Showing papers on "Spacecraft propulsion published in 2021"

Electric Propulsion Methods for Small Satellites: A Review

Abstract: Over 2500 active satellites are in orbit as of October 2020, with an increase of ~1000 smallsats in the past two years. Since 2012, over 1700 smallsats have been launched into orbit. It is projected that by 2025, there will be 1000 smallsats launched per year. Currently, these satellites do not have sufficient delta v capabilities for missions beyond Earth orbit. They are confined to their pre-selected orbit and in most cases, they cannot avoid collisions. Propulsion systems on smallsats provide orbital manoeuvring, station keeping, collision avoidance and safer de-orbit strategies. In return, this enables longer duration, higher functionality missions beyond Earth orbit. This article has reviewed electrostatic, electrothermal and electromagnetic propulsion methods based on state of the art research and the current knowledge base. Performance metrics by which these space propulsion systems can be evaluated are presented. The article outlines some of the existing limitations and shortcomings of current electric propulsion thruster systems and technologies. Moreover, the discussion contributes to the discourse by identifying potential research avenues to improve and advance electric propulsion systems for smallsats. The article has placed emphasis on space propulsion systems that are electric and enable interplanetary missions, while alternative approaches to propulsion have also received attention in the text, including light sails and nuclear electric propulsion amongst others.

In-orbit demonstration of an iodine electric propulsion system.

Abstract: Propulsion is a critical subsystem of many spacecraft1–4. For efficient propellant usage, electric propulsion systems based on the electrostatic acceleration of ions formed during electron impact ionization of a gas are particularly attractive5,6. At present, xenon is used almost exclusively as an ionizable propellant for space propulsion2–5. However, xenon is rare, it must be stored under high pressure and commercial production is expensive7–9. Here we demonstrate a propulsion system that uses iodine propellant and we present in-orbit results of this new technology. Diatomic iodine is stored as a solid and sublimated at low temperatures. A plasma is then produced with a radio-frequency inductive antenna, and we show that the ionization efficiency is enhanced compared with xenon. Both atomic and molecular iodine ions are accelerated by high-voltage grids to generate thrust, and a highly collimated beam can be produced with substantial iodine dissociation. The propulsion system has been successfully operated in space onboard a small satellite with manoeuvres confirmed using satellite tracking data. We anticipate that these results will accelerate the adoption of alternative propellants within the space industry and demonstrate the potential of iodine for a wide range of space missions. For example, iodine enables substantial system miniaturization and simplification, which provides small satellites and satellite constellations with new capabilities for deployment, collision avoidance, end-of-life disposal and space exploration10–14. The successful in-orbit operation of an electric space propulsion system based on iodine, rather than the more expensive and difficult-to-store xenon, is demonstrated.

Review of state-of-the-art green monopropellants: For propulsion systems analysts and designers

Abstract: Current research trends have advanced the use of “green propellants” on a wide scale for spacecraft in various space missions; mainly for environmental sustainability and safety concerns. Small satellites, particularly micro and nanosatellites, evolved from passive planetary-orbiting to being able to perform active orbital operations that may require high-thrust impulsive capabilities. Thus, onboard primary and auxiliary propulsion systems capable of performing such orbital operations are required. Novelty in primary propulsion systems design calls for specific attention to miniaturization, which can be achieved, along the above-mentioned orbital transfer capabilities, by utilizing green monopropellants due to their relative high performance together with simplicity, and better storability when compared to gaseous and bi-propellants, especially for miniaturized systems. Owing to the ongoing rapid research activities in the green-propulsion field, it was necessary to extensively study and collect various data of green monopropellants properties and performance that would further assist analysts and designers in the research and development of liquid propulsion systems. This review traces the history and origins of green monopropellants and after intensive study of physicochemical properties of such propellants it was possible to classify green monopropellants to three main classes: Energetic Ionic Liquids (EILs), Liquid NOx Monopropellants, and Hydrogen Peroxide Aqueous Solutions (HPAS). Further, the tabulated data and performance comparisons will provide substantial assistance in using analysis tools-such as: Rocket Propulsion Analysis (RPA) and NASA CEA-for engineers and scientists dealing with chemical propulsion systems analysis and design. Some applications of green monopropellants were discussed through different propulsion systems configurations such as: multi-mode, dual mode, and combined chemical-electric propulsion. Although the in-space demonstrated EILs (i.e., AF-M315E and LMP-103S) are widely proposed and utilized in many space applications, the investigation transpired that NOx fuel blends possess the highest performance, while HPAS yield the lowest performance even compared to hydrazine.

