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.2017.7943786
Clara O’Farrell, Suman Muppidi, Joseph M. Brock, John W. Van Norman, I. Clark
The Advanced Supersonic Parachute Inflation Research and Experiments (ASPIRE) project will investigate the supersonic deployment, inflation, and aerodynamics of Disk-Gap-Band (DGB) parachutes in the wake of a slender body. The parachutes will be full-scale versions of the DGBs used by the Mars Science Laboratory in 2012 and planned for NASA's Mars 2020 project and will be delivered to targeted deployment conditions representative of flight at Mars by sounding rockets launched out of NASA's Wallops Flight Facility. The parachutes will be tested in the wake of a slender payload whose diameter is approximately a sixth that of entry capsules used for Mars missions. Models of the deployment, inflation, and aerodynamic performance of the parachutes are necessary to design key aspects of the experiment, including: determining the expected loads and applicable margins on the parachute and payload; guiding sensor selection and placement; evaluating the vehicle trajectory for targeting, range safety, and recovery operations. In addition, knowledge of the differences in the behavior of the parachutes in the wake of slender and blunt bodies is required in order to interpret the results of the sounding rocket experiment and determine how they relate to expected performance behind blunt bodies at Mars. However, modeling the performance of a supersonic DGB in the wake of a slender body is challenging due to the scarcity of historical test data and modeling precedents. This paper describes the models of the aerodynamic performance of DGBs in supersonic slender-body wakes being developed for the ASPIRE sounding rocket test campaign. Development of these models is based on the four available flight tests of DGBs deployed in supersonic slender-body wakes as well as on data from past flight and wind-tunnel experiments of DGBs deployed in the wake of blunt bodies, on the reconstructed at-Mars DGB performance during past missions, and on computational fluid dynamics simulations. Simulations of the wakes of blunt and slender bodies in supersonic flow have been conducted in order to investigate the differences in the flowfields encountered by parachutes deployed in both wake types. The simulations have allowed the project to investigate the fundamental differences between the sounding rocket tests and the flight of a DGB during a Mars mission and to assess the limitations of the sounding rocket test architecture for evaluating flight performance at Mars.
{"title":"Development of models for disk-gap-band parachutes deployed supersonically in the wake of a slender body","authors":"Clara O’Farrell, Suman Muppidi, Joseph M. Brock, John W. Van Norman, I. Clark","doi":"10.1109/AERO.2017.7943786","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943786","url":null,"abstract":"The Advanced Supersonic Parachute Inflation Research and Experiments (ASPIRE) project will investigate the supersonic deployment, inflation, and aerodynamics of Disk-Gap-Band (DGB) parachutes in the wake of a slender body. The parachutes will be full-scale versions of the DGBs used by the Mars Science Laboratory in 2012 and planned for NASA's Mars 2020 project and will be delivered to targeted deployment conditions representative of flight at Mars by sounding rockets launched out of NASA's Wallops Flight Facility. The parachutes will be tested in the wake of a slender payload whose diameter is approximately a sixth that of entry capsules used for Mars missions. Models of the deployment, inflation, and aerodynamic performance of the parachutes are necessary to design key aspects of the experiment, including: determining the expected loads and applicable margins on the parachute and payload; guiding sensor selection and placement; evaluating the vehicle trajectory for targeting, range safety, and recovery operations. In addition, knowledge of the differences in the behavior of the parachutes in the wake of slender and blunt bodies is required in order to interpret the results of the sounding rocket experiment and determine how they relate to expected performance behind blunt bodies at Mars. However, modeling the performance of a supersonic DGB in the wake of a slender body is challenging due to the scarcity of historical test data and modeling precedents. This paper describes the models of the aerodynamic performance of DGBs in supersonic slender-body wakes being developed for the ASPIRE sounding rocket test campaign. Development of these models is based on the four available flight tests of DGBs deployed in supersonic slender-body wakes as well as on data from past flight and wind-tunnel experiments of DGBs deployed in the wake of blunt bodies, on the reconstructed at-Mars DGB performance during past missions, and on computational fluid dynamics simulations. Simulations of the wakes of blunt and slender bodies in supersonic flow have been conducted in order to investigate the differences in the flowfields encountered by parachutes deployed in both wake types. The simulations have allowed the project to investigate the fundamental differences between the sounding rocket tests and the flight of a DGB during a Mars mission and to assess the limitations of the sounding rocket test architecture for evaluating flight performance at Mars.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"28 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":"133567069","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.2017.7943981
T. Cichan, S. Bailey, S. Norris, R. Chambers, S. Jolly, J. Ehrlich
