Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943863
B. Wilcox, J. Carlton, J. Jenkins, F. A. Porter
This paper describes work on a concept for a probe that would be capable of km-scale deep subsurface ice penetration on Europa or other Ocean World. Penetrating deep into or through the ice cap over the liquid ocean is the best way to establish if life has evolved there. A central thesis of this work is that we must start by addressing the Planetary Protection constraints, and not to try to add them on at the end. Specifically, all hardware in the probe would be designed to survive heat sterilization at 500C for extended periods, as required to meet the COSPAR 1-in-10,000 probability per mission of biological contamination of the ocean. The baseline concept features a heat source containing plutonium-238 encased within a stainless Dewar so that the heat is not lost by conduction into the ice. A circular saw blade sticks out through a slot in a hemispherical turret dome at the bottom of the Dewar such that the blade cuts the ice as it spins. The turret also rotates slowly to cause the saw blade to make a hemispherical cut in the ice. The ice chips would be thrown up through the slit into the turret and would be melted by the heat source. The meltwater drains into sumps on either side of the sawblade, from which the meltwater would be pumped out to the rear of the probe. The main body of the probe contains a spool of aluminum tubing that would be dispensed from within the probe all the way back to the lander. This tubing is nominally 1–3 mm in outside dimension with integral insulated electrical wires around the center hole. This tube would pneumatically transport small (mm-scale) single-use canisters containing meltwater samples from the probe to the surface for analysis. Surface analysis allows use of instruments that can be neither miniaturized nor sterilized sufficiently to go down-hole. Dry inert gas would be used to push the canister down from the lander and back up. The dry gas would be re-compressed and re-used. During a portion of the cruise to the outer solar system, the heat source would heat the entire probe, including the coiled tubing inside as well as all the canisters and inert gas, to 500C to destroy any lingering organisms and to decompose any complex organic molecules. The paper describes analysis, design, and preliminary testing, as well as plans for building a prototype and testing it in an “ice treadmill” where a plug of ice is created inside a vertical LN2 cold jacket, pushed up by water pumped below so that the ice plug rises at the same rate that the probe penetrates the ice.
{"title":"A deep subsurface ice probe for Europa","authors":"B. Wilcox, J. Carlton, J. Jenkins, F. A. Porter","doi":"10.1109/AERO.2017.7943863","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943863","url":null,"abstract":"This paper describes work on a concept for a probe that would be capable of km-scale deep subsurface ice penetration on Europa or other Ocean World. Penetrating deep into or through the ice cap over the liquid ocean is the best way to establish if life has evolved there. A central thesis of this work is that we must start by addressing the Planetary Protection constraints, and not to try to add them on at the end. Specifically, all hardware in the probe would be designed to survive heat sterilization at 500C for extended periods, as required to meet the COSPAR 1-in-10,000 probability per mission of biological contamination of the ocean. The baseline concept features a heat source containing plutonium-238 encased within a stainless Dewar so that the heat is not lost by conduction into the ice. A circular saw blade sticks out through a slot in a hemispherical turret dome at the bottom of the Dewar such that the blade cuts the ice as it spins. The turret also rotates slowly to cause the saw blade to make a hemispherical cut in the ice. The ice chips would be thrown up through the slit into the turret and would be melted by the heat source. The meltwater drains into sumps on either side of the sawblade, from which the meltwater would be pumped out to the rear of the probe. The main body of the probe contains a spool of aluminum tubing that would be dispensed from within the probe all the way back to the lander. This tubing is nominally 1–3 mm in outside dimension with integral insulated electrical wires around the center hole. This tube would pneumatically transport small (mm-scale) single-use canisters containing meltwater samples from the probe to the surface for analysis. Surface analysis allows use of instruments that can be neither miniaturized nor sterilized sufficiently to go down-hole. Dry inert gas would be used to push the canister down from the lander and back up. The dry gas would be re-compressed and re-used. During a portion of the cruise to the outer solar system, the heat source would heat the entire probe, including the coiled tubing inside as well as all the canisters and inert gas, to 500C to destroy any lingering organisms and to decompose any complex organic molecules. The paper describes analysis, design, and preliminary testing, as well as plans for building a prototype and testing it in an “ice treadmill” where a plug of ice is created inside a vertical LN2 cold jacket, pushed up by water pumped below so that the ice plug rises at the same rate that the probe penetrates the ice.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"55 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":"114445792","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.2017.7943885
