Acta Astronautica最新文献_第6页

Study on the collision characteristics between high-temperature alumina droplets and char layer 高温氧化铝液滴与炭层碰撞特性研究

IF 3.1 2区物理与天体物理 Q1 ENGINEERING, AEROSPACE

Acta Astronautica

Pub Date : 2024-10-05 DOI: 10.1016/j.actaastro.2024.10.003

The collision characteristics between high-temperature alumina droplets and the char layer in solid rocket motors are of great significance for the accuracy of slag deposition and flow-field simulations, however, the current research on the collision characteristics of the alumina droplets and char layer is still in a blank state. This study is based on the high-temperature molding method to prepare the char layer and compare the porosity with that of the char layer in solid rocket motors, indicating that the two are relatively similar in structure and can be applied to droplet impact experiments. An experimental study on the collision of alumina droplet with the char layer was conducted using a high-temperature alumina droplet impact experimental system. The experimental results show that the adhesion behavior of alumina droplets is related to the rough structure of the char layer and the high viscosity dissipation of the process of droplets impacting the char layer, and the droplets adhere during the retraction stage with violent oscillation. The rebound behavior of the droplets on the wall was characterized by “tail dragging”, “spinning” and “asymmetric rebound” phenomena due to the combination of high surface tension and the pinning effect of wall roughness. A regime map of the rebound/adhesion results of droplets impact the char layers was constructed. At the same speed, droplets with smaller particle sizes are more likely to adhere to the char layer. We established a relationship between the rebound and adhesion behavior. Based on the experimental results, the relationship for the maximum spreading factor of the droplets was established, providing a theoretical basis for the in-depth understanding and study of the droplet collision process in solid rocket motors.

固体火箭发动机中高温氧化铝液滴与炭层的碰撞特性对熔渣沉积和流场模拟的准确性具有重要意义，但目前对氧化铝液滴与炭层碰撞特性的研究仍处于空白状态。本研究基于高温成型法制备炭层，并将其孔隙率与固体火箭发动机中的炭层孔隙率进行对比，表明二者结构较为相似，可应用于液滴碰撞实验。利用高温氧化铝液滴冲击实验系统对氧化铝液滴与炭层的碰撞进行了实验研究。实验结果表明，氧化铝液滴的粘附行为与炭层的粗糙结构和液滴撞击炭层过程中的高粘度耗散有关，液滴在回缩阶段以剧烈振荡的方式粘附。由于高表面张力和壁面粗糙度的针刺效应，液滴在壁面上的反弹行为表现为 "拖尾"、"旋转 "和 "非对称反弹 "现象。绘制了液滴撞击炭层的反弹/粘附结果体系图。在相同速度下，粒径较小的液滴更容易粘附在炭层上。我们建立了反弹和粘附行为之间的关系。根据实验结果，建立了液滴最大扩散因子的关系，为深入理解和研究固体火箭发动机中的液滴碰撞过程提供了理论依据。

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引用次数: 0

Design and tests of filtering actions for an AI-based RSOs detection and tracking algorithm 设计和测试基于人工智能的 RSO 检测和跟踪算法的过滤操作

IF 3.1 2区物理与天体物理 Q1 ENGINEERING, AEROSPACE

Acta Astronautica

Pub Date : 2024-10-03 DOI: 10.1016/j.actaastro.2024.10.001

Space debris represent a huge threat for the actual and future space traffic. The continuous monitoring of the space environment around the Earth and the build-up, maintenance of Resident Space Objects (RSOs) catalogue are essential to enhance the space segment safety. In the last few years, trend towards Space Based Space Surveillance (SBSS) missions was relevant. In particular, the monitoring of space debris through on-board Star Trackers (STs) coupled with Artificial Intelligence (AI) based algorithms is being studied, modeling RSOs as clear streaks on a dark background. This work presents the advances in the development of an AI-based algorithm for streaks-like RSOs Detection & Tracking for on-board optical sensors applications like STs. The initial design of the RSOs monitoring algorithm (Mastrofini et al., 2022) is extended and tested for what concerns the tracking part with new filtering actions. The goal is the improvement of the tracking outputs quality and reliability when multiple RSOs passages occur, with different configurations. Tests both on real and simulated images from STs are shown and discussed.

