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Journal articles on the topic 'Cold-gas thruster'

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Published: 4 June 2021

Last updated: 17 June 2025

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Lu, Songcai, Yong Gao, Haibo Tu, et al. "Thrust Measurement of an Integrated Multi-Sensor Micro-Newton Cold Gas Thruster." Aerospace 12, no. 3 (2025): 210. https://doi.org/10.3390/aerospace12030210.

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In recent years, cold gas thrusters have been successfully deployed in numerous missions, showcasing their exceptional reliability and enabling ultra-precise space operations across a broad thrust range. This article introduces an integrated cold gas thruster that integrates flow, pressure, and displacement sensors. The thrust range of this thruster can exceed 1000 μN at most, and the resolution can reach up to 0.1 μN at low thrust. The results of the high-precision displacement sensor are good, showing that the thruster performs well in terms of flow control accuracy and thrust output sensitivity. The measurement accuracy of the force frame itself is also excellent, and it can detect small thrust changes of 0.1 μN. The thrust noise level of the thruster is good, comparable to the standard noise levels of the experimental environment.

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Zharikov, K. I., and M. M. Dron`. "ANALYSIS OF THE USE OF COLD GAS THRUSTERS IN SPACECRAFT PROPULSION SYSTEMS." DYNAMICS OF SYSTEMS, MECHANISMS AND MACHINES 10, no. 3 (2022): 28–35. http://dx.doi.org/10.25206/2310-9793-2022-10-3-28-35.

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In order to determine directions for improving the energy characteristics of spacecraft propulsion systems due to the combined of cold gas and electric thrusters, the authors analyzed the design schemes and energy characteristics of cold gas thruster used on spacecraft. Propellant sharing in a combined propulsion system, as well as a single engine design combining cold gas and electric thrusters, is a promising technology. At the same time, it is emphasized that at the current time the main tasks in the development of combined propulsion systems are to ensure the compatibility of the operation of an electric rocket engine with the propulsion of a cold gas thruster

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Yachmenev, P. S., I. S. Vavilov, and V. V. Fedyanin. "EXPERIMENTAL STUDIES OF AN ARCJET THRUSTER FOR CORRECTIVE PROPULSION SYSTEMS OF SMALL SPACECRAFT." DYNAMICS OF SYSTEMS, MECHANISMS AND MACHINES 12, no. 4 (2024): 48–54. https://doi.org/10.25206/2310-9793-2024-12-4-48-54.

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A prototype arcjet thruster with low energy consumption was studied in vacuum conditions to obtain thrust and mass consumption of the working fluid by experimental methods. The thrust was determined by a pendulum, the movements of the pendulum were recorded by a video camera, and the results were processed on a PC using computer vision software. The pendulum does not have a steady position, and with the inclusion of an electric discharge, the amplitude of the pendulum's oscillations increases. The thrust of the arcjet thruster with cold gas supply was 0.00112 N, and when the electric discharge was switched on, 0.00108 N. Nitrogen consumption was determined by the method of pressure drop in a container of a known volume and amounted to 1.54 mg/s with cold gas, and 1.48 mg/s with energy supply. The specific thrust impulse was calculated. It was 727 m/s with cold gas, and 737 m/s with energy supply. The total power consumption of the prototype was 3.6 watts.

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Stone, William C. "Fast variable-amplitude cold gas thruster." Journal of Spacecraft and Rockets 32, no. 2 (1995): 335–43. http://dx.doi.org/10.2514/3.26615.

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Lu, Songcai, Xuhui Liu, Xudong Wang, Shurui Zhang, Yusong Yu, and Yong Li. "Direct Simulation Monte Carlo Simulation of the Effect of Needle Valve Structures on the Rarefied Flow of Cold Gas Thrusters." Micromachines 14, no. 8 (2023): 1585. http://dx.doi.org/10.3390/mi14081585.

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The needle valve, serving as the flow control unit of the thruster system, is a crucial component of the entire thruster. Its performance directly impacts the flow state of the rarefied gas in the micro-nozzle structure of the cold gas micro-thruster, thereby exerting a significant influence on the high precision and stability of the propulsion system as a whole. This study examines the impact of different needle valve structures on the flow and thrust in micro-nozzles using the DSMC method. The analysis includes discussions on the spatial distribution, Kn distribution, slip velocity distribution, and pressure distribution of the micro-nozzle’s flow mechanism. Notably, increased curvature of the needle valve enhances the flow velocity in the throat and expansion section. The magnitude of the curvature directly affects the flow velocity, with larger curvatures resulting in higher velocities. Comparing different spool shapes, the conical spool shape minimizes the velocity gradient in the high-speed region at the junction between the spool area and the outlet pipe, particularly with a wide opening. Increasing the curvature of the spool leads to a higher velocity in the expansion section. Consequently, an arc-shaped spool valve maximizes the nitrogen flow at the nozzle during wide openings, thereby enhancing thrust. These research findings serve as a valuable reference for the structural design of the needle valve in the micro-nozzle of the cold gas micro-thruster.

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Silik, Yusuf, and Ulas Yaman. "Control of Rotary Inverted Pendulum by Using On–Off Type of Cold Gas Thrusters." Actuators 9, no. 4 (2020): 95. http://dx.doi.org/10.3390/act9040095.

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This article describes the control of a rotary inverted pendulum utilizing on–off-type cold gas thrusters as the actuators, which have high similarities with thruster actuated spacecrafts with slosh dynamics. The study is completed in three phases. Firstly, a Pulse Width Modulator (PWM) design method is utilized to obtain quasi-linear thrust output from the on–off-type thrusters. Then, a single axis angle controller is designed and tested on the setup along with the PWM scheme. Finally, a pendulum is connected to the other end of the platform and a rotary inverted pendulum (Furuta Pendulum) is constructed. In this way, an inherently unstable, under-actuated, on–off driven system is obtained. For the swing-up motion of the pendulum, an energy-based method is employed. Balancing of the pendulum is achieved by an observer-based state feedback controller under small angle assumption and quasi-linear outputs from the PWM driven thrusters. All of these control methodologies are realized on a real-time target machine. The pendulum is stabilized in seven seconds after five swings, which is comparable to the systems with electric motors.