T. Imken, J. Castillo‐Rogez, Yutao He, J. Baker, A. Marinan
The Jet Propulsion Laboratory is investing in a suite of core flight system technologies to enable CubeSats to conduct missions in deep space. These will be demonstrated on currently funded missions, such as INSPIRE, MarCO, and Lunar Flashlight, which will be among the first CubeSat missions to leave Earth's orbit and explore deep space, Mars, and the Moon, respectively. Other concepts may consider using these technologies to explore Venus, asteroids, Europa, Titan, and other areas of the solar system. These missions and concepts can be enabled by the development of miniaturized yet performant command and data handling, power, software, and communications systems specifically designed for deep space applications. JPL is pushing the state of the art in small subsystems to augment NASA's history of exploration. While the CubeSat/SmallSat component market has grown significantly to benefit LEO applications, only a few vendors are actively developing avionics and instrument interface electronics capable of meeting the stringent environmental, reliability, and performance requirements of deep space missions. These electronics and systems need to be specifically designed to handle harsh radiation and thermal environments as well as extended mission durations, where a CubeSat may begin its science observations after a multi-year cruise. Deep space missions also require additional technologies, such as radio transponders for interplanetary navigation. This paper first summarizes the systems-level developments of the enabling technologies of the JPL avionics bus, looking at maturing hardware and as well as future evolutions of technologies. Second, the paper discusses potential science instruments and applications that could be accommodated by these unique flight systems, either within a CubeSat or SmallSat form factor. Finally, the paper pairs technologies and instruments and showcases potential science missions enabled by this novel capability.
{"title":"CubeSat flight system development for enabling deep space science","authors":"T. Imken, J. Castillo‐Rogez, Yutao He, J. Baker, A. Marinan","doi":"10.1109/AERO.2017.7943885","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943885","url":null,"abstract":"The Jet Propulsion Laboratory is investing in a suite of core flight system technologies to enable CubeSats to conduct missions in deep space. These will be demonstrated on currently funded missions, such as INSPIRE, MarCO, and Lunar Flashlight, which will be among the first CubeSat missions to leave Earth's orbit and explore deep space, Mars, and the Moon, respectively. Other concepts may consider using these technologies to explore Venus, asteroids, Europa, Titan, and other areas of the solar system. These missions and concepts can be enabled by the development of miniaturized yet performant command and data handling, power, software, and communications systems specifically designed for deep space applications. JPL is pushing the state of the art in small subsystems to augment NASA's history of exploration. While the CubeSat/SmallSat component market has grown significantly to benefit LEO applications, only a few vendors are actively developing avionics and instrument interface electronics capable of meeting the stringent environmental, reliability, and performance requirements of deep space missions. These electronics and systems need to be specifically designed to handle harsh radiation and thermal environments as well as extended mission durations, where a CubeSat may begin its science observations after a multi-year cruise. Deep space missions also require additional technologies, such as radio transponders for interplanetary navigation. This paper first summarizes the systems-level developments of the enabling technologies of the JPL avionics bus, looking at maturing hardware and as well as future evolutions of technologies. Second, the paper discusses potential science instruments and applications that could be accommodated by these unique flight systems, either within a CubeSat or SmallSat form factor. Finally, the paper pairs technologies and instruments and showcases potential science missions enabled by this novel capability.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"150 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":"131820570","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.2017.7943726
M. Gernhardt, O. Bekdash, H. Litaker, S. Chappell, K. Beaton, Carolyn E Newton, E. Crues, A. Abercromby
