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Date 3-10 March 2012

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Displaying Results 1 - 25 of 467
  • Schedule

    Publication Year: 2012 , Page(s): 1 - 22
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  • Committe

    Publication Year: 2012 , Page(s): 1
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  • Maps

    Publication Year: 2012 , Page(s): 1 - 2
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  • Table of contents

    Publication Year: 2012 , Page(s): 1 - 35
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  • Author index

    Publication Year: 2012 , Page(s): 1 - 70
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  • [Copyright notice]

    Publication Year: 2012 , Page(s): 1
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  • Call for papers

    Publication Year: 2012 , Page(s): 1 - 28
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  • 2013 IEEE Aerospace Conference [Call for papers]

    Publication Year: 2012 , Page(s): 1
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  • System verification of MSL Skycrane using an integrated ADAMS simulation

    Publication Year: 2012 , Page(s): 1 - 11
    Cited by:  Papers (1)
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (1019 KB) |  | HTML iconHTML  

    Mars Science Laboratory (MSL) will use the Skycrane architecture to execute final descent and landing maneuvers. The Skycrane phase uses closed-loop feedback control throughout the entire phase, starting with rover separation, through mobility deploy, and through touchdown, ending only when the bridles have completely slacked. The integrated ADAMS simulation described in this paper couples complex dynamical models created by the mechanical subsystem with actual GNC flight software algorithms that have been compiled and linked into ADAMS. These integrated simulations provide the project with the best means to verify key Skycrane requirements which have a tightly coupled GNC-Mechanical aspect to them. It also provides the best opportunity to validate the design of the algorithm that determines when to cut the bridles. The results of the simulations show the excellent performance of the Skycrane system. View full abstract»

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  • Terrain safety assessment in support of the Mars Science Laboratory mission

    Publication Year: 2012 , Page(s): 1 - 8
    Cited by:  Papers (1)
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (825 KB) |  | HTML iconHTML  

    In August 2012, the Mars Science Laboratory (MSL) mission will pioneer the next generation of robotic Entry, Descent, and Landing (EDL) systems by delivering the largest and most capable rover to date to the surface of Mars. The process to select the MSL landing site took over five years and began with over 50 initial candidate sites from which four finalist sites were chosen. The four finalist sites were examined in detail to assess overall science merit, EDL safety, and rover traversability on the surface. Ultimately, the engineering assessments demonstrated a high level of safety and robustness at all four finalist sites and differences in the assessment across those sites were small enough that neither EDL safety nor rover traversability considerations could significantly discriminate among the final four sites. Thus the MSL landing site at Gale Crater was selected from among the four finalists primarily on the basis of science considerations. View full abstract»

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  • Performance of a conical ribbon drogue parachute in the wake of a subscale Orion command module

    Publication Year: 2012 , Page(s): 1 - 11
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (1577 KB) |  | HTML iconHTML  

    Ten percent of full-scale Orion conical ribbon drogue parachutes in the wake of a 10% command module were tested in a subsonic 3×2.1 m (10'×7') cross section, subsonic atmospheric wind tunnel. The motivation behind the test program is to provide a cost-efficient means to provide both parachute performance data in the wake of the Command Module (CM), and a comprehensive computational fluid dynamics validation dataset, including velocity field measurement in the wake. The subscale parachutes are constructed with the full-scale geometric porosity, number of gores and ribbons, using a novel laser cutting technique. The wind tunnel allows full exploration of the parameter space, including CM pitch plane angle, dynamic pressure (Reynolds number), and trailing distance. View full abstract»

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  • Mars Science Laboratory entry guidance improvements study for the Mars 2018 mission

    Publication Year: 2012 , Page(s): 1 - 11
    Cited by:  Papers (1)
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (3001 KB) |  | HTML iconHTML  

