Warning - Anomalies currently not usable

I notice there are some problems with the anomalous sequences. Please do not consider using the dataset with the firing sequence marked as anomalous. I am investigating what is the problem and work towards a new release. I recommend to not use this dataset for anomaly detection at the moment.

Context

Testing hardware to qualify it for Spaceflight is critical to model and verify performances. Hot fire tests (also known as life-tests) are typically run during the qualification campaigns of satellite thrusters, but results remain proprietary data, hence making it difficult for the machine learning community to develop suitable data-driven predictive models. This synthetic dataset was generated partially based on the real-world physics of monopropellant chemical thrusters, to foster the development and benchmarking of new data-driven analytical methods (machine learning, deep-learning, etc.).

Overview of chemical monopropellant thruster

A monopropellant thruster is an engine that provide thrust by usage a unique propellant, as opposed to bipropellant systems which uses the combustion of fuel and oxidizer. The unique propellant flow into the chamber is controlled by a valve, usually an integral part of the thruster. It is injected into a catalyst bed, where it decomposes. A monopropellant must be slightly unstable chemical, which will decompose exothermally to produce a hot gas. The resulting hot gases are expelled through a converging/diverging nozzle generating thrust. The gas temperature is high which require the usage of high-temperature alloys to manufacture the nozzle.

The most classical type of monopropellant thrusters are reaction control thrusters generating about 1 to 10 Newton of thrust using hydrazine as propellant. These reaction control thrusters are used, for instance to control the attitude of a spacecraft and/or to desaturate the reaction wheels.

The performance of a monopropellant thruster (and its degradation) is mostly driven by the valve performance and the s of the catalyst bed on which the propellant decomposes. The life of the catalyst bed is mainly affected by the degradation of catalyst granules. The catalyst is made of alumina-based Indium metal granules (about 1mm in diameter) that are carefully designed and selected to optimize its lifetime. However, catalyst granules are easily damaged by thermoelastic shocks, collisions with other granules, and so on, thus hey are broken up into fine particles which reduces their efficiency. After the long duration of firing, large voids are formed in the catalyst bed and induce unstable decomposition of hydrazine and degradation of thruster performance.

The properties of this simulated thruster fire tests are fictious and not necessarily equivalent to a real-world thruster available on the market. Nevertheless, it provides sufficient granularity and challenge to benchmark algorithm that may then be tested on real fire test sequences. This is possible because the simulator is based, partially, on real-world physics of such reaction control thrusters. The details of the simulator are not provided on purpose to avoid leakage into feature engineering methods and modelling approaches developed.

Tasks and use cases

Regression for Performance Prediction

• Performance Modelling: Prediction of the thruster performances (target can be thrust, mass flow rate, and/or the average specific impulse of over a given sequence that can be calculated using the first two). This task may also be referred to as “time seriesforecasting with exogenous inputs”, or “System Dynamics modelling with control”, where the control is the command sent to the thruster (on/off) in the column “ton”. It is very unlikely in practice to have more than a few complete SN qualification datasets, so we suggest a 50:50 split between the train and test sets.
  • Use SN01 to SN12 to build your train and validation sets. These are data assumed to be collected on-ground.
  • Use SN13 to SN24 as the final test set. Warning: given that these thrusters are expected to be mounted on a flying satellite, you cannot retrain the model progressively using SN13 to SN24 data. The data is the ground truth as if we were able to measure it on-board.
• Acceptance Test for Individualised Performance Model refinement:
  • In practice, every newly manufactured thruster is tested via acceptance tests: test_id 1 to 12. So these measurement are actually available for each thruster serial number. For this reason, it is possible to use the acceptance test data (test_id 1 to 12) of SN13 to SN24 as input to the model to predict the flight performance of these SN over test_id 13 to 112. Taking into account the acceptance test of individual thruster might be helpful to generate individualized thruster predictive model. This use case propose to investigate...

