Shock Characteristics and Protective Design of Equipment During Spacecraft Docking Process

The shock loads generated by spacecraft during docking can cause functional failure and structural damage to aerospace electronic equipment and even lead to catastrophic flight accidents.There is currently a lack of systematic and comprehensive research on the shock environment of spacecraft electronic equipment due to the diversity and complexity of the shock envi-ronment.In this paper,the validity of the finite element model is verified based on the sinusoidal vibration experiment results of the spacecraft reentry capsule.The method of shock dynamic response analysis is used to obtain the shock environment of electronic equipment under different shock loads.The shock response spectrum is used to describe the shock environment of aerospace electronic equipment.The results show that the resonance frequency error between the sinusoidal vibration experiment and the model is less than 4.06％.When the docking relative speed of the reentry capsule is 2 m/s,the shock response spectrum values of one of the equipment are 30 m2/s,0.67 m/s,and 0.059 m,respectively.The wire rope spring on the mating surface can provide vibration isolation and shock resistance.An increase in spring damping coefficient results in a decrease in the amplitude and time of the vibration generated.An increase in spring stiffness reduces the input of shock load within a certain range.These research results can provide guidance for the design and evaluation of shock environmental adaptability of aerospace electronic equipment.     

航天器多层绝热材料的非稳态传热分析

根据反射屏能量平衡建立了多层绝热材料的非稳态传热方程。以双面镀铝涤纶薄膜为研究对象,在不进行预处理、表面印花处理两种实验条件下,得到了绝热材料中各反射屏温度随航天器飞行过程中的具体数值变化规律。最后,在特定的两个时刻分析了绝热材料的当量热导率及其分量,发现残留气体导热约占70%。     

Analytical solution of attitude motion for spacecraft with a slewing appendage

This paper investigates pitch motion and in orbital plane elastic vibration of a spacecraft with a flexible beam type appendage undergoing prescribed slew maneuver. The governing equations are transformed into a standard quasi-linear form, and then solved by Butenin's variation of parameters approach. Validity of the analytical solutions is assessed over a range of system parameters and initial conditions by comparing them with the results of numerical integration. The results show that they are very good approximations and provide extensive insight into the dynamical response of the system.     

De-wrinkling of pre-tensioned membranes

Thin membranes are used in the spacecraft industry as extremely lightweight structural components. They need to be stiffened, usually by applying discrete forces, and this increases their susceptibility to wrinkling in regions where high tensile stresses develop. We consider a regular polygonal membrane uniformly loaded at its corners by equal forces and we prevent wrinkle formation by trimming the edges of the polygon into very gentle curves. We confirm this performance through simple physical experiments using Kapton, a typical membrane material and, using computational analysis, we show how the distribution of compressive stresses, responsible for causing wrinkles, dissipates following trimming. Finally, we accurately predict the required level of trimming for any number of sides of polygon using a simple, linear model, which invokes a plate-bending analogy.     

Variations of the flux and the energy spectrum of neutrons born in the galactic cosmic ray proton interactions with the matter of Earth''s satellites and orbital stations

This paper presents the energy spectra of secondary neutrons obtained from Monte Carlo simulations of the interactions of protons of galactic cosmic rays with the material of a spacecraft, and also the variation of the fluxes of these particles as a function of latitude. The variations of the secondary neutron flux with changing geometry and orientation of the craft, and detector position were evaluated as well. The simulations were conducted over a wide range of masses of spacecraft, from 300 kg (Earth Probe Total Ozone Mapping Spectrometer) to 420,000 kg (International Space Station in its full configuration), with calculations for the ISS also being carried out for its first development stage (mass of approx. 20,400 kg) and its configuration with a mass of approx. 81,000 kg. The calculations were carried out for circular orbits at altitudes of 500 km and for the energy intervals 0.001–10 MeV and 10–10,000 MeV, for both maximum and minimum solar activity. The results obtained agree reasonably well with results from previous calculations and do not contradict experimental data.     

An improved nonlinear control strategy for deep space formation flying spacecraft

This paper aims to provide further study on the nonlinear modeling and controller design of formation flying spacecraft in deep space missions. First, in the Sun-Earth system, the nonlinear formation dynamics for the circular restricted three-body problem （CRTBP） and elliptic restricted three-body problem （ERTBP） are presented. Then, with the Floquet mode method, an impulsive controller is developed to keep the Chief on the desired Halo orbit. Finally, a nonlinear adaptive control scheme based on Nonzero set- point LQR and neural network is proposed to achieve high precision formation maneuver and keeping. The simulation results indicate that the proposed nonlinear control strategy is reasonable as it considers not only the orbit keeping of the Chief, but also the formation modeling inaccuracy. Moreover, the nonlinear adaptive control scheme is effective to improve the control accuracy of the formation keeping.     

