Space Propulsion Systems Research Articles

Experimental study and performance analysis of a dual-mode space propulsion system pressurized by electric pump

This study presents a new dual-mode propulsion system that utilizes an electrically pumped hydrogen peroxide and kerosene supply system, specifically designed for advanced spacecraft orbit and attitude control. The system incorporates two distinct electric pumps, one for hydrogen peroxide and another for kerosene, each engineered to optimize the propellant feed for the orbit control and attitude control engines. This configuration enables precise propellant management and thrust control, which are critical for the stringent demands of modern space missions. The dual-mode capability of the propulsion system allows for the use of hydrogen peroxide both as a monopropellant in the attitude control engine and as an oxidizer in combination with kerosene in the orbit control engine. This approach simplifies the propulsion architecture while maximizing the efficiency and utility of hydrogen peroxide as a propellant. The electric pumps are essential in providing high-pressure propellant delivery, which is necessary to achieve the desired combustion characteristics and thrust parameters. Performance evaluations of the propulsion system demonstrated significant results. The orbit control engine achieved a combustion chamber pressure of 5.9 MPa and a combustion efficiency of 98.42 %. The attitude control engine, operating at a peak pressure of 4.9 MPa, exhibited a stable pulsative mode, maintaining consistent pressure levels throughout its operational cycle. These performance metrics highlight the effectiveness of the electric pump technology in enhancing propulsion system responsiveness and efficiency. This propulsion system is expected to improve spacecraft maneuverability and reduce operational costs, making it suitable for long-duration space missions.

Confined Z-Pinch Magnetic Nozzle : A New Approach to Space Propulsion System that Utilizes Trapped Charged Particles

This paper provides a theoretical overview of a Z-pinch magnetic nozzle plasma propulsion system designed for Low Earth Orbit (LEO). The system utilizes trapped charged particles from the spacecraft's environment and magnetic confinement to generate thrust, reducing the need for onboard propellants. By employing a magnetic mirror configuration, the plasma is confined and accelerated to create propulsion. The proposed propulsion system is tailored to function in the dynamic space radiation environment of LEO, particularly in polar orbits, accounting for variations in solar activity that affect trapped protons and electrons. The mechanism involves converting magnetic plasma energy into directed kinetic energy, facilitating efficient plasma detachment and momentum transfer to the spacecraft. This study examines the feasibility, design overview, and constraints of implementing this propulsion technology. Key challenges include the variability of the space environment, technological complexities, radiation effects, plasma detachment efficiency, energy requirements, and material durability. Overcoming these challenges through innovative engineering and rigorous testing is essential for successful implementation. The Z-pinch magnetic nozzle plasma propulsion system promises to advance space propulsion technology, offering a more efficient and cost-effective method for satellite operations in LEO. With further research and development, it has the potential to support future satellite missions and space exploration initiatives.

Enhancement effect of Mo electrodes on the ignition characteristics of HAN-based gel propellant fuel and its combustion mechanisms

Hydroxylammonium nitrate (HAN) serves as a monopropellant fuel, renowned for its low toxicity and high-performance in space propulsion systems. HAN-based gel propellant may offer additional advantages, including long-term stability and excellent throttling capabilities. To better understand the ignition performance and combustion characteristics of this gel propellant, we prepared HAN-based gel propellant, and monitored its decomposition products using thermogravimetry (TG), Fourier Transform Infrared Spectroscopy (FTIR) and mass spectrometry (MS). Electrode materials of Mo, Cu, Ti and SUS 304 were fabricated to investigated their impact on ignition performance. It was found that the Mo electrode exhibited the shortest ignition delay time (ti) compared to Cu, Ti, and SUS 304. Meanwhile, the role of surface roughness of Mo electrode on the ignition delaytime (ti) and the ignition energy (Qi) were also studied. Our results revealed that, as the voltage increased from 100 V to 200 V, the fine-surfaced Mo electrode demonstrated a lower sensitivity to changes in ti compared to the rough-surfaced electrode. Additionally, the Qi of the rough-surfaced electrode decreased from 43.5 J to 16.7 J, while the Qi of the fine-surfaced electrode decreased from 52.5 J to 17.6 J. Combining a thermocouple, the combustion wave structure of the gel propellant was divided into pre-heating zone, condensed phase, electro chemical reaction zone, microexplosion-gases mixing zone and flame zone. Finally, the combustion model of HAN-based gel propellant was inferred to explain its combustion mechanism.

