Model predictive and reallocation problem for CubeSat fault recovery and attitude control

Recently, thanks to the miniaturization and high performance of commercial-off-the-shelf (COTS) computer systems, small satellites get popular. However, due to the very expensive launching cost, it is critical to reduce the physical size and weight of the satellite systems such as cube satellites (CubeSats), making it infeasible to install high capacity batteries or solar panels. Thus, the low-power design is one of the most critical issues in the design of such systems. In addition, as satellites make a periodic revolution around the Earth in a vacuum, their operating temperature varies greatly. For instance, in a low earth orbit (LEO) CubeSats, the temperatures vary from 30 to −30 degrees Celsius, resulting in a big thermal cycle (TC) in the electronic parts that is known to be one of the most critical reliability threats. Moreover, such LEO CubeSats are not fully protected by active thermal control and thermal insulation due to the cost, volume, and weight problems. In this paper, we propose to utilize temperature sensors to maximize the lifetime reliability of the LEO satellite systems via multi-core mapping and dynamic voltage and frequency scaling (DVFS) under power constraint. As conventional reliability enhancement techniques primarily focus on reducing the temperature, it may cause enlarged TCs, making them even less reliable. On the contrary, we try to maintain the TC optimal in terms of reliability with respect to the given power constraint. Experimental evaluation shows that the proposed technique improves the expected lifetime of the satellite embedded systems by up to 8.03 times in the simulation of Nvidia’s Jetson TK1.

Small satellites or CubeSats orbiting in low Earth orbit (LEO) have become increasingly popular in Earth Observation missions, where high-resolution imaging is essential. Due to the lower mass of these spacecrafts, they are more sensitive to vibrations, and image quality can be particularly negatively affected by micro-vibrations. These vibrations originate from on-board subsystems, such as the Attitude Determination and Control System (ADCS), which uses reaction wheels to change the orientation of the satellite. The main goal of our research was to analyze these micro-vibrations so that the acquired data could be used for post-correction of camera images. Obuda University, as a participant in a research project, was tasked with designing and building a micro-vibration measuring device for the LEO CubeSat called WREN-1. In the first phase of the project, the satellite was launched into orbit, and test data were collected and analyzed. The results are presented in this article. Based on the data obtained in this way, the next step will be to analyze the images taken at the same time as the vibration measurements and to search for a correlation between the image quality and the vibrations. Based on the results of the entire project, it could be possible to improve the image quality of the onboard cameras of microsatellites.

Three-axis attitude control by two-step rotations using only magnetic torquers in a low Earth orbit near the magnetic equator

Magnetic torquers are an effective and reliable technology for the attitude control of small satellites in low Earth orbit. Such actuators operate by generating a magnetic dipole which interacts with the magnetic field of the Earth. The main difficulty in the design of attitude control laws based on magnetic torquers is that the torques they generate are instantaneously constrained to lie in the plane orthogonal to the local direction of the geomagnetic field vector, which varies according to the current orbital position of the spacecraft. This implies that the attitude regulation problem is formulated over a time-varyingmodel. In recent years, this control problem has been studied extensively, either using methods based on averaged models or via approaches which exploit the quasi-periodic variability of the geomagnetic field. With the exception of other approaches based on Model Predictive Control, none of the above actually exploits at the design stage the fact that the geomagnetic field can be reliably measured on board and, therefore, the above mentioned time-variability of the attitude dynamics can be represented in LPV form. Therefore, in this chapter an LPV approach to the problem of magnetic attitude control law design is proposed. To this purpose, an LPV model of the attitude dynamics is first derived, LPV control laws suitable for on board implementation are synthesized and eventually tested in simulation.

In this paper we report on the application of a model predictive control-based approach to the design of a controller for formation keeping and formation attitude control, with applications to spacecraft formation flight problems such as NASA's DS3 mission. Control laws for formation keeping and attitude control are designed using a combined approach of feedback linearization and model predictive control. Actuator saturation is incorporated into the controller design. Switching between coordinated frames is incorporated to overcome singularities associated with local feedback linearization.

