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Acta Astronautica
Journal Prestige (SJR): 0.758  Citation Impact (citeScore): 2 Number of Followers: 411  Hybrid journal (It can contain Open Access articles)ISSN (Print) 0094-5765 Published by Elsevier  [3161 journals]
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- Modeling the equations of state using a flamelet approach in LRE-like
conditions- Abstract: Publication date: Available online 17 September 2018Source: Acta AstronauticaAuthor(s): Pasquale Eduardo Lapenna, Giuseppe Indelicato, Rachele Lamioni, Francesco CretaAbstractIn this study, we investigate the averaging and modeling of real fluid equations of state (EoS) used in the context of numerical simulations of flows under conditions relevant to liquid rocket engine (LRE) thrust chambers. Direct numerical simulation (DNS) data of supercritical and transcritical mixing are used, in an a-priori fashion, to glean insight on the effect of averaging the EoS in the context of Reynolds averaged Navier Stokes (RANS). The widely employed no-model approach for the EoS, which neglects residual effects, is shown to generate large errors under transcritical conditions due to the extreme non-linearity of the pseudo-boiling processes. Further a-priori analysis, performed by filtering the DNS data in a large eddy simulation (LES) context, shows that highly resolved LES mitigates such sub grid scale errors. We discuss such EoS modeling related issues in the context of flamelet based turbulent combustion approaches and low Mach number assumptions, under which they acquire an unambiguous thermodynamic consistency. We show a-posteriori that a steady laminar flamelet formulation, implemented in a low Mach number unsteady RANS solver, is capable of correctly capturing the characteristic flow field and structure of a reference liquid oxygen (LOx) - gaseous methane (GCH4) cryogenic flame at supercritical pressure.
- Abstract: Publication date: Available online 17 September 2018Source: Acta AstronauticaAuthor(s): Pasquale Eduardo Lapenna, Giuseppe Indelicato, Rachele Lamioni, Francesco CretaAbstractIn this study, we investigate the averaging and modeling of real fluid equations of state (EoS) used in the context of numerical simulations of flows under conditions relevant to liquid rocket engine (LRE) thrust chambers. Direct numerical simulation (DNS) data of supercritical and transcritical mixing are used, in an a-priori fashion, to glean insight on the effect of averaging the EoS in the context of Reynolds averaged Navier Stokes (RANS). The widely employed no-model approach for the EoS, which neglects residual effects, is shown to generate large errors under transcritical conditions due to the extreme non-linearity of the pseudo-boiling processes. Further a-priori analysis, performed by filtering the DNS data in a large eddy simulation (LES) context, shows that highly resolved LES mitigates such sub grid scale errors. We discuss such EoS modeling related issues in the context of flamelet based turbulent combustion approaches and low Mach number assumptions, under which they acquire an unambiguous thermodynamic consistency. We show a-posteriori that a steady laminar flamelet formulation, implemented in a low Mach number unsteady RANS solver, is capable of correctly capturing the characteristic flow field and structure of a reference liquid oxygen (LOx) - gaseous methane (GCH4) cryogenic flame at supercritical pressure.
- On-orbit assembly mission planning considering topological constraint and
attitude disturbance- Abstract: Publication date: Available online 15 September 2018Source: Acta AstronauticaAuthor(s): Yuchen She, Shuang Li, Bin Du, Kai CaoAbstractThis paper investigates the on-orbit assembly mission planning problem of a large space antenna structure. The study is significantly important for the construction of large-scale spatial structures. The flexible connection between the antenna pieces and the topological constraint are taken into consideration. Several innovative approaches are presented to realize the mission planning process. First, the mesh model from the fly net dynamics is adopted and modified to create a mapping between the vibrations of the antenna plate during the assembly process and the assembly path. Second, the geometrical model is developed to implement the topological constraint into the system. Finally, a hybrid method combining the branch-and-bound method and the modified ant colony algorithm is proposed to solve the optimization problem. The simulation results show that the proposed algorithm can efficiently obtain a solution which respects the topological constraint and suppresses the vibration disturbance as much as possible.
- Abstract: Publication date: Available online 15 September 2018Source: Acta AstronauticaAuthor(s): Yuchen She, Shuang Li, Bin Du, Kai CaoAbstractThis paper investigates the on-orbit assembly mission planning problem of a large space antenna structure. The study is significantly important for the construction of large-scale spatial structures. The flexible connection between the antenna pieces and the topological constraint are taken into consideration. Several innovative approaches are presented to realize the mission planning process. First, the mesh model from the fly net dynamics is adopted and modified to create a mapping between the vibrations of the antenna plate during the assembly process and the assembly path. Second, the geometrical model is developed to implement the topological constraint into the system. Finally, a hybrid method combining the branch-and-bound method and the modified ant colony algorithm is proposed to solve the optimization problem. The simulation results show that the proposed algorithm can efficiently obtain a solution which respects the topological constraint and suppresses the vibration disturbance as much as possible.
- Muscle blood content and muscle oxygen saturation in response to head down
and head UP tilt- Abstract: Publication date: Available online 15 September 2018Source: Acta AstronauticaAuthor(s): Hasan Birol Çotuk, Adil Deniz Duru, Orkun Pelvan, Savaş AkbaşAbstractAs microgravity alters blood distribution and flow in muscle tissue, in the present study, blood content and oxygen saturation in extremity muscles were evaluated by an analogous body tilting model. For this purpose, the supine resting posture was compared with successive head down tilt (HDT) and head up tilt (HUT) in ten well trained male volunteers. Muscle oxygen saturation and total hemoglobin content were measured using near infrared spectroscopy sensors attached on the gastrocnemius, vastus lateralis and biceps brachii muscles of each body side. Simultaneously, continuous and noninvasive recordings of arterial blood pressure and heart rate variability were performed. The test protocol consisted of the following stages (5 min each): supine, 6o HDT, 15o HDT, 30o HDT, supine, 6o HUT, 15o HUT, and 30o HUT postures in successive order. The evolution of the rhythmic components in the recorded time series were analyzed by wavelet based multiscale time-frequency distributions, and synchronization between measurement sites by wavelet phase synchronization indexing. A support vector machine (SVM) algorithm was employed for classification of posture in relation to muscle total hemoglobin content or oxygen saturation. Increasing or decreasing gravitational impact due to the posture changes resulted in significant increases or decreases of total hemoglobin content in the muscles, but showed no linear relation to muscle oxygen saturation. This effect was stronger pronounced as the respective extremity part was located more centrifugal to the body tilting axis. Blood pressure and heart rate did not influence muscle oxygen saturation. The two major rhythmic components in the cardiovascular system, the 0.1 Hz and respiratory rhythm, were only weakly displayed in the muscle oxygen saturation time series. Oscillations of muscle oxygen saturation in the very low frequency band displayed left-right synchrony. SVM was able to classify postural changes of muscle blood content with good accuracy but not those of muscle oxygen saturation. These results imply that muscle oxygen saturation during acute postural changes may be mainly regulated by neural drive to the micro-vascular circulation and not by systemic cardiovascular kinetics.
- Abstract: Publication date: Available online 15 September 2018Source: Acta AstronauticaAuthor(s): Hasan Birol Çotuk, Adil Deniz Duru, Orkun Pelvan, Savaş AkbaşAbstractAs microgravity alters blood distribution and flow in muscle tissue, in the present study, blood content and oxygen saturation in extremity muscles were evaluated by an analogous body tilting model. For this purpose, the supine resting posture was compared with successive head down tilt (HDT) and head up tilt (HUT) in ten well trained male volunteers. Muscle oxygen saturation and total hemoglobin content were measured using near infrared spectroscopy sensors attached on the gastrocnemius, vastus lateralis and biceps brachii muscles of each body side. Simultaneously, continuous and noninvasive recordings of arterial blood pressure and heart rate variability were performed. The test protocol consisted of the following stages (5 min each): supine, 6o HDT, 15o HDT, 30o HDT, supine, 6o HUT, 15o HUT, and 30o HUT postures in successive order. The evolution of the rhythmic components in the recorded time series were analyzed by wavelet based multiscale time-frequency distributions, and synchronization between measurement sites by wavelet phase synchronization indexing. A support vector machine (SVM) algorithm was employed for classification of posture in relation to muscle total hemoglobin content or oxygen saturation. Increasing or decreasing gravitational impact due to the posture changes resulted in significant increases or decreases of total hemoglobin content in the muscles, but showed no linear relation to muscle oxygen saturation. This effect was stronger pronounced as the respective extremity part was located more centrifugal to the body tilting axis. Blood pressure and heart rate did not influence muscle oxygen saturation. The two major rhythmic components in the cardiovascular system, the 0.1 Hz and respiratory rhythm, were only weakly displayed in the muscle oxygen saturation time series. Oscillations of muscle oxygen saturation in the very low frequency band displayed left-right synchrony. SVM was able to classify postural changes of muscle blood content with good accuracy but not those of muscle oxygen saturation. These results imply that muscle oxygen saturation during acute postural changes may be mainly regulated by neural drive to the micro-vascular circulation and not by systemic cardiovascular kinetics.
