Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943683
M. Ritter, D. Barnhart
Applications of electroadhesion include automation and inspection robots, consumer gripper devices, anchoring tools used in the military and biomedical industry, and more recently, mechanisms for spacecraft docking. The purpose of this study is to characterize geometries of electroadhesion samples for application in spacecraft docking and propose a metric to predict the interaction between geometry and captured object. Shear forces of electroadhesion samples composed of Kapton(R)Polyimide insulating material with embedded aluminum foil electrodes and three common space-rated substrate materials were measured. Responses of the electroadhesion samples configured in three geometries were identified using substrates attached to dynamic two-dimensional air bearing platforms. Geometries included a flat plate design as a prototype for cubesats, a concave, cylindrical design for potential application to circular, cylindrical spacecraft capture and torque mitigation, and a soft four-arm claw design as a prototype for docking to variable shaped objects with full coverage of object surface area. Quantitative and qualitative results were analyzed to characterize the optimal geometry for spacecraft docking. Surface area of each geometry, defined as the area of contact between electroadhesion samples implemented on the geometry and the substrate rigidly attached on air bearing platform, was compared to the stop time, defined as the time required for the geometry to mitigate both initial and residual motion of the air bearing platform. In summary, aluminized mylar substrate is identified as a superior type to achieve the highest attainable shear adhesion forces, and one electroadhesion geometry may be superior to others depending on specific docking scenarios in a space environment in agreement with the proposed metric.
{"title":"Geometry characterization of electroadhesion samples for spacecraft docking application","authors":"M. Ritter, D. Barnhart","doi":"10.1109/AERO.2017.7943683","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943683","url":null,"abstract":"Applications of electroadhesion include automation and inspection robots, consumer gripper devices, anchoring tools used in the military and biomedical industry, and more recently, mechanisms for spacecraft docking. The purpose of this study is to characterize geometries of electroadhesion samples for application in spacecraft docking and propose a metric to predict the interaction between geometry and captured object. Shear forces of electroadhesion samples composed of Kapton(R)Polyimide insulating material with embedded aluminum foil electrodes and three common space-rated substrate materials were measured. Responses of the electroadhesion samples configured in three geometries were identified using substrates attached to dynamic two-dimensional air bearing platforms. Geometries included a flat plate design as a prototype for cubesats, a concave, cylindrical design for potential application to circular, cylindrical spacecraft capture and torque mitigation, and a soft four-arm claw design as a prototype for docking to variable shaped objects with full coverage of object surface area. Quantitative and qualitative results were analyzed to characterize the optimal geometry for spacecraft docking. Surface area of each geometry, defined as the area of contact between electroadhesion samples implemented on the geometry and the substrate rigidly attached on air bearing platform, was compared to the stop time, defined as the time required for the geometry to mitigate both initial and residual motion of the air bearing platform. In summary, aluminized mylar substrate is identified as a superior type to achieve the highest attainable shear adhesion forces, and one electroadhesion geometry may be superior to others depending on specific docking scenarios in a space environment in agreement with the proposed metric.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"6 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":"114697754","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.2017.7943809
Kshitij Sadasivan, Srinivasan N. Shalini, B. Cheela, Nirav Annavarapu
This paper describes the simulations, practical tests and analysis carried out for monopole and dipole antennas for nano-satellites. The antennas are designed for a 2U nano-satellite. With the monopole and dipole antenna designed to operate in the amateur VHF & UHF bands respectively. The antennas are made of steel tapes, which are obtained from measuring tapes. The antennas have a width of 6mm and thickness 0.2mm. The length of the monopole is 570mm and that of the dipole is 203mm for each arm, with a feed gap of 11mm. The paper further describes the simulations and modelling carried out for the antennas using a CAD software: Computer Simulation Technology (CST). A thermal simulation was done using the System Assembly and Modeling (SAM) module of the CST software to understand the effects of the varied temperature range in space on the antennas. After the intended design of the antennas is achieved using the CAD software, the antennas are practically tested, and the antenna length is altered to obtain the required results. For the intended frequency of operation, measured experimentally, the gain of the monopole and dipole antenna towards the earth facing side is −0.47dBi and −0.8dBi respectively. The return loss of the antennas was experimentally measured and found to be −30.836dB for the monopole antenna and −28.672dB for the dipole antenna. The paper also analyzes the effect of thermal protection tape on the antennas.
