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At least 73 records · Page 4

Analyzing the Impacts of Natural Environments on Launch and Landing Availability for NASA's Exploration Systems Development Programs

The National Aeronautics and Space Administration (NASA) is developing new capabilities for human and scientific exploration beyond Earth orbit. Natural environments information is an important asset for NASA's development of the next generation space transportation system as part of the Exploration Systems Development (ESD) Programs, which includes the Space Launch System (SLS) and Multi-Purpose Crew Vehicle (MPCV) Programs. Natural terrestrial environment conditions - such as wind, lightning and sea states - can affect vehicle safety and performance during multiple mission phases ranging from pre-launch ground processing to landing and recovery operations, including all potential abort scenarios. Space vehicles are particularly sensitive to these environments during the launch/ascent and the entry/landing phases of mission operations. The Marshall Space Flight Center (MSFC) Natural Environments Branch provides engineering design support for NASA space vehicle projects and programs by providing design engineers and mission planners with natural environments definitions as well as performing custom analyses to help characterize the impacts the natural environment may have on vehicle performance. One such analysis involves assessing the impact of natural environments to operational availability. Climatological time series of operational surface weather observations are used to calculate probabilities of meeting/exceeding various sets of hypothetical vehicle-specific parametric constraint thresholds. Outputs are tabulated by month and hour of day to show both seasonal and diurnal variation. This paper will discuss how climate analyses are performed by the MSFC Natural Environments Branch to support the ESD Launch Availability (LA) Technical Performance Measure (TPM), the SLS Launch Availability due to Natural Environments TPM, and several MPCV (Orion) launch and landing availability analyses - including the 2014 Orion Exploration Flight Test 1 (EFT-1) mission.

Altino, Karen M.↗

Analyzing the Impacts of Natural Environments on Launch and Landing Availability for NASA's Eploration Systems Development Programs

NASA is developing new capabilities for human and scientific exploration beyond Earth orbit. Natural environments information is an important asset for NASA's development of the next generation space transportation system as part of the Exploration Systems Development Program, which includes the Space Launch System (SLS) and MultiPurpose Crew Vehicle (MPCV) Programs. Natural terrestrial environment conditions - such as wind, lightning and sea states - can affect vehicle safety and performance during multiple mission phases ranging from prelaunch ground processing to landing and recovery operations, including all potential abort scenarios. Space vehicles are particularly sensitive to these environments during the launch/ascent and the entry/landing phases of mission operations. The Marshall Space Flight Center (MSFC) Natural Environments Branch provides engineering design support for NASA space vehicle projects and programs by providing design engineers and mission planners with natural environments definitions as well as performing custom analyses to help characterize the impacts the natural environment may have on vehicle performance. One such analysis involves assessing the impact of natural environments to operational availability. Climatological time series of operational surface weather observations are used to calculate probabilities of meeting or exceeding various sets of hypothetical vehicle-specific parametric constraint thresholds.

Altino, Karen M.↗

Applied Multi-Mission Telemetry Processing and Display for Operations, Integration, Training, Playback and Event Reconstruction

Conveying spacecraft health and status information to mission engineering personnel during various mission phases, including mission operations, is a requirement to achieve a successful mission. For NASA/JPL spacecraft, that often means displaying hundreds of telemetry channels from a variety of sensors and components emitting data at rates varying from 1hz-100hz (and faster) in a way that allows the operations team to quickly evaluate the health of the vehicle, identify any off-nominal states and resolve any issues. In this paper we will discuss the system design, requirements and use cases of three telemetry processing and visualization systems recently developed and deployed by our team for NASA's Low Density Supersonic Decelerator (LDSD) test vehicle, NASA's Soil Moisture Active/Passive (SMAP) orbiter, and JPL's Sampling Lab Universal Robotic Manipulator (SLURM) test bed.

