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  • Biochemistry and Biotechnology
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  • LUNAR AND PLANETARY EXPLORATION
  • Spacecraft Design, Testing and Performance
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  • Articles  (5)
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  • 101
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    In:  CASI
    Publication Date: 2019-12-07
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN71869 , WFIRST Formulation Science Working Group; Jul 31, 2019; Greenbelt, MD; United States
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  • 102
    Publication Date: 2020-01-24
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: SSTI-2200-0178
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  • 103
    Publication Date: 2019-12-17
    Description: NASAs Entry, Descent and Landing Architecture Study uses a trajectory simulation framework to evaluate various technologies and concepts of operations for human scale EDL at Mars. The study results inform agency technology investments. This paper summarizes the design assumptions and analysis of two deployable entry concepts performed in Phase 2 of the study. The entry concepts include a rigid deployable called the Adaptable Deployable Entry Placement Technology and an inflatable concept called the Hypersonic Inflatable Aerodynamic Decelerator. This paper describes the concept operations of these vehicles to deliver a 20-metric ton payload to the surface of Mars. Details of vehicle design and flight performance are summarized along with results of analysis on the aft body heating and its effect on the payload. Finally, recommended technology investments based on the results are presented.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NF1676L-31056 , 2018 AIAA SPACE and Astronautics Forum and Exposition; Sep 17, 2018 - Sep 19, 2018; Orlando, FL; United States
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  • 104
    Publication Date: 2019-12-17
    Description: An examination of the Hubble Space Telescope Wide Field Planetary Camera 2 (WFPC-2) radiator assembly was conducted at NASA Goddard Space Flight Center during the summer of 2009. Immediately apparent was the predominance of impact features, identified as simple or complex craters, resident only in the thermal paint layer; similar features were observed during a prior survey of the WFPC-1 radiator. Larger impact features displayed spallation zones, darkened areas, and other features not observed in impacts onto bare surfaces. Craters were extracted by coring the radiator in the NASA Johnson Space Centers Space Exposed Hardware cleanroom and were subsequently examined using scanning electron microscopy/energy dispersive X-ray spectroscopy to determine the likely origin, e.g., micrometeoritic or orbital debris, of the impacting projectile. Recently, a selection of large cores was re-examined using a new technique developed to overcome some limitations of traditional crater imaging and analysis. This technique, motivated by thin section analysis, examines a polished, lateral surface area revealed by cross-sectioning the core sample. This paper reviews the technique, the classification rubric as extended by this technique, and results to date.
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN73928 , International Orbital Debris Conference (IOC); Dec 09, 2019 - Dec 12, 2019; Sugar Land, TX; United States
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  • 105
    Publication Date: 2019-12-05
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: M19-7680 , International Astronautical Congress; Oct 21, 2019 - Oct 25, 2019; Washington, D. C.; United States
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  • 106
    Publication Date: 2019-12-05
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN70954 , EAA AirVenture Oshkosh 2019; Jul 19, 2019 - Jul 26, 2019; Oshkosh, WI; United States
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  • 107
    Publication Date: 2019-08-14
    Description: Bigelow Expandable Activity Module's (BEAM) life has been extended and in addition to being a test bed of the first human-rated expandable space module, BEAM will be utilized as an ISS stowage module. Hear about BEAM's on-orbit performance as well as highlights of BEAM's use on ISS as a technology demonstration.
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN71575 , ISS Research & Development 2019; Jul 29, 2019 - Aug 01, 2019; Atlanta, GA; United States
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  • 108
    Publication Date: 2019-08-23
    Description: At NASA's Langley Research Center in Hampton, Virginia, and Ames Research Center in California's Silicon Valley, researchers and engineers are planning a mission to demonstrate the next generation of solar sail technology for small interplanetary spacecraft. As part of this development effort, the Advanced Composite Solar Sail System (ACS3) will demonstrate deployment of an approximately 800 square foot (74 square meter) composite boom solar sail system in low-Earth orbit. This will be the first use of composite booms as well as sail packing and deployment systems for a solar sail in orbit. Also developed for ACS3 is an innovative tape-spool boom extraction system to minimize blossoming, or jamming, of the coiled booms during deployment.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN71764 , International Symposium on Solar Sailing; Jul 30, 2019 - Aug 02, 2019; Aachen; Germany
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  • 109
    Publication Date: 2019-08-15
    Description: The capability of future X-ray telescopes depends on the quality of their Point Spread Function (PSF) and the size of their field of view. Traditional designs, such as Wolter, and Wolter-Schwarzschild telescopes are stigmatic on the optical axis but their PSF degrades rapidly off-axis. At the optimal focal surface, their PSFs can be significantly improved. We present a simple optimization process for Wolter (W), Wolter-Schwarzschild (WS) and Hyperboloid-Hyperboloid (HH) telescopes that substantially improves the off-axis PSF for either narrow or wide field of view applications. In this paper, we will compare the optical performance of conventional and optimized W-, WS-, and HH-telescopes for a wide range of telescope diameters that can be used to build up future x-ray telescopes.
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN70843-1 , SPIE Optics + Photonics; Aug 11, 2019 - Aug 15, 2019; San Diego, CA; United States
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  • 110
    Publication Date: 2019-07-13
    Description: A series of wind tunnel tests were conducted to characterize the force-and-moment, and aeroacoustic environment of several configurations of the Space Launch System during ascent. The tests were conducted in the 11-by-11 foot transonic and 9-by-7 foot supersonic test sections at NASA Ames research center. Throughout these experiments data was collected from several types of instrumentation including: multicomponent force-and-moment strain gage balances, dynamic and steady-state pressure sensors, unsteady and steady pressure-sensitive paint, time-resolved shadowgraph and infrared imaging. The following details results and analysis from the time-resolved shadowgraph and infrared imaging data systems. The time-resolved shadowgraph and infrared imaging provided a qualitative measurement of the near-field turbulent fluctuations. These results helped provide context to the relative magnitude and frequency content of the fluid-structure-interaction driving the surface pressure phenomena characterized by the discrete pressure transducers and unsteady pressure sensitive paint.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN63676 , American Institute of Aeronautics and Astronautics (AIAA) SciTech Forum; Jan 07, 2019 - Jan 11, 2019; San Diego, CA; United States
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  • 111
    Publication Date: 2019-07-13
    Description: Numerical evaluation of a Whipple shield has yielded a ballistic model for cylindrical projectiles. A numerical model of a Whipple shield covered with a thermal-blanket that is representative of operational shields has been developed using CTH: The shield model has been evaluated against existing test data; The projectile model has been developed from preexisting models for graphite-epoxy materials. The model has been used to identify projectile characteristics at the ballistic limit for spherical and cylindrical projectiles; Considered normal impacts of the shield at 7 kilometers per second; Varied the angle between the axis-of-symmetry and velocity vector; Varied the length to the diameter ratio over a broad range. The results have been consolidated into a generalized model that can be adapted to existing spherical models
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN68146 , Interagency Debris Cordination (IADC) Meeting; May 07, 2019 - May 10, 2019; Rome; Italy
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  • 112
    Publication Date: 2019-07-13
    Description: There is growing interest for utilizing Small Satellites beyond low Earth orbit. A number of secondary CubeSat payload missions are planned at Mars, cis-Lunar Space, near Earth objects, and moons of the Gas Giants. Use of smaller systems may enable utilization of otherwise unused capacity of larger "host" missions. Development of re-entry systems that leverage and accommodate Small Satellite technology will substantially expand the range of mission applications by offering the capability for high speed entry or aerocapture at destinations with atmospheres. Deployable entry vehicles (DEVs) offer benefits over traditional rigid aeroshells including volume, mass and payload form factor. The Adaptive Deployable Entry and Placement Technology (ADEPT) offers such a delivery capability for Small Sat or CubeSat orbiter(s), in-situ elements, or landers. The ADEPT system can package with off the shelf CubeSat deployment systems (1U-16U) to offer a delivery capability for a single CubeSat or constellations. Furthermore, ADEPT can deliver the same science payload to a destination with a stowed diameter a factor of 3-4 times smaller than an equivalent rigid aeroshell, alleviating volumetric constraints on the secondary payload accommodation or primary carrier spacecraft bus. This paper will describe ADEPT's current development status and define various interplanetary mission concepts in order to provide guidelines for potential Small Satellite payload developers and mission implementers.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN68201 , Interplanetary Small Satellite Conference; Apr 29, 2019 - Apr 30, 2019; San Luis Obispo, CA; United States
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  • 113
    Publication Date: 2019-09-07
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: M19-7604 , AIAA Propulsion and Energy Forum; Aug 19, 2019 - Aug 22, 2019; Indianapolis, IN; United States
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  • 114
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    In:  CASI
    Publication Date: 2019-09-06
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: M19-7510 , International Conference on Environmental Systems (ICES 2019); Jul 07, 2019 - Jul 11, 2019; Boston, MA; United States
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  • 115
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    In:  CASI
    Publication Date: 2019-09-06
    Description: The Near-Earth Asteroid Scout (NEA Scout) is a 6U CubeSat that will fly to a near earth asteroid using a solar sail. The mission is a joint project between NASAs Marshall Space Flight Center and the Jet Propulsion Laboratory. The CubeSat will be deployed as a secondary payload during the Space Launch System (SLS) Exploration Mission 1 (EM-1). The CubeSat will use an 85 sq m (915 sq ft) aluminized polyimide solar sail for deep space propulsion. A multispectral camera will be used to characterize a small asteroid (〈300 feet in diameter). The primary thermal architecture is a passive design with heaters to keep temperatures above the minimum allowable. Thermal vacuum testing was done on subsystems where possible. However for some long lead subsystems thermal vacuum testing will not be done until the final assembly.
    Keywords: Spacecraft Design, Testing and Performance
    Type: M19-7413 , International Conference on Environmental Systems (ICES 2019); Jul 07, 2019 - Jul 11, 2019; Boston, MA; United States
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  • 116
    Publication Date: 2019-09-07
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: M19-7606 , AIAA Propulsion and Energy Forum; Aug 19, 2019 - Aug 22, 2019; Indianapolis, IN; United States
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  • 117
    Publication Date: 2019-11-26
    Description: In the CLASP / CLASP2 experiments, the contamination control was conducted mainly for the purpose of preventing molecular contamination. In particular, the organic contamination was successfully prevented. However, in the CLASP observation, the Ly intensity decreased due to water molecules. Thus, we obtained the importance to evaluate the influence of molecules that can be easily removed and to quantitatively evaluate the vent path during the flight.
    Keywords: Spacecraft Design, Testing and Performance
    Type: MSFC-E-DAA-TN74867 , Space Sciences and Technology Conference; Nov 06, 2019 - Nov 08, 2019; Tokushima; Japan
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  • 118
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    In:  CASI
    Publication Date: 2019-10-08
    Description: Space science missions based on a swarm implementation strategy offers benefits for future science missions. Achieving viable swarm missions depends on key enabling technologies. This presentation discusses the enabling technologies and how they can be developed and tested.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN71009 , Space Studies Program 2019; Jun 24, 2019 - Aug 23, 2019; Strasbourg; France
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  • 119
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    In:  CASI
    Publication Date: 2019-11-22
    Description: With increased emphasis on lunar exploration and scientific investigation, there is a desire to deliver a wide variety of payloads to the lunar surface. Many of these payloads will require the use of surface mobility capability such as a rover. NASA has combined spacecraft and subsystem engineers from across the Agency to develop a pallet lander design intended to deliver and easily deploy a medium-sized payload (~300 kg) to the polar regions of the Moon. The lander provides power to the payload from transit soon after lunar landing. The lander is not intended to survive the lunar night. The design of the lander was based on a minimum set of level 1 requirements where traditional risk, mass, and performance trade parameters were weighed lower than cost. In other words, the team did not sacrifice good enough for better or best. As a NASA class D spacecraft (as defined in NPR 8705.4, Risk Classification for NASA Payloads, the lander employs single-string (i.e., zero-fault-tolerant) systems as a baseline. The design utilizes existing technologies and components where possible, though some enhancements have been targeted in areas such as precision autonomous landing and low-cost structural design/fabrication. It is important to note that these and other derived technologies are extensible to other lander designs and missions. This TP describes the requirements and approaches upon which the lander design is based; discusses key design decisions, analyses, and trades used to derive the design; provides a snapshot of each major subsystem; and identifies open items, issues, and challenges for which work is continuing.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NASA/TP-2019–220391 , M-1492 , MSFC-E-DAA-TN73753
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  • 120
    Publication Date: 2019-11-15
    Description: Severe entry environments at Venus are a key challenge for all missions employing probes, landers, areal platforms, aerocapture and atmospheric skimming. Three specific mature technologies, PICA, HEEET, and ADEPT, are enablers for Venus in-situ missions but are at risk of atrophy or loss if not maintained. All three technologies were NASA-developed in partnership with US industry and rely on both organizations for intellectual property. These technologies are needed only for NASA missions and lack applicability elsewhere. NASA has experienced the loss of prior TPS technologies due to lack of use, including Apollos Avcoat (re-created at enormous expense for Orion) and Pioneer-Venus heritage carbon phenolic. Given the low flight cadence for planetary entry missions overall and the lack of non-NASA uses for these technologies, there is a real concern for the sustainment of key entry technologies.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN74856 , Meeting of the Venus Exploration and Analysis Group (VEXAG); Nov 06, 2019 - Nov 08, 2019; Boulder, CO; United States
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  • 121
    Publication Date: 2019-12-28
    Description: This report documents findings from a Small Satellite (SmallSat) Industrial Base Study conducted by The Aerospace Corporation between November 2018 and September 2019. The primary objectives of this study were a) to gain a better understanding of the SmallSat communitys technical practices, engineering approaches, requirements flow-downs, and common processes and b) identify insights and recommendations for how the government can further capitalize on the strengths and capabilities of SmallSat offerings. In the context of this study, SmallSats are understood to weigh no more than 500 kg, as described in State of the Art Small Spacecraft Technology, NASA/TP-2018- 220027, December 2018. CubeSats were excluded from this study to avoid overlap and duplication of recently completed work or other studies already under way. The team also touched on differences between traditional space-grade and the emerging mid-grade and other non-space, alternate-grade EEEE (electrical, electronic, electromechanical, electro-optical) piece part categories. Finally, the participants sought to understand the potential effects of increased use of alternate-grade parts on the traditional space-grade industrial base. The study team was keenly aware that there are missions for which non-space grade parts currently are infeasible for the foreseeable future. National security, long-duration and high-reliability missions intolerant of risk are a few examples. The team sought to identify benefits of alternative parts and approaches that can be harnessed by the government to achieve greater efficiencies and capabilities without impacting mission success.
