Orbit Control Research Articles

Study on radiation characteristics of multi-phase plumes containing ice crystals in orbit-control engines

In order to obtain the radiation characteristics of the multiphase plume containing ice crystals in the high-altitude orbit control engine and analyze its influence on the telemetry effect, this paper deduces the radiation characteristics parameters of gases based on Einstein radiation theory and radiation transfer equation. Based on the latest HITRAN2020 spectral line library solving gas radiation and the Mie theory of particle scattering solving particle radiation, the radiation characteristic parameters of the multi-phase plume containing ice crystals in the high-altitude orbital control engine and the radiation intensity distribution in the near-infrared and mid-long wave infrared bands are simulated, and the main influencing factors of the radiation are analyzed. The results show that, in general, the main radiation of the engine plume includes the scattered light in the near infrared band and the thermal radiation in the middle long wave infrared band. In the near-infrared band, considering solar background radiation, the total radiation amount in the calculation domain is about 10 −6 W/sr, and the thermal radiation is dominant within the axial distance of 0.5 m, while the ice crystal scattering is dominant beyond the axial distance of 0.5 m . In the mid-long wave infrared band, the total radiation amount in the calculation domain is about 10 W/sr, and the thermal radiation of gas dominates within 1 m axial distance, while the thermal radiation of ice crystal dominates beyond 1 m axial distance.

Reliability modeling of multi-state phased mission systems with random phase durations and dynamic combined phases

Random phase durations and dynamic combined phases challenge the application of existing reliability models in the reliability analysis of multistate-phased mission systems (MS-PMSs). To this end, this paper presents a new reliability modeling method for multi-state phased mission systems with random phase durations and dynamic combined phases. Initially, a multi-state multi-valued decision diagram-based (MMDD-based) reliability modeling method is created to efficiently map random phase durations and the dynamic combined phase nature of MS-PMSs into the reliability model. To solve the MMDD-based reliability model, a path probability evaluation method is subsequently constructed with the assistance of the Markov regenerative function. The effectiveness and the superior performance of the proposed MMDD-based reliability model and its solving algorithm are validated by the application to the reliability modeling and analysis of an attitude and orbit control system with multiple modes. Overall, this paper provides the reliability sector with a new reliability model and its solving algorithm to enhance the reliability and safety of multi-state phased mission systems.

Microstructure characterization and formability assessment of electron beam welded Nb-10Hf-1Ti (wt.%) refractory alloy sheets

Formability study on welded blanks of Nb-based refractory alloy C-103 (Nb-10Hf-1Ti (wt.%)) has tremendous potential for utilizing smaller sheets for realization of large-size divergent portions of upper-stage liquid engine nozzle of satellite launch vehicles and thrusters of attitude orbital control systems. In this work, vacuum-annealed monolithic C-103 sheets were electron beam welded successfully to obtain defect-free welds and detailed microstructural and mechanical characterizations were investigated. Columnar epitaxial grains were found in fusion zone (FZ) instead of equiaxed grains in base metal (BM). Detailed SEM and HRTEM analyses revealed spherical shape of nano-sized HfO2 precipitates with monoclinic crystal structure in the C-103 sheet. Further, microhardness across the weld cross-section indicated that the FZ had a uniform and maximum hardness of ∼195 HV0.5 due to formation of finer nano-sized HfO2 precipitates with higher number density, followed by a lowering of hardness in a narrow heat-affected zone. A marginal reduction in tensile strength of ∼6% with a considerable decrease in ductility of ∼29% was noticed for the welded sample. However, fracture occurred in the BM region, indicating good weld integrity. Furthermore, formability of the welded blanks was evaluated in terms of limiting dome height (LDH), forming limit diagram, and deep drawn cup depth. The effect of the presence of WZ on the formability of the C-103 sheet was estimated under uniaxial, plane strain, and biaxial tensile strain paths for the first time, and the results were compared with the monolithic sample. It was noted that the failure limit of the welded specimens decreased compared to that of the monolithic sample, resulting in a lower LDH of approximately 69% and 62% when deformed along plane strain and biaxial tensile strain paths, respectively. However, a marginal variation in LDH was observed for the uniaxial strain path. Also, the welded blank had more resistance to material flow into the die cavity due to the harder weld region, resulting in a reduction of deep drawn cup depth by 52% than that of the monolithic blank. The overall results concluded that the presence of the WZ significantly affected the formability of the C-103 sheet, especially in plane strain and biaxial strain paths. However, the fracture was never found to be propagating along the weld line, indicating the ability of the weld to sustain large plastic strains. This study provides insightful information on the formability of electron beam welded C-103 sheets for the successful fabrication of space components.

