Pub Date : 2025-01-02DOI: 10.1016/j.actaastro.2024.12.029
Giulio Polato , Matteo Urbinati , Andrea Valmorbida , Giovanni Anese , Alice Brunello , Samantha Salmistraro , Sebastiano Chiodini , Giacomo Colombatti , Enrico C. Lorenzini
The E.T.PACK-F project, funded by the European Innovation Council (EIC), seeks to investigate and to advance electrodynamic tether technologies (EDT) by developing two flight modules connected through an aluminum tether to demonstrate space debris deorbiting. In this context the In Line Damper (ILD) plays an important role for the stability of the entire system during the deployment of the tether and the deorbiting phase. In this paper we want to present a study on the ILD behavior when subjected to an external load, due to the initial separation of the two modules. In particular, we tested the device at the SPARTANS facility of the University of Padova, utilizing a low-friction glass table which can accommodate the experimental setup. The collected data were correlated and studied numerically using an optimization process, which enables to tune the parameters needed for describing the ILD motion, using a 2D model. In conclusion, the results successfully demonstrate modeling of the ILD dynamics, enabling the extraction of characteristic parameters, revealing non-linear tendencies and providing more insights into the mechanical response of the device.
{"title":"Experimental test and numerical validation for evaluating the dynamics of the In-Line Damper for the E.T.PACK-F project","authors":"Giulio Polato , Matteo Urbinati , Andrea Valmorbida , Giovanni Anese , Alice Brunello , Samantha Salmistraro , Sebastiano Chiodini , Giacomo Colombatti , Enrico C. Lorenzini","doi":"10.1016/j.actaastro.2024.12.029","DOIUrl":"10.1016/j.actaastro.2024.12.029","url":null,"abstract":"<div><div>The E.T.PACK-F project, funded by the European Innovation Council (EIC), seeks to investigate and to advance electrodynamic tether technologies (EDT) by developing two flight modules connected through an aluminum tether to demonstrate space debris deorbiting. In this context the In Line Damper (ILD) plays an important role for the stability of the entire system during the deployment of the tether and the deorbiting phase. In this paper we want to present a study on the ILD behavior when subjected to an external load, due to the initial separation of the two modules. In particular, we tested the device at the SPARTANS facility of the University of Padova, utilizing a low-friction glass table which can accommodate the experimental setup. The collected data were correlated and studied numerically using an optimization process, which enables to tune the parameters needed for describing the ILD motion, using a 2D model. In conclusion, the results successfully demonstrate modeling of the ILD dynamics, enabling the extraction of characteristic parameters, revealing non-linear tendencies and providing more insights into the mechanical response of the device.</div></div>","PeriodicalId":44971,"journal":{"name":"Acta Astronautica","volume":"228 ","pages":"Pages 1056-1065"},"PeriodicalIF":3.1,"publicationDate":"2025-01-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143179562","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":2,"RegionCategory":"物理与天体物理","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"OA","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2025-01-02DOI: 10.1016/j.actaastro.2024.12.057
MohammadSaleh Hedayati, Afshin Rahimi
Data-driven methods, including Artificial Intelligence (AI) and Machine Learning (ML) techniques, have been becoming more prominent in the field of satellite Fault Diagnosis and Prognosis (FDP) owing to their exceptional pattern recognition capabilities. On the other hand, they have some glaring accompanying issues other than their data dependency that have not been explored in the literature on satellite fault diagnosis. These issues include their inability to accommodate real-time fault diagnosis requirements, failure to account for the fault diagnosis and fault-tolerant modules’ interactions, and being prone to getting overfit due to manually injected faults. Therefore, this work proposes a hybrid framework for real-time fault diagnosis of a single Reaction Wheel (RW) onboard a satellite that capitalizes on both data-driven and model-based methods’ strong suits. The proposed methodology can also be applied to other satellite sub-systems. The presented hybrid framework comprises a Morkov jump-adjusted RW model, a Markov Jump-Adjusted Particle Filter (MJAPF), and a One Dimensional (1D) sliding window Residual Network (ResNet). The Morkov jump-adjusted RW model addresses the under-represented issues of data-driven methods, the MJAPF provides a means of estimating the non-linear RW’s hidden states under non-Gaussian noise conditions while accounting for malfunction dynamics, and the 1D sliding window ResNet model ensures online diagnosis performance. Experiments showed that the hybrid framework can achieve accurate and timely results, even reaching accuracy rates as high as 99% in low-noise conditions. The proposed MJAPF algorithm proved to be a capable estimation technique. However, the proposed MJAPF and ResNet frameworks were incompatible due to the gap in their perceptions of fault dynamics but proved effective on their own merits. Future remarks for making the proposed hybrid framework more robust to noise are also discussed.