Atomic-scale investigation of the mechanisms of deformation-induced martensitic transformation at ultra-cryogenic temperatures

Liquefied natural gas storage and transportation as well as space propulsion systems have sparked interest in the martensitic transformation and behaviours of 316 L stainless steels (SS) under ultra-cryogenic deformation. In this study, high-resolution transmission electron microscopy (HRTEM) and molecular dynamics (MD) simulations were used to investigate the atomic arrangements and crystalline defects of deformation-induced γ-austenite → ε-martensite → α′-martensite and γ → α′ martensitic transformations in 316 L SS at 15 and 173 K. The γ → ε transformation involves the glide of Shockley partial dislocations on (111)γ planes without a change in atomic spacing. The formation of an α′ inclusion in a single ε-band is achieved by a continuous lattice distortion, accompanied by the formation of a transition zone of α′ and the expansion of the average atomic spacings due to dislocation shuffling. As α′ grows further into γ, the orientation relationship (OR) of the α′ changes by lattice bending. This process follows the Bogers-Burgers-Olson-Cohen model despite it not occurring on intersecting shear bands. Stacking faults and twins can also serve as nucleation sites for α′ at 173 K. We also found that direct transformation of γ → α′ occurs by the glide of 6aγ[112¯]/12 dislocations on every (111)γ plane with misfit dislocations. Overall, this study provides, for the first time, insights into the atomic-scale mechanisms of various two-step and one-step martensitic transformations induced by cryogenic deformation and corresponding local strain, enhancing our understanding of the role of martensitic transformation under ultra-cryogenic-temperature deformation in controlling the properties.

Calculation procedure for design parameters of transport channels of capillary in -tank devices for space propulsion systems propellant tanks

Calculation procedure for design parameters of transport channels of capillary in -tank devices for space propulsion systems propellant tanks

Удосконалення методів нагріву ксенону для запобігання потраплянню рідкої фази робочої речовини до системи подачі

Among the various types of electric propulsion, the Hall thruster type is becoming the most common. This is due to the fact that the use of a Hall thruster makes it possible to obtain high values of the thruster characteristics with a simple design compared to other types of space propulsion systems. For Hall electric propulsion thrusters, the main working substance is xenon because of its fairly high atomic weight, low ionization energy, and unreactiveness, which makes it possible to obtain high thruster characteristics with ease of operation. The use of xenon as a working substance features a peculiarity involving its critical temperature (289.74 K), which gives rise to the liquid phase in the tank and, accordingly, pressure jumps, thus making it impossible to use the xenon feed system. To exclude the ingress of the liquid phase of xenon into the accumulator tank in electric propulsion systems, heaters are placed on the xenon tank to maintain its temperature within a given range. However, this approach has the following disadvantages: the low thermal conductivity of composite tanks impairs heater-to-xenon heat transfer; warming up the whole of the tank before starting the thruster increases the thruster start-up preparation time; the continuous maintenance of the tank temperature increases energy consumption by the propulsion system; and it is impractical to maintain the temperature of the whole of the xenon, while only a few grams of it are consumed for one thruster start-up. The problem that was solved in this work consists in changing the approach to heating the working substance that enters the feed system. The analysis of literary sources showed that this problem is relevant and offers ways to improve existing methods. To solve this problem, theoretical calculations were carried out and verified by experiment. As a result, a method was proposed to calculate the gasifier so that it may maintain the temperature of the working substance entering the accumulator tank within the range from 293 K to 298 K, thus eliminating the possible ingress of the liquid phase of xenon into the accumulator tank of the feed system. This study allows one to use the proposed structural element (gasifier) instead of tank heaters, which significantly reduces power consumption and maintains the stable operation of the working substance feed system. The conclusions drawn from the study may be useful to most developers of storage and feed systems for electric propulsion systems.

Public–private partnerships in fostering outer space innovations

As public and private institutions recognize the role of space exploration as a catalyst for economic growth, various areas of innovation are expected to emerge as drivers of the space economy. These include space transportation, in-space manufacturing, bioproduction, in-space agriculture, nuclear launch, and propulsion systems, as well as satellite services and their maintenance. However, the current nature of space as an open-access resource and global commons presents a systemic risk for exuberant competition for space goods and services, which may result in a “tragedy of the commons” dilemma. In the race among countries to capture the value of space exploration, NASA, American research universities, and private companies can avoid any coordination failures by collaborating in a public–private research and development partnership (PPRDP) structure. We present such a structure founded upon the principles of polycentric autonomous governance, which incorporate a decentralized autonomous organization framework and specialized research clusters. By advancing an alignment of incentives among the specified participatory members, PPRDPs can play a pivotal role in stimulating open-source research by creating positive knowledge spillover effects and agglomeration externalities as well as embracing the nonlinear decomposition paradigm that may blur the distinction between basic and applied research.

