This paper presents the feasibility of performing multiple spacecraft per aperture (MSPA) ground operations by using a low-cost software-defined baseband (SDB). The SDB is a Consultative Committee for Space Data Systems (CCSDS)-compliant baseband that employs a personal computer for signal processing and a low-cost commercial-off-the-shelf RF front end for RF signal sampling. The SDB is customized to offer traditional telemetry, tracking, and command (TT&C) services for near-Earth missions operating in S band. The study starts by reviewing MSPA methods already studied by space agencies such as NASA and ESA before going to MSPA methods proposed in the latest CCSDS blue book on RF and modulation systems. The feasibility of operating the CCSDS-proposed multiple uplink carrier MSPA method using the SDB is assessed after evaluating in-band interference and out-of-band emissions from the uplink signals radiated by the low-cost RF front end employed by the SDB. Furthermore, we present a case study where the SDB is used in MSPA operations on a typical solid-state amplifier (SSPA) amplifier typically used in near-Earth missions. 

This paper deals with a performance comparison of the control algorithm for a variable-thrust solid-propellant rocket motor (VTSRM). To do this, we develop a simulation model of a VTSRM considering characteristic changes in the combustor and design control systems for pressure and thrust. We use three types of control algorithms for the pressure control: classical PID control, feedback linearization control, and fuzzy PID control, and two control algorithms for thrust control: classical PID control and fuzzy PID control. Finally, we compare the performance of each control system through a numerical simulation using step responses. Through this work, we check that feedback linearization is better in pressure control, and fuzzy PID control is more appropriate in thrust control. Especially using fuzzy PID control, we can get fast settling with a small undershoot even if the system is a nonminimum phase system.

Federated remote laboratories allow for the execution of experiments ex situ. The coordination of several laboratories can be used to perform concurrent experiments of combined space operations. However, the latency of the communications between facilities is critical to performing adequate real-time experiments. This paper presents an approach for conducting coordinated experiments between floating platforms at two remote laboratories. Two independently designed platforms, one at Luleå University of Technology and the other at La Sapienza University of Rome, were established for this purpose. A synchronization method based on the Simple Network Time Protocol was created, allowing the offset and delay between the agents to be measured.Both platforms exchange data about their measured time and pose through a UDP/IP protocol over the internet. This approach was validated with the execution of simulated operations. A first demonstrative experiment was also performed showing the possibility to realize leader/follower coordinated operations. The results of the simulations and experiments showed communication delays on the order of tens of milliseconds with no significant impact on the control performance. Consequently, the suggested protocol was proven to be adequate for conducting coordinated experiments in real time between remote laboratories.

This paper presents the research that is being conducted with the hardware-in-the-loop simulation testbed laying at the Nano Satellite Lab at the Kiruna Space Campus of Luleå University of Technology. Particularly, a brief introduction to the frictionless platform and both robot manipulators that comprise such a testbed is given. Additionally, this paper summarizes the simulations that are currently performed on the testbed. Finally, the paper also discusses some future activities that are planned to be conducted to improve the way systems and components of small satellites are currently validated.

The market of small satellites has experienced substantial growth in the last decade due to the high level of miniaturization in its components, allowing missions that only a decade ago were the exclusive domain of much larger satellites. This tendency has opened a new field for the development of low cost small/micro launchers (Light Launcher Vehicles - LLVs) to act on this speciffc market, mostly private-led initiatives, called nowadays as NewSpace Launch Systems. This new approach has the advantage to fulfill with high quality the mission requirements, once these launchers are dedicated to LEO missions of small, micro, and nano satellites. In the design of launchers, the use of optimization methods to maximize the payload mass and the propellant storability, as well as minimize the structural mass for a given mission or range of missions is mandatory. During the optimization process, a method to maximize the overall vehicle performance, in general, expressed by the payload capability, is handling a combination of different propulsion systems to form the vehicle stages in association with an optimum trajectory analysis. Sometimes, manipulation of the propulsion system characteristics is also required, which results in modiffcations in the original motor design. In this work, the focus will be put on the Solid Rocket Motor (SRM) S-50, baseline of the current analyzes. The S-50 is under development to compose the first and second stages of the Brazilian Micro-satellite Launch Vehicle (VLM), and the first stage of the Suborbital Rocket VS-50. This SRM burns 12 tons of solid propellant during 80 seconds, and uses a composite (CFRP) motor case presenting an attractive structural ratio. This paper addresses the stage optimization problem by using gradient method algorithm to calculate the second and third optimal stages for certain mission considering the reference SRM S-50 as launcher first stage. In the procedures to find the optimum solution for the proposed problem, concepts of optimal trajectory, maximum payload capability, structure optimization, and properly ight dynamics and control analysis are applied. The methodology developed here will be applied on reference missions for consistency checking, and the preliminary design achieved is checked by a trajectory optimization software.

The Auroral Acceleration Region (AAR) is a key region in understanding the Magnetosphere-Ionosphere interaction. To understand the physical, spatial and temporal features of the region, multi-point measurements are required. Distributed small-satellite missions such as constellations of multiple nano satellites (for example multi-unit CubeSats) would enable such type of measurements. The capabilities of such a mission will highly depend on the number of satellites - one reason that makes low-cost platforms like CubeSats a very promising choice. In a previous study, the state-of-the-art of miniaturized payloads for AAR measurements was analyzed and evaluated and capabilities of different multi-CubeSat configurations equipped with such payloads in addressing different open questions in AAR were discussed. In this paper the mission analysis and possible mission design, as well as necessary technology developments of such multi-CubeSat mission are identified and presented.