{"title":"A hybrid framework for real-time satellite fault diagnosis using Markov jump-adjusted models and 1D sliding window Residual Networks","authors":"MohammadSaleh Hedayati, Afshin Rahimi","doi":"10.1016/j.actaastro.2024.12.057","DOIUrl":"10.1016/j.actaastro.2024.12.057","url":null,"abstract":"<div><div>Data-driven methods, including Artificial Intelligence (AI) and Machine Learning (ML) techniques, have been becoming more prominent in the field of satellite Fault Diagnosis and Prognosis (FDP) owing to their exceptional pattern recognition capabilities. On the other hand, they have some glaring accompanying issues other than their data dependency that have not been explored in the literature on satellite fault diagnosis. These issues include their inability to accommodate real-time fault diagnosis requirements, failure to account for the fault diagnosis and fault-tolerant modules’ interactions, and being prone to getting overfit due to manually injected faults. Therefore, this work proposes a hybrid framework for real-time fault diagnosis of a single Reaction Wheel (RW) onboard a satellite that capitalizes on both data-driven and model-based methods’ strong suits. The proposed methodology can also be applied to other satellite sub-systems. The presented hybrid framework comprises a Morkov jump-adjusted RW model, a Markov Jump-Adjusted Particle Filter (MJAPF), and a One Dimensional (1D) sliding window Residual Network (ResNet). The Morkov jump-adjusted RW model addresses the under-represented issues of data-driven methods, the MJAPF provides a means of estimating the non-linear RW’s hidden states under non-Gaussian noise conditions while accounting for malfunction dynamics, and the 1D sliding window ResNet model ensures online diagnosis performance. Experiments showed that the hybrid framework can achieve accurate and timely results, even reaching accuracy rates as high as 99% in low-noise conditions. The proposed MJAPF algorithm proved to be a capable estimation technique. However, the proposed MJAPF and ResNet frameworks were incompatible due to the gap in their perceptions of fault dynamics but proved effective on their own merits. Future remarks for making the proposed hybrid framework more robust to noise are also discussed.</div></div>","PeriodicalId":44971,"journal":{"name":"Acta Astronautica","volume":"228 ","pages":"Pages 1066-1087"},"PeriodicalIF":3.1,"publicationDate":"2025-01-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143178237","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":2,"RegionCategory":"物理与天体物理","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"OA","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2025-01-02DOI: 10.1016/j.actaastro.2024.12.062
Liangliang Zhao , Rui Zhang , Yupei Du , Guifen Zhou , Lige Wen , Hua Zhang
As lunar exploration progressed, unmanned lunar rovers encountered harsher and more variable working environments, increasingly complex operating conditions, and greater exploration range requirements. Existing unmanned lunar mesh wheels could not meet the demands of future lunar missions, which required higher traction performance, greater reliability, and lower power consumption. In this study, inspired by the functional characteristics of ostrich toes, we designed a bionic wheel and a comparison wheel. By interchanging the wheel surfaces and lugs (grousers), four types of mesh wheels (Wheels 1–4) were developed and tested. Various loads and slip ratios were applied to investigate how the surface shape of the mesh wheel and the shape of its lugs influenced traction performance and enhanced overall traction. The test results indicated that Wheel 1 outperformed the other wheels in terms of traction performance under the specified conditions and demonstrated greater energy efficiency at lower slip ratios. This led the compacted particles beneath the bionic wheel surface to generate a lateral forward reaction force, thereby propelling the wheel. This caused the compacted particles beneath the bionic wheel surface to generate a lateral forward reaction force, thereby propelling the wheel forward. The bionic lugs operated on a similar principle; however, their contribution to traction was less significant than that of the bionic wheel surface. The innovative design of the wheel surface and lugs effectively addressed the limitations of existing lunar mesh wheel structures.