The Mars Ascent Vehicle (MAV) is the largest indivisible payload and has the largest “gear ratio” in NASA's Evolvable Mars Campaign (EMC) architecture. For this reason, the mass and related volume of the MAV cabin drives requirements for Mars Surface In-Situ Resource Utilization generators to manufacture propellants, which: drives the surface power needs, sets the Lander payload size, drives the in-space transportation architecture, and drives the number of launches and time required to land humans on Mars' surface. Some of the architectures currently under consideration in the EMC use a chemical/solar electric propulsion hybrid that inserts into a 5 sol orbit and may require that the crew spend several days in the MAV before a rendezvous with the Mars transit vehicle. There have been no human-in-the-loop (HITL) evaluations to generate the data necessary to inform decisions on the required size of the MAV. These data are critical to begin to close the various EMC architectures. The common cabin concept consists of a core cabin with ECLSS, power, thermal and GNC systems that can be coupled to various mobility systems resulting in use of the core cabin design within a lander cabin, MAV — an in-space taxi between locations in the Mars system, as a Mars moons exploration vehicle, and a Mars surface rover. The common cabin approach could be facilitated through a standard interface design that allows for attachment of the different mobility systems. The interior of the cabin would also be outfitted differently based on the mission of the particular vehicle. The focus of the project summarized here is to determine the size of the smallest viable MAV from a sizing and habitability perspective. Multiple options were considered, two of which were studied in greater depth; one that provides maximum commonality with other cabins needed across the EMC and the other being a unique minimalist cabin. To enable the sizing and habitability assessments, the project completed analysis tasks including generation of functional requirements, development of mission timelines for all phases of MAV operations (from launch to berthing with a transit habitat), definition of required subsystems, computer aided design modeling of potential cabin layouts, and development of preliminary mass equipment lists. A maximum commonality cabin was evaluated with a 4-person crew executing all phases of the mission timeline. Additionally a minimalist cabin design was subjectively evaluated using computer generated models of vehicle layout and design. The results of the test findings and analysis are described in this paper. The results suggested that the volume and mass of the common cabin and the minimalist cabins were nearly identical. The benefits of the common cabin also include the potential use as a habitable airlock/node as part of a Cislunar habitat system, buying down risk and gaining flight experience with this Mars-forward habitation element.
{"title":"Mars ascent vehicle sizing, habitability, and commonality in NASA's Evolvable Mars Campaign","authors":"M. Gernhardt, O. Bekdash, H. Litaker, S. Chappell, K. Beaton, Carolyn E Newton, E. Crues, A. Abercromby","doi":"10.1109/AERO.2017.7943726","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943726","url":null,"abstract":"The Mars Ascent Vehicle (MAV) is the largest indivisible payload and has the largest “gear ratio” in NASA's Evolvable Mars Campaign (EMC) architecture. For this reason, the mass and related volume of the MAV cabin drives requirements for Mars Surface In-Situ Resource Utilization generators to manufacture propellants, which: drives the surface power needs, sets the Lander payload size, drives the in-space transportation architecture, and drives the number of launches and time required to land humans on Mars' surface. Some of the architectures currently under consideration in the EMC use a chemical/solar electric propulsion hybrid that inserts into a 5 sol orbit and may require that the crew spend several days in the MAV before a rendezvous with the Mars transit vehicle. There have been no human-in-the-loop (HITL) evaluations to generate the data necessary to inform decisions on the required size of the MAV. These data are critical to begin to close the various EMC architectures. The common cabin concept consists of a core cabin with ECLSS, power, thermal and GNC systems that can be coupled to various mobility systems resulting in use of the core cabin design within a lander cabin, MAV — an in-space taxi between locations in the Mars system, as a Mars moons exploration vehicle, and a Mars surface rover. The common cabin approach could be facilitated through a standard interface design that allows for attachment of the different mobility systems. The interior of the cabin would also be outfitted differently based on the mission of the particular vehicle. The focus of the project summarized here is to determine the size of the smallest viable MAV from a sizing and habitability perspective. Multiple options were considered, two of which were studied in greater depth; one that provides maximum commonality with other cabins needed across the EMC and the other being a unique minimalist cabin. To enable the sizing and habitability assessments, the project completed analysis tasks including generation of functional requirements, development of mission timelines for all phases of MAV operations (from launch to berthing with a transit habitat), definition of required subsystems, computer aided design modeling of potential cabin layouts, and development of preliminary mass equipment lists. A maximum commonality cabin was evaluated with a 4-person crew executing all phases of the mission timeline. Additionally a minimalist cabin design was subjectively evaluated using computer generated models of vehicle layout and design. The results of the test findings and analysis are described in this paper. The results suggested that the volume and mass of the common cabin and the minimalist cabins were nearly identical. The benefits of the common cabin also include the potential use as a habitable airlock/node as part of a Cislunar habitat system, buying down risk and gaining flight experience with this Mars-forward habitation element.","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":"133927999","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.2017.7943728
