Modern small satellites (MSS) are revolutionizing the space industry. They can drastically reduce the mission cost, and can make access to space more affordable. The relationship between a modern small satellite and a "conventional" large satellite is similar to that between a modern compact laptop and a "conventional" work-station computer. This paper gives an overview of antenna technologies for applications in modern small satellites. First, an introduction to modern small satellites and their structures is presented. This is followed by a description of technical challenges in the antenna designs for modern small satellites, and the interactions between the antenna and modern small satellites. Specific antennas developed for modern small-satellite applications are then explained and discussed. The future development and a conclusion are presented
Future space telescopes with diameter over 20 m will require new approaches: either high-precision formation flying or in-orbit assembly. We believe the latter holds promise at a potentially lower cost and more practical solution in the near term, provided much of the assembly can be carried out autonomously. To gain experience, and to provide risk reduction, we propose a combined micro/nano-satellite demonstration mission that will focus on the required optical technology (adaptive mirrors, phase-sensitive detectors) and autonomous rendezvous and docking technology (inter-satellite links, relative position sensing, automated docking mechanisms). The mission will involve two "3U" CubeSat-like nanosatellites ("MirrorSats") each carrying an electrically actuated adaptive mirror, and each capable of autonomous un-docking and re-docking with a small central "15U" class micro/nano-satellite core, which houses two fixed mirrors and a boom-deployed focal plane assembly. All three spacecrafts will be launched as a single ~40 kg micro-satellite package. The spacecraft busses are based on heritage from Surrey's SNAP-1 and STRaND-1 missions (launched in 2000 and 2013 respectively), whilst the optics, imaging sensors and shape adjusting adaptive mirrors (with their associated adjustment mechanisms) are provided by CalTech/JPL. The spacecraft busses provide precise orbit and attitude control, with inter-satellite links and optical navigation to mediate the docking process. The docking system itself is based on the electromagnetic docking system being developed at the Surrey Space Centre (SSC), together with rendezvous sensing technology developed for STRaND-2. On orbit, the mission profile will firstly establish the imaging capability of the compound spacecraft before undocking, and then autonomously re-docking a single MirrorSat. This will test the docking system, autonomous navigation and system identification technology. If successful, the next stage will see the two MirrorSat spacecraft undock and re-dock to the core spacecraft in a linear formation to represent a large (but sparse) aperture for high resolution imaging. The imaging of stars is the primary objective, but other celestial and terrestrial targets are being considered. Teams at CalTech and SSC are currently working on the mission planning and development of space hardware. The autonomous rendezvous and docking system is currently under test on a 2D air-bearing table at SSC, and the propulsion and precision attitude control system is currently in development. Launch is planned for 2016. This paper details the mission concept; technology involved and progress to date, focussing on the spacecraft buses
Experimental results on the performance of a control moment gyroscope cluster are presented. The goal is to design and evaluate a control moment gyroscope cluster for three-axis control for agile small satellites. The experimental data are compared with simulation (theoretical) results and both are used to verify the principles, advantages, and performance specifications of a control moment gyroscope cluster for a small satellite, in a practical way. Control moment gyroscope systems are considered in the literature to be more efficient devices, from an electrical power point of view, than current actuators such as reaction/momentum wheels. Experimental measurements are presented and then compared to two reaction wheels of different size. Control moment gyroscopes are shown to have a potential performance advantage over reaction/momentum wheels for spacecraft with agile requirements. Nomenclature H= observation matrix h = control moment gyroscope (CMG) angular momentum vector, N · m · s h 0 = CMG angular momentum, N · m · s I AB = air-bearing moment of inertia, kg · m 2 I CMG = flywheel moment of inertia, kg · m 2 I RW = reaction wheel moment of inertia, kg · m 2 I s = spacecraft moment of inertia, kg · m 2 K = Kalman gain N CMG = CMG torque vector, N · m N d = external disturbances, N · m N RW = reaction wheel torque, N · m N z = Z -axis CMG experimental torque, N · m P k = covariance matrix P k = covariance matrix update P act = actuator electrical power, W q ω = estimate of angular noise, rad/ṡ q ω = estimate of angular acceleration noise, rad/s 2 R = estimate of angular rate measurement noise, (rad/s) 2 t m = time to complete maneuver, s β = CMG pyramid skew angle, deg δ = gimbal angle vector, deg ε = scaled energy index, J/kg · m 2 θ = air-bearing rotation angle, deg ω = angular rate vector, rad/s ω AB = air-bearing angular speed, rad/s ω k = intermediate angular rate state vector, rad/ṡ ω k = intermediate angular acceleration state vector, rad/s 2 ω w = flywheel speed, rad/s ω z = angular rate of air bearing, rad/s ω 0 = angular rate initial value, rad/s
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