Small Spacecraft Research Articles

Analysis of Dynamics and Control During the Deployment of an Annular Tether Group of Spacecraft

The article considers a method for forming a rotating tether group of small spacecraft in the form of a regular polygon. To create an external torque acting on the entire system as a whole, low-thrust engines are used, and the directions of the reactive forces are unchanged in the coordinate systems associated with the spacecraft. The release of the cables is controlled by measuring their length and speed in accordance with the nominal program, which assumes that the mechanisms of the release of the cables work only for their braking. The nominal program is built according to a simplified model of the system's motion, built using the Lagrange equations. The feasibility of the deployment program is checked using a more complete model of the spatial motion of the system, written in a geocentric fixed coordinate system and taking into account the extensibility of the cables, the one-sidedness of the mechanical links corresponding to them, the operation of the control system, disturbances associated with the initial conditions of motion, etc. A more complete model of motion takes into account the motion of space vehicles relative to their centers of mass as solid bodies of finite dimensions, which leads to perturbations in the directions of action of reactive forces. Numerical examples are given of the deployment of ring tether groups of small spacecraft in the form of polygons with up to seven vertices, inclusive, under the action of disturbances.

Improving the technology for ensuring the magnetic cleanliness of small spacecraft

The object of this study is the technology of ensuring magnetic cleanliness when applied to small spacecraft (SC) with overall dimensions of ~1 m in terms of minimizing the effect of magnetic interference on the on-board magnetometer. As part of solving the general problem of increasing the reliability of spacecraft, the task of improving the technology of magnetic cleanliness is considered. It is shown that the use of the multi-dipole model for the calculation of ~50 nT magnetic interference is limited in the design of small space vehicles that have an increased installation density. The expediency of using a model of its spherical harmonics instead of a multi-dipole model to represent the magnetic field of satellite components was theoretically justified. For the practical application of the model, the budgeting of the projections of the magnetic induction of hindrance to the onboard magnetometer according to the experimentally determined coefficients of the dipole, quadrupole and octupole spherical harmonics is proposed. An algorithm for calculating the coordinates of the location and projections of the magnetic moment of the dipole source inside the satellite based on the results of measuring the coefficients of spherical harmonics of its components is proposed. The possibility of representing the coefficients of the spherical harmonics of the magnetic field of the entire satellite in the form of the sum of the listed corresponding coefficients obtained during field measurements of the components is theoretically justified. It is proposed to use the difference between the calculated and measured values of the corresponding coefficients as a criterion for the quality of the work to ensure magnetic cleanliness. If the proposed procedure is used in the space industry, the quality of ensuring the magnetic cleanliness of small space vehicles can be increased, thereby improving the reliability of their operation in orbit

The main directions of research and development of BSTU «VOENMEH» named after D. F. Ustinov on the subject of small spacecraft and some results

The development of small spacecrafts requires the solution of a number of new scientific and technical problems. The main directions of research and development of the Baltic State Technical University «VOENMEH» named after D. F .Ustinov on solving the problems of creating small spacecrafts, deploying and maintaining orbital constellations of the small spacecrafts, and managing orbital groupings are presented. Some results of these works are presented, in particular, the preliminary design of a small space tug with an innovative two-component gas rocket engine powered by oxygen and hydrogen, operated in conjunction with an ultralight class launch vehicle. Small space tug solves the problems of interorbital transportation of the small spacecrafts. It has no analogues in the russian rocket and space technology. Variants of ballistically coupled groups of small spacecrafts are presented that can ensure the formation of a stable quasi-stationary distributed aperture in orbits in order to increase the efficiency of space communication systems and remote sensing of the Earth. The development of an innovative actuator – a two-axis rotary platform based on piezoelectric drives is presented. This device significantly exceeds traditional electromechanical devices based on electric drives in terms of weight, dimensions, dynamic characteristics and other parameters. The variants of the application of this rotary platform as part of the service and target systems of the small spacecrafts are considered. It is shown how this can provide a significant increase in the tactical and technical characteristics of the small spacecrafts.

