assembly was then placed inside a vacuum bell jar, which was in turn placed inside a thermal chamber, seen in Figure It was the combination of bell jar and thermal chamber that allow for testing in a representative space environment. Temperature data was gathered from the internal thermocouples as well as the externally mounted sensors through a DAQ and heater power was provided from a dedicated power supply. Figure 4.24: Test setup for the cold preheat thermal vacuum test.

61 Chapter 4. Thruster Testing Methodology 1. The bell jar was pumped down to about 200 mtorr. 2. The chamber was set to the lower operational temperature limit of -30 C. 3. Once equilibrium was reached, the heater was supplied with 30 W of power and the timer was started. 4. The thrust chamber was allowed to heat up until the heater thermocouple measured 400 C. Note that this preheat temperature was used even though it was well above the catalytic decomposition temperature of rhodium to allow for the entire catalyst pack to reach a high enough temperature instead of just the area near the thermocouple. 5. The heater was then powered off and the system was left to cool Results Figure 4.25 shows the results from the test where the temperature of the thruster started at around -33 C and the heaters being activated at time = 0 minutes. The heater thermocouple reached the desired value of 400 C after 4 minutes and 50 seconds, proving that the heaters are capable of bringing the catalyst bed up to the required temperature even under the worst-case conditions, thus meeting the preheat time requirement. Temperature ( C) Heater Exhaust Interface Plate Time (min) Figure 4.25: Temperatures during cold preheat test Cold Propellant Operation Due to the high thermal inertia of such a large quantity of propellant (11.2 kg), the temperature of the tank was expected to be relatively moderate, however some fluctuations may occur depending on the surrounding spacecraft temperature. Initiating decomposition is a function of the temperatures of the incoming propellant as much as it is the catalyst bed; and since it was confirmed in Section 4.3.2

62 Chapter 4. Thruster Testing 50 that the heater can preheat the catalyst under worst-case cold operating conditions, it was necessary to understand how the thruster performs when the incoming propellant is cold as well Test Setup The Nitro-100 monopropellant thruster was fully assembled onto the test stand with all bolts torqued to specifications. Temperature sensors were also mounted to the thruster housing, interface plate, and test stand. The thruster assembly was then placed under a fume hood to extract the exhaust gases during thrusting. A supply tank of N 2 O was used with a flow meter downstream of the pressure regulator to supply propellant to the thruster. Before reaching the thrust chamber however, the propellant feed line was routed through a heat exchanger to cool it down from room temperature (23 C) to -10 C. Temperature data was gathered from the internal thermocouples as well as the externally mounted sensors through a DAQ and heater power was provided from a dedicated power supply Methodology 1. The thruster was operated using room temperature propellant to obtain a baseline for comparison. 2. The heat exchanger was set to cool the propellant to the minimum temperature. It allowed for the propellant to be cooled to -10 C, which was verified by placing a thermocouple in the feed tube immediately before the thruster inlet. 3. Once the heat exchanger was set, the catalyst bed was preheated to the nominal 400 C. 4. Propellant was then flowed through the thruster and the system was allowed to reach steady-state conditions Results Figure 4.26 shows the results from the baseline test at ambient temperatures of 23 C, whereas Figure 4.27 shows the temperature results for the cold propellant test where the propellant temperature was cooled to -10 C. It is important to note that the steady-state exhaust temperatures achieved during non-vacuum tests (like the two described in this section) and vacuum tests (like the performance characterization tests) cannot be directly compared due to the significant heat loss from convection that is experienced in ambient air. By examining the resulting steady-state temperatures of the heater and exhaust thermocouples from both tests, it can be seen that the propellant temperature does not have a significant effect on the performance of the thruster. In both tests the heater temperature steadied out at about 500 C, while the exhaust reached about 540 C. The main difference occurs in the transient portion of the response where the flow was initially input into the thruster and in the cold propellant test, the heater temperature dropped by a larger amount than the baseline test. The incoming propellant pulls more heat away from the inlet (where the heater thermocouple is measuring), however the temperature quickly recovers and is able to reach the same steady-state value. Because of limitations in the test hardware the propellant was not able to be cooled down to a lower temperature, however the results from this test suggest that the propellant can be even colder without impacting the steady-state performance of the thruster.

63 Chapter 4. Thruster Testing Heater Exhaust Temperature ( C) Time (min) Figure 4.26: Temperatures during baseline test at 23 C. 600 Temperature ( C) Heater Exhaust Time (min) Figure 4.27: Temperatures during cold propellant test at -10 C. 4.4 Thruster Evaluation and Conclusions The Nitro-100 monopropellant thruster exhibited exceptional performance in vacuum conditions and withstood all of the qualification testing including: vibration, cold preheat, and cold propellant operation; however the lifetime of the catalyst is still of concern. Reducing the operating temperature of the thruster did not seem to mitigate the rate of degradation; in addition, increasing the preheat temperature still resulted in the eventual failure to initiate decomposition. Despite there being a lot of potential for the monopropellant implementation the challenge in finding a catalyst that can withstand the high temperatures makes it very difficult to maintain consistent decomposition. Further reducing the operational temperature to prevent catalyst degradation would greatly reduce the performance of the thruster to the point where it is not gaining any advantage over the resistojet configuration. Alternate catalyst materials, such as Iridium, will need to be investigated and show promise due to its higher sublimation temperature than rhodium, which could indicate that it is not as susceptible to the thermal

64 Chapter 4. Thruster Testing 52 sintering that rhodium has been affected by. Further work needs to be performed to ensure that the monopropellant can be relied on for an actual mission.

65 Part II DAUNTLESS Bus Propulsion System Development 53

66 Chapter 5 Preliminary Analysis The DAUNTLESS bus can be scaled to have a dry mass of up to 500 kg, however for the current mission it was designed for 70 kg. To accommodate the avionics, the satellite has an internal panel that acts as the main mounting point for the majority of the bus components and also provides a path through which all of the loads are transmitted during launch. It has the capability of operating at much higher orbits than previous SFL missions with the nominal being a 1000 km circular orbit. Thus at the end of the mission it requires propulsive maneuvers to lower its orbit to an altitude where atmospheric drag is significant enough to complete the deorbit process. Having successfully demonstrated SFL s capability to develop a high-performance thruster, a complete propulsion system (including the tank, feed system, and fill module) can be designed to fulfill the deorbit needs of the DAUNTLESS bus. 5.1 Requirements Several important requirements drove the design of the propulsion system and are summarized in Table 5.1. The propulsion system on the DAUNTLESS bus will mainly be used for deorbiting purposes and to determine the V needed to transfer from a 1000 km circular orbit to an elliptical orbit where the perigee passes through the atmosphere (about 400 km altitude) a circular-to-elliptical transfer can be used. By calculating the velocities required at the transfer point of the two orbits, the difference: 159 m/s is what is necessary to transfer between them, and is to be performed at apogee over the course of several orbits. Accounting for orbit acquisition, correction, and collision avoidance maneuvers a final requirement of m/s was selected. As explained in the requirements section for the electrothermal thruster (Section 2.2) the minimum impulse bit allows for a trade-off between thrust level and thrust time to ensure the necessary impulse is imparted. Since the deorbit maneuver cannot be performed in a single impulsive burn, the minimum thrust requirement ensures that the required V can be achieved within a reasonable number of thrusts (translating to total deorbit duration). The specific impulse specification defines how efficient the system has to be and ultimately how much propellant is required. The preheat power requirements are divided into a time term and an energy term. By defining it in this way, designers can trade-off peak power and preheat duration to minimize the total energy used. 54