simulation↗

Enterprise Mission Integration for Artemis Lunar Missions

Mission integration is an iterative process by which a specific mission is formulated, refined, planned, and executed within the established vehicle(s), architecture, and ground systems design. Mission integration includes the people, vehicle(s) and ground hardware/software, products, processes, analyses, schedules, facilities, Certification of Flight Readiness, etc. The Artemis Mission Integration Task Team (MITT) developed a series of products and processes to support the complex mission integration across various Programs within the Artemis Mission Campaign (Orion, Space Launch Systems, Exploration Ground Systems, Gateway, Human Landing System, and Extravehicular Activity and Human Surface Mobility). The Moon to Mars (M2M) Program is referred to as ‘the enterprise’ as it includes both the M2M organization and the Programs supporting the Artemis Mission Campaign. Artemis Mission Integration has five phases: mission capability, mission definition, mission preparation, mission execution, and post-mission assessment. This paper focuses on one of the enterprise-level mission checkpoints as a kick-off to the Mission Preparation phase, the Mission Integration Review (MIR), which occurs 18-24 months prior to launch. The MIR helps to confirm the defined mission technical baseline is within the existing analyzed design envelope. Details are provided on the identification of dependencies, issues, or gaps for mission-specific objectives and requirements, as well as the definition of the analysis, training, mission execution products, facilities, and detailed supporting operations requirements. The MIR was held for both Artemis I and II and this paper aims to share with the aerospace community its value as we prepare for upcoming Artemis Missions.

Mary Anne Plaza↗

Thermal design and development of a planetary probe - Pioneer Venus large probe

The thermal control system developed for the Large Atmospheric Probe of the Pioneer Venus Multiprobe Mission is described. The scope of the thermal control task requires maintaining the probe internal equipment shelf temperatures within a nonoperating range of -40 to 122 F and an operating range of -4 to 122 F during three different mission phases: (1) preseparation, the transit phase of the mission when the probe is attached to the Multiprobe Spacecraft Bus, (2) postseparation, the free-flight cruise phase of the mission following release from the Bus, and (3) descent, the phase of the mission from preentry equipment turn-on to impact on the Venus surface. Thermal control for these phases is achieved by a combination of passive thermal finishes on the probe exterior surfaces and heaters mounted on the equipment shelves. Verification of the adequacy of the total thermal design to meet all mission requirements has been completed.

Hennis, L. A.↗

Orbit determination covariance analysis for the Deep Space Program Science Experiment mission

To define an appropriate orbit support procedure for the DSPSE mission, detailed permission orbit determination covariance analyses have been performed for the translunar and trans-Geographos mission phases. Preliminary analyses were also performed for the lunar mapping mission phase. These analyses are designed to assess the tracking patterns and the amount of tracking data needed to obtain orbit solutions of required accuracy for each mission phase and before and after each major orbit perturbation, such as orbit maneuvers and flybys of the Earth and Moon. In addition to operational orbit determination procedures, these analyses identify major error sources, estimate their contribution to orbital errors, and address possible strategies to reduce orbit determination error. For the lunar orbit phase, several lunar gravity error modeling approaches have been investigated. The covariance analysis results presented in this paper will serve as a guide for providing orbit determination support for the DSPSE mission.

Beckman, M.↗

A study of system requirements for Phobos/Deimos missions. Volume 4: Phase 3 results, combined missions to Mars and its satellites

Combined missions to Mars and Phobos/Deimos were considered in relation to large scientific returns at a low total cost. A matrix of all feasible major combined mission concepts was compiled. The most promising approach was selected from which a nominal mission profile was developed. A systems analysis was performed to trade off performance, configuration, and cost characteristics of candidate mission concepts. Configuration and subsystem optimization analysis were then conducted from which a baseline and an alternate program were selected. Program schedules and cost estimates were prepared for the recommended and alternate program.

Source record↗

Magnetospheric Multiscale Mission (MMS) Phase 2B Navigation Performance

The Magnetospheric Multiscale (MMS) formation flying mission, which consists of four spacecraft flying in a tetrahedral formation, has challenging navigation requirements associated with determining and maintaining the relative separations required to meet the science requirements. The baseline navigation concept for MMS is for each spacecraft to independently estimate its position, velocity and clock states using GPS pseudorange data provided by the Goddard Space Flight Center-developed Navigator receiver and maneuver acceleration measurements provided by the spacecraft's attitude control subsystem. State estimation is performed onboard in real-time using the Goddard Enhanced Onboard Navigation System flight software, which is embedded in the Navigator receiver. The current concept of operations for formation maintenance consists of a sequence of two maintenance maneuvers that is performed every 2 weeks. Phase 2b of the MMS mission, in which the spacecraft are in 1.2 x 25 Earth radii orbits with nominal separations at apogee ranging from 30 km to 400 km, has the most challenging navigation requirements because, during this phase, GPS signal acquisition is restricted to less than one day of the 2.8-day orbit. This paper summarizes the results from high-fidelity simulations to determine if the MMS navigation requirements can be met between and immediately following the maintenance maneuver sequence in Phase 2b.