Data from: Data-driven analysis of oscillations in Hall thruster simulations

- Authors: Davide Maddaloni, Adrián Domínguez Vázquez, Filippo Terragni, Mario Merino

- Contact email: dmaddalo@ing.uc3m.es

- Date: 2022-03-24

- Keywords: higher order dynamic mode decomposition, hall effect thruster, breathing mode, ion transit time, data-driven analysis

- Version: 1.0.4

- Digital Object Identifier (DOI): 10.5281/zenodo.6359505

- License: This dataset is made available under the Open Data Commons Attribution License

Abstract

This dataset contains the outputs of the HODMD algorithm and the original simulations used in the journal publication:

Davide Maddaloni, Adrián Domínguez Vázquez, Filippo Terragni, Mario Merino, "Data-driven analysis of oscillations in Hall thruster simulations", 2022 Plasma Sources Sci. Technol. 31:045026. Doi: 10.1088/1361-6595/ac6444.

Additionally, the raw simulation data is also employed in the following journal publication:

Borja Bayón-Buján and Mario Merino, "Data-driven sparse modeling of oscillations in plasma space propulsion", 2024 Mach. Learn.: Sci. Technol. 5:035057. Doi: 10.1088/2632-2153/ad6d29

Dataset description

The simulations from which data stems have been produced using the full 2D hybrid PIC/fluid code HYPHEN, while the HODMD results have been produced using an adaptation of the original HODMD algorithm with an improved amplitude calculation routine.

Please refer to the relative article for further details regarding any of the parameters and/or configurations.

Data files

The data files are in standard Matlab .mat format. A recent version of Matlab is recommended.

The HODMD outputs are collected within 18 different files, subdivided into three groups, each one referring to a different case. For the file names, "case1" refers to the nominal case, "case2" refers to the low voltage case and "case3" refers to the high mass flow rate case. Following, the variables are referred as:

• "n" for plasma density
• "Te" for electron temperature
• "phi" for plasma potential
• "ji" for ion current density (both single and double charged ones)
• "nn" for neutral density
• "Ez" for axial electric field
• "Si" for ionization production term
• "vi1" for single charged ions axial velocity

In particular, axial electric field, ionization production term and single charged ions axial velocity are available only for the first case. Such files have a cell structure: the first row contains the frequencies (in Hz), the second row contains the normalized modes (alongside their complex conjugates), the third row collects the growth rates (in 1/s) while the amplitudes (dimensionalized) are collected within the last row. Additionally, the time vector is simply given as "t", common to all cases and all variables.

The raw simulation data are collected within additional 15 variables, following the same nomenclature as above, with the addition of the suffix "_raw" to differentiate them from the HODMD outputs.

Citation

Works using this dataset or any part of it in any form shall cite it as follows.

The preferred means of citation is to reference the publication associated to this dataset, as soon as it is available.

Optionally, the dataset may be cited directly by referencing the DOI: 10.5281/zenodo.6359505.

Acknowledgments

This work has been supported by the Madrid Government (Comunidad de Madrid) under the Multiannual Agreement with UC3M in the line of ‘Fostering Young Doctors Research’ (MARETERRA-CM-UC3M), and in the context of the V PRICIT (Regional Programme of Research and Technological Innovation). F. Terragni was also supported by the Fondo Europeo de Desarrollo Regional, Ministerio de Ciencia, Innovación y Universidades - Agencia Estatal de Investigación, under grants MTM2017-84446-C2-2-R and PID2020-112796RB-C22.

Q-thruster technology is a mission enabling form of electric propulsion and is already being traded by NASA's Concept Architecture Team (CAT) & Human Space Flight (HSF) Architecture Team (HAT) as an electric propulsion effector for Asteroid Recovery Vehicle (ARV) mission extensibility options out to Mars.  The Nuclear Electric Propulsion mission allows for rapid transit while allowing for a heavy, more near-term reactor design and the Solar Electric Propulsion mission allows for a power starved approach with similar mission durations to Design Reference Architecture - DRA 5.0 that would not be possible without the Q-thruster technology.