Development of Green Storable Hybrid Rocket Propulsion Technology Using 98% Hydrogen Peroxide as Oxidizer

Abstract: This paper presents the development of indigenous hybrid rocket technology, using 98% hydrogen peroxide as an oxidizer. Consecutive steps are presented, which started with interest in hydrogen peroxide and the development of technology to obtain High Test Peroxide, finally allowing concentrations of up to 99.99% to be obtained in-house. Hydrogen peroxide of 98% concentration (mass-wise) was selected as the workhorse for further space propulsion and space transportation developments. Over the course nearly 10 years of the technology’s evolution, the Lukasiewicz Research Network—Institute of Aviation completed hundreds of subscale hybrid rocket motor and component tests. In 2017, the Institute presented the first vehicle in the world to have demonstrated in-flight utilization for 98% hydrogen peroxide. This was achieved by the ILR-33 AMBER suborbital rocket, which utilizes a hybrid rocket propulsion as the main stage. Since then, three successful consecutive flights of the vehicle have been performed, and flights to the Von Karman Line are planned. The hybrid rocket technology developments are described. Advances in hybrid fuel technology are shown, including the testing of fuel grains. Theoretical studies and sizing of hybrid propulsion systems for spacecraft, sounding rockets and small launch vehicles have been performed, and planned further developments are discussed.

Experimental Investigation of Rotating Detonation Rocket Engines for Space Propulsion

Abstract: The performance (specific impulse, Isp) of a modular, 150-lbf-class rotating detonation rocket engine (RDRE) was measured with three gaseous fuels (methane, ethane, and ethylene) and gaseous oxygen...

Hybrid Rocket Engine Design Optimization at Politecnico di Torino: A Review

Abstract: Optimization of Hybrid Rocket Engines at Politecnico di Torino began in the 1990s. A comprehensive review of the related research activities carried out in the last three decades is here presented. After a brief introduction that retraces driving motivations and the most significant steps of the research path, the more relevant aspects of analysis, modeling and achieved results are illustrated. First, criteria for the propulsion system preliminary design choices (namely the propellant combination, the feed system and the grain design) are summarized and the engine modeling is presented. Then, the authors describe the in-house tools that have been developed and used for coupled trajectory and propulsion system design optimization. Both deterministic and robust-based approaches are presented. The applications that the authors analyzed over the years, starting from simpler hybrid powered sounding rocket to more complex multi-stage launchers, are then presented. Finally, authors’ conclusive remarks on the work done and their future perspective in the context of the optimization of hybrid rocket propulsion systems are reported.

Mechanically Amplified Milli-Newton Thrust Balance for Direct Thrust Measurements of Electric Thrusters for Space Propulsion

Abstract: Direct thrust measurements by means of a thrust balance (TB) are the golden standard for measuring thrust and, concurrently, the specific impulse in electric thrusters. To measure these properties in the novel class of electrodeless plasma thrusters, a new TB based on the Variable Amplitude Hanging Pendulum with Extended Range (VAHPER) concept has been developed. The TB has a mechanical amplification mechanism with an angular magnification of $31^\circ /^\circ $ . Using Lagrangian mechanics, we show that the TB loaded with a 5.2-kg thruster prototype behaves like a damped harmonic oscillator with a natural frequency of 0.37 Hz. A variable damping system provides damping with an optimal damping ratio of 0.78, which corresponds to a settling time of only 1.8 s. Both the model and the damping and calibration system have been validated. To accommodate the particularities of medium power electrodeless plasma thrusters, the TB design includes the following features: an optical displacement sensor, water cooling, liquid metal connectors, and dedicated vacuum-rated electronics for autoleveling, remote (in-vacuum) calibration, and temperature monitoring. To test the TB, measurements were performed on a 500-W Helicon Plasma Thruster breadboard model. When loaded with this thruster, the measured stiffness of the system was 12.67 ± 0.01 mN/mm. For this stiffness, the thrust range is 150 mN with a 0.1-mN resolution. The relative uncertainty on the thrust measurements is found to be on the order of 2%.