Scaperoth, Paige Thomas↗

Ground Segment Operations Concept for the Orion Artemis-2 Optical Communications System

The ACCESS Project (formerly Space Network) will implement an optical communications ground segment to support the Orion Artemis II Optical Communications (O2O) demonstration as part of the next manned human spaceflight mission to the moon, Artemis II. O2O implements laser communication (lasercomm) technology for operational use on the Orion series of spacecraft, as a development test objective (DTO), in order to demonstrate the feasibility and operational utility of lasercomm for human spaceflight missions. O2O consists of three segments: Space Segment, Ground Segment, and Operations Segment. The Space Segment consists of the Space Terminal Element and the Orion spacecraft. The Space Terminal Element effort is managed by the GSFC Laser-Enhanced Mission Communications Navigation and Operational Services (LEMNOS) project in collaboration with MIT Lincoln Laboratory. The Ground Segment consists of an optical ground terminal (GT) at the White Sands Complex (WSC), which is being developed in collaboration with MIT Lincoln Laboratory, the Ground Segment Operations and Analysis (GSOA) element and Ground Data Element (GDE), and a second optical GT in the Optical Communications Telescope Laboratory (OCTL) at the JPL Table Mountain Facility. The Operations Segment consists of the Artemis II Mission Control Center (MCC), the Lasercomm Space Terminal Console (LSTC), and the Lasercomm Link Planning & Analysis Center (LPAC), all located at the Johnson Space Center (JSC). O2O utilizes pulse-position modulation (PPM) direct-to-earth services resulting in an 80 Mbps downlink data rate from lunar orbit. The O2O concept of operations is to provide optical services for a minimum of 1 hour per day for each day of the Artemis II mission. O2O will utilize a 10-20 Mbps uplink data rate and 40-260 Mbps downlink data rate, depending on the Artemis II mission phase. The ACCESS project will also provide a centralized mission data interface for user data distribution and storage to the MCC and perform planning and scheduling of services in coordination with the Operations Segment for the O2O Ground Segment. The O2O Ground Segment will support the following O2O mission phases: Pre-Mission Planning; Daily Operations Planning; Event Execution; and Post-Pass Reporting. O2O will be used to exchange data files between Orion and the MCC and to distribute real-time video through the optical downlink service to the MCC; which would not be possible without the high-bandwidth link that O2O will provide to Orion. In this paper, I will discuss the O2O Ground Segment development approach and how it will support these critical O2O functions: plan and schedule the contact; acquire and track the optical link; flow information bidirectionally; distribute information; and control and accommodate the system.

optical communications↗

On-Orbit Prospective Echocardiography on International Space Station Crew

Introduction A prospective trial of echocardiography was conducted on of six crewmembers onboard the International Space Station. The main objective was to determine the efficacy of remotely guided tele-echocardiography, including just-in-time e-training methods and determine what "space normal" echocardiographic data is. Methods Each crewmember operator (n=6) had 2-hour preflight training. Baseline echocardiographic data were collected 55 to 167days preflight. Similar equipment was used in each 60-minute in-flight session (mean microgravity exposure - 114 days (34 -- 190)). On Orbit ultrasound operators used an e-learning system within 24h of these sessions. Expert assistance was provided using ultrasound video downlink and two-way voice. Testing was repeated 5 to 16 days after landing. Separate ANOVA was used on each echocardiographic variable (n=33). Within each ANOVA, three tests were made: a) effect of mission phase (preflight, in-flight, post flight); b) effect of echo technician (two technicians independently analyzed the data); c) interaction between mission phase and technician. Results Nine rejections of the null hypothesis (mission phase or technician or both had no effect) were discovered and considered for follow up. Of these, six rejections were for significant technician effects, not as a result of space flight. Three rejections of the null hypothesis (Aortic Valve time velocity integral, Mitral E wave Velocity and heart rate) were attributable to space flight, however determined not to be clinically significant. No rejections were due to the interaction between technician and space flight. Conclusion No consistent clinically significant effects of long-duration space flight were seen in echocardiographic variables of the given group of subjects.