Data from: Analysis of a cusped helicon plasma thruster discharge

- Authors: Pedro Jimenez, Jiewei Zhou, Jaume Navarro, Pablo Fajardo, Mario Merino, Eduardo Ahedo

- Contact email: pejimene@ing.uc3m.es

- Date: 2024-02-08

- Keywords: electric propulsion, electrodeless plasma thruster, helicon plasma thruster, plasma simulation

- Version: 1.1

- License: This dataset is made available under the Open Data Commons Attribution License

Abstract

This dataset contains the data found in the plots of the journal article:

Analysis of a cusped helicon plasma thruster discharge (Plasma Sources Science and Technology)

Dataset description

The data in this repository has been extracted from the experiments and simulations as described in the manuscript.

For further information on the setup for the simulation please refer to the article.

Data files

The data files are in comma separated .csv format. Many programming languages provide functionalities to load such fields.

The files are organised following the order of the figures in the article. Therefore each file contains a different sized array.

1D plots: Each line is given as 3 colums, the first one corresponds to the abscissa coordinate (X data), the second one to the ordinate (Y data) and the last one to the error (+/- value in the same units as Y data).

2D contours: All data is provided in a structured mesh. 3 matrices are concatenated in the column dimension. The first two matrices correspond to the Z,R coordinates in a meshgrid format, as customary in Matlab and NumPy. The last matrix corresponds to the field values. A description of the dimensions (extension in columns) of the matrices is provided as a header in the first row. Example: For a field represented in a 100(z) x 50 (r) mesh. The csv will have 50+1(header) rows and 3x100 = 300 columns.

Citation

Any works using this dataset or any part of it in any form shall cite it as follows:

The prefered means of citation is to reference the publication asociated to the jounal article with DOI: 10.1088/1361-6595/ad01da

The BibTex is also provided for the sake of convinience:

@article{jimenez2023analysis,

title={Analysis of a cusped helicon plasma thruster discharge},

author={Jim{\'e}nez, Pedro and Zhou, Jiewei and Navarro, Jaume and Fajardo, Pablo and Merino, Mario and Ahedo, Eduardo},

journal={Plasma Sources Science and Technology},

volume={32},

number={10},

pages={105013},

year={2023},

publisher={IOP Publishing}

}

Optionally the dataset can be cited by referencing the DOI: 10.5281/zenodo.8154866

Acknowledgments

This dataset was created by the ERC-ZARATHUSTRA project.

The ERC-ZARATHUSTRA project has received funding from the European Research Council (ERC) under the European Union’s Horizon 2020 research and innovation programme (grant agreement No 950466).

Initial support for the activities leading to work came from the HIPATIA project, funded by the European Union’s Horizon 2020 Research and Innovation Program (grant agreement No 870542).

This study will investigate the development of an atmosphere-breathing electric propulsion solar-powered vehicle to explore planets such as Mars. The vehicle would use atmospheric gas for propellant, eliminating the need to launch and carry the propellant from Earth. The propulsion thruster would be electric where the gas is ionized in a plasma and accelerated by electromagnetic fields. The combination of high efficiency and high specific impulse of the electric propulsion thruster and free propellant in-situ will result in an exciting and enabling technology. At the completion of this development, NASA will be able to perform missions of extended lifetime and capabilities beyond those available by typical chemical rockets. Phase I will formulate feasibility of the concept through modeling, calculations and preliminary laboratory experiments and push validity into Phase II research.

WARNING

This version of the dataset is not recommended for anomaly detection use case. We discovered discrepancies in the anomalous sequences. A new version will be released. In the meantime, please ignore all sequence marked as anomalous.

CONTEXT

Testing hardware to qualify it for Spaceflight is critical to model and verify performances. Hot fire tests (also known as life-tests) are typically run during the qualification campaigns of satellite thrusters, but results remain proprietary data, hence making it difficult for the machine learning community to develop suitable data-driven predictive models. This synthetic dataset was generated partially based on the real-world physics of monopropellant chemical thrusters, to foster the development and benchmarking of new data-driven analytical methods (machine learning, deep-learning, etc.).

