Thruster

Model Philosophy

Figure 1 depicts an overview of the thruster model philosophy within each phase of the LISSA project.

 Condust Research

Since the Iris did not have a thruster subsystem, there was no heritage for LISSA in terms of regrading the thruster from Iris to use as a breadboard model. Hence, it was necessary to conduct research in order to create a simulation model for the design of the thruster subsystem. The research was done in phase A and early stage of phase B.

Simulation Model

The simulation Model of the LISSA thruster subsystem is developed based on the research concept from phase A. This model is built in the middle of phase B and will be updated as needed in phase C.

CAD Model

Based on the Simulation model was conducted on phase B, The CAD model was build in phase C.

Engineering Model (EM)

Based on the CAD model developed during Phase C, the Engineering model of the thruster is scheduled to be built early in Phase D to facilitate testing.

 Flight Model (FM)

Finally, at the end of phase D, the FM model  will be built based on EM.

Mission and Systems Level Requirements Impact Analysis

The LISSA will be built upon the original Iris design, with the addition of a thruster subsystem to enable orbital maneuvers. This design and analysis are guided by the following mission and system-level requirements.

• R-LIB-TRS-001: The thruster tank shall conform to the  1 U form factor of the bus.

• R-LIB-TRS-002: The thruster subsystem shall interface with the AODCS module over UART.

• R-LIB-TRS-005: The thruster unit shall have mass no more than 1560 g.

• R-LIB-TRS-008: Thruster shall support a two-stage activation sequence consisting of arming and firing.

For more information about requirements, take a look at the valispace registry.

Thruster Design

Propellant Selection

The choice of propellant is fundamental to the design and performance of cold gas thruster systems. It influences not only thrust and specific impulse, but also storage conditions, handling procedures, safety requirements, and overall system integration. In small satellite missions like LISSA—especially those launched as secondary payloads—propellant selection must carefully balance performance, safety, availability, and regulatory compliance.

Cold gas thruster performance is closely tied to the selected propellant. The maximum achievable thrust is governed by the nozzle size and upstream stagnation pressure, while the specific impulse is largely determined by the propellant’s molecular mass and thermodynamic properties [6]. Propellants with low molecular mass, such as hydrogen or helium, can deliver specific impulses up to approximately 300 seconds. However, their low storage density [7] necessitates large tanks, increasing structural mass—often an unacceptable trade-off for CubeSat scale missions. Additionally, light gases are more prone to leakage through small gaps or seals, which can compromise mission longevity. Therefore, while high performance is desirable, practical design factors such as storage volume, material compatibility, leakage risk, and system simplicity must also be taken into account.

Over the years, a wide variety of propellants have been used in cold gas missions, each offering different trade-offs. Noble gases like argon used in the POPSAT-HIP1 satellite[8], krypton used in the Adelis-SAMSON mission[9], and xenon used in the MEPSI and PROCYON missions[10] provide relatively high specific impulse but are often costly and limited in availability. Helium was employed as a backup propellant in the NASA MESSENGER science mission to Mercury [11]. However, its low density and high leakage potential limit its practicality for small satellites. Heavier gases such as sulfur dioxide and sulfur hexafluoride have been used in the CNAPS system on the CanX missions [12 -13]. Notably, sulfur hexafluoride poses serious environmental concerns, with a global warming potential approximately 23,900 times greater than that of carbon dioxide [14]. Flammable gases like methane and butane , are not suitable for student-built platforms such as LISSA due to safety and handling risks. Refrigerants such as R-134a and R-236FA offer promising storage and controllability characteristics but are subject to sourcing issues and environmental restrictions.

Given LISSA’s volume, mass, and safety constraints, carbon dioxide (CO₂) was selected as the most appropriate propellant. CO₂ is non-toxic, chemically stable, and has favorable thermodynamic properties. It remains gaseous over a wide range of temperatures, eliminating the need for heating systems and simplifying thermal management. Its relatively high storage density supports compact tank design, and its widespread availability and regulatory acceptance make it well-suited for educational CubeSat missions like LISSA.

Pressure Strategy and System Simplification 

In the initial design phase, a pressure regulator was considered to allow high-pressure storage and lower regulated delivery downstream. While this configuration can enhance flexibility and safety, it was ultimately deemed incompatible with the LISSA thruster’s strict volume and mass constraints. Pressure regulators are typically bulky, and including one would have reduced the space available for the propellant tank, thereby limiting the amount of fuel stored and diminishing mission performance.