Hamilton, Douglas R.↗

Operability Engineering for the Europa Clipper Mission: Formulation Phase Results and Lessons

Operability is an important factor in both success and cost of space missions, but is difficult to quantify, and can suffer if not considered as part of early formulation and design. The Europa Clipper mission has taken a deliberate approach to infusing operability into the concepts and designs since early in the formulation phase. This paper reports on that approach to operability and the results thus far, as the Europa Clipper mission approaches its Preliminary Design Review. Definitions for operability and its various aspects are presented along with a description of the engineering and organizational approach taken to infusing operability into designs. A number of examples of operability requirements are shown, and key examples of trade studies and design decisions are described, along with influence on outcomes motivated by operability considerations. We find that support from management, empowerment of a cross-discipline Operability Working Group, consistent tracking of operability aspects over time, and broad infusion and participation in operability considerations are all significant factors in implementing operability. In particular, thoughtful and deliberate consideration of operability criteria has led to design decisions that include proper consideration of concerns related to the mission operations phase, and seem likely to provide better science results and mission outcomes.

Huh, Shin M.↗

EVAL mission requirements, phase 1

The aspects of NASA's applications mission were enhanced by utilization of shuttle/spacelab, and payload groupings which optimize the cost of achieving the mission goals were defined. Preliminary Earth Viewing Application Laboratory (EVAL) missions, experiments, sensors, and sensor groupings were developed. The major technological EVAL themes and objectives which NASA will be addressing during the 1980 to 2,000 time period were investigated. Missions/experiments which addressed technique development, sensor development, application development, and/or operational data collection were considered as valid roles for EVAL flights.

Source record↗

Application of Fault Management Theory to the Quantitative Selection of a Launch Vehicle Abort Trigger Suite

The theory of System Health Management (SHM) and of its operational subset Fault Management (FM) states that FM is implemented as a "meta" control loop, known as an FM Control Loop (FMCL). The FMCL detects that all or part of a system is now failed, or in the future will fail (that is, cannot be controlled within acceptable limits to achieve its objectives), and takes a control action (a response) to return the system to a controllable state. In terms of control theory, the effectiveness of each FMCL is estimated based on its ability to correctly estimate the system state, and on the speed of its response to the current or impending failure effects. This paper describes how this theory has been successfully applied on the National Aeronautics and Space Administration's (NASA) Space Launch System (SLS) Program to quantitatively estimate the effectiveness of proposed abort triggers so as to select the most effective suite to protect the astronauts from catastrophic failure of the SLS. The premise behind this process is to be able to quantitatively provide the value versus risk trade‐off for any given abort trigger, allowing decision makers to make more informed decisions. All current and planned crewed launch vehicles have some form of vehicle health management system integrated with an emergency launch abort system to ensure crew safety. While the design can vary, the underlying principle is the same: detect imminent catastrophic vehicle failure, initiate launch abort, and extract the crew to safety. Abort triggers are the detection mechanisms that identify that a catastrophic launch vehicle failure is occurring or is imminent and cause the initiation of a notification to the crew vehicle that the escape system must be activated. While ensuring that the abort triggers provide this function, designers must also ensure that the abort triggers do not signal that a catastrophic failure is imminent when in fact the launch vehicle can successfully achieve orbit. That is, the abort triggers must have low false negative rates to be sure that real crew‐threatening failures are detected, and also low false positive rates to ensure that the crew does not abort from non‐crew‐threatening launch vehicle behaviors. The analysis process described in this paper is a compilation of over six years of lessons learned and refinements from experiences developing abort triggers for NASA's Constellation Program (Ares I Project) and the SLS Program, as well as the simultaneous development of SHM/FM theory. The paper will describe the abort analysis concepts and process, developed in conjunction with SLS Safety and Mission Assurance (S&MA) to define a common set of mission phase, failure scenario, and Loss of Mission Environment (LOME) combinations upon which the SLS Loss of Mission (LOM) Probabilistic Risk Assessment (PRA) models are built. This abort analysis also requires strong coordination with the Multi‐Purpose Crew Vehicle (MPCV) and SLS Structures and Environments (STE) to formulate a series of abortability tables that encapsulate explosion dynamics over the ascent mission phase. The design and assessment of abort conditions and triggers to estimate their Loss of Crew (LOC) Benefits also requires in‐depth integration with other groups, including Avionics, Guidance, Navigation and Control(GN&C), the Crew Office, Mission Operations, and Ground Systems. The outputs of this analysis are a critical input to SLS S&MA's LOC PRA models. The process described here may well be the first full quantitative application of SHM/FM theory to the selection of a sensor suite for any aerospace system.

