ALBERT

Description: No abstract available

Description: No abstract available

Description: Eight "Expedite the Processing of Experiments to Space Station" (EXPRESS) Rack facilities are located within the International Space Station (ISS) laboratories to provide standard resources and interfaces for the simultaneous and independent operation of multiple experiments within each rack. Each EXPRESS Rack provides eight Middeck Locker Equivalent locations and two drawer locations for powered experiment equipment, also referred to as sub-rack payloads. Payload developers may provide their own structure to occupy the equivalent volume of one, two, or four lockers as a single unit. Resources provided for each location include power (28 Vdc, 0-500 W), command and data handling (Ethernet, RS-422, 5 Vdc discrete, +/- 5 Vdc analog), video (NTSC/RS 170A), and air cooling (0-200 W). Each rack also provides water cooling for two locations (500W ea.), one vacuum exhaust interface, and one gaseous nitrogen interface. Standard interfacing cables and hoses are provided on-orbit. One laptop computer is provided with each rack to control the rack and to accommodate payload application software. Four of the racks are equipped with the Active Rack Isolation System to reduce vibration between the ISS and the rack. EXPRESS Racks are operated by the Payload Operations Integration Center at Marshall Space Flight Center and the sub-rack experiments are operated remotely by the investigating organization. Payload Integration Managers serve as a focal to assist organizations developing payloads for an EXPRESS Rack. NASA provides EXPRESS Rack simulator software for payload developers to checkout payload command and data handling at the development site before integrating the payload with the EXPRESS Functional Checkout Unit for an end-to-end test before flight. EXPRESS Racks began supporting investigations onboard ISS on April 24, 2001 and will continue through the life of the ISS.

Description: The NASA Space Launch System (SLS) vehicle is composed of four RS-25 liquid oxygen- hydrogen rocket engines in the core-stage and two 5-segment solid rocket boosters and as a result six hot supersonic plumes interact within the aft section of the vehicle during ight. Due to the complex nature of rocket plume-induced ows within the launch vehicle base during ascent and a new vehicle con guration, sub-scale wind tunnel testing is required to reduce SLS base convective environment uncertainty and design risk levels. This hot- re test program was conducted at the CUBRC Large Energy National Shock (LENS) II short-duration test facility to simulate ight from altitudes of 50 kft to 210 kft. The test program is a challenging and innovative e ort that has not been attempted in 40+ years for a NASA vehicle. This presentation discusses the various trends of base convective heat ux and pressure as a function of altitude at various locations within the core-stage and booster base regions of the two-percent SLS wind tunnel model. In-depth understanding of the base ow physics is presented using the test data, infrared high-speed imaging and theory. The normalized test design environments are compared to various NASA semi- empirical numerical models to determine exceedance and conservatism of the ight scaled test-derived base design environments. Brief discussion of thermal impact to the launch vehicle base components is also presented.

Description: Post-test examination and data analysis that followed a two week long vacuum test showed that numerous self-stick thermocouples became detached from the test article. The thermocouples were reattached with thermally conductive epoxy and the test was repeated to obtain the required data. Because the thermocouple detachment resulted in significant expense and rework, it was decided to investigate the temporary attachment methods used around NASA and to perform a test to assess their efficacy. The present work describes the original test and the analysis that showed that the thermocouples had become detached, temporary thermocouple attachment methods assessed in the retest and in the thermocouple attachment test, and makes a recommendation for attachment methods for future tests.

Description: New Horizons flew by Pluto and its moons on July 14, 2015 [1]. In the days prior to the closest approach (C/A), panchromatic and color observations of Pluto and Charon were made covering a fully complete range of longitudes. Although only a fraction of this "late-approach" data series has been transmitted to the ground, the results indicate Pluto's latitudinal coloring trends seen on the encounter hemisphere continues on the far side. Charon's red pole is visible from a multitude of longitudes and its colors are uniform with longitude at lower latitudes.

Description: The NASA Engineering and Safety Center (NESC) has sponsored a Pathfinder Study to investigate how Model Based Systems Engineering (MBSE) and Model Based Engineering (MBE) techniques can be applied by NASA spacecraft development projects. The objectives of this Pathfinder Study included analyzing both the products of the modeling activity, as well as the process and tool chain through which the spacecraft design activities are executed. Several aspects of MBSE methodology and process were explored. Adoption and consistent use of the MBSE methodology within an existing development environment can be difficult. The Pathfinder Team evaluated the possibility that an "MBSE Template" could be developed as both a teaching tool as well as a baseline from which future NASA projects could leverage. Elements of this template include spacecraft system component libraries, data dictionaries and ontology specifications, as well as software services that do work on the models themselves. The Pathfinder Study also evaluated the tool chain aspects of development. Two chains were considered: 1. The Development tool chain, through which SysML model development was performed and controlled, and 2. The Analysis tool chain, through which both static and dynamic system analysis is performed. Of particular interest was the ability to exchange data between SysML and other engineering tools such as CAD and Dynamic Simulation tools. For this study, the team selected a Mars Lander vehicle as the element to be designed. The paper will discuss what system models were developed, how data was captured and exchanged, and what analyses were conducted.

Description: Orbital debris in the millimeter size range can pose a hazard to current and planned spacecraft due to the high relative impact speeds in Earth orbit. Fortunately, orbital debris has a relatively short life at lower altitudes due to atmospheric effects; however, at higher altitudes orbital debris can survive much longer and has resulted in a band of high flux around 700 to 1,500 km above the surface of the Earth. While large orbital debris objects are tracked via ground based observation, little information can be gathered about small particles except by returned surfaces, which until the Orion Exploration Flight Test number one (EFT-1), has only been possible for lower altitudes (400 to 500 km). The EFT-1 crew module backshell, which used a porous, ceramic tile system with surface coatings, has been inspected post-flight for potential micrometeoroid and orbital debris (MMOD) damage. This paper describes the pre- and post-flight activities of inspection, identification and analysis of six candidate MMOD impact craters from the EFT-1 mission.

Description: The paper presents further development in normalized contrast processing of flash infrared thermography method by the author given in US 8,577,120 B1. The method of computing normalized image or pixel intensity contrast, and normalized temperature contrast are provided, including converting one from the other. Methods of assessing emissivity of the object, afterglow heat flux, reflection temperature change and temperature video imaging during flash thermography are provided. Temperature imaging and normalized temperature contrast imaging provide certain advantages over pixel intensity normalized contrast processing by reducing effect of reflected energy in images and measurements, providing better quantitative data. The subject matter for this paper mostly comes from US 9,066,028 B1 by the author. Examples of normalized image processing video images and normalized temperature processing video images are provided. Examples of surface temperature video images, surface temperature rise video images and simple contrast video images area also provided. Temperature video imaging in flash infrared thermography allows better comparison with flash thermography simulation using commercial software which provides temperature video as the output. Temperature imaging also allows easy comparison of surface temperature change to camera temperature sensitivity or noise equivalent temperature difference (NETD) to assess probability of detecting (POD) anomalies.

Description: How crews get into or out of their ascent vehicle has profound implications for Mars surface architecture. Extravehicular Activity (EVA) hatches and Airlocks have the benefit of relatively low mass and high Technology Readiness Level (TRL), but waste consumables with a volume depressurization for every ingress/egress. Perhaps the biggest drawback to EVA hatches or Airlocks is that they make it difficult to keep Martian dust from being tracked back into the ascent vehicle, in violation of planetary protection protocols. Suit ports offer the promise of dust mitigation by keeping dusty suits outside the cabin, but require significant cabin real estate, are relatively high mass, and current operational concepts still require an EVA hatch to get the suits outside for the first EVA, and back inside after the final EVA. This is primarily because current designs don't provide enough structural support to protect the suits from ascent/descent loads or potential thruster plume impingement. For architectures involving more than one surface element-such as an ascent vehicle and a rover or surface habitat-a retractable tunnel is an attractive option. By pushing spacesuit don/doff and EVA operations to an element that remains on the surface, ascended vehicle mass and dust can be minimized. What's more, retractable tunnels provide operational flexibility by allowing surface assets to be re-configured or built up over time. Retractable tunnel functional requirements and design concepts being developed as part of the National Aeronautics and Space Administration's (NASA) Evolvable Mars Campaign (EMC) work will add a new ingress/egress option to the surface architecture trade space.

Description: Space debris poses a major risk to spacecraft. In low earth orbit, impact velocities can be 10 - 11 km/s and as high as 15 km/s. For debris shield design, it would be desirable to be able to launch projectiles of known shape and mass to these velocities. The design of the proposed 10 - 11 km/sec gun uses, as a starting point, the Ames 1.28/0.22 two stage gun, which has achieved muzzle velocities of 10 - 11.3 km/sec. That gun is scaled up to a 0.3125 launch tube diameter. The gun is then optimized with respect to maximum pressures by varying the pump tube length to diameter ratio (L/D), the piston mass and the hydrogen pressure. A pump tube L/D of 36.4 is selected giving the best overall performance. Piezometric ratios for the optimized guns are found to be ~2.3, much more favorable than for more traditional two stage light gas guns, which range from 4 to 6. (The piezometric ratio for a gun is defined as the maximum projectile base pressure divided by the constant projectile base pressure which, acting over the entire barrel length, would produce the same muzzle velocity.) The maximum powder chamber pressures are 20 to 30 ksi. To reduce maximum pressures, the desirable range of the included angle of the cone of the high pressure coupling is found to be 7.3 to 14.6 degrees. Lowering the break valve rupture pressure is found to lower the maximum projectile base pressure, but to raise the maximum gun pressure. For the optimized gun with a pump tube L/D of 36.4, increasing the muzzle velocity by decreasing the projectile mass and increasing the powder loads is studied. It appears that saboted spheres could be launched to 10.25 and possibly as high as 10.8 km/sec, and that disc-like plastic models could be launched to 11.05 km/s. The use of a tantalum liner to greatly reduce bore erosion and increase muzzle velocity is discussed. With a tantalum liner, CFD code calculations predict muzzle velocities as high as 12 to 13 km/s.

Description: The discovery of thousands of exoplanets is generating increasing interest in the direct imaging and characterization of these planets. Starshade, and eternal occulter, could fly in formation between a telescope and distant star, blocking out the light from the star, and enabling us to focus on the light of any orbiting planets. Recent technology developments in coordination with system design, has added much needed detail to define the technology requirements for a science mission that could launch in the 2020's. This paper address the mechanical architecture, the successful efforts to date, the current state of design for the mechanical system, and upcoming technology efforts.

