The objective of this study is to determine the two-dimensional unsteady aerodynamic forces and moment acting on a heaving wing in a uniform flow using a wind tunnel. However, it is difficult to measure the aerodynamic forces acting on the heaving wing due to measuring device oscillation and the large inertial force of the wing model. In this study, a new type of wind tunnel test, named ‘‘heaving wind tunnel,’’ was developed. Here, the wing model remains stationary as the wind tunnel oscillates with a heaving motion. The advantage of this experimental method is that the measurement results are unaffected by the large inertial force acting on the oscillating wing model. Therefore, the wing model can be used in the same way as in steady state experiments. The normal force, thrust and pitching moment coefficients of a heaving airfoil were measured using the heaving wind tunnel test developed in this study. Through flow visualizations and pressure measurements, we found that the rapid drop in normal force coefficient after it reached its maximum value was due to a large growing leading-edge vortex.
{"title":"Measurement of Unsteady Aerodynamic Characteristics of a Heaving Wing in a Low Reynolds Number Flow","authors":"M. Okamoto, Shota Fukatsu, D. Sasaki","doi":"10.2322/TJSASS.64.147","DOIUrl":"https://doi.org/10.2322/TJSASS.64.147","url":null,"abstract":"The objective of this study is to determine the two-dimensional unsteady aerodynamic forces and moment acting on a heaving wing in a uniform flow using a wind tunnel. However, it is difficult to measure the aerodynamic forces acting on the heaving wing due to measuring device oscillation and the large inertial force of the wing model. In this study, a new type of wind tunnel test, named ‘‘heaving wind tunnel,’’ was developed. Here, the wing model remains stationary as the wind tunnel oscillates with a heaving motion. The advantage of this experimental method is that the measurement results are unaffected by the large inertial force acting on the oscillating wing model. Therefore, the wing model can be used in the same way as in steady state experiments. The normal force, thrust and pitching moment coefficients of a heaving airfoil were measured using the heaving wind tunnel test developed in this study. Through flow visualizations and pressure measurements, we found that the rapid drop in normal force coefficient after it reached its maximum value was due to a large growing leading-edge vortex.","PeriodicalId":54419,"journal":{"name":"Transactions of the Japan Society for Aeronautical and Space Sciences","volume":"1 1","pages":""},"PeriodicalIF":1.1,"publicationDate":"2021-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"68662625","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper presents an effective maximal-length sequence design for system identification of a continuous-time linear aircraft model. The maximum-length sequence is used as the exciting signal because it is a realistic signal for identifying the aircraft model. This paper proposes two design parameters for the maximum-length sequence, which are related to the dynamical modes of aircraft. According to the identification procedures using the proposed design parameters and the subspace identification method, a continuous-time linear aircraft model in longitudinal motion is well identified in a numerical simulation.
{"title":"An Effective Maximal-Length Sequence Design for System Identification of a Continuous-Time Linear Aircraft Model","authors":"A. Fujimori, S. Oh-hara","doi":"10.2322/tjsass.64.215","DOIUrl":"https://doi.org/10.2322/tjsass.64.215","url":null,"abstract":"This paper presents an effective maximal-length sequence design for system identification of a continuous-time linear aircraft model. The maximum-length sequence is used as the exciting signal because it is a realistic signal for identifying the aircraft model. This paper proposes two design parameters for the maximum-length sequence, which are related to the dynamical modes of aircraft. According to the identification procedures using the proposed design parameters and the subspace identification method, a continuous-time linear aircraft model in longitudinal motion is well identified in a numerical simulation.","PeriodicalId":54419,"journal":{"name":"Transactions of the Japan Society for Aeronautical and Space Sciences","volume":"1 1","pages":""},"PeriodicalIF":1.1,"publicationDate":"2021-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"68663088","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Many works have reported on various sensor network position estimation methods based on the relative distance measurement that can be used when the global navigation satellite system is environmentally denied or degraded.1–3) Among others, trilateration algorithms are widely adopted because of their simple principle.4–6) However, the algorithms possibly fail if the sensors have a low range of communication or the environment includes obstacles.7) Typically, such distance-based