DFT calculations of properties of CF3-carbocations of the thiophene series as key intermediates in the synthesis of novel CF3-thiophenes

Trimethylsilyl ethers of thiophen-2-yl-(trifluoromethyl)methanols undergo multichannel electrophilic transformations under the action of Brønsted superacid acid CF3SO3H to form a wide range of CF3-substituted thiophene structures. The key cationic intermediates of these transformations are CF3-carbocations. Electronic and orbital characteristics, global electrophilicity indices, and thermodynamic characteristics of these key intermediates have been calculated with a use of the density functional theory (DFT). Experimental data on the reactivity of intermediate cationic species have been explained on the basis of the results of quantum chemical calculations. It has been shown that reactions of the cations with aromatic nucleophiles are mainly ruled by orbital control.

An Orbital-attitude Coupled Control Framework for a Full-scale Flexible Electric Solar Wind Sail Spacecraft in Orbital Transformation Missions

In this work, a novel orbital-attitude coupled control framework is developed for the electric solar wind sail (E-sail) spacecraft, aiming to handle the orbital transformation and attitude maneuver simultaneously. In the framework, the desired attitude parameters and the slow-varying voltage component are provided by the orbital control strategy, while the current attitude parameters and the fast-varying voltage component are manipulated by the attitude control strategy to approach their desired values. The orbital-attitude coupled characteristics of the E-sail spacecraft, including the flexibility-induced coupling effect, are fully described by the referenced nodal coordinate formulation. Considering the input saturation conditions, the governing equation for the orbital control strategy is then derived, in which the in-plane and out-of-plane displacement and velocity errors are prescribed as the state variables to be eliminated. An integral sliding mode control (ISMC) scheme is proposed to improve the robustness against the unmeasurable disturbance term. A model predictive control (MPC) scheme is introduced to enhance the convergence efficiency, where a quadratic optimization is performed to plan the desired attitude parameters and voltage components within the prediction horizon. To evaluate the control performance in the orbital transformation and attitude maneuver missions on the displaced non-Keplerian orbit, a series of scenarios with complex initial conditions are simulated under different control schemes, including the ISMC-MPC compound scheme. The results show that the control strategy designed under the rigid-body assumptions may not be feasible for the flexible E-sail spacecraft, while the investigated control strategy realizes the accurate and efficient convergence of the orbital and attitude variables on both the rigid and flexible E-sail spacecraft with the tether deformation stabilized.

PDE-based anti-disturbance attitude and vibration control of flexible satellite under output constraints

In this paper, an anti-disturbance control strategy is exploited for the attitude reorientation and vibration damping of flexible satellite under disturbances and output constraints. The flexible satellite is formulated by partial differential equations (PDEs) and the exploited controller extends the existing results from the following two aspects. (1) The exploited controller is capable of handling the time-varying output constraints by incorporating the barrier Lyapunov function (BLF) with a coupling-based item through Lyapunov analysis. (2) The exploited controller possesses the excellent disturbance rejection property by introducing two disturbance observers to separately identify the disturbance torque and force. The asymptotic stability of the closed-loop system is strictly evaluated. The exploited controller can restrain the attitude reorientation error and the tip deformation always in the predefined output constraints. Lastly, comparative simulations validate and highlight the main results.

Numerical investigation of micropropulsion systems for CubeSats: Gas species and geometrical effects on nozzle performance

In the present investigation, a numerical investigation is conducted to investigate cold-gas micropropulsion devices for CubeSat orbital control. Different micronozzle geometry configurations and gas species are analyzed to assess their influence on the flowfield structure, surface properties, and microthruster performance parameters. Due to the small length scales, the Direct Simulation Monte Carlo (DSMC) method is used to simulate rarefied argon and nitrogen in rectangular and curved micronozzles. The results indicate that nitrogen gas leads to a 23% increase in specific impulse at a relatively insignificant thrust cost. On the other hand, the use of a curved geometry increases the specific impulse and thrust generated by 23% and 35%, respectively. The results indicate that using lighter gases for cold gas microthrusters increases the micropropulsion efficiency. In addition, curved geometries significantly improve the overall performance of such devices, in contrast to rectangular geometries.