Description: No abstract available

Description: NASA STMD Centennial Challenges Program operates government prize programs for the public benefit. Cube Quest Challenge awards prizes to citizen inventors who advance CubeSat state of the art, enabling affordable NASA science and exploration missions. Cube Quest will take place in lunar orbit or at 4M km. CubeSat developers will make advancements in communications, propulsion and radiation tolerance suitable for future deep space missions. Cube Quest may inspire other ambitious government challenges.

Description: In-space assembly (ISA), the ability to build structures in space, has the potential to enable or support a wide range of advanced mission capabilities. Many different individual assembly technologies would be needed in different combinations to serve many mission concepts. The many-to-many relationship between mission needs and technologies makes it difficult to determine exactly which specific technologies should receive priority for development and demonstration. Furthermore, because enabling technologies are still immature, no realistic, near-term design reference mission has been described that would form the basis for flowing down requirements for such development and demonstration. This broad applicability without a single, well-articulated mission makes it difficult to advance the technology all the way to flight readiness. This paper reports on a study that prioritized individual technologies across a broad field of possible missions to determine priority for future technology investment.

Description: No abstract available

Description: MODIS-DT Collection 6 - Aqua/Terra level 2, 3; entire record processed - "Trending" issues reduced - Still a 15% or 0.02 Terra vs Aqua offset. - Terra/Aqua convergence improved with C6+, but bias remains. - Other calibration efforts yield mixed results. VIIRS-DT in development - VIIRS is similar, yet different then MODIS - With 50% wider swath, VIIRS has daily coverage - Ensures algorithm consistency with MODIS. - Currently: 20% NPP vs Aqua offset over ocean. - Only small bias (%) over land (2012-2016) - Can VIIRS/MODIS create aerosol CDR? Calibration for MODIS - VIIRS continues to fundamentally important. It's not just Terra, or just Aqua, or just NPP-VIIRS, I really want to push synergistic calibration.

Description: The goal of interface management is to identify, define, control, and verify interfaces; ensure compatibility; provide an efficient system development; be on time and within budget; while meeting stakeholder requirements. This paper will present a successful seven-step approach to interface management used in several NASA flight projects. The seven-step approach using Model Based Systems Engineering will be illustrated by interface examples from the Materials International Space Station Experiment-X (MISSE-X) project. The MISSE-X was being developed as an International Space Station (ISS) external platform for space environmental studies, designed to advance the technology readiness of materials and devices critical for future space exploration. Emphasis will be given to best practices covering key areas such as interface definition, writing good interface requirements, utilizing interface working groups, developing and controlling interface documents, handling interface agreements, the use of shadow documents, the importance of interface requirement ownership, interface verification, and product transition.

Description: To increase the number of single point in-situ measurements of thermosphere and exosphere ion and neutral composition and density, miniaturized instrumentation is in high demand to take advantage of the increasing platform opportunities available in the smallsat/cubesat industry. The INMS (Ion-Neutral Mass Spectrometer) addresses this need by providing simultaneous measurements of both the neutral and ion environment, essentially providing two instruments in one compact model. The 1.3U volume, 570 gram, 1.8W nominal power INMS instrument makes implementation into cubesat designs (3U and above) practical and feasible. With high dynamic range (0.1-500eV), mass dynamic range of 1-40amu, sharp time resolution (0.1s), and mass resolution of MdM16, the INMS instrument addresses the atmospheric science needs that otherwise would have required larger more expensive instrumentation. INMS-v1 (version 1) launched on Exocube (CalPoly 3U cubesat) in 2015 and INMS-v2 (version 2) is scheduled to launch on Dellingr (GSFC 6U cubesat) in 2017. New versions of INMS are currently being developed to increase and add measurement capabilities, while maintaining its smallsat/cubesat form.

Description: This work describes the direct simulation Monte Carlo (DSMC) investigation of Saturn entry probe scenarios and the influence of non-equilibrium phenomena on Saturn entry conditions. The DSMC simulations coincide with rarefied hypersonic shock tube experiments of a hydrogen-helium mixture performed in the Electric Arc Shock Tube (EAST) at the NASA Ames Research Center. The DSMC simulations are post-processed through the NEQAIR line-by-line radiation code to compare directly to the experimental results. Improved collision cross-sections, inelastic collision parameters, and reaction rates are determined for a high temperature DSMC simulation of a 7-species H2-He mixture and an electronic excitation model is implemented in the DSMC code. Simulation results for 27.8 and 27.4 km/s shock waves are obtained at 0.2 and 0.1 Torr, respectively, and compared to measured spectra in the VUV, UV, visible, and IR ranges. These results confirm the persistence of non-equilibrium for several centimeters behind the shock and the diffusion of atomic hydrogen upstream of the shock wave. Although the magnitude of the radiance did not match experiments and an ionization inductance period was not observed in the simulations, the discrepancies indicated where improvements are needed in the DSMC and NEQAIR models.

Description: A sweeping jet (SWJ) actuator operating over a range of nozzle pressure ratios (NPRs) was characterized with femtosecond laser electronic excitation tagging (FLEET), single hot-wire anemometry (HWA) and high-speed/phase-averaged schlieren. FLEET velocimetry was successfully demonstrated in a highly unsteady, oscillatory flow containing subsonic through supersonic velocities. Qualitative comparisons between FLEET and HWA (which measured mass flux since the flow was compressible) showed relatively good agreement in the external flow profiles. The spreading rate was found to vary from 0.5 to 1.2 depending on the pressure ratio. The precision of FLEET velocity measurements in the external flow field was poorer (is approximately equal to 25 m/s) than reported in a previous study due to the use of relatively low laser fluences, impacting the velocity fluctuation measurements. FLEET enabled velocity measurements inside the device and showed that choking likely occurred for NPR 2.0, and no internal shockwaves were present. Qualitative oxygen concentration measurements using FLEET were explored in an effort to gauge the jet's mixing with the ambient. The jet was shown to mix well within roughly four throat diameters and mix fully within roughly eight throat diameters. Schlieren provided visualization of the internal and external flow fields and showed that the qualitative structure of the internal flow does not vary with pressure ratio and the sweeping mechanism observed for incompressible NPRs also probably holds for compressible NPRs.

Description: At the end of James Webb Space Telescope (JWST) OTIS (Optical Telescope Element-OTE-Integrated Science Instrument Module-ISIM) cryogenic vacuum testing in NASA Johnson Space Centers (JSCs) thermal vacuum (TV) Chamber A, contamination control (CC) engineers are mooting the idea that chamber particulate material stirred up by the repressurization process may be kept from falling into the ISIM interior to some degree by activating instrument purge flows over some initial period before opening the chamber valves. This memo describes development of a series of models designed to describe this process. These are strung together in tandem to estimate overpressure evolution from which net outflow velocity behavior may be obtained. Creeping flow assumptions are then used to determine the maximum particle size that may be kept suspended above the ISIM aperture, keeping smaller particles from settling within the instrument module.

Description: Program to Optimize Simulated Trajectories II (POST2) was utilized to develop trajectory simulations characterizing all flight phases from drop to splashdown for the Low-Density Supersonic Decelerator (LDSD) project's first and second Supersonic Flight Dynamics Tests (SFDT-1 and SFDT-2) which took place June 28, 2014 and June 8, 2015, respectively. This paper describes the modeling improvements incorporated into the LDSD POST2 simulations since SFDT-1 and presents how these modeling updates affected the predicted SFDT-2 performance and sensitivity to the mission design. The POST2 simulation flight dynamics support during the SFDT-2 launch, operations, and recovery is also provided.

Description: The experimental investigation of a broadband far-infrared meta-material absorber is described. The observed absorptance is greater than 0.95 from 1 to 20 terahertz (300-15 microns) over a temperature range spanning 5-300 degrees Kelvin. The meta-material, realized from an array of tapers approximately 100 microns in length, is largely insensitive to the detailed geometry of these elements and is cryogenically compatible with silicon-based micro-machined technologies. The electromagnetic response is in general agreement with a physically motivated transmission line model.

Description: The Tethered Satellite System (TSS) Space Shuttle missions, TSS-1 in 1993 and TSS-1R in 1996, were the height of space tether technology development in the U.S. Altogether, the investment made by NASA and the Italian Space Agency (ASI) over the thirteen-year period of the TSS Program totaled approximately $400M-exclusive of the two Space Shuttle flights provided by NASA. Since those two pioneering missions, there have been several smaller tether flight experiments, but interest in this promising technology has waned within NASA as well as the DOD agencies. This is curious in view of the unique capabilities of space tether systems and the fact that they have been flight validated in earth orbit and shown to perform better than the preflight dynamic or electrodynamic theoretical predictions. While it is true that the TSS-1 and TSS-1R missions experienced technical difficulties, the causes of these early developmental problems are now known to have been engineering design flaws, material selection, and procedural issues that (1) are unrelated to the basic viability of space tether technology, and (2) can be readily corrected. The purpose of this paper is to review the dynamic and electrodynamic fundamentals of space tethers and the unique capabilities they afford (that are enabling to certain types of space missions); to elucidate the nature, cause, and solution of the early developmental problems; and to provide an update on progress made in development of the technology.

Description: No abstract available

Description: Human-scale landers require the delivery of much heavier payloads to the surface of Mars than is possible with entry, descent, and landing (EDL) approaches used to date. A conceptual design was developed for a 10 m diameter crewed Mars lander with an entry mass of approx.75 t that could deliver approx.28 t of useful landed mass (ULM) to a zero Mars areoid, or lower, elevation. The EDL design centers upon use of a high ballistic coefficient blunt-body entry vehicle and throttled supersonic retro-propulsion (SRP). The design concept includes a 26 t Mars Ascent Vehicle (MAV) that could support a crew of 2 for approx.24 days, a crew of 3 for approx.16 days, or a crew of 4 for approx.12 days. The MAV concept is for a fully-fueled single-stage vehicle that utilizes a single pump-fed 250 kN engine using Mono-Methyl Hydrazine (MMH) and Mixed Oxides of Nitrogen (MON-25) propellants that would deliver the crew to a low Mars orbit (LMO) at the end of the surface mission. The MAV concept could potentially provide abort-to-orbit capability during much of the EDL profile in response to fault conditions and could accommodate return to orbit for cases where the MAV had no access to other Mars surface infrastructure. The design concept for the descent stage utilizes six 250 kN MMH/MON-25 engines that would have very high commonality with the MAV engine. Analysis indicates that the MAV would require approx.20 t of propellant (including residuals) and the descent stage would require approx.21 t of propellant. The addition of a 12 m diameter supersonic inflatable aerodynamic decelerator (SIAD), based on a proven flight design, was studied as an optional method to improve the ULM fraction, reducing the required descent propellant by approx.4 t.