67 Chapter 5. Preliminary Analysis 55 Table 5.1: DAUNTLESS Bus Key Propulsion Requirements Summary Requirement Number Description PRP-101 Shall be sized to impart a minimum V of m/s to the spacecraft for an orbital altitude of 1000 km or less. PRP-102 Shall provide a minimum impulse bit of 500 mns or less. PRP-103 Shall provide a thrust magnitude of 50 mn or greater. PRP-107 Shall achieve a specific impulse of 115 s or greater. PRP-108 Shall require no more than 10 minutes to preheat. PRP-109 Shall require no more than 4.2 Wh of DC power during preheat PRP-110 Shall require no more than 75 W DC power while thrusting. PRP-111 Shall exhibit no more than 5 mg/hr propellant loss at any temperature within survival range (while not thrusting). PRP-204 Shall operate within specification over an input voltage range of 21V to 34V DC. PRP-205/206 Shall survive, though not necessarily operate, between 0V and 40V DC. PRP-209 Shall be designed to a maximum expected operating pressure (MEOP) of 1400 psi PRP-210/211 Shall be designed to withstand a proof pressure of 1.25 times MEOP (1750 psi) without yielding and a burst pressure of 1.5 times MEOP (2100 psi) without failure. PRP-215 Shall operate within specification over temperature range of -20 C to 40 C. PRP-216 Shall survive, though not necessarily operate, between -30 C and 60 C. PRP-226 Shall account for 1 cm and 5 offset in any direction of thruster mounting. The thrusting power requirement allows the resistance of the heaters to be selected and ultimately drives the allowable thermal efficiency of the thruster. During continued operation the thermal losses will define the duty-cycle of the heater, which when combined with the peak power of the heater determines the average power during thrusting. The voltage range requirements are derived from what is necessary to accommodate various battery configurations and accounts for voltage drops over the wiring harness. It also defines what levels need to be tested to during electronics testing. The total propellant can be calculated based on the worst case performance parameters but also needs to account for leaking that will be experienced throughout the mission lifetime. Thus a maximum leak rate was specified to ensure enough propellant would be remaining for the deorbit maneuvers at the End of Life (EoL). It is evident from the discussion in Section 2.2, that there is a strong relationship between temperature, pressure, and storage density of the propellant. A trade-off needed to be performed with the structural integrity of the tank at high pressures, size of the tank, and performance capabilities at both high and low pressures. With these in mind the storage density was selected and the tank was designed to a specific maximum expected operating pressure (MEOP), thus imposing temperature requirements. Finally due to uncertainties in the center of mass (CoM) location of the final spacecraft, it was necessary to have some method for adjusting the angle of the thruster (and thrust vector) to align with the CoM and minimize thrust induced torques. With the above stated requirements selected, a re-evaluation of the current RMT iteration was necessary to ensure the requirements were to be met. 5.2 Thruster Re-Evaluation Having previously concluded that as it is currently implemented, the monopropellant thruster does not have the desired performance for the DAUNTLESS point design due to catalyst limitations. Thus it

68 Chapter 5. Preliminary Analysis 56 was decided that the DAUNTLESS propulsion system would incorporate a resistojet thruster into the design. Despite being less efficient and requiring slightly more power, the consistent performance of the resistojet is desirable for deorbiting thrust maneuvers. The large size of the DAUNTLESS bus allowed for the inclusion of additional solar cells, meaning the slight increase in power for the resistojet thruster was able to be accommodated. To meet the requirements for typical missions that would use the DAUNTLESS bus, the propellant choice was re-evaluated and the thruster sizing was optimized for the new performance characteristics Propellant Selection Deciding to use a resistojet thruster over a monopropellant opens up the possibility for many more propellant types since they do not need to be decomposable. Nitrous oxide is a very energetic oxidizer with exothermic decomposition potential, thus protocol for filling, transporting, and testing must be developed prior to committing to using it in a satellite. For these reasons an evaluation into alternative propellants including ammonia, propane, butane, carbon dioxide, hydrogen peroxide, and water was performed to identify any that would offer comparable or improved performance characteristics while being safer to handle than nitrous oxide. Although ammonia, propane, and butane have similar or better performance and improved storage densities, there are safety concerns with using them making them unsuitable candidates. Hydrogen peroxide and water have promise as well although using liquid propellants adds increased complexity to the system including pumps, heat exchanger redesign, and concerns of freezing. Thus, carbon dioxide was considered the primary candidate for a replacement due to its similar performance to N 2 O and inherent safety. With all system design parameters and geometry being equal, switching to carbon dioxide only reduces the specific impulse by less than 1 s. This very small reduction in efficiency was deemed to be well worth the increased safety in handling and operation that would be gained by using CO Thruster Sizing Modifications The original resistojet was sized for a reference spacecraft with a dry mass of 150 kg that would achieve a minimum V of 100 m/s, which ultimately defined the required performance characteristics. However the baseline mission applicable for the DAUNTLESS bus requires different sizing parameters, thus the thruster needed to be redesigned and resized accordingly. A larger cartridge heater was implemented with two redundant heater elements and an integrated thermocouple. The heater elements were sized to provide the necessary heat input to the fluid at the nominal bus voltage of 28.8V while respecting the power requirements. Additional baffles, with an optimized design, were included to increase the heat transfer capabilities to the propellant, increasing efficiency, and the nozzle was also resized to meet the thrust requirements. Table 5.2 summarizes the new sizing parameters and performance characteristics for the updated Carbon dioxide Resistojet Thruster (CRT). Figure 5.1 shows the CRT. It can be seen that when comparing the CRT with the resistojet implementation of the RMT, an improvement in specific impulse is achieved. This increase of about 10 s is due to the slightly higher chamber temperature and much larger expansion ratio both of which contribute to better efficiency as described in Section 3.1. The flow rate is also much lower but this is mainly because of the lower thrust requirement.

69 Chapter 5. Preliminary Analysis 57 Table 5.2: Updated Resistojet Sizing and Performance N 2 O Resistojet CO 2 Resistojet System Performance Requirements Spacecraft Wet Mass (kg) Total V (m/s) Minimum Impulse Bit (mns) System Design Parameters Thrust (mn) Chamber Temperature ( C) Nozzle Expansion Ratio Nozzle Half-Angle ( ) Propellant Chemical Properties Specific Heat Ratio - N 2 O (@ 500 C) Specific Heat Ratio - CO 2 (@ 580 C) Molar Mass - N 2 O (g/mol) Molar Mass - CO 2 (g/mol) Performance Outputs Expected Specific Impulse (s) Minimum Thrust Duration (s) Mass Flow Rate (mg/s) Total Propellant Mass (kg) Figure 5.1: The updated Carbon dioxide Resistojet Thruster (CRT), radiation shield not shown.

70 Chapter 6 Propulsion System Design 6.1 System Architecture The thruster is just one part of the propulsion system and there is much more involved in order to get the propellant from the tank to the thruster, in addition to getting propellant into the tank in the first place. In addition to the tank, the propulsion system can be divided into two general modules: the feed system, which contains the filter, valves, plenum, thruster, and electronics; and the fill/drain module, which contains valves and ports that can be accessed from the outside of the satellite in order to fill and drain the propellant tank. An additional system is attached to the fill/drain module that is designed to mitigate over-pressure events and consists of two rupture disks in series. A block diagram of the entire system can be seen in Figure 6.1. The feed system is necessary to ensure that the propellant gets to the thruster at the appropriate time and at the desired pressure. Beginning with a feed tube from the tank, the propellant travels through a filter to ensure no particles clog small orifices further downstream. The plenum consists of the restrictor to limit the flow, a regulator valve, a small cylinder to expand the propellant (known as the secondary volume), and a pressure sensor for feedback. By carefully monitoring the pressure of the secondary volume the regulator valve can be opened to allow a certain amount of propellant through. The restrictor aids in limiting the flow through the regulator valve and thereby minimizes the cycle count of the valve. The last component before the thruster is the thrust valve which is only actuated when thrust maneuvers are commanded and provides a final seal to propellant leaking out while not in operation. The electronics board (e-board) contains all necessary hardware and software to control the heaters in the thruster, command the valves on and off, and also to read telemetry from various components including the thermocouples, thermistors, and pressure sensors. In total there are two pressure sensors in the system: one as part of the plenum (as mentioned above) and the other upstream of the filter used to monitor tank pressure. The other module that makes up the propulsion system is the fill/drain module. Diaphragm valves were used along with plugged ports to connect feed lines during filling and draining and to provide redundant seals after the satellite has been filled. These are also accessible from outside of the spacecraft to ensure that fill and drain activities can be accomplished once the satellite is fully assembled, 58

