GOES RSeriesDataBook

Transcript

1 GOES R Series - Data Book Prepared for National Aeronautics and Space Administration GOES - R Series Program Office Goddard Space Flight Center Greenbelt, Maryland 20771 Pursuant to Contract NNG09HR00C Rev - September 2018 CDRL PM - 14

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3 CONTENTS Foreword v ________________________________ ___ ________________________________ ________________________________ ___ vii Preface ________________________________ ________________________________ ________________________ ix Acknowledgements Mission Overview ________________________________ _____________________ 1 - 1 1. 2 ________________________________ ________ - 1 2. GOES Spacecraft Configuration Advanced Baseline Imager ________________________________ ___________ 3 - 1 3. 4. ________________________________ _______ 4 - 1 Geostationary Lightning Mapper Situ Suite ________________________________ _______ - 5 - 1 5. Space Environment In Magnetometer ________________________________ _______________________ 6 - 1 6. 7. - ray Irradiance Sensors Instrument _________ 7 - 1 The Extreme Ultraviolet and X ________________________________ _______________ Solar Ultraviolet Imager 8 - 1 8. 9 1 9. GOES - R Communications Subsystem ________________________________ ___ - 10. ___________________________ 10 - 1 Command and Data Handling Subsystem Electrical Power Subsystem ________________________________ _________ 11 - 9 11. 12. Guidance Navigation & Control ________________________________ _____ 12 - 1 ______________ 13. Propulsion Subsystem ________________________________ 13 - 1 14 . Thermal Control Subsystem ________________________________ _________ 14 - 1 15 15. Mechanisms ________________________________ _______________________ 1 - 16 16. ________________________________ ________ - 1 Ground System Architecture Spacecraft Mission Phases ________________________________ __________ 17 - 1 17. ________________________________ On - Orbit Mission Operations 18. ________ 18 - 1 ___ 19. Technical Performance Summary ________________________________ 19 - 1 20. Acronyms ________________________________ _________________________ 20 - 1 iii

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5 Foreword – R Series (GOES - R) is the next The Geostationary Operational Environmental Satellite generation of U.S geostationary weather satellites and is a key element in National Oceanic and Atmospheric Administration (NOAA) operations. GO gery and advanced weather ES weather ima products have been a continuous and reliable stream of environmental information used to support weather forecasting, severe storm tracking, and meteorological research. Evolutionary improvements in the geostationary satellite system since 1974 (i.e., since the first Synchronous Meteorological Satellite, SMS - 1) have been responsible for making the GOES system a mainstay of weather forecasts and environmental monitoring. The GOES - R series (GOES R, S, T, and U) represe nts the first major technological advancement in geostationary observations since 1994 and will extend the availability of the GOES system through 2036. The GOES - R series will provide critical atmospheric, hydrologic, oceanic, climatic, solar and space dat a, significantly improving the detection and observation of environmental phenomena that directly affect public safety, protection of property, and our nation’s economic health and prosperity. Designed to operate in geosynchronous orbit, 35,786 km (22,236 statute miles) above the equator, thereby remaining stationary relative to the Earth’s surface, the advanced GOES - R contiguous United States, neighboring environs of the series spacecraft will continuously view the Pacific and Atlantic Oceans, and Central and South America. The R series spacecraft bus GOES - axis stabilized and designed for 10 years of on - orbit operation preceded by up to five is three - - orbit storage. Two GOES satellites remain operational at all times while an on - orbit years of on intained to permit rapid recovery from a failure of either of the operational satellites. spare is ma GOES - R series spacecraft The Advanced Baseline Imager (ABI) is the primary instrument on the with 16 different for imaging Earth’s weather, oceans and environment. ABI views the Earth spectral bands (compared to five on the previous GOES series), including two visible channels, four near - infrared channels, and ten infrared channels. ABI’s data will enable meteorologists to pinpoint and track developing storms in much greater detail. Geostationary Lightning Mapper (GLM) is the first operational lightning mapper flown in The geostationary orbit. GLM detects and maps total lightning (in cloud and cloud - to - ground) activity - continuously over the Americas and adjacent ocean regions. Used in combination with radar, data from the ABI instrument, and surface observations, GLM data has great potential to increase lead time for severe thunderstorm and tornado warnings. GOES - R series spacecraft also carry a suite of instruments to significantly improve detection of approaching space weather hazards. The satellites provide advanced imaging of the sun and detection of solar eruptions for earlier warning of disruption to power utilities and communication ellites also more accurately monitor energetic particles and the and navigation systems. The sat magnetic field variations that are associated with space weather for better assessment of radiation hazards and mitigation of damage to orbiting satellites, communications, and power grids. v

6 round support is critical to the GOES - R series G mission. To support the large increase in spatial, spectral, and temporal resolution of the ABI and other instruments, the raw data rate increase d to 75Mbps, over 30 times the previous rate. NOAA has developed a state - of - the - art ground system to receive data from the GOES - R series spacecraft and generate real - time data products. The from two primary locations: the National Satellite Operations Facility s ground system operate (NSOF) in Suitland, Maryland, and the Wallops Command and Data Acquisition Center (WCDAS) s as the at Wallops, Virginia. A third operations facility in Fairmont, West Virginia, serve Co nsolidated Backup (CBU) in case of a systems or communications failure at either or both NSOF and WCDAS. Those desiring further information about the GOES system should contact the NOAA National Environmental Satellite, Data and Information Service (NESDIS ) and/or search the following internet addresses: https://www.goes - r.gov/ http://www.noaa.gov/ https://www.nesdis.noaa.gov/ https://www.weather.gov/ https://www.ncei.noaa.gov/ https://www.swpc.noaa.gov/ vi