{"title":"Research in innovative mesh wheel with bionic wheel surfaces and lugs for superior lunar rover performance","authors":"Liangliang Zhao , Rui Zhang , Yupei Du , Guifen Zhou , Lige Wen , Hua Zhang","doi":"10.1016/j.actaastro.2024.12.062","DOIUrl":"10.1016/j.actaastro.2024.12.062","url":null,"abstract":"<div><div>As lunar exploration progressed, unmanned lunar rovers encountered harsher and more variable working environments, increasingly complex operating conditions, and greater exploration range requirements. Existing unmanned lunar mesh wheels could not meet the demands of future lunar missions, which required higher traction performance, greater reliability, and lower power consumption. In this study, inspired by the functional characteristics of ostrich toes, we designed a bionic wheel and a comparison wheel. By interchanging the wheel surfaces and lugs (grousers), four types of mesh wheels (Wheels 1–4) were developed and tested. Various loads and slip ratios were applied to investigate how the surface shape of the mesh wheel and the shape of its lugs influenced traction performance and enhanced overall traction. The test results indicated that Wheel 1 outperformed the other wheels in terms of traction performance under the specified conditions and demonstrated greater energy efficiency at lower slip ratios. This led the compacted particles beneath the bionic wheel surface to generate a lateral forward reaction force, thereby propelling the wheel. This caused the compacted particles beneath the bionic wheel surface to generate a lateral forward reaction force, thereby propelling the wheel forward. The bionic lugs operated on a similar principle; however, their contribution to traction was less significant than that of the bionic wheel surface. The innovative design of the wheel surface and lugs effectively addressed the limitations of existing lunar mesh wheel structures.</div></div>","PeriodicalId":44971,"journal":{"name":"Acta Astronautica","volume":"228 ","pages":"Pages 853-864"},"PeriodicalIF":3.1,"publicationDate":"2025-01-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"142925310","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":2,"RegionCategory":"物理与天体物理","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2025-01-02DOI: 10.1016/j.actaastro.2024.12.061
Xiaoniu Li, Xinjian Li, Lingfeng Tang, Zheng Yang, Lin Yang
With the diversification and multifunctionality of space missions, micro-nano satellites need to carry more payloads in deep space orbits with harsh lighting conditions. Traditional built-in batteries and fixed deployable solar wings offer limited energy, making it difficult to meet these demands. Although solar wing drive mechanisms powered by stepper motors and harmonic reducers improve power output, their large size, weight, and complexity render them unsuitable for micro-nano satellites. This paper presents a solar array drive assembly driven by an ultrasonic motor, designed for micro-nano satellites operating in the halo orbit at the Lagrange L2 point of the Earth-Moon system. This design addresses the need for miniaturization and lightweight construction while enhancing energy supply. Key components, including the yaw axis pointing mechanism, solar panel assembly, angular displacement detection, energy transmission system, and ultrasonic motor, are optimized for size and weight reduction. To identify the optimal energy-saving drive method, a dynamic model of the SADA system is established, and a novel low-power driving method for the ultrasonic motor-driven solar wing is proposed. A prototype with a volume of less than 0.5U and a mass under 0.2 kg was fabricated. Experimental results show that the output torque of the ultrasonic motor exceeds 0.0581 N m, with the number of driving signal cycles n positively correlated with the total rotation angle θ of the solar wing. When n = 1000 and the interval time Tm = 1.35 h, the mechanism achieves its lowest energy consumption cost, allowing the solar wing to operate at 12.1° per day, saving over 23.8 % of energy compared to continuous drive methods.
随着航天任务的多样化和多功能化，微纳卫星需要在光照条件恶劣的深空轨道上携带更多的有效载荷。传统的内置电池和固定可展开的太阳能机翼提供有限的能量，使其难以满足这些需求。虽然由步进电机和谐波减速器驱动的太阳能翼驱动机构提高了功率输出，但它们的大尺寸、重量和复杂性使它们不适合用于微纳卫星。针对地月系拉格朗日L2点晕轨微纳卫星，设计了一种由超声电机驱动的太阳能电池阵列驱动组件。这种设计解决了小型化和轻量化结构的需要，同时提高了能源供应。关键部件，包括偏航轴指向机构、太阳能电池板组件、角位移检测、能量传输系统和超声波电机，都进行了尺寸和重量的优化。为了确定最优的节能驱动方法，建立了SADA系统的动力学模型，提出了一种新型的超声电机驱动太阳能翼的低功耗驱动方法。制作了体积小于0.5U、质量小于0.2 kg的原型机。实验结果表明，超声电机的输出转矩超过0.0581 N m，驱动信号周期数N与太阳能翼总转角θ呈正相关。当n = 1000，间隔时间Tm = 1.35 h时，该机构达到了最低的能耗成本，允许太阳能翼以每天12.1°的速度运行，与连续驱动方法相比节省了23.8%以上的能量。
{"title":"Development of a compact solar array drive assembly based on ultrasonic motor for deep space micro-nano satellites","authors":"Xiaoniu Li, Xinjian Li, Lingfeng Tang, Zheng Yang, Lin Yang","doi":"10.1016/j.actaastro.2024.12.061","DOIUrl":"10.1016/j.actaastro.2024.12.061","url":null,"abstract":"<div><div>With the diversification and multifunctionality of space missions, micro-nano satellites need to carry more payloads in deep space orbits with harsh lighting conditions. Traditional built-in batteries and fixed deployable solar wings offer limited energy, making it difficult to meet these demands. Although solar wing drive mechanisms powered by stepper motors and harmonic reducers improve power output, their large size, weight, and complexity render them unsuitable for micro-nano satellites. This paper presents a solar array drive assembly driven by an ultrasonic motor, designed for micro-nano satellites operating in the halo orbit at the Lagrange L2 point of the Earth-Moon