The PDF document "STFT Dataset Description" describes in much details the structure, context, use cases and domain-knowledge about thruster in order for ML practitioners to use the dataset.

PROPOSED TASKS

Supervised:

Performance Modelling: Prediction of the thruster performances (target can be thrust, mass flow rate, and/or the average specific impulse)

Acceptance Test for Individualised Performance Model refinement: Taking into account the acceptance test of individual thruster might be helpful to generate individualised thruster predictive model

Uncertainty Quantification for Thruster-to-thruster reproducibility verification, i.e. to evaluate the prediction variability between several thrusters in order to construct uncertainty bounds around the prediction (predictive intervals) of the thrust and mass flow rate of future thrusters that may be used during an actual space mission

Unsupervised / Anomaly Detection

Anomaly Detection: Anomalies can be detected in an unsupervised setting (outlier detection) or in a semi-supervised setting (novelty detection). The dataset includes a total of 270 anomalies. A simple approach is to predict if a firing test sequence is anomalous or nominal. A more advanced approach is trying to predict which portion of a time series is anomalous. The dataset also provide a detailed information about each time point being anomalous or nominal. In case of an anomaly, a code is provided which allows to diagnosis the detection system performance on the different types of anomalies contained in the dataset.

CubeSats enable a wide range of important, low-cost space missions, leveraging increasing densification of electronics and communications equipment to perform tasks previously in the realm of much larger spacecraft. CubeSats in sizes of 3U-12U are a promising platform for future NASA exploration and science missions, including lunar remote sensing, NEO exploration, and planetary defense, but currently lack the high ΔV propulsion required for these exciting applications. Ion thrusters are currently the best option for low power, mN range EP as required for high ΔV SmallSat missions. These thrusters are inherently more efficient than other options such as RF or Microwave, and are capable of high total impulse. This work aims to realize these advantages at the low power, miniature scale.

I propose a multi-pronged approach for the improvement of the TRL 4 MiXI thruster for miniature spacecraft and precision flying applications. By incorporating lessons learned from the UCLA Plasma and Space Propulsion Lab's MARCI experiment, the discharge efficiency of MiXI will be improved to that of larger, conventional ion thrusters, resulting in a final system with >70% total efficiency, under 100W operation, Isp >1500s, >1000m/s ΔV capability on 3-12U CubeSats, and 0.1-1.55mN thrust.

Computational simulations and experimental validations will be undertaken to optimize the MARCI discharge for a miniature ion thruster, which will then be integrated into the heritage MiXI thruster. New grids will be designed for the resulting change in discharge conditions. Flight-worthy miniature cathodes will be implemented as electron sources for the discharge and beam neutralization while minimizing propellant and power usage. The resulting enhanced MiXI will be characterized within the Plasma & Space Propulsion Lab's Miniature Electric Propulsion Test Facility to establish throttle curves and performance. Lifetime will be assessed through a 1000-hour grid life validation test and extended to 30000 hours through NSTAR models. Finally, the thruster will be assembled into a complete propulsion subsystem and tested within a spacecraft analog in simulated space conditions for TRL 6 maturation.

This development effort promises to enable low power missions with high ΔV, high propellant efficiency, and high electrical efficiency. Miniature ion thruster development is a NASA identified enabling technology for missions in planetary science, heliophysics, and astronomy, with the primary application of SmallSat and CubeSat primary propulsion. This work will address the primary obstacle for widespread adoption of miniature ion thrusters. This will be accomplished by dramatically improving the electrical efficiency of MiXI through the implementation of a novel discharge topology, and demonstration of a propulsion system capable of >1,000m/s ΔV for SmallSat missions.