- Hybrid tension control method for tethered satellite systems during large
tumbling space debris removal- Abstract: Publication date: Available online 15 September 2018Source: Acta AstronauticaAuthor(s): Zhongyi Chu, Jingnan Di, Jing CuiAbstractLarge space debris in the valuable orbit is an inevitable and serious problem which increases the risk of fatal collisions. The use of space tether is a promising method for de-orbiting large debris safely at low cost and low energy consumption. However, the risk of winding is increasing because of variations in sway motion due to underactuativity, high nonlinearity, and the strong coupling of the tethered system. Hence, in this paper, a hybrid tension control method is proposed to stabilise satellite system tethered to large space debris. The dynamic equations of a tethered satellite system are formulated with an analysis of the impact of the variation in sway motion on tether winding. Following this, the optimal commands of sway motion are planned using the Gauss pseudospectral method to preliminarily avoid the tether winding with the target. To further reduce the risk of winding, a fuzzy adaptive proportion differentiation (PD) controller is designed to stabilise the relative attitude of the tether and the large debris. And the control tension is considered as a time-varying parameter and is added using a hierarchical sliding-mode controller (HSMC), which is used to control the in-plane angle and the relative distance to implement the hybrid control of the overall tethered satellite system. Several simulations were implemented to verify the effectiveness of the proposed tension control method.
- Abstract: Publication date: Available online 15 September 2018Source: Acta AstronauticaAuthor(s): Zhongyi Chu, Jingnan Di, Jing CuiAbstractLarge space debris in the valuable orbit is an inevitable and serious problem which increases the risk of fatal collisions. The use of space tether is a promising method for de-orbiting large debris safely at low cost and low energy consumption. However, the risk of winding is increasing because of variations in sway motion due to underactuativity, high nonlinearity, and the strong coupling of the tethered system. Hence, in this paper, a hybrid tension control method is proposed to stabilise satellite system tethered to large space debris. The dynamic equations of a tethered satellite system are formulated with an analysis of the impact of the variation in sway motion on tether winding. Following this, the optimal commands of sway motion are planned using the Gauss pseudospectral method to preliminarily avoid the tether winding with the target. To further reduce the risk of winding, a fuzzy adaptive proportion differentiation (PD) controller is designed to stabilise the relative attitude of the tether and the large debris. And the control tension is considered as a time-varying parameter and is added using a hierarchical sliding-mode controller (HSMC), which is used to control the in-plane angle and the relative distance to implement the hybrid control of the overall tethered satellite system. Several simulations were implemented to verify the effectiveness of the proposed tension control method.
- Leadership roles and group climate in isolation: A case study of 4-subject
180-day mission- Abstract: Publication date: Available online 14 September 2018Source: Acta AstronauticaAuthor(s): Ruilin Wu, Qianying Ma, Jianghui Xiong, Zi Xu, Yinghui LiAbstractIn an isolated and confined environment, leadership roles are important for mission success and harmonious interpersonal relationships. The task role benefits the achievement of mission goals and operational needs while the supportive role promotes emotional support and boosts morale. This study examined the effects of time on leadership roles and leadership influence among four crew members in a 180-day isolated and confined experiment. The crew members completed the Group Environment Scale around every two weeks to record their perceptions of leader support and leader control. Overall, perceived leader support decreased over the isolation and confinement period, but the perception of leader control remained stable. Crew members perceived the role of leadership as providing support more than control. Leadership had a positive effect on the group climate by clarifying tasks and reducing aggression. Finally, the various functions of perceived leadership roles could be explained by crew composition and occupational features. The findings have implications that should be considered in future space missions and earth-based experiments when selecting leaders and arranging work duties with high autonomy. Further studies are needed to examine the generality of leadership roles in isolated and confined environments.
- Abstract: Publication date: Available online 14 September 2018Source: Acta AstronauticaAuthor(s): Ruilin Wu, Qianying Ma, Jianghui Xiong, Zi Xu, Yinghui LiAbstractIn an isolated and confined environment, leadership roles are important for mission success and harmonious interpersonal relationships. The task role benefits the achievement of mission goals and operational needs while the supportive role promotes emotional support and boosts morale. This study examined the effects of time on leadership roles and leadership influence among four crew members in a 180-day isolated and confined experiment. The crew members completed the Group Environment Scale around every two weeks to record their perceptions of leader support and leader control. Overall, perceived leader support decreased over the isolation and confinement period, but the perception of leader control remained stable. Crew members perceived the role of leadership as providing support more than control. Leadership had a positive effect on the group climate by clarifying tasks and reducing aggression. Finally, the various functions of perceived leadership roles could be explained by crew composition and occupational features. The findings have implications that should be considered in future space missions and earth-based experiments when selecting leaders and arranging work duties with high autonomy. Further studies are needed to examine the generality of leadership roles in isolated and confined environments.
- Numerical investigation on behaviors of shock train in a hypersonic inlet
with translating cowl- Abstract: Publication date: Available online 12 September 2018Source: Acta AstronauticaAuthor(s): Wen Shi, Juntao Chang, Junlong Zhang, Youyin Wang, Wenxin Hou, Wen BaoAbstractIn order to extend the range of flight Mach numbers, the variable-geometry inlet equipped with translating cowl has been proposed. To estimate the effects that the translation of cowl has on the behavior of shock train under constant backpressure in a two-ramp hypersonic inlet, numerical investigations utilizing dynamic mesh method have been conducted. The structures of background waves and trajectories of shock-impact points with variable internal contraction ratios at Mach number 5.9 are obtained. Then, the paths of shock train leading edges have been provided and studied. The results reveal that the shock train leading edge on the bottom wall gets over the shock-impact points abruptly and forces the shock train to undergo multiple back-and-forth translation motions accompanied with separation mode transition. Severe abrupt motions of shock train occur when it encounters multiple shock-impact points aroused by the separation bubble. Analogously, the shock train interacts with the separation bubble located on the top wall periodically. The probability that the oscillation takes place is enlarged due to the existence of separation bubble. Therefore, the behavior of shock train is highly associated with the variable background waves.
- Abstract: Publication date: Available online 12 September 2018Source: Acta AstronauticaAuthor(s): Wen Shi, Juntao Chang, Junlong Zhang, Youyin Wang, Wenxin Hou, Wen BaoAbstractIn order to extend the range of flight Mach numbers, the variable-geometry inlet equipped with translating cowl has been proposed. To estimate the effects that the translation of cowl has on the behavior of shock train under constant backpressure in a two-ramp hypersonic inlet, numerical investigations utilizing dynamic mesh method have been conducted. The structures of background waves and trajectories of shock-impact points with variable internal contraction ratios at Mach number 5.9 are obtained. Then, the paths of shock train leading edges have been provided and studied. The results reveal that the shock train leading edge on the bottom wall gets over the shock-impact points abruptly and forces the shock train to undergo multiple back-and-forth translation motions accompanied with separation mode transition. Severe abrupt motions of shock train occur when it encounters multiple shock-impact points aroused by the separation bubble. Analogously, the shock train interacts with the separation bubble located on the top wall periodically. The probability that the oscillation takes place is enlarged due to the existence of separation bubble. Therefore, the behavior of shock train is highly associated with the variable background waves.
- Attitude and vibration control with double-gimbal variable-speed control
moment gyros- Abstract: Publication date: Available online 12 September 2018Source: Acta AstronauticaAuthor(s): Takahiro Sasaki, Takashi Shimomura, Sam Pullen, Hanspeter SchaubAbstractThe attitude and vibration control of a flexible spacecraft with two parallel double-gimbal variable-speed control moment gyros (DGVSCMGs) is considered. The coupled nonlinear equations of motion create a complex challenge in a pointing control development. First a Gain-Scheduled (GS) controller for a 3-axis attitude control is designed by the post-guaranteed linear matrix inequalities (LMIs) method with H2/H∞ constraints. Next an H2/H∞ controller for vibration control is designed to attain both attitude and vibration control at the same time. The two controllers are combined using the dynamic inversion (DI) technique. Finally, the effectiveness of the proposed combined controller is demonstrated through a numerical example.
- Abstract: Publication date: Available online 12 September 2018Source: Acta AstronauticaAuthor(s): Takahiro Sasaki, Takashi Shimomura, Sam Pullen, Hanspeter SchaubAbstractThe attitude and vibration control of a flexible spacecraft with two parallel double-gimbal variable-speed control moment gyros (DGVSCMGs) is considered. The coupled nonlinear equations of motion create a complex challenge in a pointing control development. First a Gain-Scheduled (GS) controller for a 3-axis attitude control is designed by the post-guaranteed linear matrix inequalities (LMIs) method with H2/H∞ constraints. Next an H2/H∞ controller for vibration control is designed to attain both attitude and vibration control at the same time. The two controllers are combined using the dynamic inversion (DI) technique. Finally, the effectiveness of the proposed combined controller is demonstrated through a numerical example.