system. This design addresses the need for miniaturization and lightweight construction while enhancing energy supply. Key components, including the yaw axis pointing mechanism, solar panel assembly, angular displacement detection, energy transmission system, and ultrasonic motor, are optimized for size and weight reduction. To identify the optimal energy-saving drive method, a dynamic model of the SADA system is established, and a novel low-power driving method for the ultrasonic motor-driven solar wing is proposed. A prototype with a volume of less than 0.5U and a mass under 0.2 kg was fabricated. Experimental results show that the output torque of the ultrasonic motor exceeds 0.0581 N m, with the number of driving signal cycles <em>n</em> positively correlated with the total rotation angle <em>θ</em> of the solar wing. When <em>n</em> = 1000 and the interval time <em>T</em><sub><em>m</em></sub> = 1.35 h, the mechanism achieves its lowest energy consumption cost, allowing the solar wing to operate at 12.1° per day, saving over 23.8 % of energy compared to continuous drive methods.</div></div>","PeriodicalId":44971,"journal":{"name":"Acta Astronautica","volume":"228 ","pages":"Pages 865-874"},"PeriodicalIF":3.1,"publicationDate":"2025-01-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"142925311","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":2,"RegionCategory":"物理与天体物理","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2025-01-01DOI: 10.1016/j.actaastro.2024.12.035
Samuel T. Hart, E. Glenn Lightsey, Álvaro Romero-Calvo
CubeSat propulsion imposes unique propellant management issues. The requisite form factor often necessitates conformal tank geometries and high-density two-phase propellants. Fluid management in these saturated propellant systems cannot generally be accomplished using conventional capillary devices, and current state-of-the-art alternatives are comparatively large. Recently, approaches based on thermal phase change have been proposed. In these phase change propellant management devices (PMDs), the propellant is vaporized in one portion of the tank through the application of heat and condensed in cooler portions due to increased pressure. Experimental results presented here show that ullage bubbles can be repositioned in a thermally insulative nylon tank using less than 5 W of input power. Test data and models indicate that the thermal conductivity of the tank has a significant effect on the efficiency of a phase change PMD when the heat source makes direct contact with the wall. The position of the heater is also shown to have a major impact on performance. Appropriate heater positioning could allow phase change PMDs to be used in tanks constructed of any material in microgravity.
{"title":"Characterization of a Phase Change Propellant Management Device","authors":"Samuel T. Hart, E. Glenn Lightsey, Álvaro Romero-Calvo","doi":"10.1016/j.actaastro.2024.12.035","DOIUrl":"10.1016/j.actaastro.2024.12.035","url":null,"abstract":"<div><div>CubeSat propulsion imposes unique propellant management issues. The requisite form factor often necessitates conformal tank geometries and high-density two-phase propellants. Fluid management in these saturated propellant systems cannot generally be accomplished using conventional capillary devices, and current state-of-the-art alternatives are comparatively large. Recently, approaches based on thermal phase change have been proposed. In these phase change propellant management devices (PMDs), the propellant is vaporized in one portion of the tank through the application of heat and condensed in cooler portions due to increased pressure. Experimental results presented here show that ullage bubbles can be repositioned in a thermally insulative nylon tank using less than 5 W of input power. Test data and models indicate that the thermal conductivity of the tank has a significant effect on the efficiency of a phase change PMD when the heat source makes direct contact with the wall. The position of the heater is also shown to have a major impact on performance. Appropriate heater positioning could allow phase change PMDs to be used in tanks constructed of any material in microgravity.</div></div>","PeriodicalId":44971,"journal":{"name":"Acta Astronautica","volume":"228 ","pages":"Pages 842-852"},"PeriodicalIF":3.1,"publicationDate":"2025-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"142925286","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":2,"RegionCategory":"物理与天体物理","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2025-01-01DOI: 10.1016/j.actaastro.2024.11.030
Shuai Zhou, Jiangong Zhao, Zilong Zhao, Hongjun Liu, Wen Ao
Nanofluid fuel has garnered significant attention due to its potential to enhance combustion characteristics, energy density, and ignition properties. The study comprehensively examined the effects of aluminum nanoparticles with diverse sizes (50 nm, 100 nm, 200 nm, 500 nm, 1 μm) and concentrations (2.5 wt%, 5.0 wt%, 7.5 wt%) on the ignition and combustion characteristics of nanofluid fuel droplets, utilizing a mechanically mixed aluminum-based nanofluid fuel solution that incorporated kerosene, aluminum particles, and the surfactant oleic acid. The combustion process of the nanofluid fuel droplets encompasses phases of ignition, steady combustion, micro-explosion, and agglomerate reaction. The surface temperature of the nanofluid fuel droplets consistently exceeded that of a pure kerosene droplet, with temperature elevations correlating positively with particle concentration but not with the