Lo, Yunnhon↗

Design of a Formation of Earth-Orbiting Satellites: The Auroral Lites Mission

The National Aeronautics and Space Administration (NASA) Goddard Space Flight Center (GSFC) has proposed a set of spacecraft flying in close formation around the Earth in order to measure the behavior of the auroras. The mission, named Auroral Lites, consists of four spacecraft configured to start at the vertices of a tetrahedron, flying over three mission phases. During the first phase, the distance between any two spacecraft in the formation is targeted at 10 kilometers (km). The second mission phase is much tighter, requiring satellite interrange spacing targeted at 500 meters. During the final phase of the mission, the formation opens to a nominal 100-km interrange spacing. In this paper, we present the strategy employed to initialize and model such a close formation during each of these phases. The analysis performed to date provides the design and characteristics of the reference orbit, the evolution of the formation during Phases I and II, and an estimate of the total mission delta-V budget. AI Solutions' mission design tool, FreeFlyer, was used to generate each of these analysis elements. The tool contains full force models, including both impulsive and finite duration maneuvers. Orbital maintenance can be fully modeled in the system using a flexible, natural scripting language built into the system. In addition, AI Solutions is in the process of adding formation extensions to the system facilitating mission analysis for formations like Auroral Lites. We will discuss how FreeFlyer is used for these analyses.

Hametz, Mark E.↗

Design of a Formation of Earth Orbiting Satellites: The Auroral Lites Mission

The National Aeronautics and Space Administration (NASA) Goddard Space Flight Center (GSFC) has proposed a set of spacecraft flying in close formation around the Earth in order to measure the behavior of the auroras. The mission, named Auroral Lites, consists of four spacecraft configured to start at the vertices of a tetrahedron, flying over three mission phases. During the first phase, the distance between any two spacecraft in the formation is targeted at 10 kilometers (km). The second mission phase is much tighter, requiring satellite interrange spacing targeted at 500 meters. During the final phase of the mission, the formation opens to a nominal 100-km interrange spacing. In this paper, we present the strategy employed to initialize and model such a close formation during each of these phases. The analysis performed to date provides the design and characteristics of the reference orbit, the evolution of the formation during Phases I and II, and an estimate of the total mission delta-V budget. AI Solutions' mission design tool, FreeFlyer(R), was used to generate each of these analysis elements. The tool contains full force models, including both impulsive and finite duration maneuvers. Orbital maintenance can be fully modeled in the system using a flexible, natural scripting language built into the system. In addition, AI Solutions is in the process of adding formation extensions to the system facilitating mission analysis for formations like Auroral Lites. We will discuss how FreeFlyer(R) is used for these analyses.

Hametz, Mark E.↗

Validation Assessment for the Soil Moisture Active Passive (SMAP) Level 4 Carbon (L4_C) Data Product Version 5