RAVEN, Rocketry and Aerospace Vehicle Engineering in Norrbotten, is the first rocket project at Luleå University of Technology (LTU), Sweden. The project started in early 2020, and it aims to design, build, test and launch a hybrid propulsion rocket. The initial objectives are to reach an altitude of 10 km with an accompanying payload of 10 kg. The RAVEN team is based at the Kiruna Space Campus and consists of approximately 30 graduate students. The main goal of the first RAVEN rocket is to demonstrate its technology. The team is designing the entire rocket from scratch, including the hybrid propulsion system that uses nitrous oxide as the oxidising agent and paraffin as the solid fuel component, to produce approximately 6 kN of thrust. A modular rocket structure will offer customisation, enabling simple design changes in future iterations. Thus, the layout allows individual subsystems to be re-designed without changes directly affecting others. The project is meant to create a foundation for future student and research rocket projects at LTU. The follow-up developments could further iterate on the first rocket design, increasing hands-on education and providing more collaboration opportunities between the university and the aerospace industry – and eventually provide an in-house platform for research and education at the university or even for commercial start-ups. The infrastructure and resources, such as the proximity to the Esrange Space Center launch facility and support from the university and industry, make Northern Sweden an ideal location to establish a continuous rocket program.

The high cost of rocket technology has led to system reusability developments, particularly in the field of first-stage rockets. With the motivation of decreasing production costs, successful vertical rocket landing attempts by SpaceX and Blue Origin have led the path for autonomous recovery and reusability of rocket engines. After the success of the space companies, recovering and relaunching reusable first stages have demonstrated the possibility of building reliable and low-cost reusable first stages. In the landing process, an optimized trajectory and robust control algorithm of a reusable launch vehicle (RLV), which can be minimized propellant consuming, are needed because the RLV is affected by various uncertainties such as a wind gust, atmospheric condition, or aerodynamic problem. Therefore, it is necessary too btain an optimized trajectory and design a robust control algorithm to cope with the problems. The instantaneous impact point (IIP) is defined as “a touchdown point, following thrust termination of a launch vehicle calculated in the absence of atmospheric drag effects.” In launch operations, the IIP is considered a crash point after an abrupt end of the propelled flight, so the IIP should be calculated and monitored in real-time on the ground facility or the rocket to predict and prevent the potential risk in advance, which can occur around the touchdown point. In other words, the IIP can be used to obtain an optimized trajectory that can minimize the consuming propellants. Since the guidance and control for the landing process of reusable first stages is a specific research field that presents multiple goals and constraints, the fuzzy PID control algorithm is suitable for the system. The algorithm has several robustness strengths, bounded-input, bounded-output (BIBO) systems, and nonlinear systems. In the landing process, the control references generally are thrust magnitude and vector control for velocity control and reaction control system (RCS) or grid-fin for attitude control. We propose a fuzzy PID control algorithm for robust guidance and control of RLV with a liquid-propellant rocket engine (LPRE) in the landing process using an optimized trajectory obtained by the IIP algorithm. To do this, we develop the simplified mathematical model of SLV using LPRE, obtain optimized passage for landing using the IIP method, and design a fuzzy PID control system based on thrust magnitude and vector control to control velocity and grid-fin or RCS to control attitude. We demonstrate the performance through numerical simulation to confirm the control algorithm.

This article provides a proposal case for studying simulations of spacecraft formation maneuvers in a Hardware-in-the-Loop Simulation Testbed specifically devoted to nanosatellites. The design specifications of the setup to perform such simulations are given, as well as the methods and results of the preliminary characterization of the floating platform to be used. The intent is to create a corpus of tests to find the dynamic behavior in a frictionless simulation as part of a concurrent decentralized simulation between the NanoSat Lab in Luleå University of Technology and the Guidance and Navigation Lab at La Sapienza.

One of the challenges that satellites face is the interaction between control movement and vibration of flexible appendages such as solar arrays and antennas that can negatively affect the performance of the spacecraft. The aim of this work is to develop a numerical model of a solar panel structure for KNATTE (Kinesthetic Node and Autonomous Table-Top Emulator), a frictionless platform developed by the Space Systems group at LuleåUniversity of Technology, and develop a control law that reduces the flexible vibration of the solar arrays when attitude control manoeuvres are performed. A set of solar panel structures have been designed and tested, the mathematical model of the multibody system, which consists of KNATTE and two flexible solar panels, has been developed in MATLAB by applying the finite element method. A finite element analysis has been performed in MATLAB to extract the natural frequencies of the system. The model has been numerically verified using a commercial software, and experimentally verified by performing testing on the frictionless vehicle, KNATTE, equipped with the solar panel structures and a number of piezoelectric sensors. Once the model has been verified, a Linear Quadratic Gaussian (LQG) controller has been developed using the results from the finite element model in order to reduce the amplitude of the vibrations of the flexible solar panel structure. The behaviour of the system has been simulated when the spacecraft performs an attitude manoeuvre. The finite element model provides the modal behaviour of the multibody system, obtaining its natural frequencies with low relative error. The LQG controller reduces the amplitude of the vibrations of the flexible solar panel structure.