- Numerical investigation of a rotating detonation engine under
premixed/non-premixed conditions- Abstract: Publication date: Available online 12 September 2018Source: Acta AstronauticaAuthor(s): Jian Sun, Jin Zhou, Shijie Liu, Zhiyong LinAbstractRotating detonation engines are widely studied because of their compact configurations and high thermal cycle efficiency. For briefty, most of the numerical simulations of rotating detonation engines used premixed reactant mixtures. The rotating detonation waves under non-premixed condition are not studied enough. Here, a series of three-dimensional numerical simulations of a rotating detonation engine under both premixed and non-premixed conditions using H2/air mixture are performed. The explicit formulation of density-based solver in ANSYS Fluent is used to perform the simulations. Two total mass flow rates of 272.3 g/s and 500 g/s are selected. When the total mass flow rate is 272.3 g/s, the engine operates at single-wave mode under both premixed and non-premixed conditions. When the total mass flow rate is 500 g/s, the engine operates at single-wave mode under premixed condition. While under non-premixed condition, a spontaneous formation of dual-wave mode is observed. This case agrees well with the phenomenon observed in experiments that as the total mass flow rate increases, the number of rotating detonation waves tends to increase. Pressure waves caused by the high pressure behind the detonation waves can propagate upstream to the H2 and air plenums. The pressure feedback in the H2 plenum is much more obvious than that in the air plenum. Due to the imperfect mixing of H2/air and the more deflagration combustion caused by the hot detonation products, the thrust of the RDE under non-premixed condition is smaller than that under premixed condition.
- Abstract: Publication date: Available online 12 September 2018Source: Acta AstronauticaAuthor(s): Jian Sun, Jin Zhou, Shijie Liu, Zhiyong LinAbstractRotating detonation engines are widely studied because of their compact configurations and high thermal cycle efficiency. For briefty, most of the numerical simulations of rotating detonation engines used premixed reactant mixtures. The rotating detonation waves under non-premixed condition are not studied enough. Here, a series of three-dimensional numerical simulations of a rotating detonation engine under both premixed and non-premixed conditions using H2/air mixture are performed. The explicit formulation of density-based solver in ANSYS Fluent is used to perform the simulations. Two total mass flow rates of 272.3 g/s and 500 g/s are selected. When the total mass flow rate is 272.3 g/s, the engine operates at single-wave mode under both premixed and non-premixed conditions. When the total mass flow rate is 500 g/s, the engine operates at single-wave mode under premixed condition. While under non-premixed condition, a spontaneous formation of dual-wave mode is observed. This case agrees well with the phenomenon observed in experiments that as the total mass flow rate increases, the number of rotating detonation waves tends to increase. Pressure waves caused by the high pressure behind the detonation waves can propagate upstream to the H2 and air plenums. The pressure feedback in the H2 plenum is much more obvious than that in the air plenum. Due to the imperfect mixing of H2/air and the more deflagration combustion caused by the hot detonation products, the thrust of the RDE under non-premixed condition is smaller than that under premixed condition.
- Analysis of artificial gravity paradigms using a mathematical model of
spatial orientation- Abstract: Publication date: Available online 12 September 2018Source: Acta AstronauticaAuthor(s): Grant R. Vincent, Jason Gruber, Michael C. Newman, Torin K. ClarkAbstractArtificial gravity (AG) is a promising approach to reduce the physiological deconditioning experienced by astronauts. Here we propose the linear sled hybrid AG system as an alternative to the typical centrifuge approach to creating AG. In this paradigm, the rider is briefly linearly accelerated towards their head, then rotated 180° around, then decelerated. This sequence is repeated creating footward loading during the linear acceleration and deceleration phases, replicating standing upright on Earth, without any gravity gradient or Coriolis forces. The 180° rotation also produces gradient centripetal acceleration, for a “hybrid” approach. We simulated the well-validated observer model to predict the rider's orientation perception and potential disorientation in response to these two AG paradigms. Particularly, we simulated head tilts to investigate the cross-coupled illusion. For the centrifuge, as expected, we found head tilts caused the cross-coupled illusion and an illusory sense of tilt. As a novel prediction, we found the head tilt angle and centrifuge spin rate to interact non-linearly, producing an inflection point in the peak perceived tilt of the cross-coupled illusion. We found the linear sled paradigm to be well perceived and, as expected, head tilts did not produce the cross-coupled illusion. While the observer model predicted the linear sled paradigm to not be disorienting, future experimental work is necessary for validation. Comfort and motion sickness feasibility, as well as countermeasure efficacy, should be studied experimentally.
- Abstract: Publication date: Available online 12 September 2018Source: Acta AstronauticaAuthor(s): Grant R. Vincent, Jason Gruber, Michael C. Newman, Torin K. ClarkAbstractArtificial gravity (AG) is a promising approach to reduce the physiological deconditioning experienced by astronauts. Here we propose the linear sled hybrid AG system as an alternative to the typical centrifuge approach to creating AG. In this paradigm, the rider is briefly linearly accelerated towards their head, then rotated 180° around, then decelerated. This sequence is repeated creating footward loading during the linear acceleration and deceleration phases, replicating standing upright on Earth, without any gravity gradient or Coriolis forces. The 180° rotation also produces gradient centripetal acceleration, for a “hybrid” approach. We simulated the well-validated observer model to predict the rider's orientation perception and potential disorientation in response to these two AG paradigms. Particularly, we simulated head tilts to investigate the cross-coupled illusion. For the centrifuge, as expected, we found head tilts caused the cross-coupled illusion and an illusory sense of tilt. As a novel prediction, we found the head tilt angle and centrifuge spin rate to interact non-linearly, producing an inflection point in the peak perceived tilt of the cross-coupled illusion. We found the linear sled paradigm to be well perceived and, as expected, head tilts did not produce the cross-coupled illusion. While the observer model predicted the linear sled paradigm to not be disorienting, future experimental work is necessary for validation. Comfort and motion sickness feasibility, as well as countermeasure efficacy, should be studied experimentally.
- Stable orbital transfer of partial space elevator by tether deployment and
retrieval- Abstract: Publication date: Available online 12 September 2018Source: Acta AstronauticaAuthor(s): Gefei Shi, Zhanxia Zhu, Zheng H. ZhuAbstractThis paper investigates the stable orbital transfer of a climber in a partial space elevator system in circular orbits. The elevator includes a main satellite, an end body, and a climber, which are connected by two straight, inextensible, massless and variable length tethers. Derived from system dynamics, two closed-loop velocity control strategies are proposed to regulate the tethers' deploying and retrieving speeds directly to control the Coriolis forces acting on the climber and the end body directly. The stable region of the control strategies is analyzed, and the effectiveness of the proposed strategies is demonstrated by numerical simulation. Simulation results show that the newly proposed strategies effectively minimize the libration of a partial space elevator system during the orbital transfer of the climber and the system states successfully approach to the desired equilibrium by the end of the orbital transfer. Furthermore, the proposed deployment/retrieval control strategies can be converted to tension control laws easily.
- Abstract: Publication date: Available online 12 September 2018Source: Acta AstronauticaAuthor(s): Gefei Shi, Zhanxia Zhu, Zheng H. ZhuAbstractThis paper investigates the stable orbital transfer of a climber in a partial space elevator system in circular orbits. The elevator includes a main satellite, an end body, and a climber, which are connected by two straight, inextensible, massless and variable length tethers. Derived from system dynamics, two closed-loop velocity control strategies are proposed to regulate the tethers' deploying and retrieving speeds directly to control the Coriolis forces acting on the climber and the end body directly. The stable region of the control strategies is analyzed, and the effectiveness of the proposed strategies is demonstrated by numerical simulation. Simulation results show that the newly proposed strategies effectively minimize the libration of a partial space elevator system during the orbital transfer of the climber and the system states successfully approach to the desired equilibrium by the end of the orbital transfer. Furthermore, the proposed deployment/retrieval control strategies can be converted to tension control laws easily.
