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Acta Astronautica
Journal Prestige (SJR): 0.758
Citation Impact (citeScore): 2
Number of Followers: 403  
 
  Hybrid Journal Hybrid journal (It can contain Open Access articles)
ISSN (Print) 0094-5765
Published by Elsevier Homepage  [3162 journals]
  • Utilization of trash for radiation protection during manned space missions
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Feng Xu, Xianghong Jia, Wei Lu, Chenglong Zhou, Yaoyu Guo, Jinxue Fei, Chunxin Yang High-energy charged particles in space pose a severe threat to the health and safety of astronauts. This is especially the case in deep space exploration owing to the absence of magnetic field protection as present in low Earth orbit (LEO) manned spaceflight. This has necessitated the investigation and development of effective space radiation protection materials and methods. Manned space missions produce a significant amount of trash, which can be compressed by a heat melt compactor (HMC) to reduce space utilization. The trash would also be sterilized during the compaction process and the extracted water can be recycled. The processed trash can potentially be used for space radiation protection, with the benefit of reduced launch load. In this study, the Monte Carlo method was used to acquire information about the primary and secondary radiation particles that emerged from different radiation shields made from an HMC-processed model trash of a manned space mission, as well as those made from aluminum and water, which are common, and currently used space radiation protection materials. The types and energies of the considered incident radiation particles were based on space radiation environment spectra. A comparison of the space radiation protection capacities of the different shields revealed that HMC-processed trash was superior to aluminum and water. Processed trash thus promises to be a practicable alternative to water for the construction of radiation emergency areas for future deep space missions as the mission proceeds and the water is consumed.
       
  • Integrated vibration isolation and attitude control for spacecraft with
           uncertain or unknown payload inertia parameters
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Zixi Guo, Yao Zhang, Quan Hu To meet the spacecraft attitude control requirements with high accuracy and stability, all vibrations in the spacecraft should be reduced in appropriate ways. This paper presents an integrated control method for attitude and the vibrations in both high frequency and low frequency in the spacecraft. The integrated control method includes a vibration isolation platform and a modified adaptive attitude control method. The paper presents a vibration isolation platform with magnetic suspension to reduce high frequency vibrations and a parameter design method for the platform. An adaptive control method is presented to reduce low frequency vibrations while accounting for the bandwidth constraint due to the vibration isolation platform. Firstly, a parameter design method is proposed for the vibration isolation platform, and an entire 6×12 dimensional transformation matrix is derived for the case that the inertia of the payload is of the same order of magnitude as that of spacecraft bus. Then, an adaptive attitude controller is presented that accounts for the coupling characteristics of the spacecraft, the vibration isolation platform and the uncertain or unknown payload inertia parameters. To ensure the robustness of the attitude control system and the performance of the vibration isolation system, a method of estimating the initial value of the payload inertia is presented using classical control theory. Finally, numerical simulations demonstrate that the integrated control method presented in this paper can achieve the attitude control task for spacecraft with high accuracy and stability.
       
  • Numerical investigation of bleeding control method on section-controllable
           wavecatcher intakes
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Fengyuan Zuo, Guoping Huang The bleeding control method and analysis of advantages, challenges on wavecatcher (inward turning) intakes are investigated in this paper. Firstly, a quantitative analysis on the substantial advantages of wavecatcher intakes for ramjet is presented comparing to two-dimensional planar symmetry compression. According to the results, with the same parameters in the entrance and exit section, the total pressure recovery of Internal Conical Flow of C increases by 28.0%; the compression surface length decreases by 5%; the wetted area decreases by 15.7% and the pressure drag decreases by 12.1%. However, due to high compressive efficiency of wavecatcher intake, the boundary layer experiences a higher adverse pressure gradient, contributing to enhance shock wave/boundary layer interaction (SBLI). Secondly, the effects of bleeding control on wavecatcher intake are elucidated by numerical simulations. The bleeding control improves the flow structures by decreasing the boundary layer thickness to weaken the SBLI and bleeding the spanwise vortex out. Furthermore, due to the weak interaction, the terminal shock wave is stable closely behind the throat section, increasing the resistance against the back pressure, decreasing the Mach number before the terminal shock wave and improving the total pressure recovery of the exit section by 3.73% relatively.
       
  • Relative control of an ion beam shepherd satellite using the impulse
           compensation thruster
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): A. Alpatov, S. Khoroshylov, C. Bombardelli The “ion beam shepherd” is a recently proposed concept for removing space debris in a contactless manner. A shepherd satellite must be controlled to move at a certain small distance in front of a space debris object during the de-orbiting phase. Because of the considerable duration of this phase, the propellant consumption is a key requirement for the control design. In this paper, the in-plane relative position of the shepherd is maintained using a small thrust variation of the compensation thruster. The controller is designed and analyzed considering the time-varying and parametric uncertain plant in the presence of the ion beam and orbital perturbations, sensor noise, actuation errors, taking into account limitations on the controller output. The system robustness and specified requirements are confirmed both by a formal criteria and numerical simulations. The estimations show that this control strategy is more efficient in terms of propellant consumption than the conventional approach with chemical thrusters.
       
  • Contact dynamics and control of a space robot capturing a tumbling object
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Shuang Wu, Fangli Mou, Qian Liu, Jing Cheng Capture of a free-floating space object in orbit is a challenging task especially when the object is tumbling. In this paper, the contact dynamics modeling and control problem for capturing a fast tumbling target object by a space robot are investigated. A generic frictional contact model is developed to represent the contact forces between the robot's end-effector and the target object. The frictional contact formulation is based on the compliance contact force and bristle friction model which can simulate intermittent frictional contact situations involving multiple-point contacts between contact interfaces with complex geometries. A resolved motion admittance control method is designed to realize a good tracking for a tumbling target object while increasing the compliance of the space robot. A simulation example of a 7-joint manipulator capturing a tumbling object in three dimensions is presented. The simulation results revealed that various contact scenarios during the capture process can be well simulated with the developed contact model and a good performance of the designed control method for capturing a fast tumbling target object.
       
  • Performance evaluation methodology for multistage launch vehicles with
           high-fidelity modeling
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Marco Pallone, Mauro Pontani, Paolo Teofilatto Multistage launch vehicles of reduced size, such as ”Super Strypi” or ”Sword”, are currently investigated for the purpose of providing launch opportunities for microsatellites. Currently, microsatellites are launched according to timing and orbit requirements of the main payload. The limited costs of microsatellites and their capability to be produced and ready for use in short time make them particularly suitable for ready-on-demand requests, such as facing an emergency. As a result, launch vehicles for the exclusive use of microsatellites would be very useful. This work considers the Scout rocket, a four-stage launch vehicle of reduced size used in the past. Its aerodynamics and propulsion are modeled with high fidelity, through interpolation of reliable, accurate available data. For the purpose of reducing the rocket complexity and size, as well as the launch cost per kg of payload, simplification of the rocket subsystems is advisable, and this includes also the guidance system and the related algorithm. In fact, open-loop guidance was actually employed during real Scout flights. In this research, open-loop guidance is investigated, under the assumption that the aerodynamic angle of attack is constant for each of the first three stages. Instead, for the upper stage the terminal optimal ascent path leading to orbit injection is determined through the use of a specific implementation of firework algorithm, in conjunction with the Euler-Lagrange equations and the Pontryagin minimum principle. Firework algorithms represent a recently-introduced heuristic technique inspired by the firework explosions in the night sky. The concept that underlies this method is relatively simple: a firework explodes in the search space of the unknown parameters, with amplitude and number of sparks determined dynamically. The succeeding iterations preserve the best sparks. The firework algorithm has several original features that can ensure satisfactory performance in parameter optimization problems, because both local search and global search are effectively performed through combination of various stochastic operators. With regard to the problem at hand, the unknown parameters are (i) the aerodynamic angles of attack of the first three stages, (ii) the coast time interval and (iii) the initial values of the adjoint variables conjugate to the upper stage dynamics. The numerical results unequivocally prove that the methodology at hand is rather robust, effective, and accurate, and definitely allows evaluating the performance attainable from multistage launch vehicles with accurate aerodynamic and propulsive modeling.
       
  • Investigation of self-pulsation characteristics for a liquid-centered
           swirl coaxial injector with recess
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Xiao Bai, Qinglian Li, Peng Cheng, Liyong Sheng, Zhongtao Kang A combined experimental and numerical investigation was conducted to explore the self-pulsation characteristics generated by a liquid-centered swirl coaxial injector with an inner post recess of 5 mm. A back-lighting photography technique was employed to capture the instantaneous flow patterns in recess chamber with a high speed camera. Pressure oscillations in recess chamber were also measured by the miniature pressure transducer. 2-D unsteady numerical simulations basing on swirl axi-symmetric model were performed. Good agreements were generally achieved between numerical simulation and experiment in aspect for the mechanism and characteristic frequency of self-pulsation. It was found that the blocking actions of the conical liquid sheet play a crucial role in self-pulsation for the coaxial injector with recess. Self-pulsation occurs coinciding with strong pressure oscillation in recess chamber. The frequencies of spray oscillation and gas pressure in recess chamber correspond well with each other. Self-pulsation frequencies are approximately linearly proportional to the liquid Reynolds number under certain conditions. In addition, the spray oscillation transforms from high frequency pulsation (at about 3000 Hz) to ultra-high frequency pulsation (ranging from 8000 Hz to 9000 Hz) with an increase in the gas Reynolds number when the mass flow rate of liquid is large enough.
       
  • Design methodology of the waverider with a controllable planar shape
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Jifei Wang, Chuanzhen Liu, Peng Bai, Jinsheng Cai, Yuan Tian A novel design method of waverider with a controllable planar shape is proposed as a further development of the conventional osculating cone/flowfield method. The present method is run based on a newly established geometric relationship, represented by a differential equation set, involving the flow capture curve (FCC), the inlet capture curve (ICC) and the planar shape curve (PSC). As two of the three curves are known, the last one is easily determined by following the relationship. Therefore, the couples of FCC-PSC or ICC-PSC are introduced as design-driving parameters whereas the conventional methods just employ the couple of FCC-ICC, and then the waverider planar shape is directly specified in the design process instead of other indirect parameters. Two predefined planar shapes are employed to generate waverider configurations as test cases. The planar shapes of the design results are precisely controlled by the predefined curves, verifying the correctness of the geometric relationship. Furthermore, the numerical simulations show that customizing the planar shape does not destroy the excellent characteristics of waverider, and thus the high lift-to-drag ratio on hypersonic conditions is maintained. Since the used planar shapes are suitable for low-speed flight to the engineering point of view, the low-speed performance is significantly improved as well. The present method improves the waverider design flexibility by introducing the planar shape as a design parameter, and the ideal of planar shape customization also inspires to the design of wide-speed-range configurations.
       
  • An easy-to-implement thermal test system for large deployable antennas
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Hao Wu, Meng Liu, Jun Wang, Yuansheng Zhang, Gang An Large deployable antennas (LDAs), having a wide range of applications in aerospace engineering, encounter extreme thermal conditions when subjected to the space environment. Ground thermal test facilities are used to validate LDA deployability and accuracy under extreme thermal conditions. General antenna thermal test facility is thermal vacuum test facility, which is complex and costly as the vacuum test chamber has to be pumped down and large enough to accommodate LDAs. In this paper, an easy-to-implement thermal test system is presented, which simulates atmospheric thermal environment for LDAs using an air cycle refrigeration system and electric heaters. Additionally, a series of measures are used to ensure uniform temperature and limit airflow turbulence. Test results show that the test system can provide dry, uniform temperature and small disturbance thermal environment for LDAs.
       
  • The influence of coolant jet direction on heat reduction on the nose cone
           with Aerodome at supersonic flow
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): R. Moradi, M. Mosavat, M. Barzegar Gerdroodbary, A. Abdollahi, Younes Amini Reduction of aerodynamic heating is a critical issue for the development of the hypersonic vehicles. In this study, a computational fluid dynamic is applied to study the effect of location of coolant jet in the vicinity of the aerodome on the heat reduction of the nose cone at M = 5. In addition, the influence of the gas types (Air, He and CO2) on the cooling performance is investigated. This research mainly focused the flow feature and mass distributions of various coolant jets. In order to study these effects, a two-dimensional model with spike is chosen to simulate the various shocks in the vicinity of the nose cone. The effect of significant parameters is studied by using the Reynolds-averaged Navier–Stokes equations with Menter's Shear Stress Transport (SST) turbulence model. Results show that the injection of the coolant gas from the top of aerodome significantly decreases the heat load on the nose cone. In addition, injection of the coolant jet from the top is more efficient on the recirculation region on the top of spike. The obtained results reveal that the injection of coolant from the front of the aerodome does not reduce the heat load substantially. In addition, the cooling performance of helium jet as the lateral jet is 15% more than other gases.
       
  • An investigation of millimeter wave reflectarrays for small satellite
           platforms
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Ghulam Ahmad, Tim W.C. Brown, Craig I. Underwood, Tian H. Loh This article reports two contributions related to reflectarray antenna design at millimeter waves (mm-waves). First, a closed form analytical formulation is provided for the prediction of reflection properties of square/rectangular mm-waves reflectarray unit cells based on various quality factors and the theory of waveguide coupled resonators. To ensure a high accuracy at mm-waves, the effects of fringing fields, surface waves, metal conductivity, and metal surface roughness are included in the analysis. This analysis program greatly facilitates the parametric studies of a unit cell's constituting parameters to converge on an optimum design solution. Secondly, the concept of phase quantization is proposed for a cost effective realization of mm-waves reflectarrays. The developed formulation in the first contribution was used to design two 3 bit phase quantized, single layer, 19 wavelength, passive reflectarrays at 60 GHz. The test results are compared with simulations and a very good agreement was observed. These findings are potentially useful for the realization of high gain antennas for mm-wave inter-satellite links in small satellite platforms.
       