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Zaberchik, Michael, Dan R. Lev, Eviatar Edlerman, and Avner Kaidar. "Fabrication and Testing of the Cold Gas Propulsion System Flight Unit for the Adelis-SAMSON Nano-Satellites." Aerospace 6, no. 8 (2019): 91. http://dx.doi.org/10.3390/aerospace6080091.

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Adelis-SAMSON is a nano-satellite mission aimed at performing geo-location of target signals on Earth using a tight three-satellite formation in space. To maintain formation, each nano-satellite is equipped with a cold gas propulsion system. The design, qualification, and integration of the Adelis-SAMSON nano-satellite propulsion system is presented in this paper. The cold gas propulsion system mass is approximately 2 kg, takes a volume of 2U, and generates a thrust of 80 mN from four thrusters using krypton as a propellant. We first present the propulsion system requirements and corresponding system configuration conceived to meet the mission requirements. Subsequently, we present the system architecture while listing all the components. We overview the particular role and qualification process of four of the propulsion system’s components: the propellant tank, thruster assembly, pressure regulators, and fill and vent valve. We detail the tests performed on each component, such as proof pressure tests, vibration tests, and external leak tests. Finally, we present the propulsion system level tests before delivery for satellite integration and discuss the propulsion system’s concept of operations.

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Wachs, Benjamin N., and Benjamin A. Jorns. "Sub-millinewton thrust stand and wireless power coupler for microwave-powered small satellite thrusters." Review of Scientific Instruments 93, no. 8 (2022): 083507. http://dx.doi.org/10.1063/5.0088831.

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The design and performance of a thrust stand for characterizing low-power electric propulsion thrusters are presented. The thrust stand is capable of sub-millinewton resolution for devices on the order of 1 kg. The architecture is based on a counter-weighted hanging pendulum design, a variant of the standard hanging pendulum that employs a counterweight to increase force resolution. Thrust is measured in a displacement mode using the change in position of the pendulum arm as measured by an optical displacement sensor. Passive eddy-current damping is used to offset oscillations and decrease setting time. An in situ calibration rig using known masses is used to calculate thrust. The thrust stand features an adjustable counterweight for in-vacuum sensitivity adjustment. In addition, the design of a broadband (600–2490 MHz) wireless microwave power coupler is presented. The device eliminates stiffness and thermal drift introduced by coaxial cables—typically the leading source of error in testing low-power microwave and radio frequency-powered thrusters. The thrust stand and coupler were tested using an electron cyclotron resonance magnetic nozzle thruster operating with xenon at flow rates from 1 to 10 sccm and powers ranging from zero (cold gas thrust) to 40 W. The resulting measurements showed a force resolution of [Formula: see text]N over a range of thrusts from [Formula: see text] to 600 µN.

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Tajmar, M., P. Laufer, and D. Bock. "Chip based MEMS Ion Thruster to significantly enhance Cold Gas Thruster Lifetime for LISA." Journal of Physics: Conference Series 840 (May 2017): 012012. http://dx.doi.org/10.1088/1742-6596/840/1/012012.

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JEON, Sangwoon, Seul JUNG, Jihun KIM, and Joonyun KIM. "Thrust Measurement of a Cold Gas Thruster for KSLV-I under Vacuum Conditions." TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 58, no. 2 (2015): 108–9. http://dx.doi.org/10.2322/tjsass.58.108.

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Xu, Hao, Qiangbing Mao, Yong Gao, et al. "A newly designed decoupling method for micro-Newton thrust measurement." Review of Scientific Instruments 94, no. 1 (2023): 014504. http://dx.doi.org/10.1063/5.0120130.

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A decoupling method is proposed for micro-Newton thrust measurement with a torsion pendulum. The basic approach is to reduce the influences introduced by the propellant tube and wires of the thruster. A hollow aluminum tube is used to hang the torsion pendulum and is also chosen as the transport pipe for the propellant of the thruster. The electric control box of the thruster is mounted on the pendulum body, which is powered by an externally installed power supply through a liquid metal conductive unit. The control of the electric control box is performed through wireless transmission. With this design, the influences of the propellant tube and connection wires between the torsion pendulum and the outside device are reduced and the stability of the torsion spring constant of the system can be improved. The use of the liquid metal conductive unit reduces the coupling between the wires and the measurement system. The feasibility of the wireless transmission is analyzed. The error sources during the thrust measurement are analyzed, and the expected three σ uncertainty of the thrust is [Formula: see text]N for the measurement of the cold gas thruster. The scheme provides a thrust measurement with higher precision and stability.

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Xu, Shu-Yan, Lu-Xiang Xu, Lin-Xiao Cong, Yong-Gui Li, and Cong-Feng Qiao. "First result of orbit verification of Taiji-1 hall micro thruster." International Journal of Modern Physics A 36, no. 11n12 (2021): 2140013. http://dx.doi.org/10.1142/s0217751x21400133.

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The Hall Micro Thrusters (HMTs) use cold gas or accelerated plasma dual mode to provide ultra-precise spacecraft altitude control. They were operated in space for the first time as part of the demonstration payloads on Chinese Academy of Science’s (CASs) Taiji-1 spacecraft since September 2019. Hall Micro Thruster Assemblies (HMTAs) were the actuators in drag-free control, and will compensate the nonconservative force for gravity wave observatories. The HMTAs meet the requirements of operating at 5–100 [Formula: see text] N of thrust with 0.7 [Formula: see text] N resolution and [Formula: see text]0.6 [Formula: see text] N/Hz[Formula: see text] (0.01–1 Hz) noise to deliver the nanometer-level precision control as fast as 30 ms measured by Gravitational Reference Sensor (GRS). A transfer function model in z-domain was fit and used to filter HMTs cathode voltage to predict GRSs thrust noise response. Simulations of a single or dual-frequency disturbance and the corresponding compensation demonstrated that HMTAs could deliver the required thrust profile expected. The capability to meet the requirements of thruster noise in drag-free control is critical for future missions because the acceleration noise on test mass directly relates to the gravity wave signa l. Preliminary in-orbit verification of Taiji-1 has showed HMTAs’ great potential in future, and the data in the experiments are presented in this paper.

