Pico-satellite Autonomous Navigation with Magnetometer and Sun Sensor Data

This article presents a near-Earth satellite orbit estimation method for pico-satellite applications with light-weight and low-power requirements. The method provides orbit information autonomously from magnetometer and sun sensor, with an extended Kalman filter （EKF）. Real-time position/velocity parameters are estimated with attitude independently from two quantities： the measured magnitude of the Earth’s magnetic field, and the measured dot product of the magnetic field vector and the sun vector. To guarantee the filter’s effectiveness, it is recommended that the sun sensor should at least have the same level of accuracy as magnetometer. Furthermore, to reduce filter’s computation expense, simplification methods in EKF’s Jacobian calculations are introduced and testified, and a polynomial model for fast magnetic field calculation is developed. With these methods, 50% of the computation expense in dynamic model propagation and 80% of the computation burden in measurement model calculation can be reduced. Tested with simulation data and compared with original magnetometer-only methods, filter achieves faster convergence and higher accuracy by 75% and 30% respectively, and the suggested simplification methods are proved to be harmless to filter’s estimation performance.

Space Flight Validation of Design and Engineering of the ZDPS-1A Pico-satellite

The ZDPS-1A pico-satellites are the first satellites in China within the 1-10 kg mass range that are successfully operated on orbit. Unlike common pico-satellites, they are designed to be ＂larger but stronger＂ with more powerful platforms and unique payloads so as to bear a better promise for real applications. Through their space flight mission, the functionality and perform- ance of the two flight models are tested on orbit and validated to be mostly normal and in consistency with design and ground tests with only several inconforming occasions. Moreover, they have worked properly on orbit for one year so far, well exceed- ing their life expectancy of three months. Therefore, the space flight mission has reached all its goals, and verified that the design concept and the engineering process of the pico-satellites are sufficient in allowing them the desired functionality and perform- ance in, and the adaption to the launch procedure and the low-Earth orbit space environment. In the foreseeable future, the plat- form together with the design concept and the engineering process of the pico-satellites are expected to be applied to more com- plicated real space applications.

New technique: Design and implementation of the highly-reliable, low-cost housekeeping system in the ZDPS-1A pico-satellite

The ZDPS-1A pico-satellite designed in Zhejiang University with a mass of 3.5 kg and a power consumption of less than 3.5 W is the smallest satellite in China up to now. The housekeeping system (HKS) is the core part of ZDPS-1A. The reliability of HKS has an important influence on the safety of the satellite. Traditional fault-tolerant methods do not apply to ZDPS-1A due to such pico-satellite characteristics as light weight, compactness in size, energy saving, and high integration. This paper deals with a highly-reliable, low-cost design for HKS using industrial devices. The reliable strategies of HKS include a dual modular redundancy scheme, CPU warm backup, a static triple modular redundancy scheme, and two-level watchdogs. Recursive experiments, special tests, and environmental tests show that this system meets the design target. This design has already been applied to ZDPS-1A, which was launched to execute in-orbit tasks on Sept. 22, 2010. To date, the satellite has been in a proper state for more than 15 months.

Magnetometer-only linear attitude estimation for bias momentum pico-satellite

Satellite attitude information is essential for pico-satellite applications requiring light-weight,low-power,and fast-computation characteristics.The objective of this study is to provide a magnetometer-only attitude estimation method for a low-altitude Earth orbit,bias momentum pico-satellite.Based on two assumptions,the spacecraft spherical symmetry and damping of body rates,a linear kinematics model of a bias momentum satellite's pitch axis is derived,and the linear estimation algorithm is developed.The algorithm combines the linear Kalman filter(KF) with the classic three-axis attitude determination method(TRIAD).KF is used to estimate satellite's pitch axis orientation,while TRIAD is used to obtain information concerning the satellite's three-axis attitude.Simulation tests confirmed that the algorithm is suited to the time-varying model errors resulting from both assumptions.The estimate result keeps tracking satellite attitude motion during all damping,stable,and free rotating control stages.Compared with nonlinear algorithms,such as extended Kalman filer(EKF) and square root unscented Kalman filer(SRUKF),the algorithm presented here has an almost equal performance in terms of convergence time and estimation accuracy,while the consumption of computing resources is much lower.