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN75866 , OTR-2019-01165
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  • 122
    Publication Date: 2020-01-16
    Description: This presentation provides mission operations status for the Earth Observing System (EOS) Aqua satellite for the past six-months (June 2019 through November 2019). It only contains information that is of interest to the International Earth Science Constellation (ESC) Mission Operations Working Group (MOWG) member missions. It will be presented at the bi-annual MOWG Meeting in Gilbert, Arizona on Tuesday, December 3, 2019. These meetings have been occurring twice a year since the MOWG was formed in 2003.
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN76042 , Constellation Mission Operations Working Group; Dec 03, 2019; Gilbert, AZ; United States
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  • 123
    Publication Date: 2019-08-26
    Description: Exemplary deployable sheet material systems may be configured to stow and deploy sheet material. The systems may include one or more masts, one or more extendable booms, and one or more guys wires configured to function in conjunction with each other to deploy the sheet material and then to maintain the sheet material in the deployed configuration.
    Keywords: Spacecraft Design, Testing and Performance
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  • 124
    Publication Date: 2019-09-21
    Description: One of the unique features of spacecraft entering the atmosphere of Mars is the possibility of a major dust storm occurring during the entry. Design of the thermal protection system (TPS) for Mars missions have to account for the possibility of dust erosion when estimating the thickness of the TPS. Because weight is always a critical factor in designing entry vehicles, accurate assessment of dust erosion is necessary to avoid over-design of the TPS.This study will present computational results of heatshield erosion due to dust particle impacts on the ExoMars Schiaparelli capsule if it had encountered a dust storm during its October 2016 entry. A one-way coupling approach is used where particle trajectories are computed based on an underlying CFD flow solution. Based on a distribution of particle sizes ranging from 1 to 18 microns in diameter, the Icarus material response solver predicted approximately 1 mm of TPS heatshield erosion due to particle impacts, which was 40% of the value due to material charring.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN72901 , Ablation Workshop; Sep 16, 2019 - Sep 17, 2019; Minneapolis, MN; United States
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  • 125
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    In:  CASI
    Publication Date: 2019-10-23
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN74083 , Association of Space Explorers (ASE) Planetary Congress 2019; Oct 14, 2019 - Oct 18, 2019; Houston, TX; United States
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  • 126
    Publication Date: 2019-09-17
    Description: The Imaging X-ray Polarimetry Explorer (IXPE) will add polarization to the properties (time, energy, and position) observed in x-ray astronomy. A NASA Astrophysics Small Explorer (SMEX) in partnership with the Italian Space Agency (ASI), IXPE will measure the 28-keV polarization of a few dozen sources during the first 2 years following its 2021 launch. The IXPE Observatory includes three identical x-ray telescopes, each comprising a 4-m-focal-length (grazingincidence) mirror module assembly (MMA) and a polarization-sensitive (imaging) detector unit (DU), separated by a deployable optical bench. The Observatorys Spacecraft provides typical subsystems (mechanical, structural, thermal, power, electrical, telecommunications, etc.), an attitude determination and control subsystem for 3-axis stabilized pointing, and a command and data handling subsystem communicating with the science instrument and the Spacecraft subsystems.
    Keywords: Spacecraft Design, Testing and Performance
    Type: MSFC-E-DAA-TN72945 , SPIE Optical Engineering + Applications; Aug 11, 2019 - Aug 15, 2019; San Diego, CA; United States|Proceedings of SPIE (ISSN 0277-786X) (e-ISSN 1996-756X); 11118; 111180V
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  • 127
    Publication Date: 2019-12-10
    Description: This paper offers a NASA perspective on guidance for fracture control and structural certification of "fracture critical" additive manufactured spacecraft structures. Results of a recent industry survey on the need for maturing NASA guidance for AM certification are presented for the first time. The survey results are placed in the context of a description of what NASA standards in this area already exist, where the gaps are, and what work to address the gaps is underway.
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN73448 , ASTM Symposium on Structural Integrity of Additive Manufactured Materials and Parts; Oct 07, 2019 - Oct 10, 2019; Oxon Hill, MD; United States
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  • 128
    Publication Date: 2019-11-06
    Description: Dust is a major concern for lunar exploration. To combat the effects of dust, NASA, academia, and industry are developingsolutions to the dust problem. One potential technology solution for this problem is the Electrodynamic DustShield (EDS). Many years of research and development have gone into this technology. The Materials on InternationalSpace Station Experiment - 11 (MISSE-11) provides a long term space exposure platform for this technology to verifycompatibility of materials and manufacturing processes to the space environment. The MISSE-11 EDS experimentconsists of 12 EDS panels. These panels are made of glass, polyimide, or prototype spacesuit fabric. Some panels arecovered with a lotus leaf coating while others are covered with thermal paint. They are flown in the wake position ofthe ISS to simulate the lunar environment. Two panels are in an active configuration and are energized with a highvoltage power supply, which generates high-voltage pulses to activate the dust shields. Current and voltage data arerecovered from each of these trials to compare to baseline data. Also, each of the EDS panels are imaged on a monthlybasis to track any changes with time that may occur with the EDS variants. In this paper, we report preliminary dataand analysis from this spaceflight experiment.
    Keywords: Spacecraft Design, Testing and Performance
    Type: KSC-E-DAA-TN74429 , International Astronautical Congress (IAC) 2019; Oct 21, 2019 - Oct 25, 2019; Washington, D.C.; United States
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  • 129
    Publication Date: 2019-10-23
    Description: Analytic expressions for spacecraft attitude and rate estimation performance of an attitude estimation filter in terms of sensor specifications are useful tools for spacecraft design. Farrenkopf (1978) famously found analytic expressions for steady-state pre-update and post-update attitude and gyro bias estimate error variances for an attitude estimation filter for a single-axis spacecraft with a Rate Output Gyro (ROG). Markley and Reynolds (2000) extended the analysis for a Rate-Integrating Gyro (RIG) with angle white noise. These expressions allow for the rapid evaluation of system performance during preliminary mission design phases. One contribution of this paper is the analytic calculation of the steady-state pre-update and post-update angular rate estimate uncertainty for both the ROG and RIG cases. The primary contribution of this paper is the extension of the results for both the ROG and the RIG cases to the situation of an attitude sensor outage. This situation arises frequently in practice; for example when a star sensors field of view is occluded, when a star sensors readings are unreliable during a thruster burn that vibrates the spacecraft, or during star sensor outages due to radiation upsets. Analytic expressions for the attitude estimate uncertainty, gyro bias estimate uncertainty, and angular rate estimate uncertainty are given in terms of the attitude sensor outage interval, the star tracker measurement noise, and gyro noise parameters. Validity of the analytic results is demonstrated via Monte Carlo simulation.
    Keywords: Spacecraft Design, Testing and Performance
    Type: IAC-19-C1.6.2 , GSFC-E-DAA-TN74144-1 , International Astronautical Congress; Oct 21, 2019 - Oct 25, 2019; Washington, DC; United States
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  • 130
    Publication Date: 2019-09-14
    Description: Presentation to support a Summit/TIM with various Stakeholders (Ground and Flight) to perform a GAP analysis associated with ground and spacecraft maintenance, maintainability, and availability activities in order to identify missing Agency planning and guidance materials.
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN73163
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  • 131
    Publication Date: 2021-03-29
    Description: Abstract: Current information on the potential distribution and condition of peatland soils are of great importance. This applies notably for actions with respect to climate and environmental protection. The Ministry of infrastructure and agriculture in the federal state of Brandenburg (MIL) therefore initiated a project to provide a complete and updated map of peatland soils for the federal state of Brandenburg, Germany, for the year 2013. Extensive legacy data on both, areal extent of peatland soils as well as soil profile information where made digitally available in a homogeneous map and a consistent database. Approximately 7.725 sites were randomly selected from legacy soil data and currently reinvestigated to draw inference on the current condition of peatland soils. Statistically derived peatland subsidence rates at grassland sites of 0.50 cm/yr and 0.57 cm/yr at arable sites fit well with published values in comparable regions of central Europe. Our results prove the great dynamics of soil development on agriculturally used peatlands. Concerning Brandenburg, the area of peatland soils decreased from 270.000 ha in the early 20th century (Prussian geological map) to actually 163.000 ha.
    Description: Zusammenfassung: Aktuelle Informationen zur Verbreitung und zum Zustand der Moorböden sind vor dem Hintergrund der geplanten Einführung von Agrarumwelt- und Klimamaßnahmen in der Landwirtschaft von besonderer Bedeutung. Das Ministerium für Infrastruktur und Landwirtschaft des Landes Brandenburg (MIL) hat hierzu das Projekt „Schaffung einer Datengrundlage für die Ableitung von Agrarumwelt- und Klimamaßnahmen auf Moorstandorten in Brandenburg“ initiiert. Mit dem Ziel der Bereitstellung einer auf das Jahr 2013 bezogenen Moorkarte wurden umfangreiche, bis dato nicht genutzte Datenbestände zur Verbreitung von Moorböden in Brandenburg zu einer überschneidungsfreien Karte und umfassenden Datenbank moorbodenkundlicher Bodenprofile verarbeitet. Um den aktuellen Zustand der Moorböden zu erfassen, wurden flächenrepräsentativ und zufällig an 7.725 Standorten mit ausreichender Datenbasis erneut bodenkundliche Erhebungen durchgeführt und statistisch ausgewertet. Im Ergebnis konnten Mächtigkeitsverlustraten für landwirtschaftlich genutzte Moorstandorte in Brandenburg abgeleitet werden. Für flachgründige Niedermoorstandorte unter Grünland liegen sie bei 0,50 cm/a, vergleichbare ackerbaulich genutzte Standorte liegen mit 0,57 cm/a darüber. Dies deckt sich gut mit publizierten Werten an vergleichbaren Standorten in Zentraleuropa. Die Fläche der Moorböden in Brandenburg hat sich von Anfang des 20. Jahrhunderts von 270.000 ha, ausgewiesen auf Basis der Preußisch geologisch-agronomische Karte (PGK), auf heute noch 163.000 ha reduziert.
    Description: DFG, SUB Göttingen, DGMT
    Description: research
    Keywords: 553.21 ; Moor ; peatland ; Brandenburg ; Deutschland ; Germany ; FID-GEO-DE-7
    Language: German
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  • 132
    Publication Date: 2021-03-29
    Description: Abstract: Peatlands are intensively exchanging greenhouse gases with the free atmosphere. Natural peatlands, mires, sequester carbon dioxide as carbon in the peat and emit methane. Peatlands under agriculture or forest are mostly drained and emit carbon dioxide due to peat mineralization and nitrous oxide due to fertilization and nitrogen mineralization. In order to characterize the actual situation and to evaluate measures for the reduction of greenhouse gas emissions measured values on gas exchange are needed. Within two national projects direct measurements of the greenhouse gas exchange have been conducted on a wide range of sites, covering different peatland types, land use forms and hydrological situations and using a standardized methodology. Between 2007 and 2012 22 sites have been studied in Lower Saxony over at least two years using the closed cover method. An overview on sites and results is given. Moreover, approaches to estimate the greenhouse gas emissions due to peat excavation and horticultural peat use are presented. Own emission factors and data from literature are used to estimate the contribution of peatlands and peat use to the greenhouse gas emissions of Lower Saxony. Finally measures for the reduction of greenhouse gas emissions are discussed and estimates on reduction costs are given from examples.