localization algorithms are used to construct a globally rigid network.8,9) In other words, albeit each sensor, called node herein, has a limited transmission range, unmanned agents, like unmanned aerial vehicles (UAVs), should be inside the coverage space to receive the sensors’ information.10) Therefore, the typical algorithm requires a network that can adequately cover a certain area and must be capable of communicating with at least three sensors at any point in the area. However, such a network is not always guaranteed. This study proposes a strategy to maximize UAV’s navigation in a sparse wireless sensor network (SWSN) in the manner of the shortest distance travel. The overlapping (or localizable) area, which is calculated using the positions of three disks constructed by the sensor’s transmission range, is used to characterize the possibility of localizing UAVs through trilateration. To ensure that a UAV travels from a starting point to a destination point via the localizable area, it must pass the points that are defined by a sensor set, called vertices. The keys are to find such vertices to define a graph that is flexible to various network complexities that are determined by the combination of sensors and reduce the number of search nodes or the total distance. To determine the shortest path, the Dijkstra algorithm,11,12) one of the most widely used algorithms, is applied with proper modifications. The feasibility of the proposed method is verified through twodimensional (2D) and 3D examples.
{"title":"Efficient Navigation for Unmanned Agents in Sparse Wireless Sensor Networks","authors":"Donghoon Kim","doi":"10.2322/tjsass.64.283","DOIUrl":"https://doi.org/10.2322/tjsass.64.283","url":null,"abstract":"Many works have reported on various sensor network position estimation methods based on the relative distance measurement that can be used when the global navigation satellite system is environmentally denied or degraded.1–3) Among others, trilateration algorithms are widely adopted because of their simple principle.4–6) However, the algorithms possibly fail if the sensors have a low range of communication or the environment includes obstacles.7) Typically, such distance-based localization algorithms are used to construct a globally rigid network.8,9) In other words, albeit each sensor, called node herein, has a limited transmission range, unmanned agents, like unmanned aerial vehicles (UAVs), should be inside the coverage space to receive the sensors’ information.10) Therefore, the typical algorithm requires a network that can adequately cover a certain area and must be capable of communicating with at least three sensors at any point in the area. However, such a network is not always guaranteed. This study proposes a strategy to maximize UAV’s navigation in a sparse wireless sensor network (SWSN) in the manner of the shortest distance travel. The overlapping (or localizable) area, which is calculated using the positions of three disks constructed by the sensor’s transmission range, is used to characterize the possibility of localizing UAVs through trilateration. To ensure that a UAV travels from a starting point to a destination point via the localizable area, it must pass the points that are defined by a sensor set, called vertices. The keys are to find such vertices to define a graph that is flexible to various network complexities that are determined by the combination of sensors and reduce the number of search nodes or the total distance. To determine the shortest path, the Dijkstra algorithm,11,12) one of the most widely used algorithms, is applied with proper modifications. The feasibility of the proposed method is verified through twodimensional (2D) and 3D examples.","PeriodicalId":54419,"journal":{"name":"Transactions of the Japan Society for Aeronautical and Space Sciences","volume":"1 1","pages":""},"PeriodicalIF":1.1,"publicationDate":"2021-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"68663291","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Recently, the application of a deep-learning technique to fl uid analysis has been suggested. Additionally, a deep-learning-based method called the Deep Galerkin Method (DGM) has been suggested for solving a partial di ff erential equation. In DGM, a loss function for training a deep neural network is formulated so that di ff erential operators, boundary conditions, and initial conditions of the targeted partial di ff erential equation are satis fi ed. This study aims to extend and apply DGM to solving compressible Navier-Stokes equations and examine the feasibility of using DGM for fl uid analysis. In this paper, DGM is applied to two-dimensional Burgers equations with periodic boundary conditions, one-dimen-sional Navier-Stokes equations for a shock tube problem, and two-dimensional Navier-Stokes equations for the supersonic fl ow around a blunt body. The approximate solutions obtained using DGM show generally good agreement with that obtained using a fi nite di ff erence method.
