Spacecraft Solar Array Research Articles

A novel physics-based computational framework to model spacecraft solar array power under degradation: application to European Space Agency (ESA) Cluster mission

Accurate modeling and simulation (M&S) of spacecraft solar array power under degradation is essential for mission planning, remaining useful life assessment, and lifetime extension. A relevant example is ESA’s Cluster spacecraft fleet, launched in 2000 and operated at the European Space Operation Centre (ESOC), whose solar arrays have suffered severe degradation due to space radiation that has caused challenges to routine operations and mission planning. However, currently available physics-based and machine learning models have been proven ineffective in modeling the drastic reduction in power generation over the long operational life of the spacecraft.In response to these limitations, this work introduces a framework to model solar array degradation and predict power generation. It embeds a novel simplified physics-based model and a meta-heuristic optimization algorithm which exploits domain-specific knowledge and monitoring data for robust model parameter calibration and accurate power generation predictions. The results show the effectiveness of the proposed approach in avoiding overfitting and providing an accurate estimate of Cluster solar array power evolution.

Study on plasma expansion model of primary discharge on spacecraft solar array

In the space plasma environment, primary discharge may occur on the solar array and evolve into a destructive sustained arc, which threatens the safe operation of the spacecraft. Based on the plasma expansion fluid theory, a new multicomponent plasma expansion model is proposed in this study, which takes into account the effects of ion species, ion number, initial discharge current, and Low Earth Orbit (LEO) plasma environment. The expansion simulation of single-component and multicomponent ions is carried out respectively, and the variations of plasma number density, expansion distance, and speed during the expansion process are obtained. Compared with the experimental results, the evolution of propagation distance and speed is closed and the error is within a reasonable range, which verifies the validity and rationality of the model. The propagation characteristics of the primary discharge on the solar array surface and the influence of the initial value on the maximum propagation distance and the propagation current peaks are investigated. This study can provide important theoretical support for the propagation and evolution of the primary discharge and the key behavior of the transition to secondary discharge on spacecraft solar array.

Robust and Autonomous Space Weather Charge Mitigation Coatings

National Space Weather Strategy and Action plan calls for the need to enhanceProtection of National Security, Homeland Security, and Commercial Assets and Operations against Effects of Space Weather. Extreme space weather event like solar flares, cosmic rays, and radiation belts cause ionizing radiation that can damage electronics, solar arrays, and optical systems on satellites reducing their functionality and lifetimes, by inducing an ionic charge on the spacecraft’s surface, when a spacecraft fly in and out of the ionosphere. This negative charge buildup can lead to ion sputtering/arcing, and producing irreparable damage in spacecraft components. Therefore, the local application of materials systems that can passively mitigate the negative charge build up by emitting the electron back into space, while improving resistance to erosion during ion sputtering is of interest. Accordingly, Faraday Technology is developing a low-cost, efficient and scalable manufacturing process for the local deposition of lightweight passive highly emissive and erosion resistant coating onto spacecraft viable substrates that can mitigate charging and erosion effects from ionizing radiation. The composite coating showed a ~1500% increase in maximum (Emax) total electron yield over bare aluminum, extending by 4 times the range of electron yields between crossover energies >1, and demonstrated the potential for erosion resilience in modeled International Space Station plasma erosion conditions. The composite coating could be applied to various spacecraft platforms include spacecraft skin, solar arrays, circuit boards, and emitters such that their lifetime, effectiveness and durability within LEO/GEO environments events can be enhanced. Acknowledgements:The financial support of NASA SBIR/STTR program through contract No.80NSSC22PB020 is acknowledged.

Vacuum Arcs on Solar Array Surfaces

Arcs on the solar arrays appear to be the most plausible source for the electromagnetic radiation observed with the Arecibo radio telescope. In the laboratory, electrostatic discharges (arcs) on solar array coupons were studied in a simulated geosynchronous equatorial orbit environment. Arc current pulses were recorded with a high time resolution (up to 0.2 ns). Spectral properties of these pulses demonstrated no monotonic behavior but rather separated peaks within the frequency range of 100–1000 MHz. Electromagnetic radiation from these arcs was registered and recorded with a Yagi antenna 8 m distant from the arc site, and a high-speed oscilloscope. Analysis of the spectra resulted in the discovery of well-separated “lines.” Evaluation of arc current spectra and radio-frequency spectra revealed statistically significant patterns in the frequency domain. Measured radio fluxes and theoretical estimates of the power of electromagnetic radiation from arcs on spacecraft solar arrays strongly argue in favor of the arcing hypothesis for the source of this radiation. These results have enabled considerable progress to be made in understanding the physical nature of the generation of electromagnetic radiation by transient vacuum discharges.