71 Chapter 6. Propulsion System Design 59 SPACECRAFT FILL PORT FILL VALVE PRESSURE SENSOR (SECONDARY) SECONDARY VOLUME THRUST VALVE VENT PORT PROPELLANT TANK REGULATOR VALVE THRUSTER VENT VALVE PLENUM RESTRICTOR FILTER SECONDARY RUPTURE DISK PRIMARY RUPTURE DISK PRESSURE SENSOR (PRIMARY) FILL/DRAIN MODULE FEED SYSTEM Figure 6.1: DAUNTLESS bus propulsion system block diagram showing the fill/drain module on the left side of the tank and the feed system on the right. for example at the launch site. This module also includes the over-pressure system, which consists of two rupture disks that are rated to about 1.5 times MEOP, one of which evacuates to the outside of the satellite. In the case that increased temperature results in the pressure increasing these rupture disks are intended to burst thus relieving the pressure and preventing damage to the tank and subsequently other important components in the satellite. It was decided for simplicity to not incorporate a pressure relief valve meaning the entire tank will vent once both rupture disks burst. Since the primary purpose of the propulsion system is deorbiting and nothing particularly mission critical this was deemed to be an acceptable consequence. 6.2 Flight System Configuration With the schematic of the propulsion system and its corresponding components in mind, the flight propulsion system was designed to fit into the available space within the DAUNTLESS bus and since this mission is likely the smallest realization of the bus it would be the worse case implementation. The DAUNTLESS bus has an interesting design in that the majority of the avionics are located underneath the main internal panel, with the entire upper portion being designated to the propulsion system. This is partially owing to the fact that the size of the propellant tank is required to be so large. Having such a large space to work with allowed for a flat plate design philosophy, which although being inefficient with space is much easier from an assembly standpoint. The idea is that all components are assembled onto an easily accessible flat plate. These components, if necessary, can be replaced with very minimal interference of other components, making it easy to troubleshoot the system even in later stages of the

72 Chapter 6. Propulsion System Design 60 assembly. This proved to be a crucial aspect when leak-checking the system. Figure 6.2 shows the solid model of the propulsion system mounted on the DAUNTLESS bus internal panel. The following sections outline some of the major design issues that were encountered and how they were overcome. Figure 6.2: Propulsion system mounted on the DAUNTLESS bus internal panel showing the tank, feed system, and fill/drain module Feed System The feed system is the most complex part of the propulsion system since it contains all of the electronics and critical flow components to deliver the propellant to the thruster, ultimately generating the necessary thrust. A prototype system was designed and built to perform initial testing with all components working together and was controlled using a spare electronics board from CNAPS. It also provided a platform for a number of tests later into the development process including most of the preliminary software testing and the valve lifetime test. Upon doing a series of tests, the design was iterated on and a final configuration was realized, shown in Figure 6.3. All the required components were assembled onto a small plate and then the plate was mounted to the spacecraft internal panel using standoffs. With so many different pieces working together there were many challenges that needed to be overcome to achieve successful operation including: designing the plenum to effectively regulate the pressure and flow rate to the thruster, determining valve direction (forward or reverse configuration) to optimize their fulfillment of certain tasks, ensuring the pressure sensors were accurate over the entire operational range, and mounting the thruster to both ensure thermal isolation and also achieve adjustability of the thrust vector Plenum Configuration The main purpose of the plenum is to reduce the pressure of the propellant from the stored vapour pressure down to a more usable pressure for the thruster. Since the vapour pressure of CO 2 is approximately

73 Chapter 6. Propulsion System Design 61 Figure 6.3: Feed system design showing the configuration of all components. 830 psi at room temperature and can fluctuate between 285 and 1400 psi over the entire operational temperature range, it is important for all components to be able to withstand the higher pressures. This was initially accomplished in the prototype system (Figure 6.4) by using miniature solenoid valves (that had a relatively low pressure rating) and a dual-stage regulator to reduce the pressure to around 300 psi upstream of the regulator valve. However the regulator was deemed unreliable for the temperature range of the mission, attributable to complications at cold temperatures. This prompted the switch to a design without a mechanical regulator thus requiring higher pressure valves to withstand the vapour pressure directly. Thorough testing was necessary to retire any risks of damage or leaking in the valves at such high pressures. In addition, the restrictor and secondary volume were resized to minimize the regulator valve cycle count while maintaining a reasonably steady thrust output Solenoid Valve Direction Solenoid valves were selected for the feed system and operate in a normally closed configuration meaning when unpowered a spring forces the poppet closed sealing off the input port from the output port. When the coil assembly is powered the armature is forced upwards thereby working against the spring and opening the poppet. Over many open-close cycles the poppet can begin to degrade, compromising the sealing interface and potentially causing an increase in the leak rate through the valve. From a configuration standpoint it is important to note that one port is considered to have pressure under-seat

74 Chapter 6. Propulsion System Design 62 Figure 6.4: Prototype feed system showing the initial plenum configuration. It was used for proof of concept and some risk mitigation testing. (PUS) while the other has pressure over-seat (POS) as seen in Figure 6.5. With the higher pressure being under-seat the propellant is always working against the spring and if increased high enough it can overcome the spring force causing leakage. On the other hand, higher pressure over-seat works with the spring and forces the poppet closed. If the pressure reaches a high enough value the solenoid may not be able to overcome the pressure to open the valve. Both of these aspects contribute to the pressure rating of the valve. When designing the feed system it was selected that the regulator valve, which sees significantly more cycles, would have the higher pressure under-seat (i.e., forward configuration) to limit the force of the poppet closing against the valve body in an effort to reduce the degradation of the seal. The thrust valve, being the final seal to the vacuum of space, was selected to have the higher pressure over-seat (i.e., reverse configuration) to ensure that even under higher than expected pressures upstream of the valve, there will be minimal leakage through it. Figure 6.5: A schematic of the solenoid valve describing the various internal components.

75 Chapter 6. Propulsion System Design Pressure Sensor Temperature Compensation Pressure sensors are typically susceptible to measurement errors caused by varying temperature, which often cause readings outside of the calibration point to be inaccurate. Thus it is common for many sensors to come with some sort of temperature compensation that includes a gain and offset applied to the sensor reading. A series of calibrations over pressures and temperatures were performed to find an approximately linear correlation that could be implemented into the software to give accurate readings of pressure over the entire operational range Thruster Mounting The thruster was designed to be mounted using a series of washers and spacers. To stand off the radiation shield from the main body of the thruster, ceramic spacers were used to limit the conductive heat transfer from one component to the other. This ensures that the outer radiation shield remains as cool as possible, minimizing the radiative losses to its surroundings. To mount the entire thruster assembly to the mounting tower a series of conical spring washers were used. These washers have spring-like properties and when stacked in either series or parallel can modify the spring constant or the amount of deflection in the stack to meet the specifications of the design. When mounting the thruster a high spring constant was desired to maintain an appropriate pre-load force during the vibrations of launch and to minimize any alignment drift that may have been caused by those vibrations; however for adjustability purposes it was also desirable to have a moderate amount of deflection to work with. Through trial-and-error a stack configuration was selected to meet both of these needs. In addition to the mounting screws and washer stack, an alignment wedge was added to change the nominal angle of the thrust vector. This was necessary to better align the thruster with the Center of Mass (CoM) which moved late into the design phase. The wedge, although not present in the qualification model system, was a solution to this late change for the flight system. However since it was rigidly connected to the thruster mount and the mounting system for the thruster remains unchanged, it does not invalidate the qualification testing that was performed Fill/Drain Module The fill/drain module was designed to have ports and valves accessible from the outside of the satellite to facilitate filling and draining activities even after complete assembly. This was the main design driver for the final configuration of this module, as seen in Figure 6.6. Key challenges that needed to be addressed included: selecting appropriate tube sizing and providing mitigation techniques for over pressurization. Using 1/4 tubes over smaller 1/16 tubes gives the advantage of higher flow rates as well as a lower chance of getting blockages caused by foreign particles. Thus it was selected that all tubes involved with filling and draining the tank be 1/4 to ensure that these procedures can be performed in a reasonable time frame. In addition on the feed system side all tubes upstream of the filter were selected to be 1/4 to reduce the chance of clogging, whereas downstream they were selected as 1/16 tubes. The downside to using the larger tubes over the 1/16 tubes is that they must be custom cut, which means the ends must be carefully deburred to ensure a sufficient seal and that no particles become dislodged and enter the system. They also need to be manually bent, which when using stainless steel tubes can be quite