  • An anti-saturation steering law for Three Dimensional Magnetically
           Suspended Wheel cluster with angle constraint
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Rui Zhang, Yuanjin Yu, Chao Han, Zhaohua Yang Three Dimensional Magnetically Suspended Wheel (3-DMSW) is a new kind of inertia actuator for spacecraft attitude control, which can provide a 3 degrees of freedom torque. On account of the constraint characteristics of 3-DMSW such as a small deflection saturation angle of rotor shaft and the saturation of rotor's variable rotational speed, an anti-saturation steering law based on weighted pseudo inverse is proposed for 3-DMSW cluster. A new weight adjustment method is proposed to adjust the weights of shaft deflections dynamically. A specially designed exponential function with current deflection angle and angular velocity information on the exponent position is adopted as the evaluation criterion of current torque output ability of shaft deflection. Thus the torque command can be distributed dynamically with no angle saturation. The weight adjustment method is demonstrated theoretically and the effectiveness of the anti-saturation steering law is validated by conducting several numerical simulations of attitude agile maneuver. Comparing with the 3-DMSW cluster and flywheel cluster using the traditional steering law, the results show that the 3-DMSW cluster using the proposed method makes the process of agile maneuver more rapid and accurate and the saturation angles of 3-DMSW cluster will not be reached.
       
  • Effects of solar panels on Aerodynamics of a small satellite with
           deployable aero-brake
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): S. Mungiguerra, G. Zuppardi, L. Spanò Cuomo, R. Savino This work is focused on the aerodynamic analysis of a small satellite provided with a deployable aero-brake. The satellite is intended to perform a completely aerodynamic de-orbiting maneuver from Low-Earth-Orbit. A brief discussion about the aerodynamic effects of the position of the aero-brake along the longitudinal axis of a simplified axisymmetric system is presented. Moreover, a more complex architecture, envisaging deployable solar panels for the enhancement of power generation along the orbital path, is proposed and analyzed. The present paper is aimed at the evaluation of the influence of such a configuration on the satellite aerodynamic parameters. Computations have been carried out by means of a Direct Simulation Monte Carlo (DSMC) code at altitude of 150 km, velocity of 7800 m/s and in the interval of angle of attack 0–180 deg with a spacing of 10 deg. The results verified that the deployable solar panels strongly influence Aerodynamics of the satellite. One of the most relevant aspects is the variation of the longitudinal stability equilibrium that becomes more stable. Furthermore, the deployable solar panels increase the aerodynamic drag when the aero-brake is closed, affecting the drag modulation capability.
       
  • Space debris collision probability analysis for proposed global broadband
           constellations
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): S. Le May, S. Gehly, B.A. Carter, S. Flegel Fragmentation events, caused by the collision of two objects in space, have been a significant source of space debris objects over a cumulative five decades of space activity. Current proposals by different commercial entities aim to launch constellations comprising thousands of satellites in Low Earth Orbit (LEO), which would result in an increase of more than five times the number of currently active satellites in a region where debris objects are most concentrated. The Inter-Agency Space Debris Coordination Committee (IADC) has already recognized the potential influence of large constellations on the LEO environment and the subsequent need to assess whether current mitigation guidelines will be adequate moving forward. Given developments for such constellations are already underway, independent research efforts ahead of any revision to current IADC guidelines could be of great value not only to the organizations involved in their operation, but also to policymakers and existing space users. This paper evaluates the probability of collisions for mega-constellations operating in the current LEO debris environment under best and worst-case implementation of current mitigation guidelines. Simulation studies are performed using the European Space Agency's (ESA) MASTER-2009 debris evolutionary model, and the specifications of the proposed OneWeb and SpaceX constellations as example mega-constellations. Multiple scenarios are then tested to assess mitigation measures and their ability to minimize the probability of fragmentation events and the creation of new debris in LEO.
       
  • Investigation on plume expansion and ionization in a laser ablation plasma
           thruster
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Yu Zhang, Jianjun Wu, Daixian Zhang, Sheng Tan, Yang Ou The laser ablation plasma thruster is a novel electric propulsion thruster, which combined the laser ablation and electromagnetic acceleration. In order to investigate the plume expansion and ionization in the laser ablation plasma thruster which was difficult to obtain from experiments, the two-dimensional heat conduction model and fluid dynamics model were established. The heat conduction model was established to calculate the target ablation, taking into account temperature dependent material properties, phase transition, dielectric transition and phase explosion. The fluid dynamics model was used to calculate the plume properties, taking into account ionization, plume absorption and shielding. The good agreement between calculated and experimental data validated our model, while the plume velocity, temperature and electron number density were predicted by using the numerical method. The calculated results showed that the plume uniformly expanded into the ambience with a mushroom shape, and the peak values of plume velocity, temperature and electron number density fraction were distributed at the front of the plume. The ceramic tube limited the radial expansion of the plume, and enhanced the velocity, temperature and ionization degree nearby the wall, due to the interaction between the plume and the wall. Otherwise, the effects of laser fluence on plume properties and thrust performance of the thruster were investigated utilizing the numerical model.
       
  • Design and analysis of flexure revolute joint based on four-bar mechanism
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Jing Zhang, Hong-wei Guo, Juan Wu, Gui-jun Gao, Zi-ming Kou, Anders Eriksson In order to avoid the stress concentration and increase rotational angle of a flexure joint, the method of partial separation of storage elements in the motion transmission elements is proposed. A type of flexure revolute joint with large rotational angle is designed based on the block approach. By setting a 4-bar mechanism as the intermediate block which connects the outer ring and the inner ring of the revolute joint, and replacing the rigid bar by a flexible beam, large rotational angles of the joint can be achieved. The basic size of the joint is designed by setting the initial and the constraint condition of the 4-bar mechanism. Then, influence analyses of the size of the linkage joint and large flexible beam on the stress, the torque, and the torsional stiffness are conducted by using nonlinear static analysis method. Based on the requirements for torque and rotational stiffness, the size of the flexure revolute joint is defined. Experiments on the joint, which can rotate 90°, are conducted.
       
  • NASA's eXploration Systems and Habitation (X-Hab) Academic Innovation
           Challenge
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Jason Crusan, Carol Galica, Tracy Gill Future exploration missions in the space between the Earth and the Moon or beyond will require complex operational activities to ensure that crew, cargo, and exploration systems safely reach their destination. Through the eXploration Systems and Habitation (X-Hab) Academic Innovation Challenge, NASA develops strategic partnerships and collaborations with universities to increase knowledge in technologies, capabilities, and operational approaches related to future human spaceflight missions. X-Hab activities help NASA bridge strategic knowledge gaps, better understand technology risk reduction, and combine the innovative approaches and diverse insights of university teams with unique agency expertise.The X-Hab Academic Innovation competition links with senior- and graduate-level design curricula that emphasize hands-on development of functional prototypes for deep space exploration missions. Research topics are identified and funded annually by NASA technology projects in collaboration with the National Space Grant Foundation. University teams submit proposals based on their interests and capabilities, and multiple small awards are made for the design and creation of studies or products that align with NASA strategic objectives. The selected project teams implement the design course during the fall and spring semesters using a systems engineering approach that requires formal reviews with NASA for requirements and system definition, preliminary design, and critical design. The challenges allow students to follow genuine hardware and systems engineering development processes and gain valuable experience that will extend to their professional careers.Since 2011, NASA has selected 49 X-Hab student concepts to address space habitation systems including advanced fabrication concepts, plant growth, atmosphere management, waste handling, and recycling. This paper provides a status and overview of submissions received, selected projects, success stories, and lessons learned. It also details methods employed by NASA to manage and promote the X-Hab competition, summative information on participating organizations, and next steps for the activity. The X-Hab project assists NASA in optimizing technology investments, fosters innovation and facilitates technology infusions that address specific, real-world challenges being faced by NASA as the agency works to send humans further into space than ever before.
       
  • CGR-BF: An efficient contact utilization scheme for predictable deep space
           Delay Tolerant Network
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Sangita Dhara, Scott Burleigh, Raja Datta, Sujoy Ghose In this paper we propose a scheme that uses the fragmentation scheme of bundle protocol for efficient utilization of contacts between nodes so as to increase the goodput in deep space networks. The node (e.g., planets/satellites/Orbiters) positions in deep space are predictable and messages are transmitted from one node to the other in a store-and-forward mode whenever these nodes are in contact with each other. Contact Graph Routing (CGR) has been proposed for interplanetary networks due to its delay tolerant nature and characteristics. The CGR chooses a single path between a source and a destination node to achieve the best delivery time of a bundle. However, the contacts that cannot transmit one complete bundle or have some leftover capacity after transmitting one or more complete bundles are ignored here. This leads to wastage of deep space transmission opportunities which are significant. In this paper, we propose a scheme for utilizing these contacts efficiently by fragmenting the bundles wherever possible. The bundle fragments are then routed to the destination using multiple paths. Simulation studies show that CGR-BF efficiently exploits the network's bandwidth and substantially increases the goodput incurring minimum overhead compared to that of the CGR where fixed size bundles are used.
       
  • Trajectory design and guidance for landing on Phobos
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Eric Joffre, Mattia Zamaro, Nuno Silva, Andrés Marcos, Pedro Simplício While common Descent and Landing strategies involve extended periods of forced motion, significant fuel savings could be achieved by exploiting the natural dynamics in the vicinity of the target. However, small bodies are characterised by perturbed and poorly known dynamics environments, calling for robust autonomous guidance, navigation and control. Airbus Defence and Space and the University of Bristol have been contracted by the UK Space Agency to investigate the optimisation of landing trajectories, including novel approaches from the dynamical systems theory, and robust nonlinear control techniques, with an application to the case of a landing on the Martian moon Phobos.
       
  • Form-finding of deployable mesh reflectors using dynamic relaxation method
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Xinyu Wang, Jianguo Cai, Ruiguo Yang, Jian Feng In this paper, a novel numerical form-finding method is presented based on a dynamic relaxation algorithm for cable nets in mesh reflectors. To obtain a shape with high profile efficiency, structural forces are assumed to be constant values during the computational iterations, so a perfectly uniform distribution of structural forces can be achieved. An initial shape, normally not in equilibrium, is given in advance, and nodes of the network are forced to vibrate by the unbalanced forces. Parameters of dynamic relaxation algorithms, such as the type of damping and the time interval, are tested to ensure the speed and stability of the computation process. Finally, two different types of Astromesh reflectors are used as examples to verify the developed method.
       
  • Guaranteeing prescribed performance for air-breathing hypersonic vehicles
           via an adaptive non-affine tracking controller
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Xiangwei Bu This paper investigates a prescribed performance control strategy for air-breathing hypersonic vehicles (AHVs) based on neural approximation. Different from the existing studies, the explored controllers are derived from non-affine models instead of affine ones. For the velocity dynamics, an adaptive neural controller containing only one neural network (NN) is addressed via prescribed performance control. Specially, the altitude dynamics is transformed into a pure feedback non-affine model instead of a strict feedback one. Then a novel adaptive neural controller is exploited without using back-stepping. Also, only one NN is utilized to approximate the lumped unknown nonlinearity of the altitude subsystem. By the merit of the minimal-learning parameter (MLP) scheme, only two learning parameters are required for neural approximation. The highlights are that the proposed control methodology possesses concise control structure and a low computational cost and moreover it can guarantee the tracking errors with prescribed performance. Finally, simulation results for an AHV model are provided to demonstrate the efficacy of the proposed control approach.
       
  • Analysis and reduction of skin-friction in a rocket-based combined-cycle
           engine flow path operating from Mach 1.5 to 6.0
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Shuai Wang, Guo-qiang He, De-kun Yan, Zhi-wei Huang, Fei Qin The skin-friction in a rocket-based combined-cycle engine operating from Mach 1.5 to 6.0 was analyzed in the present study. The friction proportion of different parts of the engine was investigated to offer a reference for the rearrangement of skin-friction reduction in the engine. The distribution and variation trend of the skin-friction in the flow path as well as its impacts on the engine performance were numerically compared. At three typical flight points, i.e. at 1.8Ma, 3.0Ma and 6.0Ma, the change of the skin-friction with attack angle was studied. A special focus was placed on the reduction of the skin-friction by using boundary layer combustion. It was modeled when the airstream flowed into the engine at the speed of 6 Ma. The method of hydrogen combustion in boundary layer has achieved 57.7% skin-friction reduction effect.
       
  • Disturbance rejection dynamic inverse control of air-breathing hypersonic
           vehicles
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Hao An, Qianqian Wu This paper presents a disturbance rejection controller for air-breathing hypersonic vehicles (AHVs) based on the technique of nonlinear dynamic inverse (NDI). An observer is employed to estimate the lumped disturbance on the nominal dynamics of AHVs, such as the external disturbances introduced by the changeable flight environment and the unsteady scramjet operation. With the help of this observer, an effective NDI controller is proposed to suppress the negative effect of the lumped disturbance on output channels. Under the proposed control, the input-to-state stability of the closed-loop AHV system can be ensured if the observer gain matrix is properly selected. A simulation study on the disturbed AHV model is provided to illustrate the effectiveness of this disturbance rejection NDI control.
       
  • Curing of large prepreg shell in solar synchronous Low Earth Orbit:
           Precession flight regimes
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): V.M. Pestrenin, I.V. Pestrenina, S.V. Rusakov, A.V. Kondyurin We investigate the curing of large shell construction made of epoxy resin/carbon fibers prepreg under free space conditions in solar synchronous Low Earth Orbit. The curing kinetics is described by first order kinetic equation with auto-acceleration and deceleration parameters based on the experimental data. Heating of the shell is provided by solar radiation. The heat distribution in the shell is modelled based on partial absorbance of the solar radiation, the prepreg thermal conductivity and thermal capacity, radiation heat transfer between inner surfaces of the shell and the gas thermal conductivity. The iterated algorithm of curing was developed. Three flight regimes based on the circular motion of a construction have been considered: 1 – the axis of a shell lies in the tangent plane to the orbit and makes a constant angle with the tangent; 2 – under conditions (1), a shell rotates around its axis with a constant angular velocity; 3 – under conditions (2), the axis of the shell precesses around the tangent to the orbit. It was found, that the parameters of the motion (i.e. angular velocity of the rotation around the axis, precession angular velocity and precession angle) could be optimised in such way, that the whole shell can be completely cured under the solar radiation.
       
