The present study uses a method to address microvibrations effects on an optical satellite by combining simulations and experiments based on high-precision acceleration sensors. The displacement and angular displacement of each optical component can be obtained by introducing flywheel perturbation data from a six-component test bench to the finite element model of the optical satellite. Combined with an optical amplification factor inferred from the linear optical model, the pixel offset of the whole optical system is calculated. A high accuracy and broad frequency range for a new microvibration measurement experimental system is established to validate the simulation. The pixel offset of the whole optical system can be measured by testing the acceleration signals of each optical component and calculating optical amplification factors. The results are consistent with optical imaging test results, indicating correctness of the experimental scheme and the effectiveness of the simulation. The results suggest that the effect of microvibrations on a camera can be verified by using mechanical simulators instead of a whole optical camera for the experiment scheme, which is demonstrated to be an effective way for increasing efficiency in jitter measurements.

The separation mechanism is a critical device that transports and releases satellites during launch and on-orbit. However, as satellites become smaller and more compact, the traditional belt-locking device or pyrotechnic release and separation device cannot meet the micro-nano satellite’s separation requirements. A novel separation mechanism kinematic system was designed, analyzed, and experimentally verified to achieve non-interference and non-pyrotechnic separation of the satellites from the launch vehicle while maintaining the initial separation attitude. First, an overall structural strategy for the kinematics system was proposed based on the MF (product code of the satellite) satellite’s structural properties. The structural characteristics of the separation mechanism were also confirmed by the principle of energy conservation. Then, a finite element model and a dynamic model of the kinematic coupling system between the MF satellite and the separation mechanism during launch and deployment were constructed, along with an analysis of the mechanical characteristics and a kinematic simulation. Finally, the mechanical characteristics and dependability of the separation mechanism were verified using a ground vibration test and a separation test. The prototype test results demonstrate that the separation process is essentially identical to the numerical simulation results and that the MF satellite can be deployed successfully without interference, with the MF satellite’s initial separation attitude fully satisfying the designed technical indexes.

A sun-tracking solar array is an effective solution to the increasing energy demand of spacecrafts. However, the driving torque fluctuation of the solar array driving assembly (SADA) and its oscillatory motion lead to the micro-vibration problem of the spacecraft. In this article, a disturbance torque model of the sun-tracking solar array, which takes its friction and flexibility into consideration, is established. Furthermore, a test platform was built to measure the disturbance force/torque, and a solar array simulator (SAS) was designed to replace the solar array in the ground experiments. Finally, the disturbance torque of the SADA-driven SAS is simulated, and the model is validated by comparing the simulation results with the experimental results. These results show that the frequency error of the model is less than 0.648%, and the amplitude error of the corresponding frequency is less than 22.33%, which indicates that the proposed model can effectively predict the disturbance torque generated by the sun-tracking solar array in orbit. The research provides theoretical guidance for the system optimization design and micro-vibration suppression of spacecrafts.

Body-mounted solar panels are extensively utilized in satellite construction due to their simple structure and robust vibration resistance. The quantity and arrangement of support points on the body-mounted solar panel significantly affect its natural frequency. Thus, the design of these support points is a crucial aspect of the design process for body-mounted solar panels. This study presents a method for determining the support points of body-mounted solar panels, enabling rapid and precise identification of the quantity and positioning of these points based on the stated natural frequency in the design. First, a new algorithm is proposed, based on the finite element method, to optimize the positioning of support points on the body-mounted solar panel without the need for remeshing. Utilizing this algorithm, the distinct impacts of support point positioning and stiffness on the natural frequency of the solar panel are investigated, and the practical principles are proposed for quickly and accurately identifying the optimal locations of support points to maximize the natural frequency of the solar panel, given a predetermined number of support points. Subsequently, based on Courant–Fischer’s theorem, a method to ascertain the least quantity of support points through two modal analyses is presented. By integrating the aforementioned principles and method, a two-step procedure for identifying the quantity and positioning of support points is developed. Ultimately, the proposed two-step procedure is implemented in the design of the solar panel of the Jilin-1XXX satellite. The modal test reveals that the natural frequency of the solar panel surpasses the design index criteria, hence validating the efficacy and feasibility of the optimal design technique for the support points of the body-mounted solar panel presented in this study.

The vibrational response and stress of satellites subjected to acoustic excitation are essential components to consider in the design process of large satellite constructions. To precisely forecast the vibrational response and stress subjected to acoustic excitation of a large satellite honeycomb sandwich plate, this article employs finite element modeling software to create a finite element model of satellite, equivalent honeycomb panels to orthotropic shear plates of identical stiffness and dimensions, and convert the acoustic excitation into random pressure on the surface of the flat plate, applying it to the surface of the satellite, then the vibration response and stress analysis subjected to acoustic excitation can be performed. The honeycomb structure collapsed after noise testing on a specific model of a large satellite, the aforementioned technique for response and stress validation was employed: the stress simulation analysis revealed that the maximum shear stress at the fracture site of the honeycomb panel was 0.45 MPa, exceeding the ultimate stress value of 0.44 MPa for the sparse honeycomb core at that location, leading to the collapse and fracture of the honeycomb panel. To resolve this matter, altering the honeycomb core design framework, a local dense honeycomb structure was implemented. The simulated shear stress of the honeycomb core is 0.54 MPa, which is below the stress limit of 2.43 MPa for the dense honeycomb core. To verify the feasibility of the plan, the noise test was conducted once again. Owing to the incapacity to test the shear stress of the honeycomb core, conducted strain testing on the surface of the honeycomb collapse and derive stress results by calculation, the test results deviate from the modeling values by 8.92%. And the maximum discrepancy between the simulated noise response and the experimental noise response is 0.58 g, validated the efficacy and precision of the simulation method, and successfully resolved the issue of damage to satellite honeycomb panels. This simulation method can precisely forecast the vibrational response and stress of satellites subjected to acoustic excitation.

Microvibrations are one of the main factors contributing to platform jitter and a decline in pointing stability and precision. Among various disturbance sources, reaction wheel assembly (RWA) is one of the most significant ones and has drawn the interest of numerous scholars. How to evaluate the influence of the disturbances of RWA on the pointing accuracy of a spacecraft is an arduous task because it involves multiple disciplines. The acquisition of the frequency response function (FRF) from the disturbance source to the sensitive payload is one of the most crucial stages in integrated modeling. Direct measurement of the FRF in the six directions is challenging because of the restricted room for RWA installation in a spacecraft, particularly in small satellites. In this paper, a general method based on the H_v algorithm to obtain the FRF is presented. This method only needs the RWA, itself, as an excitation source. Then, in order to acquire the FRF, we use an optical remote-sensing satellite as the research object. The peak positions of FRF obtained by different RWAs are basically the same, while the amplitudes are slightly different, indicating that this method is effective. This method takes into consideration the coupling between the RWA and the spacecraft, making it possible to multiply the RWA disturbance measured at a fixed interface with the FRF, to determine the image motion of the sensitive payload.