Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943965
M. Lisano, P. Kallemeyn
NASA's InSight Discovery mission to Mars will land its Phoenix-heritage spacecraft to the near-equatorial Elysium Planitia region of Mars in November 2018 — instead of its original planned landing in September 2016 — to collect science measurements over a period longer than one Mars year. Thus, instead of arriving in mid-Mars-global-dust-storm season in 2016 as originally planned, InSight now will arrive in 2018 during the Martian season when dust storms are typically waning. However, it must be able to withstand a global dust storm near the mission's end a Mars year later, by which point dust on the solar arrays is likely to have accumulated significantly more. This paper discusses how the change in launch date has changed the energy management challenges for InSight, and how the energy management approach for surface operations has been adapted to address those challenges. It also describes how energy balance and battery life are protected over the course of the InSight landed mission, in terms of a deliberate balance between autonomous on-board fault protection and ground commanding into reduced-load configurations that still make progress versus specific, prioritized mission success criteria. It describes the project's unique statistical analysis and usage of Mars Exploration Rovers (MER) archived data on solar energy collection to develop and validate an explicit pre-launch margin policy versus energy reductions due to environment variability over multiple-sol sequences. And finally, the paper explains how this archived energy data has influenced the modification of the Phoenix-heritage autonomous fault protection, to guard against quickly-arising inclement power-generation conditions, such as rapid onset of a local dust storm or water ice cloud front.
{"title":"Energy management operations for the Insight solar-powered mission at Mars","authors":"M. Lisano, P. Kallemeyn","doi":"10.1109/AERO.2017.7943965","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943965","url":null,"abstract":"NASA's InSight Discovery mission to Mars will land its Phoenix-heritage spacecraft to the near-equatorial Elysium Planitia region of Mars in November 2018 — instead of its original planned landing in September 2016 — to collect science measurements over a period longer than one Mars year. Thus, instead of arriving in mid-Mars-global-dust-storm season in 2016 as originally planned, InSight now will arrive in 2018 during the Martian season when dust storms are typically waning. However, it must be able to withstand a global dust storm near the mission's end a Mars year later, by which point dust on the solar arrays is likely to have accumulated significantly more. This paper discusses how the change in launch date has changed the energy management challenges for InSight, and how the energy management approach for surface operations has been adapted to address those challenges. It also describes how energy balance and battery life are protected over the course of the InSight landed mission, in terms of a deliberate balance between autonomous on-board fault protection and ground commanding into reduced-load configurations that still make progress versus specific, prioritized mission success criteria. It describes the project's unique statistical analysis and usage of Mars Exploration Rovers (MER) archived data on solar energy collection to develop and validate an explicit pre-launch margin policy versus energy reductions due to environment variability over multiple-sol sequences. And finally, the paper explains how this archived energy data has influenced the modification of the Phoenix-heritage autonomous fault protection, to guard against quickly-arising inclement power-generation conditions, such as rapid onset of a local dust storm or water ice cloud front.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"31 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2017-03-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"128712834","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943876
A. Schmidt, G. Weisz, M. French, T. Flatley, C. Villalpando
The SpaceCubeX project is motivated by the need for high performance, modular, and scalable on-board processing to help scientists answer critical 21st century questions about global climate change, air quality, ocean health, and ecosystem dynamics, while adding new capabilities such as low-latency data products for extreme event warnings. These goals translate into on-board processing throughput requirements that are on the order of 100–1,000x more than those of previous Earth Science missions for standard processing, compression, storage, and downhnk operations. To study possible future architectures to achieve these performance requirements, the SpaceCubeX project provides an evolvable testbed and framework that enables a focused design space exploration of candidate hybrid CPU/FPGA/DSP processing architectures. The framework includes ArchGen, an architecture generator tool populated with candidate architecture components, performance models, and IP cores, that allows an end user to specify the type, number, and connectivity of a hybrid architecture. The framework requires minimal extensions to integrate new processors, such as the anticipated High Performance Spaceflight Computer (HPSC), reducing time to initiate benchmarking by months. To evaluate the framework, we leverage a wide suite of high performance embedded computing benchmarks and Earth science scenarios to ensure robust architecture characterization. We report on our projects Year 1 efforts and demonstrate the capabihties across four simulation testbed models, a baseline SpaceCube 2.0 system, a dual ARM A9 processor system, a hybrid quad ARM A53 and FPGA system, and a hybrid quad ARM A53 and DSP system.