Investigation of Perturbations Arising from Temperature Shock with a Symmetrical Arrangement of Flexible Elements of a Small Spacecraft

This paper investigates the effect of a temperature shock on a small spacecraft with symmetrically arranged flexible elements. A two-dimensional thermoelasticity problem is posed. The disturbing effect of temperature shock on a small spacecraft has been determined. The assessment of the main disturbing factors arising from the temperature shock of flexible elements of a small spacecraft was carried out. Approximate dependences were obtained for the components of the displacement vector of the flexible element points. Numerical simulation was carried out for the symmetric scheme of the small spacecraft with two and four flexible elements. The dependence of the inertia force on temperature shock for the simulated small spacecraft at various initial deflections of the flexible element was constructed. Conclusions were drawn about the significance of the temperature shock influence on the dynamics of a small spacecraft. The results obtained were compared with the results of other studies and can be used in solving problems of remote sensing of the Earth and the implementation of gravity-sensitive processes on board small spacecraft.

Space integrated network: architectural and technical solutions justifica-tion of the ReshUCube-2 space mission

Space missions using small CubeSat-type spacecraft have become widespread. In this regard, integrated space networks are a promising direction for the development of this industry. In this paper, the architecture of such a network is proposed. It is planned that the first node of this network will be the ReshUCube-2 device. Functional requirements and technical limitations of the described architecture were determined. Based on the requirements, a review and comparative analysis of network protocols and technologies of different levels: physical, channel, network, transport and application. As a result, a network stack is proposed that meets these requirements. The key protocols in the stack were chosen: LoRa, 802.15.4, IPv6, 6LoWPAN, RTL, TCP, UDP and CoAP. Methods of software and hardware implementation of network nodes are proposed. A test bench based on specially developed unified devices is demonstrated, on which it is proposed to work out the functions of network interaction. A test payload board for small spacecraft nodes of the space segment is also shown. The results obtained will be used in the development of future space missions of the Reshetnev Siberian State University. In addition, the results can be useful in the design of the Internet of Things, the Internet of vehicles and similar networks. The results of the work can also serve as a basis for future research in the field of network technologies, digital signal processing, inter-satellite interaction and space information systems.

Analysis of design parameters and technical performance of onboard equipment of the Capella (Whitney) small spacecraft carrying synthesized aperture radar

Analysis of design parameters and technical performance of onboard equipment of the Capella (Whitney) small spacecraft carrying synthesized aperture radar

Generalized method of substantiation of ballistic structures of the cluster of Earth remote sensing small spacecrafts

Generalized method of substantiation of ballistic structures of the cluster of Earth remote sensing small spacecrafts

“AIST-2D”: Results of flight tests and application of earth remote sensing data for solving thematic problems

The idea of creating small spacecraft and constellations based on them, placed in low and medium earth orbits, has been formulated and widely used in the field of space systems development over the last two decades. To date, there are a large number of small satellites designed for remote sensing of the Earth.This paper describes the purpose and main technical characteristics of the small earth remote sensing satellite “AIST-2D”. It also describes the progress of the flight tests and the results obtained with the few payloads. In addition, examples of the use of the received Earth remote sensing data to solve various problems are presented. Finally, conclusions are drawn on the interim results of the spacecraft's operation and the prospects for the development of the “AIST-2” platform are described.The small spacecraft “AIST-2D” created in the shortest possible time (less than three years) is an example of successful implementation of a project to create a modern multifunctional spacecraft for remote sensing of the Earth. The experience gained in the development, construction, experimental testing and operation of the “AIST-2D” spacecraft opens up opportunities for the creation of a series of spacecraft based on the “AIST-2” platform with minimal costs and short production times, which will allow us to count on the creation of highly efficient space monitoring systems in the foreseeable future.