particle size. The surface temperature of nanofluid fuel droplets containing 7.5 wt% aluminum particles is approximately 205°C. The incorporation of oleic acid into pure kerosene prolongs the ignition delay from 0.317 s to 0.333 s. The combustion rate of the nanofluid fuel droplets escalates upon the addition of aluminum particles, with the rate escalating in tandem with the diameter and concentration of the aluminum particles. Nanofluid fuel droplets containing 5.0 wt% aluminum and 5.0 wt% oleic acid particles exhibit a combustion rate akin to that of pure kerosene droplets, with rates of 0.596 and 0.604 mm2 s−1, respectively. Concurrently, the ignition delay for nanofluid fuel droplets is longer than that of pure kerosene, yet it exhibits insensitivity to particle size. The ignition delay for nanofluid fuel droplets with the addition of 7.5 wt% aluminum particles is approximately 1.5 times that of kerosene. Nanofluid fuel droplets devoid of oleic acid yield divergent results due to particle agglomeration effects. Subsequently, as particle size increased, the surface of combustion residue develops more pronounced bulges, becoming more prone to rupture. Ultimately, a kinetic prediction model is proposed, accounting for the inhomogeneous properties within the droplet. The root mean squared errors for ignition delay time, combustion rate, and steady surface temperature are all below 8 %, indicating a strong correlation between model predictions and experimental data. This research could help accelerate the adoption of aluminum-based nanofluid fuel.
{"title":"Effect of aluminum nanoparticles size and concentration on the combustion characteristics of nanofluid fuel: Experiments and modeling","authors":"Shuai Zhou, Jiangong Zhao, Zilong Zhao, Hongjun Liu, Wen Ao","doi":"10.1016/j.actaastro.2024.11.030","DOIUrl":"10.1016/j.actaastro.2024.11.030","url":null,"abstract":"<div><div>Nanofluid fuel has garnered significant attention due to its potential to enhance combustion characteristics, energy density, and ignition properties. The study comprehensively examined the effects of aluminum nanoparticles with diverse sizes (50 nm, 100 nm, 200 nm, 500 nm, 1 μm) and concentrations (2.5 wt%, 5.0 wt%, 7.5 wt%) on the ignition and combustion characteristics of nanofluid fuel droplets, utilizing a mechanically mixed aluminum-based nanofluid fuel solution that incorporated kerosene, aluminum particles, and the surfactant oleic acid. The combustion process of the nanofluid fuel droplets encompasses phases of ignition, steady combustion, micro-explosion, and agglomerate reaction. The surface temperature of the nanofluid fuel droplets consistently exceeded that of a pure kerosene droplet, with temperature elevations correlating positively with particle concentration but not with the particle size. The surface temperature of nanofluid fuel droplets containing 7.5 wt% aluminum particles is approximately 205°C. The incorporation of oleic acid into pure kerosene prolongs the ignition delay from 0.317 s to 0.333 s. The combustion rate of the nanofluid fuel droplets escalates upon the addition of aluminum particles, with the rate escalating in tandem with the diameter and concentration of the aluminum particles. Nanofluid fuel droplets containing 5.0 wt% aluminum and 5.0 wt% oleic acid particles exhibit a combustion rate akin to that of pure kerosene droplets, with rates of 0.596 and 0.604 mm<sup>2</sup> s<sup>−1</sup>, respectively. Concurrently, the ignition delay for nanofluid fuel droplets is longer than that of pure kerosene, yet it exhibits insensitivity to particle size. The ignition delay for nanofluid fuel droplets with the addition of 7.5 wt% aluminum particles is approximately 1.5 times that of kerosene. Nanofluid fuel droplets devoid of oleic acid yield divergent results due to particle agglomeration effects. Subsequently, as particle size increased, the surface of combustion residue develops more pronounced bulges, becoming more prone to rupture. Ultimately, a kinetic prediction model is proposed, accounting for the inhomogeneous properties within the droplet. The root mean squared errors for ignition delay time, combustion rate, and steady surface temperature are all below 8 %, indicating a strong correlation between model predictions and experimental data. This research could help accelerate the adoption of aluminum-based nanofluid fuel.</div></div>","PeriodicalId":44971,"journal":{"name":"Acta Astronautica","volume":"226 ","pages":"Pages 149-161"},"PeriodicalIF":3.1,"publicationDate":"2025-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143133567","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":2,"RegionCategory":"物理与天体物理","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2025-01-01DOI: 10.1016/j.actaastro.2024.12.055
Rachit Bhatia , Gerardo Josue Rivera Santos , Jacob D. Griesbach , Piyush M. Mehta