{"title":"Design and analysis of antennas for a nano-satellite","authors":"Kshitij Sadasivan, Srinivasan N. Shalini, B. Cheela, Nirav Annavarapu","doi":"10.1109/AERO.2017.7943809","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943809","url":null,"abstract":"This paper describes the simulations, practical tests and analysis carried out for monopole and dipole antennas for nano-satellites. The antennas are designed for a 2U nano-satellite. With the monopole and dipole antenna designed to operate in the amateur VHF & UHF bands respectively. The antennas are made of steel tapes, which are obtained from measuring tapes. The antennas have a width of 6mm and thickness 0.2mm. The length of the monopole is 570mm and that of the dipole is 203mm for each arm, with a feed gap of 11mm. The paper further describes the simulations and modelling carried out for the antennas using a CAD software: Computer Simulation Technology (CST). A thermal simulation was done using the System Assembly and Modeling (SAM) module of the CST software to understand the effects of the varied temperature range in space on the antennas. After the intended design of the antennas is achieved using the CAD software, the antennas are practically tested, and the antenna length is altered to obtain the required results. For the intended frequency of operation, measured experimentally, the gain of the monopole and dipole antenna towards the earth facing side is −0.47dBi and −0.8dBi respectively. The return loss of the antennas was experimentally measured and found to be −30.836dB for the monopole antenna and −28.672dB for the dipole antenna. The paper also analyzes the effect of thermal protection tape on the antennas.","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":"128653036","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.2017.7943921
A. Shirazi
This paper proposes a systematic direct approach to carry out effective multi-objective optimization of space orbit transfer with high-level thrust acceleration. The objective is to provide a transfer trajectory with acceptable accuracy in all orbital parameters while minimizing spacecraft fuel consumption. With direct control parameterization, in which the steering angles of thrust vector are interpolated through a finite number of nodes, the optimal control problem is converted into the parameter optimization problem to be solved by nonlinear programming. Besides the thrust vector direction angles, the thrust magnitude is also considered as variable and unknown along with initial conditions. Since the deviation of thrust vector in spacecraft is limited in reality, mathematical modeling of thrust vector direction is carried out in order to satisfy constraints in maximum deviation of thrust vector direction angles. In this modeling, the polynomial function of each steering angle is defined by interpolation of a curve based on finite number of points in a specific range with a nominal center point with uniform distribution. This kind of definition involves additional parameters to the optimization problem which results the capability of search method in satisfying constraint on the variation of thrust direction angles. Thrust profile is also modeled based on polynomial functions of time with respect to solid and liquid propellant rockets. Imperialist competitive algorithm is used in order to find optimal coefficients of polynomial for thrust vector and the optimal initial states within the transfer. Results are mainly affected by the degree of polynomials involved in mathematical modeling of steering angles and thrust profile which results different optimal initial states where the transfer begins. It is shown that the proposed method is fairly beneficial in the viewpoint of optimality and convergence. The optimality of the technique is shown by comparing the finite thrust optimization with the impulsive analysis. Comparison shows that the accuracy is acceptable with respect to fair precision in orbital elements and minimum fuel mass. Also, the convergence of the optimization algorithm is investigated by comparing the solution of the problem with other optimization techniques such as Genetic Algorithm. Results confirms the practicality of Imperialist Competitive Algorithm in finding optimum variation of thrust vector which results best transfer accuracy along with minimizing fuel consumption.