Description: The Light Microscopy Module (LMM) is a microscope facility developed at Glenn Research Center (GRC) that provides researchers with powerful imaging capability onboard the International Space Station (ISS). LMM has the ability to have its hardware recongured on-orbit to accommodate a wide variety of investigations, with the capability of remotely acquiring and downloading digital images across multiple levels of magnication.

Description: The GOES-R magnetometer accuracy requirement is 1.7 nanoteslas (nT). During quiet times (100 nT), accuracy is defined as absolute mean plus 3 sigma. During storms (300 nT), accuracy is defined as absolute mean plus 2 sigma. To achieve this, the sensor itself has better than 1 nT accuracy. Because zero offset and scale factor drift over time, it is also necessary to perform annual calibration maneuvers. To predict performance, we used covariance analysis and attempted to corroborate it with simulations. Although not perfect, the two generally agree and show the expected behaviors. With the annual calibration regimen, these predictions suggest that the magnetometers will meet their accuracy requirements.

Description: The heritage thermal model for the full STO-2 (Stratospheric Terahertz Observatory II), vehicle has been updated to model the CSBF (Columbia Scientific Balloon Facility) SIP-14 (Scientific Instrument Package) in detail. Analysis of this model has been performed for the Antarctica FY2017 launch season. Model temperature predictions are compared to previous results from STO-2 review documents.

Description: The Columbia Scientific Balloon Facility is responsible for ensuring that science payloads meet the appropriate design requirements. The ultimate goal is to ensure that payloads stay within the allowable launch limits as well as survive the termination event. The purpose of this presentation is to provide some general guidelines for Gondola Design. These include rules and reasons on why CSBF has a certain preference and location for certain components within the gondola as well as other suggestions. Additionally, some recommendations are given on how to avoid common pitfalls.

Description: Missions that involve traversing through a planetary atmosphere are unique opportunities that require elements of entry, descent, and landing (EDL). Many aspects of the EDL sequence are qualified using analysis and simulation due to the inability to conduct appropriate ground tests, however validating flight data are often lacking, especially for missions not involving Earth re-entry. NASA has made strategic decisions to collect EDL flight data in order to improve future mission designs. For example, MEDLI1 and EFT-1 gathered hypersonic pressure and in-depth temperature data in the thermal protection system (TPS). However, the ability to collect EDL flight data from the smaller competed missions, such as Discovery and New Frontiers, has been limited in part due to the Principal Investigator-managed cost-caps (PIMCC). The recent NASA decision to consider EDL instrumentation earlier in the mission design cycle led to the inclusion of a requirement in the Discovery 2014 Announcement of Opportunity which requires all missions that involve EDL to include an Engineering Science Investigation (ESI).2 The ESI would involve sensors for aerothermal environment and TPS; atmosphere, aerodynamics, and flight dynamics; atmospheric decelerator; and/or vehicle structure.3 The ESI activity would be funded outside of the PIMCC.

Description: No abstract available

Description: No abstract available

Description: The In-Space Manufacturing (ISM) project at NASA Marshall Space Flight Center currently operates a 3D FDM (fused deposition modeling) printer onboard the International Space Station. In order to enable utilization of this capability by designer, the project needs to establish characteristic material properties for materials produced using the process. This is difficult for additive manufacturing since standards and specifications do not yet exist for these technologies. Due to availability of crew time, there are limitations to the sample size which in turn limits the application of the traditional design allowables approaches to develop a materials property database for designers. In this study, various approaches to development of material databases were evaluated for use by designers of space systems who wish to leverage in-space manufacturing capabilities. This study focuses on alternative statistical techniques for baseline property development to support in-space manufacturing.

Description: No abstract available

Description: Although spacecraft developers have been moving towards standardized product lines as the aerospace industry has matured, NASA's continual need to push the cutting edge of science to accomplish unique, challenging missions can still lead to spacecraft resource growth over time. This paper assesses historical mass, power, cost, and schedule growth for multiple NASA spacecraft from the last twenty years and compares to industry reserve guidelines to understand where the guidelines may fall short. Growth is assessed from project start to launch, from the time of the preliminary design review (PDR) to launch and from the time of the critical design review (CDR) to launch. Data is also assessed not just at the spacecraft bus level, but also at the subsystem level wherever possible, to help obtain further insight into possible drivers of growth. Potential recommendations to minimize spacecraft mass, power, cost, and schedule growth for future missions are also discussed.

Description: The hypersonic regime of planetary entry combines the most severe environments that an entry vehicle will encounter with the greatest amount of uncertainty as to the events unfolding during that time period. This combination generally leads to conservatism in the design of an entry vehicle, specifically that of the thermal protection system (TPS). Each planetary entry provides a valuable aerodynamic and aerothermal testing opportunity; the utilization of this opportunity is paramount in better understanding how a specific entry vehicle responds to the demands of the hypersonic entry environment. Previous efforts have been made to instrument entry vehicles in order to collect data during the entry period and reconstruct the corresponding vehicle response. The purpose of this paper is to cumulatively document past TPS instrumentation designs for applicable planetary missions, as well as to list pertinent results and any explainable shortcomings.

Description: Over a decade of work has been conducted in the development of NASA's Hypersonic Inflatable Aerodynamic Decelerator (HIAD) deployable aeroshell technology. This effort has included multiple ground test campaigns and flight tests culminating in the HIAD project's second generation (Gen-2) aeroshell system. The HIAD project team has developed, fabricated, and tested stacked-torus inflatable structures (IS) with flexible thermal protection systems (F-TPS) ranging in diameters from 3-6m, with cone angles of 60 and 70 deg. To meet NASA and commercial near term objectives, the HIAD team must scale the current technology up to 12-15m in diameter. The HIAD project's experience in scaling the technology has reached a critical juncture. Growing from a 6m to a 15m class system will introduce many new structural and logistical challenges to an already complicated manufacturing process. Although the general architecture and key aspects of the HIAD design scale well to larger vehicles, details of the technology will need to be reevaluated and possibly redesigned for use in a 15m-class HIAD system. These include: layout and size of the structural webbing that transfers load throughout the IS, inflatable gas barrier design, torus diameter and braid construction, internal pressure and inflation line routing, adhesives used for coating and bonding, and F-TPS gore design and seam fabrication. The logistics of fabricating and testing the IS and the F-TPS also become more challenging with increased scale. Compared to the 6m aeroshell (the largest HIAD built to date), a 12m aeroshell has four times the cross-sectional area, and a 15m one has over six times the area. This means that fabrication and test procedures will need to be reexamined to account for the sheer size and weight of the aeroshell components. This will affect a variety of steps in the manufacturing process, such as: stacking the tori during assembly, stitching the structural webbing, initial inflation of tori, and stitching of F-TPS gores. Additionally, new approaches and hardware will be required for handling and ground testing of both individual tori and the fully assembled HIADs. There are also noteworthy benefits of scaling up the HIAD aeroshell to 15m-class system. Two complications in working with handmade textiles structures are the non-linearity of the materials and the role of human accuracy during fabrication. Larger, more capable, HIAD structures should see much larger operational loads, potentially bringing the structural response of the materials out of the non-linear regime and into the preferred linear response range. Also, making the reasonable assumption that the magnitude of fabrication accuracy remains constant as the structures grow, the relative effect of fabrication errors should decrease as a percentage of the textile component size. Combined, these two effects improve the predictive capability and the uniformity of the structural response for a 12-15m class HIAD. In this paper, the challenges and associated mitigation plans related to scaling up the HIAD stacked-torus aeroshell to a 15m class system will be discussed. In addition, the benefits of enlarging the structure will be further explored.

Description: The Deep Space Climate Observatory (DSCOVR), formerly known as Triana, successfully launched on February 11th, 2015. To date, each of the five space-craft attitude control system (ACS) modes have been operating as expected and meeting all guidance, navigation, and control (GN&C) requirements, although since launch, several anomalies were encountered. While unplanned, these anomalies have proven to be invaluable in developing a deeper understanding of the ACS, and drove the design of three alterations to the ACS task of the flight software (FSW). An overview of the GN&C subsystem hardware, including re-furbishment, and ACS architecture are introduced, followed by a chronological discussion of key events, flight performance, as well as anomalies encountered by the GN&C team.

Description: The Space Power Facility at NASA's Plum Brook Station houses the world's largest and most powerful space environment simulation facilities, including the Mechanical Vibration Facility (MVF), which offers the world's highest-capacity multi-axis spacecraft shaker system. The MVF was designed to perform sine vibration testing of a Crew Exploration Vehicle (CEV)-class spacecraft with a total mass of 75,000 pounds, center of gravity (cg) height above the table of 284 inches, diameter of 18 feet, and capability of 1.25 gravity units peak acceleration in the vertical and 1.0 gravity units peak acceleration in the lateral directions. The MVF is a six-degree-of-freedom, servo-hydraulic, sinusoidal base-shake vibration system that has the advantage of being able to perform single-axis sine vibration testing of large structures in the vertical and two lateral axes without the need to reconfigure the test article for each axis. This paper discusses efforts to extend the MVF's capabilities so that it can also be used to determine fixed base modes of its test article without the need for an expensive test-correlated facility simulation.

Description: NASA possesses a large quantity of flammability data performed in ISS airlock (30% Oxygen 526mmHg) and ISS cabin (24.1% Oxygen 760 mmHg) conditions. As new programs develop, other oxygen and pressure conditions emerge. In an effort to apply existing data, the question arises: Do equivalent oxygen partial pressures perform similarly with respect to flammability? This paper evaluates how material flammability performance is impacted from both the Maximum Oxygen Concentration (MOC) and Maximum Total Pressures (MTP) perspectives. From these studies, oxygen partial pressures can be compared for both the MOC and MTP methods to determine the role of partial pressure in material flammability. This evaluation also assesses the influence of other variables on flammability performance. The findings presented in this paper suggest flammability is more dependent on oxygen concentration than equivalent partial pressure.