空间碎片对现实和未来的空间交通构成巨大威胁。持续监测地球周围的空间环境以及建立和维护驻留空间物体（RSOs）目录对于提高空间段的安全性至关重要。在过去几年中，天基空间监视（SBSS）任务的发展趋势与此相关。特别是正在研究通过星载跟踪器（ST）和基于人工智能（AI）的算法对空间碎片进行监测，将驻留空间物体建模为暗色背景上的清晰条纹。这项工作介绍了在开发基于人工智能的条纹状 RSOs 检测& 跟踪算法方面取得的进展，该算法适用于 STs 等星载光学传感器应用。RSOs 监测算法的初始设计（Mastrofini 等人，2022 年）通过新的过滤操作对跟踪部分进行了扩展和测试。目标是在出现多个 RSO 时，通过不同的配置提高跟踪输出的质量和可靠性。对 STs 真实图像和模拟图像的测试结果进行了展示和讨论。

引用次数: 0

Initial investigation of catalyst pack for 98 %+ hydrogen peroxide satellite monopropellant thruster 用于 98%+ 过氧化氢卫星单推进剂推进器的催化剂组的初步研究

IF 3.1 2区物理与天体物理 Q1 ENGINEERING, AEROSPACE

Acta Astronautica

Pub Date : 2024-10-02 DOI: 10.1016/j.actaastro.2024.09.059

This study aimed to assess the decomposition rate of highly concentrated hydrogen peroxide on catalyst samples, based on the various materials and the geometry of the supporting structures. The drop test method was used and involved delivering a droplet of hydrogen peroxide to a catalyst sample in a sealed chamber, and then measuring the pressure rise. Different sample groups showed various pressure increase rates, indicating differences in catalytic performance. The drop tests were followed by a material investigation to compare the morphology of the active layer produced on various substrates and the effects of the high-test peroxide (HTP) interaction with the surface after the tests. Qualitative tests of phase composition and chemical composition were also performed. Different substrates resulted in varying morphology based on material type, chemical composition, and porosity. The study was treated as a screening test to choose the best catalytic pack option for testing under thruster-like conditions.

本研究旨在根据各种材料和支撑结构的几何形状，评估高浓度过氧化氢在催化剂样品上的分解率。研究采用了液滴测试法，将过氧化氢液滴滴入密封舱中的催化剂样品，然后测量压力上升情况。不同的样品组显示出不同的压力上升率，表明催化性能存在差异。滴液测试之后还进行了材料调查，以比较在不同基底上产生的活性层的形态以及测试后过氧化氢（HTP）与表面相互作用的影响。此外，还对相组成和化学成分进行了定性测试。根据材料类型、化学成分和孔隙率的不同，不同的基底会产生不同的形态。这项研究被视为一项筛选测试，目的是选择最佳催化包方案，以便在类似推进器的条件下进行测试。

引用次数: 0

EDT demonstration for keeping low altitude orbit using carbon nanotube tether 利用碳纳米管系绳保持低空轨道的 EDT 演示