7 Preface To further enhance the utility of the GOES system, this reference presents a summary and technical overview of the GOES - R series system, its satellites, subsystems, sensor suite, and associated ground communication and data handling subsystems. The referen ce is intended to serve as a convenient and comprehensive technical reference for people working on or associated with the GOES - R series mission as well as general information suitable for public distribution. Sufficient technical information and performan ce data are presented to enable the reader to understand the importance of the - R series mission, the system’s capabilities, GOES and how it meets the needs of the users. d using Certain performance data presented herein, e.g., instrument performance, were predicte - launch analyses and ground testing. As the satellites undergo on - pre orbit operations and actual data are obtained, such technical information in this reference may not necessarily reflect current capabilities. Furthermore, this reference is not me ant to be a technical specification with absolute worst case performance numbers but rather a general document which informs the reader of nominal and typical GOES system performance and operational capabilities. program GOES is a collaborative development and acquisition effort between The - R series NOAA and the National Aeronautics and Space Administration (NASA). Program activities occur ffices at Goddard Space Flight Center in Greenbelt, - located p rogram and p roject o at the co Maryland. The GOES - R series program collaborates with industry partners across the United R States to fulfill the GOES - series m ission. Lockheed Martin is the principal s pace system , with associ ate contractors providing individual instruments to Lockheed for contractor g integration, and Harris is the prime has numerous round system contractor. Each system contractor. supporting sub vii

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9 Acknowledgements This Data Book had major contributions from the following authors at Lockheed Martin (LM): lead), Jim Chapel, Anthony Dimercurio Jr., Bonnie Birckenstaedt, Clemens Tillier, Alexis Benz ( William Nilsson III, Jesus Colon, Robert Dence, David Hansen Jr, and Pamela Camp bell. Our special thanks for their efforts. We would also like to thank the many instrument authors and NASA/NOAA personnel that , including Dennis Chesters, Ruth Cholvibul, provided inputs and reviews of the data book content yne, Lauren Gaches, Craig Keeler, Michael Kimberling, William Jonathan Chow, Gustave Come Lebair, Marc Rafal, Chris Rollins, and Michelle Smith. We would also like to thank Michael Ames Ashton for the front and back cover artwork. Armstrong for Lastly, we would like to acknowledge for er leadership in integrating the first draft of this book, and Diana Schuler h taking over as book manager, providing the final review and edits, and seeing the data book to completion and publishing. ix

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11 1. Mission Overview Mission Goals system are to: The goals for the Geostationary Operational Environmental Satellite (GOES) Maintain continuous and  reliable operational, environmental, and storm warning ems to protect life and property. syst .  Monitor the Ea rth’s surface and space environmental conditions  Introduce improved atmospheric and oceanic observations and data dissemination capabilities . To address these goals, the National Weather Service (NWS), NESDIS, and NOAA established GOES - R series mission requirements for the 21st century that are the basis for the design of the system and its capabilities. The GOES - R series system functions to accomplish an environmental mission serving the needs of operational meteorological, space environmental and rese arch users. Figure 1 - 1 illustrates the GOES - R series satellite and its instruments. S R Series e atellit Figure 1 - 1. GOES - 1 - 1

12 GOES System GOES - R series of spacecraft perform s three major To accomplish the GOES mission, the functions: • Environmental sensing : a cquisition, processing and dissemination of imaging data, ment of near - Earth space weather. space environment monitoring data, and measure • Data collection : i nterrogation and reception of data from Earth surface - based data collection platforms (DCPs) and relay of such data to the NESDIS command and data acquisition stations . • : GOES Rebroadcast service (GRB) and Product Distribution and Access Data broadcast (PDA) of environmental sensor data. The relay of distress signals fr om aircraft or marine vessels to the S earch and R escue S atellite - A ided T racking system (SARSAT). The continuous relay of weather facsimile to the Emergency Managers Weather Information Network ( EMWIN) and other meteorological data to other users and the re lay of . anagers m emergency weather information to c ivil e mergency The three major mission functions are supported or performed by the following components of GOES - R series payloads: the Environmental sensing: • Advanced Baseline Imager (ABI) • Geostationary Lightning Mapper (GLM) • Extreme Ultraviolet and X - ray Irradiance Sensors (EXIS) • Magnetometer (MAG) • Space Environment In - Situ Suite (SEISS) Solar Ultraviolet Imager (SUVI) • Data collection: • Data collection system (DCS) • Search and Rescue (SAR) Data broadcast: • Advanced Weather Interactive Processing System (AWIPS) • Comprehensive Large Array - Data Stewardship System (CLASS) • Emergency Managers Weather Information Network (EMWIN) • GOES Rebroadcast (GRB) • High Rate In formation Transmission (HRIT) Product Distribution and Access (PDA) • • Search and Rescue Satellite Aided Tracking (SARSAT) System 1 - 2

