2016 IEEE Aerospace Conference最新文献

{"title":"Mission and system design for the manipulation of PHOs with space-borne lasers","authors":"Nicolas Thiry, M. Vasile, E. Monchieri","doi":"10.1109/AERO.2016.7500610","DOIUrl":"https://doi.org/10.1109/AERO.2016.7500610","url":null,"abstract":"Owing to their ability to move a target in space without requiring propellant, laser-based deflection methods have gained attention among the research community in the recent years. With laser ablation, the vaporized material is used to push the target itself allowing for a significant reduction in the mass requirement for a space mission. Specifically, this paper addresses two important issues which have remained unanswered by previous studies: the impact of the tumbling motion of the target as well as the impact of the finite thickness of the material ablated in the case of a space debris. We developed an analytical model based on energetic considerations in order to predict the efficiency range theoretically allowed by a CW laser deflection system operating under the plasma formation threshold and in absence of the two aforementioned issues. A numerical model was then developed to solve the transient heat equation in presence of vaporization and melting and assess the efficiency reduction due to the unsteadiness induced by the tumbling motion of the potentially hazardous object (PHO). The model was translated to handle the case where the target is a piece of space debris by considering specific materials such as aluminum and titanium alloys or even carbon fiber and by adapting the finite size of the computational domain along with the propagation of the ablation front. From the results of this later model, pulsed lasers appear better suited to answer the needs of a space debris de-orbiting laser system rather than CW lasers. An empirical ablation threshold is also found that establishes a direct relation between the pulse duration or the heating time (CW case), the delivered flux and the properties of the material. Derived from theoretical consideration, this threshold matches well with the predictions of our numerical model. Moreover, the numerical results are found to agree with published data of thrust coupling coefficient on targets made of aluminium and titanium alloys. In the second part of the paper, we coupled our thrust model within an orbit propagator and considered several redirect scenarios for the case of a small(56m) and a larger(100m) asteroid as well as an 8-ton defunct satellite currently orbiting in a sun-synchronous orbit at a 765km altitude. In each scenario, the laser is assumed mounted on a spacecraft that will first rendez-vous with the target and will then operate from a safe distance (500m). Based on the results, realistic mission architectures are explored. Within the last section, the paper also highlights the advantages offered in term of redundancy and scalability by techniques such as beam combining or formation flying. We show that a medium class mission carrying a CW laser system able to generate 2.4kW of output power could ensure the deflection of a 56m asteroid while a formation of such spacecraft could also achieve the deflection of a larger threat. For the debris case, our preliminary results indicate that a","PeriodicalId":150162,"journal":{"name":"2016 IEEE Aerospace Conference","volume":"31 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2016-06-27","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134445319","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}

{"title":"Optimal control of a space-borne laser system for a 100 m asteroid deflection under uncertainties","authors":"M. Vetrisano, J. Cano, Nicolas Thiry, C. Tardioli, M. Vasile","doi":"10.1109/AERO.2016.7500677","DOIUrl":"https://doi.org/10.1109/AERO.2016.7500677","url":null,"abstract":"The paper demonstrates the technical feasibility to deflect a 100 m diameter asteroid using a moderate size spacecraft carrying a 1-20 kW solar-powered class laser. To this purpose, a recent model of the laser ablation mechanism based on the characteristics of both the laser systems and the asteroid has been used to calculate the exerted thrust in terms of direction and magnitude. This paper shows a preliminary deflection uncertainty analysis for two different control logic and assuming different laser mechanism capabilities. In particular, an optimal thrust control direction and fixed laser pointing strategies were considered with two laser optics settings: the first maintaining the focus length fixed and the second able to exactly focus on the surface. Preliminary results show that in general the fixed laser pointing strategy at low power is less able to impart high deflection. Nonetheless, when the power increases, the optimal thrust method produces undesired torques, which reduces the laser momentum coupling as side effects. However, the overall efficiency is higher in the optimal thrust case. Since the collision risk between an impacting asteroid and the Earth depends on the probability distribution of the input uncertainty parameters, it is necessary to study how the overall deflection will be affected. Both aleatory and epistemic uncertainties are taken into account to evaluate the probability of success of the proposed deflection methods.","PeriodicalId":150162,"journal":{"name":"2016 IEEE Aerospace Conference","volume":"82 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2016-06-27","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"121220760","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}