Linear time-varying fractional-order model predictive attitude control for satellite using two reaction wheels

Attitude control system plays the important role for to maintain the satellite to desired orientation. To control the satellite it is necessary to do the attitude stabilization. Attitude stabilization achieved by Star sensor, sun sensor, Earth sensors. Attitude control is mainly use for antenna pointing accuracy, camera focus to earth surface and solar panel pointing toward sun. Due to tumbling effects, satellite will rotate all the direction in the space. To maintain the orientation of the satellite it is necessary to design the attitude determination and control. Satellite consider as the rigid body. Inertia matrix describes the rigid body dynamics. The orientation of the satellite determine by Euler angle and Quaternion. The Low earth orbiting satellite will have enormous amount of aerodynamic drag stinking the satellite body and gravitational attraction another problem. Because of that satellite dwell, time is reduced. It means satellite more time spending particular part of the earth. The attitude estimation is measures by the orientation of vectors. Attitude estimation means to find the position and orientation of flying object with respect to the fixed reference of reference. Vector remains considered in the frame of reference to compute for find the orientation of the body of the satellite in the inertial reference system. The Earth is an inertial reference frame, and Satellite is a body frame. Attitude sensor used to measure the satellite orientation in the reference frame. This will help in accurately predicting the orbit deviation and a control system to correct if any by providing the satellite momentum means ‘mass in motion’ changes in a body rapidly in Low earth orbit due to centripetal force acting on a satellite. Attitude control system (ACS) need the numerical simulation to find the required torque demand by the help of difference between reference input (Attitude) signal and feedback signal measure by attitude sensor to trim the control surface maintain the actuator required orientation . The results will consist of two parts the first part consisting of the attitude estimation using Euler angle and Quaternion method, second part consist of estimate the control torque from magnetic torquer and error estimation using non-linear filter (Unscented Kalman Filter) with MATLAB simulation. Â

Control of a Benchmark Boiler Process Model with DMC and QDMC

Magnetic torquers are used for the attitude control of small satellites, such as CubeSats with Low Earth Orbit (LEO). During the design of magnetic torquers, it is necessary to confirm if its magnetic dipole moment is enough to control the satellite attitude. The magnetic dipole moment can affect the detumbling time and the satellite rotation time. In addition, it is also necessary to understand how to design the magnetic torquer for operation in a CubeSat under the space environment at LEO. This paper reports an investigation of the magnetic dipole moment and the magnetic field generated by a circular air-coil magnetic torquer using experimental measurements. The experiment testbed was built on an air-bearing under a magnetic field generated by a Helmholtz coil. This paper also describes the procedure to determine and verify the magnetic dipole moment value of the designed circular air-core magnetic torquer. The experimental results are compared with the design calculations. According to the comparison results, the designed magnetic torquer reaches the required magnetic dipole moment. This designed magnetic torquer will be applied to the attitude control systems of a 1U CubeSat satellite in the project “KNACKSAT.”

This dissertation concerns the fundamental problems of windup and process directionality in "input-output linearizing" control of multivariable nonlinear processes with actuator saturation nonlinearities. Two classes of general nonlinear but affine-in-control processes are considered: (a) delay-free processes with full state measurements, and (b) processes with incomplete state measurements and dead-times. The two fundamental problems are addressed by using a model predictive approach. The complete connections between input-output linearizing control and model predictive control are established: input-output linearizing control is simply shortest-prediction-horizon model predictive control. The model predictive control approach leads to nonlinear control laws that can be parameterized into two distinct parts: (i) an input-output linearizing controller that inherently includes an integral windup compensator and (ii) an optimal directionality compensator. The directionality compensator is a quadratic program that is trivially solvable online. In the case that the characteristic (decoupling) matrix of process is diagonal, the optimal directionality compensator is identical to "clipping". For general processes, however, neither direction preservation nor clipping can compensate for process directionality optimally. The concrete connections between the derived control laws, and modified internal model control and model state feedback, are established. When one of the derived control laws is applied to time-invariant linear processes, the resulting linear controller will exactly be a modified internal controller and in a special case, a model state feedback controller. The application and superior performance of the derived control laws and the optimal directionality compensator are demonstrated by numerical simulations of chemical and biochemical processes with input saturation nonlinearities.