An anomaly detection method for spacecraft solar arrays based on the ILS-SVM model

Solar arrays are important and indispensable parts of spacecraft and provide energy support for spacecraft to operate in orbit and complete on-orbit missions. When a spacecraft is in orbit, because the solar array is exposed to the harsh space environment, with increasing working time, the performance of its internal electronic components gradually degrade until abnormal damage occurs. This damage makes solar array power generation unable to fully meet the energy demand of a spacecraft. Therefore, timely and accurate detection of solar array anomalies is of great significance for the on-orbit operation and maintenance management of spacecraft. In this paper, we propose an anomaly detection method for spacecraft solar arrays based on the integrated least squares support vector machine (ILS-SVM) model: it selects correlated telemetry data from spacecraft solar arrays to form a training set and extracts n groups of training subsets from this set, then gets n corresponding least squares support vector machine (LS-SVM) submodels by training on these training subsets, respectively; after that, the ILS-SVM model is obtained by integrating these submodels through a weighting operation to increase the prediction accuracy and so on; finally, based on the obtained ILS-SVM model, a parameterfree and unsupervised anomaly determination method is proposed to detect the health status of solar arrays. We use the telemetry data set from a satellite in orbit to carry out experimental verification and find that the proposed method can diagnose solar array anomalies in time and can capture the signs before a solar array anomaly occurs, which reflects the applicability of the method.

Research on the Induced Electrostatic Discharge of Solar Arrays under the Action of ESD EMP

In this paper, the electrostatic discharge of solar arrays in spacecraft energy systems is taken as the research object. The influence and internal mechanism of external electromagnetic radiation on electrostatic discharge is studied. Meanwhile, the charging and discharging test platform of spacecraft solar arrays under a strong field is first established. Then, the influence of irradiation field strength, electron beam energy and beam density on electrostatic discharge of solar arrays is analyzed and summarized. The results show that during the irradiation process of solar arrays using a high-energy electron beam under vacuum conditions, the higher the electron beam energy and the beam current density, the higher the discharge frequency of the solar array. When the intensity of the external electromagnetic radiation field increases, the discharge frequency also increases. Under the action of external radiation field with the same peak field strength, the larger the gap, the smaller the discharge frequency. With the increase in the field strength, the potential difference of each part of the solar array becomes smaller, and the peak of the discharge current decreases. The research results can provide technical reference for electrostatic protection of spacecraft solar arrays.

Dynamics and control of spacecraft solar array deployment considering physical contacts between locking mechanisms

The numerical simulation of the deployment and locking process of flexible appendages is an important topic in the overall design of spacecraft. In the existing research, the connecting mechanism between the solar panels is usually regarded as an ideal revolute constraint, and the locking process is simulated by applying a virtual lock torque related to the deployment angle. However, this method cannot obtain the physical contact force history, and because the physical meaning of the stiffness term in the lock torque function is not clear, it is difficult to accurately illustrate the frequency characteristics of the flexural vibration of the panels in the post-lock phase. In fact, the locking process is a frequent contact-impact process between the lock pin and groove. In this study, a typical deployable spacecraft with torsion-spring-driven panels were investigated numerically. The locking mechanism is modeled as a physical entity, and the continuous contact force model between the lock pin and groove is established. The contact stiffness of cylindrical line-contact is given by preloading calculation using the finite element method. Based on the multibody dynamics, rigid body model and flexible body model of the solar panels are established respectively. Through the comparison between the numerical simulation results of the two models, it is found that the rigid body model will produce excessive contact force and false high-frequency vibration. Furthermore, the attitude PD controller of spacecraft is designed, and the difference of system dynamic response in the deployment process under the main-body free state and the main-body controlled state is studied. The vibration frequency under the main-body controlled state is close to the first-order natural frequency of the structural finite element analysis, which verifies the rationality of the calculation results of the model.

