光电经纬仪测量飞行器三维坐标方法及误差分析

介绍一种用于光电经纬仪精确测量飞行器三维坐标的计算方法,该方法利用空间两异面光轴公垂线估计交会点位置,克服了常规处中的模型误差,分析了观测量误差 产生的原因并提出了改正方法,对定位测量误差进行了仿真,结果表明,利用该方法并合理布站可显著提高从标测量精度,对于机动大目标的坐标测量误差小于0．20m。     

An attitude determination algorithm for a spacecraft using nonlinear filter

In this paper, the algorithm for a real time attitude estimation of a spacecraft motion is investigated. The proposed algorithm for attitude estimation is the second order nonlinear filter form not containing truncation error in estimation values. The proposed second order nonlinear filter has improved performance compared with the EKF (extended Kalman filter), because the algorithm does not contain any truncation bias and covariance of the estimator is compensated by the nonlinear terms of the system. Therefore, the proposed second order nonlinear filter is a suboptimal estimator. However, the proposed estimator requires a lot of computation because of an inherent nonlinearity and complexity of the system model. For more efficient computation, this paper introduces a new attitude estimation algorithm using the state divided technique for a real time processing which is developed to provide an accurate attitude determination capability under a highly maneuverable dynamic environment. To compare the performance of the proposed algorithm with the EKF, simulations have been performed with various initial values and measurement covariances. Simulation results show that the proposed second order nonlinear algorithm outperforms the EKF. The proposed algorithm is useful for a real time attitude estimation since it has better accuracy compared with the EKF and requires less computing time compared with any existing nonlinear filters.     

Characterizing Debris Clouds Created in Oblique Orbital Debris Particle Impact

The problem of pollution within Earth’s orbital environment has gained considerable recognition over the past decades. Determining adequate passive protection schemes is an unending process that attempts to meet different objectives for widely varying types of missions. Significant amounts of resources have been expended toward development of numerical and analytical models that model the response of a variety of target systems under high-speed orbital debris impacts. The objective of the study whose results are presented herein was to improve upon an existing oblique hypervelocity impact model that characterizes the various secondary debris clouds created in such an impact. This was accomplished by reducing the model’s dependence on empirical user-defined parameters and by correcting an error in one of its equations. Predictions of the improved model are compared with numerical simulations generated during previous impact studies under comparable conditions. It is found that the improved model does a reasonable job of predicting the characteristics of the secondary debris clouds created in an oblique hypervelocity impact.     

Rendezvous Guidance Trajectories via Multiple-Subarc Sequential Gradient-Restoration Algorithm

We consider the three-dimensional rendezvous between a target spacecraft in a circular orbit and a chaser spacecraft with an initial separation distance and an initial separation velocity. We assume that the chaser spacecraft has variable mass and that its trajectory is governed by three controls, one determining the thrust magnitude and two determining the thrust direction. We employ the Clohessy–Wiltshire equations, describing the relative motion of the chaser vis-à-vis the target, and the multiple-subarc sequential gradient-restoration algorithm to produce first optimal trajectories and then guidance trajectories for the following problems: P1—minimum time, fuel free; P2—minimum fuel, time free; P3—minimum time, fuel given; P4—minimum fuel, time given; and P5—minimum time×fuel, time and fuel free. Clearly, P1 and P2 are basic problems, while P3, P4, and P5 are compromise problems. Problem P1 leads to a two-subarc solution including a max-thrust subarc followed by another max-thrust subarc. Problem P2 leads to a four-subarc solution including two coasting subarcs alternating with two max-thrust subarcs. Problems P5 leads to a three-subarc solution including two max-thrust subarcs alternating with one coasting subarc. Problems P3 and P4 include P1, P2, and P5 as particular cases and lead to two-, three-, or four-subarcs solutions depending on the prescribed value of fuel or time. For all problems, the thrust magnitude control is saturated at one of its extreme values: in optimization studies, we determine the best thrust direction controls; in guidance studies, we force the thrust direction controls to be constant in each subarc and determine the best thrust direction parameters. Of course, the time lengths of all the subarcs must also be determined. The computational results show that, for Problems P1–P5, the performance index of the multiple-subarc guidance trajectory approximates well the performance index of the multiple-subarc optimal trajectory: the pairwise relative differences in performance index are less than 1/100 in all cases. To sum up, the produced guidance trajectories are highly efficient and yet quite simple in implementation.     