    Description: Zusammenfassung: Moore stehen in einem intensiven Gasaustausch mit der Atmosphäre. Natürliche Moore haben über Jahrtausende Kohlendioxid aufgenommen und den Kohlenstoff als Torf gespeichert. Dabei haben sie Methan abgegeben. Entwässerte und genutzte Moore setzen den Kohlenstoff als Kohlendioxid wieder frei und emittieren zudem klimarelevantes Lachgas als Ergebnis der Stickstoffdüngung und der Torfmineralisation. Zur Charakterisierung der Ist-Situation und zur Beurteilung von Maßnahmen ist es erforderlich, belastbare Daten über die Treibhausgasemissionen der Moore zu haben. Im Rahmen von zwei bundesweit angelegten Verbundvorhaben wurden an einer Vielzahl von Standorten unterschiedlichen Moortyps, unterschiedlicher Nutzung und unterschiedlicher Klima- und Wasserregime mit einer abgestimmten Methodik die Flüsse der wesentlichen Treibhausgase direkt gemessen. Im Rahmen dieser Verbundvorhaben wurden in Niedersachsen zwischen 2007 und 2012 an 22 Standorten mindestens zweijährige Messungen mit der Haubentechnik vorgenommen. Dieser Beitrag gibt einen Überblick über untersuchte Standorte und Messergebnisse. Darüber hinaus werden Ansätze zur Ermittlung der Treibhausgasemissionen aus der industriellen Torfgewinnung dargestellt. Unter Verwendung eigener Ergebnisse und von Literaturdaten wird eine Abschätzung der Treibhausgasemissionen aus Moor und Torfgewinnung für das Land Niedersachsen vorgenommen. Abschließend geht der Beitrag auf Maßnahmen zur Verminderung der Treibhausgasemissionen aus Mooren sowie beispielhaft auf Emissionsminderungskosten ein.
    Description: DFG, SUB Göttingen, DGMT
    Description: research
    Keywords: 553.21 ; Moor ; mire ; peatland ; emission ; CO2 ; N2O ; Germany ; Deutschland ; FID-GEO-DE-7
    Language: German
    Type: article , publishedVersion
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  • 133
    facet.materialart.
    Unknown
    Inst. für Phys. Geogr., Freie Univ., Berlin
    In:  Herausgeberexemplar (FU Berlin) | ZB 20559:36
    Publication Date: 2021-03-29
    Description: STÄBLEIN, G.: Regionale Geomorphologie ; KERTESZ, A.: Geomorphologische Kartierung in der Bundesrepublik Deutschland und in Ungarn. Ein Vergleich der Methodik ; BOLLMANN, I.: Geomorphologische Daten und kartographische Darstellung ; PACHUR, H.-J. & RÖPER, H.-P.: Geolimnologische Befunde des Berliner Raumes ; REMMELE, G.: Holozäne Morphodynamik an Schichtstufenhängen — Untersuchungen in Nordwest-Irland ; MÜLLER, MJ. & STRASSER, R.: Holozäne Geomorphodynamik und Landschaftsentwicklung am Ostrand der Trier-Luxemburger Mulde ; KLEBER, A.: Zur jungtertiären Reliefentwicklung im Vorland der südlichen Frankenalb ; HAMANN, C.: Windwurf als Ursache der Bodenbuckelung am Südrand des Tennengebirges, ein Beitrag zur Genese der Buckelwiesen ; ZÖLLER, L.: Neotektonik am Hunsrückrand ; SEPPÄLÄ, M.: Glazialhydrologie des Inlandeises, eine geomorphologische Interpretation der Verhältnisse in Finnland ; JÄKEL, D.: Untersuchungen und Analysen zur Entstehung der Hamada ; GARLEFF, K. & STINGL, H.: Neue Befunde zur jungquartären Vergletscherung in Cuyo und Patagonien ; STINGL, H. & GARLBFF, K.: Tertiäre und pleistozäne Reliefentwicklung an der interozeanischen Wasserscheide in Südpatagonien (Gebiet von Rio Turbio, Argentinien) ; SCHMIDT, K.H.: Nachweis junger Krustenbewegungen auf dem Colorado Plateau, USA ; KUHLE, M.: Zur Geomorphologie Tibets, Bortensander als Kennformen semiarider Vorlandvergletscherung . ; STÄBLEIN, G.: Deutscher Arbeitskreis für Geomorphologie ;
    Description: research
    Description: DFG, SUB Göttingen, FU Berlin
    Keywords: 550 ; 910.02 ; Geomorphologie ; Kartierung ; Holocene ; Irland ; Hunsrück ; Frankenalb ; Finnland ; Hamada ; Patagonien ; Cuyo ; Colorado ; Tektonik ; Gletscher ; glacier ; Tibet ; Deutschland ; Sahara ; Tibesti ; Darfur ; Sudan ; Libyen ; FID-GEO-DE-7
    Language: German
    Type: anthology_digi
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  • 134
    Publication Date: 2021-03-29
    Description: Abstract: In addition to the widespread use of peatlands for agricultural purposes, which is also common in other German states, the raised bogs in Lower Saxony are still also used up to the present day for housing, as well as for industrial peat extraction. The Peatland Conservation Programme in Lower Saxony launched in 1981 therefore only covered raised bogs to disentangle the different uses, and to systematise the nature protection activities in raised bogs. The priority areas for the extraction of peat defined in the Federal State Regional Planning Programme are to be rescinded. The new State Development Programme being formulated will not only include the nature conservation of peatlands, but also for the first time, deal with the climate protection afforded by these very important storage areas for carbon. An important aspect in this regard is the incorporation of fens, and the discharge from peatlands into underlying water bodies. The database and the conclusions which contribute to the implementation of climate protection in peatland conservation programmes, will be discussed and compared with the peatland conservation previously implemented pursuant to the Peatland Conservation Programme in Lower Saxony. An overview describes various approaches for peatland and climate protection implemented by private initiatives, nature conservation societies, EU-Life-Projects, contractual nature protection, publicly owned businesses and authorities in Lower Saxony, research projects on alternative uses, and climate certificate trading. These varied instruments are compared to highlight their relative areal effectiveness, and relevance, followed by an estimate of the potential for future peatland and climate protection in Lower Saxony.
    Description: Zusammenfassung: Neben der auch in anderen Bundesländern weit verbreiteten landwirtschaftlichen Nutzung von Mooren werden die niedersächsischen Hochmoore bis heute auch als Siedlungsraum und zur industriellen Torfgewinnung genutzt. Daher umfasste das niedersächsische Moorschutzprogramm von 1981 nur die Hochmoore, um diese verschiedenen Nutzungen zu entflechten und den Naturschutz in Hochmooren zu systematisieren. Die bisher im Landesraumordnungsprogramm ausgewiesenen Vorranggebiete für die Rohstoffgewinnung von Torf sollen gestrichen werden. In ein neu aufzustellendes Landesentwicklungsprogramm soll neben dem Naturschutz von Mooren erstmalig auch der Klimaschutz dieser wichtigen Kohlenstoffspeicher eingehen. Von besonderer Bedeutung sind hierbei die Einbeziehung der Niedermoore und der Stoffaustrag aus Mooren in unterliegende Gewässer. Die Datengrundlagen und Schlussfolgerungen, die zur Implementierung des Klimaschutzes in den Moorschutz beitragen, werden diskutiert und mit dem bisher umgesetzten Moorschutz nach dem niedersächsischen Moorschutzprogramm verglichen. Ein Überblick beschreibt verschiedene Ansätze zum Moor- und Klimaschutz von Privatinitiativen, von Naturschutzverbänden, mit EU-Life-Projekten, im Vertragsnaturschutz, von niedersächsischen Landesbetrieben und Behörden, von Forschungsprojekten zu alternativen Nutzungen und beim Handel mit Klima-Zertifikaten. Diese unterschiedlichen Instrumente werden in ihrer Flächenwirksamkeit und Relevanz in Beziehung gesetzt und das Potenzial für den künftigen Moor- und Klimaschutz in Niedersachsen abgeschätzt.
    Description: DFG, SUB Göttingen, DGMT
    Description: research
    Keywords: 553.21 ; Moor ; peatland ; mire ; nature conservation ; Germany ; Deutschland ; emission ; bog ; Niedersachsen ; FID-GEO-DE-7
    Language: German
    Type: article , publishedVersion
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  • 135
    Publication Date: 2019-06-22
    Description: A hypersonic flowfield model that treats electronic levels of the dominant afterbody radiator, N, as individual species is presented. This model allows electron-ion recombination rate and two-temperature modeling improvements, the latter which are shown to decrease afterbody radiative heating by up to 30%. This increase is primarily due to the addition of the electron-impact-excitation energy-exchange term to the energy equation governing the vibrational-electronic-electron temperature. This model also allows the validity of the often applied quasi-steady state (QSS) approximation to be assessed. The QSS approximation is shown to fail throughout most of the afterbody region for lower electronic states, although this impacts the radiative intensity reaching the surface by less than 15%. By computing the electronic state populations of N within the flowfield solver, instead of through the QSS approximation in the radiation solver, the coupling of nonlocal radiative transition rates to the species continuity equations becomes feasible. Implementation of this higher- fidelity level of coupling between the flowfield and radiation solvers is shown to increase the afterbody radiation by up to 50% relative to the conventional model.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NF1676L-28417 , Physical Review Fluids (e-ISSN 2469-990X); 3; 1; 013402
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  • 136
    Publication Date: 2019-07-20
    Description: The Orion Crew Module is a component of NASAs Multi-Purpose Crew Vehicle that will be used for future missions to low Earth orbit and beyond. Ten water impact tests of the Orion Ground Test Article (GTA) were conducted at the Hydro Impact Basin at NASA Langley Research Center in 2016 and were designed to provide data for the validation of the LS-DYNA model used to determine the Crew Module structural loads during ocean splashdown, and the determination of an acceptable Model Uncertainty Factor to apply to simulation results used to drive the design. Post-test data obtained from the onboard sensors were used to reconstruct the GTA trajectories both before and after water impact. Results from one vertical test and two swing tests are presented and compared to videos taken for each test.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NF1676L-27423 , AIAA SciTech 2018; Jan 08, 2018 - Jan 12, 2018; Kissimmee, FL; United States
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  • 137
    facet.materialart.
    Unknown
    In:  Other Sources
    Publication Date: 2019-07-20
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-7132
    Format: text
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  • 138
    Publication Date: 2019-07-12
    Description: Polymers and other oxidizable materials on the exterior of spacecraft in the low Earth orbit (LEO) space environment can be eroded due to reaction with atomic oxygen (AO). Therefore, in order to design durable spacecraft, it is important to know the LEO AO erosion yield (Ey, volume loss per incident oxygen atom) of materials susceptible to AO reaction. The Polymers Experiment was developed to determine the AO Ey of various polymers and other materials flown in ram and wake orientations in LEO. The experiment was flown as part of the Materials International Space Station Experiment 7 (MISSE 7) mission for 1.5 years on the exterior of the International Space Station (ISS). As part of the experiment, a sample containing Class 2A diamond (100 plane) and highly oriented pyrolytic graphite (HOPG, basal and edge planes) was exposed to ram AO and characterized for erosion. The materials were salt-sprayed prior to flight to provide isolated sites of AO protection. The Ey of the samples was determined through post-flight electron microscopy recession depth measurements. The experiment also included a Kapton H witness sample for AO fluence determination. This paper provides an overview of the MISSE 7 mission, a description of the flight experiment, the characterization techniques used, the mission AO fluence, and the LEO Ey results for diamond and HOPG (basal and edge planes). The data is compared to the Ey of pyrolytic graphite exposed to four years of space exposure as part of the MISSE 2 mission. The results indicate that diamond erodes, but with a very low Ey of 1.58 +/- 0.04 x 10(exp -26) cm(exp 3)/atom. The different HOPG planes displayed significantly different amounts of erosion from each other. The HOPG basal plane had an Ey of 1.05 +/- 0.08 x 10(exp -24) cm(exp 3)/atom while the edge plane had a lower Ey of only 5.38 +/- 0.90 x 10(exp -25) -cm(exp 3)/atom. The Ey data from this ISS spaceflight experiment provides valuable information for understanding of chemistry and chemical structure dependent modeling of AO erosion.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NASA/TM-2018-219756 , E-19468 , GRC-E-DAA-TN51758
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  • 139
    Publication Date: 2019-07-12
    Description: Current concepts of operations for human exploration of Mars center on the staged deployment of spacecraft, logistics, and crew. Though most studies focus on the needs for human occupation of the spacecraft and habitats, these resources will spend most of their lifetime unoccupied. As such, it is important to identify the operational state of the unoccupied spacecraft or habitat, as well as to design the systems to enable the appropriate level of autonomy. Key goals for this study include providing a realistic assessment of what "dormancy" entails for human spacecraft, exploring gaps in state-of-the-art for autonomy in human spacecraft design, providing recommendations for investments in autonomous systems technology development, and developing architectural requirements for spacecraft that must be autonomous during dormant operations. The mission that was chosen is based on a crewed mission to Mars. In particular, this study focuses on the time that the spacecraft that carried humans to Mars spends dormant in Martian orbit while the crew carries out a surface mission. Communications constraints are assumed to be severe, with limited bandwidth and limited ability to send commands and receive telemetry. The assumptions made as part of this mission have close parallels with mission scenarios envisioned for dormant cis-lunar habitats that are stepping-stones to Mars missions. As such, the data in this report is expected to be broadly applicable to all dormant deep space human spacecraft.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NASA/TM-2018-219965
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  • 140
    Publication Date: 2019-07-12
    Description: As spacecraft travel through space plasma, spacecraft surfaces become charged by the collection of charged particles. This process is referred to as Surface Charging. These charges can be detrimental to the vehicle's electronic subsystems as they present a threat of electrostatic discharge (ESD) to onboard circuitry. The process of Surface Charging is complex and is affected by many elements. The charging of each surface is unique. The potential of an individual surface is dependent upon many variables including but not limited to the surface's geometry, material and its location. Each surface also has unique interactions with the surrounding plasma. Other factors that play large roles in the charging process is the density and temperature of plasma ions and electrons. Using Nascap-2k, a model of the Freja satellite has been constructed, and its auroral plasma environment has been imitated to simulate surface charging characteristics. The charging process of the Freja satellite has been modeled iteratively with incremental changes in both the Maxwellian electron temperature (eV) as well as the Gaussian electron energy (eV). This study provides an analysis of the sensitivity between spacecraft surface charging and these two primary variables of electron differential flux.