Description: A two-phase brushless DC motor (BDCM) with pulse-width modulated (PWM) voltage drive is simulated to control the flywheel speed of a control moment gyroscope (CMG). An overview of a double-gimballed control moment gyroscope (DGCMG) assembly is presented along with the CMG torque effects on the spacecraft. The operating principles of a two-phase brushless DC motor are presented and the system's electro-mechanical equations of motion are developed for the root-mean-square (RMS) currents and wheel speed. It is shown that the system is an extremely "stiff" set of first-order equations for which an implicit Euler integrator is required for a stable solution. An adaptive proportional voltage controller is presented which adjusts the PWM voltages depending on several control modes for speed, current, and torque. The simulation results illustrate the interaction between the electrical system and the load dynamics and how these influence the overall performance of the system. As will be shown, the CMG spin motor model can directly provide electrical power use and thermal power output to spacecraft subsystems for effective (average) calculations of CMG power consumption.

Description: The propagation of thermally-generated electromagnetic emissions through stratified human tissue is studied herein using a non-coherent mathematical model. The model is developed to complement subsurface body temperature measurements performed using a close proximity microwave radiometer. The model takes into account losses and reflections as thermal emissions propagate through the body, before being emitted at the skin surface. The derivation is presented in four stages and applied to the human core phantom, a physical representation of a stomach volume of skin, muscle, and blood-fatty tissue. A drop in core body temperature is simulated via the human core phantom and the response of the propagation model is correlated to the radiometric measurement. The results are comparable, with differences on the order of 1.5 - 3%. Hence the plausibility of core body temperature extraction via close proximity radiometry is demonstrated, given that the electromagnetic characteristics of the stratified tissue layers are known.

Description: The Regolith and Environment Science and Oxygen and Lunar Volatile Extraction (RESOLVE) payload will transport the (LAVA) subsystem to hydrogen-rich locations on the moon supporting NASA's in-situ resource utilization (ISRU) programs. There, the LAVA subsystem will analyze volatiles that evolve from heated regolith samples in order to quantify how much water is present. To do this, the system needs resilient pressure transducers (PTs) to calculate the moles in the gas samples. The PT trade study includes a comparison of newly-procured models to a baseline unit with prior flight history in order to determine the PT model with the best survivability in flight-forward conditions.

Description: Second-throat diffusers serve to isolate rocket engines from the effects of ambient back pressure. As one of the nation's largest rocket testing facilities, the performance and design limitations of diffusers are of great interest to NASA's Stennis Space Center. This paper describes a series of tests conducted on four diffuser configurations to better understand the effects of inlet geometry and throat area on starting behavior and boundary layer separation. The diffusers were tested for a duration of five seconds with a 1455-pound thrust, LO2/GH2 thruster to ensure they each reached aerodynamic steady state. The effects of a water spray ring at the diffuser exits and a water-cooled deflector plate were also evaluated. Static pressure and temperature measurements were taken at multiple axial locations along the diffusers, and Computational Fluid Dynamics (CFD) simulations were used as a tool to aid in the interpretation of data. The hot combustion products were confirmed to enable the diffuser start condition with tighter second throats than predicted by historical cold-flow data or the theoretical normal shock method. Both aerodynamic performance and heat transfer were found to increase with smaller diffuser throats. Spray ring and deflector cooling water had negligible impacts on diffuser boundary layer separation. CFD was found to accurately capture diffuser shock structures and full-flowing diffuser wall pressures, and the qualitative behavior of heat transfer. However, the ability to predict boundary layer separated flows was not consistent.

Description: All the ISS models are saved in AC3D model format which is a text based format that can be loaded into blender and exported to other formats from there including FBX. The models are saved in two different levels of detail, one being labeled "LOWRES" and the other labeled "HIRES". There are two ".str" files (HIRES _ scene _ load.str and LOWRES _ scene _ load.str) that give the hierarchical relationship of the different nodes and the models associated with each node for both the "HIRES" and "LOWRES" model sets. All the images used for texturing are stored in Windows ".bmp" format for easy importing.

Description: A dual check valve includes, a housing having a cavity fluidically connecting three ports, a movable member movably engaged within the cavity from at least a first position occluding a first port of the three ports, a second position occluding a second port of the three ports, and a third position allowing flow between both the first port, the second port and a third port of the three ports.

Description: The Project Management and Engineering Branch (SF4) supports the Human Health and Performance Directorate (HH&P) and is responsible for developing and supporting human systems hardware for the International Space Station (ISS). When a principal investigator's (PI) medical research project on the ISS is accepted, SF4 develops the necessary hardware and software to transport to the ISS. The two projects I primarily worked on were the centrifuge and ultrasound projects. Centrifuge: One concern with spacecraft such as the ISS is electromagnetic interference (EMI) from onboard equipment, typically from radio waves (frequencies of ~3 kHz to ~300 GHz), which can negatively affect nearby circuitry. Standard commercial centrifuges produce EMI above safety limits, so my task was to help reduce EMI production from this equipment. Two centrifuges were tested: one unmodified as a control and one modified. To reduce EMI below safety limits, one centrifuge was modified to become a Faraday shield, in which significant electrical contact was made between all regions of the centrifuge housing. This included removing non-conductive paint, applying conductive fabric to the lid and foam sealer, adding a 10,000 F decoupling capacitor across the power supply, and adding copper adhesive-mount gaskets to the housing interior. EMI testing of both centrifuges was performed in the EMI/EMC Control Test and Measurement Facility. EMI for both centrifuges was below safety limits for frequencies between 10 MHz and 15 GHz (pass); however, between 14 kHz and 10 MHz, EMI for the unmodified centrifuge exceeded safety limits (fail) as expected. Alternatively, for the modified centrifuge with the Faraday shield, EMI was below the safely limit of 55 dBV/m for electromagnetic frequencies between 14 kHz and 10 MHz. This result indicates our modifications were successful. The successful EMI test allowed us to communicate with the vendor what modifications they needed to make to their commercial unit to meet our specifications and to understand what needs to be done in lab to the new centrifuge. Our modifications will provide a standard for readying centrifuges for future missions. Once the new modified centrifuge arrives by the vendor, it will need to undergo EMI testing again for validation. The centrifuge is also in the process of compatibility testing with a custom stowage drawer, which is an ongoing project in SF4. Both of these items will be payloads on future missions to the ISS for various research purposes. Ultrasound: ISS currently has an onboard ultrasound (Ultrasound 2 system) for research and medical purposes. Every piece of medical flight hardware has an equivalent ground-unit so instrumentation can be routinely evaluated and transported to the ISS if necessary. The ground-unit ultrasound equipment must be evaluated every six months using a task performance sheet (TPS). A TPS is a document, written by the appropriate scientists and engineers, which describes how to run equipment and is written in such a way that astronauts with unspecialized training can follow the tasks. I was responsible for performing six TPSs on a combination of three ultrasounds and two video power converters (VPCs). Performing a TPS involves checking out and computationally documenting each piece of equipment removed from storage locations, setting up hardware and software, performing tasks to verify functionality, returning equipment, and logging items back into the computerized system. My work revealed all ground-unit ultrasounds were functioning properly. Because of proper function, a discrepancy report (DR) did not have to be opened. The TPS was then passed along to the Quality Engineering (QE) for review and ultimately given to Quality Assurance (QA). Other projects: In addition to my main projects, I participated in other tasks including troubleshooting an EEG headband, volunteering for an ultrasound training research study, and conformal coating printed circuit boards. My internship at SF4 has helped me understand how space systems hardware development for the ISS fits into NASA's mission and vision.

Description: Human space exploration to date has been confined to low-Earth orbit and the Moon. The International Space Station (ISS) provides a unique opportunity for researchers to prove out the technologies that will enable humans to safely live and work in space for longer periods of time and venture beyond the Earth/Moon system. The ability to manufacture parts in-space rather than launch them from Earth represents a fundamental shift in the current risk and logistics paradigm for human spaceflight. In September 2014, NASA, in partnership with Made In Space, Inc., launched the 3D Printing in Zero-G technology demonstration mission to explore the potential of additive manufacturing for in-space applications and demonstrate the capability to manufacture parts and tools on orbit using fused deposition modeling. This Technical Publication summarizes the results of testing to date of the ground control and flight prints from the first phase of this ISS payload.

Description: A method and apparatus for docking a spacecraft. The apparatus comprises elongate members, movement systems, and force management systems. The elongate members are associated with a docking structure for a spacecraft. The movement systems are configured to move the elongate members axially such that the docking structure for the spacecraft moves. Each of the elongate members is configured to move independently. The force management systems connect the movement systems to the elongate members and are configured to limit a force applied by the each of the elongate members to a desired threshold during movement of the elongate members.

Description: A system and method of estimation motion of a machine is disclosed. The method may include determining a first point cloud and a second point cloud corresponding to an environment in a vicinity of the machine. The method may further include generating a first extended gaussian image (EGI) for the first point cloud and a second EGI for the second point cloud. The method may further include determining a first EGI segment based on the first EGI and a second EGI segment based on the second EGI. The method may further include determining a first two dimensional distribution for points in the first EGI segment and a second two dimensional distribution for points in the second EGI segment. The method may further include estimating motion of the machine based on the first and second two dimensional distributions.

Description: Designed especially for forest ecosystem studies, EcoSAR employs state-of-the-art digital beamforming technology to generate wide-swath, high-resolution imagery. EcoSARs dual antenna single-pass imaging capability eliminates temporal decorrelation from polarimetric and interferometric analysis, increasing the signal strength and simplifying models used to invert forest structure parameters. Antennae are physically separated by 25 meters providing single pass interferometry. In this mode the radar is most sensitive to topography. With 32 active transmit and receive channels, EcoSARs digital beamforming is an order of magnitude more versatile than the digital beamforming employed on the upcoming NISAR mission. EcoSARs long wavelength (P-band, 435 MHz, 69 cm) measurements can be used to simulate data products for ESAs future BIOMASS mission, allowing scientists to develop algorithms before the launch of the satellite. EcoSAR can also be deployed to collect much needed data where BIOMASS satellite wont be allowed to collect data (North America, Europe and Arctic), filling in the gaps to keep a watchful eye on the global carbon cycle. EcoSAR can play a vital role in monitoring, reporting and verification schemes of internationals programs such as UN-REDD (United Nations Reducing Emissions from Deforestation and Degradation) benefiting global society. EcoSAR was developed and flown with support from NASA Earth Sciences Technology Offices Instrument Incubator Program.