Space safety and sustainability has recently received formalized recognition in the light of proliferation by large satellite constellations operated by the commercial sector. Enhanced space operations – detection, characterization, and tracking – are critical for safety and sustainability. A large portion of the lethal (non-)trackable debris reside in low Earth orbit (LEO) while the new commercial constellations reside dominantly in the lower LEO (LLEO) regime with significant plans for exploiting very LEO (VLEO) for future missions. With the new LEO population biased toward LLEO and VLEO, operations have become significantly more sensitive to atmospheric drag, modeling of which remains a primary challenge. Under support from the Intelligence Advanced Research Projects Activity (IARPA) Space Debris Identification and Tracking (SINTRA) program and the Office of Space Commerce (OSC), we are developing the next-generation drag modeling framework that accurately characterizes atmospheric density uncertainty due to space weather in a physics- and data-driven approach. This paper introduces one of the elements of the new framework we call stochastic Unscented Transform (SUT), a mathematical formulation designed to capture the joint statistics of probabilistic atmospheric density models and their probabilistic drivers or inputs. We present the mathematical derivation of SUT and its validation with simple numerical examples of linear and non-linear systems and then apply it to the case of drag modeling by incorporating the effects of uncertainty in the solar driver and density models in real-time orbit propagation. Enabled by the generalized nature of the SUT formulation, we also apply it to uncertainty and orbit prediction. This work moves us in the direction of realistic covariance for operations and eventually space safety and sustainability.
{"title":"A novel stochastic unscented transform for probabilistic drag modeling and conjunction assessment","authors":"Rachit Bhatia , Gerardo Josue Rivera Santos , Jacob D. Griesbach , Piyush M. Mehta","doi":"10.1016/j.actaastro.2024.12.055","DOIUrl":"10.1016/j.actaastro.2024.12.055","url":null,"abstract":"<div><div>Space safety and sustainability has recently received formalized recognition in the light of proliferation by large satellite constellations operated by the commercial sector. Enhanced space operations – detection, characterization, and tracking – are critical for safety and sustainability. A large portion of the lethal (non-)trackable debris reside in low Earth orbit (LEO) while the new commercial constellations reside dominantly in the lower LEO (LLEO) regime with significant plans for exploiting very LEO (VLEO) for future missions. With the new LEO population biased toward LLEO and VLEO, operations have become significantly more sensitive to atmospheric drag, modeling of which remains a primary challenge. Under support from the Intelligence Advanced Research Projects Activity (IARPA) Space Debris Identification and Tracking (SINTRA) program and the Office of Space Commerce (OSC), we are developing the next-generation drag modeling framework that accurately characterizes atmospheric density uncertainty due to space weather in a physics- and data-driven approach. This paper introduces one of the elements of the new framework we call stochastic Unscented Transform (SUT), a mathematical formulation designed to capture the joint statistics of probabilistic atmospheric density models and their probabilistic drivers or inputs. We present the mathematical derivation of SUT and its validation with simple numerical examples of linear and non-linear systems and then apply it to the case of drag modeling by incorporating the effects of uncertainty in the solar driver and density models in real-time orbit propagation. Enabled by the generalized nature of the SUT formulation, we also apply it to uncertainty and orbit prediction. This work moves us in the direction of realistic covariance for operations and eventually space safety and sustainability.</div></div>","PeriodicalId":44971,"journal":{"name":"Acta Astronautica","volume":"228 ","pages":"Pages 985-995"},"PeriodicalIF":3.1,"publicationDate":"2025-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143178228","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":2,"RegionCategory":"物理与天体物理","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"OA","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Planetary atmospheric detection is an important way to recognize the physical and chemical properties of planets that inform about their formation and evolution, and atmospheric in-situ sampling is an ideal way to obtain high-resolution information. In this paper, a scheme for in-situ sampling of planetary atmosphere based on multi-pass aeroassisted maneuvers is given, and the corresponding design method for multiple traversal trajectories through the atmosphere is proposed. The aeroassisted maneuvering scheme achieves target-area sampling by crossing the atmosphere circularly, and is able to flexibly adjust the sampling altitude, thus having the advantage of three-dimensional and wide-area sampling. The trajectory design method involves algorithms to determine key design parameters separately. Specifically, the minimum entry periapsis altitude is determined by building its mapping relationship with path constraints to satisfy the minimum flight altitude constraint. Besides, the pass number of atmospheric flights is calculated by giving the upper bound of the energy attenuation and mission-time constraints. Then, a rapid inclination correction method via bank angle reversal is given to satisfy the inclination constraint of the maneuver. In numerical simulations, three Martian atmospheric detection scenarios, designated as high-latitude region with superficial ice water, magnetic anomalies region, and the polar region enriched with atmospheric transport properties, are established, with corresponding maneuvering sampling trajectories and characteristic parameter distributions provided. This paper introduces for the first time the use of multi-pass aeroassisted maneuvers for in-situ atmospheric sampling. Simulation results demonstrate the effectiveness and general applicability of the proposed method.