    In 2011, the Mars Science Laboratory (MSL) was launched in a mission to deliver the largest and most capable rover to date to the surface of Mars. A follow on MSL-derived mission, referred to as Mars 2018, is being proposed to launch in 2018. Mars 2018 is investigating performance enhancements of the Entry, Descent and Landing (EDL) system over that of its predecessor MSL mission of 2011. This paper will discuss the main elements of the proposed Mars 2018 EDL preliminary design that are being considered to increase performance on the entry phase of the mission. In particular, these elements are discussed with the goals of increasing the parachute deploy altitude to allow for more time margin during the subsequent descent and landing phases, increasing the entry mass, and reducing the delivery ellipse size at parachute deploy, through modifications in the entry reference trajectory design, vehicle's lift to drag ratio, parachute deploy trigger logic design, and the effect of additional navigation hardware. View full abstract»

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  • Entry, Descent, and Landing performance trades to increase landed mass for the Mars 2018 mission

    Publication Year: 2012 , Page(s): 1 - 13
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (1145 KB) |  | HTML iconHTML  

    The proposed Mars 2018 mission is a joint mission with the National Aeronautics and Space Administration (NASA) and the European Space Agency (ESA) that is currently in the early study phase. The objective would be to land a rover which caches a sample of Martian soil for retrieval and potential return to Earth on a subsequent mission. The Mars 2018 Entry, Descent, and Landing (EDL) system would largely leverage off of heritage from the Mars Science Laboratory (MSL) EDL system in an effort to minimize cost and schedule risk. Despite the desire to have high MSL heritage, a series of trade studies were performed to quantify the impact that incorporating EDL technologies would have on the landed mass. These new technologies include a trim tab, an increased diameter disk-gap-band (DGB) parachute over MSL, a large ringsail parachute, and a supersonic inflatable aerodynamic decelerator (SIAD). View full abstract»

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  • Drag modulation flight control for aerocapture

    Publication Year: 2012 , Page(s): 1 - 10
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (1023 KB) |  | HTML iconHTML  

    Hypersonic deployable aerodynamic devices, both rigid and inflatable, have the potential to enable a broad spectrum of next-generation aeroassist missions by mitigating shape and size constraints on aeroassist vehicles and providing an in-flight reconfiguration capability. Such a capability provides new options for flight control during atmospheric flight, such as drag modulation. Drag modulation is an attractive flight control option for future aerocapture missions because it requires only minimal additional system complexity for vehicles with deployable aerodynamic devices, in contrast to more conventional lift-modulation steering methods. This study expands upon previous aerocapture drag modulation studies by extending the analysis of single-event jettison systems to Earth and Mars. A single-event jettison guidance algorithm was developed and used to evaluate the feasibility of real-time targeting of apoapsis altitude during flight. Results indicate that sufficiently large ballistic coefficient ratios provide adequate aerodynamic and guided corridors for future aerocapture missions. While the preliminary guidance algorithm demonstrates only modest insertion accuracy, this level of accuracy may be tolerable for certain missions. View full abstract»

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  • Investigation of transonic wake dynamics for mechanically deployable entry systems

    Publication Year: 2012 , Page(s): 1 - 10
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (1029 KB) |  | HTML iconHTML  

    A numerical investigation of transonic flow around a mechanically deployable entry system being considered for a robotic mission to Venus has been performed, and preliminary results are reported. The flow around a conceptual representation of the vehicle geometry was simulated at discrete points along a ballistic trajectory using Detached Eddy Simulation (DES). The trajectory points selected span the low supersonic to transonic regimes with freestream Mach numbers from 1.5 to 0.8, and freestream Reynolds numbers (based on diameter) between 2.09 × 106 and 2.93 × 106. Additionally, the Mach 0.8 case was simulated at angles of attack between 0° and 5°. Static aerodynamic coefficients obtained from the data show qualitative agreement with data from 70° sphere-cone wind tunnel tests performed for the Viking program. Finally, the effect of choices of models and numerical algorithms is addressed by comparing the DES results to those using a Reynolds Averaged Navier-Stokes (RANS) model, as well as to results using a more dissipative numerical scheme. View full abstract»

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  • Guided entry performance of low ballistic coefficient vehicles at Mars

    Publication Year: 2012 , Page(s): 1 - 15
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (2050 KB) |  | HTML iconHTML  