Does Phobos Reflect Solar Wind Protons? Mars Express Special Flyby Operations With and Without the Presence of Phobos

AbstractWe characterize the solar wind proton reflection (backscattering) from Phobos using a series of Mars Express operations. The plasma data obtained during the Phobos flyby of Mars Express in January 2016 showed a non‐solar wind signal possibly reflected from the Phobos surface. Similar signatures were previously reported during an earlier Phobos flyby in 2008. On the other hand, although Mars Express has encountered Phobos (within 100 km) more than 10 times, it has thus far detected only two clear cases of reflected protons. The intermittency of the reflected proton detections could indicate that these protons are not from Phobos but are produced by the spacecraft body or solar arrays under a special attitude configuration during the flyby maneuver. To rule out these artifacts as the cause, we conducted three special operations; Mars Express was operated using the identical attitude and solar panel control sequence but without Phobos nearby. All the measured plasma data during the special operations exhibit no additional plasma signatures. These fake flyby experiments indicate that the measured non‐solar wind protons during the real flyby are not from the spacecraft and that Phobos likely reflects solar wind into space. On the other hand, the intermittency of the reflected protons from Phobos remains unexplained. The reflected protons are vital additional plasma components because they inject enough energy to disturb the near‐Phobos plasma environment. In addition, we can utilize these reflected particles to remotely measure the surface precipitation flux.

Analysis of design of spacecraft solar arrays

Solar arrays supply electrical power to spacecraft equipment and also provide charging of electrochemical batteries used in the shadow sections of the orbit. Photovoltaic converters of light energy of solar radiation form the basis of the solar battery. Their principle of operation is based on the phenomenon of the photoelectric effect. The article provides an overview of the development of solar cells and the classification of modern designs of solar cells for spacecraft. The review considers stationary and deployable solar batteries used both on the first spacecraft and on space stations. The classification of solar cell designs is made taking into account their characteristic features. These features are the rigidity of the supporting structure, the method of placement in the starting position and the method of orientation towards the light source. The classification covered rigid panel solar arrays, flexible substrate solar panels, inflatable solar arrays, self-expanding solar arrays, and solar concentrator panels. In each design group of this classification, corresponding examples of solar cells are presented. The presented review and classification makes it possible to track trends in the development of solar array designs for spacecraft.

Modeling and disturbance suppression for spacecraft solar array systems subject to drive fluctuation

The vibration of flexible solar array during sun-tracking drive has been one of the challenging problems in modern space technology. In this work, the vibration disturbance of the drive induced solar array vibration (DISAV) to spacecraft platform is studied. A vibration disturbance model considering both the effect of rigid-flexible coupling and the drive unstable excitation is established for a typical solar array system. It extends the traditional dynamic model of the solar array system in terms of an explicit description for the vibration disturbance. Then, an external torque compensation scheme considering both vibration suppressing and driving reliability is proposed based on a failure-protection magnetorheological actuator. The extended minimal control synthesis algorithm is adopted to achieve the compensatory torque for vibration suppression. The numerical results indicate that the energy of the vibration disturbing the spacecraft is mainly provided by the unstable drive while the flexible vibration diffuses the disturbance. And the vibration disturbance suppression should be a main task for the vibration control during the sun-tracking drive. The reliability and effectiveness of the DISAV control with a magnetorheological actuator are theoretically demonstrated from the aspect of the failure protection and active control. The maximum disturbance torque is reduced up to 68.75% during the normal sun-tracking, which demonstrates the good suppression of DISAV by the proposed torque compensation scheme.

Equivalent Flux Function Smoothing in the Problem of Solar Cells Radiation Degradation Minimization During a Low-Thrust Geostationary Orbit Insertion

The paper describes a modified smooth approximation method for smoothing the dependence of the equivalent fluence of charged particles of the Earth’s Van Allen radiation belts on time and position in near-Earth space. It is shown that this method can be effectively used to solve the problems of minimizing radiation-induced degradation of spacecraft solar cell arrays. The latter problems involve the variation of the trajectory that the spacecraft follows during an electric propulsion transfer from an initial orbit to a geostationary one.

