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This paper aims to propose a constant-gain Kalman Filter algorithm based on the projection method and constant dimension projection, which ensures that the dimension of the observation matrix obtained is maintained when there is a satellite with multiple sensors.

First, a time-invariant observation matrix is determined with the projection method, which does not require the Jacobi matrix to be calculated. Second, the constant-gain matrix replaces the EKF (extended Kalman filter) gain matrix, which requires online computation, considerably improving the stability and real-time properties of the algorithm.

The simulation results indicate that compared to the EKF algorithm, the constant-gain Kalman filter algorithm has a considerably lower computational burden and improved real-time properties and stability without a significant loss of accuracy. The algorithm based on the constant dimension projection has better real-time properties, simpler computations and greater fault tolerance than the conventional EKF algorithm when handling an attitude determination system with three or more star trackers.

In satellite attitude determination systems, the constant-gain Kalman Filter algorithm based on the projection method reduces the large computational burden and improve the real-time properties of the EKF algorithm.

The purpose of this study is to address the concept and the step-by-step procedure of a high-precision optical alignment test for spacecrafts using digital theodolites. The proposed scheme focuses on the non-contact alignment qualification of spacecraft components during the integration and test phases until the launch event.

The proposed approach is based on the exploitation of the auto-collimation feature of theodolites and several prisms attached to the requested component and satellite configuration. As soon as the misalignment measurement including the difference between the real and desired attitude or position aberration of an instrument is made, the results must be transformed from the component level to the system level for misalignment error identification in the spacecraft dynamic model.

The paper introduces the main instruments, the defined coordinate systems and the architecture of the optical spacecraft misalignment test. Moreover, the guideline of the test implementation and the resulting data process have been presented carefully.

There is no limitation associated with this method because the procedure is applicable for high-precision typical missions.

This paper describes a fully implementable scheme to examine any possible inaccuracy in mounting of the spacecraft components both in position and orientation. The test can be performed without the need for a huge budget or complicated hardwares.

The contribution of this work revolves around illustrating the context and procedure of the spacecraft misalignment test which has remained unknown in literature despite the frequent implementation in the different satellite projects.

In this chapter, the authors review talent management in the research university sector, business schools in particular. The authors adopt an “exclusive” perspective on talent management, assuming that some scholars contribute disproportionately to organizational performance. The authors identify two particular groups of scholars likely to be the target of exclusive talent management practices in business schools: (i) faculty on a tenure track career path and (ii) “star” tenured faculty with exceptionally strong track records. Focusing on these current and potential future “stars,” the authors review and discuss talent management practices related to talent identification, recruitment and selection, performance management, talent development, benefits and rewards, and tenure, promotion, and retention. In the extant literature, these topics have been mostly examined in the general university environment and less so in the business school context. This is somewhat problematic given that business schools have their own special characteristics. Moreover, some of the reviewed topics – especially talent development – have received only marginal scholarly interest thus far. Based on this literature review, and by drawing on their own experience working in different roles in academia, the authors highlight some of the gaps in the current body of knowledge and propose an agenda for future research.

In the assembly process of the satellite, there will be multiple installation and disassembly operations for the solar wing and the main satellite body (or simulator). However, the traditional method of orientation adjustment by theodolite and two-axis turntable is difficult to coordinate three rotation angles of yaw, pitch and roll, which leads to the complexity of actual operation and dependency on manual experience. Therefore, this paper aims to propose a new method to achieve rapid and precise orientation adjustment.

The similarity relation of the orientation variation matrix in a different coordinate system is studied, and a mapping model of the similarity relation is established. By using multiple element matrices to construct the original rotation matrix, the mapping is solved in quaternion form. Taking the theodolite as a measuring instrument and the Stewart platform as a control equipment, an experiment on installing the solar wing is performed to validate the effectiveness of the algorithm.

Based on the solving algorithm, the orientation adjustment process is simplified to a three-step fixed mode, which is three adjustments to get the parameter of the mapping model, one to adjust the component in place and another to further fine tuning. The final orientation deviation is less than 0.003° and close to the level of using a laser tracker, achieving the required accuracy of 0.0115°.

This paper reveals the similarity relation of the variation matrix in the process of orientation adjustment and presents a new method to achieve rapid and precise orientation adjustment for the large-scale component.