For spacecraft containing components that survive re-entry, it is important to de-orbit the satellite over a non-populated area. Because most CubeSats (cube satellites that conform to the CubeSat form factor) do not have their own propulsion systems and cannot perform a de-orbit burn, aerodynamic drag modulation presents an attractive solution to re-enter the satellite at the desired location. The University of Florida Advanced Autonomous Multiple Spacecraft (ADAMUS) lab has developed a drag de-orbit device (D3) for CubeSats, which are affordable systems for demonstrating attitude and orbit control. The device consists of four retractable tape-spring booms that are designed and manufactured to validate the targeted re-entry of a CubeSat in Low-Earth Orbit (LEO). By modulating the D3 drag area, orbital maneuvering and controlled re-entry can be performed. This paper outlines the functional and vibration testing procedures and results for the completed device. The complete drag device was taken to NASA Ames Research Center and subjected to random vibrations at 9.6 times the force of gravity. The device survived vibration testing with no obvious damage. After vibration testing, it was determined that no major mechanical design changes to the D3 will need to be made. Future work for this project includes assembling a final, flight-ready drag device that will be attached to a CubeSat for launch.

This paper investigates the model predictive control approach to optimize the operation of hybrid microgrid system which contain hybrid energy harvesting and storage (HEHS) devices and bidirectional 3-Level Neutral Point Clamped (NPC) based 3-Phase Voltage Source Converter (VSC). The intermittent nature of renewable energy sources has direct effect over the health of the connected microgrid system. Hence the essential salient feature needed for hybrid microgrid is to encounter the discrete nature of loading conditions. Which arises the need of that system which can deal with such kind of practical problems in the microgrid system. This work revisits and utilizes the key elements of Model Predictive Control (MPC) to achieve the voltage regulation in the DC link of AC/DC microgrid, reduction in grid harmonics and managing the power for the whole system. Along with this the given topology is optimally utilizing the DC link split capacitors by maintaining the same voltage level on them. Besides, the performance of MPC based grid connected HEHS system with 3-Level neutral point clamped converter topology is demonstrated via simulation in the MATLAB/Simulink environment.

The impact of uneven heating on a satellite structure in low Earth orbit has been considered using the example of 3U CubeSat. The calculations showed that the thermal deformation of CubeSat structure in orbit caused a deviation between normals to opposite small satellite sides of about 0.03°. Such a deviation is commensurate with the required satellite pointing accuracy approximately 0.1° necessary for satellite laser communication. It means that to solve similar problems in the CubeSat designing that require such or better CubeSat pointing accuracy, it is necessary to take into account the expected satellite structure thermal deformation.

Aerodynamic and gravity gradient based attitude control for CubeSats in the presence of environmental and spacecraft uncertainties

The measurement of air density in the Earth’s thermosphere has a wide range of scientific applications from space weather to upper atmosphere dynamics, but also technical applications from satellite control to predictions of atmospheric reentry of space debris. This study models the torques applying on a three-unit CubeSat in low Earth orbit to infer the capability of such platforms to measure the air density along their orbit. Realistic noise levels of available CubeSat components are used, and sensitivity to the various noise sources is presented. The precise knowledge of the spacecraft attitude, angular acceleration, residual magnetic dipole, and center of gravity is critical to allow proper air density retrieval. Winds in the thermosphere also have a significant impact on the thermosphere density retrieval, suggesting that this parameter can also be constrained. Attitude control is not necessary if the attitude itself is properly known. The application to the EntrySat CubeSat predicts that such retrieval is possible at altitudes lower than 200 km with errors lower than 30%. The air density retrieval from CubeSat platforms will open new capabilities to infer upper atmosphere dynamics.