Orion, the Multi-Purpose Crew Vehicle, is a key piece of the NASA human exploration architecture for beyond earth orbit (BEO). Lockheed Martin was awarded the contracts for the design, development, test, and production for Orion up through the Exploration Mission 2 (EM-2). Lockheed Martin is also working on defining the cis-lunar Proving Ground mission architecture, in partnership with NASA. In addition, Lockheed Martin is exploring the definition of Mars missions as the horizon goal to provide input to the plans for human exploration of the solar system. This paper describes an architecture to determine the feasibility of a Mars Base Camp architecture within about a decade. This architecture would involve human exploration of both Martian moons, and provide an opportunity for the crew to interact with pre-staged robotic assets on Mars. This study is a high-level assessment to identify architecture drivers and science opportunities. There are several key tenets for this architecture. For this first human interplanetary mission, system redundancy and a self-rescue capability is required. The number of system developments is minimized, and the use of the already developed systems like the Space Launch System and Orion is maximized. To minimize the number of events that could lead to the loss of the whole crew, the architecture does not require rendezvous and docking of pre-staged elements necessary for crew survival during the mission. This paper will describe the different enabling technologies required. The trajectory assumptions will be described, including the results of studies performed for the transit to Mars and return to Earth, in addition to mission design trades for the exploration of the Martian system. The transfer vehicle module design concept will be detailed. Possible science activities will be described. Study results for propulsion technology, assembly methods, and the mission campaign will also be addressed, as well as a discussion of planned forward work. The results of this architecture study will show that a near term Mars mission is compelling and feasible, and will highlight the required key systems.
{"title":"Mars Base Camp: An architecture for sending humans to Mars by 2028","authors":"T. Cichan, S. Bailey, S. Norris, R. Chambers, S. Jolly, J. Ehrlich","doi":"10.1109/AERO.2017.7943981","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943981","url":null,"abstract":"Orion, the Multi-Purpose Crew Vehicle, is a key piece of the NASA human exploration architecture for beyond earth orbit (BEO). Lockheed Martin was awarded the contracts for the design, development, test, and production for Orion up through the Exploration Mission 2 (EM-2). Lockheed Martin is also working on defining the cis-lunar Proving Ground mission architecture, in partnership with NASA. In addition, Lockheed Martin is exploring the definition of Mars missions as the horizon goal to provide input to the plans for human exploration of the solar system. This paper describes an architecture to determine the feasibility of a Mars Base Camp architecture within about a decade. This architecture would involve human exploration of both Martian moons, and provide an opportunity for the crew to interact with pre-staged robotic assets on Mars. This study is a high-level assessment to identify architecture drivers and science opportunities. There are several key tenets for this architecture. For this first human interplanetary mission, system redundancy and a self-rescue capability is required. The number of system developments is minimized, and the use of the already developed systems like the Space Launch System and Orion is maximized. To minimize the number of events that could lead to the loss of the whole crew, the architecture does not require rendezvous and docking of pre-staged elements necessary for crew survival during the mission. This paper will describe the different enabling technologies required. The trajectory assumptions will be described, including the results of studies performed for the transit to Mars and return to Earth, in addition to mission design trades for the exploration of the Martian system. The transfer vehicle module design concept will be detailed. Possible science activities will be described. Study results for propulsion technology, assembly methods, and the mission campaign will also be addressed, as well as a discussion of planned forward work. The results of this architecture study will show that a near term Mars mission is compelling and feasible, and will highlight the required key systems.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"15 11","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2017-03-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"133616266","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.2017.7943598
M. Langer, M. Weisgerber, J. Bouwmeester, A. Hoehn