Design Method for Automatic Assembly Production Line of Electric Valves in Space Propulsion Systems

This article proposes a design method for a valve automatic assembly production line in response to the automation assembly requirements of electric valve products in space propulsion systems and the engineering problems of inaccurate loading force control and low valve measurement accuracy in existing process methods. This method can achieve five assembly processes during the assembly process of electric valves, including pre-tightening force control, valve-core stroke measurement, performance testing, and shell structure welding. The article introduces the design of platform components such as process execution, positioning, and transportation, as well as the design and operation process of workstations. By combining the design of a three-axis motion mechanism, a small turntable, and a robotic arm, the product can achieve professional, positioning, full process automation, and equipment miniaturization design across multiple workstations. Through the design of precise control of loading force and non-contact optical measurement method of moving structure, compared with the original method, the parameters affecting product performance are precisely controlled and the precision is improved. And the multivariable decoupling of valve product performance is realized by this method. Through application verification, this automatic assembly production line can significantly improve the assembly efficiency of electric valve products and solve difficult problems in product engineering.

Measurement method and results of divergence angle of laser-controlled solid propellants used in space propulsion

The plume divergence angle is an important reference index for evaluating the thrust efficiency and propellant utilization of space propulsion systems. However, the characteristics of the dynamic variation of plume divergence angle over time cannot be measured using current methods. This paper utilizes high-speed photography and image processing methods to develop a strategy that can give a quick, non-destructive and real-time detection of the divergence angle. Effectiveness of the strategy is verified, and the characteristics of plume divergence angles of different laser-controlled solid propellants were further analyzed and fitted. The experimental results indicate that graphene could effectively reduce the divergence angle, while oxide-doped samples had larger divergence angles than alloy-doped and carbon-doped samples.

Potential negative effects of the Brazilian Space Program on coastal sharks

The Brazilian government recently announced its first partnerships with the private sector, including American and Canadian companies, to use the Alcântara Space Center (Maranhão, northern Brazil). This center is known for its privileged location, saving up to 30% of fuel in launches. Its operationalization is an offshoot of the Technological Safeguards Agreement, which is important for the Brazilian Space Program due to greater space sector investments and environmentally relevant projects. In 2003, a major fire at the Alcântara Space Center destroyed a rocket and killed 21 workers, halting Brazilian Space Program activities. Recently, our research group reported serious environmental consequences of this accident, i.e., extremely high amounts of Rubidium (Rb) in apex predatory sharks near the Alcântara Space Center. This element is used in fuels and in space propulsion systems and is potentially toxic, displaying bioaccumulating and biomagnifying capabilities. The observed concentrations are the highest ever detected in any living organism (up to 24.65 mg kg-1 dry weight). The launch base is located on the Brazilian Amazon Coast, and population recruitment impacts may compromise biota conservation and biodiversity. Local shark meat consumption is also worrying, as consumers may be exposed to Rb, whose effects in humans are unknown. We, therefore, indicate an urgent need for biomonitoring efforts in the area, as the Alcântara Space Center is about to operate at its maximum capacity.

Cluster of electric thrusters for astronautic and robotic INPPS flagship space flights to Mars and Europa moon

This review deals with the selection of the electric propulsion system (EPS) for the internationally developed and designed, primary nuclear-electric space tug International Nuclear Power and Propulsion System (INPPS). INPPS is scheduled for interplanetary missions to Mars and Jupiter moon Europa missions by the end of decade 2020. Regarding specific technical and mission parameters preselected electric thruster (ET) types, developed by international companies and institutions, are analysed, evaluated and investigated for a possible application as propulsion system (PS), the so-called CET (Cluster of Electric Thrusters). It is analysed whether solely electric thrusters, combined in an adequate CET, enable the envisaged interplanetary missions—robotic and astronautic/crewed with the INPPS flagship.Thruster clusters with strategic consortium considerations are analysed as a feasible PS of the INPPS. The studied CET consists of the following: (a) only European ETs, (b) combination of German and European ETs, (c) Japanese and European ETs or at least (d) Japanese, European and US thrusters. The main results are (1) Robotic and crewed INPPS mission to Mars/Europa are realizable with EPS only (no chemical propulsion is needed), (2) that every CET, except (c) of only Japanese and part of European thrusters, is capable to perform the main part of envisaged INPPS flagship mission orbit to Mars, back to Earth and to Jupiter/Europa moon.