{"title":"Application of Deep Galerkin Method to Solve Compressible Navier-Stokes Equations","authors":"M. Matsumoto","doi":"10.2322/tjsass.64.348","DOIUrl":"https://doi.org/10.2322/tjsass.64.348","url":null,"abstract":"Recently, the application of a deep-learning technique to fl uid analysis has been suggested. Additionally, a deep-learning-based method called the Deep Galerkin Method (DGM) has been suggested for solving a partial di ff erential equation. In DGM, a loss function for training a deep neural network is formulated so that di ff erential operators, boundary conditions, and initial conditions of the targeted partial di ff erential equation are satis fi ed. This study aims to extend and apply DGM to solving compressible Navier-Stokes equations and examine the feasibility of using DGM for fl uid analysis. In this paper, DGM is applied to two-dimensional Burgers equations with periodic boundary conditions, one-dimen-sional Navier-Stokes equations for a shock tube problem, and two-dimensional Navier-Stokes equations for the supersonic fl ow around a blunt body. The approximate solutions obtained using DGM show generally good agreement with that obtained using a fi nite di ff erence method.","PeriodicalId":54419,"journal":{"name":"Transactions of the Japan Society for Aeronautical and Space Sciences","volume":"1 1","pages":""},"PeriodicalIF":1.1,"publicationDate":"2021-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"68664111","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
An improved algorithm is proposed to manage with the huge computation burden of the quaternion particle filter in aircraft attitude estimation. Based on the particle filtering frame, the new filter provides robust performance for nonlinear and non-Gaussian stochastic systems. And the posterior distribution of the new estimator is approximated as a new quaternion distribution to realize parallel computation. In addition, similar to the extended Kalman filter, this new method implements time update by replacing particles update with linear transformation to reduce computational complexity. Numerical simulations are carried out to compare the new algorithm to the extended Kalman filter and to quaternion particle filter in simulation results. The simulation results indicate that this estimation technique has faster convergence rate than the extended Kalman filter and takes less computation times than quaternion particle filter under the same accuracy as quaternion particle filter.
{"title":"Attitude Estimation Using Parallel Quaternion Particle Filter Based on New Quaternion Distribution","authors":"Zhaihe Zhou, Zhong Yulu, Chuanwei Zeng, Xiangrui Tian","doi":"10.2322/tjsass.64.249","DOIUrl":"https://doi.org/10.2322/tjsass.64.249","url":null,"abstract":"An improved algorithm is proposed to manage with the huge computation burden of the quaternion particle filter in aircraft attitude estimation. Based on the particle filtering frame, the new filter provides robust performance for nonlinear and non-Gaussian stochastic systems. And the posterior distribution of the new estimator is approximated as a new quaternion distribution to realize parallel computation. In addition, similar to the extended Kalman filter, this new method implements time update by replacing particles update with linear transformation to reduce computational complexity. Numerical simulations are carried out to compare the new algorithm to the extended Kalman filter and to quaternion particle filter in simulation results. The simulation results indicate that this estimation technique has faster convergence rate than the extended Kalman filter and takes less computation times than quaternion particle filter under the same accuracy as quaternion particle filter.","PeriodicalId":54419,"journal":{"name":"Transactions of the Japan Society for Aeronautical and Space Sciences","volume":"1 1","pages":""},"PeriodicalIF":1.1,"publicationDate":"2021-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"68663633","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Efficient operation is important to make full use of the capabilities of China’s space station. Determining the stochastic impacts of emergencies on the operational scenario of the space station is critical for successful implementation. However, few studies have assessed the uncertainties in the operational processes of the space station. To fill this gap, discrete event simulation (DES) is used to develop an evaluation method for the contingent operational plan of a space station. First, DES is used to develop a model framework of the space station operations, and the launch delay of cargo vehicles is introduced into the integrated simulation procedure. Second, the precision of the results and the computational efficiency are improved using the variance reduction technique. The corresponding effect on the number of simulation trials is confirmed using four constraints and three measurable metrics. Finally, the proposed method is applied to a two-year space station operational plan. The results show that maintaining a short interval between the launch date of the cargo vehicles and the start of the launch windows can decrease prolonged duration after a launch delay. A statistical analysis can be used to determine a safe interval between the dates of the events and vehicle launch.