{"title":"Scheme and trajectory design of in-situ atmospheric sampling with multi-pass aeroassisted maneuvers","authors":"Xiangdong Feng, Dong Qiao, Hongwei Han, Ruifeng Lv","doi":"10.1016/j.actaastro.2024.11.038","DOIUrl":"10.1016/j.actaastro.2024.11.038","url":null,"abstract":"<div><div>Planetary atmospheric detection is an important way to recognize the physical and chemical properties of planets that inform about their formation and evolution, and atmospheric in-situ sampling is an ideal way to obtain high-resolution information. In this paper, a scheme for in-situ sampling of planetary atmosphere based on multi-pass aeroassisted maneuvers is given, and the corresponding design method for multiple traversal trajectories through the atmosphere is proposed. The aeroassisted maneuvering scheme achieves target-area sampling by crossing the atmosphere circularly, and is able to flexibly adjust the sampling altitude, thus having the advantage of three-dimensional and wide-area sampling. The trajectory design method involves algorithms to determine key design parameters separately. Specifically, the minimum entry periapsis altitude is determined by building its mapping relationship with path constraints to satisfy the minimum flight altitude constraint. Besides, the pass number of atmospheric flights is calculated by giving the upper bound of the energy attenuation and mission-time constraints. Then, a rapid inclination correction method via bank angle reversal is given to satisfy the inclination constraint of the maneuver. In numerical simulations, three Martian atmospheric detection scenarios, designated as high-latitude region with superficial ice water, magnetic anomalies region, and the polar region enriched with atmospheric transport properties, are established, with corresponding maneuvering sampling trajectories and characteristic parameter distributions provided. This paper introduces for the first time the use of multi-pass aeroassisted maneuvers for in-situ atmospheric sampling. Simulation results demonstrate the effectiveness and general applicability of the proposed method.</div></div>","PeriodicalId":44971,"journal":{"name":"Acta Astronautica","volume":"226 ","pages":"Pages 135-148"},"PeriodicalIF":3.1,"publicationDate":"2025-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143133566","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":2,"RegionCategory":"物理与天体物理","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2025-01-01DOI: 10.1016/j.actaastro.2024.11.036
Longfei Du, Yajun Luo, Linwei Ji, Fengfan Yang, Yahong Zhang, Shilin Xie
Large spaceborne optical load necessitates an exceptionally serene on-orbit environment to achieve high precision but is subject to micro-vibration and attitude adjustments. At present, the Stewart platform is often adopted and implemented in micro-vibration isolation. However, based on the common model design, the actual decoupling performance of the Stewart platform is still problematic and significantly affects the image quality of the optical load when performing attitude adjustments. Therefore, this work develops a comprehensive parametric model of the Stewart platform to analyze its coupling property. Firstly, a Stewart platform is introduced and its comprehensive dynamic model is established while considering more influential parameters such as the positioning of flexible joints and the detailed equivalent modeling of the legs. Secondly, by comparing the common model and the comprehensive model, the latter significantly reduces errors in the stiffness of the system and modal frequencies. Therefore, relevant parameters are redesigned to meet decoupling requirements and target modal frequency based on the comprehensive model. Using the finite element model and the numerical simulation model of the platform, it is demonstrated that parameters designed based on the comprehensive model effectively decrease the multi-degree-of-freedom coupling degree, and ensure effective micro-vibration control during attitude adjustments.