Priyank Pradeep, Peng Wei
The current research focuses on predictability, variability and operational feasibility aspect of Continuous Descent Arrival/Approach (CDA), which is among the key concepts of the Next Generation Air Transportation System (NextGen). The idle-thrust CDA is a fuel economical, noise and emission abatement procedure, but requires increased separation to accommodate for variability and uncertainties in vertical and speed profiles of arriving aircraft. Although a considerable amount of researches have been devoted to the estimation of potential benefits of the CDA, only few have attempted to explain the predictability, variability and operational feasibility aspect of CDA. The analytical equations derived using flight dynamics and Base of Aircraft Data (BADA) Total Energy Model (TEM) in this research gives insight into dependency of vertical profile of CDA on various factors like wind speed and gradient, weight, aircraft type and configuration, thrust settings, atmospheric factors (deviation from ISA (DISA), pressure and density of air) and descent speed profile. Application of the derived equations to idle-thrust CDA gives an insight into sensitivity of its vertical profile to multiple factors. This suggests fixed geometric flight path angle (FPA) CDA has a higher degree of predictability and lesser variability at the cost of non-idle and low thrust engine settings. However, with optimized design this impact can be overall minimized. The CDA simulations were performed using Future ATM Concept Evaluation Tool (FACET) based on radar-track and aircraft type data of the real air-traffic to some of the busiest airports in the USA (ATL, SFO and New York Metroplex (JFK and EWR)). The statistical analysis of the vertical profiles of CDA shows 1) mean geometric FPAs derived from various simulated vertical profiles are consistently shallower than 3° glideslope angle and 2) high level of variability in vertical profiles of idle-thrust CDA even in absence of uncertainties in external factors. The present investigation also suggests that prediction and guidance of fixed FPA descent trajectory by the performance based Flight Management System (FMS) would help in reduction of unpredictability and variability associated with vertical profile of aircraft guided by the FMS coupled with auto-pilot (AP) and auto-throttle (AT). The statistical analysis of the vertical profiles of CDA also suggests that for procedure design; ‘AT or above’, ‘AT or below’ and ‘Window’ type altitude constraints and FPA constraints are more realistic and useful compared to obsolete ‘AT’ type altitude constraint because of variability in vertical profiles.
{"title":"Predictability, variability and operational feasibility aspect of CDA","authors":"Priyank Pradeep, Peng Wei","doi":"10.1109/AERO.2017.7943728","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943728","url":null,"abstract":"The current research focuses on predictability, variability and operational feasibility aspect of Continuous Descent Arrival/Approach (CDA), which is among the key concepts of the Next Generation Air Transportation System (NextGen). The idle-thrust CDA is a fuel economical, noise and emission abatement procedure, but requires increased separation to accommodate for variability and uncertainties in vertical and speed profiles of arriving aircraft. Although a considerable amount of researches have been devoted to the estimation of potential benefits of the CDA, only few have attempted to explain the predictability, variability and operational feasibility aspect of CDA. The analytical equations derived using flight dynamics and Base of Aircraft Data (BADA) Total Energy Model (TEM) in this research gives insight into dependency of vertical profile of CDA on various factors like wind speed and gradient, weight, aircraft type and configuration, thrust settings, atmospheric factors (deviation from ISA (DISA), pressure and density of air) and descent speed profile. Application of the derived equations to idle-thrust CDA gives an insight into sensitivity of its vertical profile to multiple factors. This suggests fixed geometric flight path angle (FPA) CDA has a higher degree of predictability and lesser variability at the cost of non-idle and low thrust engine settings. However, with optimized design this impact can be overall minimized. The CDA simulations were performed using Future ATM Concept Evaluation Tool (FACET) based on radar-track and aircraft type data of the real air-traffic to some of the busiest airports in the USA (ATL, SFO and New York Metroplex (JFK and EWR)). The statistical analysis of the vertical profiles of CDA shows 1) mean geometric FPAs derived from various simulated vertical profiles are consistently shallower than 3° glideslope angle and 2) high level of variability in vertical profiles of idle-thrust CDA even in absence of uncertainties in external factors. The present investigation also suggests that prediction and guidance of fixed FPA descent trajectory by the performance based Flight Management System (FMS) would help in reduction of unpredictability and variability associated with vertical profile of aircraft guided by the FMS coupled with auto-pilot (AP) and auto-throttle (AT). The statistical analysis of the vertical profiles of CDA also suggests that for procedure design; ‘AT or above’, ‘AT or below’ and ‘Window’ type altitude constraints and FPA constraints are more realistic and useful compared to obsolete ‘AT’ type altitude constraint because of variability in vertical profiles.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"307 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":"132932493","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.2017.7943626