{"title":"Manufacturing challenges and benefits when scaling the HIAD stacked-torus aeroshell to a 15m-class system","authors":"Greg Swanson, N. Cheatwood, Keith Johnson, A. Calomino","doi":"10.1109/AERO.2016.7500773","DOIUrl":"https://doi.org/10.1109/AERO.2016.7500773","url":null,"abstract":"Over a decade of work has been conducted in the development of NASA's Hypersonic Inflatable Aerodynamic Decelerator (HIAD) deployable aeroshell technology. This effort has included multiple ground test campaigns and flight tests culminating in the HIAD project's second generation (Gen-2) aeroshell system. The HIAD project team has developed, fabricated, and tested stacked-torus inflatable structures (IS) with flexible thermal protection systems (F-TPS) ranging in diameters from 3-6m, with cone angles of 60 and 70 deg. To meet NASA and commercial near term objectives, the HIAD team must scale the current technology up to 12-15m in diameter. Therefore, the HIAD project's experience in scaling the technology has reached a critical juncture. Growing from a 6m to a 15m-class system will introduce many new structural and logistical challenges to an already complicated manufacturing process. Although the general architecture and key aspects of the HIAD design scale well to larger vehicles, details of the technology will need to be reevaluated and possibly redesigned for use in a 15m-class HIAD system. These include: layout and size of the structural webbing that transfers load throughout the IS, inflatable gas barrier design, torus diameter and braid construction, internal pressure and inflation line routing, adhesives used for coating and bonding, and F-TPS gore design and seam fabrication. The logistics of fabricating and testing the IS and the F-TPS also become more challenging with increased scale. Compared to the 6m aeroshell (the largest HIAD built to date), a 12m aeroshell has four times the cross-sectional area, and a 15m one has over six times the area. This means that fabrication and test procedures will need to be reexamined to account for the sheer size and weight of the aeroshell components. This will affect a variety of steps in the manufacturing process, such as: stacking the tori during assembly, stitching the structural webbing, initial inflation of tori, and stitching of F-TPS gores. Additionally, new approaches and hardware will be required for handling and ground testing of both individual tori and the fully assembled HIADs. There are also noteworthy benefits of scaling up the HIAD aeroshell to a 15m-class system. Two complications in working with handmade textile structures are the non-linearity of the material components and the role of human accuracy during fabrication. Larger, more capable, HIAD structures should see much larger operational loads, potentially bringing the structural response of the material components out of the nonlinear regime and into the preferred linear response range. Also, making the reasonable assumption that the magnitude of fabrication accuracy remains constant as the structures grow, the relative effect of fabrication errors should decrease as a percentage of the textile component size. Combined, these two effects improve the predictive capability and the uniformity of the structural response for a 12-","PeriodicalId":150162,"journal":{"name":"2016 IEEE Aerospace Conference","volume":"27 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2016-06-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"129145902","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}

{"title":"A countermeasure for loss of situation awareness: Transitioning from the laboratory to the aircraft","authors":"B. Lawson, B. McGrath, A. Rupert, Linda-Brooke I. Thompson, J. Christopher Brill, Amanda M. Kelley","doi":"10.1109/AERO.2016.7500811","DOIUrl":"https://doi.org/10.1109/AERO.2016.7500811","url":null,"abstract":"Loss of situation awareness (SA) is a major contributor to aircraft mishaps. This paper describes a technological (display) countermeasure for loss of situation awareness in flight and considers its key remaining transition challenges. The display countermeasure is a tactile situation awareness system (TSAS) that provides cues concerning aircraft motion. For example, if a helicopter drifts upwards, forwards, or downwards away from its desired hover, the pilot would feel a vibrotactile pulse on top of his/her shoulders, the front of his/her torso, or beneath his/her buttocks, respectively. The key challenge remaining for the TSAS is to transition from the research laboratory science and technology (S&T) setting to routine use aboard manned aircraft, which requires extensive flight testing. We present research evidence supporting the utility of the cues provided by TSAS, the safety benefits of TSAS, and the robustness of TSAS under demanding conditions relevant to flight. However, the research setting differs greatly from the operational setting it serves. Therefore, we conclude by sharing seven practical technology transition lessons we have learned from our efforts to transition TSAS from S&T to the very different world of flight operations. We discuss how the differing procedures, standards, timelines, priorities, incentives, and expectations of scientific versus flight testing raise significant challenges to the efficient transition of new technological inventions to the aircraft. Our hope is that describing our ongoing efforts with TSAS will aid similar display technology transition efforts and provide inventors information that could foster government innovation and implementation.","PeriodicalId":150162,"journal":{"name":"2016 IEEE Aerospace Conference","volume":"25 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2016-03-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"123146112","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}