Spacecraft formation flying orbital control for earth observation mission

Synthetic Aperture Radar (SAR) sensors have gained much attention for multiple civil use cases, where NewSpace startups have launched and planning to launch sensors to orbit to harvest the unique capabilities of SAR for the Earth Observation (EO) market. Unlike optical, SAR can capture images in all lighting conditions and penetrate cloud cover, generating exceptional use cases, which makes SAR the most reliable sensor. For example, monitoring soil deformation with millimeter accuracy, oil spills, crop lodging, and typical EO use cases of optical sensors that are unable to penetrate the clouds or see at night. However, SAR is still limited in civil use cases because it is tough to interpret. In addition, unlike government users who can consume the SAR data directly, most civil end-users need analytics on top of the data. To unlock the potential civil use cases, many experts believe that optical EO data should be fused with SAR at the EO upstream, thus suggesting optical and SAR sensors fly in formation. Spacecraft Formation Flying (SFF) is one of the key technology enablers for space exploration and EO space missions. This paper presents an SFF EO mission (SAR and Optical) design considering orbit requirements to maximize user objectives inspired by OptiSAR™ Constellation.

Robust neuro-adaptive command-filtered back-stepping fault-tolerant control of satellite using composite learning

This study proposes an adaptive neural command-filtered backstepping control system for precise and fast attitude control of a satellite considering different uncertain dynamics. Mission success crucially depends on robust attitude control resilient to model uncertainties, unmodeled dynamics, external disturbances, and actuator faults. The proposed approach synergistically combines a neural network for handling model uncertainties, unmodeled dynamics, and actuator faults, with a disturbance observer for compensating external disturbances and neural network estimation errors. The command-filtered backstepping technique avoids the explosion of complexity inherent to traditional backstepping, while integrating integral action into the design results in the elimination of the steady-state tracking error. Besides, a composite learning method optimizes the update laws for the neural network and disturbance observer weights, enhancing control performance. Despite the presence of uncertainties, the closed-loop system stability is guaranteed by the Lyapunov stability theorem. Simulation results demonstrate the proposed controller’s ability to handle severe actuator faults, unmodeled dynamics, and measurement noise without requiring explicit fault detection and isolation schemes.

Attitude Determination and Control in Small Satellites: A Review

Attitude Determination and Control in Small Satellites: A Review

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.

Development of a low-power Hall thruster with permanent magnets and a dual trigger electrode hollow cathode for the Qilu satellite constellation

A low-power Hall effect thruster (LpHet-100) has been installed on the Qilu-3 satellite and has been launched by the Long March 2D rocket in January 2023. The specific features of LpHet-100 are presented in this article. An innovative dual trigger electrode poison-resistant hollow cathode and an effective magnetic shielding configuration with permanent magnets, additive manufacturing gas distributor are designed and experimentally verified. The prominent characteristics of a shorter starting time, lower ignition shock, and self-healing have been shown in tests of the dual trigger electrode hollow cathode. Thrust performance tests proved that LpHet-100 can work stably in the power range of 50–230 W and generate a thrust of 2.9–10.8 mN, with the maximum specific impulse of 1594 s and maximum thrust efficiency of 38 %. Rigorous environmental tests were conducted to verify the thruster design scheme. In addition, Torque measurement of Hall thrusters is incorporated into the ground tests for the first time. A plasma torque of 5.65–9.88 μN·m is obtained when LpHet-100 operated in 132–228 W. It provides a critical guideline for the attitude and orbit control schemes of the Qilu-3 satellite, which is an essential part of space mission planning. Generally, this study proposes new design of the thruster body and hollow cathode, providing an exhaustive reference for the development process and application of low-power Hall thrusters.

On-orbit mission overview of the low power hall thruster propulsion system aboard venμs satellite

Venμs is a satellite launched in 2017, for super-spectral Earth imaging and Electric Propulsion System (EPS) demonstration. In this paper we overview EPS design and operation throughout all five mission phases, from open/closed-loop orbit control, through orbit descent (720 → 410 km), orbit maintenance under high drag environment, to orbit raising (410 → 560 km). The EPS consisted of two throttleable Hall thrusters, PPU, Propellant Management Assembly (PMA), and a 9 L propellant tank carrying 16 kg of Xenon. Both thrusters operated in the 300–550 W power range, generated a combined total impulse of 158.1 kN-sec and consumed all propellant. Two methods are described to compute the remaining propellant mass – “Bookkeeping” and “PTV” methods. The advantages and disadvantages of each method are discussed in light of the Venµs mission. Thruster performance was measured on-orbit using the “Orbit Determination” method and compared to laboratory experiments conducted on the ground with identical thrusters. The measured performance on-orbit was found to agree within the error bars with the performance measured on the ground. Lastly, we present several repeating events in which the propulsion system suffered from sudden beam-outs or ignition difficulties. We present the methods used to construct a solution and implement it.