J. Grundmann, Jan-Gerd Mess, J. Biele, P. Seefeldt, B. Dachwald, P. Spietz, C. Grimm, Tom Sprowitz, C. Lange, S. Ulamec
In the wake of the successful Philae landing on comet 67P/Churyumov-Gerasimenko and the launch of the first Mobile Asteroid Surface Scout, MASCOT, aboard the Hayabusa2 space probe to asteroid (162173) Ryugu, small spacecraft in applications related to small solar system bodies have become a topic of increasing interest. Their unique combination of efficient capabilities, resource-friendly design and inherent robustness makes them attractive as a mission element at the frontiers of exploration of the solar system by larger spacecraft as well as stand-alone low-cost approaches to open up the solar system for a broader range of interests. The operators' requirements for cutting-edge missions compatible with available launch capabilities impose significant constraints in resources, timelines, timeliness, mass and size. To create spacecraft feasible within these constraints, the mission design teams need to accept a broad range of equipment maturity levels from fresh concepts to off-the-shelf units. The resulting Constraints-Driven Engineering (CDE) environment has led to new methods which transcend traditional evenly-paced and sequential development. We evolved and extended Concurrent Design and Engineering (CD/CE) methods originally incepted for initial studies into Concurrent Assembly, Integration and Verification (CAIV). It is applied in all phases in most of our projects to achieve convergence of asynchronous subsystem maturity timelines and to match parallel tracks of integration and test campaigns. When facing such a challenge, Model-Based Systems Engineering (MBSE) supports design trades and constant configuration evolution due to unforeseen changes. Proactive change and schedule acceleration has resulted from system-level CD/CE optimization across interface boundaries by MBSE-aided CAIV.
{"title":"Small spacecraft in small solar system body applications","authors":"J. Grundmann, Jan-Gerd Mess, J. Biele, P. Seefeldt, B. Dachwald, P. Spietz, C. Grimm, Tom Sprowitz, C. Lange, S. Ulamec","doi":"10.1109/AERO.2017.7943626","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943626","url":null,"abstract":"In the wake of the successful Philae landing on comet 67P/Churyumov-Gerasimenko and the launch of the first Mobile Asteroid Surface Scout, MASCOT, aboard the Hayabusa2 space probe to asteroid (162173) Ryugu, small spacecraft in applications related to small solar system bodies have become a topic of increasing interest. Their unique combination of efficient capabilities, resource-friendly design and inherent robustness makes them attractive as a mission element at the frontiers of exploration of the solar system by larger spacecraft as well as stand-alone low-cost approaches to open up the solar system for a broader range of interests. The operators' requirements for cutting-edge missions compatible with available launch capabilities impose significant constraints in resources, timelines, timeliness, mass and size. To create spacecraft feasible within these constraints, the mission design teams need to accept a broad range of equipment maturity levels from fresh concepts to off-the-shelf units. The resulting Constraints-Driven Engineering (CDE) environment has led to new methods which transcend traditional evenly-paced and sequential development. We evolved and extended Concurrent Design and Engineering (CD/CE) methods originally incepted for initial studies into Concurrent Assembly, Integration and Verification (CAIV). It is applied in all phases in most of our projects to achieve convergence of asynchronous subsystem maturity timelines and to match parallel tracks of integration and test campaigns. When facing such a challenge, Model-Based Systems Engineering (MBSE) supports design trades and constant configuration evolution due to unforeseen changes. Proactive change and schedule acceleration has resulted from system-level CD/CE optimization across interface boundaries by MBSE-aided CAIV.","PeriodicalId":224475,"journal":{"name":"2017 IEEE Aerospace Conference","volume":"137 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":"123591476","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.2017.7943673
N. Maille
This paper reviews processes used by airlines to analyze inflight recorded parameters in their safety management system. Then it investigates how the expected crew activity can be modeled, based on airlines' standard operating procedures, and used to get a better understanding of operational risks. The language allowing the hierarchical description of the crew activity is described as well as the process dedicated to the identification of these elements from the recorded flight parameters. As an application case study, a go-around procedure is modeled and used to analyze a set of real aviation data. Results emphasize how this new flight data analysis process can complement existing tools.