{"title":"Simulation and Evaluation of a Space Station Operational Plan Considering Launch Delay of Cargo Vehicles","authors":"Shuai Guo, Jin Zhang, Ya-zhong Luo","doi":"10.2322/tjsass.64.50","DOIUrl":"https://doi.org/10.2322/tjsass.64.50","url":null,"abstract":"Efficient operation is important to make full use of the capabilities of China’s space station. Determining the stochastic impacts of emergencies on the operational scenario of the space station is critical for successful implementation. However, few studies have assessed the uncertainties in the operational processes of the space station. To fill this gap, discrete event simulation (DES) is used to develop an evaluation method for the contingent operational plan of a space station. First, DES is used to develop a model framework of the space station operations, and the launch delay of cargo vehicles is introduced into the integrated simulation procedure. Second, the precision of the results and the computational efficiency are improved using the variance reduction technique. The corresponding effect on the number of simulation trials is confirmed using four constraints and three measurable metrics. Finally, the proposed method is applied to a two-year space station operational plan. The results show that maintaining a short interval between the launch date of the cargo vehicles and the start of the launch windows can decrease prolonged duration after a launch delay. A statistical analysis can be used to determine a safe interval between the dates of the events and vehicle launch.","PeriodicalId":54419,"journal":{"name":"Transactions of the Japan Society for Aeronautical and Space Sciences","volume":"64 1","pages":"50-62"},"PeriodicalIF":1.1,"publicationDate":"2021-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"68663794","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
We have developed a low-density wind tunnel that simulates Martian atmospheric fl ight on the ground. This wind tunnel employs a supersonic ejector-drive system to realize high-speed fl ow under low-density conditions. This study presents a general evaluation method for the ejector driver of the wind tunnel under low-pressure conditions. As an evaluation parameter for the pressure-recovery ratio, which is a representative value of the driving performance, the ejector-drive parameter (EDP) determined from the design and operating conditions is applied, verifying its e ff ectiveness under atmospheric conditions. Accordingly, we investigate the e ff ectiveness of the EDP at low pressures and its scalability to complex multiple supersonic nozzles. Our results suggest that the pressure-recovery ratio is correlated with the EDP even when the ambient pressure, system con fi guration, and operational conditions change. The EDP allows us to predict the Mach number, and can provide us with an appropriate framework for ejector design optimization.
{"title":"Supersonic Ejector-Driving System under Low Pressure: A Performance Evaluation","authors":"M. Anyoji, D. Numata, H. Nagai, K. Asai","doi":"10.2322/TJSASS.64.156","DOIUrl":"https://doi.org/10.2322/TJSASS.64.156","url":null,"abstract":"We have developed a low-density wind tunnel that simulates Martian atmospheric fl ight on the ground. This wind tunnel employs a supersonic ejector-drive system to realize high-speed fl ow under low-density conditions. This study presents a general evaluation method for the ejector driver of the wind tunnel under low-pressure conditions. As an evaluation parameter for the pressure-recovery ratio, which is a representative value of the driving performance, the ejector-drive parameter (EDP) determined from the design and operating conditions is applied, verifying its e ff ectiveness under atmospheric conditions. Accordingly, we investigate the e ff ectiveness of the EDP at low pressures and its scalability to complex multiple supersonic nozzles. Our results suggest that the pressure-recovery ratio is correlated with the EDP even when the ambient pressure, system con fi guration, and operational conditions change. The EDP allows us to predict the Mach number, and can provide us with an appropriate framework for ejector design optimization.","PeriodicalId":54419,"journal":{"name":"Transactions of the Japan Society for Aeronautical and Space Sciences","volume":"1 1","pages":""},"PeriodicalIF":1.1,"publicationDate":"2021-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"68662713","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
In recent years, there has been a steady increase in the small satellite launch market. With the rapid development of novel launchers, for small satellite owners and operators, how to e ff ectively and e ffi ciently choose appropriate launch vehicles has become a major concern. Based on updated launch records, a reliable launch data source for multi-attribute evaluation and reclassi fi cation is established. Using a statistical classi fi cation process, active launch vehicles are classi fi ed into fi ve representative-in-class launchers on the basis of their capabilities and performance. Unlike the previous categorisation based on payload ability, this method captures launch cost, technology maturity, reliability and availability of each category within the current launch vehicles in service. Moreover, representatives are selected as the baseline types for the high-level planning and designing of complex small satellite launch missions. The analysis indicates that this study provides a valid statistical classi fi cation and selection strategy of representative-in-class launch vehicles to support decision-making for rapid assessment on a large number of small satellite launch missions.