  • Numerical simulations of radiative heat effects in a plasma wind-tunnel
           flow under Mars entry conditions
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Javier García-Garrido, Adrian S. Pudsey, Christian Mundt The Mars atmosphere, consisting mainly of CO2, with a few percent of N2 and other trace gases, is of interest to future space projects. Entry into its gaseous shell is of current significant research interest. For this application, an arc jet driven plasma wind tunnel is available to simulate relevant entry conditions for the planet. Recent improvements and qualification of the test facility, enables the testing on earth of high enthalpy flows with CO2 rich compositions. In order to complement the experimental analysis, numerical simulations of the test facility running at relevant ambient pressures of 600–1000 Pa, corresponding to low altitudes, have been completed. The simulations used a density-based Navier Stokes solver and non-equilibrium chemical and thermal effects which are characteristic of these types of high enthalpy flows. Special interest is given to the radiative heat transfer mechanism. Under these high temperature conditions, radiative effects become more relevant and advanced radiation models must be used. The coupling between the Navier Stokes and radiative transfer equations favours the understanding of plasma wind tunnel flows. The radiative heat is estimated using the k-distribution spectral model, which is appropriate for non-homogeneous radiating media. The numerical results and measurements are compared in order to improve the analysis methods for Mars entry flows.
       
  • Aerodynamics and flight mechanics activities for a suborbital flight test
           of a deployable heat shield capsule
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Alberto Fedele, Stefano Mungiguerra MINI-IRENE is the Flight Demonstrator of IRENE, a new-concept capsule with a variable geometry, originally conceived by ASI to widen the range of available platforms to retrieve payloads and/or data from low Earth orbit. The main characteristics of IRENE is the “umbrella-like" deployable front structure that reduces the capsule ballistic coefficient, leading to acceptable heat fluxes, mechanical loads, stability and final descent velocity. Following the feasibility studies carried out since 2011, with also preliminary Thermal Protection System materials tests in plasma wind tunnels, the objective is now to design and build a Flight Demonstrator and a Ground Demonstrator to prove, with a suborbital flight and with a Plasma Wind Tunnel (PWT) test campaign, the functionality of the deployable heat shield. The Flight Demonstrator shall be included as a secondary payload in the interstage adapter of a VSB-30 launcher from ESRANGE, then ejected during the ascent phase of the payload section, perform a 15-min ballistic flight, re-enter the atmosphere and hit the ground. The Ground Demonstrator, representative of the Thermal Protection System of the Flight Demonstrator, shall be instead exposed to a heat flux similar to that expected for an atmospheric re-entry from low Earth orbit inside the SCIROCCO Plasma Wind Tunnel at CIRA. The paper, after a short description of the mission profile both for orbital and suborbital flights, focuses on the aerodynamics and flight mechanics activities held for the suborbital flight and PWT test campaigns.
       
  • Massive scale, long battery life, direct to orbit connectivity for the
           internet of things
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): David Haley, Andrew Beck, André Pollok, Alex Grant, Robby McKilliam Applications delivered by the internet of things have the potential to increase operational efficiency, reliability and safety. However, a challenge exists to deliver connectivity to industries with remote operations at a cost, battery life and form factor that is able to close the business case for deployment. This is especially true in cases where the system must scale to support large numbers of devices. Typical applications include sensor telemetry, low-value asset tracking, and device monitoring and control. Myriota provides global reach for the internet of things by securely delivering high-value small-data direct to a constellation of low Earth orbit satellites. This paper provides an overview of the Myriota communications architecture, and the process taken to transfer Myriota foundation technology into a highly scalable commercial product and service. Recent results from customer facing pilot deployments are also presented.
       
  • Hypersonic shock wave transitional boundary layer interactions - A review
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Doyle Knight, Mahsa Mortazavi Hypersonic shock wave transitional boundary layer interactions can result in significantly greater peak surface heat transfer than laminar or turbulent interactions. Consequently, the understanding of the flowfield structure of hypersonic shock wave transitional boundary layer interactions is important. Moreover, the capability to predict the mean and fluctuating aerothermodynamic loading due to such interactions is needed for effective design of hypersonic vehicles. A review of hypersonic shock wave transitional boundary layer interaction research since 1993 is presented. Significant progress has been achieved in the understanding of the flowfield structure. The most promising prediction methodology is Direct Numerical Simulation (DNS); however, DNS requires dynamic (i.e., time varying) inflow boundary conditions for five flow variables (i.e., three components of velocity, and two thermodynamic variables), and such experimental data is presently infeasible. Additional research is needed to understand the effect of assumed dynamic inflow boundary conditions on DNS prediction of aerothermodynamic loads.
       
  • Spacecraft angular velocity trajectory planning for SGCMG singularity
           avoidance
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Zhili Hou, Yunhai Geng, Baolin Wu, Simeng Huang A trajectory planning method for angular velocity of spacecraft is developed to avoid the impassable singular states for singular gimbal control moment gyroscope (SGCMG) systems in this paper. A new set of attitude parameters, named σ-parameters, is first developed. Based on the properties of σ-parameters, two approximate decoupled rotations are presented. To achieve a rapid attitude maneuver, both of the decoupled motions are designed as simple bang-off-bang type maneuvers. Then, a type of SGCMG singularity-free angular velocity trajectory on the conic surface is developed. Thereafter, an attitude controller based on σ-parameters is developed to track the reference trajectory. To avoid the impassable singular state, suitable axes of the approximate decoupled two rotations are chosen to achieve the fastest maneuver under the condition that the minimum distance from the angular momentum trajectory to the impassable surface is greater than a safety distance. Finally, simulations are performed to verify the effectiveness of the proposed SGCMG singularity avoidance method.
       
  • ALCIDES: A novel lunar mission concept study for the demonstration of
           enabling technologies in deep-space exploration and human-robots
           interaction
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Ghassabian G. Hady, Calzada Diaz Abigail, Hettrich Sebastian, De Quattro Nicola, Antonello Andrea, Bielicki Damian Returning to the Moon has kept gaining interest lately in the scientific community as a mandatory step for answering a cohort of key scientific questions.This paper presents a novel Lunar mission design to demonstrate enabling technologies for deep-space exploration, in accordance with the Global Exploration Roadmap and the National Research Council. This mission, named ALCIDES, takes advantage of some of the systems that are currently under development as a part of the HERACLES exploration architecture: these include the Orion module, the Space Exploration Vehicle, the Boeing Reusable Lander, the Ariane 6, the Falcon Heavy, the Space Launch System, as well as the Evolvable Deep-Space Habitat placed in EML2.A consistent part of the efforts in designing the ALCIDES mission accounts for innovative exploration scenarios: by analysing state of the art in robotics and planetary exploration, we introduce a mission architecture in which robots and humans collaborate to achieve several tasks, both autonomously and through cooperation.During this mission, high-performance mobility, extravehicular activity and habitation capabilities would be carried out and implemented. This project aims to demonstrate the human capability to live and work in the Lunar environment through the development of a long-term platform.We selected the Amundsen-Ganswindt basin as the landing site for multiple reasons: the possible presence of permanently shadowed regions, its position within the South Pole and its proximity to the Schrödinger basin. The main objectives of the ALCIDES mission are to study the Lunar cold trap volatiles, to gain understanding of the Lunar highlands geology through sampling and in-situ measurements and to study Human-Robotic interactions. In addition, factors such as psychology, legal issues and outreach regarding this mission were also considered.In particular, four traverses connecting the Amundsen crater with the Schrödinger basin were proposed, three of which to be performed by a tele-operated rover, and the remaining one to be carried out by a human crew with rover assistance. During these traverses, the rover will collect samples from several points of interest as well as perform in-situ measurements with a suite of instruments on board, helping to locate a convenient place for future human habitation.The ALCIDES mission results will help the scientific community to better understand the Moon and to take advantage of its resources for future space exploration. Gaining this knowledge will allow us to move forward in the development of systems and capabilities for manned missions to Mars and beyond.
       
  • Effects of a High Fidelity Filter on the attitude stabilization of a
           flexible spacecraft
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Marco Sabatini, Giovanni B. Palmerini, Matteo Ribet, Paolo Gasbarri, Luca Lampani The problem of stabilizing the time delayed control of a flexible space structure is analyzed in this paper. A free floating platform is used to investigate the space multibody dynamics and control. A first necessary step to develop stabilizing techniques is considered the availability of a set of measurements as complete as possible: in particular measurements of the elastic vibrations are necessary in addition to classic attitude measurements. At the scope, a net of PZT sensors have been designed and manufactured on a composite material panel, purposely built to resemble a space structure. A combined use of the PZT/optical sensor is proposed, where the role of the camera is to estimate the PZT parameters that can be changed after the manufacturing or for environmental aging. When this calibration process is performed, PZT can be used as standalone sensors for measuring also the elastic displacement of the structure. Once these measurements of attitude and elastic displacement are obtained, two stabilizing techniques have been developed, the Finite Spectrum Analysis, already known in literature, and the newly developed High Fidelity Filter approach, based on the design of a Kalman filter with large confidence on the process dynamics. It is shown that both techniques manage to increase the delay margin of the system, thus obtaining a stable maneuver, but the second approach reach this goal with very low residual vibrations and a remarkable fuel saving.
       
  • Ground-based experiments of tether deployment subject to an analytical
           control law
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): B.S. Yu, L.L. Geng, H. Wen, T. Chen, D.P. Jin Tethered satellite systems (TSSs) have shown great application potential in space missions, such as debris capture, active debris removal, and tether assisted observation. When the tether is deployed on-orbit, it may undergo a taut-slack process. This makes controlling a tether deployment more difficult than controlling a suspended tether. This paper examines a tether deployment subjected to an analytical control law in a ground-based experimental testbed. A dynamics similarity is proposed for the ground-based experiment to reproduce the dynamic environment of the tether deployment of the on-orbit TSS. Gravity compensation is used in the experiment to balance the friction forces and gravitation components that arise from the slight inclination of the testbed. The controlled stability is evaluated by the convergence of the pitch motion of the tether. The experimental results show that the controlled tether is successfully deployed along an assigned direction under a taut state during the deployment phase.
       
  • CubeSat constellation management using Ionic Liquid Electrospray
           Propulsion
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Marco Gomez Jenkins, David Krejci, Paulo Lozano The Space Propulsion Laboratory (SPL) of the Massachusetts Institute of Technology (MIT) is developing the Ion Electrospray Propulsion System (iEPS), designed to address a current need in CubeSat technology: miniaturized electric thrusters. These could be used for different applications, ranging from attitude control to interplanetary flights. In this work, performed together with the Space Systems Laboratory of the Costa Rica Institute of Technology (SETEC Lab), we explore a case study in which the iEPS is used for constellation management in Low Earth Orbit (LEO) when integrated in a 3U CubeSat. We analyze how a 180° separation in the Right Ascension of the Ascending Node (RAAN) between two CubeSats (SatA and SatB) starting in the same orbit can be achieved by modifying one of the spacecraft's orbital altitude, resulting in a difference in their rate of nodal precession (defined as the drift rate) due to the J2 effect, and therefore a difference in their relative RAAN. The method consists of SatB increasing its semi-major axis, drifting in a higher orbit with a lower drift rate, and returning to the original semi-major axis once the desired difference in RAAN in achieved relative to the other spacecraft. SatA will stay in its original orbit, using its thruster to compensate for orbital energy loss due to atmospheric drag, therefore demonstrating another application of iEPS for constellation management. Three different simulations were studied, defined as the minimum time trajectory, minimum propellant trajectory and a hybrid trajectory, consisting of reaching a higher altitude orbit, but actively changing the RAAN using the propulsion system instead of drifting. It was observed that the difference in this orbital element could be achieved using 85 g of propellant in as little as 164 days for the minimum time trajectory. The same difference could also be achieved using only 44 g of propellant in 245 days for the minimum propellant trajectory. Furthermore, the results of the hybrid trajectory showed that the goal could be achieved in 161 days, but using 158 g of propellant mass, demonstrating the benefit of using a drift orbit. The results proved the feasibility of implementing iEPS for constellation management using 3U CubeSats in LEO.
       
  • A lunar flyby for a tridimensional Earth-to-Earth mission
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Luiz Arthur Gagg Filho, Sandro da Silva Fernandes The present work formulates an orbital transfer for an Earth-to-Earth mission between non coplanar orbits with different altitudes with a special feature: the occurrence of a lunar flyby during the transfer orbit. This lunar flyby is intended to help change the plane of motion of the spacecraft without fuel consumption. Only two-impulsive trajectories are considered with the velocity increments applied at the initial and final orbits. In order to solve this problem, a 3D patched-conic approximation associated with a two-point boundary value problem is proposed. The same transfer problem is formulated considering the spatial circular restricted three-body problem (SCR3BP). The results of the patched-conic approximation is compared with the results of the SCR3BP showing a good agreement between the models. This work also determines several trajectories in order to perform a study of the fuel consumption considering several inclinations and altitudes of both initial and final orbits around the Earth. The longitude of the ascending node of the initial orbit, and, the altitude of close approach with the Moon during the flyby are also analyzed. According to the total velocity increment analysis, the changing plane assisted by a lunar flyby can be very favorable. Despite the increase of the time of flight, the saving of fuel is considerable. Indeed, the total velocity increment of this kind of maneuver is in some cases better than the velocity increment provided by the bi-parabolic transfer.
       
  • Hayabusa2-Ryugu proximity operation planning and landing site selection
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Tomohiro Yamaguchi, Takanao Saiki, Satoshi Tanaka, Yuto Takei, Tatsuaki Okada, Tadateru Takahashi, Yuichi Tsuda This paper presents the robust planning of the Hayabusa2-Ryugu proximity operation and landing site selection process considering unknown asteroid environment and the spacecraft constraints. The proximity operation scenario is described together with the relationship between the selection process and the in-situ observation. The mission constraints are summarized for the possible asteroid environment, including the rotation state, thermal condition and gravity.
       