    Current Mars entry, descent, and landing technology is near its performance limit and is unable to land payloads on the surface that exceed approximately 1 metric ton. One option for increasing landed payload mass capability is decreasing the entry vehicle's hypersonic ballistic coefficient. A lower ballistic coefficient vehicle decelerates higher in the atmosphere, providing additional timeline and altitude margin necessary for heavier payloads. This study analyzed the guided entry performance of concept low ballistic coefficient vehicles at Mars. A terminal point controller guidance algorithm was used to provide precision targeting capability. Accuracy at parachute deploy, peak deceleration, peak heat rate, and integrated heat load were assessed and compared to a traditional vehicle to determine the effects of lowering the vehicle ballistic coefficient on entry performance. Results from this study suggest that while accuracy at parachute deploy degrades with decreasing ballistic coefficient, accuracy and other performance metrics remain within reasonable bounds for ballistic coefficients as low as 1 kg/m2. As such, this investigation demonstrates that from a performance standpoint, guided entry vehicles with large diameters may be feasible at Mars. View full abstract»

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  • Mission sizing and trade studies for low ballistic coefficient entry systems to Venus

    Publication Year: 2012 , Page(s): 1 - 14
    Cited by:  Papers (1)
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (1560 KB) |  | HTML iconHTML  

    The U.S and the U.S.S.R. have sent seventeen successful atmospheric entry missions to Venus. Past missions to Venus have utilized rigid aeroshell systems for entry. This rigid aeroshell paradigm sets performance limitations since the size of the entry vehicle is constrained by the fairing diameter of the launch vehicle. This has limited ballistic coefficients (β) to well above 100 kg/m2 for the entry vehicles. In order to maximize the science payload and minimize the Thermal Protection System (TPS) mass, these missions have entered at very steep entry flight path angles (γ). Due to Venus' thick atmosphere and the steep-γ, high-β conditions, these entry vehicles have been exposed to very high heat flux, very high pressures and extreme decelerations (upwards of 100 g's). Deployable aeroshells avoid the launch vehicle fairing diameter constraint by expanding to a larger diameter after the launch. Due to the potentially larger wetted area, deployable aeroshells achieve lower ballistic coefficients (well below 100 kg/m2), and if they are flown at shallower flight path angles, the entry vehicle can access trajectories with far lower decelerations (~50-60 g's), peak heat fluxes (~400 W/cm2) and peak pressures. The structural and TPS mass of the shallow-γ, low-β deployables are lower than their steep-γ, high-β rigid aeroshell counterparts at larger diameters, contributing to lower areal densities and potentially higher payload mass fractions. For example, at large diameters, deployables may attain aeroshell areal densities of 10 kg/m2 as opposed to 50 kg/m2 for rigid aeroshells. However, the low-β, shallow-γ paradigm also raises issues, such as the possibility of skip-out during entry. The shallow-γ could also increase the landing footprint of the vehicle. Furthermore, the deployable entry systems may be flexible, so there could - e fluid-structure interaction, especially in the high altitude, low-density regimes. The need for precision in guidance, navigation and control during entry also has to be better understood. This paper investigates some of the challenges facing the design of a shallow-γ, low-β entry system. View full abstract»

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  • A Terrain Relative Navigation sensor enabled by multi-core processing

    Publication Year: 2012 , Page(s): 1 - 11
    Cited by:  Papers (1)  |  Patents (1)
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (1267 KB) |  | HTML iconHTML  

    Terrain Relative Navigation (TRN) provides accurate position estimates to spacecraft for precision planetary landing and autonomous primitive body exploration. A bolt-on instrument that provides the sensing and computing required for TRN will result in more accurate and robust position estimates and will simplify TRN validation. Multi-core processors provide the significant computational capability required for TRN, are straightforward to program and are being developed for space applications. We have implemented two versions of TRN on a multi-core processor and tested them in a laboratory setting. For primitive-body navigation we have demonstrated 4 second TRN updates with accuracies on order 1% of altitude. For a Mars landing application we have shown two second updates while taking out kilometer scale position uncertainties. View full abstract»