Proposed here is the development and testing of a long-life high-power Hall thruster. The thruster will be a 50-kW class nested-channel Hall thruster. The thruster will be designed based on a heritage of other thrusters, namely the H6MS and the X3. Both were jointly developed by the Plasmadynamics and Electric Propulsion Laboratory (PEPL) at the University of Michigan, NASA and the Air Force Office of Scientific Research. The X3 is currently undergoing a full performance characterization to gather thrust data and plume characteristics. This data will be used to develop a physics model of the thruster and investigate the interaction between channels in nested-channel Hall thrusters. The new thruster will be a two-channel magnetically-shielded thruster, ideally simulated on the physics model developed from X3. It will have a nominal power between 35 and 95 kW. New breakthroughs in life-prolonging technologies will be implemented on the thruster to enable a minimum lifetime of 10,000 hours. The development of the high power Hall thruster was a critical step forward in the field of electric propulsion and in NASA's plan to leverage them on a variety of missions, such as a manned mission to Mars. However, the technology cannot currently be used on deep space missions due to lifetime issues. Erosion is the main mechanism of failure on a Hall thruster, as once the wall thickness reaches a critical level, the magnetic circuit is exposed and shorted. The development of this thruster will mitigate this issue by significantly reducing, and ultimately stopping, the erosion of the channel. The project involves the design, fabrication, and testing of a two-channel magnetically shielded thruster. The main design changes from unshielded thrusters are the magnetic field and the chamfering of the isolation walls at thruster exit. These change allows for plasma potential near the walls equal to that of the discharge voltage, low electron energies near the wall and field lines that exit the thruster without intersecting the wall such that the ions are not accelerated into the walls. These aspects are the main requirements behind magnetic shielding of a Hall thruster. Once thruster design and fabrication is complete, testing will commence to assess performance and plasma characteristics. Measurements to estimate erosion rates will be performed to confirm that erosion has been dramatically decreased compared to conventional Hall thrusters. A suite of plasma diagnostics will be used to collect data on relevant plasma properties. These will then be used to calculate thruster efficiency and develop a physics model of the thruster similar to that of the X3 US. Additionally, thrust measurements will be taken to confirm that performance is on par with unshielded thrusters. The work proposed here has many implications on the field of electric propulsion. The project itself poses an unprecedented challenge as neither a high-power Hall thruster (50-kW class and above) nor a nested-channel Hall thruster has ever been magnetically shielded. Insight into magnetically shielding such a thruster has far reaching implications for future high-power Hall thrusters. Furthermore, NASA has particular interest in such a project, because it will fulfill the need for a cost-effective propulsion technology for human space exploration. In a recent Broad Agency Announcement, NASA called for a 50 to 300 kW thruster for a variety of missions with a lifetime of up to 10,000 hrs. Currently, this thruster is one of the few thrusters that can meet this demand in the proposed five-year time scale. Mission analysis from NASA has shown that high power Hall thrusters are ideal for near- Earth and deep-space applications. In order to support the next generation of space missions, propulsion systems must be able to meet a variety of performance metrics. Nested-channel Hall thrusters allow for this by having a variety of firing configurations, and magnetic shielding allows for Hall thrus

Busek proposes to develop a high performance, non-toxic storable, "green" monopropellant thruster suitable for in-space reaction control propulsion. The engine will deliver 100N (~25lbf) vacuum thrust with specific impulse exceeding 240sec. Estimated Isp-density is on the order of 348 sec-g/cc, a 48% increase from the state-of-the-art hydrazine systems. The most important feature that sets this thruster apart from other similar devices will be the use of an innovative, long-life catalyst. This proprietary catalyst, constructed without any bed plate or ceramic substrate, was recently demonstrated in Busek's 0.5N micro thruster. It has shown the ability to suppress catalyst-related performance degradation problems that often plague green monopropellant thrusters.
The proposed Phase I program will focus on developing a 5N green monopropellant thruster by scaling up the long-life catalyst design from the 0.5N thruster. Both empirical and modeling works are proposed to validate the scaling theory. Thruster performance will be evaluated based on hot-firing test results that include c* and vacuum thrust measurements. The Phase I findings will lead to the design of a full-scale, 100N green monopropellant thruster to be developed in Phase II.