Furthermore, many small satellites are launched as secondary payloads on rideshare missions alongside larger, more valuable spacecraft. As a result, stringent constraints may be imposed on allowable propellants and storage conditions. For instance, earlier CubeSat standards limited maximum storage pressure to 120 kPa (1.2 atm) [15]. launching LISSA from the United States introduced cross-border compliance requirements with U.S. and Canadian regulations. According to the SpaceX Rideshare Payload User’s Guide, a system is considered a pressure vessel if it stores more than 20,000 J of energy or exceeds 100 psid MEOP [16]. The guide strongly recommends using pressure vessels certified by the U.S. DOT. If custom vessels are used, their design, fabrication, and verification must comply with the standards of the American Institute of Aeronautics and Astronautics (AIAA), based on the classification and material of the vessel [16]. U.S. DOT classifies compressed gases exceeding 200 kPa gauge pressure (29.0 psig or 43.8 psia) as Division 2.2 hazardous materials [17], while Canadian TDG regulations apply a threshold of 280 kPa absolute [18]. Exceeding these limits would require DOT-certified containers and full compliance with ASME BPVC design codes. Complying with these standards imposes stringent design requirements on pressure vessels. For instance, in the case of propellant storage, adherence to DOT and ASME regulations often necessitates a higher design factor, resulting in increased wall thickness. This, in turn, reduces the available internal volume for storing propellant. Additionally, the added structural reinforcement leads to an increase in mass, which exceeds the LISSA thruster's allocated mass and volume budgets. These constraints further limit the feasibility of incorporating a regulated high-pressure storage system within the 1U propulsion volume.

To avoid these complications, the propulsion system was redesigned to operate at a maximum absolute pressure of 2.67 atm, remaining safely below regulatory thresholds from the U.S. DOT, Canadian TDG, and SpaceX [16–18]. This allowed the system to bypass classification as a hazardous pressure vessel, eliminating the need for additional certification and enabling easier transport, testing, and integration. Eliminating the pressure regulator further simplified the architecture, freeing up volume for a larger tank and increasing the propellant mass. In the final configuration, the CO₂ flows directly from the tank through a solenoid valve into the nozzle, where it expands to produce thrust. This streamlined setup reduces the number of components, lowers failure risk, and enhances reliability—critical advantages for a CubeSat-scale mission. A schematic of the updated thruster configuration is shown in Figure 2, and performance metrics are discussed in the analyses section.

Solenoid valve

The solenoid valve regulates the flow of CO₂ from the storage tank to the nozzle, operating in response to electrical signals from the control board. To conserve the spacecraft's limited power resources, the valve is selected for low power consumption. For the LI Bus, an IEP solenoid valve   from The Lee Company has been chosen. This normally closed valve supports an operating pressure range of 0–800 psi, ensuring that in the event of power loss, it remains closed to prevent unintended propellant release. Its fast response time of 0.5 ms enables precise flow control, with actuation commands issued at 0.5 ms intervals. A diagram of the valve and a summary of its compliance with system requirements are provided below.

Pressurized Storage Tank

The fuel tank is designed to store the maximum allowable quantity of propellant and represents the largest component of the thruster system. To comply with the design pressure limit of 2.67 atm absolute, multiple off-the-shelf cartridges are stacked to construct the storage tank. This modular approach enables efficient integration of the tank and propulsion components within a dedicated 1U volume. The tank includes three output ports for direct connection to the solenoid valve, fill/drain (FD) valve, and pressure sensor, reducing the need for additional tubing and simplifying the overall assembly. It is securely mounted to the structural plate using a support frame and base plate, ensuring stability during launch and operation. Figure below shows the tank design and mounting configuration along with its conformation to 1 U (Compliance with requirement R-LIB-TRS-001).

Fill and Drain Valve

The fill and drain valve(FDV) is used for loading propellant. Since the propellant loading is conducted after the entire satellite has been integrated, it was crucial to allocate the FDV so that it was easily accessible. For this reason the FDV extended via a fitting to the satellite's faces. The FDV is selected from LEE company. Because a separate pilot tool is used to open the valve prior to launch, its in-flight configuration is extremely small and light weigh. The high pressure valve, capable of up to 4,000 psia operating pressure, features a low weight of less than 55 grams. the pilot tool can be used to mechanically open the valve to drain the system .A diagram of the FD valve and the requirements compliance for this component can be found below. 

Thruster controller Board

The primary function of the thruster controller board is to serve as the command and control interface between the Attitude and Orbit Determination and Control System (AODCS) and the thruster unit. It enables the execution of thrust commands by controlling the solenoid valve that regulates the flow of propellant. To ensure reliability and prevent unintended valve activation, the board employs an armed-fire strategy. In this approach, after receiving a valid command signal from the AODCS, the controller first enters an "armed" state, confirming system readiness and validating the command. Only then does it proceed to the "fire" stage, in which the solenoid valve is activated (Compliance with requirement R-LIB-TRS-008). The valve is maintained in the open state for a predetermined duration, allowing a controlled release of propellant, and then deactivated to terminate the thrust event. This two-stage strategy enhances operational safety and ensures that the valve cannot be triggered accidentally by electrical noise or spurious signals.