Singularly perturbed feedback linearization with linear attitude deviation dynamics realization

A new approach for feedback linearization of attitude dynamics for rigid gas jet-actuated spacecraft control is introduced. The approach is aimed at providing global feedback linearization of the spacecraft dynamics while realizing a prescribed linear attitude deviation dynamics. The methodology is based on nonuniqueness representation of underdetermined linear algebraic equations solution via nullspace parametrization using generalized inversion. The procedure is to prespecify a stable second-order linear time-invariant differential equation in a norm measure of the spacecraft attitude variables deviations from their desired values. The evaluation of this equation along the trajectories defined by the spacecraft equations of motion yields a linear relation in the control variables. These control variables can be solved by utilizing the Moore–Penrose generalized inverse of the involved controls coefficient row vector. The resulting control law consists of auxiliary and particular parts, residing in the nullspace of the controls coefficient and the range space of its generalized inverse, respectively. The free null-control vector in the auxiliary part is projected onto the controls coefficient nullspace by a nullprojection matrix, and is designed to yield exponentially stable spacecraft internal dynamics, and singularly perturbed feedback linearization of the spacecraft attitude dynamics. The feedback control design utilizes the concept of damped generalized inverse to limit the growth of the Moore–Penrose generalized inverse, in addition to the concepts of singularly perturbed controls coefficient nullprojection and damped controls coefficient nullprojection to disencumber the nullprojection matrix from its rank deficiency, and to enhance the closed loop control system performance. The methodology yields desired linear attitude deviation dynamics realization with globally uniformly ultimately bounded trajectory tracking errors, and reveals a tradeoff between trajectory tracking accuracy and damped generalized inverse stability. The paper bridges a gap between the nonlinear control problem applied to spacecraft dynamics and some of the basic generalized inversion-related analytical dynamics principles.     

Chaotic attitude motion of a magnetic rigid spacecraft and its control

This paper deals with chaotic attitude motion of a magnetic rigid spacecraft with internal damping in a circular orbit near the equatorial plane of the earth. The dynamical model of the problem is established. The Melnikov analysis is carried out to prove the existence of a complicated non-wandering Cantor set. The dynamical behaviors are numerically investigated by means of time history. Poincare map, power spectrum and Lyapunov exponents. Numerical simulations indicate that the onset of chaos is characterized by the intermittency as the increase of the torque of the magnetic forces and decrease of the damping. The input-output feedback linearization method is applied to control chaotic attitude motions to the given fixed point and periodic motion.     

Evolution of morphologic properties on the preparation of Ir/Al2O3 catalysts with high metallic contents

Ir/Al₂O₃ catalysts with high metallic contents are applied on satellite thruster to decompose hydrazine. The present work has as principal aim the study of the morphologic evolution of Ir/Al₂O₃ catalysts with metallic contents from 12 to 30 wt.%. The catalysts were prepared through consecutive impregnations from the H₂IrCl₆ precursor, using three different types of aluminas. The specific surface area, volume and distribution of pore size, specific metallic area and metallic particles average diameter, as well as the mechanical resistance were determined. Results show that the Ir addition leads to a decrease of the specific surface area and the pores volumes, while increases the mechanical resistance. Values for average diameter of metallic particles are comprised between 1.4 and 2.4 nm when the metallic content increases from 12 to 30 wt.%. Catalysts containing 30 wt.% of Ir presents specific metallic areas around 30 m²/g, although pores volumes and distributions of pore size were considerably different for the three supports. Their metallic particles dispersion and size values are very close to those of a commercial catalyst Shell 405, even though the preparation methods were different. These results show that there is a strong interaction between the alumina and the iridium precursor.     

Minimum Time Control of Flexible Spacecraft by Hamilton's Principle

Modern space vehicles structure requisites are getting more and more stringent and complex as mission tasks become more sophisticated. This leads to the necessity of developing analysis methods that take into account structure flexibility and the need of reducing manoeuvre time as much as possible. In this work, a method based on the Hamilton Principle in its weak mixed form is developed, in which co-ordinates derivatives do not appear, but only their virtual variations. The proposed formulation is able to take into account system flexibility and saturation constraints on control torques and forces. A non-linear variational condition is obtained, which can be solved by means of a time-finite-element technique to give the minimum-time solutions of the control problem. The solutions for slewing manoeuvres are given, along with a new solution of the distributed optimal control problem.