  • Fault-tolerant attitude control of miniature satellites using reaction
           wheels
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Krishna Dev Kumar, Godard, Noel Abreu, Manoranjan Sinha An adaptive fault-tolerant nonlinear control scheme is proposed for precise 3-axis attitude tracking of miniature spacecraft in the presence of control input saturation, model uncertainties, external disturbances, and reaction wheel faults. Two configurations of reaction wheel assembly are examined in this paper, (A1) Traditional four wheel setup where three reaction wheels are in orthogonal configuration along with one oblique wheel; and (A2) Four wheels in a pyramid configuration. Multiplicative reaction wheel faults are considered along with complete failure of one wheel (A1) and two wheels (A2). The proposed control algorithm does not require an explicit fault detection and isolation mechanism and therefore failure time instants, patterns, and values of actuator failures remain unknown to the designer. The stability conditions for robustness against model uncertainties and external disturbances are derived using Lyapunov stability theory to establish the regions of asymptotic stabilization. The benefits of the proposed control methodology are analytically authenticated and also validated using hardware-in-the-loop simulations. The experimental results clearly establish the robustness of the proposed autonomous control algorithm for precise attitude tracking in the event of reaction wheel faults and failures.
       
  • Libration dynamics of electrodynamic tether system for 13 degrees
           International Geomagnetic Reference Field
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Yu-wei Yang, Hong Cai The libration dynamics of the electrodynamic tether system is studied for 13° International Geomagnetic Reference Field. Using the International Geomagnetic Reference Field including up to 13 t h order and 13° terms to describe the geomagnetic field, the attitude dynamic equations of the system in the elliptical orbits are built. The generalized forces produced by this magnetic model are derived. The generalized forces related to the in-plane and out-of-plane angles are sum of generalized forces for nontilted dipole model and generalized forces for higher order geomagnetic model terms. In the analysis of the libration dynamic characteristics, the generalized forces for higher order geomagnetic model terms are regarded as perturbations to the dynamic equations for the nontilted dipole model. The simulation results show that differences of components of these two geomagnetic model and differences of generalized forces related to them are all small. Failure time of the libration motion is defined to measure the influence of the perturbation to the system. Examples for different electrodynamic parameters and orbital parameters are simulated. The results show that the perturbations have obvious effects on the attitude dynamics. The influences of perturbations caused by higher order terms of 13° International Geomagnetic Reference Field for different parameters are all obtained.
       
  • Predictive visual servo kinematic control for autonomous robotic capture
           of non-cooperative space target
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Gangqi Dong, Zheng H. Zhu This paper presents a predictive visual servo kinematic control scheme for a robotic manipulator with the eye-in-hand configuration to perform autonomous capture of a non-cooperative space target with unknown motion. The unknown motion is estimated by an integrated algorithm of the photogrammetry and adaptive extended Kalman filter, in which the eye-in-hand configuration of the vision system improves the accuracy of the motion estimation as it approaches the target. Based on the vision feedback, a dynamic trajectory of robotic manipulator is planned in real time at each sampling instant and the end-effector of the manipulator moves towards the predicted position of the target at the next time instant incrementally. In this way, the multiple solutions problem of inverse kinematics in the joint space is effectively avoided and the robotic manipulator could intercept the non-cooperative target for a fast rendezvous. Validation experiments are performed on a custom built robotic manipulator with an eye-in-hand configuration. The experimental results demonstrate the effectiveness and robustness of the proposed control scheme.
       
  • Structural design and optimization of large cable–rib tension deployable
           antenna structure with dynamic constraint
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Ruiwei Liu, Hongwei Guo, Rongqiang Liu, Hongxiang Wang, Dewei Tang, Zongquan Deng The large deployable antenna has continuously received research interest in space technology. The design of such large structure has certain inherent challenges, such as limited mass and volume because of the inadequate capabilities of launchers. This constraint affects different aspects, including shapes, dimensions, and stiffness requirements. This study explores a new large cable–rib deployable antenna structure with radial ribs and tensioned cables. This structure has the advantages of high stiffness/mass ratio, which is suitable for constructing large-scale deployable antennas. A structural optimization method with a dynamic constraint for the maximum stiffness/mass ratio is proposed; this method is based on structural design formulas and the dynamic model of the deployable antenna structure. A genetic algorithm is introduced for parameter optimization with frequency constraint. Numerical examples are conducted to demonstrate the effectiveness of the proposed optimization method. By using these analysis methods, a 1.8 m prototype is fabricated and tested. Afterward, the feasibility and dynamic characteristics of the proposed cable–rib tension deployable structure are validated.
       
  • J 2 +perturbations&rft.title=Acta+Astronautica&rft.issn=0094-5765&rft.date=&rft.volume=">Non-iterative angles-only initial relative orbit determination with J 2
           perturbations
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): A.C. Perez, D.K. Geller, T.A. Lovell An approximate solution to the angles-only initial relative orbit determination problem is developed and evaluated in the context of the two-body problem with J2 perturbations. The algorithm is non-iterative and requires the singular value decomposition of a 6 × 6 matrix and the solution of a fifteenth-order polynomial. The performance of the algorithm is investigated for non-circular low-Earth orbits and near-circular geostationary orbits with and without measurement error.
       
  • Experimental investigations on ethylene-air Continuous Rotating Detonation
           wave in the hollow chamber with Laval nozzle
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Haoyang Peng, Weidong Liu, Shijie Liu, Hailong Zhang The ethylene-air Continuous Rotating Detonation (CRD) has been experimentally achieved with large operating domain, little velocity deficit and notable pressure rise in the hollow chamber with Laval nozzle. The results show that the lean limit increases while operating domain decreases with contraction ratio increasing. Deflagration flame in recirculation zone and larger width combustor enable the ethylene-air CRD to be readily achieved. Three different propagation modes are presented. Most of the achieved CRD experiments are single-wave mode. The highest frequency and velocity are 6.10 kHz and 1915.40 m/s respectively. Two-waves mode can be obtained when contraction ratio is 12. Sawtooth wave mode appears around lean limit when contraction ratio is 1,2 and 4. Sawtooth wave, as a critical condition, can be transformed into typical CRD wave or extinguish. For contraction ratios of 1,2,4 and 6, the propagation stability increases with equivalence ratio (ER) increasing. For contraction ratios of 8,10 and 12, the stability decreases with a concomitant increase of ER. The contraction ratios of 2 and 4 are beneficial for CRD wave to propagate with high frequency and stability. The study will deepen the understanding of ethylene-air CRD and enrich the combustor design theory of CRD Engine fueled by hydrocarbon fuels.
       
  • Momentum enhancement factor estimation for asteroid redirect missions
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Simon Delchambre, Tobias Ziegler, Albert Falke, Klaus Janschek An asteroid mitigation demonstration mission is gaining interest among the planetary defense community to better understand the challenges and the dynamics of a small solar system body (SB) impact scenario. The Kinetic Impactor (KI) deflection technique, considered the most mature and cost effective approach for deflecting SBs, gained credibility following both the numerous studies performed (Don Quijote, NEOShield-2, preparations for DART mission, …) as well as the successful targeting of the Deep Impact (DI) spacecraft (S/C) into comet 9 P/Tempel 1. A dual-satellite concept AIDA with KI (DART) and an Explorer S/C (AIM/HERA) is currently under study by the ESA and NASA. While one of the more mature deflection options, there are still a significant number of poorly constrained aspects of the KI deflection technique. Of particular interest are the complex ejecta cloud dynamics that can have a considerable impact on the deflection efficiency and the according β-factor. Understanding the momentum enhancement β-factor is considered paramount as it bears the potential of overall mission cost reduction and is inherently linked to the SB geotechnical properties. Therefore, estimating this β-factor is one of the top-level scientific requirements for future demonstration missions. First, this work presents a β-factor estimation technique with the focus on an SB orbit determination (OD) filter where radioscience tracking data of an Explorer S/C at the close proximity is fused with optical navigation information. Second, an extensive error analysis is presented where the major drivers of the β-factor error budget are identified based on a breakdown tree. The paper shows the estimation filter architecture and explicitly addresses the data fusion process. An extensive, high fidelity test campaign has been conducted to conclude on the achievable β-factor estimation performance for a KI impactor reference scenario with the SB 2001 QC34. An end-to-end momentum enhancement factor estimation technique is presented and it was found that the β-factor uncertainty is reduced to 0.33 (3σ) after only 1 week of monitoring with 67% availability of the tracking stations and a station-keeping manoeuver once a day. This estimation performance has shown that the momentum enhancement factor uncertainties can be constrained considerably and thus further advocates a KI demonstration mission.
       
  • Ignition mechanism in ablative pulsed plasma thrusters with coaxial
           semiconductor spark plugs
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Guorui Sun, Zhiwen Wu Ignition process, as the initiation of the entire discharge, plays an important role in ablative pulsed plasma thrusters. While spark plug exactly how initiate discharge is achieved is still under review. This study did some experiments with two kinds of propellant surfaces (normal or inclined) and without propellant to explain the ignition process. The experimental results showed: when the thruster discharge without propellant, it is essentially a surface flashover process on ceramics; when the propellant was loaded, the main discharge occurs after the initial conductive path composed of electrons emitted by spark plug forming. This study provides a reference for the high performance pulsed plasma thrusters.
       
  • Aging constitutive model of hydroxyl-terminated polybutadiene coating in
           solid rocket motor
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Ke Li, Jian Zheng, Jianzhuang Zhi, Kailun Zhang To accurately describe the tensile mechanical properties of aged hydroxyl-terminated polybutadiene (HTPB) coating, the aging constitutive model was studied. The single-step and multi-step relaxation tests were performed on the unaged samples, and the tensile mechanical properties of the aged HTPB coating were tested, while the crosslink density was obtained by nuclear magnetic resonance (NMR) experiments. The model parameters were solved using the experimental data. The crosslink density was used to characterize the aging degree of the HTPB coating, and combined with the modified Arrhenius equation, a model of crosslink density variation with aging time was built. Multiply the hyper-elastic model with the aging characteristic function, an aging constitutive model of HTPB coating was established, which can be used to describe the tensile mechanical properties of aging HTPB coating. The verification tests show that the predicted value of the crosslink density under the test of 313.15 K is in good agreement with the test value. The aging constitutive model can predict the tensile mechanical behavior of HTPB coating well, which is of important engineering significance.
       
  • SpooQySats: CubeSats to demonstrate quantum key distribution technologies
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): James A. Grieve, Robert Bedington, Zhongkan Tang, Rakhitha C.M.R.B. Chandrasekara, Alexander Ling Satellite-based quantum key distribution (QKD) offers the potential to share highly secure encryption keys between optical ground stations all over the planet. SpooQySats is a programme for establishing the space worthiness of highly-miniaturized, polarization entangled, photon pair sources using CubeSat nanosatellites. The sources are being developed iteratively with an early version in orbit already and improved versions soon to be launched. Once fully developed, the photon pair sources can be deployed on more advanced satellites that are equipped with optical links. These can allow for very secure uplinks and downlinks and can be used to establish a global space-based quantum key distribution network. This would enable highly secure symmetric encryption keys to be shared between optical ground stations all over the planet.
       
  • Uncertain surface accuracy evaluation based on non-probabilistic approach
           for large spacecraft
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Chen Yang, Xinbin Hou, Li Wang The influence of the surface shape state upon the performance and function of large spacecraft should be considered in design and analyzed in the extreme space environment. Therefore, in this paper, considering the measurement uncertainties and errors exist in large high-precision spacecraft, an interval surface accuracy evaluation method is proposed based on a non-probabilistic approach. To overcome the limitations of insufficient statistical quantification of uncertain parameters, this paper treats uncertainties as non-probabilistic intervals. The conventional root mean square index is extended to uncertain interval numbers, which can be used to evaluate the surface accuracy with the measurement uncertainties and errors. Moreover, to improve the interval expansion problem, subinterval technology is applied to the uncertainty propagation process for surface accuracy evaluation. As long as the bounds of the uncertainties and errors are known, the interval bound for uncertain surface accuracy can be estimated conveniently by interval analysis. Finally, three engineering examples are separately proposed to evaluate the interval surface accuracy thereby validating the effectiveness and veracity of the proposed method. The result obtained in this paper can be regarded as an interval estimator, offering more detailed evaluations and suggestions for large spacecraft design and analysis than deterministic methods.
       
  • Historical-orbital-data-based method for monitoring the operational status
           of satellites in low Earth orbit
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Tao Li, Lei Chen A method is developed for monitoring the operational status of satellites in low Earth orbit (LEO) based on their historical two-line element (TLE) data. In this method, whether the satellite still has the maneuverability to maintain the orbital altitude is used as the criteria for judging the operational status, the whole judgment process includes two steps. The first step is to design a specialized algorithm to detect orbit maintenance maneuvers from the satellite's TLE time-history. The algorithm uses abnormal data segments of the TLE derived semi-major axis time series to identify the orbit maintenance maneuver, and various measures are taken to eliminate the noise interference and to ensure the detection accuracy. The second step is to use the detected maneuvering history to determine the current operational status of the satellite. In this step, the statistical technique is used to get the temporal regularity of the satellite to implement orbit maintenance maneuvers and the allowable range of the natural variation of the semi-major axis, so then the criteria for determining the satellite operational status is developed. Analysis of typical LEO satellites indicates that this method can accurately determine the current operational status of the satellite and provide an approximate estimation interval of the satellite retiring time, which is of practical value.
       
  • Multi-objective integrated robust H∞ control for attitude tracking
           of a flexible spacecraft
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Shunan Wu, Weimeng Chu, Xue Ma, Gianmarco Radice, Zhigang Wu This paper investigates the multi-objective attitude tracking problem of a flexible spacecraft in the presence of disturbances, parameter uncertainties and imprecise collocation of sensors and actuators. An integrated robust H∞ controller, including an output feedback component and a feedforward component, is proposed, and its gains are calculated by solving Linear Matrix Inequalities. The output feedback component stabilizes the integrated control system while the feedforward component can drive the attitude motion to track the desired angles. The system robustness against disturbances, parameter uncertainties and imprecise collocation is addressed by the H∞ approach and convex optimization. Numerical simulations are finally provided to assess the performance of the proposed controller.
       