76 Chapter 6. Propulsion System Design 64 Figure 6.6: Fill module design. difficult, especially when trying to achieve abnormal angles to line up various components. The other main area of concern that was addressed in the fill module was that of over pressurization of the tank. This could really only be caused by an increase in temperature well beyond the survival limit. Although the tank was designed not to burst explosively but rather to crack and leak, it was still necessary to prevent this by incorporating a rupture disk directly exposed to the tank pressure. By selecting a rupture disk with a burst pressure lower than that of the tank, it can be ensured that the tank won t leak propellant into other parts of the satellite. Instead any over-pressure would be relieved and directed to a second redundant rupture disk on the outside of the satellite through a flexible tube. If the pressure continued building and burst the secondary rupture disk it would get ejected outwards in a dispersive fashion aiding in minimizing any undesired torques that may get imparted to the spacecraft as well as minimizing unintended V by venting in opposite directions. Figure 6.7 shows the design of the over-pressurization system Tank Selection The largest component of the propulsion system is the tank that the propellant is stored in throughout the satellite s life. Although many high pressure propellant tanks used in the past have been made out of monolithic metals such as titanium, recently hybrid options adding composite materials have become more popular as lighter weight options. For even more mass savings a custom linerless composite tank was designed using carbon fiber as the main component (shown in Figure 6.8). A storage density of 600 kg/m 3 was selected and knowing the necessary propellant mass of 10.8 kg results in a tank volume of 18 L. Aluminum end bosses on either side of the tank incorporate connection points to mount it

77 Chapter 6. Propulsion System Design 65 Figure 6.7: Over-pressure system. The secondary rupture disk is mounted to the external side panel of the satellite. to the internal panel, while also providing ports to attach plumbing connections. The challenge in manufacturing a linerless tank is ensuring that it can withstand the launch environment since there is no metal to provide added strength. The validity of this design was verified through a rigorous vibration test campaign described in Section Figure 6.8: Custom linerless composite tank.

78 Chapter 6. Propulsion System Design Flight Assembly After the individual modules were assembled independently, including the 1/4 tubes that attach to the modules, they were mounted to the internal panel of the spacecraft. One of the most difficult aspects of the assembly was making the final connection with the tee leading to the tank. Three manually bent tubes all must align in the exact orientation to ensure the tee was connected with a proper, leak-tight seal, yet it was difficult to bend the tubes to the exact angles and they were nearly impossible to modify once formed. It is recommended that future designs using this methodology incorporate simpler tubes (possibly even pre-bending them using dedicated equipment at external facilities) to ensure that they line up effectively. The final component to be integrated was the propellant tank, which had been fitchecked previously to line up with the fill and vent tubes. As a final note, some of the screws mounting the various modules to the internal panel were inaccessible with a standard torque wrench, requiring the employment of more exotic tools to deliver the required torques. For simplicity and to ensure that all screws are torqued to the appropriate specification, proper access must be considered for future iterations of the design. Figure 6.9 shows the final assembly of the DAUNTLESS bus.

79 Chapter 6. Propulsion System Design 67 Figure 6.9: Final DAUNTLESS bus assembly with internals visible (upper) and with the side panels mounted (lower).

80 Chapter 7 Testing The following sections outline the propulsion system testing that the author was most involved with throughout the development process including: solenoid valve testing, performance characterization of the thruster with the associated feed system, system-level qualification vibration test, and leak testing of the flight assembly. 7.1 Solenoid Valve Testing Finding high pressure valves that are small enough to incorporate into the propulsion system can be difficult to source at a reasonable cost, which resulted in the selection for COTS valves to be part of the feed system. However these valves were to be used outside of the specification ranges for pressure, cycle count, voltage, and temperature. Thus, a series of tests were conducted to confirm that the solenoid valves could withstand the conditions that they would be exposed to throughout their mission lives. This included extensive lifetime testing, thermal cycle testing, vacuum qualification, vibration testing on a unit-level, hydrostatic pressure testing, and finally acceptance testing prior to integration. The following sections describe in detail the tests the author was heavily involved with in addition to a brief summary of the results from the other tests Lifetime Test The valve seal and coil assembly of the solenoid valves selected to control the flow through the feed system are considered to have a limited life. Each valve is rated for a lifetime of cycles before the specified leak rate can no longer be guaranteed. In the DAUNTLESS bus the spike portion of the control signal to the valve is planned to be from the bus voltage of 28.8 V maximum, which is higher than the 24 V that the valve is designed for, and it is suspected that this may reduce the cycle life. Furthermore, not only is the valve seal expected to experience wear after many repeated open and close cycles, but it is also necessary to ensure that the seal material does not degrade significantly after being exposed to a lifetime s worth of CO 2. This comes from the fact that CO 2, especially when compressed, sometimes has compatibility concerns with specific seal materials, although this is not expected to be an issue with the seal material selected for this system s valves. 68

81 Chapter 7. Testing Test Setup The prototype feed system was used for this test, albeit with a slightly different plenum design. It was modified to accommodate leak tests in between each segment of regulation testing. After the regulator valve, instead of there just being a single volume and pressure sensor prior to the thrust valve, an isolation hand valve was placed between two volumes separating the regulator valve section from the thrust valve section. A schematic of this configuration can be seen in Figure 7.1. Figure 7.1: Modified prototype feed system schematic with updated plenum to accommodate leak testing of both valves. The plenum is the entire portion from the regulator valve to the thrust valve and is the volume that the control algorithm will be regulating. The volume just downstream of the regulator valve is known as the secondary volume and the attached pressure sensor, P1, allows for leaks through the valve to be detected during the leak tests. The intermediate volume is the volume downstream of the isolation valve but upstream of the thrust valve and was intended to be pressurized during leak tests to subject the thrust valve to high pressure. The volume downstream of the thrust valve and plugged at the end of the downstream restrictor is referred to as the tertiary volume, which has pressure sensor, P2, sensing any leaks through the thrust valve. Figure 7.2 shows the actual test setup Test Parameters Through preliminary testing with the prototype feed system and EM thruster a set of values were established that define the operational parameters of the regulator and thrust valves, including the definition of an operational lifetime. Table 7.1 summarizes these values and provides a brief description of how they were derived. Based on the operational lifetime of the propulsion system, 67.8 hours, a regulation frequency was selected to minimize the total cycle count while also attempting to limit thrust variability. Thus a frequency of 0.4 Hz was chosen resulting in a total cycle count of about cycles per lifetime, which is twice the specified cycle life. Since the valve specification stated the maximum number of cycles to be , it was crucial to perform a lifetime test to observe the performance and leak rates throughout a valve s life.

82 Chapter 7. Testing 70 Figure 7.2: Valve lifetime test flow control setup. Blue arrows indicate the direction of propellant flow. Table 7.1: Nominal Regulator and Thrust Valve Operational Parameters Description Symbol Value Notes Upstream Pressure p tank 830 psi Based on nominal vapour pressure of CO 2 at room temperature. Regulated Pressure p R 125 psi Required pressure in secondary volume to achieve the desired mass flow rate in a flight-like representation. Regulated Pressure Required to maintain acceptable thrust variability, while p Deadband R 50 psi reducing total cycle count. Nominal Propellant Mass ṁ p 44 mg/s Corresponds to 50 mn thrust at 115 s specific impulse. Flow Rate Regulator Valve Actuation Frequency f R 0.4 Hz Based on the frequency required to maintain the output pressure within the bounds. Determined by testing with the EM feed system in the nominal setup. Average Thrust Duration Total Thrust Duration Number of Regulation Cycles Maximum Bus Voltage Total Propellant Internal Volume of Plenum t thrust T thrust 2 minutes 67.8 hours Based on a conservative estimate of an average thrust during de-orbit phase (could be as low as 1 min up to 10 mins). 2 mins seemed like a reasonable balance between test time and thrust valve cycle count. Based on m/s V, 50 mn thrust, and 115 s specific impulse. N R = T thrust f R Total number of cycles over the entire thrust duration. V busmax 28.8 V Based on spacecraft specifications. m p 10.8 kg Mass of propellant needed for a V of m/s, not including margin. V 27.5 cm 3 Confirmed by filling plenum completely with liquid water and massing before and after.

83 Chapter 7. Testing 71 To make the lifetime test more realistic to perform in an appropriate amount of time, the regulation frequency (f R ) and the mass flow rate (ṁ p ) can be scaled up to achieve a lifetimes worth of cycles and propellant quantity in a much shorter period of time than nominal operation would permit. This was done by iterating various downstream restrictor sizes until one was selected that gave the desired mass flow rate. Once this was determined a series of tests were performed at varying deadbands about the constant regulated pressure of 125 psi to observe the relationship between regulation frequency and deadband as seen in Figure 7.3. The associated trend allows for the test deadband to be calculated. Figure 7.3: The regulation frequency as a function of the selected regulated pressure deadband, determined experimentally. Using the number of regulation valve cycles per 120 second thrust (48 cycles), the required thrust duration can be calculated for this test. Table 7.2 summarizes the conditions that will be set to during the valve lifetime testing. It was determined that to achieve a mass flow rate of 165 mg/s, the downstream restrictor needed to be a 20 cm long, ID tube, which was verified with a flow meter. It is important to note that due to limitations on the software inputs the regulated pressure deadband was set to 7 psi (the input is in fact half of the actual deadband) and the test thrust duration was set to 20 seconds.