Description: No abstract available

Description: A sealing construct for a space environment includes a seal-bearing object, a seal on the seal-bearing object, and a seal-engaging object. The seal includes a seal body having a sealing surface, and a textured pattern at the sealing surface, the textured pattern defining at least one shaded channel surface. The seal-engaging object is selectively engaged with the seal-bearing object through the seal. The seal-engaging object has a sealing surface, wherein, when the seal-engaging object is selectively engaged with the seal-bearing object, the sealing surface of the seal-engaging object engages the sealing surface of the seal, and the seal is compressed between the seal-bearing object and the seal-engaging object such that at least one shaded channel surface engages the sealing surface of the seal-engaging object.

Description: NASA Marshall Space Flight Center (MSFC) and the Agency as a whole are currently engaged in a number of in-space manufacturing (ISM) activities that have the potential to reduce launch costs, enhance crew safety, and provide the capabilities needed to undertake long-duration spaceflight. The recent 3D Printing in Zero-G experiment conducted on board the International Space Station (ISS) demonstrated that parts of acrylonitrile butadiene styrene (ABS) plastic can be manufactured in microgravity using fused deposition modeling (FDM). This project represents the beginning of the development of a capability that is critical to future NASA missions. Current and future ISM activities will require the development of baseline material properties to facilitate design, analysis, and certification of materials manufactured using in-space techniques. The purpose of this technical interchange meeting (TIM) was to bring together MSFC practitioners and experts in materials characterization and development of baseline material properties for emerging technologies to advise the ISM team as we progress toward the development of material design values, standards, and acceptance criteria for materials manufactured in space. The overall objective of the TIM was to leverage MSFC's shared experiences and collective knowledge in advanced manufacturing and materials development to construct a path forward for the establishment of baseline material properties, standards development, and certification activities related to ISM. Participants were asked to help identify research and development activities that will (1) accelerate acceptance and adoption of ISM techniques among the aerospace design community; (2) benefit future NASA programs, commercial technology developments, and national needs; and (3) provide opportunities and avenues for further collaboration.

Description: Human space exploration to date has been limited to low Earth orbit and the moon. The International Space Station (ISS), an orbiting laboratory 200 miles above the earth, provides a unique and incredible opportunity for researchers to prove out the technologies that will enable humans to safely live and work in space for longer periods of time and venture farther into the solar system. The ability to manufacture parts in-space rather than launch them from earth represents a fundamental shift in the current risk and logistics paradigm for human spaceflight. In particularly, additive manufacturing (or 3D printing) techniques can potentially be deployed in the space environment to enhance crew safety (by providing an on-demand part replacement capability) and decrease launch mass by reducing the number of spare components that must be launched for missions where cargo resupply is not a near-term option. In September 2014, NASA launched the 3D Printing in Zero G technology demonstration mission to the ISS to explore the potential of additive manufacturing for in-space applications and demonstrate the capability to manufacture parts and tools on-orbit. The printer for this mission was designed and operated by the company Made In Space under a NASA SBIR (Small Business Innovation Research) phase III contract. The overarching objectives of the 3D print mission were to use ISS as a testbed to further maturation of enhancing technologies needed for long duration human exploration missions, introduce new materials and methods to fabricate structure in space, enable cost-effective manufacturing for structures and mechanisms made in low-unit production, and enable physical components to be manufactured in space on long duration missions if necessary. The 3D print unit for fused deposition modeling (FDM) of acrylonitrile butadiene styrene (ABS) was integrated into the ISS Microgravity Science Glovebox (MSG) in November 2014 and phase I printing operations took place from November through December of that year. Phase I flight operations yielded 14 unique parts (21 total specimens) that could be directly compared against ground-based prints of identical geometry manufactured using the printer prior to its launch to ISS. The 3DP unit functioned safely and produced specimens necessary to advance the understanding of the critical design and operational parameters for the FDM process as affected by the microgravity environment. From the standpoint of operations, 3DP demonstrated the ability to remove parts from the build-tray on-orbit, teleoperate the printer from the ground, perform critical maintenance functions within defined human factors limits, produce a functional tool that could be evaluated for form/fit/function, and uplink a new part file from the ground and produce it on the printer. The flight parts arrived at NASA Marshall Space Flight Center in Huntsville, Alabama in April 2015, where they underwent months of testing in the materials and processes laboratory. Ground and flight prints completed the following phases of testing: photographic/visual inspection, mass and density evaluation, structured light scanning, XRay and CT, mechanical testing, optical microscopy, scanning electron microscopy, and chemical analysis. This presentation will discuss the results of this testing as well as phase II operations for the printer, which took place in June and July of 2016. Lessons learned from the tech demo and their impacts on the design and development of the second generation 3D printer for ISS, the Additive Manufacturing Facility (AMF) by Made In Space will also be presented. In addition, progress in other elements of NASA's In Space Manufacturing (ISM) initiative such as the on-demand ISM utilization catalog, in-space Recycler ISS Technology Demonstration development, launch packaging recycling, in-space printable electronics, development of higher strength polymeric materials for 3D printing and Additive Construction by Mobile Emplacement (ACME) will also be addressed.

Description: Debris fragments from the hypervelocity impact testing of DebriSat are being collected and characterized for use in updating existing satellite breakup models. One of the key parameters utilized in these models is the ballistic coefficient of the fragment which is directly related to its areatomass ratio. However, since the attitude of fragments varies during their orbital lifetime, it is customary to use the average crosssectional area in the calculation of the areatomass ratio. The average crosssectional area is defined as the average of the projected surface areas perpendicular to the direction of motion and has been shown to be equal to onefourth of the total surface area of a convex object. Unfortunately, numerous fragments obtained from the DebriSat experiment show significant concavity (i.e., shadowing) and thus we have explored alternate methods for computing the average crosssectional area of the fragments. An imaging system based on the volumetric reconstruction of a 3D object from multiple 2D photographs of the object was developed for use in determining the size characteristic (i.e., characteristics length) of the DebriSat fragments. For each fragment, the imaging system generates N number of images from varied azimuth and elevation angles and processes them using a spacecarving algorithm to construct a 3D point cloud of the fragment. This paper describes two approaches for calculating the average crosssectional area of debris fragments based on the 3D imager. Approach A utilizes the constructed 3D object to generate equally distributed crosssectional area projections and then averages them to determine the average crosssectional area. Approach B utilizes a weighted average of the area of the 2D photographs to directly compute the average crosssectional area. A comparison of the accuracy and computational needs of each approach is described as well as preliminary results of an analysis to determine the "optimal" number of images needed for the 3D imager to accurately measure the average cross sectional area of objects with known dimensions.

Description: Spacecraft charging induced by high voltage solar arrays can result in power losses and degradation of spacecraft surfaces. In some cases, it can even present safety issues for astronauts performing extravehicular activities. An understanding of the dominant processes contributing to spacecraft charging induced by solar arrays is important to current space missions, such as the International Space Station, and to any future space missions that may employ high voltage solar arrays. A common method of analyzing the factors contributing to spacecraft charging is the current balance model. Current balance models are based on the simple idea that the spacecraft will float to a potential such that the current collecting to the surfaces equals the current lost from the surfaces. However, when solar arrays are involved, these currents are dependent on so many factors that the equation becomes quite complicated. In order for a current balance model to be applied to solar array operations, it must incorporate the time dependent nature of the charging of dielectric surfaces in the vicinity of conductors1-3. This poster will present the factors which must be considered when developing a current balance model for high voltage solar array operations and will compare results of a current balance model with data from the Floating Potential Measurement Unit4 on board the International Space Station.

Description: On December 5, 2014 NASA conducted the first flight test of its next generation human-class Orion spacecraft. The flight was called the Exploration Flight Test -1 (EFT-1) which lasted for 4 hours and culminated into a re-entry trajectory at 9 km/s. This flight test of the 5-meter Orion Crew Module demonstrated various sub-systems including the Avcoat ablative thermal protection system (TPS) on the heat shield. The Avcoat TPS had been developed from the Apollo-era recipe with a few key modifications. The engineering for thermal sizing was supported by modeling, analysis, and ground tests in arc jet facilities. This paper will describe a postlfight analysis plan and present results from post-recovery inspections, data analysis from embedded sensors, TPS sample extraction and characterization in the laboratory. After the recovery of the vehicle, a full photographic survey and surface scans of the TPS were performed. The recovered vehicle showed physical evidence of flow disturbances, varying degrees of surface roughness, and excessive recession downstream of compression pads. The TPS recession was measured at more than 200 locations of interest on the Avcoat surface. The heat shield was then processed for sample extraction prior to TPS removal using the 7-Axis Milling machine at Marshall Space Flight Center. Around 182 rectangular TPS samples were extracted for subsequent analysis and investigation. The final paper will also present results of sample analysis. The planned investigation includes sidewall imaging, followed by image analysis to characterize TPS response by quantifying different layers in the char and pyrolysis zones. A full postmortem of the instrumentation and sensor ports will also be performed to confirm no adverse effects due to the sensors themselves. A subset of the samples will undergo structural testing and perform detailed characterization of any cracks and integrity of gore seams. Finally, the material will be characterized with layer-by-layer density measurements and SEM investigations to evaluate material morphology at microstructural level including identification of elements and compounds.

Description: Over a decade of work has been conducted in the development of NASA's Hypersonic Inflatable Aerodynamic Decelerator (HIAD) deployable aeroshell technology. This effort has included multiple ground test campaigns and flight tests culminating in the HIAD project's second generation (Gen-2) aeroshell system. The HIAD project team has developed, fabricated, and tested stacked-torus inflatable structures (IS) with flexible thermal protection systems (F-TPS) ranging in diameters from 3-6 meters, with cone angles of 60 and 70 degrees. To meet NASA and commercial near-term objectives, the HIAD team must scale the current technology up to 12-15 meters in diameter. Therefore, the HIAD project's experience in scaling the technology has reached a critical juncture. Growing from a 6-meter to a 15-meter class system will introduce many new structural and logistical challenges to an already complicated manufacturing process. Although the general architecture and key aspects of the HIAD design scale well to larger vehicles, details of the technology will need to be reevaluated and possibly redesigned for use in a 15-meter-class HIAD system. These include: layout and size of the structural webbing that transfers load throughout the IS, inflatable gas barrier design, torus diameter and braid construction, internal pressure and inflation line routing, adhesives used for coating and bonding, and F-TPS gore design and seam fabrication. The logistics of fabricating and testing the IS and the F-TPS also become more challenging with increased scale. Compared to the 6-meter aeroshell (the largest HIAD built to date), a 12-meter aeroshell has four times the cross-sectional area, and a 15-meter one has over six times the area. This means that fabrication and test procedures will need to be reexamined to account for the sheer size and weight of the aeroshell components. This will affect a variety of steps in the manufacturing process, such as: stacking the tori during assembly, stitching the structural webbing, initial inflation of tori, and stitching of F-TPS gores. Additionally, new approaches and hardware will be required for handling and ground testing of both individual tori and the fully assembled HIADs. There are also noteworthy benefits of scaling up the HIAD aeroshell to a 15m-class system. Two complications in working with handmade textile structures are the non-linearity of the material components and the role of human accuracy during fabrication. Larger, more capable, HIAD structures should see much larger operational loads, potentially bringing the structural response of the material components out of the non-linear regime and into the preferred linear response range. Also, making the reasonable assumption that the magnitude of fabrication accuracy remains constant as the structures grow, the relative effect of fabrication errors should decrease as a percentage of the textile component size. Combined, these two effects improve the predictive capability and the uniformity of the structural response for a 12-15-meter HIAD. In this presentation, a handful of the challenges and associated mitigation plans will be discussed, as well as an update on current manufacturing and testing that addressing these challenges.