The post-launch Cal/Val phase of the SMAP mission is guided by two primary objectives for each science product team: 1) to calibrate, verify, and improve the performance of the science algorithms, and 2) validate accuracies of the science data products as specified in the SMAP Level-1 mission science requirements. Algorithm science and product maintenance activities during the SMAP extended mission phase have also involved periodic algorithm calibration and product refinements to maintain or enhance product consistency and performance as well as science utility. This report provides an assessment of the latest (Version 5) SMAP Level 4 Carbon (L4_C) product. The L4_C Version 5 (v5) global record now spans more than six years (March 2015 – present) of SMAP operations and has benefited from five major reprocessing updates to the operational product. These reprocessing events and L4_C product release updates have incorporated various algorithm refinements and calibration adjustments to account for similar refinements to the upstream GEOS land model assimilation system, SMAP brightness temperatures, and MODIS vegetation inputs used for L4_C processing. The SMAP L4_C algorithms utilize a terrestrial carbon flux model informed by daily surface and root zone soil moisture information contributed from the SMAP Level 4 Soil Moisture (L4_SM) product along with optical remote sensing-based (e.g. MODIS-based) land cover and canopy fractional photosynthetic active radiation (fPAR), and other ancillary biophysical data. The carbon flux model estimates global daily net ecosystem CO2 exchange (NEE) and the component carbon fluxes, namely, vegetation gross primary production (GPP) and soil heterotrophic respiration (Rh). Other L4_C product elements include surface (~0-5 cm depth) soil organic carbon (SOC) stocks and associated environmental constraints to these processes, including soil moisture-related controls on GPP and ecosystem respiration (Kimball et al. 2014, Jones et al. 2017). The L4_C product addresses SMAP carbon cycle science objectives by: 1) providing a direct link between terrestrial carbon fluxes and underlying freeze/thaw and soil moisture-related constraints to these processes, 2) documenting primary connections between terrestrial water, energy and carbon cycles, and 3) improving understanding of terrestrial carbon sink activity. The SMAP L4_C algorithms and operational product are mature and at a CEOS Validation Stage 4 level (Jackson et al. 2012) based on extensive validation of the multi-year record against a diverse array of independent benchmarks, well characterized global performance, and systematic refinements gained from five major reprocessing events. There are no Level-1 mission science requirements for the L4_C product; however, self-imposed requirements have been established focusing on NEE as the primary product field for validation, and on demonstrating L4_C accuracy and success in meeting product science requirements (Jackson et al. 2012). The other L4_C product fields also have strong utility for carbon science applications (e.g., Liu et al. 2019, Endsley et al. 2020); however, analysis of these other fields is considered secondary relative to primary validation activities focusing on NEE. The L4_C targeted accuracy requirements are to meet or exceed a mean unbiased root-mean-square error (ubRMSE, or standard deviation of the error) for NEE of 1.6 g C m-2 d-1 and 30 g C m-2 yr-1, emphasizing northern (≥45°N) boreal and arctic ecosystems; this accuracy is similar to that of tower eddy covariance measurement-based observations (Baldocchi 2008). Methods used for the latest v5 L4_C product performance and validation assessment have been established from the SMAP Cal/Val plan and previous studies (Jackson et al. 2012, Jones et al. 2017) and include: 1) consistency evaluations of the product fields against earlier product releases (version 4 or earlier); 2) comparisons of daily carbon flux estimates with independent tower eddy 4covariance measurement-based daily carbon (CO2) flux observations from core tower validation sites (CVS); and 3) consistency checks against other global carbon products, including soil carbon inventory records, global GPP records derived from tower observation upscaling methods, and satellite-based observations of canopy solar induced chlorophyll fluorescence (SIF) as a surrogate for GPP. Metrics used to evaluate relative agreement between L4_C product fields and observational benchmarks include correlation (r-value), RMSE differences, bias and model sensitivity diagnostics. Following these validation criteria, the present report provides a validation assessment of the latest L4_C product release (v5). Detailed descriptions of the L4_C algorithm and additional global product accuracy and performance results are given elsewhere (Jones et al. 2017, Endsley et al. 2020). The v5 L4_C product replaces earlier product versions and continues to show: (i) accuracy and performance levels meeting or exceeding SMAP L4_C science requirements; (ii) improvement over the previous product version (version 4); and (iii) suitability for a diversity of science applications. Example L4_C applications from the recent literature include clarifying environmental trends and controls on the northern terrestrial carbon sink (Liu et al. 2019), diagnosing drought-related impacts on ecosystem productivity (Li et al. 2020), and regional monitoring of cropland conditions for projecting annual yields (Wurster et al. 2020). 2EXPECTED L4_C ALGORITHM AND PRODUCTPERFORMANCE The L4_C algorithm performance, including variance and uncertainty estimates of model outputs, was determined during the mission pre-launch phase through spatially explicit model sensitivity studies using available model inputs similar to those currently being used for operational production and evaluating the resulting model simulations over the observed range of northern (≥45 °N) and global conditions (Kimball et al. 2012, Entekhabi et al. 2014). The L4_C algorithm options were also evaluated during the mission prelaunch phase, including deriving canopy fPAR from lower order NDVI (Normalized Difference Vegetation Index) inputs in lieu of using MODIS (MOD15) fPAR; and including an explicit model representation of boreal fire disturbance recovery impacts. These results indicated that the L4_C accuracy requirements (i.e., NEE ubRMSE ≤ 30 g C m-2 yr-1or ≤ 1.6 g C m-2 d-1) could be met from the baseline algorithms over more than 82% and 89% of global and northern vegetated land areas, respectively (Yi et al. 2013, Kimball et al. 2014). The global L4_C algorithm error budget for NEE derived during the mission prelaunch phase indicated that the estimated NEE ubRMSE uncertainty is proportional to GPP and is therefore larger in higher biomass productivity areas, including forests and croplands (Kimball et al. 2014). Likewise, NEE ubRMSE uncertainty is expected to be lower in less-productive areas, including grasslands and shrublands. Expected model NEE ubRMSE levels were also generally within targeted accuracy levels for characteristically less-productive boreal and Arctic biomes, even though relative model error as a proportion of total productivity (NEE RMSE / GPP) may be large in these areas. The estimated NEE uncertainty was lower than expected in some warmer tropical high biomass productivity areas (e.g. Amazon rainforest) because of reduced low temperature and moisture constraints to the L4_C respiration calculations so that the bulk of model uncertainty is contributed by GPP in these areas. Model NEE uncertainty in the African Congo was estimated to