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-6709
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  • 141
    Publication Date: 2019-07-12
    Description: As spacecraft travel through plasma, spacecraft surfaces become charged by the collection of charged particles. This process is referred to as Surface Charging. These charges can be detrimental to the vehicle's electronic subsystems as they present a threat of electrostatic discharge (ESD) to onboard circuitry. The process of Surface Charging is complex and is affected by many elements. The charging of each surface is unique. The potential of an individual surface is dependent upon many variables including but not limited to the surface's geometry, material and its location. Each surface also has unique interactions with the surrounding plasma. Other factors that play large roles in the charging process is the density and temperature of plasma ions and electrons. Using Nascap-2k, a model of the Freja satellite has been constructed, and its auroral plasma environment has been imitated to simulate surface charging characteristics. The charging process of the Freja satellite has been modeled iteratively with incremental changes in both the Maxwellian electron temperature (eV) as well as the Gaussian electron energy (eV). This study provides an analysis of the sensitivity between spacecraft surface charging and these two primary variables of electron differential flux.
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-6708
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  • 142
    Publication Date: 2019-07-20
    Description: A wealth of literature exists on control allocation algorithms for over-actuated air vehicles, launch vehicles, and spacecraft's. Most of these algorithms focus primarily on minimizing some objective function such as command tracking error and/or control effector usage. Linear allocators (pseudo inverses) are usually the conventional choice due to their simplicity and the ability to achieve a significant portion of the theoretical moment/impulse space. Generally, it is assumed that there exists minimal interaction effects between control effectors. In fact, very few studies address the problem of control effector interactions in the context of control allocation, especially for small spacecraft's with a reaction control system (RCS). This paper presents a CubeSat RCS design with a four thruster tetrahedral layout such that when two or more thrusters re, the resultant impulse differs noticeably compared to the sum of the contributions from individual thruster rings. This undesirable effect is caused by the design of the propellant tank and regulator. To mitigate this issue, an innovative modified pseudo inverse (MPI) control allocation algorithm was developed that adjusts the pseudo inverse solution based on test data. The algorithm is iteration-free and superior to the standard pseudo inverse in minimizing the command tracking error.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NF1676L-27385 , AIAA Science and Technology Forum and Exposition; Jan 08, 2018 - Jan 12, 2018; Kissimmee, FL; United States
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  • 143
    Publication Date: 2019-07-20
    Description: Here we describe the Primitive Object Volatile Explorer (PrOVE), a smallsat mission concept to study the surface structure and volatile inventory of comets in their perihelion passage phase when volatile activity is near peak. CubeSat infrastructure imposes limits on propulsion systems, which are compounded by sensitivity to the spacecraft disposal state from the launch platform and potential launch delays. We propose circumventing launch platform complications by using waypoints in space to park a deep space SmallSat or CubeSat while awaiting the opportunity to enter a trajectory to flyby a suitable target. In our Planetary Science Deep Space SmallSat Studies (PSDS3) project, we investigated scientific goals, waypoint options, potential concept of operations (ConOps) for periodic and new comets, spacecraft bus infrastructure requirements, launch platforms, and mission operations and phases. Our payload would include two low-risk instruments: a visible image (VisCAM) for 5-10 m resolution surface maps; and a highly versatile multispectral Comet CAMera (ComCAM) will measure 1) H2O, CO2, CO, and organics non-thermal fluorescence signatures in the 2-5 m MWIR, and 2) 7-10 and 8-14 m thermal (LWIR) emission. This payload would return unique data not obtainable from ground-based telescopes and complement data from Earth-orbiting observatories. Thus, the PrOVE mission would (1) acquire visible surface maps, (2) investigate chemical heterogeneity of a comet nucleus by quantifying volatile species abundance and changes with solar insolation, (3) map the spatial distribution of volatiles and determine any variations, and (4) determine the frequency and distribution of outbursts.
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN65939 , Proceedings Volume 10769, CubeSats and NanoSats for Remote Sensing II; 10769; 107690J-7|SPIE Optical Engneering + Appliactions; Aug 11, 2018 - Aug 15, 2018; San Diego, California; United States
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  • 144
    Publication Date: 2019-07-20
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-6827-2 , AIAA Propulsion and Energy Forum; Jul 09, 2018 - Jul 11, 2018; Cincinnati, OH; United States
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  • 145
    Publication Date: 2019-07-20
    Description: Much effort has been made to enhance exploration on Mars. In addition to a rover and Mars-orbiting satellites, a Mars helicopter (MH) was proposed in order to augment planetary research. Computational Fluid Dynamics (CFD) simulations have been performed to have a better understanding of the behavior and performance of vertical lift Planetary Aerial Vehicles (PAV). Due to the large differences in atmospheric conditions between Mars and Earth, predicting and testing rotorcraft performance is a complex task. The goal of this project is to understand the capability of the mid-fidelity CFD software RotCFD to predict rotor performance in terms of thrust at 1013.25 milibar and 14 milibar corresponding to Terrestrial and Martian conditions, respectively. Also, in order to characterize the wind tunnel wall effects free field and wind tunnel simulations were performed, analyzed and compared. Different analytical tools have been used in order to aid with the design process for the future vertical lift planetary aerial vehicles. One of them includes experimental tests performed on a rotor in the Aeolian Wind Tunnel (AWT) facility at NASA Ames Research Center under different pressure conditions ranging from Terrestrial to Martian atmospheric conditions. Other software was used as well in order to capture the aerodynamic coefficients of the corresponding rotor sections based on the Mach and Reynolds numbers used for the experimental tests. The aerodynamic coefficients were input into RotCFD, and various simulations were performed under Terrestrial and Martian conditions in order to mimic the experimental test. Then, the obtained results from RotCFD were compared with the AWT collected data.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NASA/CR-2018-219780 , ARC-E-DAA-TN53293
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  • 146
    Publication Date: 2019-07-20
    Description: Aerocapture has been extensively studied and these studies have shown the benefit for planetary exploration missions. While the traditional approach to aerocapture with lifting configurations and lift-guided modulations have been assessed to be technologically feasible, aerocapture using purely drag modulation was proposed and studied by Prof. Braun and his students. These studies show that if one can assess the feasibility of aerocapture using drag modulation at Venus, and develop tall pole technologies needed at Venus, then this concept is much easier to execute at all other relevant destinations. Based on the above finding, partnered proposals were submitted by Adam Nelessen at JPL and Ethiraj Venkatapathy at Ames in collaboration with Prof. Braun at the University of Colorado, Boulder (UCB). Under this partnership, Ames Research Center (ARC) is working to address some of the key entry technology challenges associated with drag modulation aerocapture at Venus. Drag modulation aerocapture is a simple, scalable, and likely cost-effective way to enhance planetary science missions. The approach envisioned is to design a small spacecraft, that would most likely be a secondary payload, with a removable drag skirt. The vehicle would enter the atmosphere at Venus with a low ballistic coefficient, decelerate rapidly, drop the skirt resulting in a smaller vehicle with a higher ballistic coefficient which would skip out of the atmosphere and enter into a desired orbit. ARC's role in this collaboration is multifold. First of which is to perform design studies on various pre- and post-jettison geometries utilizing a 3-DOF trajectory code to determine the aerodynamics and aerothermodynamics of the vehicles and evaluate viable thermal protection material system designs. Once these design studies are complete, Ames will then perform higher fidelity CFD and TPS sizing to further design the vehicles. Second, the multi-body separation dynamics of the drag modulation event will be explored using both CFD simulations (CART3-D and US3D) as well as possible ballistic range testing. ARC's tools and expertise have been used to assess and advise on the selection of the separating configuration. In addition to the preliminary evaluation, ARC will provide tools and expertise to UCB team members to further assess aerodynamic interactions between the separating bodies and provide guidance as to the feasibility of stable transition.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN57402 , International Planetary Probe Workshop; Jun 11, 2018 - Jun 15, 2018; Boulder, CO; United States
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  • 147
    Publication Date: 2019-07-20
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN55686 , Annual CubeSat Developers Workshop; Apr 30, 2018 - May 02, 2018; San Luis Obispo, CA; United States
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  • 148
    Publication Date: 2019-07-13
    Description: The NASA Glenn Research Center (GRC) in Cleveland, Ohio designs and develops innovative technologies to advance NASA's missions in aeronautics and space exploration. The center's expertise includes that in power, energy storage, and conversion; in-space chemical and electric propulsion; communications; and instrumentation technologies. GRC is currently managing and/or developing a number of these technologies for Small Spacecraft applications. Small spacecraft propulsion efforts include efforts with Tethers Unlimited, Inc. (TUI) and Busek. Power systems technology efforts include the Advanced Electrical Bus (ALBus) CubeSat inhouse development as well as efforts with Rochester Institute of Technology (RIT), the Kennedy Space Center & the University Miami. In the area of communications, NASA-GRC continues to explore the potential capabilities and advantages of using Ka-band for LEO (Low Earth Orbit) spacecraft communications with both NASA and commercially owned GEO (Geosynchrous Earth Orbit) relays and direct-to-ground terminal networks. GRC has also proposed a number of small spacecraft instrumentation technology demonstration such as SPAGHETI (Solar Proton Anisotropy and Galactic cosmic ray High Energy Transport Instrument) and CFIDS (Compact Full-Field Ion Detector System).
    Keywords: Spacecraft Design, Testing and Performance
    Type: GRC-E-DAA-TN59063 , AIAA/USU Conference on Small Satellites; Aug 04, 2018 - Aug 09, 2018; Logan, UT; United States
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  • 149
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-6841 , AIAA Propulsion and Energy Conference; Jul 09, 2018 - Jul 11, 2018; Cincinnatti, OH; United States
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  • 150
    Publication Date: 2019-07-13
    Description: A regeneratively-cooled nozzle for liquid rocket engine applications is a significant cost of the overall engine due to the complexities of manufacturing a large thin-walled structure that must operate in extreme temperature and pressure environments. NASA has been investigating and advancing methods for fabrication of liquid rocket engine channel wall nozzles to realize further cost and schedule improvements. The methods being evaluated are targeting increased scale required for current NASA and commercial space programs. Several advanced rapid fabrication methods are being investigated for forming of the inner liner, producing the coolant channels, closeout of the coolant channels, and fabrication of the manifolds. NASA Marshall Space Flight Center (MSFC) completed process development and subscale hot-fire testing of a series of these advanced fabrication channel wall nozzle technologies to gather performance data in a relevant environment. The primary fabrication technique being discussed in this paper is Laser Wire Deposition Closeout (LWDC). This process has been developed to significantly reduce time required for closeouts of regeneratively-cooled slotted liners. It allows for channel closeout to be formed in place in addition to the structural jacket without the need for channel fillers or complex tooling. Additional technologies were also tested as part of this program including water jet milling and arc-based additive manufacturing deposition. Each nozzle included different fabrication features, materials, and methods to demonstrate durability in a hot-fire environment. The results of design, fabrication and hot-fire testing performance is discussed in this paper.