Hypervelocity Impact Penetration Phenomena in Aluminum Space Structures

All long‐duration spacecraft are susceptible to high‐speed impacts by meteoroids and pieces of orbiting space debris. Damage to critical spacecraft systems caused by such impacts can lead to spacecraft failure and loss of life. In order to develop adequate protection against penetration for crew compartments and other critical spacecraft systems, an aerospace design engineer must possess a full understanding of the penetration mechanics involved in the hypervelocity impact loading of a variety of structural components. This paper describes the results of an experimental investigation of the penetration phenomena associated with oblique hypervelocity projectile impact of aluminum dual‐wall structures. Equations that quantitatively describe these phenomena are obtained through a regression of hypervelocity impact test data. These equations characterize observed penetration phenomena as functions of the geometric and material properties of the impacted structure and the diameter, obliquity, and velocity of the impacting projectile. A review of the test data shows that oblique hypervelocity impact penetration phenomena are strongly dependent on impact obliquity and therefore can differ significantly from those associated with normal high‐speed impacts. It is concluded that the possibility of non‐normal impacts and their effects on structural integrity must be considered in the design of any structure that is to be exposed to the hazardous meteoroid and space debris environment.     

Design and analysis of the composite lattice frame of a spacecraft solar array

A novel design of the composite structural lattice frame for the spacecraft solar arrays is presented in the paper. The frame is composed of two flat lattice composite plates assembled into the three-dimensional panel using frame-like connectors. Design, fabrication, modelling and modal analysis of the panel solar arrays based on the proposed technology are discussed. The lattice panels are modelled as three-dimensional frame structures composed of beam elements subjected to the tension/compression, bending and torsion using the specialised finite-element model generator/design modeller. Results of the calculations of the frequencies and vibration forms for the lattice panels with various types of supports imitating the ways the panels can be attached to the spacecraft body, deployment must, and adjacent solar panels are presented and discussed. The lattice frame design for maximum fundamental frequency is performed subject to constraints imposed on the geometrical parameters of the solar panel.     

Cost effective honeycomb and multi-layer insulation debris shields for unmanned spacecraft

Ways to improve the tolerance of unmanned spacecraft to hypervelocity impact are presented. Two new honeycomb and multi-layer insulation (MLI) shields were defined: (1) double honeycomb, and (2) enhanced or toughened MLI (with additional Kevlar 310 and/or Betacloth layers). Following hypervelocity impact testing, a new ballistic limit threshold was defined, based on rear facesheet perforation and witness plate damage characteristics. At 12 km/s, the ballistic limit of single honeycomb was 0.58 mm (aluminium sphere), rising to 0.91 mm for double honeycomb, 1.00 mm for double honeycomb with MLI and 1.17 mm for double honeycomb with toughened MLI. A damage equation, based on the modified Cour-Palais equation with ESA constants, was compared with the data and found to be conservative. The impact angle exponent was increased in order to reduce the equation under-prediction for the oblique incidence data. An equivalent rear wall thickness was defined in order to distinguish between shield types above 7 km/s. The spacecraft survivability analysis showed that the double honeycomb and toughened MLI significantly reduced the number of perforating particles over the baseline single honeycomb design. The mass increase of these shields is approximately 1.2 kg/m² for double honeycomb and 0.8 kg/m² for toughened MLI.     

Space flight tests of attitude determination using GPS

Preliminary space flight results of attitude determination using GPS are presented from a spacecraft in low Earth orbit. Relative position measurements accurate to the sub-centimetre level are made among multiple GPS antennas mounted on the space vehicle. A Trimble Navigation TANS Quadrex (a GPS receiver specially adapted for attitude determination by Stanford University) is used as a differential carrier phase sensor for the flight. Four GPS antennas are mounted on the zenith face of RADCAL, a polar orbiting, gravity-gradient-stabilized Air Force Space Test Program Satellite, built by Defense Systems, Inc. The four antennas are equally spaced about the perimeter of the 30 inch diameter cylindrical spacecraft bus. The Quadrex receiver measures the phase of the L-band GPS carrier (1575 MHz) at each of up to four antennas for up to six GPS satellites simultaneously. From these measurements, an initial assessment of attitude determination in space is performed in post-processing. For RADCAL, the attitude solution is greatly overdetermined. In a preliminary evaluation of system performance, the system accuracy is determined through measurement self-consistency. Analysis of the attitude motion in the context of a gravity gradient dynamic model yields further insight into the system performance.     

Nonlinear vibration analysis of spacecraft with local nonlinearity

Most spacecraft are usually assembled from some simple substructures by different kinds of connectors, which include various kinds of joints and hinges. Most of the connectors have properties of nonlinearity, and can strongly affect the dynamic characteristics of spacecraft. Mathematical models of such spacecraft usually have a large number of degrees of freedom (DOFs), but their nonlinear connectors are generally spatially localized. In general, it is impractical and time-consuming to directly calculate the frequency response of the spacecraft using current methods. To enhance the calculation efficiency of the frequency response, an improved approach is proposed in the present paper. With describing functions (DFs) and linear receptance data, the kinetic equations are firstly converted into a set of complex algebraic equations whose dimension is only associated with nonlinear DOFs and interested DOFs. Subsequently the number of iterative equations is reduced and only related to nonlinear DOFs. Hence the improved approach can be applicable to large-scale and complicated finite element (FE) models. An FE model of a satellite with some nonlinear joints is used to show and demonstrate the usefulness of the proposed method. Besides, the effects induced by nonlinear joints on payloads’ vibration of the satellite are discussed.