{"title":"Launch Vehicle Classification for Decision-Making of Small Satellite Launch Options","authors":"Mengying Zhang, Q. Xu, Qingbin Zhang","doi":"10.2322/tjsass.64.234","DOIUrl":"https://doi.org/10.2322/tjsass.64.234","url":null,"abstract":"In recent years, there has been a steady increase in the small satellite launch market. With the rapid development of novel launchers, for small satellite owners and operators, how to e ff ectively and e ffi ciently choose appropriate launch vehicles has become a major concern. Based on updated launch records, a reliable launch data source for multi-attribute evaluation and reclassi fi cation is established. Using a statistical classi fi cation process, active launch vehicles are classi fi ed into fi ve representative-in-class launchers on the basis of their capabilities and performance. Unlike the previous categorisation based on payload ability, this method captures launch cost, technology maturity, reliability and availability of each category within the current launch vehicles in service. Moreover, representatives are selected as the baseline types for the high-level planning and designing of complex small satellite launch missions. The analysis indicates that this study provides a valid statistical classi fi cation and selection strategy of representative-in-class launch vehicles to support decision-making for rapid assessment on a large number of small satellite launch missions.","PeriodicalId":54419,"journal":{"name":"Transactions of the Japan Society for Aeronautical and Space Sciences","volume":"1 1","pages":""},"PeriodicalIF":1.1,"publicationDate":"2021-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"68662779","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Yuhei Kikuya, Yohei Iwasaki, Y. Yatsu, S. Matunaga
This paper describes a new algorithm to determine the attitude of micro- / nano-satellites using an Earth sensor. For recent micro- / nano-satellites, the requirements for attitude determination accuracy are becoming more stringent, despite its limited volume. Since Earth sensors have the advantage of smaller size, some studies have presented using them as attitude sensors; however, they could not achieve fully automatic processing in real-time. Therefore, we have developed an algorithm that e ff ectively combines geometrical consideration and image recognition technology, thus realizing high autonomy, robustness, and real-time processing. The validity of this algorithm is con fi rmed through ground experiments. The algorithm operates at a rate of 0.2Hz and achieves an accuracy of 0.1 – 1deg, which is similar to the accuracy of a coarse sun sensor. Furthermore, it is capable of determining the three-axis attitude using only an Earth sensor and a GNSS receiver for position information. This study proves that the bus equipment required for attitude determination systems in micro- / nano-satellites can be reduced, thereby contributing to increased design freedom.