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Zhang, Shurui, Yong Li, Xudong Wang, Songcai Lu, Yusong Yu, and Jun Yang. "Effects of the Wall Temperature on Rarefied Gas Flows and Heat Transfer in a Micro-Nozzle." Micromachines 15, no. 1 (2023): 22. http://dx.doi.org/10.3390/mi15010022.

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When the satellite is in orbit, the thruster will experience drastic temperature changes (100–1000 K) under solar radiation, which will affect the rarefied gas flow state in the micro-nozzle structure of the cold gas micro-thruster. In this study, the effect of different wall temperatures on the rarefied flow and heat transfer in the micro-nozzle is investigated based on the DSMC method. The micro-nozzle structure in this paper has a micro-channel with a large length-to-diameter ratio of 10 and a micro-scale needle valve displacement (maximum needle valve displacement up to 4 μm). This leads to more pronounced multiscale flow characteristics in the micro-nozzle, which is more influenced by the change in wall temperature. At wall temperatures ranging from 100 K to 1000 K, the spatial distribution of local Kn distribution, slip velocity distribution, temperature, and wall heat flux distribution in the micro-nozzle were calculated. The slip flow region is located in the flow channel and transforms into transition flow as the slip velocity reaches approximately 50 m/s. The spatial distribution of the flow pattern is dominated by the wall temperature at small needle valve opening ratios. The higher the wall temperature, the smaller the temperature drop ratio in the low-temperature region inside the micro-nozzle. The results of the study provide a reference for the design of temperature control of micro-nozzles in cold gas micro-thrusters.

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Armano, M., H. Audley, G. Auger, et al. "A Strategy to Characterize the LISA-Pathfinder Cold Gas Thruster System." Journal of Physics: Conference Series 610 (May 11, 2015): 012026. http://dx.doi.org/10.1088/1742-6596/610/1/012026.

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Wang, Zhaoli, Changbin Guan, Xudong Wang, Weijie Zheng, and Longfei Su. "Study on Characteristics of a High-Precision Cold Gas Micro Thruster." Engineering 16, no. 01 (2024): 38–45. http://dx.doi.org/10.4236/eng.2024.161005.

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Moriai, Isamu, Kentaro Yamauchi, Hiroyuki Koizumi, and Kimiya Komurasaki. "Performance Validation of the Six-Degree-of-Freedom Thrust Stand Using a Cold Gas Jet Thruster." JOURNAL OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 72, no. 2 (2024): 59–63. http://dx.doi.org/10.2322/jjsass.72.59.

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Reid, Michael R., David B. Scharfe, Faraz A. M. Saleem, and Rebecca N. Webb. "Preheating Cold Gas Thruster Flow Through a Thermal Energy Storage Conversion System." Journal of Propulsion and Power 29, no. 6 (2013): 1488–92. http://dx.doi.org/10.2514/1.b34898.

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Gatsonis, Nikos A., R. E. Erlandson, P. K. Swaminathan, and C. K. Kumar. "Analysis of Pressure Measurements During Cold-Gas Thruster Firings Onboard Suborbital Spacecraft." Journal of Spacecraft and Rockets 36, no. 5 (1999): 688–92. http://dx.doi.org/10.2514/2.3480.

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Lugini, Claudio, and Marcello Romano. "A ballistic-pendulum test stand to characterize small cold-gas thruster nozzles." Acta Astronautica 64, no. 5-6 (2009): 615–25. http://dx.doi.org/10.1016/j.actaastro.2008.11.001.

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Alikhani, َAlireza, and Mohammad Reza Salimi. "Design, Construction and Performance Evaluation of a Cold Gas Thruster Test Stand." Journal of Space Science and Technology 15, English Special Issue (2022): 55–64. http://dx.doi.org/10.30699/jsst.2021.1305.

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Lekzian, E., A. Ebrahimi, and H. Parhizkar. "Performance analysis of microelectromechanical thrusters using a direct simulation Monte Carlo solver." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 232, no. 7 (2017): 1212–22. http://dx.doi.org/10.1177/0954410017691066.

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Abstract In the present study, a direct simulation Monte Carlo solver is utilized to simulate the effects of heater plates on the performance parameters of microelectromechanical propulsion devices. The simulation is two dimensional. Proper cell dimensions, number of particles per cell, and grid study are used to guarantee the accuracy of simulations. Three types of microthrusters including cold gas as type 1, a propulsion device with heaters in the walls as type 2, and a microthruster with heater plates inside the domain as type 3 are studied. Type 1 is considered as a reference case and two other types are compared with type 1. It is observed that heater plates inside the microelectromechanical thruster enhance the downstream temperature due to conversion of pressure drop occurred by plates into temperature. In type 3, the specific impulse is enhanced but the thrust force is decreased in comparison with type 1. Heating the walls in type 2 accelerates the flow while there is no considerable pressure reduction. Moreover, all performance parameters are increased in this type. It is also demonstrated that increasing of wall temperature increases thrust and specific impulse and decreases the sensitivity of thruster due to rarefaction effects.

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Grm, Aleksander, Tor‐Arne Grönland, and Tomaž Rodič. "Numerical analysis of a miniaturised cold gas thruster for micro‐ and nano‐satellites." Engineering Computations 28, no. 2 (2011): 184–95. http://dx.doi.org/10.1108/02644401111109222.

