Aerotecnica Missili & Spazio最新文献

The recent success of the Ingenuity Mars helicopter developed by the jet propulsion laboratory (JPL) demonstrated the feasibility of the Martian flight. Low pressure (660 Pa) and temperature (210 K) characterize the ground-level Martian atmosphere. Since such conditions are difficult and expensive to mimic on Earth, it is necessary to have reliable simulation tools that can correctly reproduce Martian aerodynamics. In the case of unmanned aerial systems (UAS), the latter is characterized by a high subsonic Mach number at the tip of the blades and an Ultra-low Reynolds number regime ((1000< hbox {Re} < 10000)). To this purpose, the laminar solver embedded in the commercial CFD code STAR CCM+ was validated by reproducing experiments carried out in the Martian Wind Tunnel at Tohoku University using a triangular airfoil wing at Reynolds 3000 and a Mach number of 0.5. Simulations are performed at angles of attack ranging from 0 to 16 degrees showing a satisfactory agreement with experimental results for very different flow conditions.

This work aims to contribute to the innovation in the urban air mobility and delivery sector and represents a starting point for air logistics and its future scenarios. The dissertation focuses on modeling, simulation, and control of a formation of multirotor aircraft for cooperative load transportation, with particular attention to the stabilization of payload swing motion. Starting from the mathematical model of two identical multirotors, formation-flight-keeping and collision-avoidance algorithms are implemented to ensure the safety of the vehicles within the formation and that of the payload. Then, a mathematical model for the suspended load is implemented, as well as an active controller for its stabilization. The focus of this section is thus represented by the analysis of payload oscillatory motion, whose kinetic energy decay is investigated. Several test cases are presented to establish the most effective and safe strategy in light of future aerospace applications.

Fiber-reinforced composites (FRC) have gained widespread recognition in the aerospace, automotive, and energy industries due to their exceptional strength to-weight ratio. However, comprehending their performance within varying environmental contexts poses a multifaceted challenge. Specifically, the influence of humidity, temperature fluctuations, and freeze–thaw cycles on the structural integrity of FRC components requires careful examination. This research work seeks to provide insights into the effects of humidity, temperature, and freeze thaw cycles on FRC inter-laminar regions and the critical matrix/fiber interface. The experimental methodology employed includes a comprehensive array of techniques, such as thermal analysis, X-Ray tomography, and ILSS mechanical testing. Through these methods, an effort is made to discern the material’s response to the environmental variables. Carbon-reinforced composites exhibited a shear strength reduction of 16.9% at 80 °C, and glass-reinforced composites displayed a reduction of 18.4%. Further increasing the temperature to 125 °C resulted in a reduction of 32.5% for carbon-reinforced composites and 38.8% for glass-reinforced composites. In hot-wet conditions, which combine humidity saturation and a testing temperature of 80 °C, the shear strength reductions were the most pronounced, with a reduction of 48.7% for carbon-reinforced composites and 60.2% for glass-reinforced composites. Moreover, freeze–thaw cycle has been performed. The findings of this research endeavor hold profound implications for both the design and maintenance of FRC components. As FRCs continue to gain prominence in critical applications, an enhanced understanding of their behavior in diverse environmental conditions becomes increasingly imperative.

In situ resource utilization (ISRU) will be the key for the success of many future space missions which are especially time or cost demanding. This is particularly true for the next Moon settlement mankind has decided to establish, and even further, for Mars colonization. Focusing on Moon operations, this paper presents a study carried out to assess the benefits of on-orbit servicing (OOS) exploiting lunar resources for the resupply of the Herschel Space Observatory (HSO). Herschel ended its operations in 2013 as a consequence of the depletion of its coolant, therefore an experimental mission has been envisaged to refill it. An adapted cargo spacecraft (s/c), employed in lunar gateway operations, will be supposed to depart from the Moon and reach Herschel for the resupply. Trajectory design and optimization for rendezvous, as well as attitude control and a methodology to simplify the dynamics equations through linearization, are the topics addressed in this study in order to obtain some preliminary data on the feasibility of these kind of missions. The results found list different trajectories that could be taken on for the mission, their (Delta)V cost and time of flight (ToF) and show the advantage of relying on simplified dynamics for the calculations. Moreover, it is presented the methodology used to approach the observatory during rendezvous (RdV) while ensuring minimum thrusting errors of the cargo spacecraft and a continuous visibility of the space observatory. The conclusion displays some ideas on how the next studies for the mission could be carried on.