{"title":"SpaceCubeX: A framework for evaluating hybrid multi-core CPU/FPGA/DSP architectures","authors":"A. Schmidt, G. Weisz, M. French, T. Flatley, C. Villalpando","doi":"10.1109/AERO.2017.7943876","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943876","url":null,"abstract":"The SpaceCubeX project is motivated by the need for high performance, modular, and scalable on-board processing to help scientists answer critical 21st century questions about global climate change, air quality, ocean health, and ecosystem dynamics, while adding new capabilities such as low-latency data products for extreme event warnings. These goals translate into on-board processing throughput requirements that are on the order of 100–1,000x more than those of previous Earth Science missions for standard processing, compression, storage, and downhnk operations. To study possible future architectures to achieve these performance requirements, the SpaceCubeX project provides an evolvable testbed and framework that enables a focused design space exploration of candidate hybrid CPU/FPGA/DSP processing architectures. The framework includes ArchGen, an architecture generator tool populated with candidate architecture components, performance models, and IP cores, that allows an end user to specify the type, number, and connectivity of a hybrid architecture. The framework requires minimal extensions to integrate new processors, such as the anticipated High Performance Spaceflight Computer (HPSC), reducing time to initiate benchmarking by months. To evaluate the framework, we leverage a wide suite of high performance embedded computing benchmarks and Earth science scenarios to ensure robust architecture characterization. We report on our projects Year 1 efforts and demonstrate the capabihties across four simulation testbed models, a baseline SpaceCube 2.0 system, a dual ARM A9 processor system, a hybrid quad ARM A53 and FPGA system, and a hybrid quad ARM A53 and DSP system.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"1 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2017-03-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"125687102","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943774
Shawn C. Johnson, A. Pini, D. Reeves, A. S. Martin, Keith Deweese, J. Brophy
Electric propulsion may play a crucial role in the implementation of the gravity tractor planetary defense technique. Gravity tractors were devised to take advantage of the mutual gravitational force between a spacecraft flying in formation with the target celestial body to slowly alter the celestial body's trajectory. No physical contact is necessary, which bypasses issues associated with surface contact such as landing, anchoring, or spin compensation. The gravity tractor maneuver can take several forms, from the originally proposed constant thrust in-line hover to the offset halo orbit. Both can be enhanced with the collection of mass at the asteroid. The form of the gravity tractor ultimately impacts the required thrust magnitude to maintain the formation, as well as constraints on the vectoring of the thrust direction. Solar electric propulsion systems provide an efficient mechanism for tugging the spacecraft-asteroid system due to their high specific impulse. Electric propulsion systems can generate thrust continuously at high efficiency, which is an ideal property for gravity tractors that may require years of operation to achieve the desired deflection because of the very low coupling force provided by the gravitational attraction. The performance and feasibility of the deflection are predicated on having the propulsion capability to maintain the gravity tractor. This paper describes the impacts of constraining the solar electric propulsion thrust magnitude and thrust vectoring capability. It is shown that uncertainty in asteroid density and size, when combined with the enforcement of the electric propulsion constraints, can preclude the feasibility of certain gravity tractor configurations. Additionally, odd thruster configurations are shown to drive the gimbal performance and to have major impacts on eroding incident spacecraft surfaces due to plume interaction. Center of gravity movement further exacerbates issues with gimbaling and plume interaction. A tighter plume divergence angle is therefore always desired, but this paper shows that there is an optimal momentum balance between plume interaction and asteroid-plume avoidance. Several gravity tractor techniques are compared based on metrics of time efficacy, as measured by the induced asteroid delta-V per unit time, and mass efficiency, as measured by the induced asteroid delta-V per unit mass of fuel. Given the propulsion constraints, halo orbits can be infeasible for smaller asteroids unless the mass of the spacecraft is augmented with collected material through a technique called the Enhanced Gravity Tractor. Another proposed method is to alter the halo period by canting the thrusters. In-line hover gravity tractors can always be moved along the net thrust direction to conform to the given propulsion system at the expense of performance, except in the case of smaller asteroids with propulsion systems that are limited in lower throttle range or maximum gimbal angle. Alternative str