Analysis and Synthesis of Control Systems for Spacecraft Solar Arrays

The paper is devoted to the problem of creating highly reliable power supply systems for spacecrafts intended for long-term autonomous flights. Within its framework, the problem of synthesizing a control system for solar arrays is being solved. To solve this problem, a mathematical model of a solar panel was compiled, and a study of its static and dynamic characteristics was carried out. It was found that when the solar panel is controlled using a shunt switch with pulse-width modulation, resonance phenomena appear in the system, leading to an unacceptable change in the polarity of voltages on the photocells. The operating conditions of the solar panels, which exclude the occurrence of the indicated alternating voltages, are found, and appropriate recommendations are given for the choice of the quantization frequency in the system. On the basis of the recommendations received, the transition to a quasi-continuous representation of the control system was carried out, and a graphic-analytical synthesis of the controller providing the required quality indicators of the system was carried out. To ensure the survivability of the power supply system, a method is proposed for the hierarchical organization of the interaction of solar panels, which reproduces the homeostatic properties of biological structures in the system. This property is provided by automatic transfer of control to subsequent levels of the hierarchy as the energy resources of the previous levels are exhausted. In addition, selective control is applied only to that part of the total generated power, which is sufficient to counter the current disturbing influences on the system. This approach to control prevents cascading failures in the system. The paper presents simulation models on which all theoretical positions and methods proposed in the work are tested.

Power conditioning unit with independent extreme power control for two sections of small spacecraft solar arrays

the article analyzes the main approaches applied in the development process of power conditioning unit (PCU) model to ensure the efficient power use of non-orientable solar arrays for small satellites operated in low Earth orbits with altitudes ranging from 300 to 1500 km. Particular attention is paid to the description of PCU structure choice, as well as the way to achieve the technical specification goals, such as efficiency factor and specific power. A conclusion is made that the optimal structure to ensure PCU efficiency and reliability is the one with a series regulator and section-by-section solar array power extreme control. It is noted that this structure will make it possible to receive power from solar array sections regardless of the difference in temperature and illumination between them. Based on the quantity and duration analysis of the solar battery charge and discharge cycles, the charge and discharge modes of the batteries are defined as basic ones, so the efficiency factor in these modes is maximized. The results of the study described in the article can be used by enterprises specialized in developing small satellites and power conditioning units as a stepping stone for developing a PCU flight model to ensure the efficient power use of non-orientable solar arrays for small satellites.

Minimizing the Degradation of Triple Junction Solar Array of a Spacecraft during Geostationary Orbit Insertion

A method of reducing the radiation power degradation of spacecraft solar array during combined geostationary orbit insertion using a booster and an electric propulsion system is considered in this paper. The essence of the method is to optimize the shape of the transfer trajectory and the perigee argument of the intermediate orbit. The maximum principle is applied to the problem of optimizing the SA electrical power at the end of 15 year spacecraft’s operational lifetime (EOL). For this, the equation for the current SA power and the constraint on this power at the EOL is added to the equations of spacecraft motion. The closed-form solution to the adjoint equation to SA power is obtained. Calculation of SA radiation degradation was carried out using the models of charged particle fluxes of Earth’s radiation belts, AE8 MAX and AP8 MAX. To increase the EOL SA power by 0.16–0.66% of the SA power at the beginning of the transfer depending on the parameters of intermediate orbits. The additional characteristic velocity increased relative to the minimum time trajectories by 13–1087 m/s depending on the parameters of intermediate orbits.

Studying the Interaction of Hot Magnetospheric Plasma With the Protective Coating of the Spacecraft Solar Array

This paper presents the results of the simultaneous electron–proton irradiation of glass samples used as the protective coating of solar cells of high-altitude spacecraft. Combined irradiation with 20-keV electrons and protons was carried out in a vacuum chamber at a pressure of $10^{-4}$ Pa; charged particles fluxes were varied in the range of $10^{14}$ – $10^{15}\,\,\text{m}^{-2}\cdot \text {s}^{-1}$ . Structural changes of irradiated glass surfaces were studied by atomic force microscopy, and the reduction in glass transmittance due to the irradiation was measured in the wavelength range of 200–2500 nm. These changes in the surface structure and electrostatic discharge development under the combined irradiation are compared with the data obtained for separate exposure of the same glasses to electron or proton beams.

A Smoothed Eclipse Model for Solar Electric Propulsion Trajectory Optimization

Solar electric propulsion (SEP) is the dominant design option for employing low-thrust propulsion on a space mission. Spacecraft solar arrays power the SEP system but are subject to blackout periods during solar eclipse conditions. Discontinuity in power available to the spacecraft must be accounted for in trajectory optimization, but gradient-based methods require a differentiable power model. This work presents a power model that smooths the eclipse transition from total eclipse to total sunlight with a logistic function. Example trajectories are computed with differential dynamic programming, a second-order gradient-based method.