84 Chapter 7. Testing 72 Table 7.2: Test Parameters for Regulation and Thrust Valves Description Symbol Value Notes Regulation Test Duration t test 6 hours Test period. Test Regulation f Frequency R Desired regulation frequency to achieve half a valve 2.45 Hz lifetime in a standard work day. Valve Cycle Lifetimes to Propellant Mass Ratio Test Propellant Mass Flow Rate Test Regulated Pressure Deadband Number of Cycles per Thrust (Nominal) Test Thrust Duration Total Thrusts per Day z = NR m p 2 ṁ p = mp 2t testz p R n R = N R f R t thrust = nr f R 165 mg/s (7.3 SLM) 14.6 psi 48 cycles 19.6 sec N thrust = NR 2n R 1015 Number of valve lifetimes (N R ) to total propellant mass (m p ) ratio. Corresponds to flow rate needed to deplete 1 lifetime of propellant in 2 valve cycle lifetimes. Determined by plotting the experimental relationship between f R and p R (Figure 7.3). Based on the required amount of regulator valve cycles and the regulation frequency. Thrust time required to achieve correct number of cycles per thrust at the test regulation frequency. Number of thrusts to complete half a valve lifetime of cycles in a work day Methodology After the initial setup and test conditions had been determined, the system was placed in the standby configuration defined as follows: Supply tank regulator was set to upstream pressure, p tank. Electronics board was powered on. Both the regulator and thrust valves were closed but connected and ready to operate. Secondary volume was unpressurized and had equalized to ambient conditions. Secondary volume vent was closed. Isolation valve was fully open. Downstream plug was open. The test campaign began with an initial leak test on the brand new valves to observe their condition prior to testing. Following this, a thrust test was performed exposing the valves to half a spacecraft lifetime of cycles throughout a test period of 6 hours. Another leak test was performed overnight (18 hours) to observe if any leaks resulted from that day s thrust test. This was repeated until the desired number of lifetimes was performed or until a valve experienced significant leakage. Leak Testing 1. The leak test began by ensuring the e-board was turned on and was set to the maximum bus voltage and that the system was in the standby configuration. 2. Both valves were cycled on and off to reseat the valve poppets.

85 Chapter 7. Testing With the isolation valve still open, the regulator valve was commanded open to fill the plenum with CO The isolation valve was closed isolating the secondary and intermediate volumes. 5. The regulator valve was commanded closed to isolate the secondary volume from the supply tank. 6. The secondary volume vent was opened to release the pressure. 7. All pressure sensors and thermistors were checked to ensure they were reading ambient pressures and temperatures. 8. Both vents were closed with plugs. 9. The logging script was begun to log both pressure sensor and thermistor readings. Regulation Testing 1. The regulation test began by ensuring the e-board was turned on and was set to the maximum bus voltage and that the system was in the standby configuration. 2. The supply tank pressure was recorded periodically throughout the day since it decreased as the supply tank cooled from long duration thrusting. 3. The logging scripts were initiated. 4. The regulation algorithm was initiated with the desired test conditions. 5. The pressure regulation was continued until either: (a) Supply tank pressure had dropped below 700 psi, OR (b) One half-lifetime of cycles was met 6. A leak test was performed overnight. 7. The above steps were repeated until the desired number of lifetimes (5) was completed Test Results Table 7.3 summarizes the results from the lifetime test. Table 7.3: Valve Lifetime Test Results Summary One Lifetime Achieved Lifetimes Tested Regulator Valve Thrust Valve Propellant Usage 10.8 kg 36 kg 3.3 Both the regulator valve and the thrust valve achieved 5 lifetimes worth of cycles without experiencing any noticeable leaks. In addition, the valves were exposed to about 3.3 lifetimes worth of carbon dioxide which seemed to have no negative impact on the valve seal either. Knowing the internal volume of each

86 Chapter 7. Testing 74 section downstream of the valves and assuming constant temperature, the leak rate in mg/hr can be calculated from the pressure increases. The secondary and tertiary volume pressures during leak testing are shown in Figures 7.4 and 7.5, where an initial increase in pressure was observed before reaching steady-state constant pressure. Due to the fact that after several hours of operation during the day the entire system cooled down quite a bit, evident especially in the frosting of the tube downstream of the regulator valve (Figure 7.6). During the initial part of the leak test the system was heating back up to room temperature, causing a pressure increase. It was concluded that this did not represent a leak since after reaching room temperature, no further increase in pressure was observed. Figure 7.4: Pressure over time in the secondary volume for all leak tests. The frosting of the tube will not occur in space due to the lack of water vapour and also because it will never be in operation continually for this long. However because the valves have a lower operational temperature limit of 4 C it is important to monitor system temperatures to ensure they do not fall outside of their range Dry Actuation Test Based on the results from the primary valve lifetime test it was concluded that the valves would be able to successfully withstand at least 5 lifetimes worth of cycles, however it was desirable to quantify how far the valves could be pushed before noticeable damage would occur. Thus, a series of rapid open-close actuations were performed without propellant (in other words dry ). This was considered a worse case since the use of propellant typically acts to cushion the impact of the poppet on the valve body, reducing the rate of degradation. Again the test consisted of many segments of actuations with leak tests in between. Since no propellant was being used the actuations could be performed overnight as well. The on and off times of

87 Chapter 7. Testing 75 Figure 7.5: Pressure over time in the tertiary volume for all leak tests. Values were corrected to account for the uncalibrated pressure sensor. Figure 7.6: Due to continuous operation over several hours, frost began to form on the tube downstream of the regulator valve, where most of the expansion was occurring. each valve were set in software and were selected to allow for cycles per 24 hour period. This also included two leak tests, one part way and one at the end of the 24 hour period. Figures 7.7 and 7.8 show the leak test results at each stage. Although the thrust valve never displayed any signs of leakage despite experiencing over cycles, the regulator valve began to show a leak after cycles. This is shown in Leak Tests 20 and 21 in Figure 7.7, which were the last leak tests to be performed. The difference between these two leak tests was the supply pressure that was being exposed to the upstream side of the valves, where as expected the higher pressure (Leak Test 21) generated a larger leak. It is also important to note that in some cases, if a very small leak was present and some pressure built up on the downstream side of the valve, that back-pressure may actually aid in closing the valve and preventing further leakage. Thus, it is

88 Chapter 7. Testing 76 Figure 7.7: Pressure over time in the secondary volume for all leak tests after dry actuations. Figure 7.8: Pressure over time in the tertiary volume for all leak tests after dry actuations. Values were corrected to account for the uncalibrated pressure sensor. suspected that even though a leak developed after 11.7 lifetimes, the pressure build-up in the plenum would eventually stop the leak. It was concluded that despite the valves having a cycle limit of only half of one regulator valve mission lifetime, the valve was able to achieve almost 12 lifetimes of cycles before experiencing any noticeable leakage.