13 A general overview of the GOES - R series system (including space and ground elements) is shown in Figure 1 - 2. and Ground System System Series Overview Figure 1 - 2. GOES - R ystem S Space GOES - R series of spacecraft are the prime observational platforms for covering dynamic The - Earth space environment for the 21st century. These advanced weather events and the near spacecraft enhance the capability of the GOES system to continuously observe and measure mete orological phenomena in real time, providing the meteorological community and atmospheric scientists of the western hemisphere with greatly improved observational and measurement data. The key advancements realized by the GOES - R series are related to the instrument payloads and spacecraft. The advanced instruments drive improvements in the overall system, such as the processing, generation and distribution of data products. Advances in the spacecraft improve the overall operations of the satellites and imp rovements in the instruments provide greater temporal term weather resolution. These enhanced operational services improve support for short - forecasting and space environment monitoring as well as atmospheric sciences research and weather prediction models, and environmental sensor design. development for numerical 1 - 3

14 Observational Platform orbit - spacecraft bus is three - GOES - R series The axis stabilized and designed for 10 years of on - orbit storage. Two GOES satelli tes remain operation preceded by up to five years of on - operational at all times while an on orbit spare is maintained to permit rapid recovery from a failure of either of the operational satellites. The - R series spacecraft design enables the sensors GOES y image clouds and lightning, and monitor the Earth’s to stare at the Earth and thus more frequentl surface temperature and water vapor fields. Thus, the evolution of atmospheric phenomena can - time coverage of short be followed, ensuring real lived, dynamic events, especially severe local - storms and t ropical cyclones. These are meteorological events that directly affect public safety, protection of property, and, ultimately, economic health and development. Various design features of the GOES - R series spacecraft enable high volume, high quality data to be generated for the weather community. There are two important capabilities. The first is flexible scan control — a - capability that allows small area coverage for improved short term weather forecasts over local — and simultaneous, independent imaging. The second is precision on areas orbit station keeping, - coupled with three - axis stabilization, which ensures a steady observational platform for the mission sensors. The GOES - R series will permit a vast reduction over legacy GOES missions in instrument data co es. Satellite “operate - through” performance llection outages due to satellite maintenance activiti for routine housekeeping such as momentum management and e ast/ w est station keeping maneuvers precludes the need to schedule daily or monthly outage periods. C oupled with the GOES outage goal is less R series - enhanced ABI capabilities of imaging through eclipse, the than 3 hours per year compared to the hundreds of hours per year of the GOES - I/M series. Other solation for the Earth - pointed optical notable performance include: vibration i enhancements high bench speed spacecraft - to - instrument interfaces designed to maximize science data , - collection, and an improved attitude control and image navigation capability. Geographic Coverage on station 35,786 km (22,236 statute miles) above the equator and The GOES spacecraft, - stationary relative to the Earth’s surface, can view the contiguous 48 states , Alaska, the central and eastern Pacific Ocean central and western Atlantic Ocean areas , and the South American , continent. Pacific coverage includes the Hawaiian Islands and the Gulf of Alaska. Because the Atlantic and Pacific basins strongly influence the weather affecting the United States, coverage is provided by two GOES spacecraft. iometric coverage and communications range) of the two spacecraft The combined footprint (rad encompasses Earth’s full disk about the meridian approximately in the center of the contiguous nd to United States. Circles of observational limits centered at a spacecraft’s suborbital point exte 60° north/south latitudes. The radiometric footprints are determined by the limit from the beyond suborbital point, beyond which interpretation of cloud data becomes unreliable. At least one line GOES spacecraft is always within of - sight view of Eart h - based terminals and stations. The - command and data acquisition station at WCDAS has a line - of - sight to both spacecraft to uplink commands and receive downlinked data from each simultaneously, with CBU facility ready in the event of a systems or communica tions failure at WCDAS. The GOES - R series maintains the two - satellite system implemented by the previous GOES satellites. However, the locations of the ⁰ W and 137 W. The latter is a shift from previous operational GOES - R series satellites will be 75 ⁰ 1 - 4

15 ⁰ W in order to eliminate conflicts with other satellite systems. The at 135 - R series GOES GOES Figure 1 operational lifetime extends through December 2036. 3 illustrates the geographic - coverage of the - R series GOES . constellation Figure 1 - 3. Geographic Coverage of the GOES - R Series Constellation Advanced Baseline Imager (ABI) The Advanced Baseline Imager , manufactured by the Harris Corporation , is the primary GOES - instrument on the R series for imaging Earth’s weather, oceans and environment . It is a channel passive imaging radiometer designed to observe the Western Hemisphere and multi - provide variable area imagery and radiometric information of Earth’s surface, atmosphere and ABI views the Earth with 16 different spectral bands cloud cover. (compared to five on the previous GOES series), including two visible channels, four near infrared channels, and ten infrared - channels. These different channels (wavelengths) are used by models and tools to indicate various elements on the earth’s surface or in the atmosphere, such as trees, water, clouds, . ABI provides four times the resolution and five times faster temporal coverage moisture or smoke than the prior generation of GOES. ABI has three scan modes. The default mode (known as Scan Mode 3, or Flex Mode) concurrently Contiguous takes a full disk (Western Hemisphere) image every 15 minutes, an image of the U.S. (CONUS) mesoscale images of areas where storm every five minutes, two smaller, more detailed activity is present, every 60 seconds or one mesoscale image every 30 seconds . Similarly, a second flexible mode (Scan Mode 6) has been added that is the same as Scan mode 3 in all regards except that the full disk image is taken every 10 minutes. The ABI can also operate in continuous full d isk mode (known as Scan Mode 4) , providing uninterrupted scans of the full disk - 4A and Figure 1 illustrate 4B every 5 minutes. All ABI bands have on - orbit calibration. Figure 1 - 1 - 5