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Lackner, Maximilian, Alexander Löhr, Felix Schill, and Martin Van Essche. "Rim Driven Thruster as Innovative Propulsion Element for Dual Phase Flows in Plug Flow Reactors." Fluids 9, no. 7 (2024): 168. http://dx.doi.org/10.3390/fluids9070168.

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The purpose of this work was to test a new setup to pump water with entrained air for application in gas fermentation. A mixed flow, where gas is contained in a liquid to be pumped, rapidly reduces the efficiency of a conventional pump, due to the compressibility of the gas. It is not always possible to degas the fluid, for instance in gas fermentation, which is preferably carried out in tubular reactors (loop fermenters) to achieve a high conversion rate of the gaseous feedstocks. Method: In this work, a rim-driven thruster (RDT) was tested in a lab-scale, cold flow model of a loop reactor with 5–30% (by volume) of gas fraction (air) in the liquid (water) as alternative propulsion element (6 m total pipe length, ambient temperature and pressure). As a result, it was found that the RDT, in connection with a guiding vane providing swirling motion to the two-phase fluid, could pump a mixed flow with up to 25.7% of gas content (by volume) at atmospheric pressure and 25 °C and 0.5 to 2 m/s flow speed. In conclusion, an RDT is advantageous over a classic propulsion element like a centrifugal pump or axial flow pump for transporting liquids with entrained gases. This article describes the potential of rim-driven thrusters, as known from marine propulsion, in biotechnology, the chemical industry, and beyond, to handle multiphase flows.

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Glenn Lightsey, E., Terry Stevenson, and Matthew Sorgenfrei. "Development and Testing of a 3-D-Printed Cold Gas Thruster for an Interplanetary CubeSat." Proceedings of the IEEE 106, no. 3 (2018): 379–90. http://dx.doi.org/10.1109/jproc.2018.2799898.

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Boccelli, S., J. G. McDonald, and T. E. Magin. "14-moment maximum-entropy modeling of collisionless ions for Hall thruster discharges." Physics of Plasmas 29, no. 8 (2022): 083903. http://dx.doi.org/10.1063/5.0100092.

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Ions in Hall effect thrusters are often characterized by a low collisionality. In the presence of acceleration fields and azimuthal electric field waves, this results in strong deviations from thermodynamic equilibrium, introducing kinetic effects. This work investigates the application of the 14-moment maximum-entropy model to this problem. This method consists in a set of 14 partial differential equations (PDEs) for the density, momentum, pressure tensor components, heat flux vector, and fourth-order moment associated with the particle velocity distribution function. The model is applied to the study of collisionless ion dynamics in a Hall thruster-like configuration, and its accuracy is assessed against different models, including the Vlasov kinetic equation. Three test cases are considered: a purely axial acceleration problem, the problem of ion-wave trapping, and finally the evolution of ions in the axial-azimuthal plane. Most of this work considers ions only, and the coupling with electrons is removed by prescribing reasonable values of the electric field. This allows us to obtain a direct comparison among different ion models. However, the possibility to run self-consistent plasma simulations is also briefly discussed, considering quasi-neutral or multi-fluid models. The maximum-entropy system appears to be a robust and accurate option for the considered test cases. The accuracy is improved over the simpler pressureless gas model (cold ions) and the Euler equations for gas dynamics, while the computational cost shows to remain much lower than direct kinetic simulations.

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Tseng, K. C., Y. Y. Lian, Y. S. Chen, T. C. Kuo, B. R. Gu, and J. S. Wu. "Simulations of the FORMOSAT-5 Cold Gas Propulsion System by Using the Hybrid Continuum-Particle Method." Applied Mechanics and Materials 110-116 (October 2011): 707–14. http://dx.doi.org/10.4028/www.scientific.net/amm.110-116.707.

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Reaction control subsystem (RCS) is an onboard satellite propulsion system used to provide required thrusting for orbit raising, orbit maintenance, and attitude control etc.. High pressure flow with high temperature could be generated in a chamber by chemical reactions or other power resources then expelled through a convergent-divergent nozzle to obtain thrust. In order to optimize the thrusting performance, numerical simulation is an efficient method to study the physics and parameters in design phase. In the current study, a hybrid method coupled continuum and particle methods is proposed to simulate flows involving continuum and rarefied regions. The Navier-Stokes (NS) solver named UNIC is developed by Chen and his coworkers. It employs the cell-centered finite-volume method with a hybrid 2D/3D unstructured-grid topology. The proposed particle code named Parallel DSMC Code (PDSC) is a parallelized solver based on the well-known Direct Simulation of Monte Carlo (DSMC) method, which was proposed by Bird in 1976. The physical domain is decomposed into several regimes and each sub-domain executes the serial DSMC code at different processor for speeding up the computing. A practical cold gas thruster with different chamber pressures is simulated by using this hybrid code to study the potential malfunction of the pressure regulator. Then the curve-fitting thrusting equation can be referred to satellite control operations.

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Sturesson, Peter, Ragnar Seton, Lena Klintberg, Greger Thornell, and Anders Persson. "Effect of Resistive and Plasma Heating on the Specific Impulse of a Ceramic Cold Gas Thruster." Journal of Microelectromechanical Systems 28, no. 2 (2019): 235–44. http://dx.doi.org/10.1109/jmems.2019.2893359.

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Takao, Yoshinori. "(Invited) Ionic Liquid Electrospray Ion Sources for Space Propulsion." ECS Meeting Abstracts MA2023-01, no. 56 (2023): 2719. http://dx.doi.org/10.1149/ma2023-01562719mtgabs.