- The influences of aramid fibre and its compound with ceramic on shielding
performance of stuffed layer- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Fa-wei Ke, Jie Huang, Xue-zhong Wen, Qiang Song, Qing Luo, Sen LiuAbstractFour kinds of aramid fibre and three compound modes of aramid with ceramic fibre were studied by hypervelocity impact tests respectively to probe the fabric layer with higher performance in space debris shield configuration. The aramid fibres include the Aramid-II manufactured in America, the Aramid-II, the Aramid-III and the PBO manufactured in China. Based on the above test results, the Aramid-III was chosen to compound with the basalt. The compound modes include hybrid weaving, affixing the Aramid-III and basalt fabric layers with the glue, and sewing them with the aramid thread. The configurations are impacted by the 5 mm-diameter Al projectile with the velocities of 4.5 km/s to 5.0 km/s. The test results show that the fabric layer containing Aramid-III fabric has equivalent shielding performance with the PBO fabric, and they are better than the other two stuffed layers containing Aramid-II fabric respectively under the equivalent areal density. The performance of stuffed layer is influenced by the adhesion degree of basalt and Aramid-III fibre. The three compound modes all improve the fabric adhesion degree, however, the over adhesion makes against dissipating the kinetic energy of debris cloud, which causes the performance to degrade.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Fa-wei Ke, Jie Huang, Xue-zhong Wen, Qiang Song, Qing Luo, Sen LiuAbstractFour kinds of aramid fibre and three compound modes of aramid with ceramic fibre were studied by hypervelocity impact tests respectively to probe the fabric layer with higher performance in space debris shield configuration. The aramid fibres include the Aramid-II manufactured in America, the Aramid-II, the Aramid-III and the PBO manufactured in China. Based on the above test results, the Aramid-III was chosen to compound with the basalt. The compound modes include hybrid weaving, affixing the Aramid-III and basalt fabric layers with the glue, and sewing them with the aramid thread. The configurations are impacted by the 5 mm-diameter Al projectile with the velocities of 4.5 km/s to 5.0 km/s. The test results show that the fabric layer containing Aramid-III fabric has equivalent shielding performance with the PBO fabric, and they are better than the other two stuffed layers containing Aramid-II fabric respectively under the equivalent areal density. The performance of stuffed layer is influenced by the adhesion degree of basalt and Aramid-III fibre. The three compound modes all improve the fabric adhesion degree, however, the over adhesion makes against dissipating the kinetic energy of debris cloud, which causes the performance to degrade.
- An early history of the Philippine space development program
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Q. Verspieren, G. Coral, B. Pyne, H. RoyAbstractIn 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 20th 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Q. Verspieren, G. Coral, B. Pyne, H. RoyAbstractIn 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 20th 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.
- Vibration suppression control of free-floating space robots with flexible
appendages for autonomous target capturing- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Deshan Meng, Weining Lu, Wenfu Xu, Yu She, Xueqian Wang, Bin Liang, Bo YuanAbstractSome flexible appendages, such as solar panels, communication antenna and other large structures are mounted on the base of the space robot. The structure flexibilities will cause vibration during the operation of manipulators. Due to complicated dynamic coupling among manipulators, the rigid base and flexible appendages, it is very challenging to control the end-effector to track inertial trajectories, especially when the target states are constantly changing. This paper proposes a vibration suppression method for autonomous target capturing during the preimpact phase without controlling the base. Firstly, we derive the rigid-flexible coupling dynamics of a space robot system with flexible appendages. Then, the relationship among joint rates, elastic motion and the end-effector velocities is established by using the linear momentum and angular momentum conservation equations. Secondly, a closed-loop control system is designed based on the dynamic coupling model. And the control system is composed of target motion prediction, autonomous trajectory planning, energy-based joint controllers and so on. Thirdly, the energy-based joint control is detailed, which is proved to be stable by Lyapunov direct method. Finally, simulations of a planar space robot with two flexible appendages and a 3D space robot with single flexible appendage are provided to verify the effectiveness of the presented approach. The effectiveness of energy-based joint control for vibration suppression is verified by a single-degree-of-freedom space robot experimental system. The simulation and experimental results show that the space manipulator can successfully capture the moving target while suppressing the structure vibration.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Deshan Meng, Weining Lu, Wenfu Xu, Yu She, Xueqian Wang, Bin Liang, Bo YuanAbstractSome flexible appendages, such as solar panels, communication antenna and other large structures are mounted on the base of the space robot. The structure flexibilities will cause vibration during the operation of manipulators. Due to complicated dynamic coupling among manipulators, the rigid base and flexible appendages, it is very challenging to control the end-effector to track inertial trajectories, especially when the target states are constantly changing. This paper proposes a vibration suppression method for autonomous target capturing during the preimpact phase without controlling the base. Firstly, we derive the rigid-flexible coupling dynamics of a space robot system with flexible appendages. Then, the relationship among joint rates, elastic motion and the end-effector velocities is established by using the linear momentum and angular momentum conservation equations. Secondly, a closed-loop control system is designed based on the dynamic coupling model. And the control system is composed of target motion prediction, autonomous trajectory planning, energy-based joint controllers and so on. Thirdly, the energy-based joint control is detailed, which is proved to be stable by Lyapunov direct method. Finally, simulations of a planar space robot with two flexible appendages and a 3D space robot with single flexible appendage are provided to verify the effectiveness of the presented approach. The effectiveness of energy-based joint control for vibration suppression is verified by a single-degree-of-freedom space robot experimental system. The simulation and experimental results show that the space manipulator can successfully capture the moving target while suppressing the structure vibration.
- Investigation of RBCC performance improvements based on a variable
geometry ramjet combustor- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Jinying Ye, Hongliang Pan, Fei Qin, Yajun Wang, Duo ZhangAbstractThe use of a variable geometry combustor is one of the most effective methods to improve the performance of a rocket-based combined-cycle (RBCC) engine over a wide range of operating conditions. This paper aims to study the capabilities of a variable geometry combustor operating over a wide range of conditions and determine the performance of the combustor under various inflow conditions. Based on the inflow conditions, the configuration parameters of the combustor were adjusted. Ground direct-connect experiments were conducted under the inflow conditions of Ma 2, Ma 3, Ma 4, and Ma 6, and numerical simulations were performed under the conditions of Ma 3 and Ma 6. The direct-connect experiments showed that the primary rocket operating at a low flow rate can reliably ignite the secondary fuel and maintain a stable and efficient combustion in a variable geometry combustor. Under the inflow conditions of both Ma 4 and Ma 6, smooth transitions of the variable geometry combustor from rocket-ramjet mode to ramjet mode were achieved, and the specific impulses were greatly improved and reached 28.2% and 37.1% of the amplitude, respectively. The numerical simulation showed that the variable geometry combustor can effectively control the combustion heat release region and greatly improve the performance of the combustor. The specific impulse in the combustor increased by 18.6% and 26.2% compared with that in the fixed geometry combustor under Ma 3 and Ma 6 inflow conditions, respectively. It is therefore strongly believed that the variable geometry combustor has significant performance advantages under a wide range of operating conditions.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Jinying Ye, Hongliang Pan, Fei Qin, Yajun Wang, Duo ZhangAbstractThe use of a variable geometry combustor is one of the most effective methods to improve the performance of a rocket-based combined-cycle (RBCC) engine over a wide range of operating conditions. This paper aims to study the capabilities of a variable geometry combustor operating over a wide range of conditions and determine the performance of the combustor under various inflow conditions. Based on the inflow conditions, the configuration parameters of the combustor were adjusted. Ground direct-connect experiments were conducted under the inflow conditions of Ma 2, Ma 3, Ma 4, and Ma 6, and numerical simulations were performed under the conditions of Ma 3 and Ma 6. The direct-connect experiments showed that the primary rocket operating at a low flow rate can reliably ignite the secondary fuel and maintain a stable and efficient combustion in a variable geometry combustor. Under the inflow conditions of both Ma 4 and Ma 6, smooth transitions of the variable geometry combustor from rocket-ramjet mode to ramjet mode were achieved, and the specific impulses were greatly improved and reached 28.2% and 37.1% of the amplitude, respectively. The numerical simulation showed that the variable geometry combustor can effectively control the combustion heat release region and greatly improve the performance of the combustor. The specific impulse in the combustor increased by 18.6% and 26.2% compared with that in the fixed geometry combustor under Ma 3 and Ma 6 inflow conditions, respectively. It is therefore strongly believed that the variable geometry combustor has significant performance advantages under a wide range of operating conditions.
- The structural properties of paraffin wax based hybrid rocket fuels with
aluminium particles- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Kirsty Veale, Sarp Adali, Jean Pitot, Clinton BemontAbstractParaffin wax has been identified as a feasible high regression rate hybrid fuel. For this reason, paraffin wax needs to undergo stringent performance measures to qualify it to meet the requirements necessary for a large scale launch vehicle. Energetic additives such as aluminium powders have been considered for their performance enhancing possibilities. This research focusses on measuring the structural performance of the fuels similar to what would be required for a solid propellant. Thus, structural properties of both pure and 40 W t% aluminised fuels are investigated. Additionally, both elastic and plastic structural properties of the fuel need to be determined for complete material characterisation. Strain rate and temperature dependence of the material structural properties are investigated in this work through compression and tension testing. Results indicate that the addition of aluminium increases the strength of the fuel. Also, a slight increase in temperature was seen to decrease the structural performance significantly. This means that the rate of thermal propagation within the fuel grain is an important consideration. Finally, strain rate dependence is evident in paraffin wax. Higher strain rates result in higher Ultimate Tensile Strength (UTS) failure points, at lower levels of strain.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Kirsty Veale, Sarp Adali, Jean Pitot, Clinton BemontAbstractParaffin wax has been identified as a feasible high regression rate hybrid fuel. For this reason, paraffin wax needs to undergo stringent performance measures to qualify it to meet the requirements necessary for a large scale launch vehicle. Energetic additives such as aluminium powders have been considered for their performance enhancing possibilities. This research focusses on measuring the structural performance of the fuels similar to what would be required for a solid propellant. Thus, structural properties of both pure and 40 W t% aluminised fuels are investigated. Additionally, both elastic and plastic structural properties of the fuel need to be determined for complete material characterisation. Strain rate and temperature dependence of the material structural properties are investigated in this work through compression and tension testing. Results indicate that the addition of aluminium increases the strength of the fuel. Also, a slight increase in temperature was seen to decrease the structural performance significantly. This means that the rate of thermal propagation within the fuel grain is an important consideration. Finally, strain rate dependence is evident in paraffin wax. Higher strain rates result in higher Ultimate Tensile Strength (UTS) failure points, at lower levels of strain.