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-6464 , AIAA Propulsion and Energy Forum; Jul 09, 2018 - Jul 11, 2018; Cincinnati, OH; United States
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  • 151
    Publication Date: 2019-07-13
    Description: The Space Launch System (SLS) Block-1B vehicle includes a low thrust-to-weight upper stage, which presents challenges to heritage ascent guidance algorithms. A trade study was conducted to evaluate two alternative guidance algorithms: 1) Powered Explicit Guidance (PEG), based on a modified implementation of PEG used on the Block-1 vehicle, and 2) Optimal Guidance (OPGUID), an algorithm developed for Marshall Space Flight Center (MSFC) and used on Constellation and other Guidance, Navigation, and Controls (GN&C) projects. The design criteria, approach, and results of the trade study are given, as well as other impacts and considerations for Block-1B type missions.
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-6865 , 2018 AAS/AIAA Astrodynamics Specialist Conference; Aug 19, 2018 - Aug 23, 2018; Snowbird, UT; United States
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  • 152
    Publication Date: 2019-07-13
    Description: This paper details the results of an initial study to develop a certification plan for human-rated inflatable space structures, including guidelines for qualification testing. Habitable softgoods inflatables are multi-layered shell structures that use high-strength webbing, cordage and broadcloth fabric to carry the skin loads of a variety of volumetric shapes and structural architectures. The primary objectives of this study are to define the key parameters that affect these structures and propose a statistically robust approach to defining safety and knockdown factors based on test and analysis. Current NASA standards for habitable inflatable space structures use a factor of safety of 4, which was inherited from airship design criteria. An updated approach to defining a design factor, taking into account material strength variability, load variability in the article, number of test samples, and damage and degradation effects is specified. Accurate analytical modeling of these structures is hindered by the difficulty of obtaining accurate and consistent material data due to load-history- dependent, nonlinear load versus strain behavior. A building block approach to certification is detailed that uses stochastic modeling and statistical test design and analysis to address the unique challenges these high-strength softgoods structures present. Human-rated inflatable modules are a transformative capability for launching much larger habitable volumes into space than is possible with rigid shell structures. This research aims to provide the framework for certifying these structures for future human space exploration missions.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NF1676L-27608 , IEEE Aerospace Conference; Mar 03, 2018 - Mar 10, 2018; Big Sky, MT; United States
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  • 153
    Publication Date: 2019-07-13
    Description: Final Document is attached. The Robotic External Leak Locator (RELL) was deployed to the International Space Station (ISS) with the goal of detecting and locating on-orbit leaks around the ISS. Three activities to investigate and corroborate the background natural and induced environment of ISS were performed with RELL as part of the on-orbit validation and demonstration conducted in November December 2016. The first demonstration activity pointed RELL directly in the ram and wake directions for one orbit each. The ram facing measurements showed high partial pressure for mass-to-charge ratio 16, corresponding to atomic oxygen (AO), as well as the presence of mass-to-charge ratio 17. RELLs view in the wake-facing direction included more ISS structure and several Environmental Control and Life Support System (ECLSS) on-orbit vents were detected, including the Carbon Dioxide Removal Assembly (CDRA), Russian segment ECLSS, and Sabatier vents. The second demonstration activity pointed RELL at three faces of the P1 Truss segment. Effluents from ECLSS and European Space Agency (ESA) Columbus module on-orbit vents were detected by RELL. The partial pressures of mass-to-charge ratios 17 and 18 remained consistent with the first on-orbit activity of characterizing the natural environment. The third demonstration activity involved RELL scanning an Active Thermal Control System (ATCS) radiator. Three locations along the radiator were scanned and the angular position of RELL with respect to the radiator was varied. Mass-to-charge ratios 16 and 17 both had upward shifts in partial pressure when pointing toward the Radiator Beam Valve Modules (RBVMs), likely corresponding to a known, small ammonia leak.
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN58665 , SPIE Optical Engineering + Applications Symposium; Aug 19, 2018 - Aug 23, 2018; San Diego, CA; United States
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  • 154
    Publication Date: 2019-07-13
    Description: The 3rd Planetary CubeSat Science Symposium will be held at NASA Goddard Space Flight Center, with the participation of CubeSat/SmallSat scientists and developers. Discussions will include current missions, mission concepts, and opportunities for future mission selections. The sessions will also include panel discussions about strategic and technical aspects of planetary small satellite missions, and an afternoon poster session providing mission proposers the opportunity to meet with vendors and suppliers. This presentation (no paper), will provide an overview of the navigation systems avaiable for Cubesat Planetary missions.
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN59777 , Planetary CubeSat Science Symposium; Aug 16, 2018 - Aug 17, 2018; Greenbelt, MD; United States
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  • 155
    Publication Date: 2019-07-13
    Description: Phenolic Impregnated Carbon Ablator (PICA), invented in the mid 1990's, is a low-density ablative thermal protection material proven capable of meeting sample return mission needs from the moon, asteroids, comets and other "unrestricted class V destinations" as well as for Mars. Its low density and efficient performance characteristics have proven effective for use from Discovery to Flagship class missions. It is important that NASA maintain this TPS material capability and ensure its availability for future NASA use. The rayon based carbon precursor raw material used in PICA preform manufacturing required replacement and requalification at least twice in the past 25 years and a third substitution is now needed. The carbon precursor replacement challenge is twofold the first involves finding a long-term replacement for the current rayon and the second is to assess its future availability periodically to ensure it is sustainable and be alerted if additional replacement efforts need to be initiated. Rayon is no longer a viable process in the US and Europe due to environmental concerns. In the early 80's rayon producers began investigating a new method of producing a cellulosic fiber through a more environmentally responsible process. This cellulosic fiber, lyocell, is a viable replacement precursor for PICA fiberform. This presentation reviews current SMD-PSD funded PICA sustainability activities in ensuring a rayon replacement for the long term is identified and in establishing that the capability of the new PICA derived from an alternative precursor is in family with previous versions of the so called "heritage" PICA.State of the Art Low Density Carbon Phenolic AblatorsStardust SRC post flight withPICA forebody heat shield(0.8m max. diameter)PICA Processing StepsRole of Rayon/Lyocellin PICA.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN57669 , National Space and Missile Material Symposium (NSMMS); Jun 25, 2018 - Jun 28, 2018; Madison, WI; United States
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  • 156
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    In:  CASI
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-6627 , Presentation to Louisiana State University; Apr 05, 2018; Baton Rouge, LA; United States
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  • 157
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN56664 , Constellation Mission Operations Working Group (MOWG); Jun 12, 2018 - Jun 14, 2018; Sioux Falls, SD; United States
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  • 158
    Publication Date: 2019-07-13
    Description: This presentation introduces a new sizing and margin methodology for dual-layer Thermal Protection Systems (TPS). The methodology has been tailored for application to a dual-layer 3D-woven TPS called Heat-shield for Extreme Entry Environments Technology (HEEET). Sizing is performed for a reference Saturn probe mission to show how uncertainties in trajectory, aerothermal modelling and TPS response impact the sizing of each layer.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN57591 , International Planetary Probe Workshop; Jun 11, 2018 - Jun 15, 2018; Boulder, CO; United States
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  • 159
    Publication Date: 2019-07-13
    Description: The Origins, Spectral Interpretation, Resource Identification, Security, Regolith Explorer (OSIRIS-REx) Visible and Infrared Spectrometer (OVIRS) is a cryogenic instrument. At the Outbound Cruise nominal spacecraft attitude, sunlight impinges on several multilayer insulation blankets on the forward deck. It is reflected or scattered to other components on the deck. This solar illumination adds heat load to the OVIRS, and causes its detector temperature to exceed the 105K maximum operating allowable flight temperature limit by 0.8K. During the flight system thermal vacuum test, the solar simulator beam reflected or scattered from the test fixtures to the OVIRS added non-flight heat load. The detector temperature was 9K warmer than that in flight. At those temperatures, the science data was acceptable, despite its quality was not as high as that of 105K or colder.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ICES-2018-008 , GSFC-E-DAA-TN56295 , International Conference on Environmental Systems; Jul 08, 2018 - Jul 12, 2018; Albuquerque, NM; United States
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  • 160
    Publication Date: 2019-07-13
    Description: This paper summarizes the on-orbit structural dynamic data and the related modal analysis, model validation and correlation performed for the International Space Station (ISS) configuration ISS Stage ULF7, 2015 Dedicated Thruster Firing (DTF). The objective of this analysis is to validate and correlate the analytical models used to calculate the ISS internal dynamic loads and compare the 2015 DTF with previous tests. During the ISS configurations under consideration, on-orbit dynamic measurements were collected using the three main ISS instrumentation systems; Internal Wireless Instrumentation System (IWIS), External Wireless Instrumentation System (EWIS) and the Structural Dynamic Measurement System (SDMS). The measurements were recorded during several nominal on-orbit DTF tests on August 18, 2015. Experimental modal analyses were performed on the measured data to extract modal parameters including frequency, damping, and mode shape information. Correlation and comparisons between test and analytical frequencies and mode shapes were performed to assess the accuracy of the analytical models for the configurations under consideration. These mode shapes were also compared to earlier tests. Based on the frequency comparisons, the accuracy of the mathematical models is assessed and model refinement recommendations are given. In particular, results of the first fundamental mode will be discussed, nonlinear results will be shown, and accelerometer placement will be assessed.
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN52496 , International Modal Analysis Conference; Feb 12, 2018 - Feb 15, 2018; Orlando, FL; United States
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  • 161
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-7013 , Aerospace Control and Guidance Systems Committee (ACGSC); Oct 09, 2018 - Oct 12, 2018; Savannah, GA; United States
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  • 162
    Publication Date: 2019-07-13
    Description: The Mars2020 entry vehicle is currently being developed by NASA to safely land its next rover on the Martian surface in 2021. During entry, the vehicle will be protected from aerothermal environments using a PICA (Phenolic Impregnated Carbon Ablator)-tiled heatshield. PICA loses mass through surface recession and in-depth pyrolysis as it is heated. Pre-flight knowledge of heatshield mass loss is required for vehicle balancing during critical mission events. This study attempts to predict the total mass loss experienced by the Mars2020's heatshield during its entry. A grid was created over the half of the heatshield which generated 108 points across a total of 9 spokes. Aero-thermal environments were provided from CFD (Computational Fluid Dynamics) calculations that considered a baselined trajectory. The TPS (Thermal Protection System) stack was a build-up of composite, aluminum, composite, an HT-424 bond, followed by PICA. The FIAT (Fully Implicit Ablation, Thermal-response) 1-D analysis utilized this TPS stack and the CFD environments and was run at each grid point giving mass flux information from the point of atmospheric entry until parachute deployment. The mass flux due to recession and pyrolysis gas was summed and integrated first through time and then across the half heatshield using a polar integration tool. The mass loss results were mirrored to the other half of the heatshield to calculate total mass loss throughout the entry phase of flight. This total mass loss value and its distribution was used by entry vehicle designers to account for CG (Center of Gravity) offset during parachute descent when the heatshield is no longer losing significant mass.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN58301 , AIAA Aviation and Aeronautics Forum (Aviation 2018); Jun 25, 2018 - Jun 29, 2018; Atlanta, GA; United States|AIAA/ASME Joint Thermophysics and Heat Transfer Conference (2018); Jun 25, 2018 - Jun 29, 2018; Atlanta, GA; United States
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  • 163
    Publication Date: 2019-07-13
    Description: This paper presents an overview of the design optimisation measures that have been proposed and analysed in order to reduce the mass of the structure, including the MMOD (Micro-Meteoroid and Orbital Debris) protection system, of the ESM (European Service Module) for the Orion MPCV (Multi-Purpose Crew Vehicle). Under an agreement between NASA and ESA, the NASA Orion MPCV for human space exploration missions will be powered by a European Service Module, based on the design and experience of the ATV (Automated Transfer Vehicle). The development and qualification of the European Service Module is managed and implemented by ESA. The ESM prime contractor and system design responsible is Airbus Defence and Space. Thales Alenia Space Italia is responsible for the design and integration of the ESM Structure and MMOD protection system in addition to the Thermal Control System and the Consumable Storage System. The Orion Multi-Purpose Crew Vehicle is a pressurized, crewed spacecraft that transports up to four crew members from the Earths surface to a nearby destination or staging point. Orion then brings the crew members safely back to the Earths surface at the end of the mission. Orion provides all services necessary to support the crew members while on-board for short duration missions (up to 21 days) or until they are transferred to another orbiting habitat. The ESM supports the crew module from launch through separation prior to re-entry by providing: in-space propulsion capability for orbital transfer, attitude control, and high altitude ascent aborts; water and oxygen/nitrogen needed for a habitable environment; and electrical power generation. In addition, it maintains the temperature of the vehicle's systems and components and offers space for unpressurized cargo and scientific payloads. The ESM has been designed for the first 2 Lunar orbit missions, EM-1 (Exploration mission 1) is an un-crewed flight planned around mid-2020, and EM-2, the first crewed flight, is planned in 2022. At the time where the first ESM is about to be weighted, the predicted mass lies slightly above the initial requirement. For future builds, mass reduction of the Service Module has been considered necessary. This is being investigated, together with other design improvements, in order to consolidate the ESM design and increase possible future missions beyond the first two Orion MPCV missions. The mass saving study has introduced new optimised structural concepts, optimisation of the MMOD protection shields, and optimised redesign of parts for manufacturing through AM (Additive Manufacturing).