Description: The National Aeronautical Space Administration has deployed the Meter Class Autonomous Telescope (MCAT) to Ascension Island with plans for it to become fully operational by summer 2016. This telescope will be providing data in support of research being conducted by the Orbital Debris Program Office at the Johnson Space Center. In addition to the main observatory, a smaller, auxiliary telescope is being deployed to the same location to augment and support observations generated by MCAT. It will provide near-simultaneous photometry and astrometry of debris objects, independent measurements of the seeing conditions, and offload low priority targets from MCAT's observing queue. Its hardware and software designs are presented here The National Aeronautical and Space Administration (NASA) has recently deployed the Meter Class Autonomous Telescope (MCAT) to Ascension Island. MCAT will provide NASA with a dedicated optical sensor for observations of orbital debris with the goal of statistically sampling the orbital and photometric characteristics of the population from low Earth to Geosynchronous orbits. Additionally, a small auxiliary telescope, co-located with MCAT, is being deployed to augment its observations by providing near-simultaneous photometry and astrometry, as well as offloading low priority targets from MCAT's observing queue. It will also serve to provide an independent measurement of the seeing conditions to help monitor the quality of the data being produced by the larger telescope. Comprised of off-the-shelf-components, the MCAT Auxiliary Telescope will have a 16-inch optical tube assembly, Sloan g'r'i'z' and Johnson/Cousins BVRI filters, and a fast tracking mount to help facilitate the tracking of objects in low Earth orbit. Tracking modes and tasking will be similar to MCAT except an emphasis will be placed on observations that provide more accurate initial orbit determination for the objects detected by MCAT. The near-simultaneous observations will also provide the opportunity for multi-filter color information of the debris objects to be obtained. Color information can further distinguish the individual objects within the population and provide insight into the reflectance properties of their surface material. The specific hardware, software, and tasking methodology of the MCAT Auxiliary Telescope is presented here..

Description: This presentation summarizes the Marshall Space Flight Center Natural Environments Terrestrial and Planetary Environments (TPE) Team support to the NASA Orion space vehicle. The TPE utilizes meteorological data to assess the sensitivities of the vehicle due to the terrestrial environment. The Orion vehicle, part of the Multi-Purpose Crew Vehicle Program, is designed to carry astronauts beyond low-earth orbit and is currently undergoing a series of tests including Exploration Test Flight (EFT) - 1. The presentation describes examples of TPE support for vehicle design and several tests, as well as support for EFT-1 and planning for upcoming Exploration Missions while emphasizing the importance of accounting for the natural environment's impact to the vehicle early in the vehicle's program.

Description: No abstract available

Description: No abstract available

Description: In order to perform long term missions with multiple objectives using a single space vehicle, there is a need to develop a highly efficient propulsion-navigation system that enables multi-mission capabilities, point-to-point operation and an extended operational lifetime. The majority of space propulsion systems are fuel-based and require the vehicle to carry and consume fuel as part of the mission. Once the fuel is consumed, the mission is terminated. Alternatively, a method that derives its acceleration, velocity and direction from solar photon pressure using a solar sail to capture photon momentum would eliminate the requirement of fuel and all the fuel-based propulsion components. The most important factors that govern the solar sail spacecrafts characteristic acceleration are the sail loading (how much total mass the solar sail has to carry) and the exposed sail area. This paper introduces several potential mission concepts that can be achieved using heliogyro-configured solar sail spacecraft. It then presents 30 potential configurations of heliogyro small spacecraft solar sail and design concepts, based on CubeSat-scale units from 6U to 48U (1U = a cube 10 cm on each side). The area of the sail and total CubeSat masses are used to calculate their characteristic accelerations, and these accelerations are equated to those of previous spacecraft using solar sail technologies; IKAROS, NanoSail-D and LightSail. The analyses in this paper predict that out of these 30 configurations, the 12U-4B(a) configuration has the maximum and the 45U-6B(a) configuration has the minimum characteristic accelerations of 190 and 70 times higher than the IKAROS, 49 and 18 times higher than the NanoSail-D, and 16 and 6 times higher than LightSail, respectively. Several blade deployment configurations, the jelly roll, and a hybrid heliogyro-jelly roll are introduced and compared to the standard reel configuration. The hybrid configurations are predicted to produce higher characteristic accelerations than the jelly roll configurations. The analyses of heliogyro configurations suggest that the amount of payload units (non-sail) when compared to the whole spacecraft allowable units should be less than 40% and the optimized amount, i.e. no empty payload units, is approximately 33% to produce characteristic accelerations 〉 0.7 mm/sq s. For the hybrid configuration, the results suggests that the number of payload units should be between 30 40% of the total units to produce a characteristic acceleration 〉 0.8 mm/sq s.

Description: A piezoelectric sensor with a floating element was developed for direct measurement of flow induced shear stress. The piezoelectric sensor was designed to detect the pure shear stress while suppressing the effect of normal stress generated from the vortex lift-up by applying opposite poling vectors to the piezoelectric elements. During the calibration stage, the prototyped sensor showed a high sensitivity to shear stress (91.3 2.1 pC/Pa) due to the high piezoelectric coefficients (d31=1330 pC/N) of the constituent 0.67Pb(Mg13Nb23)O3-0.33PbTiO3 (PMN- 33%PT) single crystal. By contrast, the sensor showed almost no sensitivity to normal stress (less than 1.2 pC/Pa) because of the electromechanical symmetry of the sensing structure. The usable frequency range of the sensor is up to 800 Hz. In subsonic wind tunnel tests, an analytical model was proposed based on cantilever beam theory with an end-tip-mass for verifying the resonance frequency shift in static stress measurements. For dynamic stress measurements, the signal-tonoise ratio (SNR) and ambient vibration-filtered pure shear stress sensitivity were obtained through signal processing. The developed piezoelectric shear stress sensor was found to have an SNR of 15.8 2.2 dB and a sensitivity of 56.5 4.6 pC/Pa in the turbulent flow.

Description: This paper introduces the Mars Entry Descent and Landing Instrumentation 2 (MEDLI2) concept for NASAs Mars 2020 mission. Mars 2020 is a flagship-class mission, scheduled for launch in 2020, with science and technology objectives to help answer questions about habitability of Mars as well as to demonstrate technologies for future human expedition. MEDLI2 is a suite of instruments embedded in the heatshield and backshell thermal protection systems (TPS) of the Mars 2020 entry vehicle. The objectives of MEDLI2 are to gather critical aerodynamics, aerothermodynamics and TPS (Thermal Protective System) performance data during the Entry Descent and Landing (EDL) phase of the mission.

Description: This work describes the direct simulation Monte Carlo (DSMC) investigation of Saturn entry probe scenarios and the influence of non-equilibrium phenomena on Saturn entry conditions. The DSMC simulations coincide with rarefied hypersonic shock tube experiments of a hydrogen-helium mixture performed in the Electric Arc Shock Tube (EAST) at NASA Ames Research Center. To directly compare to the experimental results, the DSMC simulations are post-processed through the NEQAIR line-by-line radiation code. Improved collision cross-sections, inelastic collision parameters, and reaction rates are determined for a high temperature DSMC simulation of a 7-species H2-He mixture and an electronic excitation model is implemented in the DSMC code. Simulation results for 27.8 and 27.4 kms shock waves are obtained at 0.2 and 0.1 Torr respectively and compared to measured spectra in the VUV, UV, visible, and IR ranges. These results confirm the persistence of non-equilibrium for several centimeters behind the shock and the diffusion of atomic hydrogen upstream of the shock wave. Although the magnitude of the radiance did not match experiments and an ionization inductance period was not observed in the simulations, the discrepancies indicated where improvements are needed in the DSMC and NEQAIR models.

Description: In its twelfth year touring Saturn, the Cassini spacecraft continues to gather valuable scientific data about the planet and its moons. Cassini has executed a total of 331 propulsive maneuvers through January 23, 2016. With more than 30 maneuvers planned through July 2017 before the mission ends in September 2017, a dwindling propellant supply has become a chief concern. This manuscript will report on the analysis of Cassini maneuvers performed through December 30, 2015 and recommend execution-error models for the remainder of the mission. Maneuver performance assessment techniques and execution-error model development methods will also be outlined.

Description: No abstract available

Description: No abstract available

Description: The Atmospheric Infrared Sounder (AIRS) on the EOS Aqua Spacecraft was launched on May 4, 2002 and is currently fully operational. AIRS acquires hyperspectral infrared radiances in 2378 channels ranging in wavelength from 3.7-15.4 um with spectral resolution of better than 1200, and spatial resolution of 13.5 km with global daily coverage. The AIRS was designed to measure temperature and water vapor profiles for improvement in weather forecast and improved parameterization of climate processes. Currently the AIRS Level 1B Radiance Products are assimilated by NWP centers worldwide and have shown considerable forecast improvement. Although the calibration of AIRS (〈 200 mK 3 sigma) is sufficient for data assimilation into Numerical Weather Prediction (NWP) models, long term trends of Earths climate require radiances with stability approaching 10 mK/year, and absolute accuracies better than 100 mK. This investigation uses views of space during roll maneuvers of the Aqua spacecraft to calibrate the mirror emission (one of the largest error sources for AIRS) and reduce the residual errors in cold scenes. We also present results of a secondary study that uses MODIS data to determine the alignment of the AIRS boresight. In this study we match AIRS and MODIS data and iterate on the assumed boresight to find the minimum difference in signal. In this way we are able to confirm the boresight projections determined shortly after launch.