IF 3.1 2区物理与天体物理 Q1 ENGINEERING, AEROSPACE

Acta Astronautica

Pub Date : 2024-10-01 DOI: 10.1016/j.actaastro.2024.09.066

This paper describes an electrodynamic tether system (EDT). Conductive tethers are used to overcome atmospheric resistance in low orbit and extend orbit life using EDT propulsion. First, based on the conceptual design, the feasibility of the EDT system has been evaluated through simulation. Especially, the induced electromotive force of the conductive tether has been clarified, the amount of current required to keep the orbital altitude, and the requirements for each device. Then, feasibilities of the key mission devices have been evaluated experimentally. Those are the conductive tether, the electron emission device, and the tether extension control mechanism. The conductive tether is planned to be made of carbon nanotubes, which are attracting attention as a new material, and its properties are evaluated. The electron-emitting device has been evaluated as a device that can be mounted on a micro/nano satellite. Tether extension mechanism was developed in our past project, then it should be evaluated with the purpose of adapting to conductive tether systems. Based on the evaluation, the orbital demonstration mission plan using a micro/nano satellite has been introduced. The mission is as follows. The conductive tether is extended and stabilized by the gravity gradient. By emitting electrons from an electron emitter mounted on a high-altitude satellite and collecting electrons on bare tether at low-altitude end, a current is generated from the high-altitude side to the low-altitude side. According to Fleming's law, the Lorentz force is generated in the direction of the orbiting accelerated, so it can be propelled in the upward direction of the orbit.

本文介绍了电动系绳系统（EDT）。导电系绳用于克服低轨道的大气阻力，并利用 EDT 推进器延长轨道寿命。首先，根据概念设计，通过模拟评估了 EDT 系统的可行性。特别是明确了导电系绳的感应电动势、保持轨道高度所需的电流大小以及对每个装置的要求。然后，对关键任务装置的可行性进行了实验评估。这些装置包括导电系绳、电子发射装置和系绳延伸控制机制。导电系绳计划由碳纳米管制成，碳纳米管作为一种新材料备受关注，我们对其特性进行了评估。电子发射装置已被评估为可安装在微型/纳米卫星上的装置。系绳延伸机制是在我们过去的项目中开发的，因此应该对其进行评估，以适应导电系绳系统。在评估的基础上，提出了使用微型/纳米卫星的轨道演示任务计划。任务如下。导电系绳在重力梯度的作用下延伸并稳定。通过安装在高空卫星上的电子发射器发射电子，并在低空一端的裸系带上收集电子，从而产生从高空一侧到低空一侧的电流。根据弗莱明定律，洛伦兹力产生于轨道加速的方向，因此可以推动卫星沿轨道向上运行。

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引用次数: 0

Modular design and structural optimization of CubeSat separation mechanism 立方体卫星分离机构的模块化设计和结构优化

IF 3.1 2区物理与天体物理 Q1 ENGINEERING, AEROSPACE

Acta Astronautica

Pub Date : 2024-09-30 DOI: 10.1016/j.actaastro.2024.09.067

In order to achieve the design requirements of light weight and high stiffness of the CubeSat modular separation mechanism, a BPNN-GA-PSO size optimization method is proposed to optimize the design of the separation mechanism. Firstly, the separation mechanism of modularization and standardization is analyzed to select the optimization objects. Then, a hierarchical optimization strategy of topology optimization followed by size optimization is adopted to find the global optimum using a hybrid GA-PSO optimization algorithm. Meanwhile, the BPNN surrogate model is introduced to improve the optimization efficiency. The results show that the mass proportion of the optimized separation mechanism is reduced to 18 %, and the maximum deformation of the separation mechanism is 0.123 mm, which meets the design requirements of light weight and high stiffness of the separation mechanism. It proves the applicability of the optimization method to the optimal design of the separation mechanism. The CubeSat modular separation mechanism designed in this paper has been verified by ground verification with overload, vibration, and shock mechanical tests, and successfully deployed the BY-03 satellite in-orbit, which can provide reference for the design and development of subsequent CubeSat modular separation mechanisms.

为了实现立方体卫星模块化分离机构轻量化、高刚度的设计要求，提出了一种BPNN-GA-PSO尺寸优化方法来优化分离机构的设计。首先，分析模块化和标准化的分离机构，选择优化对象。然后，采用先拓扑优化后规模优化的分层优化策略，利用混合 GA-PSO 优化算法找到全局最优。同时，引入 BPNN 代理模型来提高优化效率。结果表明，优化后分离机构的质量比例降低到 18%，分离机构的最大变形量为 0.123 mm，满足了分离机构轻质高刚度的设计要求。这证明优化方法适用于分离机构的优化设计。本文设计的立方体卫星模块化分离机构经过过载、振动和冲击力学试验的地面验证，并成功将BY-03卫星部署在轨，可为后续立方体卫星模块化分离机构的设计和开发提供参考。