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  • Doppler lidar sensor for precision landing on the Moon and Mars

    Publication Year: 2012 , Page(s): 1 - 7
    Cited by:  Papers (1)
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (631 KB) |  | HTML iconHTML  

    Landing mission concepts that are being developed for exploration of planetary bodies are increasingly ambitious in their implementations and objectives. Most of these missions require accurate position and velocity data during their descent phase in order to ensure safe soft landing at the pre-designated sites. To address this need, a Doppler lidar is being developed by NASA under the Autonomous Landing and Hazard Avoidance Technology (ALHAT) project. This lidar sensor is a versatile instrument capable of providing precision velocity vectors, vehicle ground relative altitude, and attitude. The capabilities of this advanced technology have been demonstrated through two helicopter flight test campaigns conducted over a vegetation-free terrain in 2008 and 2010. Presently, a prototype version of this sensor is being assembled for integration into a rocket-powered terrestrial free-flyer vehicle. Operating in a closed loop with the vehicle's guidance and navigation system, the viability of this advanced sensor for future landing missions will be demonstrated through a series of flight tests in 2012. View full abstract»

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  • Improving the landing precision of an MSL-class vehicle

    Publication Year: 2012 , Page(s): 1 - 10
    Cited by:  Papers (1)
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (1284 KB) |  | HTML iconHTML  

    Prior to Mars Science Laboratory (MSL), Mars landers flew ballistic entry trajectories. Improvements in landing accuracy (from ~150 km from the target for Mars Pathfinder to ~30-40 km for Mars Exploration Rover and Phoenix) were solely due to improved approach navigation. MSL will fly the first guided-entry trajectory to Mars, further improving accuracy to ~10-12 km from the target by modifying the trajectory in the atmosphere using a lift created by slightly offsetting the vehicle center-of-mass. EDL systems of future Mars lander missions are likely to substantially resemble MSL to maximize heritage and minimize cost. Further improvements in landing accuracy are desired for these missions, motivating an investigation into improvements in landing accuracy with minimal impact to the MSL EDL system architecture. View full abstract»

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  • STP-SIV: Lessons learned through the first two Standard Interface Vehicles

    Publication Year: 2012 , Page(s): 1 - 9
    Cited by:  Papers (1)
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (637 KB) |  | HTML iconHTML  

    Spacecraft standards can enable dramatic reduction in the cost, schedule, and risk of spaceflight. The Space Development & Test Directorate (SD) of the USAF Space and Missile Systems Center (SMC) developed standards for the Space Test Program-Standard Interface Vehicle (STP-SIV) program with prime contractor Ball Aerospace & Technologies Corp. STP-SIV is designed to provide affordable, repeatable, and reliable space access to the science and technology (S&T) community. STP-SIV provides the space community well-defined standard spacecraft (SC) to-payload (PL) interface on which to base PL designs for rapid mission formation. Rather than designing unique SC for each payload; the standards provide adaptable interfaces to accommodate a range of payloads. With the first STP-SIV spacecraft, STPSat-2, operating inorbit since November 2010 and the second vehicle, STPSat-3, bus integration completed in January 2011, two data sets are available to quantitatively examine: (1) Bus integration efficiency with commercially available components (2) Payload integration and test timelines achieved with standardization (3) Improvements in efficiency from vehicle #1 to #2 (4) Relative cost savings from vehicle #1 to #2 (4) Lessons learned, successes, and drawbacks of the standardization approach These topics will be explored using examples and data from both vehicle projects, offering the reader insight into the challenges and successes surrounding the topic. The paper will also describe how the lessons learned have contributed to program efficiencies for the second vehicle, recommendations for future improvements, and how the STP-SIV approach could be further evolved to meet the aggressive demands of rapid, low-cost, responsive space. View full abstract»

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  • Integration and management challenges for multi-nanosatellite missions

    Publication Year: 2012 , Page(s): 1 - 8
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (354 KB) |  | HTML iconHTML  