16 the ABI spatia l resolution and coverage area East and and default meso locations for the GOES - West locations, respectively. - S GOE Figure - Coverage Area 4A. ABI Spatial Resolution, 1 - , and Collection Times with GOES East Operational View and default Meso locations 1 - 6

17 1 min Mesos 5 min CONUS 15 min Full Disk Coverage Area , and Collection Times with GOES Figure - 1 - 4 B . ABI Spatial Resolution, View and default Meso locations West Operational ABI is used for a wide range of applications related to weather, oceans, land, climate and The hazards (fires, volcanoes, floods, hur ricanes and storms that spawn tornadoes). It can track and monitor cloud formation, atmospheric motion, convection, land surface temperature, ocean dynamics, flow of water, fire, smoke, volcanic ash plumes, aerosols and air quality, and vegetative ABI’s data enables meteorologists to pinpoint and track developing storms in much greater health. detail. Future products will also help the aviation industry with aircraft icing threat detection and turbulent flight condition predictions. Benefits from the ABI in - related flight clude improved tropical cyclone forecasts, fewer weather delays and airline incidences with volcanic plumes, increased efficiency in irrigated water usage 7 - 1

18 ropical storm or in agriculture, and higher protection rates for recreational boats in the event of a t hurricane. Geostationary Lightning Mapper (GLM) manufactured by Lockheed Martin, is the first operational The Geostationary Lightning Mapper, - cloud and lightning mapper flown in geostationary orbit. GLM measures total lightning, both in c loud - to - ground, to aid in forecasting intensifying storms and severe weather events. GLM is unique both in how it operates and in the information it collects. The instrument is sensitive to the in uniform total - - cloud lightning that is most dominant in severe thunderstor ms and provides nearly lightning coverage over the region of interest. GLM is a single - channel, near - infrared optical transient detector that can detect the momentary changes in an optical scene, indicating the presence of lightning. GLM detects and maps total lightning activity throughout the day and night over the Americas and adjacent ocean regions with near uniform spatial resolution of approximately 10 kilometers. The instrument collects - information such as the location, brightness an d extent of lightning discharges to identify intensifying storms, which are often accompanied by increased total lightning activity. Trends in total lightning that will be available with GLM have the promise of providing critical which will allow them to focus on developing severe storms much earlier information to forecasters than they can currently, and before these storms produce damaging winds, hail or even arly tornadoes. Such storms often exhibit a significant increase in total lightning activity, particul in cloud lightning, often many minutes before radar detects the potential for severe weather. Used - in combination with radar, data from ABI, and surface observations, GLM data has great potential to increase lead time for severe thunderstorm and torna do warnings. Data from the instrument - term database to track decadal changes in lightning activity. will also be used to produce a long This is important due to lightning’s role in maintaining the electrical balance between Earth and its atmosphere and pot ential changes in extreme weather and severe storms under a changing climate. Space Weather Instr uments The GOES - R series of satellites host a suite of instruments that provide significantly improved detection of approaching space weather hazards. Changes in “space weather” can affect the operational reliability of communication and navigation systems, disrupt power lines, damage altitude satellite electrical systems, and may cause radiation damage to orbiting satellites, high - aircraft and the International Space Station , as well as harming astronauts . Two sun - pointing instruments measure solar ultraviolet light and X - rays. The Solar Ultraviolet Imager (SUVI) observes and characterizes complex active regions of the sun, and provides six - channel movies of so lar flares and the eruptions of solar filaments which may give rise to coronal mass ejections. The Extreme Ultraviolet and X - ray Irradiance Sensor (EXIS) detects solar flares and monitors solar irradiance that impacts the upper atmosphere. The satellites also carry two instruments that measure the space environment. The Space Situ Suite (SEISS) monitors proton, electron and heavy ion fluxes in the - Environment In magnetosphere. The Magnetometer (MAG) measures the magnetic field in the outer portion of agnetosphere. the m 1 - 8

19 The GOES R SUVI and EXIS i nstruments provide improved imaging of the sun and detection of - solar eruptions, while SEISS and MAG more accurately monitor, respectively, energetic particles and the magnetic field variations that are associated with space weather. Together, observations from these instruments will enable NOAA’s Space Weather Prediction Center to significantly improve space weather forecasts and provide early warning of possible impacts to Earth’s space environment and potentially disruptive events on the ground. Other Data Services Emergency radio beacons are carried on ships and planes to signal distress to satellites orbiting ues the legacy Geostationary Search and Rescue contin overhead. The GOES - R series the SARSAT system onboard NOAA’s GOES satellites which has (GEOSAR) function of contributed to the rescue of thousands of individuals in distress. The SARSAT transponder was modified slightly for the GOES - R series by operating with a lower uplink power (32 dBm), enabling - the s atellites to detect weaker signal beacons. The SARSAT transponder onboard GOES R series satellites provides the capability to immediately detect distress signals from emergency beacons and relay them to ground stations , called Local User Terminals. In turn , this signal is routed to a SARSAT Mission Control Center and then sent to a Rescue Coordination Center which dispatches a search and rescue team to the location of the distress. Ground System - ground system has a much greater product dis tribution capability over the The GOES R series legacy missions. To support the large increase in spatial, spectral, and temporal resolution of the ABI and other instruments, the raw data rate has increased to 75Mbps, over 30 times the previous rate. R series data volu me drives a large increase in processing requirements for product - GOES generation and for distribution of the products to users. nd system receives the Level 0 (L0) raw data from GOES - R series spacecraft and The grou products and makes these products available to generates Level 1b (L1b) and Level 2+ (L2+) users in a timely manner . Level 0 data is the unprocessed instrument data at full resolution. L1b processed with radiometric and geometric correction applied to produce data is the L0 data al units. L2+ data are derived environmental variables parameters in physic generated from L1b data along with other ancillary source data, such as from the National Weather Prediction forecast model output data. Table 1 - 1 shows a listing of the L1b and L2+ GOES - R series scien ce products by instrument. The GOES - R ground system (GS), developed by Harris Corporation, consists of several functional components. Mission Management (MM) provides the satellite operations (monitor and control) function of the GS (via the Harris OS/COMET ® software). Enterprise Management (EM) is distributed over all GS components and locations and provides for the ability to monitor the complete enterprise, as well as control the operations not directly associated with satellite operations. Produc t Generation (PG) provides the L1b and L2+ product generation function. Product Distribution (PD) functionality provides for direct distribution of product data to the GOES - Advanced Weather Interactive Processing System (AWIPS R Access Subsystem (GAS), NWS ) (an electronic library of NOAA environmental data) and CLASS . Long term archive and access 1 - 9