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引用次数: 0

Drag reduction mechanism based on the aerospike-jet composite approach under different incoming dynamic pressures 不同传入动压下基于气钉喷气复合方法的减阻机制

IF 3.1 2区物理与天体物理 Q1 ENGINEERING, AEROSPACE

Acta Astronautica

Pub Date : 2024-09-30 DOI: 10.1016/j.actaastro.2024.09.068

The drag experienced by the nose of an aircraft varies significantly under different incoming flow conditions. In order to reduce the aerodynamic force on the nose of the (hypersonic) supersonic vehicle, the numerical simulation approach has been used to analyze the drag reduction laws and mechanisms of a blunt-body vehicle with the aerospike-counterflowing jet composite approach under different incoming flow dynamic pressures. The effects of two environments with H = 0.1 km and H = 30 km, as well as the influence of different incoming Mach numbers on the drag reduction at H = 0.1 km were investigated. The research has shown that there exists a minimum overall drag value for the vehicle under different dynamic pressures, while the variation in the wall pressure is related to the actual operating conditions. The recirculation zones at the aerospike and in front of the blunt body are important flow structures that affect the drag reduction. When the jet total pressure ratio increases, the position of the recirculation zone at the head of the drag reduction rod moves away from the central axis of the model. When the incoming flow dynamic pressure is higher, the separation point of the recirculation zone in front of the blunt body is located further downstream, and the reattachment point is related to the incoming flow and jet conditions, aligning with the trend of variation of the position of the maximum pressure on the blunt body wall. As it can effectively push away the shock wave, the enhanced drag reduction effect of the jet is significant when the incoming flow dynamic pressure is higher. However, the high jet pressure resulting from high dynamic pressure needs to be considered in the actual design. Furthermore, the complex turbulent flow field formed by the high dynamic pressure incoming flow and the jet is worthy of further study by means of the large eddy simulation.

在不同的入流条件下，飞机机头受到的阻力有很大不同。为了降低（高超声速）超音速飞行器机头的气动力，采用数值模拟方法分析了钝体飞行器在不同来流动压下采用气刺-逆流喷气复合方法的减阻规律和机理。研究了 H = 0.1 千米和 H = 30 千米两种环境的影响，以及不同来流马赫数对 H = 0.1 千米时阻力降低的影响。研究表明，在不同的动态压力下，飞行器存在一个最小总阻力值，而壁面压力的变化与实际运行条件有关。气刺和钝体前方的再循环区是影响阻力降低的重要流动结构。当射流总压比增加时，减阻杆头部的再循环区位置会远离模型的中心轴线。当来流动压较高时，钝体前方再循环区的分离点位于更下游的位置，重新附着点与来流和射流条件有关，与钝体壁上最大压力位置的变化趋势一致。由于射流可以有效地推开冲击波，因此当来流动压较高时，射流的减阻效果会显著增强。然而，在实际设计中需要考虑高动压导致的高射流压力。此外，高动压入流与射流形成的复杂湍流场也值得通过大涡模拟进一步研究。

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引用次数: 0

Combustion flow field and thermal protection performance of a liquid oxygen/kerosene engine with pintle injectors 采用针形喷射器的液氧/煤油发动机的燃烧流场和热保护性能