- Design and high speed aerodynamic performance analysis of vortex lift
waverider with a wide-speed range- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Zhen-tao Zhao, Wei Huang, Bin-bin Yan, Li Yan, Tian-tian Zhang, R. MoradiAbstractAccording to the geometrical relationships of osculating cone waveriders, two kinds of design methods for the constant swept waverider are discussed in the current study, and both of them own more flexible design curves. The correctness and effectiveness of the proposed approaches has been verified by means of the CFD approach. The design Mach numbers of the cuspidal waverider and the general osculating cone waverider with the same volumetric efficiency of the delta-winged waverider have been determined by writing a program for calculating the volumetric efficiencies of waveriders. Furthermore, the high speed aerodynamic performance advantages of the cuspidal and delta-winged waveriders have been analyzed in detail, in comparison with the high speed aerodynamic performance of the general osculating cone waverider. The obtained results show that compared with the general osculating cone waverider, the cuspidal waverider has better high speed aerodynamic performance under any flight condition in the high speed range, while the high speed aerodynamic performance of the delta-winged waverider is worse. In addition, at the high flight Mach number, the cuspidal waverider shows the significant non-linear behavior of the vortex generated lift, while the delta-winged waverider and the general osculating cone waverider do not have this property.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Zhen-tao Zhao, Wei Huang, Bin-bin Yan, Li Yan, Tian-tian Zhang, R. MoradiAbstractAccording to the geometrical relationships of osculating cone waveriders, two kinds of design methods for the constant swept waverider are discussed in the current study, and both of them own more flexible design curves. The correctness and effectiveness of the proposed approaches has been verified by means of the CFD approach. The design Mach numbers of the cuspidal waverider and the general osculating cone waverider with the same volumetric efficiency of the delta-winged waverider have been determined by writing a program for calculating the volumetric efficiencies of waveriders. Furthermore, the high speed aerodynamic performance advantages of the cuspidal and delta-winged waveriders have been analyzed in detail, in comparison with the high speed aerodynamic performance of the general osculating cone waverider. The obtained results show that compared with the general osculating cone waverider, the cuspidal waverider has better high speed aerodynamic performance under any flight condition in the high speed range, while the high speed aerodynamic performance of the delta-winged waverider is worse. In addition, at the high flight Mach number, the cuspidal waverider shows the significant non-linear behavior of the vortex generated lift, while the delta-winged waverider and the general osculating cone waverider do not have this property.
- Improved tracklet association for space objects using short-arc optical
measurements- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Han Cai, Yang Yang, Steve Gehly, Suqin Wu, Kefei ZhangInitial orbit determination (IOD) for space objects is challenging, especially in the case where only optical observations, i.e. angles-only observations, are available and the optical observing arcs are very short (i.e. the too-short arc (TSA) problem). One approach to address the TSA problem is to associate several short-arc tracklets to targets across varying time intervals. In order to achieve better association and run-time performance, this study proposes an improvement to the traditional initial value problem (IVP) solution that determines the association by searching for the global minimum of a new loss function defined in a nonsingular canonical space. The improved IVP method was validated using optical data of space objects at different altitudes collected from the Mount Stromlo Observatory and compared with traditional IVP and another popular tracklet association method: the boundary value problem (BVP) approach. Results illustrate that the improved IVP method is superior to IVP and BVP in terms of association performance, and it also achieves good run-time performance. In addition, traditional methods suffer the drawback of incorrectly associating tracklets from different objects in the same constellation. A new approach dubbed the common ellipse method is presented to address this issue. The common ellipse method is tested with 86 Iridium constellation tracklets, and results show that it significantly improves the true negative rate for the tested scenario.Graphical abstract
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Han Cai, Yang Yang, Steve Gehly, Suqin Wu, Kefei ZhangInitial orbit determination (IOD) for space objects is challenging, especially in the case where only optical observations, i.e. angles-only observations, are available and the optical observing arcs are very short (i.e. the too-short arc (TSA) problem). One approach to address the TSA problem is to associate several short-arc tracklets to targets across varying time intervals. In order to achieve better association and run-time performance, this study proposes an improvement to the traditional initial value problem (IVP) solution that determines the association by searching for the global minimum of a new loss function defined in a nonsingular canonical space. The improved IVP method was validated using optical data of space objects at different altitudes collected from the Mount Stromlo Observatory and compared with traditional IVP and another popular tracklet association method: the boundary value problem (BVP) approach. Results illustrate that the improved IVP method is superior to IVP and BVP in terms of association performance, and it also achieves good run-time performance. In addition, traditional methods suffer the drawback of incorrectly associating tracklets from different objects in the same constellation. A new approach dubbed the common ellipse method is presented to address this issue. The common ellipse method is tested with 86 Iridium constellation tracklets, and results show that it significantly improves the true negative rate for the tested scenario.Graphical abstract
- Investigation on dynamic behaviors of thermal protection system using a
two degree-of-freedom nonlinear theoretical method- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Jie Huang, Weixing Yao, Piao Li, Danfa Zhou, Cheng Chang, Hanyu LinAbstractIn 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Jie Huang, Weixing Yao, Piao Li, Danfa Zhou, Cheng Chang, Hanyu LinAbstractIn 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.
- Structures of near-wall wakes subjected to a sonic jet in a supersonic
crossflow- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Yuan Liu, Ming-bo Sun, Chang-hai Liang, Zun Cai, Ya-nan WangAbstractNanoparticle-based Planar Laser Scattering (NPLS) technology and Oil flow visualization are employed to investigate the wake flow of a sonic jet injected into a supersonic crossflow at Ma = 2.95. Experiments are run for seven jet-to-crossflow momentum flux ratios (J) of 2.3, 5.5, 7.7, 11.2, 16.0, 20.6, and 28.9 of a sonic jet injected into a supersonic crossflow at Ma = 2.95. Experimental results suggest that the near-wall wake zone could be divided into three regions: the V-shape separation region behind the jet, followed by the reattachment region and the mixing and recovery zone further downstream. Lower jet-to-crossflow momentum flux ratio cases always have shorter distance from the orifice to the interaction position of the V-shape collision shock and the reflected shock. The angle between the main separation lines is found to be independent of J. Correlations for predicting the separation length and width are proposed based on the experiments of different momentum flux ratios.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Yuan Liu, Ming-bo Sun, Chang-hai Liang, Zun Cai, Ya-nan WangAbstractNanoparticle-based Planar Laser Scattering (NPLS) technology and Oil flow visualization are employed to investigate the wake flow of a sonic jet injected into a supersonic crossflow at Ma = 2.95. Experiments are run for seven jet-to-crossflow momentum flux ratios (J) of 2.3, 5.5, 7.7, 11.2, 16.0, 20.6, and 28.9 of a sonic jet injected into a supersonic crossflow at Ma = 2.95. Experimental results suggest that the near-wall wake zone could be divided into three regions: the V-shape separation region behind the jet, followed by the reattachment region and the mixing and recovery zone further downstream. Lower jet-to-crossflow momentum flux ratio cases always have shorter distance from the orifice to the interaction position of the V-shape collision shock and the reflected shock. The angle between the main separation lines is found to be independent of J. Correlations for predicting the separation length and width are proposed based on the experiments of different momentum flux ratios.
- Numerical study on separation shock characteristics of pyrotechnic
separation nuts- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Hongda Zhao, Wei Liu, Jifeng Ding, Yi Sun, Xiao Li, Yizhi LiuAbstractThis paper establishes a finite element model of pyrotechnic separation nuts and simulates the whole separation process based on the explicit dynamic codes LS-dyna with Arbitrary Lagrange-Euler (ALE) algorithm. The separation shock generated in the unlocking process is analyzed in detail and the results reveal that the pyrotechnic explosion shock and strain energy shock are two sources for the whole separation shock. Besides that, the influence of prestress on the separation shock and its two shock sources is also researched. The results show that the prestress could strengthen all three shocks, especially the strain energy release shock. This paper also compares the shock response spectrum (SRS) of the three shocks and finds that the SRS of separation shock is equal to the envelope of the SRS of pyrotechnic explosion shock and strain energy release shock. In the model of this paper, the pyrotechnic explosion shock plays a dominant role in the frequency range of 100 Hz to 10 KHz. The conclusions got in this work are helpful to insight the mechanism of separation shock and can provide a reference for the design of separation nuts.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Hongda Zhao, Wei Liu, Jifeng Ding, Yi Sun, Xiao Li, Yizhi LiuAbstractThis paper establishes a finite element model of pyrotechnic separation nuts and simulates the whole separation process based on the explicit dynamic codes LS-dyna with Arbitrary Lagrange-Euler (ALE) algorithm. The separation shock generated in the unlocking process is analyzed in detail and the results reveal that the pyrotechnic explosion shock and strain energy shock are two sources for the whole separation shock. Besides that, the influence of prestress on the separation shock and its two shock sources is also researched. The results show that the prestress could strengthen all three shocks, especially the strain energy release shock. This paper also compares the shock response spectrum (SRS) of the three shocks and finds that the SRS of separation shock is equal to the envelope of the SRS of pyrotechnic explosion shock and strain energy release shock. In the model of this paper, the pyrotechnic explosion shock plays a dominant role in the frequency range of 100 Hz to 10 KHz. The conclusions got in this work are helpful to insight the mechanism of separation shock and can provide a reference for the design of separation nuts.