  • Analysis of lithium-combustion power systems for extreme environment
           spacecraft
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Christopher J. Greer, Michael V. Paul, Alexander S. Rattner The longest duration mission on the Venus surface was Venera 13 at just over 2 h. This time constraint was due to limited battery power life and craft thermal management challenges. A lithium combustion based power system has been proposed to increase landed mission durations for Venus and other extreme environment targets. This paper presents a new detailed thermodynamic and heat transfer model of a conceptual lithium combustion power system. Findings are applied to specify engineering requirements for potential missions. Results indicate that a lithium combustion power system using the in-situ carbon dioxide atmosphere as an oxidizer could power a Venus lander for up five days (24 h, Earth day) with 185 kg of fuel, delivering 14 kWth thermal energy continuously. Even greater durations are possible if lower power missions are considered. The potential performances of a Li-CO2 powered Stirling engine and sulfur-sodium batteries were compared. It was found that sulfur-sodium batteries would require about 1.75–2.5 times more mass to provide 1 kW of power output for mission durations of five to ten days, respectively. A lithium combustion power system with a sulfur-hexafluoride oxidizer could power a Europa lander at 94W with a Stirling engine for up to twenty days with 43 kg of reactants mass. Lithium-combustion activated Stirling engines and TEG arrays were compared with batteries to meet this power and mission duration requirement. It was found that batteries would require less mass than either lithium-fueled system. However, for mission durations longer than twenty-six days the Stirling engine power system may require less total mass than batteries. Future work will include laboratory-based experimental studies to validate results and improve heat transfer closure models.
       
  • Benchmarking information carriers
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Michael Hippke The search for extraterrestrial communication has mainly focused on microwave photons since the 1950s. We compare other high speed information carriers to photons, such as electrons, protons, and neutrinos, gravitational waves, inscribed matter, and artificial megastructures such as occulters. The performance card includes the speed of exchange, information per energy and machine sizes, lensing performance, cost, and complexity. In fast point-to-point communications, photons are superior to other carriers by orders of magnitude. Sending probes with inscribed matter requires less energy, but has higher latency. For isotropic beacons with low data rates, our current technological level is insufficient to determine the best choice. We discuss cases where our initial assumptions do not apply, and describe the required properties of hypothetical particles to win over photons.
       
  • Lessons learned in 20 years of application of Systems Concurrent
           Engineering to space products
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): G. Loureiro, W.F. Panades, A. Silva This paper aims to present the lessons learned in 20 years of application of the SCE (Systems Concurrent Engineering) approach that evolved over the last 20 years being applied to the development of more than 200 complex system solutions. SCE is an approach to the integrated development of complex systems that applies the systems engineering process, simultaneously, to the product elements of a system solution as well as for the service elements of the system solution, recursively, at every layer of the system solution breakdown structure. The approach was born as the application of the requirements, functional and physical analysis processes to the simultaneous development of a product, its life cycle processes and their performing organizations, at every layer of the product breakdown structure. The continuous application of the approach up to 2010, showed the need to include a stakeholder analysis step, to acknowledge that the solution was comprised of product and organization elements (processes were, in fact, the functions of products and organizations), that a mission layer should be added at the top of the product breakdown structure and that the notion of circumstances should be added to the traditional notion of scenarios. With the increasing use of the approach for system of systems conception and development such as those involving multi-spacecraft solutions, the mission layer needed to be extended to include other life cycle processes (besides the operations processes) concept of service and system service architecture. This requires the development of a system solution breakdown structure that will guide the development of the overall solution. For multi-spacecraft solutions, for example, it is necessary to conceive and architect testing, launching and decommissioning services as early as operations. Also, going into more detail in the approach, modes can be derived from circumstances, interface states and internal states of the system and not only from circumstances, as initially established in the approach. These lessons to be presented were learned during the development of: 1) the Brazilian Strategic Program for Space Systems (PESE) and; 2) the TIM Project (Telematics International Mission), a satellite formation with contributions from many regions in the world.
       
  • Experimental study of near-blowoff characteristics in a cavity-based
           supersonic combustor
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Xiliang Song, Hongbo Wang, Mingbo Sun, Yanxiang Zhang Lean blowoff characteristics of an ethylene-fueled model scramjet combustor with cavity flameholder are investigated under the inflow conditions of Ma = 2.52 and T0 = 1600 K. It is observed that, lean blowoff limits increase with increasing injection distance and which for the single-orifice cases are found to be higher than those for the multiple-orifice cases. For the multiple-orifice cases studied, once the flame is ignited, it can always be stabilized by the cavity as long as the fuel supply is constant. For the single-orifice cases, however, the flame can be extinguished intermittently even if the fuel is served continuously. That is, the lean flames are more stable for the multiple-orifice cases. Near-blowoff dynamics are then analyzed for the less unstable single-orifice cases. When the lean blowoff limits are approached, the cavity flames become less and less stable and may be partially extinguished. Nevertheless, the residual flame within the cavity may reignite the combustible mixture outside the cavity and the entire flame may restabilize. When the equivalence ratio is further decreased, ultimate blowoff takes place and is found to occur in multiple steps - the shear-layer flame becomes weaker, the flame is partially extinguished near the trailing edge, the flame shrinks into the latter part of the cavity, the flame moves towards the cavity front wall and is subsequently extinguished completely.
       
  • Benchmarking inscribed matter probes
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Michael Hippke, Paul Leyland, John G. Learned We have explored the optimal frequency of interstellar photon communications and benchmarked other particles as information carriers in previous papers of this series. We now compare the latency and bandwidth of sending probes with inscribed matter. Durability requirements such as shields against dust and radiation, as well as data duplication, add negligible weight overhead at velocities v
       
  • Angular momentum management strategy of the FengYun-4 meteorological
           satellite
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Yiwu Liu, Zhuhua Si, Liang Tang, Shoulei Chen Solar radiation torque has a great impact on Chinese new generation geostationary meteorological satellite FengYun-4 since its single winged solar array configuration. Daily accumulation in angular momentum stored in reaction wheels may increase about 30Nms, and should be unloaded via bipropellant thrusters during a daily 15-min housekeeping period. How to prevent zero-crossing and saturation of wheels, how to reduce the operation speed range of wheels, and how to ensure the service continuity if any one of wheels fails are the problems to be faced. In addition, both the uncertainty of thruster torque and the variability of solar radiation during all the life span may go against with the automaticity of angular momentum management. In this paper, a null-space-based momentum management method is presented, upon which the angular momentum is managed by null motion control while service and dumped during the housekeeping period. In-flight and numerical simulation results demonstrate the reliability and validity of this strategy.
       
  • Performance evaluation and comparison of electricity generation systems
           based on single- and two-stage thermoelectric generator for hypersonic
           vehicles
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Kunlin Cheng, Duo Zhang, Jiang Qin, Silong Zhang, Wen Bao Thermoelectric generator (TEG) is a promising electricity generation technology distinguished by a direct thermoelectric conversion. The single- and two-stage TEG model, the heat source of which was the combustion heat dissipation, were developed to predict and compare the power generation performance on hypersonic vehicles at different inlet temperatures of heating channel, Tfh0. The distributions of the temperature and thermoelectric figure of merit (ZT value) were described by diagrams. Besides, some methods for performance enhancement were discussed. The results indicate that the single-stage TEG has an advantage of the maximum power density, and the two-stage TEG shows a higher conversion efficiency at the same Tfh0. The maximum power density of 16.53 kW/m2 is achieved by the single-stage thermoelectric generator. The optimal conversion efficiency is 10.78%, obtained by the two-stage TEG. Both the maximum power density and corresponding conversion efficiency increase with the inlet temperature of heating channel. In addition, the two-stage TEG has a greater potential for improving performance, by means of multiple thermoelectric materials in their optimal temperature ranges.
       
  • Experimental study on a rotating detonation combustor with an axial-flow
           turbine
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Shengbing Zhou, Hu Ma, Yuan Ma, Changsheng Zhou, Daokun Liu, Shuai Li The research on rotating detonation turbine engine is attracting much attention in recent years. In this study, experiments have been performed on a structure combining a rotating detonation combustor and an axial-flow turbine to investigate the propagation characteristics of the hydrogen-air rotating detonation wave. The stable rotating detonation wave is successfully initiated using the spark plug and pre-detonator, and there is still a velocity deficit of about 20% relative to the Chapmane-Jouguet value. There is a formation process for the stable detonation wave, and the formation time for the pre-detonator is far less than the spark plug, however the final state is independent on the ignition device. The rotating detonation wave successively appears the two-wave state with a same direction, the two-peak wave state, and the state of strong–weak alternation during the formation process. Finally, only one stable detonation wave is formed in the chamber and propagates until the operation off.
       
  • Comparison of the space bubble detector response to space-like neutron
           spectra and high energy protons
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Alexander Miller, Rachid Machrafi, Eric Benton, Hisashi Kitamura, Satoshi Kodaira To compare the response to high energy neutrons and protons of the space bubble detectors in use aboard the International Space Station (ISS), three series of experiments were conducted with high energy protons and neutrons. The first series of experiments was conducted with high-energy neutrons in the energy range expected for neutrons encountered during space flight (0.6–800 MeV) at the Los Alamos Neutron Science Center (LANSCE) using the spallation neutron source. The second series was conducted with high energy protons from 30 to 70 MeV using the cyclotron at the National Institute of Radiological Science NIRS in Japan, and the third series of experiments was performed with high energy protons from 60 to 230 MeV at the ProCure proton therapy facility, Oklahoma, USA. The bubble detectors were exposed to different fluences in different experiments and the number of bubbles was counted using a bubble detector reader. The proton response of the bubble detector (sensitivity), as a function of energy, was determined and compared to the neutron sensitivity. In addition, to adjust the neutron sensitivity of the bubble detector determined in an AmBe field, a calibration factor was obtained for space applications.
       
  • Parameterization and optimization for shape-transition curved isolator
    • Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Zewei Meng, Xiaoqiang Fan, Yi Wang, Bing Xiong In order to optimize total pressure recovery performance of a type of variable cross-section curved isolator, it is designed and parameterized by mathematical methods. Blending functions are utilized to morph cross-sections from entrance to exit and B-spline curves are used to control cross-section translation to meet offset requirement. Evolutionary algorithm (multi-island genetic algorithm) is introduced to search the optimum individual for the target of total pressure recovery coefficient based on numerical calculation results under no backpressure conditions. Firstly, to ensure accuracy and feasibility of the calculation method, it is validated by comparing with the wind tunnel experiment results. Then, the three typical curved isolators, including rectangular-to-circular isolator, circular isolator and rectangular isolator, are chosen to study. Finally, the optimized configuration performances are analyzed under both no backpressure and variable backpressure conditions. The result shows that the performances of optimal isolators are well in both states. In the no backpressure state, the extra total pressure loss is mainly determined by wetted area of the configuration when the offset line is optimized to minimize the total pressure loss. In the backpressure state, the separation mode switch induced by the changes of the backpressure condition is also observed in curved isolators. What's more, the withstanding backpressure ability of the optimized rectangular-to-circular isolator is best based on analyzing the leading edge position of shock trains. And this optimization method can be also applied to studying other variable cross-section curved isolators.
       
  • A maximal-reward preliminary planning for multi-debris active removal
           mission in LEO with a greedy heuristic method
    • Abstract: Publication date: August 2018Source: Acta Astronautica, Volume 149Author(s): Jianan Yang, Yu Hen Hu, Yong Liu, Quan Pan An active multi-debris removal mission planning algorithm for low Earth orbit (LEO) debris removal is presented. We assume that a space platform carrying multiple debris-deorbiting nano-satellites will be launched to carry out the debris-removal mission. This platform travels from one debris object to the next using a drift-orbit-transfer strategy. Each debris is assigned a reward value reflecting potential collision risk reductions if the debris is deorbited. The objective of mission planning is to determine an ordering of debris to be deorbited that maximizes the total reward due to debris removal subject to the constraints of (a) the total amount of velocity change for orbit transfer, (b) the total duration of the debris removal mission, and (c) the number of deorbiting nano-satellites carried on board the platform. In this work, a new approach is taken to approximate the response curve of drift-orbit transfers between each debris pair using a numerically interpolated continuous curve. An optimal orbit transfer strategy then is adopted to minimizes the sum of relative Δv consumption and relative transfer duration. The task of multi-debris removal mission planning is formulated as a combinatorial tree search discrete optimization problem using this optimal orbit transfer cost estimate. A greedy heuristic with polynomial time complexity is proposed to select the next debris to deorbit that maximizes the reward-to-cost ratio. The effectiveness of this novel approach is demonstrated using real-world Iridium 33 debris cloud data. It is observed that this algorithm delivers superior performance while consuming a small fraction of computing time compared to state of the art mission planning algorithms.
       
  • Being a father during the space career: Retired cosmonauts' involvement
    • Abstract: Publication date: August 2018Source: Acta Astronautica, Volume 149Author(s): Phyllis J. Johnson, Peter Suedfeld, Vadim I. Gushin The space career requires numerous absences from the cosmonaut's family during training and spaceflight. Such absences mean missing important milestones, events, and celebrations in the lives of their children. This study assesses retired cosmonauts' views of actual and desired involvement with their children during their spaceflight career. The Father Involvement Scale (adapted from Finley & Schwartz, 2004; Hawkins et al., 2002), translated into Russian, was answered by 17 retired cosmonauts. The 20 domains in the scale included 10 Expressive (e.g., intellectual, emotional, social, and spiritual development; sharing activities and interests) and 10 Instrumental (e.g., providing income, being protective, discipline, school/homework, and developing responsibility, independence, and competence). The cosmonauts' ratings of actual involvement with their children's lives was between Sometimes involved and Often involved (M = 3.66, SD = 0.42). None of the cosmonauts indicated Never involved for any of the Expressive domains or for seven of the ten Instrumental domains. Within the Expressive domains, the majority of cosmonauts said they were “often” involved in their child's spiritual development and in sharing activities/interests. Within the Instrumental domains, they were “often” involved in discipline and “always” involved in providing income. The areas in which they wished they had been “much more involved” than they had been were Expressive, rather than Instrumental: intellectual, spiritual, and physical development; sharing activities/interests, and companionship. This is the first study to measure retrospective assessments of father involvement during spaceflight careers. Space agencies should consider how Family Support personnel can enhance the parental involvement of future spacefarers.
       