{"title":"Multi-objective optimization of orbit transfer trajectory using imperialist competitive algorithm","authors":"A. Shirazi","doi":"10.1109/AERO.2017.7943921","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943921","url":null,"abstract":"This paper proposes a systematic direct approach to carry out effective multi-objective optimization of space orbit transfer with high-level thrust acceleration. The objective is to provide a transfer trajectory with acceptable accuracy in all orbital parameters while minimizing spacecraft fuel consumption. With direct control parameterization, in which the steering angles of thrust vector are interpolated through a finite number of nodes, the optimal control problem is converted into the parameter optimization problem to be solved by nonlinear programming. Besides the thrust vector direction angles, the thrust magnitude is also considered as variable and unknown along with initial conditions. Since the deviation of thrust vector in spacecraft is limited in reality, mathematical modeling of thrust vector direction is carried out in order to satisfy constraints in maximum deviation of thrust vector direction angles. In this modeling, the polynomial function of each steering angle is defined by interpolation of a curve based on finite number of points in a specific range with a nominal center point with uniform distribution. This kind of definition involves additional parameters to the optimization problem which results the capability of search method in satisfying constraint on the variation of thrust direction angles. Thrust profile is also modeled based on polynomial functions of time with respect to solid and liquid propellant rockets. Imperialist competitive algorithm is used in order to find optimal coefficients of polynomial for thrust vector and the optimal initial states within the transfer. Results are mainly affected by the degree of polynomials involved in mathematical modeling of steering angles and thrust profile which results different optimal initial states where the transfer begins. It is shown that the proposed method is fairly beneficial in the viewpoint of optimality and convergence. The optimality of the technique is shown by comparing the finite thrust optimization with the impulsive analysis. Comparison shows that the accuracy is acceptable with respect to fair precision in orbital elements and minimum fuel mass. Also, the convergence of the optimization algorithm is investigated by comparing the solution of the problem with other optimization techniques such as Genetic Algorithm. Results confirms the practicality of Imperialist Competitive Algorithm in finding optimum variation of thrust vector which results best transfer accuracy along with minimizing fuel consumption.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"34 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":"125394208","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.2017.7943889
P. Wercinski, B. Smith, B. Yount, Carl Kruger, Chad A. Brivkalns, A. Makino, A. Cassell, S. Dutta, Shakib Ghassemieh, Shang Wu, S. Battazzo, O. Nishioka, E. Venkatapathy, G. Swanson
The Adaptable, Deployable Entry and Placement Technology (ADEPT) architecture represents a novel approach for entry vehicle design utilizing a high performance carbon-fabric to serve as the primary drag surface of the mechanically deployed decelerator. The ADEPT project team is advancing this decelerator technology via systems-level testing at the one-meter diameter (nano-ADEPT) scale. A subsonic aeroloads test (May 2015) and an arc-jet aeroheating test (Sept 2015) have already been completed. The initial system-level development of the nano-ADEPT architecture will culminate in the launch of a 0.7 meter deployed diameter ADEPT sounding rocket flight experiment named, SR-1. Launch is planned for August 2017. The test will utilize the NASA Flight Opportunities Program sounding rocket platform provided by UP Aerospace to launch SR-1 to an apogee over 100 km and achieve re-entry conditions with a peak velocity near Mach 3. The SR-1 flight experiment will demonstrate most of the primary end-to-end mission stages including: launch in a stowed configuration, separation and deployment in exo-atmospheric conditions, and passive ballistic re-entry of a 70-degree half-angle faceted cone geometry. ADEPT SR-1 will determine supersonic through transonic aerodynamic stability of the unique ADEPT blunt body shape with an open-back entry vehicle configuration.