{"title":"Comprehensive parametric model and decoupling design of a Stewart platform for a large spaceborne optical load","authors":"Longfei Du, Yajun Luo, Linwei Ji, Fengfan Yang, Yahong Zhang, Shilin Xie","doi":"10.1016/j.actaastro.2024.11.036","DOIUrl":"10.1016/j.actaastro.2024.11.036","url":null,"abstract":"<div><div>Large spaceborne optical load necessitates an exceptionally serene on-orbit environment to achieve high precision but is subject to micro-vibration and attitude adjustments. At present, the Stewart platform is often adopted and implemented in micro-vibration isolation. However, based on the common model design, the actual decoupling performance of the Stewart platform is still problematic and significantly affects the image quality of the optical load when performing attitude adjustments. Therefore, this work develops a comprehensive parametric model of the Stewart platform to analyze its coupling property. Firstly, a Stewart platform is introduced and its comprehensive dynamic model is established while considering more influential parameters such as the positioning of flexible joints and the detailed equivalent modeling of the legs. Secondly, by comparing the common model and the comprehensive model, the latter significantly reduces errors in the stiffness of the system and modal frequencies. Therefore, relevant parameters are redesigned to meet decoupling requirements and target modal frequency based on the comprehensive model. Using the finite element model and the numerical simulation model of the platform, it is demonstrated that parameters designed based on the comprehensive model effectively decrease the multi-degree-of-freedom coupling degree, and ensure effective micro-vibration control during attitude adjustments.</div></div>","PeriodicalId":44971,"journal":{"name":"Acta Astronautica","volume":"226 ","pages":"Pages 119-134"},"PeriodicalIF":3.1,"publicationDate":"2025-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143133568","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":2,"RegionCategory":"物理与天体物理","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-12-31DOI: 10.1016/j.actaastro.2024.12.042
Renhe Shi , Xinhui Tai , Teng Long , Nianhui Ye , Fuxiang Dong
With the increasing demands for high-speed data transmission and global communication, GEO telecommunication satellites with large-size antenna payload have attracted much attention nowadays. To address the challenge of effective system design, this paper proposes a metamodel assisted multidisciplinary design optimization (MDO) framework for a Large-size Payload Telecommunication Satellite (LSP-TS). In the framework, the LSP-TS MDO problem is formulated to minimize the total system mass subject to several practical engineering constraints. Considering the interconnected relationship between the large-size payload and the satellite platform, the analysis models of satellite geometry configuration, power, attitude control, structure, GEO station-keeping, orbital transfer, and mass disciplines are established. To reduce the computational cost, an adaptive Kriging method using Pareto fitness-based sampling (AKM-PFS) is proposed as the optimizer integrated with the satellite MDO framework. In this approach, the Kriging metamodels of LSP-TS system are constructed and adaptively refined for optimization via exploring the Pareto frontier of objective and constraints, which leads the search to the feasible optimized satellite system design efficiently. After optimization, the total system mass is reduced by 318.53 kg (8.87 %) compared with the initial solution where all constraints being satisfied. Moreover, the optimization solution of the proposed AKM-PFS is further discussed to illustrate the practicality and effectiveness of the proposed method.
{"title":"Multidisciplinary design and metamodel assisted optimization for a telecommunication satellite with large-size payload","authors":"Renhe Shi , Xinhui Tai , Teng Long , Nianhui Ye , Fuxiang Dong","doi":"10.1016/j.actaastro.2024.12.042","DOIUrl":"10.1016/j.actaastro.2024.12.042","url":null,"abstract":"<div><div>With the increasing demands for high-speed data transmission and global communication, GEO telecommunication satellites with large-size antenna payload have attracted much attention nowadays. To address the challenge of effective system design, this paper proposes a metamodel assisted multidisciplinary design optimization (MDO) framework for a Large-size Payload Telecommunication Satellite (LSP-TS). In the framework, the LSP-TS MDO problem is formulated to minimize the total system mass subject to several practical engineering constraints. Considering the interconnected relationship between the large-size payload and the satellite platform, the analysis models of satellite geometry configuration, power, attitude control, structure, GEO station-keeping, orbital transfer, and mass disciplines are established. To reduce the computational cost, an adaptive Kriging method using Pareto fitness-based sampling (AKM-PFS) is proposed as the optimizer integrated with the satellite MDO framework. In this approach, the Kriging metamodels of LSP-TS system are constructed and adaptively refined for optimization via exploring the Pareto frontier of objective and constraints, which leads the search to the feasible optimized satellite system design efficiently. After optimization, the total system mass is reduced by 318.53 kg (8.87 %) compared with the initial solution where all constraints being satisfied. Moreover, the optimization solution of the proposed AKM-PFS is further discussed to illustrate the practicality and effectiveness of the proposed method.</div></div>","PeriodicalId":44971,"journal":{"name":"Acta Astronautica","volume":"228 ","pages":"Pages 900-917"},"PeriodicalIF":3.1,"publicationDate":"2024-12-31","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"142929277","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":2,"RegionCategory":"物理与天体物理","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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