    Keywords: Spacecraft Design, Testing and Performance
    Type: IAC-18,C2,1,11,x48504 , GRC-E-DAA-TN61395 , International Astronautical Congress (IAC); Oct 01, 2018 - Oct 05, 2018; Bremen; Germany
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  • 164
    Publication Date: 2019-07-13
    Description: The interaction of on-axis and o -axis laser discharge in front of a hemisphere cylinder in Mach 2.0 ow is investigated numerically. Details of the physics of the interaction of the laser-induced shock and the heated region with the bow shock and its e ect on drag reduction are included. The energetic eciency of the laser discharge in reducing drag is calculated.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NF1676L-28965 , AIAA SciTech; Jan 08, 2018 - Jan 12, 2018; Kissimmee, FL; United States
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  • 165
    Publication Date: 2019-07-13
    Description: Active flow control (AFC) in the form of sweeping jet (SWJ) excitation and discrete steady jet excitation is used to control the flow separation on an NACA 0015 semispan wing with a deflected, simple-hinged, trailing edge flap. This geometry has been the focus of several recent publications that investigated methods to improve the efficiency of sweeping jet actuators. In the current study, the interaction of the AFC excitation with the separated flowfields present at several flap deflection angles was examined. Previous studies with this model have been limited to a maximum flap deflection angle of 40. The flap deflection range was extended to 60! because systems studies have indicated that a high-lift system with simple-hinged flaps may require larger flap deflections than the Fowler flaps found on most high-lift systems. The results obtained at flap deflection angles of 20, 40, and 60 are presented and compared. Force and moment data, Particle Image Velocimetry (PIV) data, and steady and unsteady surface pressure data are used to describe the flowfield with and without AFC. With a flap deflection of 60, increasing the SWJ actuator momentum at the flap shoulder increased lift due to an increase in circulation but did not completely eliminate the recirculation region above the flap surface. AFC using the discrete steady jet actuators of this study increased lift as well but required more mass flow than the SWJ actuators and had a detrimental effect on lift at the highest mass flow level tested. PIV results showed that the angle between the excitation and the flap surface was not optimal for attaching the separated shear layer.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NF1676L-28928 , AIAA SciTech; Jan 08, 2018 - Jan 12, 2018; Kissimmee, FL; United States
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  • 166
    Publication Date: 2019-07-13
    Description: The FAST-MAC circulation control model was modified to test an array of steady and unsteady actuators at realistic flight Reynolds numbers in the National Transonic Facility at the NASA Langley Research Center. Previous experiments in the FAST-MAC test series used a fullspan tapered slot, and that configuration is used as a baseline for performance and weight flow requirements. The goal of the latest experiment was to reduce the weight flow required to achieve comparable performance established by the baseline FAST-MAC data. Thirty-nine interchangeable actuator cartridges of various designs were mounted into the FAST-MAC model where the exiting jet was directed over a 15% chord simple hinged-flap. These two types of actuators were fabricated using rapid prototype techniques and their design performance was optimized for a transonic cruise configuration having a 0 flap deflection. The steady actuators were found to provide an off-design drag reduction of 5.5%, nearly equaling the drag reduction of the fullspan tapered slot configuration, but with a 69% weight flow reduction. This weight flow savings is similar to the sweeping jet actuators, but with better drag performance.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NF1676L-28921 , AIAA SciTech; Jan 08, 2018 - Jan 12, 2018; Kissimmee, FL; United States
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  • 167
    Publication Date: 2019-07-13
    Description: Various methods for remote recession sensing of PICA have been developed and several seeding methods have been tested. The most recent method involved seeding the ablator with wires fed to the sample from the backside with a defined amount of PICA left towards the upstream front of the sample. This seed method mimics the installation of in-depth thermocouples as they are frequently used in ground testing and flight. Arc-jet tests were conducted in the NASA Langley HYMETS facility at a heat flux of 320 W/sq.cm. The emission of the post-shock layer was observed in spectral resolution from the side along an optical axis perpendicular to the arc-jet flow and from the front, looking at the sample surface from an upstream position. Various metallic seed materials with different melting points were used. In addition to the emission spectroscopy measurements, the samples were monitored during the tests through pyrometry and videography. The time resolved response of the seeded material is described and compared to earlier tests with different seeding methods. The combination of seed materials was found to be critical for the selection of emission signatures characteristic for the material recession which can be isolated in the final emission spectra.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NF1676L-27563 , AIAA SciTech; Jan 08, 2018 - Jan 12, 2018; Kissimmee, FL; United States
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  • 168
    Publication Date: 2019-07-13
    Description: The Molecular Adsorber Coating (MAC) is a sprayable coatings technology that was developed at NASA Goddard Space Flight Center (GSFC). The coating is comprised of highly porous, zeolite materials that help capture outgassed molecular contaminants on spaceflight applications. The adsorptive capabilities of the coating can alleviate molecular contamination concerns on or near sensitive surfaces and instruments within a spacecraft. This paper will discuss the preliminary testing of NASA's MAC technology for use on future missions to Mars. The study involves evaluating the coating's molecular adsorption properties in simulated test conditions, which include the vacuum environment of space and the Martian atmosphere. MAC adsorption testing was performed using a commonly used plasticizer called dioctyl phthalate (DOP) as the test contaminant.
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN59323 , SPIE Optics and Photonics 2018; Aug 19, 2018 - Aug 23, 2018; San Diego, CA; United States
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  • 169
    Publication Date: 2019-07-13
    Description: Final document is attached. This paper proposes an enhanced control technique for stationkeeping maneuvers to reduce delta-v costs for the Korea Pathfinder Lunar Orbiter (KPLO). A scheduled circularization control technique exploits patterns in the evolution of the line of apsides and eccentricity to achieve a significant reduction in stationkeeping delta-v costs based on spacecraft requirements. The technique is compared against previous algorithms implemented for maneuver operations of the Lunar Prospector and Lunar Reconnaissance Orbiter (LRO) missions in the USA and KAGUYA in Japan. Through Monte Carlo analysis, the efficacy and robustness of the proposed method are verified, and the technique is shown to meet the operational requirements of KPLO.
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN60023 , AAS Astrodynamics Specialists Conference; Aug 19, 2018 - Aug 23, 2018; Snowbird, Ut; United States
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  • 170
    Publication Date: 2019-07-13
    Description: Final document not an Abstract attached. The International Space Station (ISS) has been on-orbit for nearly 20 years, and there have been numerous technical challenges along the way from design to assembly to on-orbit anomalies and repairs. The Passive Thermal Control System (PTCS) management team has been a key player in successfully dealing with these challenges. The PTCS team performs thermal analysis in support of design and verification, launch and assembly constraints, integration, sustaining engineering, failure response, and model validation. This analysis is a significant body of work and provides a unique opportunity to compile a wealth of real world engineering and analysis knowledge and the corresponding lessons-learned. The PTCS lessons encompass the full life cycle of flight hardware from design to on-orbit performance and sustaining engineering. These lessons can provide significant insight for new projects and programs. Key areas to be presented include thermal model fidelity, verification methods, analysis uncertainty, and operations support.
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN59953 , Thermal and Fluids Analysis Workshop; Aug 20, 2018 - Aug 24, 2018; Galveston, TX; United States
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  • 171
    Publication Date: 2019-07-13
    Description: The flight focal plane array (FPA) for the Thermal Infrared Sensor 2 (TIRS-2) instrument, to be flown on Landsat 9, was built and characterized at NASA Goddard Space Flight Center (GSFC). The FPA was assembled using GaAs quantum well infrared photodetector (QWIP) arrays from the same lot as the TIRS instrument on Landsat 8. Each QWIP array is hybridized to an Indigo ISC9803 readout integrated circuit (ROIC) with 640 x 512, 25m by 25m pixels. Each QWIP hybrid was tested at the NASA/GSFC Detector Characterization Laboratory (DCL) as a single sensor chip assembly (SCA). The best SCAs in terms of performance were then built up into an FPA consisting of three SCAs, required to provide the necessary 15-degree field of view of the instrument. The FPA was tested to determine if project requirements were being met as a fully assembled unit. The performance of the QWIP SCAs and the fully assembled, NASA flight-qualified FPA will be reviewed.
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN60078 , SPIE Remote Sensing; Sep 10, 2018 - Sep 13, 2018; Berlin; Germany
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  • 172
    Publication Date: 2019-07-13
    Description: Direct Field Acoustic Testing (DFAT) offers potential cost and time savings over reverberant chamber acoustic testing of spacecraft. The NASA Multi-Purpose Crew Vehicle (MPCV) Program recently directed a series of acoustic tests on Orion structural test articles comparing DFAT and reverberant testing of the same test article with a view to qualifying DFAT for manned space flight vehicles. The verification process compared four parameters noise level compliance with the one third octave test specification, spatial uniformity of the acoustic field, spatial correlation of the acoustic field and vibration response of vehicle structure, including representative solar array panels. While the results of the verification were encouraging, MPCV Loads and Dynamics engaged Quartus Engineering to investigate whether alternative MIMO random control strategies might improve the spatial uniformity and/or the spatial correlation of the DFAT acoustic field. This paper presents the results of acoustic field simulations of the DFAT test and provides a better understanding of how MIMO random control systems originally developed for vibration and structural durability testing can be expected to perform in DFAT testing.
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN57215 , Spacecraft and Launch Vehicle Dynamic Environments Workshop; Jun 26, 2018 - Jun 28, 2018; El Segundo, CA; United States
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  • 173
    Publication Date: 2019-07-13
    Description: In order to optimize systems, systems engineers require some sort of measure with which to compare vastly different system components. One such measure is system exergy, or the usable system work. Exergy balance analysis models provide a comparison of different system configurations, allowing systems engineers to compare different systems configuration options. This paper presents the exergy efficiency of several Mars transportation system configurations, using data on the interplanetary trajectory, engine performance, and vehicle mass. The importance of the starting and final parking orbits is addressed in the analysis, as well as intermediate hyperbolic escape and entry orbits within Earth and Mars' spheres of influence (SOIs). Propulsion systems analyzed include low-enriched uranium (LEU) nuclear thermal propulsion (NTP), high-enriched uranium (HEU) NTP, LEU methane (CH4) NTP, and liquid oxygen (LOX)/liquid hydrogen (LH2) chemical propulsion.
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-6553 , Annual Conference on Systems Engineering Research (CSER 2018); May 08, 2018 - May 09, 2018; Charlottesville, VA; United States
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  • 174
    Publication Date: 2019-07-13
    Description: The Near Earth Asteroid (NEA) [1] Scout is a deep space CubeSat designed to use an 86 m2 solar sail to navigate to a near earth asteroid called VG 1991. The solar sail deployment mechanism aboard NEA Scout has gone through numerous design cycles and ground tests since its conception in 2014. An engineering development unit (EDU) was constructed in the spring of 2016 and since then, the NEA Scout team has completed numerous ground deployments aiming to mature the deployment system and the ground test methods used to validate that system. Testing a large, non-rigid gossamer system in 1G environments has presented its difficulties to numerous solar sailing programs before, but NEA Scout's size, sail configuration, and budget has led the team to develop new deployment techniques and uncover new practices while improving their test methods. The program has planned and completed 5 separate full scale sail deployments to date, with a flight sail deployment test scheduled for FY18. The paper entitled "Design and Development of NEA Scout Solar Sail Deployer Mechanism" [2] was presented at the 43rd Aerospace Mechanisms Symposia. Since then, the system has matured and completed ascent vent, random vibration, boom deployment and sail deployment tests. This paper will discuss the lessons learned and advancements made while working on solar sail deployment testing and mechanical redesign cycles.
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-6541 , Aerospace Mechanisms Symposium; May 16, 2018 - May 18, 2018; Cleveland, OH; United States
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  • 175
    Publication Date: 2019-07-13
    Description: For the last 5 years, NASA Goddard has been investigating Distributed Spacecraft Missions (DSM) system architectures, surveying past, current and potential mission concepts, developing several taxonomies and identifying some key technologies that will enable future DSM mission design, development, operations and management. This paper summarizes this Initiative and the talk will provide details about specific Goddard DSM projects that are currently underway and that are relevant to future Earth Science missions.