{"title":"The effects of constrained electric propulsion on gravity tractors for planetary defense","authors":"Shawn C. Johnson, A. Pini, D. Reeves, A. S. Martin, Keith Deweese, J. Brophy","doi":"10.1109/AERO.2017.7943774","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943774","url":null,"abstract":"Electric propulsion may play a crucial role in the implementation of the gravity tractor planetary defense technique. Gravity tractors were devised to take advantage of the mutual gravitational force between a spacecraft flying in formation with the target celestial body to slowly alter the celestial body's trajectory. No physical contact is necessary, which bypasses issues associated with surface contact such as landing, anchoring, or spin compensation. The gravity tractor maneuver can take several forms, from the originally proposed constant thrust in-line hover to the offset halo orbit. Both can be enhanced with the collection of mass at the asteroid. The form of the gravity tractor ultimately impacts the required thrust magnitude to maintain the formation, as well as constraints on the vectoring of the thrust direction. Solar electric propulsion systems provide an efficient mechanism for tugging the spacecraft-asteroid system due to their high specific impulse. Electric propulsion systems can generate thrust continuously at high efficiency, which is an ideal property for gravity tractors that may require years of operation to achieve the desired deflection because of the very low coupling force provided by the gravitational attraction. The performance and feasibility of the deflection are predicated on having the propulsion capability to maintain the gravity tractor. This paper describes the impacts of constraining the solar electric propulsion thrust magnitude and thrust vectoring capability. It is shown that uncertainty in asteroid density and size, when combined with the enforcement of the electric propulsion constraints, can preclude the feasibility of certain gravity tractor configurations. Additionally, odd thruster configurations are shown to drive the gimbal performance and to have major impacts on eroding incident spacecraft surfaces due to plume interaction. Center of gravity movement further exacerbates issues with gimbaling and plume interaction. A tighter plume divergence angle is therefore always desired, but this paper shows that there is an optimal momentum balance between plume interaction and asteroid-plume avoidance. Several gravity tractor techniques are compared based on metrics of time efficacy, as measured by the induced asteroid delta-V per unit time, and mass efficiency, as measured by the induced asteroid delta-V per unit mass of fuel. Given the propulsion constraints, halo orbits can be infeasible for smaller asteroids unless the mass of the spacecraft is augmented with collected material through a technique called the Enhanced Gravity Tractor. Another proposed method is to alter the halo period by canting the thrusters. In-line hover gravity tractors can always be moved along the net thrust direction to conform to the given propulsion system at the expense of performance, except in the case of smaller asteroids with propulsion systems that are limited in lower throttle range or maximum gimbal angle. Alternative str","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"132 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2017-03-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"121568022","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943881
Alexander W. Raymond, B. Drouin, A. Tang, E. Schlecht, Yanghyo Kim, M. Chang
A new instrument for in situ rotational spectroscopy of gases is presented. The design is based on the pulsed Fourier transform method of Balle-Flygare but operates at higher frequency than traditional microwave implementations. A semi-confocal cavity is an essential part of the new technology, which builds field strength for pumping rotational transitions. Details about the cavity quality factor and design are discussed. The cavity is combined with custom CMOS integrated circuits that synthesize, amplify, and mix the transmitter and receiver signals. Proof-of-concept laboratory measurements of molecular gases are presented. Incorporation in a comet surface sample return mission concept is discussed in detail. The sensor could be used in number of different planetary missions.