IF 3.1 2区物理与天体物理 Q1 ENGINEERING, AEROSPACE

Acta Astronautica

Pub Date : 2024-09-30 DOI: 10.1016/j.actaastro.2024.09.065

A pintle injector is a convenient method for adjusting propellant combustion and has received extensive attention in recent years. However, few studies have examined the effects of the injection direction and momentum ratio (ratio of liquid oxygen momentum to kerosene momentum) on the flow field, temperature field, and dual-belt liquid film cooling effect in the combustion chamber. This study verified a new computational fluid dynamics (CFD) engine combustion model based on a pintle injector through a hot-test experiment. A simulation was conducted using this model. With different injection directions, when the momentum ratio was increased, the trends in the combustion efficiency and chamber pressure growth were entirely opposite. For the axial kerosene and radial liquid oxygen injection direction of liquid film cooling, when the momentum ratio was 0.73–0.9, the proportion of the first ring belt liquid film flow was selected as 5–10 %, and the second ring belt liquid film flow proportion was selected as 5–7%. When the momentum ratio was less than 0.73, the liquid film flow proportions for both the first and second ring belts were selected as 5 %. When the total liquid film flow proportion was 10 %, the chamber pressure decreased by 0.2–0.3 MPa, and the combustion efficiency decreased by 8–10 %. This numerical study provides guidance for studying the combustion characteristics and cooling performance of pintle-type rocket combustion chambers.

凤尾喷射器是一种调整推进剂燃烧的便捷方法，近年来受到广泛关注。然而，很少有研究探讨喷射方向和动量比（液氧动量与煤油动量之比）对燃烧室流场、温度场和双带液膜冷却效果的影响。本研究通过热试实验验证了基于针形喷油器的新型计算流体动力学（CFD）发动机燃烧模型。使用该模型进行了模拟。在不同的喷射方向上，当动量比增加时，燃烧效率和燃烧室压力的增长趋势完全相反。对于液膜冷却的轴向煤油和径向液氧喷射方向，当动量比为 0.73-0.9 时，第一环带液膜流比例选择为 5-10%，第二环带液膜流比例选择为 5-7%。当动量比小于 0.73 时，第一环带和第二环带的液膜流量比例均为 5%。当总液膜流比例为 10 % 时，燃烧室压力降低了 0.2-0.3 MPa，燃烧效率降低了 8-10%。这项数值研究为研究品特型火箭燃烧室的燃烧特性和冷却性能提供了指导。

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引用次数: 0

Semi-analytic approximate time-optimal asteroid landing with dimensionality reduction shooting 半解析近似时间最优小行星着陆降维射击

IF 3.1 2区物理与天体物理 Q1 ENGINEERING, AEROSPACE

Acta Astronautica

Pub Date : 2024-09-28 DOI: 10.1016/j.actaastro.2024.09.032

Pinpoint landing on asteroid is very challenging due to the uncertainties of the gravitational field, which highlights the urgent need for autonomous landing algorithms under such uncertainties. Some well-known guidance algorithms, such as APDG and real-time convex programming, either ignore thrust boundary constraints or are severely time consuming. In addition, these algorithms have the same difficulty in analyzing landing errors due to model uncertainty propagation, making the reliability of the algorithms in actual landings questionable. To address these challenges, we propose a landing framework that combines model identification, trajectory analytical solution and closed-loop corrections to achieve near-optimal real-time landing control. Firstly, we simplified the landing scenario, focusing on reducing the dimension of shooting variables to enable rapid trajectory calculation. Secondly, we derived error propagation equations and established criteria for trajectory replanning based on landing error prediction. Finally, we validated our approach with practical examples of the Shoemaker probe landing on Eros 433. Through real-time landing error estimation, our proposed framework enables spacecraft to achieve near time-optimal land within a given error threshold.

由于重力场的不确定性，在小行星上进行精确定位着陆非常具有挑战性，这凸显了在这种不确定性下对自主着陆算法的迫切需求。一些众所周知的制导算法，如 APDG 和实时凸编程，要么忽略了推力边界约束，要么非常耗时。此外，这些算法在分析模型不确定性传播导致的着陆误差时同样存在困难，使得算法在实际着陆中的可靠性受到质疑。为解决这些难题，我们提出了一种结合模型识别、轨迹分析求解和闭环修正的着陆框架，以实现近乎最优的实时着陆控制。首先，我们简化了着陆场景，重点是减少射击变量的维度，以实现快速轨迹计算。其次，我们推导了误差传播方程，并建立了基于着陆误差预测的轨迹重新规划标准。最后，我们以 "舒梅克 "探测器在厄洛斯 433 上着陆的实际案例验证了我们的方法。通过实时着陆误差估算，我们提出的框架可使航天器在给定误差阈值内实现接近时间最优的着陆。