{"title":"Modeling airline crew activity to improve flight safety analysis","authors":"N. Maille","doi":"10.1109/AERO.2017.7943673","DOIUrl":"https://doi.org/10.1109/AERO.2017.7943673","url":null,"abstract":"This paper reviews processes used by airlines to analyze inflight recorded parameters in their safety management system. Then it investigates how the expected crew activity can be modeled, based on airlines' standard operating procedures, and used to get a better understanding of operational risks. The language allowing the hierarchical description of the crew activity is described as well as the process dedicated to the identification of these elements from the recorded flight parameters. As an application case study, a go-around procedure is modeled and used to analyze a set of real aviation data. Results emphasize how this new flight data analysis process can complement existing tools.","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":"130095899","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.2017.7943834
T. Steiner, Robert Truax, Kristoffer M. Frey
Small, lightweight flight vehicles, such as consumer-grade quadrotors, are becoming increasingly common. These vehicles' on-board state estimators are typically reliant upon frequent and accurate updates from external systems such as the Global Positioning System (GPS) to provide state estimates required for stable flight. However, in many cases GPS signals may be unavailable or unreliable, and loss of GPS can cause these vehicles to go unstable or crash, potentially putting operators, bystanders, and property in danger. Thus reliance on GPS severely limits the robustness and operational capabilities of lightweight flight vehicles. This paper introduces the Smoothing And Mapping With In-ertial State Estimation (SAMWISE) navigation system. SAM-WISE is a vision-aided inertial navigation system capable of providing high-rate, low-latency state estimates to enable high-dynamic flight through obstacle-laden unmapped indoor and outdoor environments. SAMWISE offers a flexible framework for inertial navigation with nonlinear measurements, such as those produced by visual feature trackers, by utilizing an incremental smoother to efficiently optimize a set of nonlinear measurement constraints, estimating the vehicle trajectory in a sliding window in real-time with a slight processing delay. To overcome this delay and consistently produce state estimates at the high rates necessary for agile flight, we propose a novel formulation in which the smoother runs in a background thread while a low-latency inertial strapdown propagator outputs position, attitude, and velocity estimates at high-rate. We additionally propose a novel measurement buffering approach to seamlessly handle delayed measurements, measurements produced at inconsistent rates, and sensor data requiring significant processing time, such as camera imagery. We present experimental results high-speed flight with a fully autonomous quadrotor using SAMWISE for closed-loop state estimation from flight demonstrations during the DARPA Fast Lightweight Autonomy (FLA) program in April and November of 2016. SAMWISE achieved less than 1% position error and up to 5.5 m/s (12 mph) flight in a simulated indoor warehouse environment using a scanning-lidar, inertial measurement unit, and laser altimeter during the first FLA milestone event in April 2016. In November 2016, SAMWISE achieved approximately 3% error and up to 20 m/s (45 mph) flight in an open outdoor environment with large obstacles during the second FLA milestone event. The results of these flight tests demonstrate that our navigation system works robustly at high speed across multiple distinct environments.