Propulsion Systems, Propellants, Green Propulsion Subsystems and their Applications: A Review

A wide range of propellants, and propulsion systems in space exploration by aircrafts or space vehicles was studied, developed, investigated, and commercialized. Liquid, solid, or hybrid propellants have been used for rocket’s launches. In this review, a consistent definition of space propulsion systems, including solid, liquid and hybrid has been given with up-to-date state of developments. A comparison of their performances was made by theoretical and experimental specific impulses. On the other hand, ammonium perchlorate and hydrazine were used as propellants for rocket’s launches and for satellite’s maneuverings; respectively. However, their high toxicity and their storage problem pushed researchers and scientists to investigate and develop other eco-friendly, propellant systems, so called “green propellants”, for launch or reaction control systems of satellites.

New model offers 3D insights into plasma-based space propulsion systems

A 3D model of Hall Thrusters more accurately represents electron drift instabilities.

Thrust Performance of Converging Rotating Detonation Engine Compared with Steady Rocket Engine

Rotating detonation engines (RDEs) have been actively researched around the world for application to next-generation aerospace propulsion systems because detonation combustion has theoretically higher thermal efficiency than conventional combustion. Moreover, because cylindrical RDEs have simpler combustors, further miniaturization of conventional combustors is expected. Therefore, in this study, with the aim of applying RDEs to space propulsion systems, a cylindrical RDE with a converging–diverging nozzle was manufactured; the combustor length was changed to 0, 10, 30, 50, and 200 mm; and the thrust performance and combustion mode with the different combustor lengths were compared. As a result, four combustion modes were confirmed. Detonation combustion occurred with a combustor length of : that is, a converging rotating detonation engine. The thrust performance of this engine was 94 to 100% of the theoretical rocket thrust performance, which is equivalent to the thrust performance of conventional rocket combustion generated at . This study shows that detonation combustion can significantly reduce engine weight while maintaining thrust performance.

A data-driven approach for analyzing Hall thruster discharge instability leading to plasma blowoff

The Hall-effect thruster is a prominent space propulsion system, providing a high specific impulse and thrust-to-power ratio. However, detrimental phenomena occurring in Hall thrusters, such as plasma instability and blowoff, are less understood. In this study, we employ a data-driven approach for analyzing the blowoff phenomenon of the Hall-effect propulsion system. From a 600-W class Hall thruster, we observe the oscillatory behavior of the anode current with a frequency of 34–38 kHz, which sometimes leads to plasma blowoff. We employ an output-only system identification framework that uses the noise-induced dynamics at this frequency, extracting linear and nonlinear coefficients of the low-order system model. The computed coefficients suggest that the plasma blowoff is characterized by large linear and small nonlinear parameters of the stochastic Van der Pol equation. From the obtained results, we conclude that this framework can be exploited either to identify a reliable operating condition with a low possibility of plasma blowoff or to diagnose the blowoff susceptibility of the thruster. This research constitutes the first application of stochastic system identification to the Hall-effect thruster, opening new possibilities for avoiding unwanted oscillations and blowoff phenomena in electric propulsion systems.

静圧と動圧を考慮した磁気プラズマセイルの電磁流体解析

Magsail is a space propulsion system using the interactions between the solar wind and the magnetic field generated by the onboard coils. Magnetoplasma sail is a propulsion system that increases thrust by expanding the magnetosphere through plasma injection from the spacecraft. There are two mechanisms on the magnetospheric inflation: method using frozen-in of magnetic field to carry magnetic field lines by high dynamic pressure plasma and method using the diamagnetic current by thermal plasma, which is called the ring current. We investigated the effect of the dynamic pressure and thermal pressure on the MPS thrust performance used electromagnetic hydrodynamic simulation. It was shown that the ring current is enhanced by adding dynamic pressure to the thermal plasma and increases thrust gain. The high thrust gain over 2.25 was obtained at βth = 0.5 - 2 and βk = 4 - 8. However, the thrust is reduced because the super magneto acoustic wave region is generated in the magnetosphere, which prevents the propagation of thrust in large β conditions. The wide parameter survey reveals injection plasma parameter regions where thrust reduction is restrained and high thrust gain is obtained.