Beyond valve actuation, the controller board also fulfills several additional roles. It provides electrical interfaces for onboard sensors and actuators, including a pressure transducer to monitor tank pressure, a thermistor to track temperature near the propellant line, and connections for the reaction wheels to support attitude control. These interfaces enable real-time health monitoring and system coordination.

figure 10 illustrates the PCB layout of the controller board. Further technical specifications  data protocols used for communication with other subsystems, are detailed in the Interface Control Document (ICD) available on the ICD page.

Pressure transducer

The pressure transducer monitors the propellant pressure within the system, measuring it in the tank and sending the data to the microcontroller. This ensures the amount of propellant is available for thrust. By preventing over-pressurization and under-pressurization, the transducer safeguards the system, ensuring the cold gas thruster operates safely and efficiently. In the event of a transducer failure, a software algorithm estimates the residual pressure based on the consumed mass from the previous burn and the remaining mass within a constant volume. The flowchart illustrating this method is shown in Figure 12.

LI Bus will use a MSP300 a miniature pressure transducer from the ETE Connectivity company . A figure of the PT and the requirements compliance for this component can be found below. 

Nozzle

The primary nozzle concept for the propulsion system was based on a convergent-divergent (CD) geometry. In this configuration, subsonic gas enters the converging section, where the cross-sectional area decreases until it reaches the throat—the point of minimum area. At the throat, the gas reaches sonic speed (Mach =1). Beyond this point, the flow enters the divergent section, where the increasing area allows the gas to expand and accelerate into the supersonic regime, increasing the exhaust velocity and thereby enhancing thrust output [20].

The amount of thrust produced by the nozzle depends on the mass flow rate, the exit velocity of the exhaust, and the pressure differential at the nozzle exit. All of these parameters are directly influenced by the nozzle geometry [21]. As such, the nozzle design must carefully balance performance optimization, propellant efficiency, and manufacturing feasibility, especially within the strict volume and fabrication constraints of a CubeSat platform.

The minimum feasible throat diameter was selected as 0.5 mm, based on the limitations of metal additive manufacturing. This diameter represents the smallest reliably manufacturable hole while preserving structural integrity. In addition to fabrication considerations, the throat diameter also plays a critical role in setting the mass flow rate, as this is directly proportional to the throat area [21]. In CubeSat-scale missions where the onboard propellant supply is limited, a smaller throat enables controlled and gradual propellant release, allowing multiple short-duration thrust events rather than depleting the tank in a single high-impulse burst.

The ideal nozzle geometry is one that minimizes mass flow rate while maximizing exit velocity, as this combination leads to higher specific impulse and effective thrust. The exit velocity V_e is a function of the exit Mach number, which in turn depends on the area ratio A_e/A* , where A_ e is the exit area and A*  is the throat area. To evaluate this relationship, a detailed parametric simulation was conducted, as presented in the Analyses section. The results demonstrate that exit velocity increases with area ratio, reaching its highest value at an extremely large (theoretically infinite) exit diameter. This finding supports that a nozzle with an infinitely large expansion section would yield the maximum achievable exhaust velocity. While the simulation suggests that an abrupt expansion to infinity would provide optimal performance, such a geometry is not practical. A sharp transition from the throat to a large exit area can induce shock waves, flow separation, and instability within the nozzle. These adverse effects degrade performance, reduce thrust, and may compromise system reliability. To mitigate these issues, the final nozzle design incorporates a gentle chamfer immediately after the throat instead of a sudden expansion. This smooth, gradual transition promotes more stable supersonic flow, minimizes the likelihood of shock formation, and maintains nozzle efficiency, all while remaining compatible with the constraints of metal manufacturing and CubeSat integration. Figure 14 shows the CAD model of the nozzle. 

Requirement R-LIB-TRS-006 Compliance

Most components in the system are made of metal, which exhibit negligible outgassing and are inherently compliant with outgassing requirements. In the solenoid valve, the default sealing material is EPDM, which may be replaced with FFKM (Perfluoroelastomer) to further reduce outgassing. As shown in the table below, FFKM has a Total Mass Loss (TML) of less than 1%, meeting the required threshold.

In the fill and drain valve, the total mass of EPDM used is only 0.0024 g, making both its TML and Collected Volatile Condensable Material (CVCM) values negligible. The table below presents the TML and CVCM values for EPDM, confirming that its limited usage still complies with relevant outgassing requirements.

Additionally, PTFE tape—used for thread sealing—has been independently verified to have a TML less than 1%, as illustrated in the referenced figure.

All data are sourced from the NASA Outgassing Database, and collectively demonstrate that the materials used in each component are compliant with Requirement R-LIB-TRS-006.

Reference

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