Tracking Model Predictive Control for Docking Maneuvers of a CubeSat with a Big Spacecraft

The release and retrieval of a CubeSat from a big spacecraft is useful for the external inspection and monitoring of the big spacecraft. However, docking maneuvers during the retrieval are challenging since safety constraints and high performance must be achieved, considering the small dimensions and the actual small satellites technology. The trajectory control is crucial to have a soft, accurate, quick, and propellant saving docking. The present paper deals with the design of a tracking model predictive controller (TMPC) tuned to achieve the stringent docking requirements for the retrieval of a CubeSat within the cargo bay of a large cooperative vehicle. The performance of the TMPC is verified using a complex model that includes non-linearities, uncertainties of the CubeSat parameters, and environmental disturbances. Moreover, 300 Monte Carlo runs demonstrate the robustness of the TMPC solution

Validation of a test platform to qualify miniaturized electric propulsion systems

Miniaturized electric propulsion systems are one of the main technologies that could increase interest in CubeSats for future space missions. However, the integration of miniaturized propulsion systems in modern CubeSat platforms presents some issues due to the mutual interactions in terms of power consumption, chemical contamination and generated thermal and electro-magnetic environments. The present paper deals with the validation of a flexible test platform to assess the interaction of propulsion systems with CubeSat-technologies from mechanical, electrical, magnetic, and chemical perspectives. The test platform is a 6U CubeSat hosting an electric propulsion system and able to manage a variety of electric propulsion systems. The platform can regulate and distribute electric power (up to 60 W), exchange data according to several protocols (e.g. CAN bus, UART, I2C, SPI), and provide different mechanical layouts in 4U box completely dedicated to the propulsion system. Moreover, the data gathered by the onboard sensors are combined with the data from external devices and tools providing unprecedented information about the mutual behavior of a CubeSat platform and an electric propulsion system

Design and verification of Guidance, Navigation and Control systems for space applications

In the last decades, systems have strongly increased their complexity in terms of number of functions that can be performed and quantity of relationships between functions and hardware as well as interactions of elements and disciplines concurring to the definition of the system. The growing complexity remarks the importance of defining methods and tools that improve the design, verification and validation of the system process: effectiveness and costs reduction without loss of confidence in the final product are the objectives that have to be pursued. Within the System Engineering context, the modern Model and Simulation based approach seems to be a promising strategy to meet the goals, because it reduces the wasted resources with respect to the traditional methods, saving money and tedious works. Model Based System Engineering (MBSE) starts from the idea that it is possible at any moment to verify, through simulation sessions and according to the phase of the life cycle, the feasibility, the capabilities and the performances of the system. Simulation is used during the engineering process and can be classified from fully numerical (i.e. all the equipment and conditions are reproduced as virtual model) to fully integrated hardware simulation (where the system is represented by real hardware and software modules in their operational environment). Within this range of simulations, a few important stages can be defined: algorithm in the loop (AIL), software in the loop (SIL), controller in the loop (CIL), hardware in the loop (HIL), and hybrid configurations among those. The research activity, in which this thesis is inserted, aims at defining and validating an iterative methodology (based on Model and Simulation approach) in support of engineering teams and devoted to improve the effectiveness of the design and verification of a space system with particular interest in Guidance Navigation and Control (GNC) subsystem. The choice of focusing on GNC derives from the common interest and background of the groups involved in this research program (ASSET at Politecnico di Torino and AvioSpace, an EADS company). Moreover, GNC system is sufficiently complex (demanding both specialist knowledge and system engineer skills) and vital for whatever spacecraft and, last but not least the verification of its behavior is difficult on ground because strong limitations on dynamics and environment reproduction arise. Considering that the verification should be performed along the entire product life cycle, a tool and a facility, a simulator, independent from the complexity level of the test and the stage of the project, is needed. This thesis deals with the design of the simulator, called StarSim, which is the real heart of the proposed methodology. It has been entirely designed and developed from the requirements definition to the software implementation and hardware construction, up to the assembly, integration and verification of the first simulator release. In addition, the development of this technology met the modern standards on software development and project management. StarSim is a unique and self-contained platform: this feature allows to mitigate the risk of incompatibility, misunderstandings and loss of information that may arise using different software, simulation tools and facilities along the various phases. Modularity, flexibility, speed, connectivity, real time operation, fidelity with real world, ease of data management, effectiveness and congruence of the outputs with respect to the inputs are the sought-after features in the StarSim design. For every iteration of the methodology, StarSim guarantees the possibility to verify the behavior of the system under test thanks to the permanent availability of virtual models, that substitute all those elements not yet available and all the non-reproducible dynamics and environmental conditions. StarSim provides a furnished and user friendly database of models and interfaces that cover different levels of detail and fidelity, and supports the updating of the database allowing the user to create custom models (following few, simple rules). Progressively, pieces of the on board software and hardware can be introduced without stopping the process of design and verification, avoiding delays and loss of resources. StarSim has been used for the first time with the CubeSats belonging to the e-st@r program. It is an educational project carried out by students and researchers of the “CubeSat Team Polito” in which StarSim has been mainly used for the payload development, an Active Attitude Determination and Control System, but StarSim’s capabilities have also been updated to evaluate functionalities, operations and performances of the entire satellite. AIL, SIL, CIL, HIL simulations have been performed along all the phases of the project, successfully verifying a great number of functional and operational requirements. In particular, attitude determination algorithms, control laws, modes of operation have been selected and verified; software has been developed step by step and the bugs-free executable files have been loaded on the micro-controller. All the interfaces and protocols as well as data and commands handling have been verified. Actuators, logic and electrical circuits have been designed, built and tested and sensors calibration has been performed. Problems such as real time and synchronization have been solved and a complete hardware in the loop simulation test campaign both for A-ADCS standalone and for the entire satellite has been performed, verifying the satisfaction of a great number of CubeSat functional and operational requirements. The case study represents the first validation of the methodology with the first release of StarSim. It has been proven that the methodology is effective in demonstrating that improving the design and verification activities is a key point to increase the confidence level in the success of a space mission