{"title":"In situ gas sensing with a 100 GHz CMOS spectrometer","authors":"Alexander W. Raymond, B. Drouin, A. Tang, E. Schlecht, Yanghyo Kim, M. Chang","doi":"10.1109/AERO.2017.7943881","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943881","url":null,"abstract":"A new instrument for in situ rotational spectroscopy of gases is presented. The design is based on the pulsed Fourier transform method of Balle-Flygare but operates at higher frequency than traditional microwave implementations. A semi-confocal cavity is an essential part of the new technology, which builds field strength for pumping rotational transitions. Details about the cavity quality factor and design are discussed. The cavity is combined with custom CMOS integrated circuits that synthesize, amplify, and mix the transmitter and receiver signals. Proof-of-concept laboratory measurements of molecular gases are presented. Incorporation in a comet surface sample return mission concept is discussed in detail. The sensor could be used in number of different planetary missions.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"23 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2017-03-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"132536985","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943648
Brett J. Streetman, J. Shoer, L. Singh
As a spacecraft becomes smaller, a number of physical effects scale both favorably and unfavorably for passive stabilization of the craft. Unfortunately, two separate quantities both scale unfavorably for the use of traditional spinning rotor actuators (e.g. reaction wheels, momentum wheels, control moment gyros) for momentum and attitude control. First, the dominant disturbance torques on small spacecraft in low earth orbit, aerodynamic drag and solar radiation pressure, both become relatively larger as spacecraft size decreases. Second, the effectiveness of spinning rotors reduces as the rotor inertia decreases with the square or the wheel radius. These two factors conspire to greatly reduce the effectiveness of rotor-based momentum control systems at small scales. This reduction requires small spacecraft designers to either devote a significantly larger mass fraction to momentum control or adopt alternative momentum control systems. In this study we examine this problem from two viewpoints. First, empirical data is used to find a relationship between spacecraft size and mass fraction devoted to attitude control. While the International Space Station can devote less than 1% of its mass fraction to momentum control effectors, GEO telecom spacecraft tend to need around 1–2% of available mass, and some CubeSats must devote greater than 50% of their mass fraction. Second, we derive an expression for the smallest spacecraft that can use a reaction wheel for effective momentum management. For reasonable assumptions, this lower limit is on the order of 1 cm length scale, which is in good agreement with the empirical trend. Finally, we list some alternative momentum management strategies and discuss how they apply to spacecraft at the smallest size: the centimeter scale ChipSat.
{"title":"Limitations of scaling momentum control strategies to small spacecraft","authors":"Brett J. Streetman, J. Shoer, L. Singh","doi":"10.1109/AERO.2017.7943648","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943648","url":null,"abstract":"As a spacecraft becomes smaller, a number of physical effects scale both favorably and unfavorably for passive stabilization of the craft. Unfortunately, two separate quantities both scale unfavorably for the use of traditional spinning rotor actuators (e.g. reaction wheels, momentum wheels, control moment gyros) for momentum and attitude control. First, the dominant disturbance torques on small spacecraft in low earth orbit, aerodynamic drag and solar radiation pressure, both become relatively larger as spacecraft size decreases. Second, the effectiveness of spinning rotors reduces as the rotor inertia decreases with the square or the wheel radius. These two factors conspire to greatly reduce the effectiveness of rotor-based momentum control systems at small scales. This reduction requires small spacecraft designers to either devote a significantly larger mass fraction to momentum control or adopt alternative momentum control systems. In this study we examine this problem from two viewpoints. First, empirical data is used to find a relationship between spacecraft size and mass fraction devoted to attitude control. While the International Space Station can devote less than 1% of its mass fraction to momentum control effectors, GEO telecom spacecraft tend to need around 1–2% of available mass, and some CubeSats must devote greater than 50% of their mass fraction. Second, we derive an expression for the smallest spacecraft that can use a reaction wheel for effective momentum management. For reasonable assumptions, this lower limit is on the order of 1 cm length scale, which is in good agreement with the empirical trend. Finally, we list some alternative momentum management strategies and discuss how they apply to spacecraft at the smallest size: the centimeter scale ChipSat.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"81 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2017-03-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"133505429","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943701
Brian J. Barritt, T. Kichkaylo, K. Mandke, Adam Zalcman, Victor Lin
In this paper we describe an application of Tem-porospatial Software Defined Networking (TS-SDN) to UAV networks. Airborne platforms (airplanes, balloons, airships) can be used to carry wireless communication nodes to provide direct-to-user as well as backhaul connections. Such networks also include ground nodes typically equipped with highly directional steerable transceivers. The physics of flight as well as state of the atmosphere lead to time-dynamic link metrics and availability. As nodes move around, the network topology and routing need to adjust to maintain connectivity. Further, mechanical aspects of the system, such as time required to mechanically steer antennas, makes the reactive repair approach more costly than in terrestrial applications. Instead, TS-SDN incorporates reasoning about physical evolution of the system to proactively adjust the network topology in anticipation of future changes. Using airborne networks under development at Google as an example, we discuss the benefits of the TS-SDN approach compared to reactive repair in terms of network availability. We also identify additional constraints one needs to account for when computing the network topology, such as noninterference with other stationary and moving sources. Existing SDN standards do not support scheduled updates necessary in a TS-SDN. We describe our extensions to control messages and software implementation used in field tests.