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The number of microspacecraft launched has significantly increased since 2013, and more than 300 nano/microsatellites (< 50 kg) were launched in 2017. Although some microspacecraft already have a propulsion system, a commonly used propulsion system is a cold-gas thruster, which has a low delta-v (a few 10s of m/s), and thus, high delta-v (> 1 km/s) thrusters are required for more advanced missions. One of the candidates for such thrusters is ion thrusters, which were already mounted on two 50 kg-class microsatellites and successfully operated in space. However, it is not easy to miniaturize ion thrusters so that they can be mounted on even 10 kg-class nanosatellites unless high-pressure gas-feeding systems are removed. Moreover, ion thrusters require electron sources to neutralize ion beams, which consume additional propellants and power. Since most recently launched microspacecraft are less than 10 kg, more compact thrusters are required. Ionic liquid electrospray ion sources are attractive devices for such electric propulsion, especially for microspacecraft. Electrospray thrusters typically consist of many emitters, an extractor electrode, and an accelerator electrode optionally. When a high voltage between the emitter and the extractor is applied, the ionic liquid is transported to the emitter tips, and a strong electric field deforms the ionic liquid into a conical shape known as a Taylor cone. When the force of the electric field is stronger than the surface tension pressure of the ionic liquid, ions evaporate directly from the liquid surface. The extracted ions are then accelerated by the potential difference between the emitter and extractor to produce the thrust. Since the ionic liquid is composed of only cations and anions without solvent, its vapor pressure is negligible; therefore, it can exist in the liquid phase in vacuum, simplifying the propellant feed system and reducing the size of the entire propulsion system. Moreover, the direct acceleration of ions without discharges can be possible, and no neutralizers are required, which also contributes to the power and size reduction together with high efficiency. The key component of electrospray thrusters is the emitter, where there are typically three types of structures: externally-wetted (or needle-shaped), internally-wetted (or capillary), and porous ones. Each has ion emission characteristics depending on the structure. Here, we have fabricated these three types of emitter structures with a wide range of scale: 1 to 100 micrometers. Using these emitters, we have conducted ion emission experiments and beam diagnostics. We obtained a high current density over a few 10 mA/cm2 with 1-micrometer scale emitters, a high current of approximately 1 mA with porous emitters, and a high efficiency with externally-wetted emitters. The emitter structures as mentioned earlier and experimental results will be presented at the conference site.

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van der Veek, Bartel, Hector Gutierrez, Brian Wise, Daniel Kirk, and Leon van Barschot. "Vibration Control of Flexible Launch Vehicles Using Fiber Bragg Grating Sensor Arrays." Sensors 25, no. 1 (2025): 204. https://doi.org/10.3390/s25010204.

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The effects of mechanical vibrations on control system stability could be significant in control systems designed on the assumption of rigid-body dynamics, such as launch vehicles. Vibrational loads can also cause damage to launch vehicles due to fatigue or excitation of structural resonances. This paper investigates a method to control structural vibrations in real time using a finite number of strain measurements from a fiber Bragg grating (FBG) sensor array. A scaled test article representative of the structural dynamics associated with an actual launch vehicle was designed and built. The main modal frequencies of the test specimen are extracted from finite element analysis. A model of the test article is developed, including frequency response, thruster dynamics, and sensor conversion matrices. A model-based robust controller is presented to minimize vibrations in the test article by using FBG measurements to calculate the required thrust in two cold gas actuators. Controller performance is validated both in simulation and on experiments with the proposed test article. The proposed controller achieves a 94% reduction in peak–peak vibration in the first mode, and 80% reduction in peak–peak vibration in the second mode, compared to the open loop response under continuously excited base motion.

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Wei, Yanming, Hao Yan, Xuhui Liu, et al. "The View of Micropropulsion Technology for China’s Advanced Small Platforms in Deep Space." Space: Science & Technology 2022 (August 24, 2022): 1–9. http://dx.doi.org/10.34133/2022/9769713.

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In this paper, micropropulsion systems are analyzed in conjunction with the various mission requirements of China’s deep space exploration. As a great challenge facing the world, deep space exploration can be enabled only in a few countries with a success rate of around 50%. With the advancement of spacecraft and scientific instruments, it is now feasible to build small and low-cost spacecraft for a variety of deep space missions. As spacecraft become smaller, there is a need for proper micropropulsion systems. Examples of propulsion system selections for deep space exploration are discussed with a focus on products developed by Beijing Institute of Control Engineering (BICE). The requirements for propulsion systems are different in lunar/interplanetary exploration and gravitational wave detection. Chemical propulsion is selected for fast orbit transfer and electric propulsion for increasing scientific payloads. Cold gas propulsion and microelectric propulsion are good choices for space-based gravitational wave detection due to the capability of variable thrust output at the micro-Newton level. The paper also introduces the sub-1-U micropropulsion modules developed by BICE with satisfactory performance in flight tests, which are promising propulsion systems for small deep space platforms. A small probe with an electric sail propulsion system has been proposed for the future solar system boundary exploration of China. The electric sail serves as not only a propellant-free thruster but also a detector probing the properties of the space medium.

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Modechang, S., K. Chaiyapoch, A. Pinthavanich, and J. Wongwiwat. "The study of regression rate parameters on ABS fuel grain for hybrid micro thruster." IOP Conference Series: Earth and Environmental Science 1500, no. 1 (2025): 012018. https://doi.org/10.1088/1755-1315/1500/1/012018.

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Abstract CubeSat propulsion systems can be categorized based on the type of fuel used, including cold gas propulsion, electric propulsion and chemical propulsion, each possessing unique characteristics. Chemical propulsion generates thrust through chemical reactions, providing significantly higher thrust than other propulsion systems. This research focused on the development and optimization of chemical propulsion systems, with particular attention to hybrid propulsion systems. Hybrid propulsion combines the advantages of solid and liquid fuel systems, enabling start-stop operation while ensuring high safety and system simplicity. This research utilized high-test peroxide (HTP) as the oxidizer for hybrid propulsion due to its high reactivity, safety and ease of storage. In the section on solid fuel, acrylonitrile butadiene styrene (ABS) was used as the solid fuel due to its high energy content, ease of fabrication and low cost. The objective of this research was to study parameters affecting the regression rate coefficients of solid fuel made from ABS and to investigate the port shape to enhance combustion efficiency. The experiment was set up to measure the regression rate of solid fuel by burning HTP as an oxidizer flow rate ranging from 5 to 25 mL/min. Circular port and star port were used in this study. Regression rate results were obtained by measuring fuel weight loss during the test. The results showed that the circular port exhibited a higher regression rate than the star port, which seemed to be attributed to more efficient mixing of the oxidizer and fuel within the circular port compared to the star port. The reaction constant (a) and reaction exponent (n) for the circular port were determined to be 0.00829 and 0.5125, respectively, while the star port exhibited lower values of 0.000237 and - 0.249, respectively. These findings indicated that ABS was a viable alternative as a solid fuel for hybrid propulsion systems and enhanced space exploration capabilities, particularly for small spacecrafts and CubeSats.