Attitude stabilization of a pico-satellite by momentum wheel and magnetic coils

The three-axis active attitude control method with a momentum wheel and magnetic coils for a pico-satellite is considered. The designed satellite is a 2.5 kg class satellite stabilized to nadir pointing. The momentum wheel performs a pitch-axis momentum bias, nominally spinning at a particular rate. Three magnetic coils are mounted perpendicularly along the body axis for precise attitude control through the switch control mechanism. Momentum wheel start up control, damping control and attitude acquisition control are considered. Simulation results show that the proposed combined control laws for the pico-satellite is reliable and has an appropriate accuracy under different separation conditions. The proposed strategy to start up the wheel after separation from the launch vehicle shows that its pitch momentum wheel can start up successfully to its nominal speed from rest, and the attitude convergence can be completed within several orbits, depending on separation conditions.

Magnetometer Compensation Scheme and Experimental Results on ZDPS-1A Pico-satellite

In a pico-satellite with small volume, measurements from on-board three-axis magnetometer （TAM） are not accurate, as it can be easily disturbed by other electronic systems. To improve its accuracy, a scheme of compensation methods is introduced in this article. The scheme is based on an improved meast, rement model of pico-satellite TAM, and it mainly consists of three steps. First, in satellite design stage, several techniques are recommended to simplify the afterwards compensations. Then after satellite assembly, TAM ground tests and pre-launch calibration with least-square batch filter are introduced to improve magnetometer performance. At the end, a post-launch calibration with unscented Kalman filter （UKF） is implemented with in-orbit data. The compensation scheme is used in the development of Chinese pico-satellite ZDPS-1A made by Zhejiang University. Results show that with the introduced compensation scheme, the maximum error of ZDPS-1A TAM can be reduced from 80 mG to 6 naG （1 G=10^-4T）.

Graph-tree-based software control flow checking for COTS processors on pico-satellites

This paper proposes a generic high-performance and low-time-overhead software control flow checking solution, graph-tree-based control flow checking （GTCFC） for space-borne commercial-off-the-shelf （COTS） processors. A graph tree data structure with a topology similar to common trees is introduced to transform the control flow graphs of target programs. This together with design of IDs and signatures of its vertices and edges allows for an easy check of legality of actual branching during target program execution. As a result, the algorithm not only is capable of detecting all single and multiple branching errors with low latency and time overheads along with a linear-complexity space overhead, but also remains generic among arbitrary instruction sets and independent of any specific hardware. Tests of the algorithm using a COTS-processor-based onboard computer （OBC） of in-service ZDPS-1A pico-satellite products show that GTCFC can detect over 90% of the randomly injected and all-pattern-covering branching errors for different types of target programs, with performance and overheads consistent with the theoretical analysis; and beats well-established preeminent control flow checking algorithms in these dimensions. Furthermore, it is validated that GTCGC not only can be accommodated in pico-satellites conveniently with still sufficient system margins left, but also has the ability to minimize the risk of control flow errors being undetected in their space missions. Therefore, due to its effectiveness, efficiency, and compatibility, the GTCFC solution is ready for applications on COTS processors on pico-satellites in their real space missions.

A composite optimization method for separation parameters of large-eccentricity pico-satellites

A spacecraft＇s separation parameters directly affect its flying trace. If the parameters exceed their limits, it will be difficult to adjust the flying attitude of the spacecraft, and the spacescraft may go off-track or crash. In this paper, we present a composite optimization method, which combines angular velocities with external moments for separation parameters of large-eccentricity pico-satellites. By changing the positions of elastic launch devices, the method effectively controls the popping process under the condition of less change in the separation mechanism. Finally, the reasons for deviation of angular velocities and unreliable optimization results are presented and analyzed. This optimization method is proved through a ground test which offsets the gravity. Simulation and test results show that the optimization method can effectively optimize the separation parameters of large-eccentricity pico-satellites. The proposed method adapts particularly to the fixed and non-stable status elastic parameters, the distribution of all kinds of elastic devices, and large-eccentricity spacecrafts for which attitude corrections are difficult. It is gen- erally applicable and easy to operate in practical applications.