{"title":"An SET-free, fully-digital point-of-load regulator for next-generation spacecraft power systems","authors":"Nijad Anabtawi, Rabih Chamoun","doi":"10.1109/AERO.2016.7500865","DOIUrl":"https://doi.org/10.1109/AERO.2016.7500865","url":null,"abstract":"This paper presents a digitally controlled point-of-load regulator for next-generation power systems. It is intended for spacecraft with limited energy harvesting capability and on board battery storage such as micro- and nano-satellites as well as multi-tier power distribution networks of conventional satellite and spacecraft subsystems. The novel control loop was designed to minimize radiation induced single-event effects (SEE) and resulting transients. The design was implemented in 14nm bulk complimentary metal-oxide semiconductor (CMOS) process and validated with post layout simulations. It attains a peak efficiency of 95% at heavy load conditions and 79% at light loads with a maximum voltage ripple of 25mV at light loads. The most susceptible elements of the proposed regulator have a relatively high energy threshold (~ 15pJ corresponding to 40 MeV.cm2/mg) indicating a small probability of occurrence in harsh environments and no catastrophic failure.","PeriodicalId":150162,"journal":{"name":"2016 IEEE Aerospace Conference","volume":"33 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2016-03-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"127336160","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}

{"title":"Flight control system for guided rolling-airframe missile","authors":"Saeb AmirAhmadi Chomachar, A. M. Fard","doi":"10.1109/AERO.2016.7500499","DOIUrl":"https://doi.org/10.1109/AERO.2016.7500499","url":null,"abstract":"A sampled-data system associated with the guided rolling-airframe missile (RAM) is digitally controlled. The digital control system is open-loop and missile dynamics (pitch-rate-to-elevator) is assumed to be of second-order type. The square-wave input, corresponding to the elevator deflections, stabilizes system online output that is the rate of the line-of-sight (LOS) angle. The output stabilization results in two-point guidance-law to be actively realized, hence the missile approaches the target until a hit. The guidance strategy is open-loop (it doesn't require active homing), whereas the missile can hit dynamic targets moving uniformly on a linear pathway. Moreover, only the initial triggering of the missile is target-oriented and requires active target data, to be provided by a visual device at the launch stage where the missile is to be directly pointed towards the target before being fired. During the engagement, the missile is assumed to have constant roll-rate (obtained at launch stage) and also constant forward velocity. The missile is also slightly asymmetric due to actuating fins geometry. Moreover, it could be assumed almost symmetric hence the linear motion theory analysis is valid. Meanwhile, as long as the roll-rate is not close to the pitch natural frequency, the unwanted tricyclic motion is avoided and this is preferred. In the current study, and through the numerical simulations, the roll-rate is so much higher than the pitch-rate, hence normally, the Magnus-moment and its accompanied tricyclic motion are not present. Novel ideas for technology development of surface-to-air RAM (SARAM) and seaborne RAM (SEARAM) are presented. The simulations is performed in Matlab Simulink software environment with discrete-time blocks. The miss-distance is almost zero and the simulations outcome is satisfactory.","PeriodicalId":150162,"journal":{"name":"2016 IEEE Aerospace Conference","volume":"65 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2016-03-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"124980372","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}

{"title":"Analytical performance estimation and optimization of cooperative multi-channel phase locked loops","authors":"G. Poberezhskiy","doi":"10.1109/AERO.2016.7500942","DOIUrl":"https://doi.org/10.1109/AERO.2016.7500942","url":null,"abstract":"Cooperative phase tracking enhances the mitigation of phase errors common for all channels, such as errors caused by phase noise of a local oscillator (LO) in multichannel receivers. Analysis and optimization of phase-locked loops (PLLs) in cooperative mode is significantly more complex than that of standalone PLLs, and to date their performance characteristics and optimal parameters are obtained numerically or by simulations. This complicates trade studies and implementation of cooperative phase tracking. In this paper, simple closed-form estimates for tracking errors and thresholds, as well as optimal noise bandwidths and filter coefficients of PLLs in cooperative mode have been derived for the case of mitigation of 1/f3 LO phase noise. The obtained equations can also be used to evaluate the applicability of cooperative phase tracking concept to a given scenario. An example of signal tracking with strong LO phase noise is analyzed, and good agreement between analytical results and simulations is demonstrated.","PeriodicalId":150162,"journal":{"name":"2016 IEEE Aerospace Conference","volume":"2673 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2016-03-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"124996638","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}