{"title":"ADEPT sounding rocket one (SR-1) flight experiment overview","authors":"P. Wercinski, B. Smith, B. Yount, Carl Kruger, Chad A. Brivkalns, A. Makino, A. Cassell, S. Dutta, Shakib Ghassemieh, Shang Wu, S. Battazzo, O. Nishioka, E. Venkatapathy, G. Swanson","doi":"10.1109/AERO.2017.7943889","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943889","url":null,"abstract":"The Adaptable, Deployable Entry and Placement Technology (ADEPT) architecture represents a novel approach for entry vehicle design utilizing a high performance carbon-fabric to serve as the primary drag surface of the mechanically deployed decelerator. The ADEPT project team is advancing this decelerator technology via systems-level testing at the one-meter diameter (nano-ADEPT) scale. A subsonic aeroloads test (May 2015) and an arc-jet aeroheating test (Sept 2015) have already been completed. The initial system-level development of the nano-ADEPT architecture will culminate in the launch of a 0.7 meter deployed diameter ADEPT sounding rocket flight experiment named, SR-1. Launch is planned for August 2017. The test will utilize the NASA Flight Opportunities Program sounding rocket platform provided by UP Aerospace to launch SR-1 to an apogee over 100 km and achieve re-entry conditions with a peak velocity near Mach 3. The SR-1 flight experiment will demonstrate most of the primary end-to-end mission stages including: launch in a stowed configuration, separation and deployment in exo-atmospheric conditions, and passive ballistic re-entry of a 70-degree half-angle faceted cone geometry. ADEPT SR-1 will determine supersonic through transonic aerodynamic stability of the unique ADEPT blunt body shape with an open-back entry vehicle configuration.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"48 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":"122993606","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.2017.7943704
Alison Gibson, Andrea K. Webb, L. Stirling
The future of human space exploration will involve extra-vehicular activities (EVA) on foreign planetary surfaces (i.e. Mars), an activity that will have significantly different characteristics than the common exploration scenarios on Earth. These activities become challenging due to restricted visual cues and other limitations placed on sensory feedback from altered gravity and the pressurized suit. The use of a bulky, pressurized EVA suit perceptually disconnects human explorers from the hostile foreign environment, increasing the navigation workload and risk of collision associated with traversing through unfamiliar terrain. Due to the hazardous nature of this work, there is a critical need to design multimodal interfaces for optimizing task performance and minimizing risks; in particular, an information presentation device that can aid in obstacle avoidance during surface exploration and way-finding. Previous research has shown that multimodal cues can communicate risk more efficiently than cues to a single modality. This paper presents a wearable interface system to examine human performance when visual, vibratory, and visual-vibratory cues are provided to aid ground obstacle avoidance. The wearable system applies vibro-tactile cues to the feet and visual cues through augmented reality glasses to convey obstacle location and proximity during an approach. This study examined participants stepping over a randomly placed obstacle in a path while wearing the multimodal interface. Measures of performance included path completion time, subjective workload, head-down time, collisions, as well as gait parameters. Differences in obstacle avoidance performance were analyzed across conditions and results provide implications for presenting multimodal information during active tasks such as obstacle avoidance.
{"title":"Analysis of a wearable, multi-modal information presentation device for obstacle avoidance","authors":"Alison Gibson, Andrea K. Webb, L. Stirling","doi":"10.1109/AERO.2017.7943704","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943704","url":null,"abstract":"The future of human space exploration will involve extra-vehicular activities (EVA) on foreign planetary surfaces (i.e. Mars), an activity that will have significantly different characteristics than the common exploration scenarios on Earth. These activities become challenging due to restricted visual cues and other limitations placed on sensory feedback from altered gravity and the pressurized suit. The use of a bulky, pressurized EVA suit perceptually disconnects human explorers from the hostile foreign environment, increasing the navigation workload and risk of collision associated with traversing through unfamiliar terrain. Due to the hazardous nature of this work, there is a critical need to design multimodal interfaces for optimizing task performance and minimizing risks; in particular, an information presentation device that can aid in obstacle avoidance during surface exploration and way-finding. Previous research has shown that multimodal cues can communicate risk more efficiently than cues to a single modality. This paper presents a wearable interface system to examine human performance when visual, vibratory, and visual-vibratory cues are provided to aid ground obstacle avoidance. The wearable system applies vibro-tactile cues to the feet and visual cues through augmented reality glasses to convey obstacle location and proximity during an approach. This study examined participants stepping over a randomly placed obstacle in a path while wearing the multimodal interface. Measures of performance included path completion time, subjective workload, head-down time, collisions, as well as gait parameters. Differences in obstacle avoidance performance were analyzed across conditions and results provide implications for presenting multimodal information during active tasks such as obstacle avoidance.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"43 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":"126682306","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.2017.7943736