- Fixed-time attitude tracking control for spacecraft without unwinding
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Bing Huang, Ai-jun Li, Yong Guo, Chang-qing WangAbstractThis paper investigates the fixed-time attitude tracking control problem for rigid spacecraft based on rotation matrix. Two anti-unwinding control schemes are developed such that the desired attitude can be tracked with bounded convergence time regardless of the initial conditions. The first controller is established on basis of a novel sliding mode surface without considering the external disturbance. In addition, the explicit description of the settling time can be provided under this controller. By revising the sliding mode surface and utilizing an adaptive law, the fixed-time stability can still be achieved under the second controller even in the presence of external disturbance. Theoretical analysis and numerical simulations are presented to demonstrate the validity of the proposed controllers.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Bing Huang, Ai-jun Li, Yong Guo, Chang-qing WangAbstractThis paper investigates the fixed-time attitude tracking control problem for rigid spacecraft based on rotation matrix. Two anti-unwinding control schemes are developed such that the desired attitude can be tracked with bounded convergence time regardless of the initial conditions. The first controller is established on basis of a novel sliding mode surface without considering the external disturbance. In addition, the explicit description of the settling time can be provided under this controller. By revising the sliding mode surface and utilizing an adaptive law, the fixed-time stability can still be achieved under the second controller even in the presence of external disturbance. Theoretical analysis and numerical simulations are presented to demonstrate the validity of the proposed controllers.
- A passive camera based determination of a non-cooperative and unknown
satellite's pose and shape- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Renato Volpe, Giovanni B. Palmerini, Marco SabatiniAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Renato Volpe, Giovanni B. Palmerini, Marco SabatiniAbstractThe 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.
- Linear time-varying model predictive control of magnetically actuated
satellites in elliptic orbits- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Jongbum Kim, Youeyun Jung, Hyochoong BangAbstractThree-axis attitude control using only magnetic actuators has been considered a difficult task due to earth's time-varying magnetic field, the inherent under-actuation associated with this task, and constraints on control dipole moments. In this study, a linear time-varying model predictive control approach is applied to magnetically actuated satellites in elliptic low-earth orbits for nadir and inertial pointing. A linear time-varying model predictive control problem is formulated using an augmented state-space model based on small-angle approximations for both nadir and inertial-pointing nonlinear dynamics models. To reduce the on-line computational load, the model predictive control design for an inertial-pointing problem with Laguerre functions is proposed, and exponential data weighting is used to improve the numerically ill-conditioned problem. Nonlinear simulation results demonstrate the effectiveness of the proposed method.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Jongbum Kim, Youeyun Jung, Hyochoong BangAbstractThree-axis attitude control using only magnetic actuators has been considered a difficult task due to earth's time-varying magnetic field, the inherent under-actuation associated with this task, and constraints on control dipole moments. In this study, a linear time-varying model predictive control approach is applied to magnetically actuated satellites in elliptic low-earth orbits for nadir and inertial pointing. A linear time-varying model predictive control problem is formulated using an augmented state-space model based on small-angle approximations for both nadir and inertial-pointing nonlinear dynamics models. To reduce the on-line computational load, the model predictive control design for an inertial-pointing problem with Laguerre functions is proposed, and exponential data weighting is used to improve the numerically ill-conditioned problem. Nonlinear simulation results demonstrate the effectiveness of the proposed method.
- Mission analysis for Earth to Mars-Phobos distant Retrograde Orbits
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Davide Conte, David B. SpencerAbstractThis paper focuses on the trajectory design for missions destined to explore Mars and/or Phobos departing from Low Earth Orbit (LEO) and arriving into a Mars-Phobos Distant Retrograde Orbit (DRO). Lunar DROs are also briefly explored as an alternative departure location. A Mars-Phobos DRO is a relatively stable environment which would make both the surfaces of Mars and Phobos available for a reasonable propellant expenditure. This paper presents the methodology used to compute LEO to Mars-Phobos DRO trajectories and results regarding required C3 at launch, v∞ at arrival, Time-of-Flight (TOF), and total ΔV for various Mars-Phobos DROs using full ephemeris planetary data. The results show that propellant-optimal trajectories from LEO to a specified Mars-Phobos DRO could be used as a staging location between Mars and Phobos. Assuming that refueling is available at the targeted DRO, LEO to Low Mars Orbits (LMO) trajectories would have higher total ΔV due to the additional stop at the Mars-Phobos DRO. However, the aformentioned trajectories would have lower Initial Mass in LEO (IMLEO) and thus a lower gear ratio thanks to the added “pit stop” located at the given DRO. This results in a lower overall spacecraft dry mass that needs to be launched into space from Earth's surface.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Davide Conte, David B. SpencerAbstractThis paper focuses on the trajectory design for missions destined to explore Mars and/or Phobos departing from Low Earth Orbit (LEO) and arriving into a Mars-Phobos Distant Retrograde Orbit (DRO). Lunar DROs are also briefly explored as an alternative departure location. A Mars-Phobos DRO is a relatively stable environment which would make both the surfaces of Mars and Phobos available for a reasonable propellant expenditure. This paper presents the methodology used to compute LEO to Mars-Phobos DRO trajectories and results regarding required C3 at launch, v∞ at arrival, Time-of-Flight (TOF), and total ΔV for various Mars-Phobos DROs using full ephemeris planetary data. The results show that propellant-optimal trajectories from LEO to a specified Mars-Phobos DRO could be used as a staging location between Mars and Phobos. Assuming that refueling is available at the targeted DRO, LEO to Low Mars Orbits (LMO) trajectories would have higher total ΔV due to the additional stop at the Mars-Phobos DRO. However, the aformentioned trajectories would have lower Initial Mass in LEO (IMLEO) and thus a lower gear ratio thanks to the added “pit stop” located at the given DRO. This results in a lower overall spacecraft dry mass that needs to be launched into space from Earth's surface.
- GomX-3+approach&rft.title=Acta+Astronautica&rft.issn=0094-5765&rft.date=&rft.volume=">Mastering operational limitations of LEO satellites – The GomX-3
approach- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Gilles Nies, Marvin Stenger, Jan Krčál, Holger Hermanns, Morten Bisgaard, David Gerhardt, Boudewijn Haverkort, Marijn Jongerden, Kim G. Larsen, Erik R. WognsenAbstractWhen working with space systems the keyword is resources. For a satellite in orbit all resources are sparse and the most critical resource of all is power. It is therefore crucial to have detailed knowledge on how much power is available for an energy harvesting satellite in orbit at every time – especially when in eclipse, where it draws its power from onboard batteries. This paper addresses this problem by a two-step procedure to perform task scheduling for low-earth-orbit (LEO) satellites exploiting formal methods. It combines cost-optimal reachability analyses of priced timed automata networks with a realistic kinetic battery model capable of capturing capacity limits as well as stochastic fluctuations. The procedure is in use for the automatic and resource-optimal day-ahead scheduling of GomX-3, a power-hungry nanosatellite currently orbiting the earth. We explain how this approach has overcome existing problems, has led to improved designs, and has provided new insights.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Gilles Nies, Marvin Stenger, Jan Krčál, Holger Hermanns, Morten Bisgaard, David Gerhardt, Boudewijn Haverkort, Marijn Jongerden, Kim G. Larsen, Erik R. WognsenAbstractWhen working with space systems the keyword is resources. For a satellite in orbit all resources are sparse and the most critical resource of all is power. It is therefore crucial to have detailed knowledge on how much power is available for an energy harvesting satellite in orbit at every time – especially when in eclipse, where it draws its power from onboard batteries. This paper addresses this problem by a two-step procedure to perform task scheduling for low-earth-orbit (LEO) satellites exploiting formal methods. It combines cost-optimal reachability analyses of priced timed automata networks with a realistic kinetic battery model capable of capturing capacity limits as well as stochastic fluctuations. The procedure is in use for the automatic and resource-optimal day-ahead scheduling of GomX-3, a power-hungry nanosatellite currently orbiting the earth. We explain how this approach has overcome existing problems, has led to improved designs, and has provided new insights.