{"title":"Attitude Determination Algorithm Using Earth Sensor Images and Image Recognition","authors":"Yuhei Kikuya, Yohei Iwasaki, Y. Yatsu, S. Matunaga","doi":"10.2322/TJSASS.64.82","DOIUrl":"https://doi.org/10.2322/TJSASS.64.82","url":null,"abstract":"This paper describes a new algorithm to determine the attitude of micro- / nano-satellites using an Earth sensor. For recent micro- / nano-satellites, the requirements for attitude determination accuracy are becoming more stringent, despite its limited volume. Since Earth sensors have the advantage of smaller size, some studies have presented using them as attitude sensors; however, they could not achieve fully automatic processing in real-time. Therefore, we have developed an algorithm that e ff ectively combines geometrical consideration and image recognition technology, thus realizing high autonomy, robustness, and real-time processing. The validity of this algorithm is con fi rmed through ground experiments. The algorithm operates at a rate of 0.2Hz and achieves an accuracy of 0.1 – 1deg, which is similar to the accuracy of a coarse sun sensor. Furthermore, it is capable of determining the three-axis attitude using only an Earth sensor and a GNSS receiver for position information. This study proves that the bus equipment required for attitude determination systems in micro- / nano-satellites can be reduced, thereby contributing to increased design freedom.","PeriodicalId":54419,"journal":{"name":"Transactions of the Japan Society for Aeronautical and Space Sciences","volume":"1 1","pages":""},"PeriodicalIF":1.1,"publicationDate":"2021-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"68664064","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Mariko Akiyama, K. Nishii, Yoshihito Mannami, Masaya Murohara, H. Koizumi, K. Komurasaki
High thrust propulsion systems that can be used in small spacecraft are urgently needed to expand the use of small spacecraft. This study proposes a hybrid thruster for small spacecraft using wire-shaped magnesium as a fuel and water as an oxidizer. “Hybrid thruster” means that the in-space propulsion system generates thrust using a chemical reaction between a solid fuel and a vapor oxidizer, such as those utilized for hybrid rockets. Both magnesium and water are very safe, highly available, and very storable. To assess the feasibility of a hybrid thruster, we carried out experiments utilizing magnesium-wire combustion in water vapor at a pressure lower than the atmosphere, and estimated the input power for ignition, ignitability, and the average magnesium mass consumption rate. Then, propulsion performances were calculated using the experimental results, which show there is an input power of 2–3W, with higher ignitability and a lower average mass consumption rate for thinner magnesium wires. The calculated specific impulse achieved its maximum at 323 s, with a mass mixture ratio of 1.25. At that mass mixture ratio, the calculated thrust using magnesium wires having diameters of 0.5mm and 0.8mm was 49.8mN and 68.5mN, respectively. Additionally, the applicability of the hybrid thruster was shown through a case study considering an existing deep-space misson.
{"title":"Feasibility Study of a Hybrid Thruster using Wire-Shaped Magnesium and Water for Application to Small Spacecraft","authors":"Mariko Akiyama, K. Nishii, Yoshihito Mannami, Masaya Murohara, H. Koizumi, K. Komurasaki","doi":"10.2322/tjsass.64.223","DOIUrl":"https://doi.org/10.2322/tjsass.64.223","url":null,"abstract":"High thrust propulsion systems that can be used in small spacecraft are urgently needed to expand the use of small spacecraft. This study proposes a hybrid thruster for small spacecraft using wire-shaped magnesium as a fuel and water as an oxidizer. “Hybrid thruster” means that the in-space propulsion system generates thrust using a chemical reaction between a solid fuel and a vapor oxidizer, such as those utilized for hybrid rockets. Both magnesium and water are very safe, highly available, and very storable. To assess the feasibility of a hybrid thruster, we carried out experiments utilizing magnesium-wire combustion in water vapor at a pressure lower than the atmosphere, and estimated the input power for ignition, ignitability, and the average magnesium mass consumption rate. Then, propulsion performances were calculated using the experimental results, which show there is an input power of 2–3W, with higher ignitability and a lower average mass consumption rate for thinner magnesium wires. The calculated specific impulse achieved its maximum at 323 s, with a mass mixture ratio of 1.25. At that mass mixture ratio, the calculated thrust using magnesium wires having diameters of 0.5mm and 0.8mm was 49.8mN and 68.5mN, respectively. Additionally, the applicability of the hybrid thruster was shown through a case study considering an existing deep-space misson.","PeriodicalId":54419,"journal":{"name":"Transactions of the Japan Society for Aeronautical and Space Sciences","volume":"1 1","pages":""},"PeriodicalIF":1.1,"publicationDate":"2021-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"68663220","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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