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Jarrige, J., P. Thobois, C. Blanchard, et al. "Thrust Measurements of the Gaia Mission Flight-Model Cold Gas Thrusters." Journal of Propulsion and Power 30, no. 4 (2014): 934–43. http://dx.doi.org/10.2514/1.b35091.

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Yachmenev, P. S., V. V. Fedyanin, and I. S. Vavilov. "DEVELOPMENT OF A THRUST MEASUREMENT STAND BASED ON THE AERODYNAMIC METHOD FOR ELECTRIC THRUSTERS OF SMALL SPACECRAFT." DYNAMICS OF SYSTEMS, MECHANISMS AND MACHINES 11, no. 2 (2023): 51–57. http://dx.doi.org/10.25206/2310-9793-2023-11-2-51-57.

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A traction measurement stand based on the aerodynamic method has been developed. To test the stand, a prototype electric rocket engine was used as a cold gas engine. The working medium was nitrogen gas. At a flow rate of 1.2 mg/s, the thrust value was from 0.674 to 0.736 mN, at a flow rate of 1.7 mg/s, the thrust value was 0.934-1 mN and at a flow rate of 2.6 mg/s, the thrust value was from 1.59 to 1.634 mN. The experimental specific thrust impulse was at a flow rate of 1.2 mg/s from 574 to 613 m/s, at a flow rate of 1.7 mg/s from 549 to 588 m/s and at a flow rate of 2.6 mg/s from 612 to 623 m/s. According to the measured pressure values in the chamber of the prototype electric rocket engine, the values of the ideal flow rate of the working fluid from the nozzle were obtained, which amounted to 661 m/s at a flow rate of 1.2 mg/s, 667 m/s at a flow rate of 1.7 mg/s and 674 m/s at a flow rate of 2.6 mg/s. The obtained values of the specific thrust impulse do not contradict the previously obtained experimental data on cold gas jet engines using nitrogen as the working fluid, and also do not exceed the ideal design flow rate

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Kvell, U., M. Puusepp, F. Kaminski, et al. "Nanosatellite orbit control using MEMS cold gas thrusters." Proceedings of the Estonian Academy of Sciences 63, no. 2S (2014): 279. http://dx.doi.org/10.3176/proc.2014.2s.09.

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Kumar Meena, Lokesh, A. Anand, and A. S. Gour. "Thrust estimation for HTS-magnet based Magneto Plasma Dynamic Thrusters (MPDT)." IOP Conference Series: Materials Science and Engineering 1240, no. 1 (2022): 012002. http://dx.doi.org/10.1088/1757-899x/1240/1/012002.

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Abstract At present time, electric propulsion is being considered for outer deep space missions. Magneto Plasma Dynamic Thrusters are capable of accelerating semi- neutral plasma gas (Xenon, Iodine, Argon, Ammonia, or Lithium) to a high exhaust velocity using combination of RF power & high magnetic fields. High Temperature Superconducting (HTS) coils can generate high magnetic field due to high operating current density at temperatures below its critical temperature(Tc). These superconducting coils can be cooled by using the cold of outer dark space. The use of 2nd Generation (2G) HTS coils can increase thrust, by increasing high magnetic field with very low power consumption. Thus, specific impulse due to the combination of RF plasma and HTS based superconducting high magnetic field is increased drastically. In this paper, the thrust calculation are carried out for MPDT for varying magnetic fields produced by superconducting coils. The estimation of exhaust velocity and thrust generated by plasma for an MPDT is of great challenge and is required to determine for calculating the life of mission as well as the distance of travel. These estimations are required to support the indigenous development of MPDT.

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Liu, Xuhui, Dong Li, Xinju Fu, Yong Gao, and Xudong Wang. "MODELING of Rarefied Gas Flows Inside a Micro-Nozzle Based on the DSMC Method Coupled with a Modified Gas–Surface Interaction Model." Energies 16, no. 1 (2023): 505. http://dx.doi.org/10.3390/en16010505.

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In this study, we first considered the influence of micro-nozzle wall roughness structure on molecular collision and reflection behavior and established a modified CLL model. The DSMC method was used to simulate and analyze the flow of the micro-nozzle in the cold gas micro-propulsion system, and the deviation of simulation results before and after the improvement of CLL model were compared. Then, the rarefied flow characteristics under a small needle valve opening (less than 1%) were focused on the research, and the particle position, molecular number density, and spatial distribution of internal energy in the micro-nozzle were calculated. The spatial distributions of the flow mechanism in the micro-nozzle under different needle valve openings were compared and analyzed. It was found that when the needle valve opening is lower than 1%, the slip flow and transition flow regions move significantly upstream of the nozzle, the free molecular flow distribution region expands significantly, and the relationship between thrust force and needle valve opening is obviously different from that of medium and large needle valve openings. The effect of nitrogen temperature on the rarefied flow and thrust force is also discussed in this research. The numerical results showed that as gas temperature increases, the molecular internal energy, momentum, and molecular number density near the nozzle exit are enhanced. The thrust at small needle valve openings was significantly affected by the temperature of the working mass. The results of this study will provide key data for the design and development of cold gas micro-thrusters.

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Chen, Kuangshi, Jinglei Xu, Qihao Qin, and Guangtao Song. "Three issues on nozzle thrust performance in cold-to-hot correlation considering variable specific heat effect." Physics of Fluids 34, no. 7 (2022): 076114. http://dx.doi.org/10.1063/5.0098894.