Thrust measurements and evaluation of asymmetric infrared laser resonators for space propulsion

Abstract: Since modern propulsion systems are insufficient for large-scale space exploration, a breakthrough in propulsion physics is required. Amongst different concepts, the EMDrive is a proposed device claiming to be more efficient in converting energy into propulsive forces than classical photon momentum exchange. It is based on a microwave resonator inside a tapered cavity. Recently, Taylor suggested using a laser instead of microwaves to boost thrust by many orders of magnitude due to the higher quality factor of optical resonators. His analysis was based on the theory of quantised inertia by McCulloch, who predicted that an asymmetry in mass surrounding the device and/or geometry is responsible for EMDrive-like forces. We put this concept to the test in a number of different configurations using various asymmetrical laser resonators, reflective cavities of different materials and size as well as fiber-optic loops, which were symmetrically and asymmetrically shaped. A dedicated high precision thrust balance was developed to test all these concepts with a sensitivity better than pure photon thrust, which is the force equivalent to the radiation pressure of a laser for the same power that is used to operate each individual devices. In summary, all devices showed no net thrust within our resolution at the Nanonewton range, meaning that any anomalous thrust must be below state-of-the-art propellantless propulsion. This puts strong limits on all proposed theories like quantised inertia by at least 4 orders of magnitude for the laboratory-scale geometries and power levels used with worst case assumptions for the theoretical predictions.

Modular impulsive green monopropellant propulsion system (Mimps-g) : For cubesats in leo and to the moon

Abstract: Green propellants are currently considered as enabling technology that is revolutionizing the development of high-performance space propulsion, especially for small-sized spacecraft. Modern space missions, either in LEO or interplanetary, require relatively high-thrust and impulsive capabilities to provide better control on the spacecraft, and to overcome the growing challenges, particularly related to overcrowded LEOs, and to modern space application orbital maneuver requirements. Green monopropellants are gaining momentum in the design and development of small and modular liquid propulsion systems, especially for CubeSats, due to their favorable thermophysical properties and relatively high performance when compared to gaseous propellants, and perhaps simpler management when compared to bipropellants. Accordingly, a novel high-thrust modular impulsive green monopropellant propulsion system with a micro electric pump feed cycle is proposed. MIMPS-G500mN is designed to be capable of delivering 0.5 N thrust and offers theoretical total impulse Itot from 850 to 1350 N s per 1U and >3000 N s per 2U depending on the burnt monopropellant, which makes it a candidate for various LEO satellites as well as future Moon missions. Green monopropellant ASCENT (formerly AF-M315E), as well as HAN and ADN-based alternatives (i.e., HNP225 and LMP-103S) were proposed in the preliminary design and system analysis. The article will present state-of-the-art green monopropellants in the (EIL) Energetic Ionic Liquid class and a trade-off study for proposed propellants. System analysis and design of MIMPS-G500mN will be discussed in detail, and the article will conclude with a market survey on small satellites green monopropellant propulsion systems and commercial off-the-shelf thrusters.

Mechanism analysis and experimental verification of electromagnetic sail, a new solar propulsion system without propellent

Abstract: The electromagnetic sail, a new space propulsion system without propellant, which propels a spacecraft by deflecting solar wind with an electromagnetic field, is proposed. The thrust mechanism and the performance variations of the electromagnetic sail are analyzed with both particles in cell and magnetohydrodynamic models, and the interaction between the incoming plasma fluid and the electromagnetic field is investigated by tests on the ground experimental system. The simulation results demonstrate that a magnetosphere is created in front of the sail and its area increases with magnetic field intensity; with appropriate dipole moment, the thrust of the electromagnetic sail is higher than a pure magnetic sail or a pure electric sail. In test experiments, the plume structure of a helicon plasma source changes obviously with coil current and metal mesh voltage, which verifies the repulsion of the magnetic field and electric field to the incoming plasma. These results indicate that the combination of the magnetic sail and electric sail can increase the utilization efficiency of solar wind, which means the electromagnetic sail is a feasible propulsion system.

Design and Testing of a Propellant Management System for Bimodal Chemical-Electrospray Propulsion

Abstract: The most prominent modes of propulsion for spacecraft are chemical and electric propulsion, and missions that require maneuvering need to select one mode over the other, or in some cases need to carry two separate propulsion subsystems, including management of incompatible propellants. This is particularly limiting for small spacecraft, which lack the volume, mass and power resources to accommodate high-performance propulsion components. Most systems are then confined to either the chemical mode (typically high thrust, low specific impulse) or the electric mode (typically low thrust, high specific impulse). However, recent development of green monopropellants as replacements to hydrazine in chemical thrusters has created a family of propellants that are also compatible with electric thrusters. Notably, the monopropellant AF-M315E/ASCENT and other hydroxylammonium nitrate (HAN) based ionic liquids can be used in electrospray thrusters. This work proposes a bimodal propulsion system that integrates a chemical mono-propellant thruster with electrospray thrusters into a unified system with common propellant. The design, fabrication, and validation of a propellant feed line connecting the central tank to electrospray thrusters is presented here. Key elements of this design include solenoid valves for flow control and capillary tubes for pressure conditioning. A subsequent section of nonconductive tubing allows for water (present in the composition of many HAN-based propellants) to evaporate at its vapor pressure and form bubbles in the propellant line. These bubbles physically segment the conductive ionic liquid and provide electrical isolation between the electrospray thruster and the central tank; this is required to apply high voltage to operate the electrospray thrusters without current leaks or shorts. A prototype propellant line was fabricated and its flow characteristics and electrical impedance were measured in a series of tests to validate the design. These tests successfully demonstrated the feasibility of this design for a bimodal spacecraft propulsion system.