Discrete search-based determination of a local orbital frame in unknown environments

A local orbital frame is an important and often applied coordinate frame in attitude and orbit control. Usually, a navigation filter estimates the position and velocity vector, which are used to build up this reference frame. It is beneficial for many space applications, such as communication or Earth observation missions, or for conducting orbit control maneuvers to change the orbit plane. Furthermore, satellite formation flying can require an orbital frame to describe relative motion. This work investigates the feasibility of determining the local-vertical local horizontal (LVLH) frame without full knowledge of the orbit. This scenario often applies to non-characterized small bodies like asteroids or comets. A discrete search method is used to determine the complete LVLH frame in real-time, assuming that one axis, a line-of-sight vector to the body’s center of mass, is known. By minimizing an objective function, the remaining axes can be determined. The estimated parameters are the inclination and the right ascension of the ascending node of the osculating orbit plane. This study investigates the feasibility of this concept and then applies it to perturbed orbits and sensor noise. The proposed method demonstrates that a discrete search can determine the LVLH frame. The approach can be seen as a step toward increasing the autonomy of a spacecraft near small bodies with unknown environments.

Comparison of Optimization Methods for the Attitude Control of Satellites

The definition of multiple operational modes in a satellite is of vital importance for the adaptation of the satellite to the operational demands of the mission and environmental conditions. In this work, three optimization methods were implemented for the initial calibration of an attitude controller based on fuzzy logic with the purpose of performing an initial exploration of optimal regions of the design space: a multi-objective genetic algorithm (GAMULTIOBJ), a particle swarm optimization (PSO), and a multi-objective particle swarm optimization (MOPSO). The performance of the optimizers was compared in terms of energy cost, accuracy, computational cost, and convergence capabilities of each algorithm. The results show that the PSO algorithm demonstrated superior computational efficiency compared to the others. Concerning the exploration of optimum regions, all algorithms exhibited similar exploratory capabilities. PSO’s low computational cost allowed for thorough scanning of specific interest regions, making it ideal for detailed exploration, whereas MOPSO and GAMULTIOBJ provided more balanced performance with constrained Pareto front elements.

ORBIT SELECTION OF THE SPACE INDUSTRIAL PLATFORM WITH DISTRIBUTED ELECTRICAL-POWER SYSTEM MODULES

Space industrialization is one of the prospective directions in modern aerospace science and engineering for space exploration of new resources and habitats. The key issue is to provide industrial space modules with the required amount of electricity needed. One type of power supply for such modules is the use of distributed power systems, which consist of constellations of spacecraft with contactless power transmission. Given this, the problem of rational orbit selection for their dislocation arises. Considering these problems, the methodology for orbits selection of the space industrial platform with distributed electrical-power system modules is proposed in the paper. This methodology includes orbital translation, attitude, relative dynamics estimation for each power satellite, and its corresponding orbit optimization algorithm. The orbit optimization algorithm includes statistical processing and elements of gradient and coordinate descent methods, allowing us to determine the most significant parameter influencing the duration of the contactless power transmission session. Also, quaternion mathematics is used to estimate the dynamics in the program parameters for targeting the transmitter spacecraft antenna to the receiver spacecraft rectenna. With the approaches mentioned above, the methodology proposed in this paper allows us to form the requirements for the power satellites’ attitude and orbit control system to improve the process of selecting corresponding design parameters of such systems. Thus, the usage of the proposed methodology can allow the designing of the power satellites’ attitude and orbit control system in the conceptual stages of designing.

Kinetic Simulation of Nozzle Flow in a Micronewton-Class Cold Gas Thruster

Many scientific space missions need highly precise attitude and orbit control or ultrafine drag compensation, which relies on the variable-thrust propulsion technology operating at the micronewton level. Cold gas thruster (CGT) is a very promising solution, mainly because of its high reliability. One of the keys to the success of micronewton variable-thrust CGT is to understand the flow in its nozzle, whose configuration is much more complex than traditional CGT nozzles. This paper applies kinetic-based multiscale models to investigate the gas flow in the complex nozzle of a micronewton variable-thrust CGT. The simulations reveal that the flow simultaneously experiences all kinds of regimes from continuum to free molecular. The continuum breakdown is likely to occur near the throat region due to large gradients of flow variables and in the expander due to low gas density. Frictional choking is observed, and the nozzle length can be optimized to improve the thruster performance. Nozzle performance measures such as thrust, discharge coefficient, and thrust efficiency are found to change only with the throat Knudsen number Knt. The performance curves can be divided into two sections at Knt≃0.1, and thereby an empirical piecewise formula for thrust prediction is proposed.