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A high-temperature effect is crucial in cold-to-hot correlation for thrust nozzles to employ experimental data of cold flow to predict the real flight performance of hot gas. The high-temperature nozzle flow behaves beyond the classical gas dynamics and restricts the feasibility of cold-flow experiments, and it becomes more severe due to the species transformation from cold air to hot gas when safety and cost are considered. For an in-depth awareness, this work refines three fundamental issues regarding the high-temperature variable specific heat effect on nozzle flow characteristics. A comprehensive analysis is performed from theory to applications. First, the flow properties of calorically perfect gas (CPG) and thermally perfect gas (TPG) are distinguished and connected via the basic flow equations. One-dimensional flow theory is extended by the generalized stagnation–static gas functions for TPG. The unanticipated intersections within pressure are discovered, which could produce substantial perplexities in nozzle performance determinations. Second, the pros and cons of two homologous nozzle thrust coefficients are clarified on application objects, definition methods, and solution manners. It is proved that temperature has no influence on thrust coefficients for CPG, while the variable specific heat effect might induce three types of false-positive thrust coefficients, to make flow state unidentified, and further shake the baseline of nozzle performance. Third, for the aggravated variable specific heat effect in cold-to-hot correlation from air to hot gas, two methods are proposed with reliable verifications to solve this issue through introducing a novel concept of relative nozzle operating conditions.

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Lee, Thimthana, Mahdi Davoodianidalik, Dimitrios Tsifakis, Roderick W. Boswell, and Christine Charles. "Exploring potential candidates of alternative solid hydrocarbon propellants for cold-gas thrusters." Acta Astronautica 226 (January 2025): 427–38. http://dx.doi.org/10.1016/j.actaastro.2024.10.047.

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Shan, Jinjun, and Piotr Wenderski. "Hardware-in-the-Loop Simulation for Spacecraft Formation Flying." Journal of Control Science and Engineering 2010 (2010): 1–13. http://dx.doi.org/10.1155/2010/572526.

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This paper presents a hardware-in-the-loop (HITL) simulation approach for multiple spacecraft formation flying. Considering a leader-follower formation flying configuration, a Fuzzy Logic controller is developed first to maintain the desired formation shape under external perturbations and the initial position offsets. Cold-gas on/off thrusters are developed to be introduced to the simulation loop, and the HITL simulations are conducted to validate the effectiveness of the proposed simulation configuration and Fuzzy Logic control.

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LEE, Yang-Suk, and Jun Hwan JANG. "The design and performance on 200N-class bipropellant rocket engine using decomposed H2O2 and Kerosene." INCAS BULLETIN 11, no. 3 (2019): 99–110. http://dx.doi.org/10.13111/2066-8201.2019.11.3.9.

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Mono-propellant thrusters are widely utilized in satellites and space launchers. In many cases, they are using hydrazine as a propellant. However, hydrazine has high toxicity and high risks in using for launch campaign. Recently, low-toxic (green) propellant is being highlighted as a replacement for hydrazine. In this paper, 200N bi-propellant engine using hydrogen peroxide/kerosene was designed/manufactured, and the spray or atomization characteristic and inflation pressure were determined by cold flow test, and combustion and pulse tests in a single cycle same as previous methods were conducted. As uniformly supplying hydrogen peroxide through plate-type orifice to a catalyst bed, the hot gas was created as a reaction with hydrogen and catalyst. And then, it was confirmed that the ignition is possible on the wide range of O/F ratio without additional ignition source. The liquid rocket engine with bi-propellant of hydrogen peroxide/kerosene and design/test methods which developed in this study are expected to be utilized as an essential database for designing of the ignitor/injector of bi-propellant liquid rocket engine using hydrogen peroxide/kerosene with high-thrust/performance in near future.

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SUN, Boao, Shencheng DOU, Xiaoqing WANG, et al. "Design and Calibration of High-Resolution Low-Noise Micro Flow Sensors for Cold Gas Thrusters." Chinese Journal of Space Science 45, no. 2 (2025): 1. https://doi.org/10.11728/cjss2025.02.2024-0147.

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Bayir, Şerif Bilalhan, and Hürrem Akbıyık. "INVESTIGATION OF THE VARIOUS TAPER ANGLE OF THE TIP AND ANGLE OF THE CONICAL SHAPE OF AN AEROSPIKE NOZZLE." Konya Journal of Engineering Sciences 12, no. 4 (2024): 908–19. https://doi.org/10.36306/konjes.1474579.

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This study considers the design of aerospike nozzles and the thrust measurement experiments conducted with various designs, taking into account existing research in the literature. Unlike other studies, this experimental study examines the effects of the angle of the conical shape and the taper angle at the tip of the aerospike nozzle on thrust. Thrust measurement experiments were conducted using a cold gas system. All measurement results were compared, revealing that the design parameters for aerospike nozzle efficiency in this study were a conical shape angle of 45° following the main flow passage and a tapering angle of 30° at the tip of the aerospike nozzle. Additionally, it appears that the taper angle and conical angle of the aerospike nozzle are interrelated in thrust production.

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Dentis, Matteo, Elisa Capello, and Giorgio Guglieri. "A Novel Concept for Guidance and Control of Spacecraft Orbital Maneuvers." International Journal of Aerospace Engineering 2016 (2016): 1–14. http://dx.doi.org/10.1155/2016/7695257.

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The purpose of this paper is the design of guidance and control algorithms for orbital space maneuvers. A 6-dof orbital simulator, based on Clohessy-Wiltshire-Hill equations, is developed in C language, considering cold gas reaction thrusters and reaction wheels as actuation system. The computational limitations of on-board computers are also included. A combination of guidance and control algorithms for an orbital maneuver is proposed: (i) a suitably designed Zero-Effort-Miss/Zero-Effort-Velocity (ZEM/ZEV) algorithm is adopted for the guidance and (ii) a linear quadratic regulator (LQR) is used for the attitude control. The proposed approach is verified for different cases, including external environment disturbances and errors on the actuation system.