О практических методах определения порометрических характеристик тонких образцов из комбинированного пористого сетчатого материала

On the example of a porous structure operation in the gas-liquid flow as the phase separator, the paper demonstrates the porometric characteristics effect to assess its performance in the capillary-type intra-tank devices of the space propulsion system fuel tanks. It is shown that for manufacture of a permeable phase separator taking into account the stable capillary structure and the possibility to obtain the hydraulic resistance given value, it is advisable to introduce a combined porous mesh material. For practical determination of one of the porous structure characteristics, i.e. the capillary holding capacity, a method is presented for preparing the permeable structure surface of the phase separator material for testing based on the “bubble” technique, and principles of using such wetting liquids as water, glycerin or ethyl alcohol are discussed. Taking into account small thicknesses of the phase separators made of combined porous materials, types of the circuit solutions of the test installations are provided for determining the capillary holding capacity of porous mesh materials differing from the traditional ones used in orosimetry of soils, rocks and filter materials obtained on the basis of powder metallurgy.

Підхід до чисельного моделювання просторових рухів газорідинного сере-довища в паливному баку космічного ступеня в умовах мікрогравітації з урахуванням гарячої зони

Space propulsion systems ensure multiple startups and shutdowns of the main liquid-propellant rocket engines in microgravity conditions for spacecraft preset motions and reorientation control. During the passive flight of a space stage (after its main engine shutdown), the liquid propellant in the tanks continues moving by inertia in microgravity and moves as far away from the propellant management device as possible. In this case, the pressurization gas is displaced to the propellant management device, which creates the potential danger of the gas entering the engine inlet in quantities unacceptable for multiple reliable engine restarts. In this regard, the determination of the parameters of fluid movement in propellant tanks under microgravity conditions is a pertinent problem to be solved in the designing of liquid-propellant propulsion systems. This paper presents an approach to the theoretical calculation of the parameters of motion of the gas–liquid system in the propellant tanks of today’s space stages in microgravity conditions. The approach is based on the use of the finite element method, the Volume of Fluid method, and up-to-date computer tools for finite-element analysis (Computer Aided Engineering - CAE systems). A mathematical simulation of the spatial motion of the liquid propellant and the formation of free gas inclusions in passive flight was performed, and the motion parameters and shape of the free liquid surface in the tank and the location of gas inclusions were determined. The liquid motion in a model spherical tank in microgravity conditions was simulated numerically with and without account for the hot zone near the tank head. The motion parameters of the gas-liquid interface in a model cylindrical tank found using the proposed approach are in satisfactory agreement with experimental data. The proposed approach will significantly reduce the extent of experimental testing of space stages under development.

Diversity of Physical Processes: Challenges and Opportunities for Space Electric Propulsion

The renewed interest in space exploration has led to the growth in research efforts pertaining to advanced space propulsion systems, including highly efficient electric propulsion systems. Although already tested in space many decades ago and being currently employed on various space platforms and thousands of satellites, these systems are yet to reach their full potential for applications on orbit and in deep space. One specific feature of space electric propulsion is the large diversity of physical processes used in this technology, which is not typical for many other types of propulsion systems used in transport, such as those used by airplanes or automobiles. Various physical processes and mechanisms underpin different electric propulsion technologies and should be integrated to drive the future science and technology of space electric propulsion systems. This opinion article briefly highlights this feature of space electric propulsion and outlines some challenges and opportunities that follow from this diversity.

Validation of a torsional balance for thrust measurements of Hall effect and microwave-based space propulsion systems.

A torsional thrust balance has been designed and validated by Surrey Space Centre and Added Value Solutions UK Ltd. in collaboration with the UK Space Agency. The thrust stand has been tested with two electric propulsion (EP) systems operating with xenon: the Halo thruster and the XJET thruster. The first consists of a low-power (<1 kW) Hall effect-based thruster, whose thrust level is between 3 and 20mN, depending on the power of the system. The second is an electron cyclotron resonance thruster whose operative point is in the 0.3-1.5mN thrust range. The thruster is mounted on a titanium rotating beam, whose movement is measured by an optical fiber displacement sensor. The thrusters' direct current electrical connections are routed through room temperature liquid metal pots and microwave power is transmitted via a wireless transfer system, minimizing friction effects. To reduce thermal issues during long thruster operations, the torsional thrust balance is designed with a water-cooling hub around the flex pivot. Noise from the laboratory environment is lessened by using four vibration-dampening spring systems as thrust balance feet. The tests on the two EP systems have shown accurate and repeatable results, demonstrating that the balance can be used to characterize different EP systems in the μN-mN thrust range.