Optimized sleeve monopole antenna for detection of electrostatic discharge radiation of spacecraft solar array.

The spacecraft operating on a geosynchronous orbit (GEO) is affected by the plasma environment and is prone to electrostatic discharges (ESD) that affect the safe operation of the spacecraft, especially solar array, radiating electromagnetic (EM) signals of a few MHz to tens of MHz. The capture of radiated EM signals helps to better study the characteristics and suppress the occurrence of ESD. In this paper, based on the frequency band characteristics of the EM wave signal radiated by ESD in space, a novel optimized compact sleeve monopole antenna is designed by using three-dimensional EM field simulation software High Frequency Structure Simulator. The length of the proposed antenna is only about 52 mm and its center frequency is 30 MHz. The frequency band of the voltage standing wave ratio (VSWR) less than 2 is 28.8-31.1 MHz and that of less than 5 is 26.8 MHz-33.2 MHz, and the bandwidth is 21.3%. The sample preparation and VSWR test by using the vector network analyzer were carried out. Finally, the sensitivity test was carried out in an EM shielding room by the piezoelectric ceramic ESD simulation experiment. The test results show that the designed antenna can receive the EM wave signal radiated by the ESD stably and can be used for spacecraft discharge detection.

Control strategy of optimal deployment for spacecraft solar array system with initial state uncertainty

A control strategy combining feedforward control and feedback control is presented for the optimal deployment of a spacecraft solar array system with the initial state uncertainty. A dynamic equation of the spacecraft solar array system is established under the assumption that the initial linear momentum and angular momentum of the system are zero. In the design of feedforward control, the dissipation energy of each revolute joint is selected as the performance index of the system. A Legendre pseudospectral method (LPM) is used to transform the optimal control problem into a nonlinear programming problem. Then, a sequential quadratic programming algorithm is used to solve the nonlinear programming problem and offline generate the optimal reference trajectory of the system. In the design of feedback control, the dynamic equation is linearized along the reference trajectory in the presence of initial state errors. A trajectory tracking problem is converted to a two-point boundary value problem based on Pontryagin’s minimum principle. The LPM is used to discretize the two-point boundary value problem and transform it into a set of linear algebraic equations which can be easily calculated. Then, the closed-loop state feedback control law is designed based on the resulting optimal feedback control and achieves good performance in real time. Numerical simulations demonstrate the feasibility and effectiveness of the proposed control strategy.

Considerations of Flashover Propagation

Flashover propagation of plasma on a solar panel is considered from a physics standpoint. Although useful as a first attempt to describe the flashover phenomenon, the previous perimeter theory is shown to have many areas in need of improvement before it can be used as a guide for spacecraft solar array design. A new, more physical picture is developed here. Results from the U.S. Round-Robin Experiments on Plasma Propagation Speed are used to elaborate the physical principles. The new model explains simultaneous current onset, peaks, and cutoffs at all distances; multiple peaks moving at different velocities; and nonthermal directed velocities. Further elaboration is needed before such a model can be used as a reliable guide to flashover pulse peaks, widths, and electrostatic discharge; but, as it stands, the model corrects many of the perimeter theory’s deficiencies.

Residual life prediction for complex systems with multi-phase degradation by ARMA-filtered hidden Markov model

The performance of certain critical complex systems, such as the power output of ground photovoltaic (PV) modules or spacecraft solar arrays, exhibits a multi-phase degradation pattern due to the redundant structure. This pattern shows a degradation trend with multiple jump points, which are mixed effects of two failure modes: a soft mode of continuous smooth degradation and a hard mode of abrupt failure. Both modes need to be modeled jointly to predict the system residual life. In this paper, an autoregressive moving average model-filtered hidden Markov model is proposed to fit the multi-phase degradation data with unknown number of jump points, together with an iterative algorithm for parameter estimation. The comprehensive algorithm is composed of non-linear least-square method, recursive extended least-square method, and expectation–maximization algorithm to handle different parts of the model. The proposed methodology is applied to a specific PV module system with simulated performance measurements for its reliability evaluation and residual life prediction. Comprehensive studies have been conducted, and analysis results show better performance over competing models and more importantly all the jump points in the simulated data have been identified. Also, this algorithm converges fast with satisfactory parameter estimates accuracy, regardless of the jump point number.