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.

This paper addresses the degradation of imaging quality in high-resolution CubeSats caused by micro-vibrations from attitude control flywheels. It proposes a micro-vibration suppression scheme that incorporates multi-disciplinary integrated modeling, dual passive vibration isolation, and multi-level verification. A comprehensive model encompassing flywheel disturbance, optics, attitude control, and structure is developed to elucidate the transmission dynamics of micro-vibrations from the source to the optical payload. A dual suppression system utilizing silicone rubber isolators is engineered for both the disturbance source (flywheel) and the payload (optical camera). By optimizing stiffness matching and damping, it achieves a balance between isolation efficiency and stability in attitude control. A three-tier verification system comprising “numerical simulation–ground microgravity testing–on-orbit imaging” has been established. The findings indicate that the dual isolation system diminishes the pixel offset amplitude of the optical payload to under 0.1 pixels (down to the 0.02 pixel level in the high-frequency band), with an isolation efficiency of 80%. Consistent outcomes from terrestrial and orbital validation affirm the engineering viability of the plan. This research offers theoretical backing for the precise control of micro-vibrations in micro-nano satellites, thereby enhancing their utility in high-resolution remote sensing applications.

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.

The ground modal test is an important approach to the natural frequency of solar arrays to support the attitude control of spacecraft. However, for the batch production of small satellites, the accuracy and efficiency of traditional ground modal testing methods are limited. This shortcoming restricts the development of satellite constellations. Based on the encapsulation method widely used in the computer field, this paper proposed a natural frequency identification method of deployable solar arrays with multiple plates. This method is of high accuracy and efficiency that meets the demand of attitude control and makes sense to accelerate the batch production of small satellites. First, a suspended modal test system with gravity compensation function is designed. Second, the mathematical model of the test system is established. Abstracting parts of the parameters of the test object into an encapsulated entity, the mathematical model is simplified by equivalent variables. Thus, the direct mapping relationship between the ground test result and the true natural frequency is proposed. Finally, to verify the identification accuracy, finite element analysis (FEA) and the ground modal test of a two‐folder solar array simulant are carried out. The results show that the relative error of the first‐order natural frequency after correction and the theoretical value is less than 3%. Meanwhile, the identification accuracy of the ground modal test is improved by more than 50%. This method improves the availability of ground test results and reduces the calculation amount, so that it is convenient for engineering applications.