小型、轻型飞行器，如消费级四旋翼飞行器，正变得越来越普遍。这些车辆的机载状态估计器通常依赖于来自外部系统(如全球定位系统(GPS))的频繁和准确的更新，以提供稳定飞行所需的状态估计。然而，在许多情况下，GPS信号可能不可用或不可靠，失去GPS可能会导致这些车辆不稳定或坠毁，潜在地使操作员、旁观者和财产处于危险之中。因此，对GPS的依赖严重限制了轻型飞行器的稳健性和操作能力。介绍了基于惯性状态估计的SAMWISE(平滑映射)导航系统。SAM-WISE是一种视觉辅助惯性导航系统，能够提供高速率、低延迟状态估计，使高动态飞行能够通过充满障碍物的未映射的室内和室外环境。SAMWISE为具有非线性测量的惯性导航提供了一个灵活的框架，例如那些由视觉特征跟踪器产生的测量，通过利用增量平滑器有效地优化一组非线性测量约束，实时估计滑动窗口中的车辆轨迹，并具有轻微的处理延迟。为了克服这种延迟并始终如一地以敏捷飞行所需的高速率产生状态估计，我们提出了一种新的公式，其中平滑运行在背景线程中，而低延迟惯性捷联传播器以高速率输出位置、姿态和速度估计。我们还提出了一种新的测量缓冲方法，以无缝地处理延迟的测量，以不一致的速率产生的测量，以及需要大量处理时间的传感器数据，如相机图像。我们展示了2016年4月和11月DARPA快速轻量级自治(FLA)项目飞行演示中使用SAMWISE进行闭环状态估计的全自动四旋翼高速飞行的实验结果。在2016年4月的第一次FLA里程碑事件中，SAMWISE使用扫描激光雷达、惯性测量单元和激光高度计，在模拟室内仓库环境中实现了小于1%的位置误差和高达5.5 m/s (12 mph)的飞行。2016年11月，在第二次FLA里程碑事件中，SAMWISE在开放的室外环境中实现了约3%的误差和高达20米/秒(45英里/小时)的飞行。这些飞行测试的结果表明，我们的导航系统在多种不同环境下高速运行。
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Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943984
M. Goldflam, S. Hawkins, S. Parameswaran, A. Tauke-Pedretti, L. Warne, D. Peters, S. Campione, W. Coon, G. Keeler, E. Shaner, J. Wendt, E. Kadlec, T. Fortune, J. Klem, P. Davids, J. Kim
High-quality infrared focal plane arrays (FPAs) are used in many satellite, astronomical, and terrestrial applications. These applications require highly-sensitive, low-noise FPAs, and therefore do not benefit from advances made in low-cost thermal imagers where reducing cost and enabling high-temperature operation drive device development. Infrared detectors used in FPAs have been made for decades from alloys of mercury cadmium telluride (MCT). These infrared detectors are nearing the believed limit of their performance. This limit, known in the infrared detector community as Rule 07, dictates the dark current floor for MCT detectors, in their traditional architecture, for a given temperature and cutoff wavelength. To overcome the bounds imposed by Rule 07, many groups are working on detector compounds other than MCT. We focus on detectors employing III-V-based gallium-free In As Sb superlattice active regions while also changing the basic architecture of the pixel to improve signal-to-noise. Our architecture relies on a resonant, metallic, subwavelength nanoantenna patterned on the absorber surface, in combination with a Fabry-Perot cavity, to couple the incoming radiation into tightly confined modes near the nanoantenna. This confinement of the incident energy in a thin layer allows us to greatly reduce the volume of the absorbing layer to a fraction of the free-space wavelength, yielding a corresponding reduction in dark current from spontaneously generated electron-hole pairs in the absorber material. This architecture is detector material agnostic and could be applied to MCT detector structures as well, although we focus on using superlattice antimonide-based detector materials. This detector concept has been applied to both mid-wave (3–5 μm) and longwave (8–12 μm) infrared detectors and absorbers. Here we examine long-wave devices, as these detectors currently have a larger gap between desired device performance and that of currently existing detectors. The measured structures show an external quantum efficiency exceeding 50%. We present a comparison of the modeled and measured photoresponse of these detectors and compare these detectors to currently available commercial detectors using relevant metrics such as external quantum efficiency. We also discuss modeling of crosstalk between adjacent pixels and its influence on the potential for a dual-wavelength detector. Finally, we evaluate potential advances in these detectors that may occur in the near future.