20 services to retrospective users of - R series data will be provided by CLASS, which is GOES GOES - R series GS, but is part of considered an external interface to the the NOAA infrastructure. 1 - 1 Table GOES - R L1b and L2+ Science Products . Instrument Level1b Products ABI Level 2+ Products Cloud and Moisture Imagery (CMI)  ABI  Radiances and Sectorized CMI (KPP)  Aerosol Detection (Smoke & Dust)  Aerosol Optical Depth (AOD)  Clear Sky Mask  Cloud Optical Depth  Cloud Particle Size Distribution Cloud Top Height  SEISS Energetic Heavy Ions  Cloud Top Phase  - Magnetospheric e  /p+: Cloud Top Pressure  Low Energy Cloud Top Temperature  -  /p+: Magnetospheric e Derived Motion Winds  High Energy  Derived Stability Indices  Downward S/W Radiation: Surface ar & Galactic l So  n Fire/Hot Spot Characterizatio  Protons  Hurricane Intensity Estimation EXIS Solar Flux: EUV   Land Surface Temperature  Legacy Vertical Moisture Profile  Legacy Vertical Temperature Profile  ray - Solar Flux: X  Rainfall Rate/QPE Irradiance Reflected S/W Radiation: TOA   Sea Surface Temperature  Snow Cover  Total Precipitable Water  on and Height Volcanic Ash: Detecti SUVI Solar EUV Imagery  MAG Geomagnetic Field  2+ Products GLM Level  Lightning: Events, Groups, & Flashes Network Architecture The GS operates from two primary locations: NSOF in Suitland, Maryland, and WCDAS in as the CBU in case s Wallops, Virginia. A third operations facility in Fairmont, West Virginia, serve of a systems or communications failure at either or both NSOF and WCDAS. The satellites are commanded throughout their mission lifetime from the NSOF with the ground station radio s needed). The engineering telemetry frequency (RF) interface located at WCDAS (or the CBU, a - 10 1

21 streams are received by both the WCDAS and CBU, then ground relayed to the NSOF for processing and monitoring at all locations. In addition to the redundant operational environments at each ground location, t he GS incl udes separate development and integration and test (I&T) environments for the purposes of ongoing - R mission. Two on development and I&T throughout the GOES site Development Environments - (DE) (one at NSOF and one at WCDAS) and three Integration and Test Env ironments (ITE) (one at NSOF and two at WCDAS) are provided by the GS for software maintenance. Local DE and ITE workstations are provided at WCDAS and NSOF. In addition, DE and ITE workstations are E and ITE functions. The CBU provided at NSOF to accommodate remote use of the WCDAS D does not provide a DE or ITE and relies on WCDAS for software maintenance. Raw instrument data (L0) is received at WCDAS. It is then processed by the Product Generation products. (PG) function at WCDAS to create and some Level 2+ ( L2+ ) These L1b Level 1b (L1b) and L2+ products are then rebroadcast through the GRB transponder. The GRB data are then received at NSOF where the rest of the L2+ products are created. Ancillary data used in generating the L2+ products are ingested fro m the Ancillary Data Relay System (ADRS). Applicable products are directly distributed to 1) NWS AWIPS where key NWS Weather Forecast Offices (WFO) and other AWIPS users get their data, and 2) GAS for use by the Environmental Satellite Processing Center (E SPC) and other GOES data users. - ion and access data products will be available using new product distribut R series GOES ground system architecture technologies. An overview of the - R series GOES as well as more e found in the Ground System Architecture tion can b of section information about product distribu this book. GOES Rebroadcast (GRB) GRB provides the primary relay of full resolution, calibrated, near real - time direct broadcast L1b - VARiable (GVAR) data from each instrument and L2 data from GLM. GRB replaces the GOES service . GRB contains the ABI, GLM, , and is a significant increase in capability from that service space environment, and solar data which drive data flow in the NOAA space and Earth environment research and operational framework. GRB us es two digital streams, each at 15.5 Mbps, compared to the GVAR standard of a single 2.11 Mbps stream. A dual polarization approach is used to accommodate the 31 Mbps data rate within a frequency bandwidth of 9.8 or 10.9 MHz per polarization, using a stand ard downlink a full disk image in either five modulation at 1686.6 MHz (L - band). GRB is able to deliver , ten, or fifteen minutes, depending on mode, compared to GVAR’s thirty minutes. The GRB processed instrument data is packetized compliant with Consultative Committee for Space Data Systems (CCSDS) S tandard 133.0 - B - l and utilizes lossless data compression to fit within allocated bandwidth. Data blocking and accompanying header metadata are used to loss due to link errors and allow for user verification of data integrity. data minimize the risk of 1 - 11