- Uncertainty calculation for spacecraft thermal models using a generalized
SEA method- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Alejandro Gómez-San-Juan, Isabel Pérez-Grande, Angel Sanz-AndrésAbstractUncertainty calculation in spacecraft thermal control and design is generally performed using two methods: Statistical Error Analysis (SEA) and Monte Carlo Simulation (MCS). These two methods present differences both in accuracy and in time of execution. Both features are compared in this paper, and the sources of possible divergence between their results are identified. Having these sources of divergence in mind, a new methodology has been developed. In it, temperature uncertainty is obtained as a linear combination of the probability density functions including non-linear effects due to independent variation of each parameter. In order to compare the new method with the previous ones, all three have been applied to a practical case: the EPD-STEP particle detector onboard ESA's Solar Orbiter mission. In this example this new method is proved to be comparable to SEA in terms of computing time and to MCS in terms of accuracy.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Alejandro Gómez-San-Juan, Isabel Pérez-Grande, Angel Sanz-AndrésAbstractUncertainty calculation in spacecraft thermal control and design is generally performed using two methods: Statistical Error Analysis (SEA) and Monte Carlo Simulation (MCS). These two methods present differences both in accuracy and in time of execution. Both features are compared in this paper, and the sources of possible divergence between their results are identified. Having these sources of divergence in mind, a new methodology has been developed. In it, temperature uncertainty is obtained as a linear combination of the probability density functions including non-linear effects due to independent variation of each parameter. In order to compare the new method with the previous ones, all three have been applied to a practical case: the EPD-STEP particle detector onboard ESA's Solar Orbiter mission. In this example this new method is proved to be comparable to SEA in terms of computing time and to MCS in terms of accuracy.
- Experimental investigation of self-excited combustion instabilities with
injection coupling in a cryogenic rocket combustor- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Wolfgang Armbruster, Justin S. Hardi, Dmitry Suslov, Michael OschwaldAbstractThe cryogenic LOX/H2 multi-injector research combustor “BKD” allows the investigation of high-frequency combustion instabilities under realistic conditions. Two different types of self-excited instabilities were observed. For one instability the underlying coupling mechanism was already identified as LOX injection-driven. The second type of combustion instability was experienced for different operating conditions and is characterized by higher amplitudes, more than 75% of the static chamber pressure. Analysis of the pressure data showed that amplitude and frequency of the acoustic field vary strongly over time, which complicates interpretation of the coupling mechanism. A normalization of the shifting frequency shows that the increase of oscillation frequency depends on transverse acoustic velocity. This is an effect that has also been noticed in other experiments and simulations and is explained by improved mixing leading to a reduced length of the combustion zone. This observation suggests that during the large amplitude pressure oscillations the mode shape remains a first tangential (1T) mode with shifting frequency. By using highly resolved information of the acoustic field in the time domain, both from the combustion chamber and the injector volumes, in combination with acoustic modelling of the injector elements, insights into the coupling mechanism could be gained. For periods of lower amplitudes the pressure oscillations are LOX injector-driven, similar to the first type of instability. With increasing amplitude also the frequency of the unstable mode increases and shifts into a region, where interaction with the hydrogen injector 1L mode becomes possible.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Wolfgang Armbruster, Justin S. Hardi, Dmitry Suslov, Michael OschwaldAbstractThe cryogenic LOX/H2 multi-injector research combustor “BKD” allows the investigation of high-frequency combustion instabilities under realistic conditions. Two different types of self-excited instabilities were observed. For one instability the underlying coupling mechanism was already identified as LOX injection-driven. The second type of combustion instability was experienced for different operating conditions and is characterized by higher amplitudes, more than 75% of the static chamber pressure. Analysis of the pressure data showed that amplitude and frequency of the acoustic field vary strongly over time, which complicates interpretation of the coupling mechanism. A normalization of the shifting frequency shows that the increase of oscillation frequency depends on transverse acoustic velocity. This is an effect that has also been noticed in other experiments and simulations and is explained by improved mixing leading to a reduced length of the combustion zone. This observation suggests that during the large amplitude pressure oscillations the mode shape remains a first tangential (1T) mode with shifting frequency. By using highly resolved information of the acoustic field in the time domain, both from the combustion chamber and the injector volumes, in combination with acoustic modelling of the injector elements, insights into the coupling mechanism could be gained. For periods of lower amplitudes the pressure oscillations are LOX injector-driven, similar to the first type of instability. With increasing amplitude also the frequency of the unstable mode increases and shifts into a region, where interaction with the hydrogen injector 1L mode becomes possible.
- Vibration analysis of functionally graded porous shear deformable tubes
excited by moving distributed loads- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Yuewu Wang, Ke Xie, Tairan FuAbstractThe present work focuses on the free and forced vibrations of the functionally graded (FG) porous tubes subjected to the continuously distributed loads, which move along their axial directions with a constant velocity. The modulus of elasticity of porous composite is assumed to be graded in the radial direction of the tube. The open-cell metal foam provides a typical mechanical feature to determine the relationship between coefficients of density and porosity. A refined beam theory in conjunction with the Lagrange method is employed to derive the governing equations of motion. The Newmark-β method is adopted to obtain the response of the tube in the time domain. Two types of graded porosity distributions are taken into account in the present work. A simply-supported FG porous tube is considered as an example to illustrate the effects of the porosity distribution, geometrical parameters, and load length on the vibration behaviors of the FG porous tube.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Yuewu Wang, Ke Xie, Tairan FuAbstractThe present work focuses on the free and forced vibrations of the functionally graded (FG) porous tubes subjected to the continuously distributed loads, which move along their axial directions with a constant velocity. The modulus of elasticity of porous composite is assumed to be graded in the radial direction of the tube. The open-cell metal foam provides a typical mechanical feature to determine the relationship between coefficients of density and porosity. A refined beam theory in conjunction with the Lagrange method is employed to derive the governing equations of motion. The Newmark-β method is adopted to obtain the response of the tube in the time domain. Two types of graded porosity distributions are taken into account in the present work. A simply-supported FG porous tube is considered as an example to illustrate the effects of the porosity distribution, geometrical parameters, and load length on the vibration behaviors of the FG porous tube.
- Re-planning strategies for space station on-orbit activities executed in
emergencies- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Shuai Mu, Ya-zhong Luo, Dong-yang Qiu, Jun LiangAbstractIn this paper, we develop strategies for the problem of space station on-orbit activity re-planning. Considering the effects of emergencies, a preprocessed strategy is firstly proposed to reconstruct conflicting missions. Furthermore, a greedy algorithm with interval-based strategies is used to re-plan the activities of the reconstructed missions by utilizing mission priority and preplanned activity schedule intervals. Taking into account the propagation of complicated constraints, a time backtrack iteration strategy is adopted to solve constraint conflicts. We demonstrate the proposed approach with a notional operation scenario of the Chinese future space station. The results indicate that our method can successfully re-plan space station on-orbit activities when emergencies occur, all constraint conflicts can be solved, and mission accomplishment rate is increased, in comparison with a method that does not include re-planning strategies.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Shuai Mu, Ya-zhong Luo, Dong-yang Qiu, Jun LiangAbstractIn this paper, we develop strategies for the problem of space station on-orbit activity re-planning. Considering the effects of emergencies, a preprocessed strategy is firstly proposed to reconstruct conflicting missions. Furthermore, a greedy algorithm with interval-based strategies is used to re-plan the activities of the reconstructed missions by utilizing mission priority and preplanned activity schedule intervals. Taking into account the propagation of complicated constraints, a time backtrack iteration strategy is adopted to solve constraint conflicts. We demonstrate the proposed approach with a notional operation scenario of the Chinese future space station. The results indicate that our method can successfully re-plan space station on-orbit activities when emergencies occur, all constraint conflicts can be solved, and mission accomplishment rate is increased, in comparison with a method that does not include re-planning strategies.