Prescribed modal vibration control and disturbance load analysis of rigid-flexible satellites

We investigate the vibration control issues of rigid-flexible satellites and present a novel control methodology for vibration suppression during attitude maneuvers. Within the vibration model, we derive the terms of the disturbance load on each vibration mode, meticulously analyzing their effects. To reduce the impact of the disturbance load on vibration modes, the prescribed modal vibration control is proposed for vibration suppression, encompassing model adjustment, optimization of actuator/sensor deployment, and controller design. The effectiveness and the advantage of the prescribed modal vibration control are demonstrated through numerical simulations. This methodology achieves proper controlled vibration modes selection and optimized deployment, thereby amplifying the effect of vibration suppression in the presence of disturbances relative to extant vibration control methodologies.

Characterization and verification of the optimal feedback gain of a satellite magnetorquer-based attitude control system

In spacecraft mission planning and operation, the attitude determination and control subsystem (ADCS) of a satellite provides information about the orientation of the satellite in the inertial reference frame. Furthermore, this subsystem produces the control actions required to adjust the orientation of the satellite, especially in the low-Earth orbit (LEO) regime. This paper focuses on the satellite’s three-axis attitude control problem within the context of active and passive control, which includes detumbling control, pointing control, magnetic control, and attitude stabilization after solar panel wing deployment using magnetorquers as the primary actuators. The objective is to stabilize and reduce the angular rate while orienting the satellite to the desired attitude. The proposed satellite attitude control system (ACS) strategies are designed, developed, characterized, and verified. These strategies encompass the B-dot control algorithm for detumbling control along with pointing control and attitude stabilization after solar panel wing deployment. hardware-in-the-loop simulation (HiLs) tests are conducted to assess the performance of the satellite magnetorquer-based ACS in the presence of noise. These tests involve a relative Earth’s magnetic field (EMF) generator in conjunction with SGP-4-based satellite orbital propagator high-level control software. Additionally, cascade proportional-integral-derivative (PID) and state-dependent Riccati equation (SDRE) controllers are implemented to generate sufficient torque using three-axis magnetorquers on a frictionless air-bearing platform. The platform is balanced to closely simulate the dynamic motion of a spacecraft in space. The testing includes a single initial condition and three inertia conditions for stabilization after solar panel wing deployment. Finally, the effectiveness of the cosimulation as a primary experiment through an integrated HiLs process is validated. This comprehensive approach confirms the control system’s performance and its ability to meet mission requirements.

Onboard guidance and control algorithms for increased autonomous aerobraking capabilities at Mars

Aerobraking represents a valuable option for interplanetary missions thanks to its capability of reducing the orbital semi-major axis through multiple atmospheric passages, thus with limited fuel consumption. However, due to structural and thermal loads the spacecraft undergoes, mission risks and ground operations costs increase. This study aims to tackle this aerobraking drawback by enhancing its autonomy. The main activities required for aerobraking execution, including atmospheric prediction, orbital state estimation, and orbit control, are tailored for onboard execution. Regarding orbit control, two different methodologies, differentiated by working principle and prediction horizon breadth, are investigated. The first plans orbit control manoeuvres based on the prediction of the atmospheric conditions for a single upcoming pericentre, while the second schedules and optimizes the manoeuvres considering the recent atmospheric conditions experienced by the spacecraft and the altitude trend of multiple upcoming pericentres. The whole operational concept is tested within a simulated environment, with the primary objective of accurately reproducing the intense atmospheric variations of the Martian atmosphere, which represents the most challenging aspect during aerobraking execution. The numerical testing, involving the simulation of different aerobraking regimes under varied conditions, shows that the natural dynamics of the orbital motion can be exploited to minimize both fuel consumption for orbit control manoeuvres and spacecraft load variations, thus enabling more efficient aerobraking and more stable flight conditions, even with the limited predictive capabilities that characterize onboard resources. Both control logics permit the aerobraking execution without exceeding the maximum allowed thermal limits, hence showing their effectiveness. Furthermore, a Monte Carlo analysis conducted on a complete aerobraking simulation reveals a 25.5% reduction in propellant consumption and a 9.4% reduction in heat rate variability when the orbit control decision process uses information relative to multiple future pericentres.