{"title":"Operating a UAV mesh & internet backhaul network using temporospatial SDN","authors":"Brian J. Barritt, T. Kichkaylo, K. Mandke, Adam Zalcman, Victor Lin","doi":"10.1109/AERO.2017.7943701","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943701","url":null,"abstract":"In this paper we describe an application of Tem-porospatial Software Defined Networking (TS-SDN) to UAV networks. Airborne platforms (airplanes, balloons, airships) can be used to carry wireless communication nodes to provide direct-to-user as well as backhaul connections. Such networks also include ground nodes typically equipped with highly directional steerable transceivers. The physics of flight as well as state of the atmosphere lead to time-dynamic link metrics and availability. As nodes move around, the network topology and routing need to adjust to maintain connectivity. Further, mechanical aspects of the system, such as time required to mechanically steer antennas, makes the reactive repair approach more costly than in terrestrial applications. Instead, TS-SDN incorporates reasoning about physical evolution of the system to proactively adjust the network topology in anticipation of future changes. Using airborne networks under development at Google as an example, we discuss the benefits of the TS-SDN approach compared to reactive repair in terms of network availability. We also identify additional constraints one needs to account for when computing the network topology, such as noninterference with other stationary and moving sources. Existing SDN standards do not support scheduled updates necessary in a TS-SDN. We describe our extensions to control messages and software implementation used in field tests.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"33 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2017-03-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134585289","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943565
A. Mahran
When selecting an error correcting code, it is desired to fulfill a data error rate criterion, but also the code that is selected does this without being excessively complicated. For specific channel conditions it is quite difficult to optimize the error correcting code parameters' analytically. This work proposes multi-objective optimization by applying the Genetic Algorithm (GA) in the selection of Turbo Product Codes (TPC) parameters' that are used for transmission of data over an AWGN channel. The results show that the GA is capable of converging to a set of sensible solutions and giving the pareto-optimum set for error performance against code complexity.