- 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 YangAbstractHigh-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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Feng Xu, Xianghong Jia, Wei Lu, Chenglong Zhou, Yaoyu Guo, Jinxue Fei, Chunxin YangAbstractHigh-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 HuAbstractTo 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Zixi Guo, Yao Zhang, Quan HuAbstractTo 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 HuangAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Fengyuan Zuo, Guoping HuangAbstractThe 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. BombardelliAbstractThe “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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): A. Alpatov, S. Khoroshylov, C. BombardelliAbstractThe “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 ChengAbstractCapture 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Shuang Wu, Fangli Mou, Qian Liu, Jing ChengAbstractCapture 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 TeofilattoAbstractMultistage 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Marco Pallone, Mauro Pontani, Paolo TeofilattoAbstractMultistage 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 KangAbstractA 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Xiao Bai, Qinglian Li, Peng Cheng, Liyong Sheng, Zhongtao KangAbstractA 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 TianAbstractA 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Jifei Wang, Chuanzhen Liu, Peng Bai, Jinsheng Cai, Yuan TianAbstractA 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 AnAbstractLarge 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Hao Wu, Meng Liu, Jun Wang, Yuansheng Zhang, Gang AnAbstractLarge 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 AminiAbstractReduction 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): R. Moradi, M. Mosavat, M. Barzegar Gerdroodbary, A. Abdollahi, Younes AminiAbstractReduction 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. LohAbstractThis 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Ghulam Ahmad, Tim W.C. Brown, Craig I. Underwood, Tian H. LohAbstractThis 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 YangAbstractThree 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Rui Zhang, Yuanjin Yu, Chao Han, Zhaohua YangAbstractThree 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. SavinoAbstractThis 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): S. Mungiguerra, G. Zuppardi, L. Spanò Cuomo, R. SavinoAbstractThis 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. FlegelAbstractFragmentation 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): S. Le May, S. Gehly, B.A. Carter, S. FlegelAbstractFragmentation 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 OuAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Yu Zhang, Jianjun Wu, Daixian Zhang, Sheng Tan, Yang OuAbstractThe 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 ErikssonAbstractIn 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Jing Zhang, Hong-wei Guo, Juan Wu, Gui-jun Gao, Zi-ming Kou, Anders ErikssonAbstractIn 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 GillAbstractFuture 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Jason Crusan, Carol Galica, Tracy GillAbstractFuture 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 GhoseAbstractIn 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Sangita Dhara, Scott Burleigh, Raja Datta, Sujoy GhoseAbstractIn 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 SimplícioAbstractWhile 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Eric Joffre, Mattia Zamaro, Nuno Silva, Andrés Marcos, Pedro SimplícioAbstractWhile 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 FengAbstractIn 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Xinyu Wang, Jianguo Cai, Ruiguo Yang, Jian FengAbstractIn 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 BuAbstractThis 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Xiangwei BuAbstractThis 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 QinAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Shuai Wang, Guo-qiang He, De-kun Yan, Zhi-wei Huang, Fei QinAbstractThe 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 WuAbstractThis 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Hao An, Qianqian WuAbstractThis 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. KondyurinAbstractWe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): V.M. Pestrenin, I.V. Pestrenina, S.V. Rusakov, A.V. KondyurinAbstractWe 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 MundtAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Javier García-Garrido, Adrian S. Pudsey, Christian MundtAbstractThe 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 MungiguerraAbstractMINI-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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Alberto Fedele, Stefano MungiguerraAbstractMINI-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 McKilliamAbstractApplications 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): David Haley, Andrew Beck, André Pollok, Alex Grant, Robby McKilliamAbstractApplications 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 MortazaviAbstractHypersonic 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Doyle Knight, Mahsa MortazaviAbstractHypersonic 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 HuangAbstractA 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Zhili Hou, Yunhai Geng, Baolin Wu, Simeng HuangAbstractA 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 DamianAbstractReturning 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Ghassabian G. Hady, Calzada Diaz Abigail, Hettrich Sebastian, De Quattro Nicola, Antonello Andrea, Bielicki DamianAbstractReturning 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 LampaniAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Marco Sabatini, Giovanni B. Palmerini, Matteo Ribet, Paolo Gasbarri, Luca LampaniAbstractThe 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. JinAbstractTethered 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): B.S. Yu, L.L. Geng, H. Wen, T. Chen, D.P. JinAbstractTethered 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 LozanoAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Marco Gomez Jenkins, David Krejci, Paulo LozanoAbstractThe 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 FernandesAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Luiz Arthur Gagg Filho, Sandro da Silva FernandesAbstractThe 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 TsudaAbstractThis 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Tomohiro Yamaguchi, Takanao Saiki, Satoshi Tanaka, Yuto Takei, Tatsuaki Okada, Tadateru Takahashi, Yuichi TsudaAbstractThis 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 SinhaAbstractAn 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Krishna Dev Kumar, Godard, Noel Abreu, Manoranjan SinhaAbstractAn 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 CaiAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Yu-wei Yang, Hong CaiAbstractThe 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. ZhuAbstractThis 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Gangqi Dong, Zheng H. ZhuAbstractThis 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 DengAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Ruiwei Liu, Hongwei Guo, Rongqiang Liu, Hongxiang Wang, Dewei Tang, Zongquan DengAbstractThe 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.
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J 2
perturbations- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): A.C. Perez, D.K. Geller, T.A. LovellAbstractAn 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): A.C. Perez, D.K. Geller, T.A. LovellAbstractAn 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 ZhangAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Haoyang Peng, Weidong Liu, Shijie Liu, Hailong ZhangAbstractThe 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 JanschekAbstractAn 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Simon Delchambre, Tobias Ziegler, Albert Falke, Klaus JanschekAbstractAn 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 WuAbstractIgnition 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Guorui Sun, Zhiwen WuAbstractIgnition 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 ZhangAbstractTo 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Ke Li, Jian Zheng, Jianzhuang Zhi, Kailun ZhangAbstractTo 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 LingAbstractSatellite-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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): James A. Grieve, Robert Bedington, Zhongkan Tang, Rakhitha C.M.R.B. Chandrasekara, Alexander LingAbstractSatellite-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 WangAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Chen Yang, Xinbin Hou, Li WangAbstractThe 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 ChenAbstractA 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Tao Li, Lei ChenAbstractA 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 WuAbstractThis 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Shunan Wu, Weimeng Chu, Xue Ma, Gianmarco Radice, Zhigang WuAbstractThis 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. RattnerAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Christopher J. Greer, Michael V. Paul, Alexander S. RattnerAbstractThe 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 HippkeAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Michael HippkeAbstractThe 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. SilvaAbstractThis 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): G. Loureiro, W.F. Panades, A. SilvaAbstractThis 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 ZhangAbstractLean 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Xiliang Song, Hongbo Wang, Mingbo Sun, Yanxiang ZhangAbstractLean 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. LearnedAbstractWe 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
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Michael Hippke, Paul Leyland, John G. LearnedAbstractWe 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 ChenAbstractSolar 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Yiwu Liu, Zhuhua Si, Liang Tang, Shoulei ChenAbstractSolar 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 BaoAbstractThermoelectric 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Kunlin Cheng, Duo Zhang, Jiang Qin, Silong Zhang, Wen BaoAbstractThermoelectric 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 LiAbstractThe 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Shengbing Zhou, Hu Ma, Yuan Ma, Changsheng Zhou, Daokun Liu, Shuai LiAbstractThe 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 KodairaAbstractTo 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Alexander Miller, Rachid Machrafi, Eric Benton, Hisashi Kitamura, Satoshi KodairaAbstractTo 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 XiongAbstractIn 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.
- Abstract: Publication date: October 2018Source: Acta Astronautica, Volume 151Author(s): Zewei Meng, Xiaoqiang Fan, Yi Wang, Bing XiongAbstractIn 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 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 JieAbstractHigh-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.
- Abstract: Publication date: Available online 11 July 2018Source: Acta AstronauticaAuthor(s): Wang Dingjie, Lv Hanfeng, An Xueying, Wu JieAbstractHigh-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.
- 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 LovellAbstractThis 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.
- Abstract: Publication date: Available online 10 July 2018Source: Acta AstronauticaAuthor(s): Jingwei Wang, Eric A. Butcher, T. Alan LovellAbstractThis 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.
- 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 FangAbstractIn 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.
- Abstract: Publication date: Available online 7 July 2018Source: Acta AstronauticaAuthor(s): Sichen Yuan, Bingen Yang, Houfei FangAbstractIn 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.
- Pros and cons of relativistic interstellar flight
- Abstract: Publication date: Available online 6 July 2018Source: Acta AstronauticaAuthor(s): Oleg G. SemyonovAbstractTwo 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.
- Abstract: Publication date: Available online 6 July 2018Source: Acta AstronauticaAuthor(s): Oleg G. SemyonovAbstractTwo 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. HelmanAbstractConceptualizing 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.
- Abstract: Publication date: Available online 6 July 2018Source: Acta AstronauticaAuthor(s): Daniel S. HelmanAbstractConceptualizing 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 WangAbstractThis 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.
- Abstract: Publication date: Available online 5 July 2018Source: Acta AstronauticaAuthor(s): Qi Li, Jianping Yuan, Huan WangAbstractThis 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.
- 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 QinAbstractThe 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.
- Abstract: Publication date: Available online 5 July 2018Source: Acta AstronauticaAuthor(s): Hongkang Liu, Chao Yan, Yatian Zhao, Yupei QinAbstractThe 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.
- 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 HeAbstractThe 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.
- Abstract: Publication date: Available online 4 July 2018Source: Acta AstronauticaAuthor(s): Guobiao Cai, Hongru Zheng, Lihui Liu, Xiang Ren, Bijiao HeAbstractThe 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. McDowellAbstractIn 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.
- Abstract: Publication date: Available online 3 July 2018Source: Acta AstronauticaAuthor(s): Jonathan C. McDowellAbstractIn 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.
- Optimal injection point for launch trajectories with parametric thrust
profile- Abstract: Publication date: Available online 2 July 2018Source: Acta AstronauticaAuthor(s): Max CerfAbstractThe 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.
- Abstract: Publication date: Available online 2 July 2018Source: Acta AstronauticaAuthor(s): Max CerfAbstractThe 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.
- 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. KochubeiAbstractThis 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.
- 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. KochubeiAbstractThis 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.