B. Kantsiper
Mitigation of a hazardous NEO can be accomplished by deflecting it so that it misses the Earth. Strategies to deflect an asteroid include impacting it with a spacecraft (a kinetic impactor), pulling it with the gravity of the mass of a spacecraft (a gravity tractor), using the blast of a nearby nuclear explosion, and modifying the surface or causing ablation by various means including lasers or particle beams. None of these approaches has been tested on a NEO. The AIDA mission is a proposed international collaboration to demonstrate kinetic deflection, the most mature technique for mitigating the impact hazard of a Near Earth Object (NEO). AIDA consists of two mission elements, the NASA Double Asteroid Redirection Test (DART) mission and the ESA Asteroid Impact Mission (AIM). The main objectives of the DART mission, which includes the spacecraft kinetic impact and an Earth-based observing campaign, are to: • Perform a full scale demonstration of the spacecraft kinetic impact technique for deflection of an asteroid, by targeting an object large enough to qualify as a Potentially Hazardous Asteroid (that is, larger than 100 m); • Measure the resulting asteroid deflection, by targeting the secondary member of a binary NEO and measuring the period change of the binary orbit; • Understand the hypervelocity collision effects on an asteroid, including the long-term dynamics of impact ejecta; validate models for momentum transfer in asteroid impacts, inferring physical properties of the asteroid surface and sub-surface. The DART target is the secondary member of the binary asteroid 65803 Didymos, with the impact scheduled to occur in September, 2022. The DART impact on the secondary member of the Didymos binary at ∼7 km/s will alter the binary orbit period by at least 4 minutes, assuming a simple transfer of momentum to the target. The period change may be significantly greater, as the momentum transferred to the target asteroid may exceed the incident momentum of the kinetic impactor, possibly by a large factor. The AIM spacecraft will characterize the asteroid target and monitor results of the impact in situ at Didymos, but the period change can be determined accurately solely with ground-based observatories, an approach that is only feasible because of the choice of a binary system as target. DART held its Mission Concept Review on May 21, 2015. At MCR, the DART concept had only expensive and potentially risky launch options. During Phase A, the project explored the possibility of being a secondary payload on a commercial Geosynchronous Transfer Orbit (GTO) launch with electric propulsion (EP) as an approach to reduce mission cost, eliminate the launch vehicle risk, and demonstrate the NASA Evolutionary Xenon Thruster (NEXT) engine. NASA determined that DART would use this approach, and the EP-based concept was presented at the DART System Requirements Review on Aug 30, 2016. This paper summarizes the trade that resulted in adoption of the new design.
{"title":"The Double Asteroid Redirection Test (DART) mission electric propulsion trade","authors":"B. Kantsiper","doi":"10.1109/AERO.2017.7943736","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943736","url":null,"abstract":"Mitigation of a hazardous NEO can be accomplished by deflecting it so that it misses the Earth. Strategies to deflect an asteroid include impacting it with a spacecraft (a kinetic impactor), pulling it with the gravity of the mass of a spacecraft (a gravity tractor), using the blast of a nearby nuclear explosion, and modifying the surface or causing ablation by various means including lasers or particle beams. None of these approaches has been tested on a NEO. The AIDA mission is a proposed international collaboration to demonstrate kinetic deflection, the most mature technique for mitigating the impact hazard of a Near Earth Object (NEO). AIDA consists of two mission elements, the NASA Double Asteroid Redirection Test (DART) mission and the ESA Asteroid Impact Mission (AIM). The main objectives of the DART mission, which includes the spacecraft kinetic impact and an Earth-based observing campaign, are to: • Perform a full scale demonstration of the spacecraft kinetic impact technique for deflection of an asteroid, by targeting an object large enough to qualify as a Potentially Hazardous Asteroid (that is, larger than 100 m); • Measure the resulting asteroid deflection, by targeting the secondary member of a binary NEO and measuring the period change of the binary orbit; • Understand the hypervelocity collision effects on an asteroid, including the long-term dynamics of impact ejecta; validate models for momentum transfer in asteroid impacts, inferring physical properties of the asteroid surface and sub-surface. The DART target is the secondary member of the binary asteroid 65803 Didymos, with the impact scheduled to occur in September, 2022. The DART impact on the secondary member of the Didymos binary at ∼7 km/s will alter the binary orbit period by at least 4 minutes, assuming a simple transfer of momentum to the target. The period change may be significantly greater, as the momentum transferred to the target asteroid may exceed the incident momentum of the kinetic impactor, possibly by a large factor. The AIM spacecraft will characterize the asteroid target and monitor results of the impact in situ at Didymos, but the period change can be determined accurately solely with ground-based observatories, an approach that is only feasible because of the choice of a binary system as target. DART held its Mission Concept Review on May 21, 2015. At MCR, the DART concept had only expensive and potentially risky launch options. During Phase A, the project explored the possibility of being a secondary payload on a commercial Geosynchronous Transfer Orbit (GTO) launch with electric propulsion (EP) as an approach to reduce mission cost, eliminate the launch vehicle risk, and demonstrate the NASA Evolutionary Xenon Thruster (NEXT) engine. NASA determined that DART would use this approach, and the EP-based concept was presented at the DART System Requirements Review on Aug 30, 2016. This paper summarizes the trade that resulted in adoption of the new design.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"2015 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":"121600997","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.2017.7943761