Description: JPL has traditionally performed system level vibration testing of flight spacecraft. The benefits and potential issues of fully assembled flight spacecraft vibration testing are discussed herein. The following specific topics, which may be complementary to the special session on Virtual Vibration Testing, are discussed: spacecraft workmanship, functional and structural integrity testing to uncover workmanship problems, force- and moment-limited vibration testing, potential issues with structural frequency identification using base shake test data, and several failures related to vibration shaker testing. The information provided in this paper is complementary to the special session on Virtual Shaker Testing, and attention is given to issues that virtual shaker testing may face.

Description: In this paper, we present some ideas regarding the modeling and the control aspects of liquid crystal-based imaging systems. We address the challenges of building and deploying a liquid crystal imager by: a) considering the dynamics and control aspects, such as pointing and wavefront error control of large lightweight apertures synthesized by liquid crystal elements, and b) conducting experimental characterization of a prototype electrically-switchable liquid crystal lens.

Description: Proposed is a scalable, six-degree-of-freedom, pressurized docking adapter that can connect multiple volumes while resolving all forces within itself. In a large space outpost pass-through connection is needed between multiple volumes to maintain a continuous pressurized cabin for crew access, translation, and egress. Zero-g docking and berthing of elements can be done using robotic arms, thrusters, and simple docking interface hardware because orthogonal mating is only governed by position, orientation, and momentum, but soft capture / hard docking techniques would not work in a gravity environment because modules cannot be brought in square with each other. Gravity docking is problematic in that any two elements have a gravity vector and it is not practical to provide a perfectly flat surface for them to rest on. Any stretch of natural or graded terrain still has surface fluctuations - maneuvering one element in respect to another would constantly be working against a gravity vector, where uneven surfaces would cause modules to come to rest in odd configurations in respect to each other. Manipulation of heavy elements, such as habitats will be difficult to do with precision -- elements may be placed as close as the mobility system can handle but would still leave the elements not in square with each other. The proposed Pressurized Adapter for "Shirt-Sleeve" Transfer and Universal Base Expansion (PASSTUBE) element will connect non-square and skewed elements while resolving all forces internal to itself.

Description: The Thermal Energy Conversion Technologies Group at the Jet Propulsion Laboratory has been pursuing the development of thermopiles capable of higher temperature operation in support of terrestrial and space sensing applications in extreme environments. The next-generation sensor technology would rely on filled skutterudite (SKD) compounds, the same materials currently being considered for use in the enhanced Multi-Mission Radioisotope Thermoelectric Generator (eMMRTG). We report on the design, development and fabrication of SKD-based prototype thermopile devices. The beginning-of-life (BOL) experimental and predicted performance of the SKD-based prototype thermopile in terms of electrical resistance, output voltage, current, power output and efficiency in the 423 to 723 K operating temperature range is reported and discussed.

Description: This paper describes the electromagnetic compatibility (EMC) requirements on the NASA SMAP mission, the implementation of EMC best practices at various levels of development in subsystems packaging and system level cabling harnesses, and the testing and result s of flight hardware at the sub system and spacecraft system levels.

Description: The Wallops Star Tracker system provides a means of attitude knowledge during balloon missions based on stars detected in the systems field of view. It was developed as an attitude reference source for the Wallops Arc-Second Pointer (WASP) system, but can be utilized as a stand-alone system to support other requirements. The system, known as the Celestial Attitude Reference and Determination System (CARDS), will be described, and the results of the CARDS test flights and the current development status will be presented.

Description: We have developed an optical design for a high resolution spectrograph in response to NASA's call for an extreme precision Doppler spectrometer (EPDS) for the WIYN telescope. Our instrument covers a wavelength range of 380 to 930 nm using a single detector and with a resolution of 100,000. To deliver the most stable spectrum, we avoid the use of an image slicer, in favor of a large (195 mm diameter) beam footprint on a 1x2 mosaic R4 Echelle grating. The optical design is based on a classic white pupil layout, with a single parabolic mirror that is used as the main and transfer collimator. Cross dispersion is provided by a single large PBM2Y glass prism. The refractive camera consists of only four rotationally symmetric lenses made from i-Line glasses, yet delivers very high image quality over the full spectral bandpass. We present the optical design of the main spectrograph bench and discuss the design trade-offs and expected performance.

Description: No abstract available

Description: Numerous mission support hardware systems and their spares are maintained outside of the habitable volume of the International Space Station (ISS), and are arranged covered by a multi-layer insulation (MLI) thermal blanket which provides both thermal control and a measure of protection from micrometeoroids and orbital debris (MMOD). The NASA Hypervelocity Impact Technology (HVIT) group at the Johnson Space Center in Houston Texas has assessed the protection provided by MLI in a series of hypervelocity impact tests using a 1 mm thick aluminum 6061-T6 rear wall to simulate the actual hardware behind the MLI. HVIT has also evaluated methods to enhance the protection provided by MLI thermal blankets. The impact study used both aluminum and steel spherical projectiles accelerated to speeds of 7 km/s using a 4.3 mm, two-stage, light-gas gun at the NASA White Sands Test Facility (WSTF).

Description: No abstract available

Description: Human-scale landers require the delivery of much heavier payloads to the surface of Mars than is possible with entry, descent, and landing (EDL) approaches used to date. A conceptual design was developed for a 10 m diameter crewed Mars lander with an entry mass of approx. 75 t that could deliver approx. 28 t of useful landed mass (ULM) to a zero Mars areoid, or lower, elevation. The EDL design centers upon use of a high ballistic coefficient blunt-body entry vehicle and throttled supersonic retro-propulsion (SRP). The design concept includes a 26 t Mars Ascent Vehicle (MAV) that could support a crew of 2 for approx. 24 days, a crew of 3 for approx.16 days, or a crew of 4 for approx.12 days. The MAV concept is for a fully-fueled single-stage vehicle that utilizes a single pump-fed 250 kN engine using Mono-Methyl Hydrazine (MMH) and Mixed Oxides of Nitrogen (MON-25) propellants that would deliver the crew to a low Mars orbit (LMO) at the end of the surface mission. The MAV concept could potentially provide abort-to-orbit capability during much of the EDL profile in response to fault conditions and could accommodate return to orbit for cases where the MAV had no access to other Mars surface infrastructure. The design concept for the descent stage utilizes six 250 kN MMH/MON-25 engines that would have very high commonality with the MAV engine. Analysis indicates that the MAV would require approx. 20 t of propellant (including residuals) and the descent stage would require approx. 21 t of propellant. The addition of a 12 m diameter supersonic inflatable aerodynamic decelerator (SIAD), based on a proven flight design, was studied as an optional method to improve the ULM fraction, reducing the required descent propellant by approx.4 t.

Description: NASA's Cassini Spacecraft, launched on October 15th, 1997 which arrived at Saturn on June 30th, 2004, is the largest and most ambitious interplanetary spacecraft in history. As the first spacecraft to achieve orbit at Saturn, Cassini has collected science data throughout its four-year prime mission (200408), and has since been approved for a first and second extended mission through 2017. As part of the final extended missions, Cassini will begin an aggressive and exciting campaign of high inclination, low altitude flybys within the inner most rings of Saturn, skimming Saturns outer atmosphere, until the spacecraft is finally disposed of via planned impact with the planet. This final campaign, known as the proximal orbits, requires a strategy for managing the Sun Sensor Assembly (SSA) health, the details of which are presented in this paper.

Description: Model-Based Systems Engineering (MBSE) can augment existing Systems Engineering (SE) processes to more efficiently deliver enhanced products over the project life cycle. Using a multi-user accessible System Model, MBSE has been successfully deployed for the conceptual and preliminary design development of the Asteroid Redirect Robotic Mission (ARRM). The paper provides an overview and examples of the targeted MBSE deployment for development of the mission operational concept, system description, and functional requirements. The paper also includes description of the challenges and lessons learned.

Description: The paper will describe the key technical drivers on the Sylph mission concept to explore a plume at Europa as a secondary free-flyer as a part of the planned Europa Mission. Sylph is a radiation-hardened smallsat concept that would utilize terrain relative navigation to fly at low altitudes through a plume, if found, and relay the mass spectra data back through the flyby spacecraft during its 24-hour mission. The second topic to be discussed will be the mission design constraints of the Near Earth Asterioid (NEA) Scout concept. NEAScout is a 6U cubesat that would utilize an 86 sq. m solar sail as propulsion to execute a flyby with a near-Earth asteroid and help retire Strategic Knowledge Gaps for future human exploration. NEAScout would cruise for 24 months to reach and characterize one Near-Earth asteroid that is representative of Human Exploration targets and telemeter that data directly back to Earth at the end of its roughly 2.5 year mission.

Description: Are NASAs space flight instruments becoming cheaper or more expensive as time marches forward? After analyzing the costs of hundreds of instruments launched over the last 30 years, the short answer to this question is no and yes. This paper gives a visual analysis of the cost time trends for various NASA space flight instrument types, such as optical, particles detectors, fields detectors and microwave instruments. In addition to the statistical approaches utilized, such as significance tests, cluster analysis and principle components analysis (PCA), we will also discuss the intangibles which are likely at play, including technological progress, NASA policy and the luck of the draw associated with mission manifests. This analysis was performed as the main driver for the NASA Instrument Cost Model (NICM) recent cost estimating model redesign. Started in 2004, the first version of NICM was based off of instruments launched from 1985-2005, or 20 years worth of data. As NICM hit its 10-year anniversary, we wanted to know: should NICM continue to only use the most recent 20 years worth of data (1995-2015)? Are instruments becoming cheaper or more expensive as time marches forward? There is evidence in favor of a drop in the median dollar-per-kg value across some instrument types, but little in others. Whereas further research is needed to substantiate, Particles and Optical-Planetary instrument types show moderate to strong evidence of a downward trend in dollar-per-kg. Further research is required to study the nature of this trend (shift, taper, cyclic, etc.). Little evidence for a similar downward trend was detected for Fields or Microwave instruments, or Optical instruments on Earth Orbiting spacecraft. We presented evidence in favor of a drop in the median dollar-per-kg value for Particles and Optical-Planetary instrument types. While similar evidence was weak at best for Fields and Microwave instruments. We can speculate as to the causes for this effect, but we are also equipped to begin to rule out, or at least prioritize, some of the suspected drivers. We observed, for Particles and Optical-Planetary instruments, that perhaps a launch manifest effect was playing part of the role in the observed decrease in dollar-per-kg over the years, noting that the more flagship class missions, which have more money to spend on their instruments, were seen in the earlier years in our data, versus the later years which were dominated by less expensive class missions. However, if this were a dominating driver, would we not have seen the downward trend in the Fields and Microwave instruments as well, which were drawn from that same launch manifest? The fact that we did not observe this helps us rule out the launch manifest effect, and other drivers, such as advances in technology, that seem to be more likely suspects. In that case, however, why would technology advances be helping the Particles and Optical-Planetary instruments only? Why would it not be impacting Optical-Earth Orbiting instruments? Further suspects were looked at as well and ruled out, such as the Faster, Better, Cheaperera of NASA development which did not seem to actually impact trends by instrument type on a dollar-perkg scale. VI. Future Work A. Time Series Detailed Statistical Assessment The analysis discussed above sets the foundation for a more rigorous time series analysis of the data. Time series analysis will further explore evidence to-date of time trends for the instrument types which showed the strongest indicators for a decrease in dollar-per-kg: Optical (Planetary) and Particles instruments. More than providing evidence and top-level significance tests, time series analysis would help elucidate what kind of trend that exists in the data, their significance and allow statistically based forecasting (see Figure 10)

Description: This work presents focal plane modules for a Radiation Budget Instrument, a passive remote-sensing instrument that follows the legacy of the Clouds and Earths Radiant Energy System (CERES) to measure short and longwave Earths radiation budget. The focal plane arrays are micromachined at JPL (Jet Propulsion Laboratory) and integrated into sub-assembly modules to be mounted on the optical telescope of the instrument.

