We discuss the possibility to exploit electromagnetic cloaking for enhancing the performances of communication antennas installed on nanosatellite platforms. As a case study, we consider a CubeSTAR system and show that properly designed metasurfaces can be effectively used to minimize the impact of its deployable payload sensors on both radiation and electrical characteristics of the antennas. The results discussed here pave the way to new strategies for the design of extremely compact nanosatellite platforms equipped with multifunctional antennas and sensors.

Satellite Applications of Electromagnetic Cloaking

Challenges of direct-to-Earth (DTE) laser communications (lasercom) between spacecraft in low-Earth orbit (LEO) or medium-Earth orbit (MEO) and ground terminals can include short duration transmission windows, long time gaps between the transmission windows, deleterious effects of atmospheric turbulence, and the inability to operate in cloudy weather. Direct-link optical communications systems described herein can have data rates that are high enough to empty high-capacity on-board buffer(s) (e.g., having a capacity of at least about 1 Tb to hundreds of Tb) of a spacecraft in a single pass lasting only tens of seconds to a few minutes (e.g., 1-15 minutes), and overprovisioning the buffer capacity accounts for variations in the latency between links. One or more distributed networks of compact optical ground terminals, connected via terrestrial data networks, receive and demodulate WDM optical data transmissions from a plurality of orbiting spacecraft (e.g., satellites).

Ground terminal design for high rate direct to earth optical communications

Described herein are systems and methods for attitude determination using infrared Earth horizon sensors (EHSs) with Gaussian response characteristics. Attitude information is acquired by detecting Earth's infrared electromagnetic radiation and, subsequently, determining the region obscured by Earth in the sensors' fields of view to compute a nadir vector estimation in the spacecraft's body frame. The method can be applied when two sensors, each with known and distinct pointing directions, detect the horizon, which is defined as having their fields of view partially obscured by Earth. The method can be implemented compactly to provide high-accuracy attitude within small spacecraft, such as CubeSat-based satellites.

/pdf/attitude-determination-using-infrared-earth-horizon-sensors-xcfs9tl60b.pdf

Attitude determination using infrared earth horizon sensors

Hybrid actuator optimal angular momentum management with analytical solution for spacecraft attitude agile maneuvering mission

Because they are inexpensive platforms for satellites, CubeSats have become a low-cost way for universities and even developing countries to have access to space technology. This paper presents the ITASAT design, particularly the Attitude Determination and Control Subsystem, the Onboard Software, and the Assembly,Integration and Testing program. The ITASAT is a 6U CubeSat nano-satellite in development at the Instituto Tecnologico de Aeronautica, in Sao Jose dos Campos, Brazil. The platform and its subsystems will be provided by industry while the payloads are being designed and developed by the principal investigators. The ITASAT Attitude Determination and Control Subsystem will rely on a 3-axis magnetometer, 6 analog cosine sun sensors, 3-axis MEMS gyroscopes, 3 magnetic torque coils, and 3 reaction wheels. The Attitude Determination and Control Subsystem operating modes, control laws, and embedded software are under the responsibility of the Instituto Tecnologico de Aeronautica. A Kalman filter shall be employed to estimate the quaternion attitude and gyroscope biases from sensor measurements. The Attitude Determination and Control Subsystem operating modes are the nominal mode, with geocentric pointing attitude control and the stabilization mode, in which only the satellite angular velocity is controlled. The nominal mode will be split into 2 sub-modes: reaction wheel control plus magnetic wheel desaturation and 3-axis magnetic attitude control. Simulation results have shown that the attitude can be controlled with 1-degree accuracy in nominal mode with the reaction wheels, but these errors grow as much as 20 degrees or higher with the 3-axis magnetic control.

http://www.jatm.com.br/ojs/index.php/jatm/article/download/614/617

The ITASAT CubeSat Development and Design

We present a satellite attitude control system design using low-cost hardware and software for a 1U CubeSat. The attitude control system architecture is a crucial subsystemfor any satellite mission since precise pointing is often required to meet mission objectives. The accuracy and precision requirements are even more challenging for small satellites where limited volume, mass, and power are available for the attitude control system hardware. In this proposed embedded attitude control system design for a 1U CubeSat, pointing is obtained through a two-stage approach involving coarse and fine control modes. Fine control is achieved through the use of three reaction wheels or three magnetorquers and one reaction wheel along the pitch axis. Significant design work has been conducted to realize the proposed architecture. In this paper, we present an overview of the embedded attitude control system design; the verification results fromnumerical simulation studies to demonstrate the performance of a CubeSat-class nanosatellite; and a series of air-bearing verification tests on nanosatellite attitude control systemhardware that compares the performance of the proposed nonlinear controller with a proportional-integral-derivative controller.

https://downloads.hindawi.com/journals/jcse/2013/657182.pdf

Thomas Wright

Papers

Design of attitude control systems for cubesat-class nanosatellite