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引用次数: 0

Accelerated biological evolution in outer space: Insights from numerical analysis 加速外太空生物进化：数值分析的启示

IF 3.1 2区物理与天体物理 Q1 ENGINEERING, AEROSPACE

Acta Astronautica

Pub Date : 2024-09-28 DOI: 10.1016/j.actaastro.2024.09.044

As humanity continues its space exploration, understanding biological evolution in extraterrestrial environments will become crucial. On Earth, organisms have adapted to new environments, and some genetic data indicate positive natural selection. This paper investigates the impact of space environments, such as high radiation and microgravity, which may lead to high mutation rates and positive selection, on biological evolution, using numerical analysis. It quantifies the evolutionary rates and the time until a new mutation reaches fixation (100 % frequency within population) beyond Earth. The findings reveal accelerated evolution rates, 1,000 to 10,000 times faster than on Earth for beneficial mutations, with the time until fixation being 0.002 to 0.004 times shorter, assuming mutation rates are 10–100 times higher. These results offer insights into various areas, including space facility design, space agriculture, astrobiological exploration, and life sustainability beyond Earth and Solar System, illuminating the potential for a ‘Big Bang of Evolution’ in outer space.

随着人类继续进行太空探索，了解地外环境中的生物进化将变得至关重要。在地球上，生物已经适应了新的环境，一些遗传数据显示了积极的自然选择。高辐射和微重力等太空环境可能导致高突变率和正选择，本文通过数值分析研究了这些环境对生物进化的影响。它量化了地球以外的进化速度和新突变达到固定（种群内频率达到 100%）所需的时间。研究结果表明，有益突变的进化速度比地球快 1000 到 10000 倍，假设突变率比地球高 10-100 倍，则突变达到固定的时间缩短 0.002 到 0.004 倍。这些结果为包括太空设施设计、太空农业、天体生物学探索以及地球和太阳系以外的生命可持续性在内的各个领域提供了启示，揭示了外太空 "进化大爆炸 "的潜力。

引用次数: 0

Learning-based control for deployment and retrieval of a spinning tethered satellite formation system 基于学习的控制，用于部署和回收旋转系留卫星编队系统

IF 3.1 2区物理与天体物理 Q1 ENGINEERING, AEROSPACE

Acta Astronautica

Pub Date : 2024-09-28 DOI: 10.1016/j.actaastro.2024.09.061

This paper investigates the nonlinear dynamics and control of the deployment and retrieval for a spinning tethered satellite formation system via artificial intelligent method. A dynamic model of the spinning tethered formation system is developed to describe the attitude motions of the system, involving the relative rotations of the tethers to the central main satellite. Considering the system with symmetric and asymmetric configurations, a learning-based control strategy with low time cost is proposed to achieve the stable deployment and retrieval of tethers. In the strategy, a nonlinear model predictive control law accounting for the control constraints and nonlinear dynamics is developed to achieve the control goal. Based on a deep learning method, a dataset including control input and state output obtained offline is trained to form deep neural networks. An online feedback control of the system can be achieved by conducting real-time mapping from the system state to the control input using the neural networks. Finally, numerical simulations for deployment and retrieval of the system with different configurations are presented to demonstrate the computational efficiency and to validate the effectiveness of the control strategy.

本文通过人工智能方法研究了旋转系留卫星编队系统部署和回收的非线性动力学和控制。建立了旋转系留编队系统的动态模型来描述系统的姿态运动，包括系留卫星与中央主卫星的相对旋转。考虑到系统的对称和非对称配置，提出了一种基于学习的、时间成本低的控制策略，以实现系绳的稳定布放和回收。在该策略中，开发了一种考虑控制约束和非线性动力学的非线性模型预测控制法则，以实现控制目标。基于深度学习方法，将离线获得的控制输入和状态输出数据集训练形成深度神经网络。利用神经网络从系统状态到控制输入进行实时映射，从而实现系统的在线反馈控制。最后，介绍了不同配置下系统部署和检索的数值模拟，以展示计算效率并验证控制策略的有效性。

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引用次数: 0