Description: The Dawn mission, part of NASAs Discovery Program, has as its goal the scientific exploration of the two most massive main-belt objects, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H- 9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, transfer between Ceres science orbits, and orbit maintenance maneuvers. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional V of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. From July 2011 through September 2012 the IPS was used to transfer to all the different science orbits at Vesta and to escape from Vesta orbit. Cruise for a rendezvous with Ceres began in August 2012 and completed in late December 2014. From December 2014 through June 2016 the IPS was used for transiting the spacecraft to the Approach phase, survey orbit, the high altitude mapping orbit (HAMO), and the low altitude mapping orbit )LAMO) with arrival to LAMO on December 13, 2015, almost eight years after the start of deterministic thrusting to Vesta. The LAMO orbit, at a mean altitude above Ceres of approximately 385 km, is the spacecrafts final destination and there are no plans to move the spacecraft from LAMO once science operations there are completed. Since arrival at LAMO Dawns IPS has been used for occasional orbit maintenance maneuvers while the spacecraft performs scientific investigations. Dawn has successfully completed its science goals and Dawns primary mission is scheduled to end in the summer of 2016. To date the IPS has been operated for over 48,454 hours, consumed approximately 401 kg of xenon, and provided a delta-V of over 11.0 km/s, a record for an on-board propulsion system. The IPS performance characteristics are close to the expected performance based on analysis and testing performed pre-launch. Dawns IPS continues to be fully operational as of June 2016. This paper provides an overview of Dawns mission objectives and the results of Dawn IPS mission operations from Survey orbit through June 2016.

Description: NASAs Mars Science Laboratory (MSL) spacecraft successfully performed its Entry, Descent & Landing (EDL) phase on August 6, 2012. This paper presents the thermal response of the MSL spacecraft from EDL Initialization (5 days prior to Entry) to Rover touchdown on the surface of Mars. Temperature telemetry recorded during EDL is used to reconstruct the thermal response of the spacecraft to each EDL event. Temperature profiles for the Descent Stage and Rover hardware are presented and explained in the context of the changing EDL environments (aerothermal heating and convective cooling) and power states.

Description: Although NASA has no official plans at this time for a mission to return samples from Mars, the Program Formulation Office of the Mars Exploration Program sponsors ongoing mission concept studies, systems analyses, and technology investments which explore different strategies for the potential return of samples from Mars, consistent with the charter of the program and stated priorities of the science community. Maintaining the thermal integrity of collected samples would be very important. In general, samples would be collected, sealed inside tubes, and left on the surface for later retrieval. They would then be inserted into an OS (Orbiting Sample), and carried to a Mars or Solar orbit via a MAV (Mars Ascent Vehicle). Subsequently, an Earth return vehicle would rendezvous with the OS and bring it back to Earth. During ascent from Mars, the OS could serve as the nose cone of the MAV and would be subjected to significant aerodynamic heating from the Martian atmosphere. Once the OS is released from the MAV, its external surface would be exposed to potentially several years of sunlight, eclipse, planetary IR, albedo, and space. The challenge is to ensure that these samples are kept at thermally moderate conditions to preserve their integrity in these widely different environments. Various thermal techniques have been investigated to achieve sample thermal control: use of thermal protection shields and surfaces (ablative and non-ablative) to protect them from adverse exposure to ascent heating, as well combinations of thermo-optical coatings during the orbital phase. The work described herein is part of this ongoing effort & will describe the key challenges related to the thermal control of the potential Mars samples during these phases and the corresponding schemes to overcome them.

Description: In spacecraft that have propulsion lines that are located externally with open bus architecture, the lines are typically insulated by Multi Layer Insulation (MLI) blankets to protect them thermally from the cold space environment. In addition to heat loss through the insulation, mechanical supports used to attach the lines to the spacecraft structure also create heat leaks from the lines. These lines typically have very low thermal conduction in the axial direction, so the heat balance in the lines tends to be very local without much heat spreading. The typical allowable temperature range for hydrazine-based lines is +15/+50C. This tight temperature range has to be maintained for every location on these lines. For typical spacecraft, these lines can be several meters long. Temperature control is typically achieved by closed loop monitoring of temperatures along the lines and the corresponding powering of the heaters in a bang-bang approach to maintain the temperatures within the dead band of the control loop. The temperatures of propulsion lines are a function of several parameters with heat loss characteristics of the MLI being the key one. Unfortunately, this same key characteristic (MLI effective emittance) has a large variation along its length due to its dependence on workmanship, which in turn leads to large uncertainties in the propulsion lines local temperatures. Because of the poor conduction along the axial direction, heat balance along the length varies dramatically from one location to the next, even few inches apart, depending on the combination of the controlling parameters. This paper describes various robust design and implementation approaches that have been investigated to greatly reduce the randomness associated with predicting the temperature of these propulsion lines.

Description: Ocean surface wind (U) is air in motion and stress () is the turbulent transport of momentum between the ocean and the atmosphere. While the strong wind of a tropical cyclone (TC) causes destruction at landfall, it is the surface stress that drags down the TC. There was almost no stress measurement except in dedicated field campaigns and the stress we used was almost entirely derived from wind through a drag coefficient (C(sub D)), as defined by C(sub D) = / ( U(exp 2)). In TC, there is difficulty in measuring strong wind and large uncertainty in the drag coefficient.

Description: In its twelfth year touring Saturn, the Cassini spacecraft continues to gather valuable scientific data about the planet and its moons. Cassini has executed a total of 331 propulsive maneuvers through January 23, 2016. With more than 30 maneuvers planned through July 2017 before the mission ends in September 2017, a dwindling propellant supply has become a chief concern. This manuscript will report on the analysis of Cassini maneuvers performed through December 30, 2015 and recommend execution-error models for the remainder of the mission. Maneuver performance assessment techniques and execution-error model development methods will also be outlined.

Description: NASAs Low-Density Supersonic Decelerator Project is developing and testing the next generation of supersonic aerodynamic decelerators for planetary entry. A key element of that development is the testing of full-scale articles in conditions relevant to their intended use, primarily in the tenuous Mars atmosphere. To achieve this testing, the LDSD project developed a new test architecture for the qualification of their supersonic parachute. A large, helium filled scientific balloon is used to hoist a 4.7 m blunt body test vehicle to an altitude of approximately 32 kilometers. The test vehicle is released from the balloon, spun up for gyroscopic stability, and accelerated to over four times the speed of sound and an altitude of 50 kilometers using a large solid rocket motor. Once at those conditions, the vehicle is despun and the test period begins.

Description: The Regolith and Environment Science Oxygen and Lunar Volatile Extraction (RESOLVE) payload is part of Resource Prospector (RP) along with a rover and a lander that are expected to launch in 2020. RP will identify volatile elements that may be combined and collected to be used for fuel, air, and water in order to enable deeper space exploration. The Resource Prospector mission is a key part of In-Situ Resource Utilization (ISRU). The demand for this method of utilizing resources at the site of exploration is increasing due to the cost of resupply missions and deep space exploration goals. The RESOLVE payload includes the Lunar Advanced Volatile Analysis (LAVA) subsystem. The main instrument used to identify the volatiles evolved from the lunar regolith is the Gas Chromatograph-Mass Spectrometer (GC-MS). LAVA analyzes the volatiles emitted from the Oxygen and Volatile Extraction Node (OVEN) Subsystem. The objective of OVEN is to obtain, weigh, heat and transfer evolved gases to LAVA through the connection between the two subsystems called the LOVEN line. This paper highlights the work completed during a ten week internship that involved the integration, testing, data analysis, and procedure documentation of two candidate mass spectrometers for the LAVA subsystem in order to aid in determining which model to use for flight. Additionally, the examination of data from the integrated Resource Prospector '15 (RP' 15) field test will be presented in order to characterize the amount of water detected from water doped regolith samples.

Description: The Visible Infrared Imaging Radiometer Suite (VIIRS) on-board the first Joint Polar Satellite System (JPSS) completed its sensor level testing on December 2014. The JPSS-1 (J1) mission is scheduled to launch in December 2016, and will be very similar to the Suomi-National Polar-orbiting Partnership (SNPP) mission. VIIRS instrument has 22 spectral bands covering the spectrum between 0.4 and 12.6 m. It is a cross-track scanning radiometer capable of providing global measurements twice daily, through observations at two spatial resolutions, 375 m and 750 m at nadir for the imaging and moderate bands, respectively. This paper will briefly describe J1 VIIRS characterization and calibration performance and methodologies executed during the pre-launch testing phases by the government independent team to generate the at-launch baseline radiometric performance and the metrics needed to populate the sensor data record (SDR) Look-Up-Tables (LUTs). This paper will also provide an assessment of the sensor pre-launch radiometric performance, such as the sensor signal to noise ratios (SNRs), radiance dynamic range, reflective and emissive bands calibration performance, polarization sensitivity, spectral performance, response-vs-scan (RVS), and scattered light response. A set of performance metrics generated during the pre-launch testing program will be compared to both the VIIRS sensor specification and the SNPP VIIRS pre-launch performance.

Description: No abstract available

Description: No abstract available