K. Mann, D. Holker, N. Conn
The Hosted Payload Office (HPO) in the Advanced Systems and Development Directorate (AD) at the Space and Missile Systems Center (SMC), is designing, developing and fielding the Air Force's first Low Earth Orbit (LEO) commercially hosted payload constellation. Satellite operations are potentially impacted by space weather hazards such as Single Event Effects (SEE), radiation dose effects, and deep dielectric (“internal”) charging. With its global coverage and one hertz sampling rate, REACH would provide satellite operators the ability to rapidly determine active space environments which could induce anomalies. If the environment is ruled out, then the possibility of hostile actions as a causative factor in the anomaly is more likely, a goal of the Space Enterprise Vision (SEV). This project provides an unprecedented example of how the commercially hosted payload construct enables deployment of a responsive, affordable, distributive, and proliferated space weather capability. The REACH space segment operates 32 sensors in 6 planes to reduce revisit rates to less than 20 minutes and measure radiation levels as low as 50keV. The ground segment leverages a commercially owned and fully automated network for continuous data collection and dissemination without the need for a standalone infrastructure. REACH is the embodiment of rapid acquisition principles, going from Preliminary Design Review (PDR) to on-orbit operations in less than three years. Defending on-orbit space systems from natural and hostile acts is critical to the US and its Allies to ensure persistent access to key warfighting capabilities. REACH provides a unique opportunity to demonstrate a significant improvement over existing capabilities to enable global access, persistence, and awareness.
{"title":"Responsive environmental assessment commercially hosted (REACH) payloads","authors":"K. Mann, D. Holker, N. Conn","doi":"10.1109/AERO.2017.7943761","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943761","url":null,"abstract":"The Hosted Payload Office (HPO) in the Advanced Systems and Development Directorate (AD) at the Space and Missile Systems Center (SMC), is designing, developing and fielding the Air Force's first Low Earth Orbit (LEO) commercially hosted payload constellation. Satellite operations are potentially impacted by space weather hazards such as Single Event Effects (SEE), radiation dose effects, and deep dielectric (“internal”) charging. With its global coverage and one hertz sampling rate, REACH would provide satellite operators the ability to rapidly determine active space environments which could induce anomalies. If the environment is ruled out, then the possibility of hostile actions as a causative factor in the anomaly is more likely, a goal of the Space Enterprise Vision (SEV). This project provides an unprecedented example of how the commercially hosted payload construct enables deployment of a responsive, affordable, distributive, and proliferated space weather capability. The REACH space segment operates 32 sensors in 6 planes to reduce revisit rates to less than 20 minutes and measure radiation levels as low as 50keV. The ground segment leverages a commercially owned and fully automated network for continuous data collection and dissemination without the need for a standalone infrastructure. REACH is the embodiment of rapid acquisition principles, going from Preliminary Design Review (PDR) to on-orbit operations in less than three years. Defending on-orbit space systems from natural and hostile acts is critical to the US and its Allies to ensure persistent access to key warfighting capabilities. REACH provides a unique opportunity to demonstrate a significant improvement over existing capabilities to enable global access, persistence, and awareness.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"43 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":"127833642","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.2017.7943897
A. Babuscia, J. Sauder, A. Chandra, J. Thangavelautham, L. Feruglio, N. Bienert
Interplanetary1 CubeSats and small satellites have potential to provide means to explore space and to perform science in a more affordable way. As the goals for these spacecraft become more ambitious in space exploration, the communication systems currently implemented will need to be improved to support those missions. One of the bottlenecks is the antennas' size, due to the close relation between antenna gain and dimensions. Hence, a possible solution is to develop inflatable antennas which can be packaged efficiently, occupying a small amount of space, and they can provide, once deployed, large dish dimension and correspondent gain. A prototype of a 1 m inflatable antenna for X-Band has been developed in a joint effort between JPL and ASU. After initial photogrammetry tests and radiation tests, it was discovered that the design was not able to meet the required gain. As a result, a new design, based on a spherical inflatable membrane, is proposed. This new design will allow reaching a more stable inflatable surface, hence improving the electromagnetic performance. This paper will detail the principle challenges in developing this new antenna focusing on: design, EM analysis, fabrication and tests.