The endeavor of the pure-pursuit guidance is aligning the missile velocity vector with the line-of-sight to the target. The classical pure-pursuit guidance is not the preferred choice for a guidance law as it does not perform well against moving targets. Albeit this, its appealing feature and the main advantage are that it needs measurement of angle only for implementation, thus reducing the cost at the price of the performance. To this day, the implemented guidance law for classical pure-pursuit is mostly a constant proportional control law, and acceptable miss distance is achieved for stationary and very slow targets, thus the use for engagement of moving targets is limited. This paper's objective is to use the optimal control theory in the design of the guidance law for pure-pursuit guidance and to assess the performance against non-stationary/moving targets and deterministic disturbances. The main conclusion is that applying the optimal control theory to design optimal guidance laws for pure pursuit improves performance and reduces the miss distance below a meter for a moderate target’s velocity. The Optimal Pure-Pursuit Guidance Law for stationary targets is shown to realize the Proportional Navigation guidance law.

In aerospace industrial and commercial scenario, the reusable launch vehicles (RLV) evolution works constantly toward the lowering of payload conveyance expenses. The thermal protection system (TPS) preserves the integrity of the space vehicle surfaces exposed to huge thermal shock during the re-entry phase: its advanced design and manufacturing, aimed at both reusing and withstanding harsh space environment, result in increasing the production and maintenance charges. The present study introduces a cost-saving concept of TPS component made of carbon/carbon (C/C) tiles coated by a commercial refractory varnish reinforced with ceramic nanoparticles. Using a reliable computing method, known as inverse method, the thermophysical properties such as heat capacity and thermal conductivity of the manufactured materials are assessed in a broad range of temperatures, with the input aid of an in-house developed experimental setup. The described technique is especially suited for approaching such kind of issues, thanks to the capability of taking into account several physical variables simultaneously, with the aim of gaining a robust knowledge of materials’ thermal behavior for potential use in spacecraft TPS.

In this work, the effects of different drag laws regarding the ice crystal impingement on the fuselage of a regional aircraft are investigated. Different probes are considered on the surface of interest and simulated like simple segments normal to the fuselage. Along each of these instrumentations, the collection efficiency has been calculated by using a RANS structured solver named UZEN and an Eulerian impingement code IMP3D, both developed internally at CIRA. The solvers are parallelized and well-assessed. The computational grid has been generated with ICEM CFD. Results show the strong influence of the shape considered for the ice crystal particles and how different laws can guide different water concentrations on the probe surface.

In the next years, the space debris population is expected to progressively grow due to in-space collisions and break-up events; in addition, anti-satellite tests can further affect the debris environment by generating large clouds of fragments. The simulation of these events allows identifying the main parameters affecting fragmentation and obtaining statistically accurate populations of generated debris, both above and below detection thresholds for ground-based observatories. Such information can be employed to improve current fragmentation models and to reproduce historical events to better understand their influence on the non-detectable space debris population. In addition, numerical simulation can also be used as input to identify the most critical objects to be removed to reduce the risk of irreversible orbit pollution. In this paper, the simulation of historical in-orbit fragmentation events is discussed and the generated debris populations are presented. The presented case-studies include the COSMOS-IRIDIUM collision, the COSMOS 1408 anti-satellite test, the 2022-151B CZ-6A in-orbit break-up, and a potential collision of ENVISAT with a spent rocket stage; for these events, results are presented in terms of cumulative fragments distributions and debris orbital distributions.

The Space Rider program falls within the framework of ESA activities for the design of affordable and sustainable reusable aerospace vehicles. Among the greatest challenges for this mission is the design of the Guidance, Navigation and Control (GNC) subsystem for the re-entry phase. The final stage of the latter consists of an autonomous flight under parafoil that must guarantee a smooth and precise landing. To ensure compliance with the requirements in terms of landing accuracy and ground speed constraints at touchdown, the GNC subsystem must be able to counterbalance the effect of the wind during the flight and guarantee an upwind landing. A key role in this regard is played by what is commonly referred to as the Terminal Guidance phase i.e., the final part of the descent under parafoil where the vehicle performs the final approach to the designated landing point. The study presented in this work has been developed at the AOCS/GNC department of SENER Aeroespacial and the objective is to design a complete solution for managing the Terminal Guidance phase of a Space Rider-type case. This includes a guidance algorithm based on a direct method to generate an optimal solution for the trajectory, a path-tracking procedure, and a guidance logic that allows for a correct implementation within the whole GNC software. The optimal terminal guidance algorithm has then been implemented within the six-degrees-of-freedom simulator developed by SENER Aeroespacial demonstrating an excellent functioning for the proposed problem.