Failure Identification Method and LMI-Based Control System for Small Spacecraft

This study aimed to design a linear matrix inequality (LMI)-based robust control system and a failure identification method for small spacecraft. The key feature of the proposed method is its capability to withstand failure in the actuation system by means of the observer and controller gain definition. An effective approach for withstanding failure is required due to the small capabilities of the momentum exchange device (MED) and the system’s sensitivity to external perturbations. Passive fault-tolerant approaches can be included in the design process of the control algorithm. In particular, the main objective of this study was the design of an H∞ controller that started from an LMI formulation and considered the parametric uncertainties and matched failure of the actuation system. In addition, a fault detection method based on sliding mode observers was proposed to include an active disturbance correction algorithm and improve the system robustness and performance during spacecraft stabilization. The closed-loop system was evaluated for different initial conditions, including attitude positions that are far from the desired condition. The effectiveness of the proposed approach was demonstrated using extensive simulations that considered a pyramidal actuation configuration and hardware constraints.

Investigating the Temperature Shock of a Plate in the Framework of a Static Two-Dimensional Formulation of the Thermoelasticity Problem

The paper investigates the stress–strain state of a homogeneous rectangular plate after a temperature shock. It is believed that the plate is the first approximation of the solar panel model of a small spacecraft. To study the stress–strain state of the plate, a two-dimensional thermoelasticity problem is posed. The problem has a static formulation, since it does not take into account the dynamics of the plate’s natural oscillations. These oscillations affect the stress–strain state through the initial deflection of the plate at the time of the temperature shock. This deflection changes the parameters of the temperature shock and does not allow the use of a one-dimensional formulation of the thermoelasticity problem. As a result of solving the static two-dimensional thermoelasticity problem, approximate solutions are obtained for the components of the plate point’s displacement vector after the temperature shock. An approximation of the temperature field is presented. A numerical simulation is carried out. The correspondence of the obtained approximate analytical dependencies of the components of the plate point’s displacement vector to the numerical simulation data is analyzed. The proposed method can be used to assess the significance of the influence of the small spacecraft’s solar panels temperature shock on the dynamics of its rotational motion.

Space Panoramic Radio Altimeter Model: Imaging of the Developing Tsunami Wave Field in a Bistatic Quasi-Mirror Radar

A model of the formation of the gradient sea-level field by the space bi-static radar measurements is developed. An interferometer radar with cross-sectional base is used as a receiver. The relatively small antenna 5–10 m of cross-sectional base gives a huge gain in energy when working for quasi-mirror scattering in the range of short wind waves. The efficiency of the system is evaluated by converting the dynamical model of tsunami wave evolution for the case of the Kuril earthquake (October 4, 1994) into a panoramic radar image of the sea level. The panoramic image of the front wave allows predicting the direction, amplitude and, finally, the expected time of tsunami arrival to a given point. The obtained radar image confirms the main feature of the quasi-mirror method: the fluctuation-level sensitivity varies within the radar swath (~2000 km) and is the worst near the mirror point. For the chosen radar parameters, the average sensitivity in the swath is ~5 cm for a site (15 × 15) km. Without accounting the time required to transmit information from the receiver to the tsunami–prone sites, the minimal time interval between the appearance of the wave front and the tsunami alert is determined by the number of sequentially launched small tandem spacecrafts. For a single tandem this time is about 45 min.

A Compact Full-disk Solar Magnetograph Based on Miniaturization of the GONG Instrument* *Submitted January XX, 2023.

The design of compact instruments is crucial for the scientific exploration by smaller spacecraft such as CubeSats and deep space missions, as these missions require minimal instrument mass. In this proof-of-concept study, we demonstrate the miniaturization of the Global Oscillation Network Group (GONG) instrument. GONG instruments routinely produce full-disk maps of the Sun’s photosphere using the Ni i 676 nm absorption line, measuring Doppler shifts and magnetic fields. To miniaturize the GONG optical design, we propose replacing the bulky Lyot filter with a narrow-band interference filter. We validate this concept through numerical modeling and proof-of-concept observations. Finally, we propose a simple optical design for building a compact version of GONG.