Technologies include, but are not limited to, electric and advanced chemical propulsion, propellantless propulsion such as aerocapture and solar sails, sample return ascent vehicles, and Earth return systems. ISP will enable access to more challenging and interesting science destinations, including enabling sample return missions. ISP continues to advance several propulsion technologies in support of future Flagship, Discovery, Mars, and New Frontiers missions. The ISP portfolio continues to invest in high-priority technology areas such as the electric propulsion and aerocapture/Earth entry, descent, and landing technologies identified in the Solar System Exploration Roadmap, the 2010 SMD Science Plan, and the 2011 Planetary Decadal Survey. The ISP project is highly responsive to the Decadal Survey. The ISP project will complete the 7kW NASA's Evolutionary Xenon Thruster (NEXT) Power Processing Unit (PPU) repair in 2012, and will complete NEXT PPU characterization and integration testing and long duration validation testing of the NEXT thruster in 2013. ISP is completing the electric propulsion 4kW High Voltage Hall Accelerator (HiVHAC) thruster development task, is assessing commercial Hall systems, and will start long duration testing of the HIVHAC thruster in 2012. The Hall system power processing unit (PPU) and other subsystem technology development starts development in FY 2012. High Voltage Hall Accelerator (HiVHAC) thruster technology is applicable to Earth return vehicles (ERV), transfer stages, and low-cost electric propulsion systems for Discovery-class missions. In FY 2012 ISP will continue development of NDI techniques and a detailed design for a lightweight propellant tank applicable to the Skycrane. ISP will continue completing Earth Entry Vehicle (EEV ) heat shield micro-meteoroid/orbital debris characteristics studies, a preliminary design of a multi-mission Earth entry vehicle (MMEEV) concept and continuing MMEEV technology development.

Computational tools that accurately predict the performance of electric propulsion devices are highly desirable and beneficial to NASA and the broader electric propulsion community. The current state-of-the-art in electric propulsion modeling relies heavily on empirical data and on numerous computational "knobs". In Phase I of this project, we developed the most detailed ion engine discharge chamber model that currently exists. This is a kinetic model that simulates all particles in the discharge chamber along with a physically correct simulation of the electric fields. In addition, kinetic erosion models are included for modeling the ion-impingement effects on thruster component erosion. For Phase II of this project, the goal is to make this sophisticated computer program a user friendly program that NASA and other governmental and industrial customers are able to utilize. In Phase II we will implement a number of advanced numerical routines to bring the computational time down to a commercially acceptable level. At the end of Phase II, NASA will have a highly sophisticated, user friendly ion engine discharge chamber modeling tool that will save time and expense in designing new and different size ion engines, as well as analyzing existing ion engine performance.

Proposed here is a full performance characterization of the X3 Nested-channel Hall Thruster (NHT), a 100-kW class thruster developed jointly by the Plasmadynamics and Electric Propulsion Laboratory (PEPL) at the University of Michigan, NASA, and the Air Force Office of Scientific Research. The thruster has been built and run through a burn-in procedure, but a performance characterization is the vital next step in its development. This characterization will include gathering thrust measurements, a detailed analysis of the plasma plume of the thruster, calculating efficiencies, a study of the interactions between multiple discharge channels running at one time, and an investigation of the effect of magnetic field shape on thruster performance. The characterization will occur at multiple points across the entire operating envelope of the thruster, ranging from 2 kW to over 200 kW of discharge power.

The discharge power levels of the X3 cause unique facility issues that will be addressed by testing at two different facilities during this characterization. The low half of the performance envelope of the thruster will be investigated inside the Large Vacuum Test Facility at PEPL. At higher power levels, the X3 will be investigated in Vacuum Facility 5 (VF-5) at NASA's Glenn Research Center. VF-5 offers the ideal facility in which to run the X3 in the upper half of its performance envelope. This has been planned since the genesis of the project, and much of the infrastructure related to the X3 was designed with testing at VF-5 in mind.