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23 GOES Spacecraft Configuration 2. - R series satellite is based on the Lockheed Martin A2100 bus. The The GOES satellite with the magnetometer, solar array, X - band antenna, and antenna wing fully dep loyed (on - orbit ed in this fig configuration) ure as it is can be viewed in Figure 2 - 1. ( Note that SEISS is not identifi - isometr ic view. ) not viewable in this The 6,280 pound, three axis stabilized GOES - R spacecraft orbit life of 15 years. The spacecraft bus provides mechanical support was designed for an on - nt payloads, communications payloads and other bus and alignment of the various instrume components. Figure - 1 : Fully Deployed GOES 2 R Satellite - The satellite houses three classifications of instruments: Nadir - pointing, Solar - pointing, and In - Situ (near envi ronment) . The Nadir pointing instruments include the ABI and the GLM. These are mounted on a highly stable, precision earth instruments pointed platform, and are dynamically - isolated f The Solar - pointing instruments, which include the EXIS, rom the rest of the spacecraft. SUVI, and the Unique Payload Services (UPS) are mounted on a Sun Pointing Platform (SPP) housed on a solar array yoke. The SPP provides a stable platform that tracks the seasonal and daily movement of the sun relative to the spacecraft. The In - Situ instruments include the SEISS and th e Magnetometer. The Magnetometer is mounted on a boom that deploys on ce the spacecraft reaches orbit. The boom provides relative magnetic isolation for this instrument. Each - chapters R Series instruments are described in detail in subsequent . of the GOES 2 - 1

24 spacecraft overview system block diagram, depicting each of the sp acecraft subsystems, is A - shown in Figure 2 More details 2. The following paragraphs briefly describe each subsystem. can be found in subsequent chapters . Figure 2 2: GOES - R Series Satellite System Block Diagram - Communications Subsystem The communications subsystem provides the antennas and transponders used for communicating with the ground system and for data relay services. The spacecraft uses S - band, Band and UHF links to provide the command, telemetry, and track - Band, X t also ing functions. I L - s services (Unique Payload Services) which: support the communication Data provides a set of Collection Platforms ; relay High Rate Information Transmission (HRIT) and imaging data between Earth terminal s and relays the Emergency Managers Weather Inf ormation Network ( EMWIN ) (SAR) broadcast on the HRIT/EMWIN; detects and communicates Search and Rescue distress signals; and rebroadcast s processed GOES - R sensor Emergency Locator Transmitter data via the GOES Re - broadcast (GRB). Command and Data Handling Subsystem to process, needed The command and data handling (C&DH) system includes the hardware route, and deliver commands, telemetry, and instrument data on board the spacecraft, including the processing resources needed for the flight software to functi on. The C&DH system includes 2 - 2

25 - board computers, command and telemetry processors, 6 component types, including the on sun pointing platform interface , pyro relay assemblies, and a remote interface units, electronics command decryption unit assembly. Electrical Power Subsystem The primary function of the electrical power subsystem (EPS) is to provide power to operate the a Scalable Power Regulation Unit (SPRU) satellite for 15 years in geostationary orbit. It consists of to control the flow of power fr om the solar arrays and batteries onto the busses, two batteries for energy storage, a solar array for power generation, 70 volt and 28 volt power busses to distribute . power to the loads, and two Fuse Board Assemblies to provide load overcurrent protection Guidance, Navigation and Control Subsystem The guidance, navigation and control (GN&C) subsystem provides guidance, navigation, and attitude & articulation co ntrol for the GOES R spacecraft . Inertial Measurement Unit s (IMU) and - itude determination, a global positioning system (GPS) receiver provides star trackers provide att orbit determination, and attitude control is provided using reaction wheels and the propulsion system. The GN&C subsystem is also responsible for controlling th e spacecraft’s gimbals for the X - B and antenna, solar panel, and sun - pointing platform. Propulsion Subsystem The propulsion subsystem provides the impulse to perform the thrusting maneuvers required during each mission phase, from launch and orbit raising through end of life man euvers. It consists t rocket engine low thrus monopropellant engine (LAE), sixteen of a single liquid apogee assemblies (REA) , propellant thrusters, 4 Arcjets, eight monopropellant REAs, two hydrazine bi - and associated tanks, valves, and regulators. Therma l Control Subsystem The thermal control subsystem consists of all spacecraft thermal elements and coatings required to control the onboard temperature of spacecraft elements. The subsystem uses active and inishes, insulators, multi - layer insulation (MLI), passive thermal control techniques, and includes f heaters, thermostats and other heater controls, temperature sensors, heat transferring apparatus, radiators, sun shields and radiation shields to accomplish its function. Mechanical Subsystem l subsystem provides the mechanical interface and The Mechanica support for structural spacecraft components as well as the mechanisms that must be stowed and restrained for launch and later deployment . The structure possesses sufficient strength, stiffness, and damping for subsystems and payloads to survive the load conditions that exist within the envelope of the mission as well as test requirements. The mechanisms include hinges, retention and release loy and point the Solar (R&R) hardware, motors, gimbals and dampers. These are used to dep Wing Subsystem and the antennas. Launch and Ascent Configuration During launch and ascent, the Sun Pointing Subsystem (SPS) , as well as the Solar Array Wing Assembly (SAWA) , is folded against the +Y side of the spacecraft bus (as seen in a Body Reference Frame (BRF)) and each are held in place with restraint and release (R&R) mechanisms. The stowed configuration is shown in Figure 2 - 3 . The SPS utilizes six R&Rs to the yoke/frame assembly. and two R&Rs to support The un Pointing Platform (SPP) support the S 2 - 3