  • Onset of detonation in hydrogen-air mixtures due to shock wave reflection
           inside a combustion chamber
    • Abstract: Publication date: August 2018Source: Acta Astronautica, Volume 149Author(s): N.N. Smirnov, O.G. Penyazkov, K.L. Sevrouk, V.F. Nikitin, L.I. Stamov, V.V. Tyurenkova Hydrogen fuel, as for now, is the most energy efficient and clean fuel for aviation engines. At present time RAM engines based on chemical burning of fuels and accelerating exhaust gases practically reached the top of their efficiency in terms of specific impulse. New schemes are under development. As it is known, complex geometry of combustion chambers could bring to compression waves focusing and changing the mode of flame propagation from slow combustion to detonation. Keeping this process under control is of major importance for developing new generation of engines. The paper presents results of numerical and experimental investigation of mixture ignition and detonation onset in shock wave reflected from inside a wedge. Comparison of numerical and experimental results made it possible to validate the developed 3-D transient mathematical model of chemically reacting gas mixture flows incorporating hydrogen – air mixtures. Kinetic schemes were improved based on comparison of numerical and experimental results. A tool was developed making it possible to set a criterion for obtaining either safe combustion regime, or transition to detonation regime in shock wave focusing.
       
  • Data authentication, integrity and confidentiality mechanisms for
           federated satellite systems
    • Abstract: Publication date: August 2018Source: Acta Astronautica, Volume 149Author(s): Olga von Maurich, Alessandro Golkar This work addresses a critical topic in federated satellites development: the lack of trust between stakeholders that would prevent any stakeholder joining a satellite federation owned and operated by multiple parties. A characterisation of security needs for federated satellite systems is proposed, showing that in order for a federation to offer an environment for a beneficial cooperation, a notion of identity, both user identity and data authentication, has to be introduced, and stakeholders' security requirements have to be satisfied. This paper presents a public key infrastructure (PKI) based protocol for addressing stakeholders' security requirements and ensuring data authentication, integrity and confidentiality in data transfer operations within satellite federations. The performance and cost overheads of the proposed security protocol are first characterised with an experimental implementation on a Raspberry Pi 2 platform, used as a representative proxy testbed of commercial off-the-shelf avionics for small satellites, and then with a benchmark on a range of CPUs to analyse which platforms achieve set performance goals with radio-based and laser-based communications. Recommendations for implementing security mechanisms in federated satellite systems are thus derived.
       
  • Exogeoconservation: Protecting geological heritage on celestial bodies
    • Abstract: Publication date: August 2018Source: Acta Astronautica, Volume 149Author(s): Jack J. Matthews, Sean McMahon Geoconservation is an increasingly widely adopted theoretical, practical and administrative approach to the protection of geological and geomorphological features of special scientific, functional, historic, cultural, aesthetic, or ecological value. Protected sites on Earth include natural rocky outcrops, shorelines, river banks, and landscapes, as well as human-made structures such as road cuts and quarries exposing geological phenomena. However, geoconservation has rarely been discussed in the context of other rocky and icy planets, rings, moons, dwarf planets, asteroids, or comets, which present extraordinarily diverse, beautiful, and culturally, historically and scientifically important geological phenomena. Here we propose to adapt geoconservation strategies for protecting the geological heritage of these celestial bodies, and introduce the term ‘exogeoconservation’ and other associated terms for this purpose. We argue that exogeoconservation is acutely necessary for the scientific exploration and responsible stewardship of celestial bodies, and suggest how this might be achieved and managed by means of international protocols. We stress that such protocols must be sensitive to the needs of scientific, industrial, and other human activities, and not unduly prohibitive. However, with space exploration and exploitation likely to accelerate in coming decades, it is increasingly important that an internationally agreed, holistic framework be developed for the protection of our common ‘exogeoheritage’.
       
  • Natural formation flying on quasi-halo orbits in the photogravitational
           circular restricted three-body problem
    • Abstract: Publication date: August 2018Source: Acta Astronautica, Volume 149Author(s): Tong Luo, Ming Xu, Yunfeng Dong Formation flying near a libration point is an effective method of the multi-point in-situ detection of deep-space environment and the high resolution observation of astronomy etc. This paper proposes a novel numerical searching method based on ergodic Poincaré mappings to seek natural formation flying on quasi-halo orbits in a photogravitational circular restricted three-body problem (PCR3BP). The reduced Hamiltonian system using the Lie series method gives a qualitative description of the phase space near the libration points. Then, an algorithm to return numerical solutions of invariant tori back to the synodic reference coordinate is presented. This research proposes a new type of Poincaré mapping from center manifolds of quasi-halo orbits onto the defined characteristic indexes in formation flying. Appropriate orbit groups, including two quasi-halo orbits, a halo orbit and a quasi-halo orbit, or multiple quasi-halo orbits, are all feasible for formation flying through numerically ergodic searching. Finally, the relative trajectory is considered unstable and a control scheme is required in practical formation flying.
       
  • Experimental investigation on the surface wave characteristics of conical
           liquid film
    • Abstract: Publication date: August 2018Source: Acta Astronautica, Volume 149Author(s): Zhongtao Kang, Xiangdong Li, Xiongbing Mao The disintegration of the liquid sheet is due to the growth of the unstable surface wave. Research on the surface wave characteristics mainly uses the linear stability analysis, and the surface wave characteristics, for example the frequency and wave length of the surface wave, have not been analyzed well experimentally. In the present study, the proper orthogonal decomposition is adopted to analyze the instantaneous spray images and the extracted surface waves on the interface of the gas and liquid. The wavelength and frequency of the surface wave are obtained, and the influence of the injector mass flow rate is investigated. The results show that the spray pattern moves from onion stage to tulip stage and fully developed wavy cone stage with the increase of liquid mass flow rate. In the tulip stage, the wave frequency decreases slightly and then increases with the increase of liquid mass flow rate. On the contrary, the wavelength increases gradually with the increase of liquid mass flow rate and then decreases slightly. This variation trend is caused by the competition effect of the film thickness and the axial film velocity. The decrease of film thickness decreases the wave frequency and increases the wavelength, while the increase of the axial film velocity increases the wave frequency and decreases the wavelength. In the fully developed wavy cone stage, the wave frequency increases gradually with the increase of liquid mass flow rate. And the bandwidth of the surface wave frequency increases gradually, which means that the leading role of the dominant surface wave is weakened and the conical film is dominated by surface waves with multiple frequencies. The wavelength decreases with the increase of liquid mass flow rate because of the significant increased axial film velocity.
       
  • Continuous discharge in micro ablative pulsed plasma thrusters
    • Abstract: Publication date: August 2018Source: Acta Astronautica, Volume 149Author(s): Zhiwen Wu, Guorui Sun The mechanism of continuous discharge in micro ablative pulsed plasma thrusters has always been ignored. A series of experiments on discharge completion rate were conducted, and theoretical analyses were performed to understand the results. These results showed that as initial voltage decreased to a certain value, a portion of energy was not released in a short time, which caused overall efficiency to greatly decrease. As the initial voltage decreased sequentially, this phenomenon occurred more frequently. This study provides a reference for high-frequency micro ablative pulsed plasma thrusters.
       
  • Numerical and experimental investigation on the influence of inlet
           contraction ratio for a rocket-based combined cycle engine
    • Abstract: Publication date: August 2018Source: Acta Astronautica, Volume 149Author(s): Zhengze Zhang, Peijin Liu, Fei Qin, Lei Shi, Yajun Wang, Chao Huo The influences of inlet contraction ratio on both the inlet and combustor of the rocket-based combined cycle engine are investigated in this paper by experimental and numerical approaches. A translating spike was employed to simulate the inlet counter pressure in the wind tunnel experiment. Steady-state pressure distributions were recorded and schlieren photos were obtained simultaneously. The combined numerical-experimental results show that the increased inlet contraction ratio always results in the increasing of the compression ratio and capability of anti-counter pressure, while its impact on the mass flow coefficient is negligible. Besides, the reactive numerical results show that the coefficients of inlet drag and combustor thrust increase with the increasing of the inlet contraction ratio, while the nozzle thrust coefficient demonstrates an opposite trend. These suggest that the selection of an appropriate inlet contraction ratio can effectively improve the total thrust coefficient of the engine, while the appropriate inlet contraction ratio is to be set to the minimum that satisfies the robust combustion requirement.
       
  • Lunar orbit dynamics and maneuvers for Lunisat missions
    • Abstract: Publication date: August 2018Source: Acta Astronautica, Volume 149Author(s): Mauro Pontani, Riccardo Di Roberto, Filippo Graziani Lunisat represents a next-generation microsatellite aimed at orbiting the Moon, and equipped with dispensers for the release of nanosatellites. This research is focused on the orbital dynamics of both the main microsatellite and the nanosatellites after release. Due to the irregular concentrations of mass in the Moon, low altitude, near-circular lunar orbits are affected by a considerable number of harmonics of the Moon gravitational field. Nonsingular equinoctial orbit elements are employed for orbit propagations, in conjunction with numerical averaging, which is a numerical technique that allows substantial computational improvements. The dynamical model considers a large number of harmonics of the lunar gravitational field, as well as the Earth and Sun perturbing influence as third bodies. Low-altitude lunar satellites turn out to impact the lunar surface after a few weeks or months. In case of an unsatisfactory lifetime, two simple orbit maintenance strategies are evaluated, together with the related propellant budget. Nanosatellites will be released from Lunisat by means of springs. The orbit of each nanosatellite depends on the release conditions, i.e. the relative velocity magnitude and direction. This work investigates the nanosatellite lifetimes as functions of these conditions. Lastly, minimum-time low-thrust transfers for reducing the orbit altitude are investigated, both for nanosatellite release and for the conclusive phase of the Lunisat mission, which finally terminates with the impact on the lunar surface.
       
  • IFC - Publication Information
    • Abstract: Publication date: August 2018Source: Acta Astronautica, Volume 149Author(s):
       
  • A high-accuracy constrained SINS/CNS tight integrated navigation for
           high-orbit automated transfer vehicles
    • Abstract: Publication date: Available online 11 July 2018Source: Acta AstronauticaAuthor(s): Wang Dingjie, Lv Hanfeng, An Xueying, Wu Jie High-accuracy and reliable autonomous navigation is increasingly crucial for automated transfer vehicles (ATV). This paper proposes a novel strapdown inertial navigation system/celestial navigation system (SINS/CNS) tight integration scheme aided by dynamic model constraints for high-orbit ATV to realize accurate and autonomous navigation. In this scheme, the complete weightlessness constraint in orbit is used to address the divergence of position and velocity caused by inaccurate accelerometer bias estimation problem encountered in the traditional SINS/CNS integration method, and the image point position-based tight integration model is derived to handle the adverse influence of time-varying attitude measurement noise due to changes of star geometry observed by a large-view-filed star sensor. Moreover, an information filter is devised to fuse the multi-rate measurements. The proposed algorithm is evaluated by a representative high-orbit ATV trajectory simulation, which indicates significant improvements in navigation accuracy compared with its traditional counterparts. The proposed algorithm can realize navigation accuracy enhancements without introducing additional sensors, strengthening its potentials in engineering application.
       
  • Investigation on dynamic behaviors of thermal protection system using a
           two degree-of-freedom nonlinear theoretical method
    • Abstract: Publication date: Available online 11 July 2018Source: Acta AstronauticaAuthor(s): Jie Huang, Weixing Yao, Piao Li, Danfa Zhou, Cheng Chang, Hanyu Lin In order to study the nonlinear dynamic behaviors of Thermal Protection System (TPS) and the nonlinear dynamic strength of the strain-isolation-pad (SIP), a two degree-of-freedom nonlinear dynamic theoretical model was presented under the acoustic excitation and base excitation. The tile is simplified as a mass point, a linear spring and a damping element, and the SIP is simplified as a mass point, a nonlinear spring and a damping element. On this basis, the solving process of the nonlinear theoretical model and the iterative process of the equivalent linear stiffness coefficient of SIP were derived by the statistical linearization method. The dynamic responses analyzed by the nonlinear theoretical model and linear theoretical model are compared. The nonlinear stiffness of SIP shows obvious influence on behaviors of TPS and dynamic stress of SIP, and the equivalent linear stiffness of SIP is related to the types of excitations. Finally, the influences on above dynamic responses by the nonlinear stiffness level of SIP were studied. The equivalent linear stiffness coefficient of SIP, acceleration of TPS and dynamic stress of SIP decrease with the increase of the nonlinear level for the stiffness of SIP.
       
  • Ambiguous relative orbits in sequential relative orbit estimation with
           range-only measurements
    • Abstract: Publication date: Available online 10 July 2018Source: Acta AstronauticaAuthor(s): Jingwei Wang, Eric A. Butcher, T. Alan Lovell This paper describes a manifold of ambiguous spacecraft relative orbits that arise in sequential relative orbit estimation. The development herein assumes linear relative dynamics, a circular reference orbit, and range-only measurements. Using a formulation based on relative orbit elements, the ambiguous orbits are categorized into two cases: mirror orbits, which conserve the size and shape but transform the orientation of the true relative orbit, and deformed orbits, which both distort the shape and change the orientation. A special case, that of central ambiguous relative orbits, which are geometrically symmetric relative to the chief's local-vertical-local-horizontal frame is also discussed. The multiplicity of mirror ambiguous orbits, deformed ambiguous orbits and central ambiguous orbits are shown to be three, four and infinity, respectively. Numerical results using an extended Kalman filter are provided to confirm the existence of these ambiguous orbits. Furthermore, the observability is studied analytically with a nonlinear observability criterion using Lie derivatives. It is also shown by numerical results that the inclusion of nonlinearities in the filter model can help resist the tendency of an extended Kalman filter to converge to the ambiguous relative orbits. Finally, the persistence of these ambiguous orbits under unmodeled chief eccentricity error and J2 perturbation is studied.
       