For many years, traditional satellite design philosophy was dominated by highly reliable components, conservative designs and extensive performance testing at subsystem and integrated system levels to achieve long lifetimes in the harsh space environment. CubeSats attempted to choose a different philosophy, utilizing suitable state-of the art, commercial-off-the shelf products, yielding, if successful, an increased performance per mass figure of merit for those small vessels at potentially higher risk but lower cost. CubeSats seemed to promise universities and companies to be faster, better and cheaper — once more in history. Unfortunately, many CubeSat missions, especially university-built ones, never achieved a detectable functional state or failed shortly after the satellites were ejected from their deployer. Data based on our developed CubeSat Failure Database (CFD) and research carried out by others suggest, that a great percentage of those early failure cases could have been detected and avoided by more careful and adequate system-level functional testing on the ground. However, many university teams still fail to plan with adequate resources for system level functional testing or are confronted with hard deadlines, thus unable to complete appropriate integrated system testing on a sufficient level, and launching a satellite that never was adequately functional. Ongoing work on a novel reliability estimation tool using Bayesian methods is introduced to fill this gap and to provide meaningful data for all developers on the achievable reliability and required functional testing time of their CubeSats. Using test data and reliability goals for their actual mission, merging that data with statistical data from past missions and a database of subjective developer's beliefs, CubeSat developers should now be able to estimate their required functional testing time on subsystem and system level at an early project stage, as a function of the targeted reliability goal for their CubeSat. Alternatively, if the required resources (testing time, money, knowledge) are not available, CubeSat developers and program managers can still use the tool to now quantify a resulting realistic lower boundary for the expected system reliability of the mission, and decide, if their mission goals can be fulfilled or not with a certain probability. To evolve CubeSats into more reliable and accepted platforms for scientific payloads and commercial applications, it is utmost important to avoid or reduce the many infant mortality cases, where no or little useful data is produced by the satellite. To guide developers towards higher success rates without losing the spirit of using novel, state of the art technology in fast mission timelines, the reliability estimation tool should ensure higher reliability of CubeSat missions without drawing too much resources nor imposing too many burdens on the CubeSat teams.
{"title":"A reliability estimation tool for reducing infant mortality in Cubesat missions","authors":"M. Langer, M. Weisgerber, J. Bouwmeester, A. Hoehn","doi":"10.1109/AERO.2017.7943598","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943598","url":null,"abstract":"For many years, traditional satellite design philosophy was dominated by highly reliable components, conservative designs and extensive performance testing at subsystem and integrated system levels to achieve long lifetimes in the harsh space environment. CubeSats attempted to choose a different philosophy, utilizing suitable state-of the art, commercial-off-the shelf products, yielding, if successful, an increased performance per mass figure of merit for those small vessels at potentially higher risk but lower cost. CubeSats seemed to promise universities and companies to be faster, better and cheaper — once more in history. Unfortunately, many CubeSat missions, especially university-built ones, never achieved a detectable functional state or failed shortly after the satellites were ejected from their deployer. Data based on our developed CubeSat Failure Database (CFD) and research carried out by others suggest, that a great percentage of those early failure cases could have been detected and avoided by more careful and adequate system-level functional testing on the ground. However, many university teams still fail to plan with adequate resources for system level functional testing or are confronted with hard deadlines, thus unable to complete appropriate integrated system testing on a sufficient level, and launching a satellite that never was adequately functional. Ongoing work on a novel reliability estimation tool using Bayesian methods is introduced to fill this gap and to provide meaningful data for all developers on the achievable reliability and required functional testing time of their CubeSats. Using test data and reliability goals for their actual mission, merging that data with statistical data from past missions and a database of subjective developer's beliefs, CubeSat developers should now be able to estimate their required functional testing time on subsystem and system level at an early project stage, as a function of the targeted reliability goal for their CubeSat. Alternatively, if the required resources (testing time, money, knowledge) are not available, CubeSat developers and program managers can still use the tool to now quantify a resulting realistic lower boundary for the expected system reliability of the mission, and decide, if their mission goals can be fulfilled or not with a certain probability. To evolve CubeSats into more reliable and accepted platforms for scientific payloads and commercial applications, it is utmost important to avoid or reduce the many infant mortality cases, where no or little useful data is produced by the satellite. To guide developers towards higher success rates without losing the spirit of using novel, state of the art technology in fast mission timelines, the reliability estimation tool should ensure higher reliability of CubeSat missions without drawing too much resources nor imposing too many burdens on the CubeSat teams.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"9 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":"117351738","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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