This paper aims to present an attitude determination and control system for a nano‐astrometry satellite which requires precise angular rate control. Focus of the research is methods to achieve the requirement.

In order to obtain astrometry data, the satellite attitude should be controlled to an accuracy of 0.05°. Furthermore, attitude spin rate must be controlled to an accuracy of 4×10⁻⁷ rad/s during observation. In this paper the following unique ideas to achieve these requirements are introduced: magnetic disturbance compensation and rate estimation using star blurred images.

This paper presents the feasibility of a high accurate attitude control system in nano‐ and micro‐satellite missions.

This paper presents a possibility of the application of nano‐satellites to remote‐sensing and astronomy mission, which requires accurate attitude control.

Originalities of the paper are the methods to achieve the high accurate attitude control: magnetic disturbance compensation and angular rate estimation using star images.

This paper seeks to provide an insight into the design and development of the thermal test model (TTM) of X‐Sat, a 120 kg class micro‐satellite, being developed at the Centre. This model was specifically constructed for carrying out a thermal balance test (TBT) in a 4 m diameter vertical thermal vacuum chamber.

The construction of the thermal model followed a structural mock‐up model which was modified thermally to suit the purpose. Specific and careful consideration was given to the geometry and, more importantly, thermal characteristics such as thermal mass, surface properties, etc. to mimic the actual satellite configuration as closely as possible. Test plans were devised to qualify the fabricated components to meet the out‐gassing and other thermal requirements for the model. Design and qualification of supporting frame and linkages for TBT are also covered.

It is possible to simulate the thermal characteristics of a micro‐satellite in orbit under a different mission scenario through proper scaling and using alternative material options while developing TTM.

The paper discusses in detail the simplified cost‐effective approach of constructing TTM and also outlines the various issues to be considered for a TBT. It provides valuable information needed for micro‐satellite designers.

Autonomous orbit determination using geomagnetic measurements is an important backup technique for safe spacecraft navigation with a mere magnetometer. The geomagnetic model is used for the state estimation of orbit elements, but this model is highly nonlinear. Therefore, many efforts have been paid to developing nonlinear filters based on extended Kalman filter (EKF) and unscented Kalman filter (UKF). This paper aims to analyze whether to use UKF or EKF in solving the geomagnetic orbit determination problem and try to give a general conclusion.

This paper revisits the problem and from both the theoretical and engineering results, the authors show that the EKF and UKF show identical estimation performances in the presence of nonlinearity in the geomagnetic model.

While EKF consumes less computational time, the UKF is computationally inefficient but owns better accuracy for most nonlinear models. It is also noted that some other navigation techniques are also very similar with the geomagnetic orbit determination.

The intrinsic reason of such equivalence is because of the orthogonality of the spherical harmonics which has not been discovered in previous studies. Thus, the applicability of the presented findings are not limited only to the major problem in this paper but can be extended to all those schemes with spherical harmonic models.

The results of this paper provide a fact that there is no need to choose UKF as a preferred candidate in orbit determination. As UKF achieves almost the same accuracy as that of EKF, its loss in computational efficiency will be a significant obstacle in real-time implementation.

The purpose of this paper is to overcome the limitations of existing celestial horizon references, and improve the navigation accuracy of the strap‐down inertial navigation system/celestial navigation system (SINS/CNS) integrated system with an innovative scheme of deep integration.

First, a novel conception of mathematical horizon reference (MHR) provided by the strap‐down matrix of SINS is proposed. Then, the realization mechanism of the MHR‐based vertical vector is introduced from the viewpoint of vector rotation. Moreover, the MHR implementation scheme of high precision and reliability is presented, and on this basis, the method which utilizes vertical vector to achieve celestial navigation is introduced. In addition, with considering the characteristics of SINS and the MHR‐based CNS, the SINS/CNS deep integrated navigation system and its specific realization are proposed. Finally, simulation tests are implemented to validate this SINS/CNS deep integrated navigation method.

The innovative SINS/CNS deep integrated system could make full use of SINS and CNS navigation information to achieve higher navigation accuracy for the long‐duration and high‐altitude vehicles.

This paper provides a novel realization method of high precision MHR and the MHR‐based SINS/CNS deep integration.

APPLIED Technology, Middle East and European marketing and technical support representative of PF Industries Inc, will exhibit ground support equipment supplied to airlines worldwide.