  • XXI century tower: Laser orbital debris removal and collision avoidance
    • Abstract: Publication date: Available online 7 July 2018Source: Acta AstronauticaAuthor(s): Max Calabro, Loïc Perrot A tall “tower” may have many useful applications in scientific domain: energy production, telecommunications, and entertainment. Sometimes space towers are proposed for an easier access to space such as the Thoth project culminating to an altitude of more than 20 km. Nevertheless, access to space is not the most promising use of a high-altitude tower, among them is the orbital debris removal, major point to keep a safe access to space. Ground based laser has been studied several times, laser on-board of a satellite also (chaser), but never a laser at the top of a tower.Such a solution may combine the advantages of the ground based system (i.e. maintenance, power supply, vast number of debris sightings, versatility) and of the space based system. Nevertheless, building a tall tower is a huge investment and must be profitable, but multiple applications of this tower may be envisaged such to have a good return of investment and so, the others potential utilisation may generate a bulk of revenues. As a driver, this tower supporting a powerful laser can be assimilated to a weapon and so must be kept under international control on the European territory.
       
  • The Projecting Surface Method for improvement of surface accuracy of large
           deployable mesh reflectors
    • Abstract: Publication date: Available online 7 July 2018Source: Acta AstronauticaAuthor(s): Sichen Yuan, Bingen Yang, Houfei Fang In traditional form-finding of a deployable mesh reflector (DMR), the nodes of the DMR mesh are placed on the desired working surface and the surface accuracy of the DMR is measured either by the deviation of the nodes from the desired working surface or by the deviation of the mesh from its best-fit surface. Placement of nodes on working surface and inaccurate measures of surface accuracy cause non-negligible surface errors that cannot be further reduced. To deal with these issues and to further improve surface accuracy of DMRs, a new mesh geometry design method, called the Projecting Surface Method (PSM), is presented in this paper. The highlight of the PSM is that it purposely places the nodes of a DMR off its working surface, to achieve higher surface accuracy. To this end, a direct RMS error measuring the deviation of a DMR mesh from its desired working surface is introduced and a projecting surface for hosting the nodes of the DMR mesh is defined. By the direct RMS error and projecting surface, an optimization process produces a mesh geometry with its best-fit surface closest to the desired working surface, leading to significant surface error reduction. As shown in numerical examples of DMRs with 37, 271 and 817 nodes, the PSM can reduce surface errors by 50% or more. The proposed method is usable with existing form-finding methods for further improvement of surface accuracy of DMRs.
       
  • Timing performance evaluation of Radio Determination Satellite Service
           (RDSS) for Beidou system
    • Abstract: Publication date: Available online 7 July 2018Source: Acta AstronauticaAuthor(s): Dongxia Wang, Rui Guo, Tianqiao Zhang Radio Determination Satellite Service (RDSS) is the advantage and particular characteristics of Beidou, which is different from other satellite navigation systems. According to the rare existing researches of its timing service, this article evaluates timing performance of one-way and two-way using the time series analysis method. Moreover, this paper systematically studies the one-way timing and two-way timing principle and introduces Beidou measured data and analysis method. By analyzing the clock error total curve, the mean value segment, noise situation and timing accuracy, we have the conclusions: (1) one-way timing accuracy is less than 30 ns, and its Root Mean Square (RMS) is less than 6.81 ns; (2) two-way timing accuracy is less than 20 ns, and its Root Mean Square (RMS) is less than 3.60 ns; (3) there exist period switching phenomenon of timing data of one-way and two-way in each beam, and stratification of one-way timing data. These conclusions can be used for the difference compensation of Radio Determination Satellite Service (RDSS), which can provide reference for the clock error consistency of Beidou system, and then improve the system service precision.
       
  • Pros and cons of relativistic interstellar flight
    • Abstract: Publication date: Available online 6 July 2018Source: Acta AstronauticaAuthor(s): Oleg G. Semyonov Two technological problems must be solved before daring to interstellar flight: fuel and propulsion. The highest energy-density ‘fuel’ is antimatter in its solid or liquid state and this fuel is likely to be our primary choice for multi-ton relativistic rockets. High-energy ion thrusters powered by annihilation reactors promise superior performance in comparison with direct propulsion by annihilation products. However the power generator onboard can significantly enlarge the rocket dry mass thus limiting the achievable speed. Two physical factors that stand against our dream of the stars are thermodynamics and radiation hazard. Heat-disposing radiator also increases the rocket dry mass. Interstellar gas turns into oncoming flux of hard ionizing radiation at a relativistic speed of the rocket while the oncoming relativistic interstellar dust grains cause mechanical damage. Economy and psychology will play a decisive role in voting for or against the manned interstellar flights.
       
  • Galactic distribution of chirality sources of organic molecules
    • Abstract: Publication date: Available online 6 July 2018Source: Acta AstronauticaAuthor(s): Daniel S. Helman Conceptualizing planetary habitability depends on understanding how living organisms originated and what features of environments are essential to foster abiogenesis. Estimates of the abundance of life's building blocks are confounded by incomplete knowledge of the role of chirality and racemization in organic compounds in the origination of living organisms. Chirality is an essential feature of enzymes as well as many lock-and-key type structures. There are four known processes that can act on complex organic molecules to promote racemization for abiogenesis: quantum-tunneling effects; selection via interaction with circularly polarized light (CPL); templating processes; and interactions with electrical and magnetic (EM) fields. These occur in different places, respectively: cold interstellar space; regions of space with energetic photons, dust and/or magnetic fields; and mineral surfaces (for both templating and EM fields). Chirality as a feature of terrestrial life suggests neither a special place for local development of homochirality nor for extra-terrestrial enrichment and delivery. The presence of these molecules in three competing scenarios for life's origin—chemical gardens, geothermal fields, and ice substrates—relies on a framework of hypothesis and estimation. An easily-modified worksheet is included in the supplemental material that allows a user to generate different scenarios and data related to the distribution of chiral organic molecules as building blocks for living organisms within the galaxy. A simple hypothetical mechanism for planetary magnetic field reversals, based on a high-density plasma inner core, is also presented as a means to aid in estimating field polarity and hence the orientation of racemization processes based on planetary magnetic fields.
       
  • Sliding mode control for autonomous spacecraft rendezvous with collision
           avoidance
    • Abstract: Publication date: Available online 5 July 2018Source: Acta AstronauticaAuthor(s): Qi Li, Jianping Yuan, Huan Wang This paper studies the relative position tracking and attitude synchronization problem of spacecraft rendezvous with the requirement of collision avoidance. To achieve the implementation of the rendezvous procedure, the docking port of the chaser is required to direct towards the counterpart of the target, while the relative distance between the two spacecraft should be larger than the radius of the danger zone during close proximity phase. In order to address the concerned problem, a novel sliding mode control strategy based on artificial potential function is developed, and more specifically, the sliding manifold of the close-loop system is chosen along the negative gradient of the artificial potential function. Within the Lyapunov framework, the proposed control laws are proved to guarantee the convergence of relative position and attitude errors while avoiding any accidental collision between the two spacecraft, even in the presence of external disturbance. Numerical simulations are carried out to demonstrate the effectiveness of the designed control laws.
       
  • Optical changes of molecular contamination thin-film outgassed from
           epoxy-based resin during deposition and desorption process
    • Abstract: Publication date: Available online 5 July 2018Source: Acta AstronauticaAuthor(s): Kazunori Shimazaki, Eiji Miyazaki, Yugo Kimoto Molecular contaminants outgassed from organic materials used for the spacecraft degrade the performance of optical surfaces of spacecraft. The influence of contaminants outgassed from epoxy resin on the spectral transmittance of the quartz substrate was investigated with an in-situ measurement system. The system can deposit the contaminants on temperature-controlled quartz substrates and the transmittance spectra were measured immediately after deposition in vacuum ambient. We obtained the optical constants of the contaminant using transmittance spectrum and simple optical models for optical calculations. The optical constants were described with a harmonic oscillator model and an effective medium approximation model. This paper reports the in-situ measurement results of transmittance spectra of the epoxy-resin-induced contaminants in deposition and desorption process. The thin contamination layer decreased the transmittance in the ultraviolet region. However, the contamination was entirely desorbed at −20 °C and transmittance recovered to the initial value. In addition, the results of optical calculations using the obtained optical constants were compared to the measurement results.
       
  • Uncertainty and sensitivity analysis of flow parameters on aerodynamics of
           a hypersonic inlet
    • Abstract: Publication date: Available online 5 July 2018Source: Acta AstronauticaAuthor(s): Hongkang Liu, Chao Yan, Yatian Zhao, Yupei Qin The performance of the inlet is crucial to the cruise flight of a hypersonic air-breathing propulsion vehicle. The objective of this work is to investigate the uncertainty and sensitivity of pressure field and the performance parameters for a hypersonic inlet due to the uncertainty of five flow parameters, including freestream Mach number, Reynolds number, angle of attack, temperature and wall temperature. The steady Reynolds Averaged Navier-Stokes equations are solved to predict the inlet start and unstart flows within the hysteresis loop. Then, a point-collocation non-intrusive polynomial chaos method (NIPC) is utilized to quantify the uncertainty and sensitivity in the output quantities of interest. The uncertainty analysis in pressure field shows that Mach number and angle of attack of freestream make dominant contributions to the total uncertainty, and the Mach number has remarkable impacts in the isolator. In the start flow, the angle of attack exerts its prominent influence in the post-shock regions, while Mach number mainly dominates these regions ahead of and around the shocks. The reason may be interpreted as the much greater pressure derivatives with respect to angle of attack in the post-shock regions. Significant discrepancies are presented for the unstart flow. The reflected shock waves in the unstart flow are less sensitive to the variations of flow parameters. The external flow field, separation bubble and reflected shocks are significantly affected by angle of attack. Besides, the uncertainties of the performance parameters in the start flow are about twice those in the unstart flow. The sensitivity analysis further reveals that Mach number is the major contributor to the total uncertainty of performance parameters. The correlation coefficients via linear regression method clearly illustrate the relationships between the five input parameters and the performance parameters.
       
  • A passive camera based determination of a non-cooperative and unknown
           satellite's pose and shape
    • Abstract: Publication date: Available online 4 July 2018Source: Acta AstronauticaAuthor(s): Renato Volpe, Giovanni B. Palmerini, Marco Sabatini The relevance of autonomy in space systems during rendezvous and docking operations has been lately increasing. At the scope, a robust GNC architecture is required, which strictly relies on the navigation system's performance and must assure both high efficiency and safety, i.e. low errors and no collisions with the target satellite. One of the most explored fields is the optical navigation one. Using passive optical sensors such as cameras can give high benefit in terms of characterization of the observed scene, thus enlarging the consciousness of what is going on in the mission scenario. The present research investigates the development of a filter which can estimate the shape and relative attitude, position and velocity of a non-cooperative, possibly unknown satellite orbiting around Earth, observed by a camera and a distance sensor mounted on a chaser satellite, whose objective is to successfully complete a docking maneuver. The image taken at a certain time is processed, features are extracted from it and matched with the ones extracted from the image at the previous time step. The matched features along with the relative distance measured by the distance sensor are merged inside an unscented Kalman filter, which predicts, updates and improves the state's estimate throughout the iterations. The expedient used in the filter is to give a 3D characterization to the 2D features used as measurements. The filter estimates the 3D coordinates of these points, i.e. the target's shape, in the camera reference frame, which depend on the target's attitude dynamics and the chaser's relative orbital dynamics. Thus, the target's attitude parameters, i.e. the quaternions, and angular velocity vector, the relative position and velocity vectors and the tracked 3D points are all included in the state vector and estimated by the filter. Subsequently, the 3D point coordinates are determined in the body reference frame. By doing this for all the tracked points, a 3D map of the target can be built.
       
  • A model for understanding and managing cost growth on joint programs
    • Abstract: Publication date: Available online 4 July 2018Source: Acta AstronauticaAuthor(s): Morgan Dwyer, Zoe Szajnfarber, Bruce Cameron, Edward Crawley Although joint programs are typically established to save the government money, recent studies suggest that instead of reducing program cost, jointness may actually induce cost growth. Motivated by three case studies that explored the cost of acquiring systems jointly, this paper presents a model that explains why joint programs often experience large cost growth and how jointness itself may induce it. Specifically, our proposed Agency Action Model suggests that on joint programs, the collaborating agencies' institutional interest in retaining or regaining their autonomy induces cost growth. After explaining the basic components of the model, we demonstrate its ability to explain the cost growth observed in our case studies. Finally, we use the model and our case study data to generate recommendations for managing joint programs in the future.
       
  • Three-dimensional particle simulation of ion thruster plume impingement
    • Abstract: Publication date: Available online 4 July 2018Source: Acta AstronauticaAuthor(s): Guobiao Cai, Hongru Zheng, Lihui Liu, Xiang Ren, Bijiao He The interaction between the high-energy particles in the plume and the spacecraft surfaces will produce interference torque that affects the operating state of the spacecraft in orbit. The thrust of an electric propulsion system is quite small, so it is difficult to be measured directly. The Vacuum Plume Laboratory (VPL) measured the LIPS-200 type ion thruster plume force in an order of 10−3N using a fully elastic micro thrust measuring device. In this paper, the particle in cell (PIC) method and the direct simulation Monte Carlo (DSMC) method are employed to analyze the three-dimensional plasma environments under specified experimental conditions. The Maxwell model is used to calculate the plume force on a 300 mm diameter plate. Simulation results of the plume force give good agreements with the experimental data. Moreover, the effects of the 300 mm diameter aluminium plate on the flow field in vacuum conditions are analyzed. The results show that the number density of atoms is greatly increased before the plate, which has a further impact on the distribution of charge exchange ions (CEX). The enhanced CEX ions moving towards the solar battery panels or sensitive optical components may cause possible damage or interference, which should be avoided by the designers.
       
  • The edge of space: Revisiting the Karman Line
    • Abstract: Publication date: Available online 3 July 2018Source: Acta AstronauticaAuthor(s): Jonathan C. McDowell In this paper I revisit proposed definitions of the boundary between the Earth's atmosphere and outer space, considering orbital and suborbital trajectories used by space vehicles. In particular, I investigate the inner edge of outer space from historical, physical and technological viewpoints and propose 80 km as a more appropriate boundary than the currently popular 100 km Von Kármán line.
       