The thrust measurements will be collected using an inverted-pendulum thrust stand. Thrust measurements are an important metric on their own, but also are used to calculate a number of important parameters including anode specific impulse. The analysis of the plasma plume will be done with an array of probes, including a Faraday probe, a Langmuir probe, an ExB probe, and a Retarding Potential Analyzer. The data from these probes additionally will be used to calculate a number of thruster efficiencies. These efficiencies are indicators of thruster performance, and will be used as a comparative tool across multiple operating points.

The work proposed here is important on its own, but has further-reaching implications as well. The two biggest technical challenges with Hall Thrusters, as identified by NASA, are increasing discharge power and increasing lifetime. The X3's main contribution is its power level, but what is proposed here with it will also have influence on thruster lifetime. Magnetic shielding has shown great promise in the area of extending lifetimes, and the data collected in this work can be used to explore the possibility of designing the world's first magnetically-shielded NHT.

The X3 is a monumental step forward for the field of electric propulsion. Goals stated by NASA put targeted discharge power levels in the 100s of kW, and this thruster should achieve that. However, if future, flight-qualified thrusters are to be built, the X3 must be interrogated fully. Understanding not only the performance values but the phenomena from which they come is essential engineering knowledge to apply to future thruster designs.

Flight Works is proposing to expand its technology in micropump-fed propulsion, including 1U CubeSat green propulsion, to the development and demonstration of a low cost, pump-fed, cooled, non-catalytic 1-5 N-class AF-M315E thruster for secondary payload propulsion. Typically, requirements imposed by the primary mission have led secondary payloads to have very limited propulsion capability. For earth orbiting spacecraft, the requirements to reenter within 25 years can be an issue. For lunar or interplanetary missions, lack of significant ΔV capability limits the science potential. For example, the system in the Mars Cube One 6U spacecraft is only capable of a few tens of m/s. Many such nanosats, including CubeSats slated to accompany the primary spacecraft towards Europa, could greatly benefit from real delta-V capability (> 1 km/s) while reducing risks to the primary payload. Flight Works is proposing to develop such capability and focus on the development of a pump-fed AF-M315E thruster. In the novel concept, propellant atomization is improved, conventional materials can be used for the injector and the valve, resulting in a more compact, lower cost, high performance thruster. Also, since the approach to ignition and combustion sustainment does not involve catalysts, the thruster life-limiting component is removed. This thruster is integrated into a micropump-fed system: there is no need for a separate pressurization, the propellant storage and feed system operates at low pressures, and lighter, conformal tanks can be used. This combination decreases system overall size and mass by 20-40% depending on the mission while reducing risks to the primary payload. More generally, the technology is applicable to any propulsion system, whether primary or for attitude control, where hydrazine is currently used, and is competitive with bipropellant systems for microsats due to the reduced system mass.

The global monopropellant thruster market is experiencing robust growth, driven by increasing demand for satellite propulsion systems in the burgeoning space exploration and communication sectors. The market's expansion is fueled by several key factors, including the miniaturization of satellites, rising investments in space-based infrastructure, and the growing adoption of electric propulsion systems that often incorporate monopropellant thrusters for critical maneuvers. The period from 2019 to 2024 witnessed significant advancements in thruster technology, leading to improved efficiency, reliability, and reduced costs. This trend is expected to continue throughout the forecast period (2025-2033), leading to further market penetration. Major players like Busek, ArianeGroup, Moog, IHI Aerospace, Nammo Space, Rafael, Northrop Grumman, and T4i Technology for Propulsion and Innovation are actively involved in developing and deploying advanced monopropellant thruster solutions, fostering innovation and competition within the market. The market segmentation is likely diverse, encompassing various thruster types based on propellant (e.g., hydrazine), size, and application (e.g., attitude control, orbit raising). Regional variations in market share will likely reflect the distribution of space agencies, private space companies, and research institutions. The predicted Compound Annual Growth Rate (CAGR) points to a substantial increase in market value over the forecast period. This growth will be influenced by technological advancements focused on increasing thrust-to-weight ratios, enhancing longevity, and improving overall system performance. Factors such as stringent regulatory requirements related to propellant handling and environmental concerns regarding the use of hydrazine may pose some constraints on the market's growth. However, the ongoing research and development into alternative, more environmentally benign propellants are expected to mitigate this challenge and sustain the positive trajectory of the monopropellant thruster market in the long term. This overall positive outlook suggests that the market will continue its expansion, driven by innovation, increasing demand, and growing investment across various global regions.