{"title":"FiSI: Fiberscope sample imaging system for robotic comet surface sample return missions","authors":"R. Toda, Y. Bae, Jesse Grimes-York, M. Badescu, P. Vieira, S. Moreland, P. Backes, H. Manohara","doi":"10.1109/AERO.2016.7500715","DOIUrl":"https://doi.org/10.1109/AERO.2016.7500715","url":null,"abstract":"This paper discusses the Fiberscope Sample Imaging (FiSI) system currently being developed for a potential robotic comet surface sample return mission. In this mission concept, the spacecraft would perform touch-and-go maneuver at a small body to collect a comet surface sample. Immediately after the sample is captured the FiSI would perform in situ verification of the comet sample. Sample volume would be estimated and images of the collected sample acquired and evaluated. If the captured sample volume were deemed insufficient, the sample collection maneuver would be re-attempted, multiple times if necessary, until a baseline sample volume was positively confirmed. This repeatability would improve the potential science outcome of the sample return mission. Our proof-of-concept FiSI hardware consists of nine imaging fiberscopes integrated into a single bundle. The nine fiberscopes are designed to provide wide swath coverage of overlapping fields of view within a sample measurement station. The achieved image resolution is in excess of 4 linepair/mm at 20 mm working distance. Surface color and texture of a comet sample simulant would clearly be discernible at this fidelity. The distal end of these fiberscopes are designed to tolerate harsh temperature and radiation environments near a comet while sensitive electronics and optical components at the proximal end can be placed in a more benign electronics bay of the notional spacecraft. An early FiSI prototype was tested in a -50°C chamber and showed no image degradation. To study the FiSI proof-of-concept system response in a microgravity-like environment, a preliminary experiment was attempted using a neutral buoyancy sample. The test result was consistent with Monte Carlo simulation.","PeriodicalId":150162,"journal":{"name":"2016 IEEE Aerospace Conference","volume":"78 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2016-03-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"125021250","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}

{"title":"An affordable and flexible architecture for deep space exploration","authors":"J. Engle, Travis Moseman","doi":"10.1109/AERO.2016.7500590","DOIUrl":"https://doi.org/10.1109/AERO.2016.7500590","url":null,"abstract":"This paper describes a cislunar proving ground architecture as part of an affordable, evolvable program that maximizes return on investment in the International Space Station (ISS), the Space Launch System (SLS), and Orion spacecraft. The cislunar proving ground architecture is capable of performing/supporting key missions and objectives, including deep space habitation stays for increasing durations, an asteroid redirect mission, lunar excursions, and other Mars mission enabling demonstrations. Key approaches to affordability include maximizing the use of SLS/Orion co-manifested payload launch mass and assuming incorporation of International Partner (IP) contributions and commercial logistics support. Additional focus on affordability was placed in the areas of standardizing key interfaces, incorporating commonality where feasible, and making use of ISS as a technology demonstration facility. The primary conclusion from the study is that an incremental architecture can support meaningful cislunar proving ground missions/objectives as well as provide a framework for international and commercial cooperation. The underlying long duration mission capabilities and partnerships developed with the ISS in concert with the transportation system enabled by the SLS and Orion make this architecture feasible.","PeriodicalId":150162,"journal":{"name":"2016 IEEE Aerospace Conference","volume":"125 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2016-03-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"126093308","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}

{"title":"TDRSS narrow-band simulator and test system","authors":"A. Dissanayake, A. Zarembowitch, K. Hogie, Xun Yang, Jeffry Lubelczyk, H. Safavi","doi":"10.1109/AERO.2016.7500752","DOIUrl":"https://doi.org/10.1109/AERO.2016.7500752","url":null,"abstract":"Introducing new communication equipment or system to the Tracking and Data Relay Satellite System (TDRSS) ground segment is a complicated process that involves a lengthy development and test cycle. Complexity arises due to the presence of the variety of signaling formats, operation and user constraints, monitor and control parameters, and the requirement to ensure the new piece of ground equipment or system meets the functional and performance requirements for all different combinations of system configuration parameters and operational scenarios. The development cycle starts with extensive software simulations to establish the performance parameters of the new equipment or system followed by a prototyping stage. The prototype system undergoes a battery of tests in a laboratory environment to verify that it is capable of meeting the performance levels derived from the software simulations. Testing must be performed using simulated signals generated either using software or hardware. The current practice is to use a chain of test equipment, each piece in the chain dedicated to a single function. This requires manual setup between different test configurations and test reporting is also handled manually. This is costly both in terms of test hardware investment and the manpower involved. The options available to reduce the cost are to use either a software or hardware simulator that can combine multiple functions in a single unit. Software simulation involves generating signal files that can be played back repeatedly. Depending on the data rate being simulated, each test file can occupy several Terabytes (TBs); in addition, it requires a playback system which tends to be relatively expensive to acquire and maintain. On the other hand, with more advanced signal processing technology the hardware simulation can be achieved with low cost Field Programmable Gate Array (FPGA). A low cost hardware simulator that can support data rates up to 25 Mbps is selected for the simulation system. The simulator supports baseband digital data formats and different modulation and coding schemes used in the TDRSS. Channel impairments applicable for communicating with orbiting platforms and signal impairments generated by the TDRSS itself are also simulated by the system; these include stressing profiles of delay, Doppler, and multipath. Hardware distortions of customer platforms are simulated by a set of software defined distortion filters designed according to the user constraints specified in the Space Network user's guide. In order to reduce the time spent during the prototype testing phase, an automated test system is being developed. Simulator architecture and the test automation approach will be presented.","PeriodicalId":150162,"journal":{"name":"2016 IEEE Aerospace Conference","volume":"24 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2016-03-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"126619888","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}