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An onboard autonomous technique can significantly reduce the costs of the mission. The purpose of this paper is to deal with the autonomous orbit determination and attitude determination of a satellite based on the sun, the earth and the moon sensors.

The models of the conical earth scanner are presented, and its measurement with information from the sun and the moon sensors is processed to simultaneously acquire the orbit and attitude of the satellite via extended Kalman filter.

The numerical simulation shows that the presented method can obtain the orbit and attitude information precisely; even in the new moon period, it can be used to get the satisfied results.

Autonomous orbit determination and attitude determination based on direction information of celestial objects, such as sun, earth and moon, are put forward. The method improves the survival ability of the satellite and decreases its reliance on the ground stations.

The purpose of this paper is to propose an attitude determination and control scheme for a low‐cost Micro‐satellite with defective inertia. Restricted by the payload design, the z‐axis inertia of this satellite is larger than the x and y axes, which is unstable for natural attitude dynamics.

An original operation mode is designed to avoid z axis from long‐time pointing to the sun during damping, which avoids some unexpected damage. In attitude determination design, EKF and UKF algorithms are compared on estimation accuracy, convergence time and computation complexity in attitude estimation design, which is referred to determine the final estimation scheme. A DSP‐based hardware solution is achieved and a semi‐physical testing and simulation system is built.

Simulation results show the 3‐axis stable mode can be built with the proposed scheme, and the unprotected facet of the satellite can be kept away from long‐time pointing to the sun.

The proposed ADCS scheme can be a reference for the future Micro‐satellite programs which share the similar configuration.

Attitude determination and control subsystem (ADCS) is a vital part of earth observation satellites (EO-Satellites) that governs the satellite’s rotational motion and pointing. In designing such a complicated sub-system, many parameters including mission, system and performance requirements (PRs), as well as system design parameters (DPs), should be considered. Design cycles which prolong the time-duration and consequently increase the cost of the design process are due to the dependence of these parameters to each other. This paper aims to describe a rapid-sizing method based on the design for performance strategy, which could minimize the design cycles imposed by conventional methods.

The proposed technique is an adaptation from that used in the aircraft industries for aircraft design and provides a ball-park figure with little engineering man-hours. The authors have shown how such a design technique could be generalized to cover the EO-satellites platform ADCS. The authors divided the system requirements into five categories, including maneuverability, agility, accuracy, stability and durability. These requirements have been formulated as functions of spatial resolution that is the highest level of EO-missions PRs. To size, the ADCS main components, parametric characteristics of the matching diagram were determined by means of the design drivers.

Integrating the design boundaries based on the PRs in critical phases of the mission allowed selecting the best point in the design space as the baseline design with only two iterations. The ADCS of an operational agile EO-satellite is sized using the proposed method. The results show that the proposed method can significantly reduce the complexity and time duration of the performance sizing process of ADCS in EO-satellites with an acceptable level of accuracy.

Rapid performance sizing of EO-satellites ADCS using matching diagram technique and consequently, a drastic reduction in design time via minimization of design cycles makes this study novel and represents a valuable contribution in this field.

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.

The understandable recent trend in sensor design has been to exploit the rapid advance in digital electronics to reduce reliance on analogue circuits. In contrast to this general trend some researchers have been inspired by biological systems to design smart imaging sensors based upon collective analogue computation in networks of resistors. This has resulted in sensor designs which efficiently extract information from a large volume of data whilst reducing manufacturing costs by improving yield.

The measurement of heat flux is of importance to the development of aerospace engine as basic physical quantities in extreme environment. Heat radiation is one of the basic forms of heat transfer phenomenon. The structure optimizing can improve the performance and infrared absorptivity of the thin film sensor.

This paper designed one kind of thin film heat flux sensor (HFS) with antireflective coating based on transparent conductive oxide thermopile. The introduced membrane structure is so thin that it has little impact on sensor performance. Fabrication of thin film sensors were fabricated by physical vapor deposition (PVD) process.

The steady-state and dynamic response characteristics of the HFS were investigated by calibration platform. The experimental results shown that the absorptivity of the membrane structure (for1070nm) improved compared with that before optimization. The sensitivity of heat flux gauge was 48.56 µV/ (kW/m2) and its frequency response was determined to be about 1980 Hz.

The thin film HFS uses thermopile based on Indium Tin Oxid and In2O3. The antireflective coating is introduced to hot endpoint of HFS to improve sensitivity on laser thermal source. The infrared optical properties of membrane layer structure were investigated. The steady-state and the transient response characteristics of the heat flux sensor were also investigated.

Amperometric gas sensors are commonly used in air quality monitoring in long-term measurements. Baseline shift of sensor responses and power failure may occur over time, which is an obstacle for reliable operation of the entire system. The purpose of this study is to check the possibility of using PCA method to detect defected samples, identify faulty sensor and correct the responses of the sensor identified as faulty.

In this work, the authors present the results obtained with six amperometric sensors. An array of sensors was exposed to sulfur dioxide at the following concentrations: 0 ppm (synthetic air), 50 ppb, 100 ppb, 250 ppb, 500 ppb and 1000 ppb. The damage simulation consisted in adding to the sensor response a value of 0.05 and 0.1 µA and replacing the responses of one of sensors with a constant value of 0 and 0.15 µA. Sensor validity index was used to identify a damaged sensor in the matrix, and its responses were corrected via iteration method.

The results show that the methods used in this work can be potentially applied to detect faulty sensor responses. In the case of simulation of damage by baseline shift, it was possible to achieve 100% accuracy in damage detection and identification of the damaged sensor. The method was not very successful in simulating faults by replacing the sensor response with a value of 0 µA, due to the fact that the sensors mostly gave responses close to 0 µA, as long as they did not detect SO₂ concentrations below 250 ppb and the failure was treated as a correct response.

This work was inspired by methods of simulating the most common failures that occurs in amperometric gas sensors. For this purpose, simulations of the baseline shift and faults related to a power failure or a decrease in sensitivity were performed.