89 Chapter 7. Testing Thermal Test The solenoid valves are rated to operate between 4 C and 105 C which is acceptable on the upper limit but not sufficient for the lower limit that the spacecraft is expected to experience, with the range being -20 C to 60 C. Thus it was necessary to determine if the valves could operate at both temperature extremes, in each of the forward and reverse configurations, and at the minimum and maximum operating voltages. With the prototype feed system (with valves) in the thermal chamber and supplying input pressures of 1000, 1250, and 1400 psi all combinations of the above stated conditions were tested. In each case the valves successfully overcame the pressure and opened, showing that even outside of their specified operational temperatures, voltages, and pressures, the valves will still actuate. In addition to performing actuations at varying pressures and voltages at the temperature extremes, regulation tests were also performed where the valves were set to regulate (i.e., thrust) for 10 minute durations with an input supply pressure of 830 psi. The valves were able to regulate the flow to the desired setpoint pressure of 125 psi in all cases. However, at -20 C after the regulation test was complete a leak was observed through the regulator valve, manifesting as an increase in secondary volume pressure to about 230 psi. This was attributed to the fact that below the specified lower temperature limit, thermal expansion within the valve may be preventing the poppet from reseating itself in the sealed position. The leak was mitigated after manually actuating the valve and reseating it, which means that operational limitations need to be implemented on-orbit to either ensure the valves do not get colder than their lower limit of 4 C; or after each series of thrusts a manual actuation command is necessary to reseat the poppet. Despite the valve successfully actuating under all potential operating conditions, the observed leak at cold temperatures needs to be carefully monitored and may be able to be mitigated through a modification of the thermal design, although further investigation is required Acceptance Prior to integrating the solenoid valves into the flight feed system it was necessary to perform acceptance tests on the units to ensure that they were all operating normally and were sufficient candidates for the flight system. This test would examine the valves performance near the lower temperature limit (5 C), the operational upper limit (40 C), and the survival upper limit (60 C). At each of these temperatures the valves (in their respective configurations, forward and reverse) will attempt to actuate at the maximum (28.8 V), minimum (25 V), and below the minimum (18 V) spacecraft bus voltages when exposed to the maximum (1400 psi) and the minimum (200 psi) pressure limits. Table 7.4 summarizes these test parameters. Table 7.4: Valve Acceptance Test Parameters Test Limits Temperature 60 C, 40 C, and 5 C Voltage 28.8 V, 25 V, and 18 V Pressure 1400 psi and 200 psi

90 Chapter 7. Testing Test Setup The test setup for the solenoid valve acceptance test was quite simple, a nitrogen supply tank with a filter to prevent debris from contaminating the valves was connected to a tee. The regulator and thrust valve were set up inside a thermal chamber and connected to the tee with the regulator valve being in the forward and the thrust valve being in the reverse configurations. The outputs of the valves were routed outside of the chamber for the operator to verify if the valves had in fact actuated and thus allowed gas through them. The electronics board that was controlling the valves was also inside the chamber to ensure that its temperature was equivalent to that of the valves. Figure 7.9 shows the setup of the valves inside the thermal chamber. Figure 7.9: Valve acceptance test setup inside a thermal chamber Methodology Once the valves and e-board had been set up inside the thermal chamber, the following steps were performed: 1. The chamber temperature was set to 60 C and using telemetry from the valve thermistors the valves were allowed to reach steady-state temperature. 2. First, a leak test was performed by setting the supply tank regulator to 850 psi to pressurize the section upstream of the valves. The supply tank was then closed off and the output from the digital pressure gauge was monitored to identify any drop in pressure that could indicate a leak through the valves. 3. Assuming no leaks, the pressure was set to 1400 psi. 4. The supply voltage to the e-board was set to 28.8 V and each valve was commanded open ten times consecutively while the output tubes were monitored to ensure the valves were actually opening. 5. Step 4 was repeated at 25 V and 18 V.

91 Chapter 7. Testing 79 If the valves were unsuccessful in opening at 18 V, the voltage was increased incrementally until a voltage was found that did actuate both valves. This voltage was recorded. 6. Once all voltages were successfully tested on both valves, the pressure was set to 200 psi. 7. Steps 4 and 5 were repeated at this pressure. 8. After both pressures and all voltages were tested at 60 C, the temperature of the chamber was set to 40 C and again the valves were allowed to reach steady-state conditions. 9. Steps 3-7 were repeated at 40 C. 10. A leak test was then performed after the successful actuations at 40 C. 11. The chamber temperature was then set to 5 C and the valves reached steady-state. 12. Steps 3-7 were repeated. 13. A final leak test was performed at the end of the acceptance test to ensure both valves were still operating as expected Results The above procedure was performed on the flight and spare valves in two separate tests and in both tests 87% of the scenarios resulted in the valves successfully actuating ten consecutive times at the specified temperatures, pressures, and voltages. The only failures to actuate occurred at 60 C and 40 C, and at 1400 psi, when both the flight and spare thrust valves were unable to actuate at 18 V. This was expected since the higher temperatures were increasing the resistance which caused a lower power output to the solenoid. At such a low voltage the power was insufficient to overcome the high pressure acting against the solenoid armature forcing the poppet closed. For the flight thrust valve, 20 V was required at both 60 C and 40 C, whereas for the spare thrust valve, 23 V was required at 60 C and 24 V at 40 C. Despite this it is extremely rare that the thrust valves will experience such high pressures, in addition the minimum voltage requirement of 25 V was still maintained. Thus, all valves were concluded to be sufficient for use in the flight system Additional Valve Testing Unit-level Vibration Four solenoid valves were exposed to qualification level vibration loads including sinusoidal burst, random vibration, and shock profiles. Figure 7.10 shows the setup of the valves mounted to the vibration slip table, where two of the valves were configured with the Pressure Under Seat (PUS) (regulator valve orientation) and two were configured with the Pressure Over Seat (POS) (thrust valve orientation). After performing tests in all three axes with an upstream pressure of 1400 psi exposed to the valves, the only times that propellant leaked through were during the random vibration tests of the x- and z-axes for the valves in the PUS configuration. It was noticed that the leak continued even after the vibrational loads were removed and was only stopped when the valves were manually actuated causing the poppet to reseat itself into the correct sealing position (as was seen during the valve thermal testing). The important fact to note is that at no point did any pressure leak through the POS valves (i.e., thrust

92 Chapter 7. Testing 80 valves), thus providing confidence that no propellant will be lost during launch, even if the regulator valve leaks slightly into the secondary volume. In addition, actuations performed before, in between axes, and after the test at spacecraft minimum voltage levels (25 V) indicated that the vibrations did not damage the valve from an operational standpoint either. Figure 7.10: Unit-level valve vibration test setup Hydrostatic Testing Although the operational pressure limit of the solenoid valves was 1250 psi, the maximum expected operating pressure (MEOP) was 1400 psi meaning there is a good chance that the valves, specifically the regulator valve will be exposed to a pressure higher than its operational limit. Thus a hydrostatic pressure test was performed to observe the cracking limit of the valve poppet (the point at which the upstream pressure overcomes the spring s closing force). Also the valves have a specified proof pressure of 1875 psi, thus the test was meant to confirm if they could withstand up to 2x MEOP in the POS configuration. Table 7.5 summarizes the test results.

93 Chapter 7. Testing 81 Item Solenoid Valves Table 7.5: Hydrostatic Pressure Test Results Number of Pressure Tested Units Rating (psi) Pressure (psi) Result (operational) 2200 Pass (proof) 2800 Pass 7.2 Performance Characterization To verify the CRT s performance against the expected results a series of thrust tests were performed. aiming to quantify the thrust output and specific impulse of the redesigned thruster. The test setup and procedure were nearly identical to that of the RMT performance testing, described in detail in Section Test Setup A vacuum chamber was used to simulate the space environment, which is important because the interactions between exhaust particles from the thruster and ambient air molecules will skew the data. A microbalance was used for measuring thrust, on which the thruster and thrust stand were mounted. Exhaust and internal heater cartridge temperatures were monitored. Temperatures, mass flow rate, and force (mass) were logged to evaluate the thruster s performance. Figure 7.11 shows the setup described above. In addition the feed system was setup outside of the vacuum chamber to deliver the required propellant to the thruster through the feed. Figure 7.11: CRT performance test setup in the vacuum chamber Methodology The following procedure was used to perform the thrust tests: 1. The vacuum chamber was pumped down to about Torr.