Determination of Coordinates of Radio Emission Sources for a Satellite Radio Monitoring System Based on a Small Spacecraf

To date, within the framework of existing ground-based radio monitoring facilities, it is impossible to qualitatively perform the functions and tasks of radio monitoring. For countries with a large territory, such as the Republic of Kazakhstan, it is advisable to use low-orbit small spacecraft as radio monitoring stations to increase the efficiency of the radio monitoring system for the use of the radio frequency spectrum. The paper considers the developed method for determining the coordinates of a radio emission source based on the angle-measuring method. Algorithms are proposed and a computer program is compiled to determine the location of the radio source. The developed method can be used at the initial stages of the development of a radio monitoring system based on a single small spacecraft.

An experimental maneuver control for rendezvous and autonomous docking of a small spacecraft

This paper deals with maneuver control for the autonomous docking of two small spacecraft in a rendezvous flight. Due to hardware constraints, maneuver control suppressing fuel consumption and computational cost is a significant issue for small spacecraft. Here, the maneuver control for the approaching motion used a model predictive control system. The spacecraft's maneuvers in two-dimensional plane motion were performed in a frictionless environment with air bearings to verify the control performance.

Aerocapture: Enabling Small Spacecraft Direct Access to Low-Circular Orbits for Planetary Constellations

Small satellite constellations in multiple-inclination, low-circular orbits around Mars and Venus have the potential to perform a range of high-value science investigations within cost-constrained missions. A major challenge for small satellites is that they require large ΔV to enter low-circular orbits, which can drive up both spacecraft mass and cost. Compared to chemical propulsion, which requires large amounts of propellant, and electric propulsion, which requires large solar arrays and comes with long flight times, aerocapture enables direct access to low-circular orbits at Mars and Venus with minimal ΔV. The study shows how drag-modulation aerocapture, when combined with small B-plane targeting maneuvers, allows the delivery of multiple small satellites to various-inclination, low-circular orbits to establish a constellation. Preliminary cost estimates indicate that by reducing the required ΔV for orbit insertion, aerocapture can potentially reduce the cost of a small satellite going to a low-circular Mars orbit compared to propulsive insertion. The ability of low-cost spacecraft to enter planetary orbits will enable a new paradigm of interplanetary missions using small dedicated launch vehicles and planetary constellations at Mars and Venus.

An information system to support preliminary design of small spacecraft

With the development of information technology, the role of support systems for the technical products design process has increased significantly. The capabilities of knowledge management systems make it possible to ensure the commonality of terminology, object definitions and relations between them. In addition, the use of formal semantics is necessary to support the reuse of knowledge. The article discusses ways to improve the efficiency of the small spacecraft design process and a creation of tools for the knowledge exchange about the functional characteristics of typical subsystems of such vehicles. The creation of an information system to support the preliminary design of small spacecraft, which contains information linked to the relationship between the functions and components of subsystems and their characteristics, is considered. The created system facilitates the synthesis of various product configurations. When developing the information system, modern approaches in the field of structural and functional classifications of small spacecraft systems and approaches existing in the field of taxonomy to create database structures were used. The information system database contains a set of characteristics for a wide range of small spacecraft equipment. As an example of the use of an information system to support the design of small spacecraft, the process of choosing the characteristics of a power supply system for a promising small spacecraft designed to provide communication between mobile ground terminals is considered..

Postflight Recovery of the Rotational Motion of a Small Space Vehicle from Solar Sensor Information

The use of the developed technology for solving inverse problems of parametric identification of the rotational motion of a spacecraft (SC) based on the accumulated sample of measurements of various compositions is expanded. On the example of processing telemetry data (measuring the direction vector to the Sun) of the BugSat-1 small spacecraft (SSC), the possibility of determining not only the characteristics of rotational motion but also its inertia coefficient is demonstrated. The rotational motion is determined based on the data obtained for five time intervals from January to May 2022, during which BugSat-1 made an uncontrolled movement around the direction vector toward the Sun. This type of motion does not allow applying traditional approaches to determine the dynamics of its rotation. In the process of data processing, a simplified model of the rotation of a SSC is used. At the same time, the initial conditions of rotational motion are estimated—the angular velocity vector and orientation angles, which made it possible to draw a conclusion about the precessional motion of a SSC. Using the analytical model of a regular precession, it is possible to estimate the inertia coefficient of a SSC. The results of flight information processing are presented and the errors of the obtained results are estimated.