高质量的红外焦平面阵列(fpa)用于许多卫星、天文和地面应用。这些应用需要高灵敏度、低噪声的fpa，因此无法从低成本热成像仪的进步中获益，因为低成本热成像仪可以降低成本并实现高温操作，从而驱动设备的开发。几十年来，fpa中使用的红外探测器一直是由碲化汞镉(MCT)合金制成的。这些红外探测器正接近其性能的极限。这个限制，在红外探测器界被称为规则07，规定了MCT探测器的暗电流底限，在他们的传统架构中，对于给定的温度和截止波长。为了克服规则07规定的限制，许多团体正在研究除MCT以外的检测化合物。我们专注于采用iii - v基无镓In As Sb超晶格有源区域的探测器，同时也改变了像素的基本结构以改善信噪比。我们的结构依赖于在吸收器表面设计的谐振金属亚波长纳米天线，结合法布里-珀罗腔，将入射辐射耦合到纳米天线附近的紧密受限模式中。将入射能量限制在薄层中，使我们能够将吸收层的体积大大减小到自由空间波长的一小部分，从而使吸收材料中自发产生的电子-空穴对产生的暗电流相应减少。这种结构与探测器材料无关，也可以应用于MCT探测器结构，尽管我们主要使用超晶格锑基探测器材料。该探测器概念已应用于中波(3-5 μm)和长波(8-12 μm)红外探测器和吸收器。在这里，我们研究长波器件，因为这些探测器目前在期望的设备性能和目前现有的探测器之间有较大的差距。所测结构的外量子效率超过50%。我们将这些探测器的模拟光响应和测量光响应进行了比较，并将这些探测器与目前可用的商用探测器进行了比较，使用相关指标(如外部量子效率)。我们还讨论了相邻像素之间串扰的建模及其对双波长检测器电位的影响。最后，我们评估了这些探测器在不久的将来可能出现的潜在进展。
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Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943874
O. Sindiy, Brian Weatherspoon, Raffi P. Tikidjian, Tanaz Mozafari
Model-based Systems Engineering can be employed beyond management of the technical architecture development of a system to also manage the programmatics associated with Systems Engineering activities of a project. On NASA's Asteroid Redirect Robotic Mission, MBSE has been successfully employed to manage, generate, and interact with the documentation-based deliverables associated with System Engineering activities. This has been involved in defining and tracking project document, milestone, and personnel metadata via the same modeling framework used for the technical architecture management. Additionally, it has focused on improving overall user experiences through linkage of documentation to technical content in the system model, automation of manually intensive tasks, and others stakeholder-oriented features.
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Pub Date : 2017-03-04DOI: 10.1109/AERO.2017.7943651
A. Schmidt, M. French, T. Flatley
We present our work on implementing Radiation Hardening by Software (RHBSW) techniques on the Xilinx Virtex5 FPGAs PowerPC 440 processors on the SpaceCube 2.0 platform. The techniques have been matured and tested through simulation modeling, fault emulation, laser fault injection and now in a flight experiment, as part of the Space Test Program-Houston 4-ISS SpaceCube Experiment 2.0 (STP-H4-ISE 2.0). This work leverages concepts such as heartbeat monitoring, control flow assertions, and checkpointing, commonly used in the High Performance Computing industry, and adapts them for use in remote sensing embedded systems. These techniques are extremely low overhead (typically <1.3%), enabling a 3.3x gain in processing performance as compared to the equivalent traditionally radiation hardened processor. The recently concluded STP-H4 flight experiment was an opportunity to upgrade the RHBSW techniques for the Virtex5 FPGA and demonstrate them on-board the ISS to achieve TRL 7. his work details the implementation of the RHBSW techniques, that were previously developed for the Virtex4-based SpaceCube 1.0 platform, on the Virtex5-based SpaceCube 2.0 flight platform. The evaluation spans the development and integration with flight software, remotely uploading the new experiment to the ISS SpaceCube 2.0 platform, and conducting the experiment continuously for 16 days before the platform was decommissioned. The experiment was conducted on two PowerPCs embedded within the Virtex5 FPGA devices and the experiment collected 19,400 checkpoints, processed 253,482 status messages, and incurred 0 faults. These results are highly encouraging and future work is looking into longer duration testing as part of the STP-H5 flight experiment.
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