94 Chapter 7. Testing The thruster was then preheated for the nominal preheat duration. 3. The regulation algorithm was initiated at the desired setpoint pressure and deadband to allow CO 2 to flow towards the thruster at the required flow rate. 4. The system was allowed to reach steady-state conditions. 5. The regulation algorithm was stopped after the desired thrust duration Test Parameters and Results The parameters used in the feed system software to operate the thruster were a result of iterations through a variety of tests. The regulation pressure setpoint in the plenum that the algorithm tries to maintain was initially 50 psi. However after the mechanical regulator was removed from the feed system in favour of higher pressure valves the setpoint was increased to 125 psi. The valves have a minimum duration for which they can open and close and despite the duration being short, because the regulator valve is exposed to the vapour pressure of CO 2 (830 psi at room temperature) it can be quite difficult to limit the ingress of propellant into the secondary volume. This makes it challenging to maintain a low pressure setpoint. In addition the pressure sensors selected for the system have non-linearities at pressures approaching ambient conditions (14.7 psia) as seen in Figure Since the calibration model in the software is treated as linear, the errors get slightly larger at low pressures. By regulating the pressure in the plenum to a setpoint above the non-linear region (above 90 psi), it can be assured that the pressure being distributed to the thruster is accurate. As identified earlier the high pressure solenoid valves selected for the feed system have a low cycle lifetime specification, thus it was necessary to minimize the cycle frequency wherever possible. From a fluid perspective this was achieved by increasing the size of the secondary volume where the regulator valve needs to stay open longer to allow an equivalent amount of pressure in compared to a smaller volume. Also, adding a restrictor upstream of the regulator valve limits the amount of flow that can get to the valve further increasing the time needed to fill the secondary volume to the setpoint pressure. In conjunction with these methods, increasing the deadband on the setpoint pressure can further increase the cycle time of the valve (and subsequently reduce the cycle frequency) as shown graphically in Figure The trade-off when using a larger deadband is that there is also a higher variability in the thrust output. However since the majority of thrusting will be for deorbiting and no fine-pointing or precise maneuvers are required it was determined that the variability was acceptable since the average thrust output is the more important value for determining the deorbit V. The test parameters and the subsequent results for the flight thruster testing have been summarized in Table 7.6 below and it can be seen that the thruster performed very close to expected. Figure 7.14 shows the heater and exhaust temperature profiles for the performance test, where there is a 10 minute preheat of the heater followed by a 10 minute thrust period. Figure 7.15 shows the thrust output from that 10 minute duration, clearly exemplifying the thrust variation described above.

95 Chapter 7. Testing 83 Figure 7.12: A preliminary pressure sensor calibration that identifies the non-linear section below 90 psi. Table 7.6: Flight CRT Performance Parameters and Results Parameter Flight CRT Heater Setpoint Temperature 800 C Preheat Duration 10 min Regulation Setpoint Pressure 125 psi Regulation Deadband 50 psi Steady-State Exhaust Temperature 660 C Mass Flow Rate 47 mg/s Thrust 53 mn Specific Impulse 115 s Thrust Variability 10% 7.3 System Qualification A qualification model (QM) propulsion system was assembled to provide a platform in which vibration testing could be performed without compromising a fully assembled satellite. This system was nearly identical to the final flight system with only a few subtle modifications being made as a result of the vibration test described below.

96 Chapter 7. Testing 84 Figure 7.13: An example of the pressure regulation algorithm controlling the pressure of the plenum at a setpoint of 125 psi with two different deadbands Tank Hydrostatic Test Prior to integrating the propellant tank into the propulsion system it was necessary to clean the inside and remove any potential debris remnants left over from the fabrication process. This procedure, performed at an external facility, involved a wash of the inside of the tank with a slightly alkaline solution, followed by a series of rinses with de-ionized (DI) water until all debris was removed, and then a final rinse with iso-propyl alcohol (IPA) before being blow-dried with filtered air. Examining pull tests from the tank before and after indicated that the cleaning process succeeded in removing essentially all of the debris. However, there was a concern that the procedure may have compromised the resin holding the composite fibers of the tank together. Thus, a hydrostatic test was performed on the tank to ensure that it could withstand pressure before exposing it to vibrational loads. A hand pump was used to inject DI water into the QM tank at a rate of 260 psi/minute until the pressure increased to 1570 psig, at which point it was held for over 15 minutes. This was selected since it is above MEOP (1400 psi) but below the proof pressure (1750 psi) ensuring that the tank can withstand the test pressure. During the test no loss in pressure was observed and there was no visible leakage thus confirming that the cleaning procedure did not compromise the tank s strength System-level Vibration Test The purpose of performing a vibration test on the QM propulsion system was primarily to qualify the propellant tank to the levels specified by the launch vehicle provider and to provide confidence that the

97 Chapter 7. Testing 85 Figure 7.14: Typical CRT performance test heater and exhaust temperature profiles outlining the preheat period (0-600 s) and the thrust period ( s). tank will not get damaged during launch. In addition, this was the perfect opportunity to also qualify the rest of the propulsion system including the feed system and the fill/drain module to those same levels Test Setup The QM propulsion system was fully assembled on the vibration fixture plate using the same bolt patterns that would be used on the internal panel of the spacecraft. All screws were torqued to the required specification and secured with RTV silicone to prevent loosening. The fitting nuts were also secured with RTV for the same reason. The screws of most importance, mainly on the tank brackets, were marked with a line to clearly indicate any rotation and is something that was monitored throughout the test. It was necessary to match the mass of the propellant as closely as possible to what it would be for the actual spacecraft, however due to safety concerns and timeliness it was not desirable to perform a complete fill with liquid CO 2 as it would be for flight. Thus DI water was used as the bulk of the propellant mass. Since the main purpose of the test was to examine the tank s capabilities to withstand launch loads while pressurized it was also necessary to pressurize the DI water to MEOP. Thus a small amount of nitrogen (N 2 ) was injected to increase the pressure. This translated to 10 kg of DI water and 0.8 kg of nitrogen, resulting in the tank having a total propellant mass of 10.8 kg and a pressure of about

98 Chapter 7. Testing 86 Figure 7.15: Typical CRT performance test thrust output psi. The primary difference between this setup and what it will be for flight is the sloshing that will occur by having a partially filled tank with liquid water. CO 2 will be in either a gas/liquid mixture or a supercritical state but in both cases the behaviour is different than what will be tested during the vibration test, however it was deemed that sloshing is a worse case scenario. The fixture was then secured to the vibration slip table (or head expander for the z-axis) as seen in Figure Three-axis accelerometers were fixed to the tank bracket, feed plate, and fill module to monitor the response of each component, as well as a control accelerometer directly attached to the fixture to monitor the control input Methodology The vibration test was split up into three segments, one for each of the principal axes (x, y, and z). In each axis sinusoidal burst, extended sinusoidal, random vibration, and shock profiles were to be performed. Before and after each of these profiles a low-level sinusoidal (LLS) test and a long form functional test (LFFT) was performed. They were meant to monitor the structural and electronic status of the system respectively throughout the vibration test. The goal of the LLS test was to examine the natural frequencies of the test article before and after each profile to allow for comparison at various stages of the vibration test. Any change in natural frequency or amplification could indicate damage of the components or loosening of screws. The amplification factor was defined as the measured acceleration divided by the input acceleration. The LFFT monitored the pressure inside the tank and obtained telemetry from the sensors thus ensuring all of the

99 Chapter 7. Testing 87 Figure 7.16: QM propulsion system mounted to the vibration table in all three axes: y-axis (upper left), x-axis (upper right), z-axis (lower). electronics were still functional and that the e-board was communicating as expected. A sinusoidal burst or quasi-static acceleration test was performed to evaluate the strength of the system and ensure that nothing will brake when under static loads at relatively high accelerations. The extended sinusoidal test was similar to the sinusoidal burst except it was performed at a lower acceleration and for a much longer duration. This is not a typical profile that is performed on the spacecraft-level but was required for this mission by the launch vehicle provider. The random vibration test is usually the most vigorous and is the one that best represents the actual launch environment. The DAUNTLESS bus uses a pyrotechnic separation system and thus the vibration test must also verify that the tank and propulsion system can withstand this sudden shock. Table 7.7 summarizes all of the test parameters for each of the profiles that were used based on GEVS classification Results The first axis that was tested was the y-axis, being the tank axial direction. After the y-axis random vibration test, an LLS test was performed which revealed an increase in the amplification factor of the tank brackets from 35 times in previous tests to 75 times amplification. After further examination the

100 Chapter 7. Testing Test Low-Level Sine Sine Burst Extended Sine Random Shock 88 Table 7.7: QM Propulsion System Vibration Test Levels Frequency Acceleration Duration Description Range (Hz) (g) (sec) Track changes in natural frequency and amplification Quasi-static acceleration (strength) test Quasi-static acceleration, long duration 17.7 (x,y) Simulates complex launch (z) environment 75 (x,y) 1 ms (x,y) Simulates pyrotechnic events 55 (z) 0.8 ms (z) M4 screws mounting the tank to the tank brackets as well as the M8 screws mounting the tank brackets to the vibration fixture plate (as indicated by Figure 7.17) had all lost their pre-loads (meaning they were loose). However, the RTV remained intact and the markings were still aligned, thus confirming that the loosening was not caused by the screws backing out but rather by degradation of the interface between components. This was discovered after the system was disassembled upon the completion of the test. Examples of this degradation are evident in Figures 7.18 and It was expected that this was caused by the pre-load (i.e., torque) specifications not being sufficient to keep the components together. The interfaces were rubbing and material was getting worn away which further reduced the pre-load causing even more degradation. Figure 7.17: Tank bracket screws that experienced a loss of pre-load. This occurred on both sides of the tank. During the test it was decided that the cause of the pre-load loss and the resulting amplification increase was a structural issue not related to the tank or propulsion system components directly. Thus the testing was continued. The screws were tightened as necessary and the random vibration profile levels were modified to prevent the unrealistically high amplifications of the tank bracket around the natural frequency. This was achieved by introducing notching to the control input. Using feedback from the

