This paper presents a new algorithm for multi-input multi-output (MIMO) system identification in the time domain using impulse responses. The algorithm is suitable for the on-orbit system identification of spacecraft using the responses to thruster impulse inputs measured by typical satellite on-board sensors. The Eigensystem Realization Algorithm (ERA) realizes a multi-input multi-output (MIMO) system using asynchronous impulse responses in the time domain. Our new method identifies the input and output matrices of a MIMO collocated system by applying a recursive least-squares iteration scheme to refine the matrices obtained from conventional ERA. In this manner, the input matrix is considered to be constructed by the mode shape vectors and the actuator sensitivity matrix. A numerical simulation of an actual spacecraft, the Engineering Test Satellite-VI (ETS-VI), is performed to verify the algorithm. The nominal dynamics model of ETS-VI, which has six rigid body modes and fourteen elastic modes due to large flexible solar panels, is excited by six body-mounted thrusters, and the translational velocities and attitude rates are measured simultaneously. Our method successfully identifies all of the fourteen natural frequencies, damping ratios and mode shape vectors, confirming its validity.
Acoustically induced random vibration of satellite equipment mounted on honeycomb panels is a critical design consideration in satellite equipment development. Prediction of this random vibration is performed in the early stage of satellite design to specify the design limit value of random vibration excitation for satellite equipment. Various prediction methods for response prediction using Statistical Energy Analysis (SEA) have been developed: (i) NASA Lewis method, (ii) point-mass impedance method, and (iii) area-coupling impedance method. However, the first method has limited accuracy for heavy and concentrated equipment, the second one often overestimates, and the third one requires a detailed parameter. A new method combining the asymptotic apparent mass of specific equipment with NASA Lewis method is proposed herein. This proposed method takes the elastic behavior of satellite equipment rather than a rigid mass. The acoustic excitation experiments for nine real satellites (404 equipments in all) were conducted to compare existing methods to the proposed method statistically. Results show that the proposed method provides the most accurate prediction in the important frequency range.
This paper proposes a basic method for formation flight based on Hill's equations of motion. It proposes a trajectory design of minimum fuel consumption for spacecraft relative position change by two impulses. Focusing on enlarging the formation size, we consider a problem to obtain the time and phase of the second impulse when the phase of the first impulse is given. First, the phase of the second impulse is obtained as a function of the time of the second impulse. Subsequently, the time of the second impulse is obtained to minimize the total fuel consumption. There are five cases to apply two impulses for minimum fuel consumption which depends on the phase of the first impulse. Numerical studies for enlarging the formation size by applying two impulses are carried out in order to verify the trajectory design.
A satellite antenna alignment technique is proposed to ensure terrestrial service quality for users. The antenna bore sight orientation is calculated directly from measured data acquired from general ground receivers, which intercept the communication radio waves from any position on the earth's surface. The method coordinates the satellite pointing parameters with signal strength at the receivers while considering location-specific geographical and antenna radiation characteristics and control accuracy. The theoretical development and its validity are examined in the course of equation derivation. Actual measured data of an existing satellite at the maneuver was applied to the method, and the capability was demonstrated and verified. With the wide diversity of satellite usage, such as for mobile communications, temporary network deployment or post-launch positioning accommodations, the proposed method provides a direct evaluation of satellite communication performance at the service level, in conjunction with using high frequency spot beam antennas, which are highly susceptible to pointing gain. This can facilitate swift and flexible satellite service planning and deployment for operators.
A simplified estimating method for the Shock Response Spectrum (SRS) envelope at the spacecraft interface near the V-band clamp separation device has been established. This simplified method is based on the pyroshock analysis method with a single degree of freedom (D.O.F) model proposed in our previous paper. The parameters required in the estimating method are only geometrical information of the interface and a tension of the V-band clamp. According to the use of these parameters, a simplified calculation of the SRS magnitude at the knee frequency is newly proposed. By comparing the estimation results with actual pyroshock test results, it was verified that the SRS envelope estimated with the simplified method appropriately covered the pyroshock test data of the actual space satellite systems except some specific high frequency responses.
Specific cutting resistance was determined through airplane experiments under low relative gravity conditions such as μ G, 0.15 G, 0.3 G, 0.5 G, and 1 G. Results showed that the relationship between specific cutting resistance and relative gravity could be expressed as a linear function. As for numerical analysis by discrete element method (DEM), the data of spring constant in a contact model of DEM could be treated as constant in the analysis of specific cutting resistance under low gravity conditions from the viewpoint of stress-oriented soil-machine interaction. Moreover, the numerical analysis by DEM with change of relative gravity and the corresponding modification of consolidation time is found to be sufficient to obtain a specific cutting resistance at a given low gravity condition below 1 G.
In this study, we propose a basic method for the construction of a formation flight around an eccentric reference orbit under the influence of gravitational disturbances (J2). The relative motion between two spacecraft in the formation flight is affected by J2. The effects of this perturbing force are analyzed, and the initial conditions of the formation flight for the suppression of these effects are derived on the basis of these analytical results.
A simplified analysis model based on the frequency response analysis and the wave propagation analysis was established for predicting Shock Response Spectrum (SRS) on the composite panel subjected to pyroshock loadings. The complex composite panel was modeled as an isotropic single layer panel defined in NASA Lewis Method. Through the conductance of an impact excitation test on a composite panel with no equipment mounted on, it was presented that the simplified analysis model could estimate the SRS as well as the acceleration peak values in both near and far field in an accurate way. In addition, through the simulation for actual pyroshock tests on an actual satellite system, the simplified analysis model was proved to be applicable in predicting the actual pyroshock responses, while bringing forth several technical issues to estimate the pyroshock test specifications in early design stages.