26 When in the stowed configuration, instruments on the SAWA is also supported with six R&Rs. SPP face outboard and solar cells on the outboard SAWA panel face outward. SPS SAWA - 3 . Launch and Ascent Configuration in Body Reference Frame Figure 2 Orbit Raising Configuration During transfer orbit operations, the solar array is in the first stage deployed configuration. In the first stage deployed configuration, the inboard SAWA panel (panel 1) and the SPS remain stowed ls 2 through 5 are fully deployed and in against the +Y side of the satellite. The solar array pane the same plane as panel 1. All panel solar cells face outward (+Y direction). The inboard panel utilizes two R&Rs to attach to the spacecraft bus in this configuration. The solar cells on the SAWA rection shown in - Figure 2 panels face the di 4 . - 4 2

27 st Figure 2 4 . Transfer Orbit Configuration (1 - Stage Deployment) - On Orbit Configuration is achieved, the , as is commanded to fully deploy Af ter G eosynchronous Earth Orbit ( GEO ) SPS d maintains the SPP and the solar array in a sun shown in Figure 2 - 5 . The SPS articulates an pointing orientation. Occasionally the SPP may be slewed 16 degrees off pointing from the sun to allow for calibration of the science instrument on the SPP. 2 - 5

28 nd 2 Orbit Solar Array Configuration (2 Final Stage Deployment) Figure / - 5 . On - 6 - 2

29 3. Advanced Baseline Imager The Advanced Baseline Imager, manufactured by Harris Corporation, is a multi - spectral imaging radiometer for the R series of satellites . It provides nearly continuous imagery of the GOES - Western Hemisphere from geostationary orbit for weather prediction and other Earth science μ applications. ABI measures Earth’s radiance in 16 spectral channels ranging from visible (0.47 m) to longwave infrared (13.3 μ m). A view of the ABI is shown in Figure 3 - 1. Figure 3 - ABI 1. orth n est (EW) and one outh s ABI scans the E arth via two orthogonal scan mirrors: one e ast - w - called a (NS). The EW mirror scans the Earth at 1.4° (optical) per second; a single EW scan is swath. The NS mirror is then stepped to a new location to begin another EW swath. In this manner, ABI can scan the full Earth image in five minutes (known as Mod e 4 – Continuous Full minute images of Disk mode ) or alternatively scan the full Earth in 15 minutes w ith interleaved 5 - or one 30 1000 x 1000 km the Contiguous United States (CONUS) and two 60 - second ) ( second - mesoscale images . This mode is known as Mode 3, or Flex mode . Mesoscale images are helpful A third scan mode (known and Mode 6) has been added that is the for viewing storm activity. same as Mode 3 except that the full Earth scan is performed in 10 minutes instead of 15 minutes. xis The scanning motion of the two scan mirrors direct Earth’s radiance into a four - mirror, off - a telescope that converges the energy into the aft optics. There the energy is separated into three bands, ongwave infrared visible and near infrared (VNIR), midwave infrared (MWIR), and l h of the 16 spectral bands. (LWIR), and ultimately sensed by a unique set of detectors for eac The nadir spatial resolution of the collected imagery ranges from 0.5 to 1 km in the visible channels and 1 to 2km in the infrared channels. A summary of the characteristics for each 3 - 1

30 - 1 . A fur ther description of the key performance spectral channel is shown i n Table 3 is shown in - 2 . requirements Table 3 3, along Representative Instrument radiometric noise and dynamic range are shown in Table 3 - The with typical applications for each ABI bands. The nominal center wavelength is shown. exact wavelength center can vary slightly by flight model (FM). Each ABI FM was designed to collect radiometric data both for the earth, and for the needed reference sources to calibrate the hot blackbody source flying on - earth scenes. For ABI, this includes cold space (~4 K) and a board ABI. Although ABI observes cold space to help remove its own warm telescope background, earth scenes are warmer. The brightness temperature of the internal blackbody emitter used for calibration is warmer than that of t ypical earth scenes, and can be configured. Achieved radiometric noise performance varies by detector and FM, but is often better than these levels. The maximum range capability for each FM to observe radiance (expressed in ten exceeds the physical brightness temperature ranges for the brightness temperature units) of earth. In the case of fires though, the need to report brightness temperature of the fire was balanced against the need to monitor cold clouds. Typical applications of these radiance ts are shown in Table 3 3 as well. measuremen - The Radiances product produced by the Ground Segment (GS) from the ABI observations has 2 2 - 1 sr cm ) in bands greater than 3 um. sr um) in bands less than 3 um and mW/(m units of W/(m nd Moisture Imagery product that converts the The Ground Segment produces the Cloud a radiance output for bands greater than 3 um to Brightness temperatures in units of Kelvin, so maximum ranges are often listed in K. The product range, configurable in the GS, was required of interest and start at about 180 K. The products, ranges, and associated to cover earth scenes - R Product Users Guide (PUG), bit depths are detailed in a separate document called the GOES https://www.goes - r.gov . which can be found at Table - 1 . ABI Channel Characteristics 3 Center NS IFOV Channel Pixel EW ASD EW IFOV Wavelength / Band (μrad) (μrad) (μrad) Size*(km) (μm) 1 0.47 22.9 22.9 22 1 10.5 0.64 12.4 11 0.5 2 22.9 3 0.865 22.9 22 1 2 44 42.0 4 1.378 51.5 VNIR 5 1.61 22.9 22.9 22 1 51.5 42.0 2.25 44 2 6 2 7 3.90 51.5 47.7 44 8 6.185 51.5 47.7 44 2 6.95 51.5 9 47.7 44 2 MWIR 44 2 10 7.34 51.5 47.7 44 11 51.5 47.7 2 8.5 12 9.61 51.5 47.7 44 2 13 10.35 34.3 38.1 44 2 38.1 11.2 34.3 44 2 14 LWIR 38.1 15 12.3 34.3 44 2 2 16 13.3 34.3 38.1 44 resampled image pixel spacing at nadir - *Pixel size refers to post - 2 3