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN59192 , International Geoscience and Remote Sensing Symposium (2018 IGARSS); Jul 22, 2018 - Jul 27, 2018; Valencia; Spain
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  • 176
    Publication Date: 2019-07-13
    Description: This paper presents an overview of the development and qualification test campaign for the primary structure of the European Service Module of ORION, the NASA spacecraft which will serve the future human exploration missions to the Moon, Mars and beyond. Under an agreement between NASA and ESA, the ORION will be powered by a European Service Module (ESM), providing also water and oxygen for astronauts' life sustainability. The development and qualification of the European Service Module (ESM) is under ESA responsibility with Airbus Defense and Space as the prime contractor. Thales Alenia Space Italia is responsible for design development, manufacturing, assembly and qualification of the Structure subsystem. The European Service Module, installed onto the launch adapter, shall support the crew module with its adapter and a launch abort system. It shall sustain: - A combination of global and local launch loads during lift off and ascent phases, - On orbit loads induced by engine firing for orbital transfers and attitude control. The ESM structure is based on a core made of Composite Fiber Reinforced Polymer (CFRP) sandwich panels complemented by aluminum alloy platforms, longerons and secondary structures. A development campaign has been implemented in order to define and validate composite parts' strength allowable values for design: coupon tests at material level, test at component level up to breadboards tests performed on main structural components (composite to metallic joints, and at panels' discontinuities). An incremental approach as defined in [1] has been followed. A qualification static test campaign at primary structure assembly level has been implemented in order to validate the design against static stiffness and ultimate strength as well as to correlate the structural Finite Element Model (FEM) used for sizing and confirm the margins of safety. The tests have been performed successfully by Thales Alenia Space Italia (TAS-I) on two flight representative structural models (STA1, STA2), in Turin facilities (Italy) between August 2015 and March 2017, with engineering support of technical representatives from Airbus, ESA, NASA and LMCO. The main development and qualification test activities and associated results are presented and discussed in the paper
    Keywords: Spacecraft Design, Testing and Performance
    Type: GRC-E-DAA-TN53178 , European Conference on Spacecraft Structures, Materials and Environmental Testing(ECSSMET); May 28, 2018 - Jun 01, 2018; Noordwijk; Netherlands
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  • 177
    Publication Date: 2019-07-13
    Description: As NASA looks towards human missions to Mars, an effort has started to advance the technology of a Mars in situ resource utilization (ISRU) Propellant Production Plant to a flight demonstration. This paper will present a design study of the Sabatier subsystem. The Sabatier subsystem receives carbon dioxide, CO2, and hydrogen, H2, and converts them to methane, CH4, and water, H2O. The subsystem includes the Sabatier reactor, condenser, thermal management, and a recycling system (if required). This design study will look at how the choice of reactor thermal management, number of reactors, and recycling system affect the performance of the overall Sabatier system. Different schemes from the literature involving single or cascading reactors will be investigated to see if any provide distinct advantages for a Mars propellant production plant.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ICES-2018-155 , KSC-E-DAA-TN57348 , International Conference on Environmental Systems; Jul 08, 2018 - Jul 12, 2018; Albuquerque, NM; United States
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  • 178
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: MSFC-E-DAA-TN58825 , AIAA Propulsion and Energy Forum; Jul 09, 2018 - Jul 11, 2018; Cincinatti, OH; United States
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  • 179
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN55613 , Aerospace Mechanisms Symposium; May 16, 2018 - May 18, 2018; Cleveland, OH; United States
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  • 180
    Publication Date: 2019-07-13
    Description: Propellant slosh was analyzed for both the oxidizer and the fuel for the Europa Clipper propulsion system. Slosh was examined for various fill fractions for cases where acceleration was on the order of magnitude of 10(exp -2) m/sq. s using the computational fluid dynamics software package STAR-CCM+ and at various fill fractions for cases where acceleration was on the order of magnitude of 10(exp -5) m/sq. s using Surface Evolver. Equivalent mechanical model parameters were derived from the CFD data using MATLAB for both the higher and the lower acceleration slosh cases. These parameters were plotted and can be used to interpolate mechanical model parameters at fill fractions not analyzed by CFD or Surface Evolver.
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN57194 , AIAA/SAE/ASEE Joint Propulsion Conference; Jul 09, 2017 - Jul 11, 2017; Cincinnati, OH; United States
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  • 181
    Publication Date: 2019-07-13
    Description: System engineering of launch vehicles and spacecraft is a challenging and complex undertaking. There are many diverse systems which must be integrated and balanced to produce an effective design. This involves a multiplicity of individual engineering relationships that are difficult to integrate and even more difficult to define in a best balance. Integration efforts involve many different approaches, from process management to mass balance. But these approaches either do not directly address the launch vehicle or spacecraft performance or require many adjustments to be made to discover a balance. The system integrating physics, derived from the fundamental physics of the system, is the key to identifying a fully integrated system performance measure. Launch vehicles and spacecraft are thermodynamic systems with performance defined by thermodynamic properties. Thus, thermodynamic exergy, which integrates all of the systems thermodynamic properties, provides the system integrating relationships. This provides a basis for determining the most efficient design from among many different configuration options and for guiding the design activities from an integrated system level. This paper explores the current physics relationships used in launch vehicle system design and demonstrates that thermodynamic exergy provides a more explicit and complete approach to system integration.
    Keywords: Spacecraft Design, Testing and Performance
    Type: M17-6439 , Journal of Spacecraft and Rockets (ISSN 0022-4650) (e-ISSN 1533-6794); 55; 2; 451-461
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  • 182
    Publication Date: 2019-07-13
    Description: Atmospheric probes have been successfully flown to planets and moons in the solar system to conduct in situ measurements. They include the Pioneer Venus multi-probes, the Galileo Jupiter probe, and Huygens probe. Probe mission concepts to five destinations, including Venus, Jupiter, Saturn, Uranus, and Neptune, have all utilized similar-shaped aeroshells and concept of operations, namely a 45-degree sphere cone shape with high density heatshield material and parachute system for extracting the descent vehicle from the aeroshell. Each concept designed its probe to meet specific mission requirements and to optimize mass, volume, and cost. At the 2017 International Planetary Probe Workshop (IPPW), NASA Headquarters postulated that a common aeroshell design could be used successfully for multiple destinations and missions. This "common probe" design could even be assembled with multiple copies, properly stored, and made available for future NASA missions, potentially realizing savings in cost and schedule and reducing the risk of losing technologies and skills difficult to sustain over decades. Thus the NASA Planetary Science Division funded a study to investigate whether a common probe design could meet most, if not all, mission needs to the five planetary destinations with extreme entry environments. The Common Probe study involved four NASA Centers and addressed these issues, including constraints and inefficiencies that occur in specifying a common design. Study methodology: First, a notional payload of instruments for each destination was defined based on priority measurements from the Planetary Science Decadal Survey. Steep and shallow entry flight path angles (EFPA) were defined for each planet based on qualification and operational g-load limits for current, state-of-the-art instruments. Interplanetary trajectories were then identified for a bounding range of EFPA. Next, 3-degrees-of-freedom simulations for entry trajectories were run using the entry state vectors from the interplanetary trajectories. Aeroheating correlations were used to generate stagnation point convective and radiative heat flux profiles for several aeroshell shapes and entry masses. High fidelity thermal response models for various Thermal Protection System (TPS) materials were used to size stagnation-point thicknesses, with margins based on previous studies. Backshell TPS masses were assumed based on scaled heat fluxes from the heatshield and also from previous mission concepts. Presentation: We will present an overview of the study scope, highlights of the trade studies and design driver analyses, and the final recommendations of a common probe design and assembly. We will also indicate limitations that the common probe design may have for the different destinations. Finally, recommended qualification approaches for missions will be presented.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN53719 , International Planetary Probe Workshop (IPPW-2018); Jun 11, 2018 - Jun 15, 2018; Boulder, CO; United States
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  • 183
    Publication Date: 2019-07-13
    Description: NASA's Orion exploration spacecraft will fly more demanding mission profiles than previous NASA human flight spacecraft. Missions currently under development are destined for cislunar space. The EM-1 mission will fly unmanned to a Distant Retrograde Orbit (DRO) around the Moon. EM-2 will fly astronauts on a mission to the lunar vicinity. To fly these missions, Orion requires powered flight guidance that is more sophisticated than the orbital guidance flown on Apollo and the Space Shuttle. Orion's powered flight guidance software contains five burn guidance options. These five options are integrated into an architecture based on a proven shuttle heritage design, with a simple closed-loop guidance strategy. The architecture provides modularity, simplicity, versatility, and adaptability to future, yet-to-be-defined, exploration mission profiles. This paper provides a summary of the executive guidance architecture and details the five burn options to support both the nominal and abort profiles for the EM-1 and EM-2 missions.
    Keywords: Spacecraft Design, Testing and Performance
    Type: AAS 18-084 , JSC-E-DAA-TN50474-1 , Annual AAS Guidance and Control Conference; Feb 02, 2018 - Feb 07, 2018; Breckenridge, CO; United States
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  • 184
    Publication Date: 2019-07-13
    Description: This poster provides an overview of the requirements, design, development and testing of the 3D (Three Dimensional) Woven TPS (Thermal Protection System) being developed under NASA's Heatshield for Extreme Entry Environment Technology (HEEET) project. Under this current program, NASA is working to develop a TPS capable of surviving entry into Saturn. A primary goal of the project is to build and test an Engineering Test Unit (ETU) to establish a Technical Readiness Level (TRL) of 6 for this technology by 2017.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN52838 , Outer Planet Advisory Group (OPAG) Spring Meeting; Feb 21, 2018 - Feb 22, 2018; Hampton, VA; United States
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  • 185
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN61922-2 , Space Simulation Conference; Nov 05, 2018 - Nov 08, 2018; Annapolis, MD; United States
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  • 186
    Publication Date: 2019-07-13
    Description: The aim of the Distributed Attitude Control and Maneuvering for Deep Space SmallSats project is to advance a multi-purpose, deep space mission-enabling technology for low-power attitude and thermal control of small satellites to a flight demonstration technology readiness level (TRL). The film-evaporation microelectromechanical systems tunable array (FEMTA) small satellite technology combines innovative microelectromechanical systems (MEMS) microfabrication and microscale effects in fluid surface tension to produce a thermally actuated capillary valve. Using water as the propellant, the FEMTA thruster can generate finely controllable thrust at a thrust to power ratio of about 200 microNewton per Watt (W).