A Computational Tool for the Design of Hybrid Rockets

Abstract: A computational tool able to perform a fast analysis of hybrid rocket engines is presented, describing briefly the mathematical and physical models used. Validation of the code is also shown: 16 different static firing tests available in the open literature are used to compare measured operational parameters such as chamber pressure, thrust, and specific impulse with the code’s output. The purpose of the program is to perform rapid evaluation and assessment on a possible first design of hybrid rockets, without relying on computationally expensive simulations or onerous experimental tests. The validated program considers as benchmark and study case the design of a liquid-oxygen/paraffin hybrid rocket engine to be used as the upper stage of a small launcher derived from VEGA building blocks. A full-factorial parametric analysis is performed for both pressure-fed and pump-fed systems to find a configuration that delivers the equivalent total impulse of a VEGA-like launcher third and fourth stage as a first evaluation. This parametric analysis is also useful to highlight how the oxidizer injection system, the fuel grain design, and the nozzle features affect the performance of the rocket.

Acceleration characteristics of laser ablation Cu plasma in the electrostatic field

Abstract: As a new concept of space propulsion system, laser-ablation propulsion has attracted more and more attention due to its characteristics of low power consumption, high specific impulse, variable and controllable thrust. With an aim to further raise up the movement velocity of plasma, we combine the laser with high-voltage electrostatic field to accelerate the Cu plasma induced by laser ablation. To demonstrate the acceleration characteristics of plasma under different electric field intensity, the plasma conductivity, plasma shockwave intensity and plasma plume movement process were tested using parallel electrode plate device, self-made torsion pendulum impulse test bench and high-speed ICCD camera. The results showed that the conductive current and impulse formed by the plasma obviously increased under the applied electric field. The images captured by high-speed ICCD camera showed the plasma cross-sectional area was 0.194 mm2 at 900 ns and 0.217 mm2 at 1600 ns when the electric field intensity was 0 V/mm. With the electric field intensity increased to 30 V/mm, the plasma cross-sectional area elevated to 0.280 mm2 at 900 ns and 0.288 mm2 at 1600 ns. The acquisitions prove that the idea of this paper is feasible and favorable, which provide a theoretical basis for the combination of laser ablation propulsion and electric field.

Optimization of Regression Rate of Solid Fuel Grain in Hybrid Rocket Motor Physical Method

Abstract: Rocket propulsion unit is very essential to for launching a space mission. Rocket propulsion has various types. In that hybrid rocket propulsion is an emerging and innovative technology with strong features such as economical, safe, and simple propulsion devices. On the other hand, few issues such as mixing ratio changes, low regression rates and inefficient mixing, restricts their application as a the major area. This research represents an improvement of an important issue, namely the low fuel regression rate. If an oxidizer velocity increases at the surface of fuel, the regression rate of fuel may increase due to the oxidizer flow above the burning surface produces shear stress, which erodes the fuel surface and causing the missing of fuel and air. Methods to increase the shear stress are to induce an oxidizer vortex inside the combustion chamber. In the current analysis, a diaphragm has been used to create the air vortex inside the solid fuel grain. A sequence of static rocket firing was accompanied for different position diaphragm. Operating condition has been kept constant for all the run. The results of six laboratory scale hybrid rocket test firing using polyvinyl chloride and gaseous oxygen are presented here. It has been found that the regression rate of solid fuel grain has been improved for the diaphragm cases.