Calibration Method of MEMS Gyroscopes Using a Robot Manipulator

Inertial measurement units composed of microelectromechanical systems (MEMS) gyroscopes are widely used sensors for small spacecraft and nanosatellite attitude estimation, control, and orientation. This is due to the fact that when developing nanosatellites, it is convenient to use devices of small dimensions and low energy consumption, which are advantages of MEMS sensors. However, one of the significant disadvantages of these measuring devices is their relatively low accuracy, which is related to systematic and random errors. For this reason, this article proposes a method for calibrating MEMS-based angular rate sensors using a robot manipulator for partial elimination of these errors. There are high-precision rotary calibration stands on the market that are designed to set angular rates and orientations when testing products of various types and accuracy classes. Similarly, a robot manipulator represents a possible option for performing this task. As a result, a modified six-position method as a sequence for performing laboratory calibration using a robot joint for high-precision rotation is proposed. The features of the results of processing experimental measurement data of tests have been validated on MEMS angular rate sensors. After calibration, the quality of MEMS gyroscope output data using this calibration approach was significantly improved.

Подход к снижению виброактивности малых космических аппаратов

Flawless spacecraft operation and successful implementtation of their research in orbit mainly depend on technical and operational characteristics of the spacecraft orientation and stabilization control systems in space relative to the base frame of reference axes achieved through stabilizing properties of certain external and internal perturbations. The research was aimed at determining static and dynamic imbalances and analyzed the levels of influence of forces and moments created by non-contact electric motor-flywheel operating at the various rotor speeds. In order to reduce the static and dynamic imbalance of the non-contact electric motor-flywheel obtained as a result of ongoing experimental studies, it was proposed to introduce a vibration isolator with natural frequency significantly lower than that in the conventional systems in order to statically hold the non-contact electric motor-flywheel during ground and flight operation. The vibration isolator should have small mass and rigidity, as well as a wide range of operating temperatures, etc. Depending on the orbit altitude, the moments of external disturbing forces acting on the Iskra-5 spacecraft were determined. Electric DC motor-flywheels of various types were analyzed. It was shown that introduction of the non-contact electric motor-flywheel as the executive element of the orientation and stabilization control systems on small spacecraft improved the orientation accuracy. Ways to reduce the effect of vibration disturbances on the spacecraft structure from the non-contact electric motor-flywheel were considered

LOW POWER HALL THRUSTER ST-22 WITH PERMANENT MAGNETS


 
 
 The results of the design, development and testing of the laboratory model of the Hall Thruster ST- 22 are presented. The main attention in the development of the thruster was focused on the optimization of the magnetic field in the accelerating channel of the thruster. When choosing a constructive thruster diagram, it was proposed to use constant magnets and a non -magnetic ring in a magnetic system. The proposed constructive features of the thruster made it possible to exclude a magnetic screen from its design, which led to the simplification of the structure and reduce its mass. Laboratory tests of the ST-22 thruster were carried out when changing the discharge voltage in the range of 160-260 V, as well as with three different values of the mass flow rate of Xenon submitted to the anode - 0.4, 0.5, 0.6 mg/s. In the course of laboratory investigation of the ST-22 Hall thruster, the following characteristics were determined: current-voltage characteristics of the discharge in the acceleration channel of the thruster; the dependences of thrust and specific impulse on the mass flow rate of the working substance and the discharge voltage, as well as the value of the efficiency of the anode block of the thruster on the discharge voltage. As a result of laboratory tests, it was shown that at input power of 100 W, the discharge voltage 215 V and the working substance flow through the anode unit 0.6 mg/s, the thruster provides the thrust value of 7 mN, the discharge current of 0.47 A, a specific impulse of 950 s, and efficiency 33%. The values of the specific impulse and the thruster efficiency were given taking into account the mass flow rate of the working substance in the cathode-compensator. The thruster provides sustainable operation in the range of power consumption of 50 - 100 W and it was designed to use small spacecraft on board, which on board have appropriate electric power.