  • Surrogate modeling for liquid-gas interface determination under
           microgravity
    • Abstract: Publication date: Available online 3 July 2018Source: Acta AstronauticaAuthor(s): Zongyu Wu, Yiyong Huang, Xiaoqian Chen, Xiang Zhang, Wen Yao In recent years the advent of on-orbital refueling technology and the accompanying interest in liquid management in space have rekindled attention to the study of the liquid-gas interface determination. So far, a series of numerical methods, such as the Shooting method, are used to calculate the mathematical model of liquid-gas interface. However, these methods have some drawbacks in common, such as poor convergence, dependence on initial value and instability. As a result, long calculation time and sudden calculation interruption are inevitable. Although satisfactory results can be achieved, it requires human intervention. As an important intermediary of capillary flow and liquid management device design, liquid-gas interface calculations need to be done thousands of times. Therefore, the quickness and robustness of liquid-gas interface calculations are needed. Surrogate modeling method arising with the development of aerospace technology in recent years provides a new way to solve this problem. Combining with the characteristics of liquid-gas interface calculation, a double-layer radial basis function surrogate model is proposed to approximate the mathematical model of liquid-gas interface. This surrogate model is approximately equivalent to the mathematical model of liquid-gas interface but is much easier to solve. Compared with the Shooting method, the efficiency of the surrogate model is improved by 99.46%, and the success rate is increased to 100% from 35%.
       
  • Optimal injection point for launch trajectories with parametric thrust
           profile
    • Abstract: Publication date: Available online 2 July 2018Source: Acta AstronauticaAuthor(s): Max Cerf The problem of finding the optimal thrust profile of a launcher upper stage is analyzed. The engine is non-re-ignitable and it is continuously thrusting, following either a linear or a bilevel parametric profile, until reaching the targeted coplanar orbit. This problem differs from the classical rocket problem where the thrust level is a time-dependent function varying freely between prescribed bounds. Applying the maximum principle yields an analytical closed-loop solution for the thrust direction. Furthermore the final point is found to be necessarily at an apsis, reached from above in the case of a perigee injection. The optimal control problem reduces to a nonlinear problem with only the thrust profile parameters as unknowns. This formulation eases preliminary design studies aiming at defining the optimum upper stage thrust profile. An application case targeting a geostationary transfer orbit illustrates the solution method.
       
  • Systems engineering and design of a Mars Polar Research Base with a human
           crew
    • Abstract: Publication date: Available online 30 June 2018Source: Acta AstronauticaAuthor(s): Anne-Marlene Rüede, Anton Ivanov, Claudio Leonardi, Tatiana Volkova Mars Polar Ice caps have been known ever since they were first observed by Cassini. Robotic exploration missions, starting with Mariner 9, have confirmed that they are composed of water ice. During later missions, instruments such as Mars Global Surveyor's MOLA have established a detailed topography and have estimated their depth to about 3 km in the thickest part, while detailed internal structure has been investigated by MARSIS from Mars Express and SHARAD from the Mars Reconnaissance Orbiter. This analysis proposes to establish a base near North Polar Layered Deposits, to investigate Mars'climate, hydrological processes and to test for possible traces of life. The objectives of the mission are to sustain a crew for nine months on the surface of Mars, near the North Pole, and to bring the crew back to Earth safely. During the surface mission, the crew will drill and analyze Polar Layered Deposits in ice samples. Furthermore, because the North Polar region provides an easy access to water ice, this area has the potential of sustaining a long-term human presence. The Mars Polar Research mission shall therefore prepare for long term missions, spanning over multiple crew generations. Indeed, longer duration missions and larger crews should be facilitated by this first mission. This paper describes a mission design for a Mars Research base using systems engineering approach and scenario testing. The goal of the work is to establish a strategy composed of various technologies that have been selected accordingly. The requirements related to crew composition, human requirements, science requirements, communication, habitat structure and usability requirements are derived and compiled into mass, volume, data and power consumption. A design for the base and mission scenario is also proposed. Given the identified requirements, possible technologies for life support systems, radiation protection, in-situ propellant production, thermal control, air pressure difference compensation and availability of power are discussed and solutions to focus on are recommended. Furthermore, the requirements for a long-term mission preparation are also identified and solutions to include in a first Mars mission with crew are recommended. In conclusion, approximately 110 metric tons and 160 kW are required to enable a Mars Polar mission with human crew. A two-phase mission is recommended for enabling the testing of key in-situ resource utilization technologies allowing to minimize mass, while ensuring the security of the crew. The use of optimal payload and fairing, a Mars orbit crane system and deployable structures is recommended. Finally, in preparation for a long-term presence of humans on Mars, including in-situ testing of key technologies enabling the production of consumables facilitating autonomy from Earth is recommended. The consumables that have been identified as not being able to be tested before a first crew is sent to Mars are food and energy production. These developments may serve as priorities for current Mars settlement programs.
       
  • Active slosh control and damping - Simulation and experiment
    • Abstract: Publication date: Available online 30 June 2018Source: Acta AstronauticaAuthor(s): Martin Konopka, Francesco De Rose, Hans Strauch, Christina Jetzschmann, Nicolas Darkow, Jens Gerstmann Future reignitable cryogenic upper stages perform long ballistic coasting phases in earth orbit. During those coasting phases, the tanks are loaded with liquid propellants and propellant sloshing occurs due to external disturbances or attitude change maneuvers. The sloshing propellant motion induces reaction forces and torques acting on the space vehicle structure, e.g. rocket upper stages. To keep the upper stage at the desired target attitude, the guidance, navigation, and control (GNC) algorithm commands thruster firings to counter the fluid forces. At ArianeGroup (AG), the Final Phase Simulator FiPS aims at simulating the coupling between fluid mechanics, GNC, and rigid body dynamics. To validate the coupling of GNC with linear lateral water sloshing, on ground experiments at the German Aerospace Center's Hexapod sloshing facility were performed. It was demonstrated that the developed control algorithm is able to damp the linear lateral sloshing within 4 s. FiPS simulations of the open and closed loop sloshing experiments showed that the experimental forces are matched with an uncertainty of less then 5% for open loop phases. For closed-loop phases the simulations match the experimental damping intervals with an accuracy of better than 5% and the force amplitude with an accuracy of about 20%.
       
  • A concept of hazardous NEO detection and impact warning system
    • Abstract: Publication date: Available online 30 June 2018Source: Acta AstronauticaAuthor(s): Toshinori Ikenaga, Yohei Sugimoto, Matteo Ceriotti, Makoto Yoshikawa, Toshifumi Yanagisawa, Hitoshi Ikeda, Nobuaki Ishii, Takashi Ito, Masayoshi Utashima In 2013, the well-known Chelyabinsk meteor entered the Earth's atmosphere over Chelyabinsk, Russia. It is estimated that the meteor exploded at altitude near 30 km[2], which damaged thousands of buildings and injured a thousand of residents[3–4]. The estimated size of the meteor is approximately 20 m[2]. Because the meteor approached to Earth from Sun direction, no ground-based observatories could not detect until the impact.Considering such situations, the paper proposes a concept to detect Chelyabinsk-class small Near-Earth Objects. The concept addresses a “last-minute” warning system of NEO impact, in the same manner of “Tsunami” warning.To achieve the mission objective, two locations are assumed for the space telescope installation point i.e., Sun-Earth Lagrange point 1, SEL1 and Artificial Equilibrium Point, AEP. SEL1 is one of the natural equilibrium points, on the other hand, AEP is artificially equilibrated point by Sun and Earth gravity, centrifugal force and low-thrust acceleration. The magnitude of the acceleration to keep AEP is sufficiently small near 1 au radius orbit around the Sun i.e., the order of μm/s2 which can be achieved by solar sail. Through some cases of numerical simulations considering the size of NEOs and detector capability, this paper will show the feasibility of the proposed concept.
       
  • Tentative design of SBSS constellations for LEO debris catalog maintenance
    • Abstract: Publication date: Available online 28 June 2018Source: Acta AstronauticaAuthor(s): Jianli Du, Junyu Chen, Bin Li, Jizhang Sang This paper proposes three Sun-Synchronous Orbit (SSO) Space-Based Space Surveillance (SBSS) optical constellations with the intention of building up and maintaining a catalog of 200,000 LEO space debris. The three constellations, namely CON.1, CON.2 and CON.3, are Walker analogs whose notations are “98.7°:12/4/1”, “98.7°:12/6/1”, and “98.7°:24/4/1”, respectively. The configurations can ensure a globally even distribution that is suitable for catalog maintenance of LEO debris. In addition, the visible-band sensors of the three constellations are assumed to have different sensitivity to cm-size debris and Field Of Views (FOVs). Simulation experiments are made to evaluate the detectability and maintainability performances of the three constellations and their variants. By analyzing the detectable orbital arcs, CON.1 and CON.2 show the potential capability to maintain a dynamic catalog of more than 200,000 debris when the number of satellites is 8 or 12. The CON.1 constellation with 12 satellites performs best in terms of the capacities of the dynamic catalog and continuously maintainable catalog, with 225,000 and 157,500 objects, respectively.
       
  • Asteroid de-spin and deflection strategy using a solar-sail spacecraft
           with reflectivity control devices
    • Abstract: Publication date: Available online 24 June 2018Source: Acta AstronauticaAuthor(s): Shota Kikuchi, Jun'ichiro Kawaguchi Various asteroid mitigation strategies have been proposed to avoid destructive impact events. In such a mission, the spinning motion of an asteroid can prevent a precise and effective deflection maneuver. To circumvent this problem, this study investigates a novel de-spin method using a solar sail spacecraft that is attached to the surface of an asteroid. In this approach, the solar radiation pressure torque induced by reflectivity control devices on the sail membrane is exploited to cancel out the spin rate of an asteroid without requiring fuel. In addition, once attitude control is achieved after the de-spin, the trajectory of the asteroid can be deflected by leveraging the solar radiation pressure force acting on the sail attached to the asteroid. This paper constructs general theories on the proposed asteroid de-spin and deflection methods and provides evidence that this novel strategy would be a feasible option for mitigating small and slow-spinning asteroids.
       
  • An Early History of the Philippine Space Development program
    • Abstract: Publication date: Available online 21 June 2018Source: Acta AstronauticaAuthor(s): Q. Verspieren, G. Coral, B. Pyne, H. Roy In 2018, the Congress of the Philippines is expected to pass the Philippine Space Development Act, leading to the adoption of the first national space policy in the country and the establishment of the Philippine Space Agency (or PhilSA). This historic event is the final outcome of a long process involving various stakeholders in the Philippines from government, academia and industry.This article provides the first comprehensive history of this process from its inception in the late 20 t h century until now, with a specific focus on its acceleration since 2013. It also investigates the future expectations of the national space development program, in particular regarding the development of a local space industry.Apart from solely describing the history of the Philippines, this paper presents more generally the case of a developing country willing to gain a foothold in space. In comparison with existing literature on space development programs focusing exclusively on rich western countries or powerful emerging nations, this paper provides other developing countries with highly valuable information by describing such a transparent, balanced and promising initiative as the Philippine space development program.
       
  • Development of an intrusive technique for particles collection in rockets
           plume
    • Abstract: Publication date: Available online 19 June 2018Source: Acta AstronauticaAuthor(s): Stefania Carlotti, Filippo Maggi, Alessandro Ferreri, Luciano Galfetti, Riccardo Bisin, Dominik Saile, Ali Gühlan, Christopher Groll, Tobias Langener An intrusive technique for particles capturing in supersonic-high temperature flows for the use in solid rocket motors plume characterization is proposed. A supersonic probe for the collection of the condensed combustion products in the proximity of the rocket nozzle has been sized to handle a progressive deceleration and cool down of the exhaust gas, preventing from liquid particles breakup. A quasy-1D gas dynamics software (POLIRocket-V2) based on the Shapiro method and normal shock wave theory, supported by a CFD investigation using the DLR TAU code, was employed for the feasibility and the design study. Preliminary cold flow tests have been performed in the supersonic vertical wind tunnel at DLR in relevent environment as a proof of concept of the probe working principle and the collection methodology studied.
       
  • Experimental and numerical methods for radiative wall heat flux
           predictions in paraffin–based hybrid rocket engines
    • Abstract: Publication date: Available online 18 June 2018Source: Acta AstronauticaAuthor(s): Giuseppe Leccese, Daniele Bianchi, Francesco Nasuti, Keith Javier Stober, Pavan Narsai, Brian Joseph Cantwell The paper is intended to present both experimental and numerical approaches for estimating the heat exchange through thermal radiation towards the walls of lab–scale paraffin–based thrust chambers. Two firing tests of a lab–scale gaseous–oxygen/paraffin–wax hybrid rocket engine have been performed to apply such methods. In particular, the radiative wall heat flux has been evaluated by both spectroscopic measurements and discrete transfer method computations. Details of such approaches are given together with results achieved. The significant gap between experimental and numerical estimations suggested further investigations of influencing processes, whose main outcomes and lessons learned are critically discussed.
       
  • Control of the drag on a spacecraft in the earth’s ionosphere using the
           spacecraft’s magnetic field
    • Abstract: Publication date: Available online 18 June 2018Source: Acta AstronauticaAuthor(s): Valentin A. Shuvalov, Nikolai B. Gorev, Nikolai A. Tokmak, Nikolai I. Pis'mennyi, Galina S. Kochubei This paper shows the possibility of active control of the drag on a spacecraft in the Earth's ionosphere using the electromagnetic force produced by the interaction of the spacecraft's magnetic field with the incident plasma flow. As a result of experimental simulation of the dynamic interaction of the magnetic field of a sphere with a hypersonic flow of the rarefied ionospheric plasma, the sphere drag coefficient is determined as a function of the ratio of the magnetic pressure to the dynamic pressure in a wide range of the angle between the incident flow velocity and the magnetic field and the angle between the incident hypersonic plasma flow and the velocity of a subsonic plasma jet injected from the sphere surface. It is shown that injecting a subsonic plasma jet into the mini-magnetosphere cavity provides a several-fold increase in the drag coefficient of a “magnetized” sphere (a sphere with its own magnetic field) in a hypersonic rarefied plasma flow in comparison with a “nonmagnetized” sphere. A 0.6 … 0.8 T magnetic field of a “magnetized” body may be an efficient means for its deorbiting through increasing the drag on the body in the Earth's ionosphere, which provides a way for removing space debris objects to lower orbits.
       
 
 
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