Design and development of iodine flow control systems for miniaturized propulsion systems

Abstract: The high density of iodine, and its storage in solid form with no pressurization requirements, have driven the increase on the use of the halogen for space propulsion systems. This paper presents the design and development process of a miniaturized propellant management system which can supply iodine at a stable mass flow rate to the thrusters. This system approach has been adapted to produce the flow control units of the NPT30-I2 gridded ion propulsion system and the I2T5 cold gas thruster, both of which have performed their first in-orbit demonstration flights. The paper is dedicated to the description of the design process, with a focus on the particularities of the use of iodine for space propulsion systems, mainly related to chemical interaction and deposition. The numerical analysis of rarefied flows serves as a base for the design of the flow path elements of the system. A performance characterization of the system obtained on ground and the challenges tackled during the development process are evaluated with the in-orbit data from the first flight missions of both propulsion systems.

Is laser space propulsion practical?: review.

Abstract: In this paper, we review practical limitations to laser space propulsion that have been discussed in the literature. These are as follows: (1) thermal coupling to the propelled payload, which might melt it; (2) a decrease in mechanical coupling with number of pulses, which has been observed in some cases; and (3) destruction of solar panels in debris removal proposals that might create more debris rather than less. Previously, lack of data prevented definite assessments. Now, new data on multipulse vacuum laser impulse coupling coefficient Cm on several materials at 1064 nm, at 1030 nm, and at 532 nm are available. We are now able to compare the results for single and multiple pulses on materials that have been considered for laser ablation space propulsion (LASP), or that are likely space debris constituents, and decide whether LASP is a practical idea. Laser space propulsion and debris removal concepts depend on thousands or hundreds of thousands of repetitive pulses. Repetitive pulse mechanical coupling as well as thermal coupling (which can melt the target rather than propel it) are both important considerations. Materials studied were 6061T6 aluminum, carbon-doped polyoxymethylene (POM), undoped POM, a yellow POM copolymer, and a mixture of Al and POM microparticles combined and pressed, containing a 50%/50% mixture of the two materials by mass. We address 6 and 70 ps pulses because of the availability of data at these pulse durations. We also briefly consider continuous wave (CW) laser propulsion. Finally, we consider a recent paper concerning solar panel destruction from a positive perspective.

Preliminary Investigation of Robust Reinforcement Learning for Control of an Existing Green Propellant Thruster

Abstract: Advanced engine control is an important requirement for the efficient operation of future reusable engines, facilitating a safer and more economical engine operation. For this reason, modern control strategies are extensively studied in recent years, mainly through the use of simulation environments. An important development step is to test the performance and robustness of the control algorithm at real test benches. The present paper describes the first steps towards the use of an reinforcement learning based controller on a N2O / C2H6 22 N green propellant thruster. The control objectives are given by regulating the mixture ratio and combustion pressure. The existing test bench is modelled in EcosimPro / ESPSS. Based on the simulation model deep reinforcement learning is used to train the controller and domain randomization is used to increase the robustness. The overall goal is to transfer the controller from the simulation model to the real test bench. Finally, preliminary experiments demonstrate the basic functionality of reinforcement learning based controllers for real rocket propulsion systems.

Plasma Thrusters for In-Space Propulsion; New Trends and Physical Limitations

Abstract: In-space plasma propulsion is today a necessity for the economic competitiveness of commercial satellites. In plasma thrusters the impulse is imparted by a mesothermal plasma jet where ions drift a supersonic velocities. Compact multi-electrode systems for ion acceleration are studied by a combination of analytical calculations and computer simulations. The results show exahust speeds of 140 km/s can be reached with specific impulses higher than unsual plasma thrusters. Additionally, the ion velocity distribution functions of the plasma jet exhasut obtained with retarded field energy analyzers that can be used to cross validate the results of numerical simulations.

Revolutionizing Spaceflight: A Study on Electric Propulsion and Air-Breathing MPDs

Abstract: Modern liquid-fuel rocket propulsion harbors a number of great limitations. Among those is the weight of fuel, which makes up more than 90% of the mass of the SpaceX Falcon 9 (NASA, 2018). Electric propulsion has been used for decades as an alternative to liquid-fuel rockets due to low propellant requirements and high specific impulse. Although electric thrusters have strictly been used in non-atmospheric conditions, recent innovations attempt to expand its use to airspace. This quasi-experimental study focuses on the creation of an air-breathing magnetoplasmadynamic (MPD) thruster, with attempts being made to maximize the efficiency of the engine. Immense safety concerns prevented testing from occurring after the engine was built. However, the estimated performance of the built MPD is compared to a multitude of existing forms of electric propulsion, from Hall-effect thrusters to electrodynamic tethers. The concluding evidence suggests that air-breathing MPDs are not currently viable, high-power photon thrusters being of greater use in atmospheric conditions. Further research focusing on decreasing atmospheric breakdown voltage and increasing mirror reflectance of photon thrusters is suggested.