{"title":"Optimizing the parameters of turbo product codes using genetic algorithms","authors":"A. Mahran","doi":"10.1109/AERO.2017.7943565","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943565","url":null,"abstract":"When selecting an error correcting code, it is desired to fulfill a data error rate criterion, but also the code that is selected does this without being excessively complicated. For specific channel conditions it is quite difficult to optimize the error correcting code parameters' analytically. This work proposes multi-objective optimization by applying the Genetic Algorithm (GA) in the selection of Turbo Product Codes (TPC) parameters' that are used for transmission of data over an AWGN channel. The results show that the GA is capable of converging to a set of sensible solutions and giving the pareto-optimum set for error performance against code complexity.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"45 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2017-03-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"130149973","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943789
F. Alibay, J. Kasper, T. Lazio, T. Neilsen
The authors present a space-based array designed to localize and track the radio emission associated with coronal mass ejections (CMEs) from the Sun. Radio emission from CMEs is a direct tracer of the particle acceleration in the inner heliosphere and potential magnetic connections from the lower solar corona to the larger heliosphere. These questions are among those highlighted in the current Solar Decadal Servey, e.g., “Discover and characterize fundamental processes that occur both within the heliosphere and throughout the Universe.” Furthermore, CME radio emission is quite strong, such that only a relatively small number of antennas is required, and a small mission would make a fundamental advancement in our scientific understanding. Indeed, the current state-of-the-art for tracking CME radio emission is defined by single antennas (Wind/WAVES, Stereo/SWAVES) in which the tracking is accomplished by assuming a frequency-to-density mapping. This type of heliophysics mission has been studied several times in the past, but had so far been found to be cost prohibitive, due to the inherent complexity of building multiple spacecraft and flying them in constellation. However, with the increased popularity and success of CubeSat concepts, accompanied by the miniaturization of subsystem components, a range of missions are now being enabled at lower cost than ever before. The paper presents the science requirements for a Small Explorer (SMEX)-class (typically < ∼$100M, including all lifecycle costs) mission concept, and walks through the major features of the SunRISE mission study. SunRISE is composed of six 6U (where 1U is defined as a 10 by 10 by 10cm form-factor) CubeSats placed in an orbit slightly above the Geostationary Equatorial Orbit (GEO) to achieve the aforementioned science goals. The spacecraft fly in a passive formation, which allows them to form an interferometer while minimizing the impact on operations complexity. The paper provides an overview of the mission and spacecraft design, as well as the concept of operations for the mission. Finally, it discusses how the SunRISE mission concept could serve as a stepping stone in demonstrating space-based interferometry and enable more complex mission concepts in the future.
{"title":"Sun radio interferometer space experiment (SunRISE): Tracking particle acceleration and transport in the inner heliosphere","authors":"F. Alibay, J. Kasper, T. Lazio, T. Neilsen","doi":"10.1109/AERO.2017.7943789","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943789","url":null,"abstract":"The authors present a space-based array designed to localize and track the radio emission associated with coronal mass ejections (CMEs) from the Sun. Radio emission from CMEs is a direct tracer of the particle acceleration in the inner heliosphere and potential magnetic connections from the lower solar corona to the larger heliosphere. These questions are among those highlighted in the current Solar Decadal Servey, e.g., “Discover and characterize fundamental processes that occur both within the heliosphere and throughout the Universe.” Furthermore, CME radio emission is quite strong, such that only a relatively small number of antennas is required, and a small mission would make a fundamental advancement in our scientific understanding. Indeed, the current state-of-the-art for tracking CME radio emission is defined by single antennas (Wind/WAVES, Stereo/SWAVES) in which the tracking is accomplished by assuming a frequency-to-density mapping. This type of heliophysics mission has been studied several times in the past, but had so far been found to be cost prohibitive, due to the inherent complexity of building multiple spacecraft and flying them in constellation. However, with the increased popularity and success of CubeSat concepts, accompanied by the miniaturization of subsystem components, a range of missions are now being enabled at lower cost than ever before. The paper presents the science requirements for a Small Explorer (SMEX)-class (typically < ∼$100M, including all lifecycle costs) mission concept, and walks through the major features of the SunRISE mission study. SunRISE is composed of six 6U (where 1U is defined as a 10 by 10 by 10cm form-factor) CubeSats placed in an orbit slightly above the Geostationary Equatorial Orbit (GEO) to achieve the aforementioned science goals. The spacecraft fly in a passive formation, which allows them to form an interferometer while minimizing the impact on operations complexity. The paper provides an overview of the mission and spacecraft design, as well as the concept of operations for the mission. Finally, it discusses how the SunRISE mission concept could serve as a stepping stone in demonstrating space-based interferometry and enable more complex mission concepts in the future.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"14 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2017-03-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"123762947","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943765
A. Ho, Eric Shea, Alan D. George, A. Gordon-Ross
Requirements for higher video quality in space applications continuously calls for increased resolution in imaging sensors, higher bit-depth codecs, more creative solutions for preprocessing and compression techniques, and faster, yet resilient, space-grade platforms. Understanding how these variables interact and affect each other on different platforms is crucial in system development when trying to meet requirements and constraints, such as compression speed, compression ratio (CR), image quality, bandwidth, etc. To analyze this interaction, we present a comparative analysis between compression speed and compression ratio using serial and parallel compression codes on different platforms and architectures, focusing upon video data from overhead-persistent infrared (OPIR) sensors on spacecraft. Previous research allowed us to compare CR and image quality with new preprocessing techniques, but it did not evaluate and address the challenges of compression speed on space-grade processors. Performance is critical, since of course the preprocessing and compression codes plus downlink of compressed data must require less total time than downlink of the raw data, in order for compression to be fully effective.