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN55820 , FS #2018-03-07-ARC , Interplanetary Small Satellite Conference; May 07, 2018 - May 09, 2018; Pasadena, CA; United States
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  • 187
    Publication Date: 2019-07-13
    Description: The Starling series of demonstration missions will test technologies required to achieve affordable, distributed spacecraft ("swarm") missions that: are scalable to at least 100 spacecraft for applications that include synchronized multipoint measurements; involve closely coordinated ensembles of two or more spacecraft operating as a single unit for interferometric, synthetic aperture, or similar sensor architectures; or use autonomous or semi-autonomous operation of multiple spacecraft functioning as a unit to achieve science or other mission objectives with low-cost small spacecraft.Starling1 will focus on developing technologies that enable scalability and deep space application. The mission goals include the demonstration of a Mobile Ad-hoc NETwork (MANET) through an in-space communication experiment, vision based relative navigation through the Starling Formation-flying Optical eXperiment (StarFOX), and demonstration of autonomous spacecraft reconfiguration using technologies developed by the Distributed System Autonomy (DSA) project.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN59780 , Small Satellite Conference; Aug 04, 2018 - Aug 09, 2018; Logan, UT; United States
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  • 188
    Publication Date: 2019-07-13
    Description: Atmospheric probes have been successfully flown to planets and moons in the solar system to conduct in-situ measurements. They include the Pioneer Venus multi-probes, the Galileo Jupiter probe, and Huygens probe. Probe mission concepts to five destinations, including Venus, Jupiter, Saturn, Uranus, and Neptune, have all utilized similar-shaped aeroshells and concept of operations, namely a 45 deg sphere cone shape with high density heatshield material and parachute system for extracting the descent vehicle from the aeroshell. Each concept designed its probe to meet specific mission requirements and to optimize mass, volume, and cost. At the 2017 IPPW, NASA Headquarters postulated that a common aero-shell design could be used successfully for multiple destinations and missions. This "common probe" design could even be assembled with multiple copies, properly stored, and made available for future NASA missions, potentially realizing savings in cost and schedule and reducing the risk of losing technologies and skills difficult to sustain over decades. Thus the NASA Planetary Science Division funded a study to investigate whether a common probe design could meet most, if not all, mission needs to the five planetary destinations with extreme entry environments. The Common Probe study involved four NASA Centers and addressed these issues, including constraints and inefficiencies that occur in specifying a common design.Study methodology: First, a notional payload of instruments for each destination was defined based on priority measurements from the Planetary Science Decadal Survey. Steep and shallow entry flight path angles (EFPA) were defined for each planet based on qualification and operational g-load limits for current, state-of-the-art instruments. Interplanetary trajectories were then identified for a bounding range of EFPA. Next, 3-DoF simulations for entry trajectories were run using the entry state vectors from the interplanetary trajectories. Aeroheating correlations were used to generate stagnation point convective and radiative heat flux profiles for several aeroshell shapes and entry masses. High fidelity thermal response models for various TPS materials were used to size stagnation point thicknesses, with margins based on previous studies. Backshell TPS masses were assumed based on scaled heat fluxes from the heatshield and also from previous mission concepts.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN60861 , Outer Planets Assessment Group; Sep 11, 2018 - Sep 12, 2018; Pasadena, CA; United States
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  • 189
    Publication Date: 2019-07-13
    Description: This paper presents an overview of the Second European Service Module (ESM-2), the second in a series of European Service Modules produced as part of the Barter agreement between NASA and ESA for the Orion Program. The European Industrial consortium is led by the ESA prime contractor Airbus Defence and Space in Bremen. ESA and Airbus signed the ESM-2 contract on 16 February 2017, for this key element of the Orion Exploration Mission 2 (EM-2). EM-2 is the first crewed mission for Orion and will take astronauts farther into the solar system than humanity has ever travelled. EM-2 will also be a historic mission for Europe, as the ESM-2 will be the first European spacecraft to be part of a human transportation system carrying humans beyond low Earth orbit. ESM-2 is mainly a recurring production following ESM-1. Nevertheless, there are a number of important changes being implemented, for example, to incorporate upgrades to further enhance safety and reliability. The challenging delivery schedule for ESM-2 has driven the need to commence manufacturing prior to completion of the qualification on ESM-1. In addition, some requirement deviations and non-compliances approved for ESM-1 have resulted in modifications for ESM-2. In order to manage the competing constraints effectively, the ESM-2 Team has put in place a number of novel approaches to manage schedule, risk, and technical changes. Airbus has set up multi-functional teams according to an approach known as "Major Spacecraft Deliveries" consisting of quality assurance, engineering and procurement. The risk of starting manufacturing prior to qualification is managed through a special risk share agreement. This agreement necessitates rigorous risk reviews across the board for all manufacturing, assembly, integration and test milestones. The ESM-2 changes are managed by Configuration Management, but Airbus has also introduced the Technical Baseline Matrix to provide a transparent top-level overview of the changes from ESM-1 to ESM-2. The tool provides the basis for ESM-2 design and development needs, decisions, as well as the input for the Orion EM-2 Critical Design Review (CDR). The main technical evolutions, status of the production and the novel management approaches for ESM-2 are presented and discussed in the paper.
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN61230 , International Astronautical Congress; Oct 01, 2018 - Oct 05, 2018; Bremen; Germany
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  • 190
    Publication Date: 2019-07-13
    Description: Following a very successful year of manufacturing, assembly and testing in factories located around the globe, NASA and ESA are preparing to deliver the major Exploration Mission-1 (EM-1) Orion flight elements, including the Crew Module, ESA Service Module and Launch Abort System. This international effort to design and develop a deep space exploration capable human spacecraft is rapidly transitioning from the design, development and test phase to the early test flight and production phase. Two major flight tests, an Ascent Abort test and EM-1, Orion's first flight onboard NASA's new heavy lift Space Launch System, are planned for the near future. Further, Orion will play a crucial role in the ambitious new Deep Space Gateway human exploration Program. This paper gives a short overview of the system and subsystem configuration of the Orion spacecraft, including NASA and ESA contributions, a status of EM-1, AA-2 and EM-2 spacecraft production, and a look at Orion's role in the construction and operation of the Deep Space Gateway. The paper will also address the innovative international cooperation methods being employed to conduct Orion and Service Module integration.
    Keywords: Spacecraft Design, Testing and Performance
    Type: JSC-E-DAA-TN59421 , International Astronautical Congress; Oct 01, 2018 - Oct 05, 2018; Bremen; Germany
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  • 191
    Publication Date: 2019-07-13
    Description: A regeneratively-cooled nozzle for liquid rocket engine applications is a significant cost of the overall engine due to the complexities of manufacturing a large thin-walled structure that must operate in extreme temperature and pressure environments. The National Aeronautics and Space Administration (NASA) has been investigating and advancing methods for fabrication of liquid rocket engine channel wall nozzles to realize further cost and schedule improvements over traditional techniques. The methods being evaluated are targeting increased scale required for current NASA and commercial space programs. Several advanced rapid fabrication methods are being investigated for forming of the inner liner, producing the coolant channels, closeout of the coolant channels, and fabrication of the manifolds. NASA's Marshall Space Flight Center (MSFC) has completed process development and subscale hot-fire testing of a series of these advanced fabrication channel wall nozzle technologies to gather performance data in a relevant environment. The primary fabrication technique being discussed in this paper is Laser Wire Direct Closeout (LWDC). This process has been developed to significantly reduce the time required for closeouts of regeneratively-cooled slotted liners. It allows for channel closeout to be formed in place in addition to the structural jacket without the need for channel fillers or complex tooling. Additional technologies were also tested as part of this program including water jet milling and arc-based additive manufacturing deposition. Each nozzle included different fabrication features, materials, and methods to demonstrate durability in a hot-fire environment. The results of design, fabrication, and hot-fire testing are discussed in this paper.
    Keywords: Spacecraft Design, Testing and Performance
    Type: AIAA Paper 2018-4860 , M18-6804 , AIAA Propulsion and Energy Forum,; Jul 09, 2018 - Jul 11, 2018; Cincinnati, OH; United States
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  • 192
    Publication Date: 2019-07-13
    Description: In liquid propellant rocket engines, spark igniters are often used indirectly to light preburners, gas generators, and main chambers [1]. Attraction for spark igniters is strongly influenced by their ability for repeatable engine starts and high reliability. In the case of spark igniters, however, ignition is reliant upon an ignitable mixture passing near the spark tip very early in the engine start transient, prior to pressure quenching of the spark. While direct ignition of rocket engine combustion chambers is possible and has been successfully implemented in engines such as RL-10, the development time can be significant since ignition requires precise and repeatable control of the propellant mixture ratio within the very small volume and short duration of the spark plasma. Generally, the preferred method of implementing spark igniters within rocket engines - especially larger engines, is to design a smaller "augmented spark igniter" pre-chamber in which propellant injection and mixture ratio near the spark plasma can be controlled independent of the engine injector. The resultant combustion products within the small pre-chamber are directed into the larger engine chamber via a torch tube. An augmented spark igniter is advantageous because the output torch flame that is much larger and more energetic than a discrete train of small spark plasmas.
    Keywords: Spacecraft Design, Testing and Performance
    Type: M17-6460 , 2018 AIAA Propulsion and Energy Forum and Exposition; Jul 09, 2018 - Jul 11, 2018; Cincinnati, OH; United States
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  • 193
    Publication Date: 2019-07-13
    Description: The James Webb Space Telescope Primary Mirror Segment Assemblies (PMSAs) and Secondary Mirror Assembly (SMA) were cleaned at the Johnson Space Center (JSC) in January 2018. In order to quantify the effectiveness of the cleaning, the same cleaning process was performed on the PMSA and SMA traveling witness wafers. These wafers have accompanied their respective mirror segments from their arrival at the Goddard Space Flight Center, through transport to JSC, and ultimately their exposure in Chamber A for cryogenic testing. The traveling wafers were analyzed using an Image Analysis automated microscope both prior to and after the cleaning. The resulting data showed that the PMSA wafers' Percent Area Coverage (PAC) reduced by 83.5% on average, from 0.1524 PAC to 0.0251 PAC. The SMA wafer's PAC decreased by 97.2%, from 0.1194 PAC to 0.0034 PAC. Further analysis of the particle size bins was completed in order to calculate their particle distribution slopes. The slope of the PMSA wafers increased by 0.025 on average, and the SMA wafer slope increased by 0.066. This indicates that the ratio of large to small particles slightly increased after the cleaning across all mirror segments. Visual inspections of the wafers and the flight PMSAs and SMA showed considerable and comparable particulate coverage improvements, thus leading to the conclusion that the average PAC on the PMSAs and SMA improved by the same factor as their respective wafers.
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN58353 , SPIE Optics+Photonics Conference; Aug 19, 2018 - Aug 23, 2018; San Diego, CA; United States
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  • 194
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: AIAA Paper 2018-4860 , M18-6827 , AIAA Propulsion and Energy Forum; Jul 09, 2018 - Jul 11, 2018; Cincinnati, OH; United States
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  • 195
    Publication Date: 2019-07-13
    Description: Presently, most CubeSat components and buses are generally not appropriate for missions where significant or indeterminate risk of failure is unacceptable. This has precluded their use in many cases where their attributes could otherwise enable or enhance mission objectives. However, in the future, CubeSats and SmallSats, which deviate from CubeSat form factors but often incorporate CubeSat components and subsystems, will address challenges that many presently consider to be beyond the platform's capabilities. This growing potential utility, combined with the limited volume of successful CubeSat flight heritage, is driving an interagency effort to improve small satellite mission confidence.
    Keywords: Spacecraft Design, Testing and Performance
    Type: GSFC-E-DAA-TN58616 , AIAA Small Satellite Conference; Aug 04, 2018 - Aug 09, 2018; Logan, UT; United States
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  • 196
    Publication Date: 2019-07-13
    Description: Recent introduction of Coaxial Thermocouple type calorimeters into the NASA Ames arc jet facilities has inspired an analysis of 2D conduction effects internal to this type of calorimeter. The 1D finite slab inverse analysis (which is typically used to deduce the heat transfer to the calorimeter) relies on the assumption that lateral conduction (i.e., 2D effects) is negligible. Most calorimeter bodies have a spherical nose, which in itself is a violation of the 1D finite slab analysis assumption. Secondly most calorimeters experience a variation in heating across the face of the body which is also a violation of the 1D finite slab analysis assumption. It turns out that these two effects tend to cancel each other to some extent. This paper shows the extent to which error exists in the analysis of the Coaxial Thermocouple type calorimeters, and also offers analysis strategies for reducing the errors.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN58319 , AIAA Aviation Forum; Jun 25, 2018 - Jun 29, 2018; Atlanta, GA; United States
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  • 197
    Publication Date: 2019-07-13
    Description: The Mars2020 entry vehicle is currently being developed by NASA to safely land its next rover on the Martian surface in 2021. The vehicle will be protected from entry aeroheating using three different TPS materials: PICA tiles on the forebody, SLA-561V on the backshell and Acusil-II on the parachute close-out cone (PCC) and its backshell interface plate (BIP). Mars2020's entry vehicle and TPS design is identical to the Mars Science Laboratory, NASA's last Mars lander; therefore, the purpose of this study is to assess the adequacy of the existing TPS design and thickness for Mars2020 predicted environments. This study focuses on sizing and margin assessment of Acusil-II TPS on the PCC and BIP. The methodology and analysis techniques that were used for assessing thermal margins are reviewed. Analysis assumptions and limitations are discussed in detail. Thermal sizing is performed at different locations and results are presented.
    Keywords: Spacecraft Design, Testing and Performance
    Type: ARC-E-DAA-TN58297 , AIAA Aviation Forum; Jun 25, 2018 - Jun 29, 2018; Atlanta, GA; United States
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  • 198
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-6809 , National Space & Missile Materials Symposium (NSMMS); Jun 25, 2018 - Jun 28, 2018; Madison, WI; United States
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  • 199
    Publication Date: 2019-07-13
    Description: Lynx is an X-Ray telescope large-mission concept for consideration in NASA's 2020 Astrophysics Decadal Survey. A conceptual structural design is evolving that leverages the success and lessons learned from Chandra and that takes into account unique needs of Lynx. Space optics systems require extreme stability. Any motion in-service (thermal effects, structural dynamics, etc.) impacts performance. An initial analysis was performed to predict the first-cut dynamic responses, jitter, at two selected points on the Lynx observatory. One point is on the Lynx X-ray Mirror Assembly (LMA) and the other, on the focal plane Integrated Science Instrument Module (ISIM). Relative motion between these two points was predicted along with vibration spectra. This information will be used in upcoming analyses of the LMA and the ISIM.
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-6781 , SPIE Astronomical Telescopes + Instruments; Jun 10, 2018 - Jun 15, 2018; Austin, TX; United States
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  • 200
    facet.materialart.
    Unknown
    In:  Other Sources
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: M18-6712 , Osher Lifelong Learning Institute Outreach Presentation; May 09, 2018; Huntsville, AL; United States
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