    The need for launch opportunities for nanosatellite missions is ever increasing as the number of nanosatellites in development increases. Nanosatellites are typically launched as secondary, also known as auxiliary payloads, on missions to minimize launch costs. This creates managerial and technical challenges for integrating multiple nanosatellite programs as auxiliary payloads with primary missions. The priority of the integration team is to minimize impacts and risks to the primary spacecraft and the launch vehicle that are present with auxiliary satellites. Early identification of requirements for nanosatellites and their deployment devices is necessary to increase the opportunity for success of a nanosatellite rideshare mission. This paper discusses management and integration challenges for nanosatellite missions that are launched as auxiliary payloads on United States launch vehicles and techniques for addressing these challenges. View full abstract»

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  • A statistical survey of rideshares (and attack of the CubeSats, part deux)

    Publication Year: 2012 , Page(s): 1 - 7
    Cited by:  Papers (2)
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (518 KB) |  | HTML iconHTML  

    In last year's conference, we presented a statistical look at the 316 rideshare missions launched from 1990-2010, examining issues of mass, nations of origin and launch and mission type. Examinations of the data indicated that the broad range of mission types, sizes and participating nations could be classified in several useful ways. For example, we were able to forecast a bifurcation of rideshares into the CubeSat-scale and ESPA-scale categories. In this paper, we will expand on last year's results in three meaningful ways. First, we will extend the analysis back to the first rideshare in 1960 and up through 2011. In doing so, we will be able to confirm what were anecdotal conjectures from the previous paper: that the changes in the numbers and demographics of rideshares can be tied to the availability of specific launch vehicles/systems (namely the Ariane, Dnepr, Shuttle and P-POD); and that the avalanche of CubeSat flights represents a significant change in the nature of rideshares. The second extension of previous work will be the further subclassification of rideshares into military, civil, commercial and educational categories. Identifying the nature of the rideshare operator will help us better correlate the launches available to different missions. For example, we will show that the large number of U.S. rideshare missions is actually a large number of DoD rideshare missions, with a handful of U.S. civil, commercial and educational flights (most of them in the last 5 years). With this new data, we will further refine our forecasts of the launches available for various mission categories in the next few years. View full abstract»

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  • Designing affordable, flexible, and resilient small sats utilizing innovative acq strategies and standards

    Publication Year: 2012 , Page(s): 1 - 6
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (399 KB) |  | HTML iconHTML  

    The Department of Defense (DoD) Space Test Program (STP) was created in 1965 to provide access to space for DoD research and development activities. STP provides spaceflight opportunities for payloads vetted through the DoD Space Experiments Review Board (SERB). Small spacecraft acquisition is one of STP's key areas of excellence; it is a key aspect to the program that enables the program to fly multi-payload missions for low costs and provides flexibility to fly on a variety of launch vehicles. To facilitate the space flight of these experiments, the Evolved Expendable Launch Vehicle (EELV) Secondary Payload Adaptor (ESPA) was developed and flown under the direction of STP. The same payload interface requirements and environments defined for ESPA are now being applied to other multi-payload adaptors (MPAs) for other launch vehicles. To fully utilize the capabilities and opportunities presented by ESPA opportunities, STP is looking to take advantage of the ESPA form factor and success of the Standard Interface Vehicle (SIV) Program through lessons learned and innovative acquisition approaches. STP's goal is to access capable, low cost spacecraft buses through partnering and utilizing the pre-existing industry base while encouraging standardized mechanical and data interfaces where appropriate to maximize responsive spaceflight opportunities. These principles are at the foundation of STP and are influencing the future of small spacecraft acquisition and access-to-space across the space community. View full abstract»

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  • LEO protons on selected optical fibers

    Publication Year: 2012 , Page(s): 1 - 12
    Save to Project icon | Request Permissions | Click to expandQuick Abstract | PDF file iconPDF (1044 KB) |  | HTML iconHTML  

    Seven optical fiber types were measured for Low Earth Orbit (LEO) proton susceptibility. All types systematically darkened for fluence up to 1012 protons/cm2. Results demonstrate traditional 60Co γ irradiation may not represent proton darkening. Measure like you fly is important and needs representative targets and particles. View full abstract»

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