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Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943953
S. Herzig, S. Mandutianu, Hongman Kim, S. Hernandez, T. Imken
In this paper, a model-based approach to exploring the trade space of multi-spacecraft mission architectures is introduced. Missions involving multiple spacecraft are inherently more complex to design than traditional single spacecraft missions. This is particularly true for fractionated mission concepts, where spacecraft have diverse roles and distributed responsibilities, and fulfilling mission goals requires communication and collaboration. In practice, this complexity and the lack of computational models and tools limit design teams to consider small design spaces and force them to quickly fixate on a single mission design. However, design fixation at such early stages often leads to sub-optimal designs. Towards overcoming this limitation, we propose an automated approach to exploring and visualizing the trade space of multi-spacecraft mission architectures that aids users in decision making, and provides a basis for identifying solutions that are Pareto-optimal with respect to user-defined science requirements, technical and resource constraints, and mission objectives. Central to our approach is the automated synthesis and analysis of mission architecture alternatives from a set of user-provided functional requirements and mission goals, as well as a library of spacecraft components and analysis models. Design rules (synthesis knowledge) are provided in the form of model-transformation rules. Sequences of endogenous model transformations are applied in-place to produce sets of candidate solutions, thereby effectively searching the design space. The search process is guided by the specified objectives, and is implemented using evolutionary algorithms. We demonstrate our approach to architectural optimization using a simplified radio interferometry mission design as a case study. We conclude that using the proposed approach, the number and diversity of candidate mission architectures available for consideration can be significantly increased. Furthermore, the automated synthesis and evaluation of mission architectures can lead to emergent and non-intuitive solutions to be discovered.
{"title":"Model-transformation-based computational design synthesis for mission architecture optimization","authors":"S. Herzig, S. Mandutianu, Hongman Kim, S. Hernandez, T. Imken","doi":"10.1109/AERO.2017.7943953","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943953","url":null,"abstract":"In this paper, a model-based approach to exploring the trade space of multi-spacecraft mission architectures is introduced. Missions involving multiple spacecraft are inherently more complex to design than traditional single spacecraft missions. This is particularly true for fractionated mission concepts, where spacecraft have diverse roles and distributed responsibilities, and fulfilling mission goals requires communication and collaboration. In practice, this complexity and the lack of computational models and tools limit design teams to consider small design spaces and force them to quickly fixate on a single mission design. However, design fixation at such early stages often leads to sub-optimal designs. Towards overcoming this limitation, we propose an automated approach to exploring and visualizing the trade space of multi-spacecraft mission architectures that aids users in decision making, and provides a basis for identifying solutions that are Pareto-optimal with respect to user-defined science requirements, technical and resource constraints, and mission objectives. Central to our approach is the automated synthesis and analysis of mission architecture alternatives from a set of user-provided functional requirements and mission goals, as well as a library of spacecraft components and analysis models. Design rules (synthesis knowledge) are provided in the form of model-transformation rules. Sequences of endogenous model transformations are applied in-place to produce sets of candidate solutions, thereby effectively searching the design space. The search process is guided by the specified objectives, and is implemented using evolutionary algorithms. We demonstrate our approach to architectural optimization using a simplified radio interferometry mission design as a case study. We conclude that using the proposed approach, the number and diversity of candidate mission architectures available for consideration can be significantly increased. Furthermore, the automated synthesis and evaluation of mission architectures can lead to emergent and non-intuitive solutions to be discovered.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"33 22","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2017-03-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"131776036","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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