101 Chapter 7. Testing 89 Figure 7.18: Degradation of the tank bracket against the vibration fixture plate. Figure 7.19: Degradation of the end bosses of the tank and the shims against the tank bracket. tank bracket measurement accelerometer the control input was limited to ensure that the power spectral density (PSD) of the tank bracket stayed below a specified value. Based on the highest levels observed during previous DAUNTLESS bus structural vibration tests, the maximum was set to 3.5 g 2 /Hz. The following images (Figures 7.20 and 7.21) clearly differentiate a control input that is not being notched

102 Chapter 7. Testing 90 compared to one that is. In addition a summary of the measured first natural frequencies of each module has been summarized in Table 7.8. Figure 7.20: Random vibration profile: control input, no notching. Table 7.8: QM Propulsion System First Natural Frequency (FNF) Y-Axis X-Axis Z-Axis Tank Bracket Fill Module Feed Plate It is also important to note that no significant leaks were observed from the tank as the pressure monitored throughout the test stayed constant ±3 psi. This proves that the tank itself can withstand the launch vibration loads without experiencing any damage. However there was a leak observed through the regulator valve into the secondary volume. After the random vibration in the z-axis, the secondary volume pressure increased from near ambient to 450 psi. This phenomenon was also observed during the unit-level valve vibration test (Section ) and was caused by the vibrations occurring in the same axis that the valve poppet opens. Despite pressure building up in the secondary volume, no propellant escaped through the thrust valve, likely because the upstream pressure was aiding in keeping the thrust valve closed. To conclude, the levels that were performed for the vibration test of the QM propulsion system were very conservative being much higher than anything that will be experienced during launch and despite the structural issues that were observed in the tank brackets, the tank itself survived with no issues.

103 Chapter 7. Testing 91 Figure 7.21: Random vibration profile: control input, with notching. To combat the structural concerns, wedge-locking washers and higher torque specifications were used to maintain the pre-load in the screws and to prevent degradation of the interfaces. This solution was confirmed to be successful after the complete spacecraft-level vibration test was performed without any issues. 7.4 Flight Assembly Leak Testing Leak testing of the flight propulsion system was necessary to ensure that the leak rate requirement was met. The requirement of 5 mg/hr was calculated based on the amount of propellant that could be lost throughout the mission while still having enough left-over to perform the deorbit maneuvers at the end of life (EoL). To verify the leak rate of the system, two different methods were used each with differing advantages. The first measured the pressure drop over time of a small test article pressurized with nitrogen. This could then be used to calculate an estimated leak rate and was more useful for subassemblies. The second method involved the use of a pump system and mass spectrometer to detect helium that could be leaking into or out of the test article (depending on the setup configuration) and was used on both the subassembly level (to verify earlier testing) and the system-level (with and without the tank) Subassembly Nitrogen Leak Testing This method was more useful for subassemblies on a smaller scale, where the test article was pressurized with nitrogen gas and the pressure was monitored over a long duration. The advantage of this was that it was extremely simple to set up and could be left logging overnight, which meant it could be performed

104 Chapter 7. Testing 92 more frequently. By performing these checks more often, any leaks in the subassemblies could be dealt with early on in the integration process, making it easier to troubleshoot. Approximating the volume of the test article and knowing the rate of pressure drop, one can estimate the leak rate in mg/hr Test Setup The test setup for the nitrogen leak test included a nitrogen supply tank with a pressure regulator, filter, and purge port attached. This fed to an isolation hand valve and pressure sensor, which was connected to a spare e-board with a laptop providing the interface. Following the pressure sensor was a fitting that could be configured to accommodate any potential connection for the various subassemblies. Figure 7.22 shows the test setup. Prior to any leak testing this setup was plugged and pressurized to ensure that there were no significant leaks within it seeing as it was used for all the future leak tests. Furthermore, a bias could be determined that was subtracted from the subassembly leak results. It was determined that the leak rate of the test setup was pretty negligible, calculated to be approximately 0.03 mg/hr. Figure 7.22: Nitrogen subassembly leak test setup Methodology This procedure was used to perform leak tests on subassemblies: 1. The pressure sensor assembly was connected to the test article using the appropriate fitting. 2. All other open ports were plugged and all fittings were securely tightened. 3. With the isolation valve open, the test article was pressurized to about 850 psi. 4. After ensuring that the pressure had stabilized in the test article, the isolation valve was closed. 5. The pressure log was initiated and the system was left logging overnight.

105 Chapter 7. Testing Results Initially many of the subassemblies had preliminary leaks that were unacceptable, prompting the use of IPA injected over certain connections to observe any bubbles that would form as a result of the leak. Once identified the fitting in question was re-tightened and the test was repeated until a sufficiently low leak rate was calculated. Table 7.9 summarizes the estimated leak rates of the various subassemblies and provides a total expected leak rate for the system. Table 7.9: Nitrogen Subassembly Leak Test Results Pressure Drop Estimated Test Test Article Rate (psi/hr) Article Volume (cc) Secondary Volume Leak Rate (mg/hr) Subassembly Restrictor Subassembly Fill Subassembly Vent Subassembly Total Leak Rate (mg/hr) 4.08 As can be seen the cumulative leak rate for these subassemblies is under the 5 mg/hr maximum requirement, however due to the nature of the test setup not all of the connections were able to be tested using this method, including the solenoid valve and tank fittings. These connections were tested after the complete propulsion system was assembled during the helium leak testing. In addition, the results calculated in Table 7.9 were very conservative, specifically because a factor was applied to the volume estimations for each subassembly. For a given pressure drop rate, having a larger volume means more mass is flowing out through leaks per unit time compared to a smaller volume. Thus to be conservative, since the volumes were uncertain, they were estimated on the large end. In addition nitrogen has a smaller molecule size than CO 2 which would allow it to fit more easily through smaller leaks, making it more of a worse case. Overall despite being overly conservative in many respects the calculated leak rate was still below the requirement Helium System-level Leak Testing The other leak test method using a helium (He) detector had the advantage of being much more accurate. The unit was composed of a roughing/turbo pump combination to pull vacuum on the test article and by injecting helium around the connections, the detector s internal mass spectrometer would be able to detect and quantify the rate of helium ingress through leaks. This method is often referred to as outside-in leak detection mode. Alternatively the test article can be pressurized with helium and placed inside a vacuum chamber, the pumps in the detector can then be used to pull vacuum on the chamber around the test article, referred to as inside-out detection mode. Any helium leaking out will be pulled in by the device and detected. It was necessary to use the latter technique when the propellant tank was finally integrated since it was not recommended to pull vacuum on the tank itself. Figures 7.23 and 7.24 show the test setup for both the outside-in and inside-out helium leak testing respectively.

106 Chapter 7. Testing 94 Figure 7.23: Outside-in helium leak testing setup. Figure 7.24: Inside-out helium leak testing setup. The test article was placed inside the vacuum chamber and pressurized with helium Results The leak rate is defined as the pressure increase/decrease in a specific volume per unit time and is displayed in units of T orrl s [24]. To comprehend this leak rate in a quantity that is relevant to the requirement it is necessary to convert this to mg/hr of CO 2. One must take into account the differences in physical properties of helium compared to carbon dioxide as well as the leak test pressures versus the actual operational pressures. In addition, different equations are used whether the leak is considered to be in the laminar or molecular flow regime, where the worse case prediction was used for the leak rate calculations. Table 7.10 summarizes the calculated leak rates for the entire system with and without the tank. The operational condition that was used for comparison was when the propulsion system, filled with CO 2, has a pressure of 1050 psi (corresponding to maximum storage density and 30 C). When the test condition is not the same as the true operating condition a correction factor must be applied to the