SMAP↗

Navigation of the Mariner Tempel 2 mission - Orbit phase and penetrator deployment

The navigation of the Mariner Mark II spacecraft about the comet Tempel 2 in the Comet Rendezvous Asteroid Flyby (CRAF) mission is described. This mission is the first of a series using the Mariner Mark II spacecraft planned for the 1990's and thereafter. The primary purpose of the mission is to orbit the nucleus of the periodic comet Tempel 2 and study it at close range. A secondary purpose is to fly by and observe one or two asteroids enroute to the comet. During the orbital phase of the mission, a penetrator will be deployed from the spacecraft to impact the surface of the comet and measure the composition and other properties of the nucleus. Navigation accuracies are given in this paper for spacecraft and penetrator trajectory prediction and control. The orbit determination performance in support of science observations while in orbit about the nucleus is also described.

Miller, James K.↗

Analysis of One-Way Laser Ranging Data to LRO, Time Transfer and Clock Characterization

We processed and analyzed one-way laser ranging data from International Laser Ranging Service ground stations to NASA's Lunar Reconnaissance Orbiter (LRO), obtained from June 13, 2009 until September 30, 2014. We pair and analyze the one-way range observables from station laser fire and spacecraft laser arrival times by using nominal LRO orbit models based on the GRAIL gravity field. We apply corrections for instrument range walk, as well as for atmospheric and relativistic effects. In total we derived a tracking data volume of approximately 3000 hours featuring 64 million Full Rate and 1.5 million Normal Point observations. From a statistical analysis of the dataset we evaluate the experiment and the ground station performance. We observe a laser ranging measurement precision of 12.3 centimeters in case of the Full Rate data which surpasses the LOLA (Lunar Orbiting Laser Altimeter) timestamp precision of 15 centimeters. The averaging to Normal Point data further reduces the measurement precision to 5.6 centimeters. We characterized the LRO clock with fits throughout the mission time and estimated the rate to 6.9 times10 (sup -8), the aging to 1.6 times 10 (sup -12) per day and the change of aging to 2.3 times 10 (sup -14) per day squared over all mission phases. The fits also provide referencing of onboard time to the TDB (Barycentric Dynamical Time) time scale at a precision of 166 nanoseconds over two and 256 nanoseconds over all mission phases, representing ground to space time transfer. Furthermore we measure ground station clock differences from the fits as well as from simultaneous passes which we use for ground to ground time transfer from common view observations. We observed relative offsets ranging from 33 to 560 nanoseconds and relative rates ranging from 2 times 10 (sup -13) to 6 times 10 (sup -12) between the ground station clocks during selected mission phases. We study the results from the different methods and discuss their applicability for time transfer.

Bauer, S.↗