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Verma, S. B., and Oskar Haidn. "Cold Gas Testing of Thrust-Optimized Parabolic Nozzle in a High-Altitude Test Facility." Journal of Propulsion and Power 27, no. 6 (2011): 1238–46. http://dx.doi.org/10.2514/1.b34320.

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Badis, Tarek, Elhassen Benfriha, and Jalal Eddine Benmansour. "Fuel Consumption Estimation via Bookkeeping Method for Geostationary Satellites: Simple Application." International Journal of Energetica 8, no. 2 (2023): 54. https://doi.org/10.47238/ijeca.v8i2.228.

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This work focuses on the Satellite Propulsion Subsystem (UPS), a critical aspect of satellite technology that can be supported by various propulsion types: electrical, chemical, cold gas, and nuclear propulsion. For communication satellites, chemical propulsion emerges as the most suitable option due to its simplicity and lower energy requirements. The chemical propulsion subsystem comprises oxidizer and fuel tanks, gas pressuring tanks utilizing helium. Wherein, Thrusters are employed for diverse tasks, encompassing tank sinking, orbital maneuvers (correction), attitude control, and deorbiting. These processes induce propellant consumption from orbit transfer to the deorbiting operation. The satellite's mission life hinges on propellant quantity, emphasizing the need to maintain sufficient reserves for deorbiting at satellite’s end of life. Thus, accurately estimating propellant mass becomes a crucial task. This work delves into propellant mass estimation methods, specifically Bookkeeping (BKP). Moreover, we introduce and test a developed tool based on the Bookkeeping method. This tool proves instrumental in estimating the remaining propellant, offering a valuable resource for satellite mission planning and longevity.

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H. A. Hamid, Ahmad, Zulkifli A. Ghaffar, and Nurulhanis C. Rus. "Roles of Atomizing Gas in Swirl Effervescent Atomization." International Journal of Engineering & Technology 7, no. 4.25 (2018): 24–28. http://dx.doi.org/10.14419/ijet.v7i4.25.22242.

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Issues of propellant atomizing, mixing and viscous loss become increasingly more important as the thrust chamber are reduced in size. The present investigation examines the behavior of resulting sprays emanating from swirl effervescent atomizers at various gas-to-liquid ratios (GLRs) and aeration tube configurations. A series of cold flow test has been conducted, where water and nitrogen were used as simulation fluids. Results show that the injection of atomizing gas tends to reduce the spray angle and the discharge coefficient. Results also indicate narrower spray angle and lower discharge coefficient at higher GLRs. A smaller total aeration hole size also leads to a narrower spray angle and a higher pressure drop for the gas injection. Interestingly, a smaller total aeration hole size produces higher discharge coefficient. In general, the atomizing gas has shown to significantly alter the resulting sprays of a swirl effervescent atomizer even at a relatively low GLR.Â

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Wang, Hongfu, Qinghua Zeng, Xianhe Chen, Zongyu Zhang, and Weide Liu. "Hybrid Modeling Method of SDR Interstage Valve Based on Mechanism and Data-Driven." International Journal of Aerospace Engineering 2022 (October 25, 2022): 1–15. http://dx.doi.org/10.1155/2022/2410681.

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Variable-throat adjustment is the most practical flow regulation method of solid ducted rocket ramjet (SDR). The high-fidelity mathematical model of the interstage valve is the basis for realizing high-precision gas flow and thrust regulation. In this paper, the complex effect of gas was divided into load and throat deformation effect. The load was mainly determined by the clearance, friction torque, and pneumatic torque that the valve was subjected to during operation. And the throat deformation was determined primarily by the deposition and ablation of the valve faced in the gas. Therefore, we could divide the valve model into three parts: the servo motor model, the load characteristic model, and the deformation model of the actual acting throat (referred to as the throat). Given, we have designed a cold-air experiment program, using cold air to equalize the valve load. Furthermore, we analyzed its mechanism of action and established the load model using the experimental data and neural network. Finally, the deformation mechanism of the throat was investigated, and simultaneously, the deformation model was shown based on the flight test data. Compared with the traditional interstage valve model, the model established in this paper is closer to the actual working conditions, which is helpful to carry out the more comprehensive and practical ground simulation. It has essential reference value for further realizing the precise regulation of gas flow.

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Mousavi, S. Fazlolah, Jafar Roshanian, and M. Reza Emami. "Quaternion-based attitude control design and hardware-in-the-loop simulation of suborbital modules with cold gas thrusters." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 229, no. 4 (2014): 717–35. http://dx.doi.org/10.1177/0954410014539294.

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Zivkovic, S., M. Milinovic, N. Gligorijevic, and M. Pavic. "Experimental research and numerical simulations of thrust vector control nozzle flow." Aeronautical Journal 120, no. 1229 (2016): 1153–74. http://dx.doi.org/10.1017/aer.2016.48.

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ABSTRACTRocket motor nozzle flow geometry is considered through its influence on the thrust vector control (TVC) performances. Extensive research is conducted using theoretical and software simulations and compared with experimental results. Cold and hot flow test equipments are used. The main objective of the research is to establish the methodology of flow geometry optimisation on the TVC hardware system. Several geometry parameters are examined in detail and their effects on the system performances are presented. The discovered effects are used as guidelines in the TVC system design process. A numerical method is presented for the determination of dynamic response time upper limit for the TVC system based on the gas flow dynamics performances.

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Lee, Ho-Sung, Do-Yoon Lee, Jong-Seung Park, and Joung-Keun Kim. "A Study on Pressure Control for Variable Thrust Solid Propulsion System Using Cold Gas Test Equipment." Journal of the Korean Society for Aeronautical & Space Sciences 37, no. 1 (2009): 76–81. http://dx.doi.org/10.5139/jksas.2009.37.1.076.

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