This dataset contains the data found in the plots of the paper:

Mario Merino, Diego García, Eduardo Ahedo, "Plasma acceleration in a magnetic arch.

Provisionally accepted in the journal Plasma Sources Science and Technology

The data in this repository has been extracted from the fluid simulations as described in the [article](link to the paper).

For further information on the setup for the simulation please refer to the article.

Rotating Magnetic Field (RMF) thrusters are a form of electrodeless plasma propulsion. This technology is a low maturity but potentially enabling candidate for high-power in-space propulsion for use with alternative propellants. The purpose of the data here, and the associated publication is to evaluate the phenomenological efficiency modes for this thruster test article to explain and understand its overall efficiency. These modes include divergence, power coupling, mass utilization, and plasma/acceleration efficiency. Additional time-resolved measurements of the internal plasma properties were performed using a triple Langmuir probe to evaluate energy loss processes within the thruster.

Alameda Applied Sciences has an opportunity (by working with Novawurks of Los Alamitos, CA) to secure flight heritage for its Metal Plasma Thruster (MPT-X). The MPT-X is an Electric Propulsion Thruster that delivers >4000 Ns of total impulse from a 1U package, with no moving parts, liquids or gases to be handled. The TRL-4 prototypes have been developed through several generations over two years. This SBIR affords an opportunity to integrate the MPT into Novawurks HISat platform and obtain flight heritage, to be followed by commercial sales to other satelliet builders and for NASA deep space missions.

Present and future spaceflight missions depend on the ability to produce high exhaust velocities while reducing the dependence on chemical fuel and its mass onboard a spaceflight vehicle. Oscillation of gaseous molecules during pre-ejection stages via embedded wave driver allows for ejection at higher velocities, increasing chemical fuel efficiency. Oscillation of granulate and liquid reagents using simple harmonic motion has been shown to excite particles, forming geometric patterns when using calibrated frequencies. Methods shown to induce geometric patterns were used to attain similar formations in the reagents Lycopodium, CO2(g) and SF6(g). Oscillation of Lycopodium was used as a proven method to target, observe, and calibrate specific sound formations for experimentation with gases. Sulfur hexafluoride (SF6) was used to simulate xenon, a dense gas used in modern electronic propulsion devices. Ten-millimeter polypropylene, air-filled mass objects were used to observe acceleration, force, and velocity for a dense gas during oscillation and resulting formations. Observation of non-zero forces within gas formations during oscillation shows that additional thrust velocity can be achieved through the oscillation of propellant gas via wave drivers embedded within experimental electronic propulsion systems. Force and velocity calculations taken during oscillation of SF6 demonstrate proof of concept for future experimentation using xenon as an oscillation and ionization medium for ejection at velocities which can be used for spaceflight. Results of this experiment introduce a novel method for achieving increased velocity during space flight using sound as a performance enhancer.

A table containing the data from this experiment. The yaw is taken from the plume center, not the emitter axis. The yields are raw current from the MCP, but these should not be taken to be absolute, as the gain of the MCP was unknown for these experiments as described in the text. The relative yields are the relative percentages of each plume constituent species. The current is the emitter output current recorded by hand off the SMU.

1967 Global export shipment records of Thruster with prices, volume & current Buyer's suppliers relationships based on actual Global export trade database.