{"title":"Inflatable antenna for CubeSat: A new spherical design for increased X-band gain","authors":"A. Babuscia, J. Sauder, A. Chandra, J. Thangavelautham, L. Feruglio, N. Bienert","doi":"10.1109/AERO.2017.7943897","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943897","url":null,"abstract":"Interplanetary1 CubeSats and small satellites have potential to provide means to explore space and to perform science in a more affordable way. As the goals for these spacecraft become more ambitious in space exploration, the communication systems currently implemented will need to be improved to support those missions. One of the bottlenecks is the antennas' size, due to the close relation between antenna gain and dimensions. Hence, a possible solution is to develop inflatable antennas which can be packaged efficiently, occupying a small amount of space, and they can provide, once deployed, large dish dimension and correspondent gain. A prototype of a 1 m inflatable antenna for X-Band has been developed in a joint effort between JPL and ASU. After initial photogrammetry tests and radiation tests, it was discovered that the design was not able to meet the required gain. As a result, a new design, based on a spherical inflatable membrane, is proposed. This new design will allow reaching a more stable inflatable surface, hence improving the electromagnetic performance. This paper will detail the principle challenges in developing this new antenna focusing on: design, EM analysis, fabrication and tests.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"38 2 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":"125874789","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.2017.7943866
M. Pugh, I. Kuperman, Fernando H. Aguirre, H. Mojaradi, Carl Spurgers, M. Kobayashi, E. Satorius, T. Jedrey
The Universal Space Transponder (UST) is a next generation transponder developed at the Jet Propulsion Laboratory to meet a large variety of telecom, navigation, and radio science needs for future deep-space and near-Earth missions. This paper details the UST software defined radio design and describes how the combination of a modular hardware architecture and in-flight reprogrammability enables a new level of flexibility and expandability for a space transponder. The UST uses common power and digital processing assemblies that can be integrated with a variety of RF modules and is capable of simultaneous, multiband operations with data rates up to 37.5 Mbps RX and 300 Mbps TX. This allows a single radio to support all the direct-to-Earth and relay communication requirements for even complex mission scenarios, reducing the total cost, mass, and power. The discussion includes a description of the current UST engineering models that have been built and tested, as well as details about the next generation capabilities supported by UST, including advanced link coding and modulation, radiometric techniques, and in-radio protocol handling. Details are also presented on RF modules and digital processing in development for radio science and astronomy purposes, including a bistatic radar receiver and broadband planetary emissions receiver. These will demonstrate the ability to integrate low-cost science instruments into the UST architecture, further expanding the versatility of the UST.