A Test Platform to Assess the Impact of Miniaturized Propulsion Systems

Miniaturized propulsion systems can enable many future CubeSats missions. The advancement of the Technology Readiness Level of this technology passes through the integration in a CubeSat platform and the assessment of the impact and the interactions of the propulsion systems on the actual CubeSat technology and vice versa. The request of power, the thermal environmental, and the electromagnetic emissions generated inside the platform require careful analyses. This paper presents the upgraded design and the validation of a CubeSat test platform (CTP) that can interface a wide range of new miniaturized propulsion systems and gather unprecedented information for these analyses, which can be fused with the commonly used ground support equipment. The CTP features are reported, and the main achievements of the tests are shown, demonstrating the effective capabilities of the platform and how it allows for the investigation of the mutual interactions at system level between propulsion systems and the CubeSat technology

Orthogonal-Array based Design Methodology for Complex, Coupled Space Systems

The process of designing a complex system, formed by many elements and sub-elements interacting between each other, is usually completed at a system level and in the preliminary phases in two major steps: design-space exploration and optimization. In a classical approach, especially in a company environment, the two steps are usually performed together, by experts of the field inferring on major phenomena, making assumptions and doing some trial-and-error runs on the available mathematical models. To support designers and decision makers during the design phases of this kind of complex systems, and to enable early discovery of emergent behaviours arising from interactions between the various elements being designed, the authors implemented a parametric methodology for the design-space exploration and optimization. The parametric technique is based on the utilization of a particular type of matrix design of experiments, the orthogonal arrays. Through successive design iterations with orthogonal arrays, the optimal solution is reached with a reduced effort if compared to more computationally-intense techniques, providing sensitivity and robustness information. The paper describes the design methodology in detail providing an application example that is the design of a human mission to support a lunar base

ENTRY, DESCENT AND IMPACT SYSTEM DESIGN AND ANALYSIS OF A SMALL PLATFORM IN MARTIAN ENVIRONMENT

Thanks to the latest Mars missions, planetary exploration has made enormous strides over the past ten years increasing the interest of the scientific community and beyond. These missions must fulfil many complex operations which are of paramount importance to mission success. Among these, a special mention goes to the Entry, Descent and Landing (EDL) functions which require a dedicated system to overcome all the obstacles of these critical phases. The goal of this study is to describe in detail the design methodology for EDL system during the preliminary phase of the design. The design is supported by a simulation tool integrating the entry trajectory algorithm. The trajectory data computed are used to size the EDL system and strategy in order to have a low aerodynamic acceleration, low dynamic pressure and low convective heat flux incoming to the spacecraft. The reference mission has the goal to find bioevidence and biohazards on Martian subsurface in order to prepare future manned missions. The mission is based on Space Penetrator Systems (SPS) that can descend on Mars surface following a ballistic fall and penetrate the ground after the impact with the surface (around 50 and 300 cm depth). The SPS contains all the instrumentation required to sample and make the required analyses. As results, an Entry Descent and Impact (EDI) system based on inflatable structure is designed, respecting the low-cost and low-mass constraints. For this mission, a solution, like the one of Finnish Meteorological Institute in the Mars Met-Net mission, is chosen, using an inflatable Thermal Protection System (TPS) called Inflatable Braking Unit (IBU) and an additional inflatable decelerator. Consequently, there are three configurations during the EDI phases: at an altitude of 125 km, the IBU is inflated at speed 5.5 km/s; at an altitude of 16 km, the IBU is jettisoned and an Additional Inflatable Braking Unit (AIBU) is inflated; at last, at about 13 km, the SPS is ejected from AIBU and it impacts on the Martian surface. In this paper, the results obtained by the application of this design methodology are presented and, the obtained system and descent strategy satisfy the requirements of the mission

Docking Manoeuvre Control for CubeSats

Rendezvous and Docking missions of small satellites are opening new scenarios to accomplish unprecedented in-obit operations. These missions impose to win the new technical challenges that enable the possibility to successfully perform complex and safety-critical manoeuvres. The disturbance forces and torques due to the hostile space environment, the uncertainties introduced by the onboard technologies and the safety constraints and reliability requirements lead to select advanced control systems. The paper proposes a control strategy based on Model Predictive Control for trajectory control and Sliding Mode Control for attitude control of the chaser in last meters before the docking. The control performances are verified in a dedicated simulation environment in which a non-linear six Degrees of Freedom and coupled dynamics, uncertainties on sensors and actuators responses are included. A set of 300 Monte Carlo Simulation with this Non-Linear system are carried out, demonstrating the capabilities of the proposed control system to achieve the final docking point with the required accuracy