{"title":"The Universal Space Transponder: A next generation software defined radio","authors":"M. Pugh, I. Kuperman, Fernando H. Aguirre, H. Mojaradi, Carl Spurgers, M. Kobayashi, E. Satorius, T. Jedrey","doi":"10.1109/AERO.2017.7943866","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943866","url":null,"abstract":"The Universal Space Transponder (UST) is a next generation transponder developed at the Jet Propulsion Laboratory to meet a large variety of telecom, navigation, and radio science needs for future deep-space and near-Earth missions. This paper details the UST software defined radio design and describes how the combination of a modular hardware architecture and in-flight reprogrammability enables a new level of flexibility and expandability for a space transponder. The UST uses common power and digital processing assemblies that can be integrated with a variety of RF modules and is capable of simultaneous, multiband operations with data rates up to 37.5 Mbps RX and 300 Mbps TX. This allows a single radio to support all the direct-to-Earth and relay communication requirements for even complex mission scenarios, reducing the total cost, mass, and power. The discussion includes a description of the current UST engineering models that have been built and tested, as well as details about the next generation capabilities supported by UST, including advanced link coding and modulation, radiometric techniques, and in-radio protocol handling. Details are also presented on RF modules and digital processing in development for radio science and astronomy purposes, including a bistatic radar receiver and broadband planetary emissions receiver. These will demonstrate the ability to integrate low-cost science instruments into the UST architecture, further expanding the versatility of the UST.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"1 11 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":"129213834","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.2017.7943963
Pamela Wheeler, R. Cobb, C. Hartsfield, B. Prince
Space Situational Awareness (SSA) is of utmost importance in today's space dependent, congested and contested environment. The health of a propulsion system is vital to ensure proper function and thus proper mission placement. Electric propulsion is gaining popularity for satellite propulsion systems due to higher efficiencies, specific impulse, and the savings it offers in both spacecraft mass and launch costs. Electron temperature is a commonly used diagnostic to determine the efficiency of a Hall thruster. Recent papers have coordinated near infrared (NIR) spectral measurements of ionization lines in xenon and krypton to electron temperature measurements. This research will characterize NIR plume emissions for a 600 Watt Hall thruster using both xenon and krypton propellants for a variety of observation angles and operating power levels. By determining spectral differences when altering these variables, it would be possible to identify angle, power level, and propellant in order to provide information on electron temperature and thus efficiency. Although they have a high specific impulse, electric propulsion systems provide lower thrust than chemical alternatives. This means that the firing times needed for spacecraft maneuvers can be on the order of hours to months. This provides an opportunity for this characterization to not only be put to use in chamber experiments but on-orbit as well. Ground-based observations of these spectral lines would allow for identification of the type of thruster as well as the health of the system while the satellite is in operation on-orbit. The current SSA architecture is limited and task saturated. If smaller telescopes, like those at universities, could successfully detect these signatures they could augment data collection for the SSA network. To facilitate data collection, precise atmospheric modeling must be used to identify the signature. Within the atmosphere, the NIR has a higher transmission rate and typical HET propellants are approximately 3x the intensity in the NIR versus the visible spectrum making it ideal for ground based observations. This research will combine emission measurements with atmospheric and plume models to develop a single end-to-end model that will determine xenon and krypton signatures through the atmosphere, discernable differences in power level and viewing angle of Hall thruster systems, and estimate the efficacy through ground-based observations.
{"title":"Satellite propulsion spectral signature detection and analysis","authors":"Pamela Wheeler, R. Cobb, C. Hartsfield, B. Prince","doi":"10.1109/AERO.2017.7943963","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943963","url":null,"abstract":"Space Situational Awareness (SSA) is of utmost importance in today's space dependent, congested and contested environment. The health of a propulsion system is vital to ensure proper function and thus proper mission placement. Electric propulsion is gaining popularity for satellite propulsion systems due to higher efficiencies, specific impulse, and the savings it offers in both spacecraft mass and launch costs. Electron temperature is a commonly used diagnostic to determine the efficiency of a Hall thruster. Recent papers have coordinated near infrared (NIR) spectral measurements of ionization lines in xenon and krypton to electron temperature measurements. This research will characterize NIR plume emissions for a 600 Watt Hall thruster using both xenon and krypton propellants for a variety of observation angles and operating power levels. By determining spectral differences when altering these variables, it would be possible to identify angle, power level, and propellant in order to provide information on electron temperature and thus efficiency. Although they have a high specific impulse, electric propulsion systems provide lower thrust than chemical alternatives. This means that the firing times needed for spacecraft maneuvers can be on the order of hours to months. This provides an opportunity for this characterization to not only be put to use in chamber experiments but on-orbit as well. Ground-based observations of these spectral lines would allow for identification of the type of thruster as well as the health of the system while the satellite is in operation on-orbit. The current SSA architecture is limited and task saturated. If smaller telescopes, like those at universities, could successfully detect these signatures they could augment data collection for the SSA network. To facilitate data collection, precise atmospheric modeling must be used to identify the signature. Within the atmosphere, the NIR has a higher transmission rate and typical HET propellants are approximately 3x the intensity in the NIR versus the visible spectrum making it ideal for ground based observations. This research will combine emission measurements with atmospheric and plume models to develop a single end-to-end model that will determine xenon and krypton signatures through the atmosphere, discernable differences in power level and viewing angle of Hall thruster systems, and estimate the efficacy through ground-based observations.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"54 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":"121107681","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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