31 3 Table 2. ABI Channel Performance Requirements Summary - Parameter Performance VNIR for Calibration Accuracy ±3% (±4% for 1.378 μm) 100% albedo scene MWIR and LWIR ±1 K for 300 K scene

32 - 3. ABI Radiometric Precision and Dynamic Range 3 able T Requirements Center Channel SNR / Dynamic Wavelength Typical Applications NEdT * / Band Range (μm) Daytime aerosol over land, coastal water 0 – 100% 1 0.47 300:1 mapping Albedo 0 – 100% 2 0.64 Daytime clouds, fog, solar flux, and winds 300:1 Albedo Daytime vegetation/burn scar and aerosol 100% – 0 0.865 3 300:1 Albedo over water, winds 100% – 0 VNIR 1.378 4 300:1 Daytime cirrus clouds Albedo 0 – 100% top phase and particle - Daytime cloud 300:1 5 1.61 size, snow Albedo 100% 0 properties, particle Daytime land/cloud – 6 2.25 300:1 Albedo size, vegetation, snow 400 K 0.1 K Surface and cloud, fog at night, fire, winds – 4 7 3.90 High - level atmospheric water vapor, 4 – 6.185 300 K 8 0.1 K winds, rainfall Mid - level atmospheric water vapor, winds, 4 9 300 K 6.95 0.1 K – and rainfall MWIR 10 7.34 0.1 K 4 – 320 K Lower - level water vapor and winds Total water for stability, cloud phase, 0.1 K 4 – 8.5 330 K 11 dust, and rainfall 300 K – 4 12 Total ozone, turbulence, and winds 9.61 0.1 K 10.35 Surface and cloud properties 13 0.1 K 4 – 330 K Sea surface temperatures, rainfall and – 330 K 14 0.1 K 4 11.2 cloud properties LWIR Total water, ash, sea surface temp., and 12.3 0.1 K 15 – 330 K 4 cloud properties 4 16 13.3 0.3 K – 305 K Atm. temps and cloud heights @ 100% albedo; NEdT at 300 K S NR * 3 - 4

33 ABI is comprised of three units as shown in - 2. The Sensor Unit (SU) is mounted on the Figure 3 spacecraft Earth Pointing Platform (EPP) and collects the scene radiance and converts it to digital counts. The Cryocooler Control Electronics (CCE) controls the active coolers used to maintain the infrared detectors at cryogenic temperatures. The two CCE units are mounted on the - Y module panel. The Electronics Unit (EU) provides command and control of the SU and CCE. It is also mou nted on the - Y module panel. U ABI 2. nits Figure 3 - Sensor Unit 3 The Sensor Unit (SU) consists of a number of subsystems , as shown in Figure 3 - . The Optical Bench is the backbone of ABI as it provides the structural support for all of the other subsystems, and controls the thermal and mechanical establishes the mechanical alignment to the spacecraft , loads to and from the spacecraft. The Optical Bench provides support for the Optical Port (SCA) Solar Calibration Assembly , Sunshield, the EW and NS Scanners, Telescope, Aft Optics, Internal Calibration Target, Active Cooler, Scanner Shrouds, the Thermal Control Radiator and . The function of each subsystem is Heat Pipes, and the Sensor Unit Electronics (SUE) . This section will 4 address most of the subassemblies in the SU. - summarized in Table 3 Discussion of the SUE can be found in the electronics section, and the two calibration assemblies are described in a later section on calibration during operations. 3 - 5

34 Figure - 3. ABI Sensor Unit Subsystem s 3 Tab le 3 - 4. Sensor Unit (SU) S ubsystems Subsystem Function Optical Bench Provides structural support for subsystems Reduces stray light via series of baffles, prevents contamination Optical Port Sunshield during storage, provides mechanical mount for solar calibration Assembly target sight pointing capability - of - EW and NS scanner Provides line Creates an image of the select scene on the focal plane array Telescope detectors Provides spectral separation and a controlled thermal environment Focal Plane Modules (FPMs) and Aft Optics for the focal plane detectors, which convert photons to electrons Solar Calibration Provides radiometric calibration target for visible and near - infrared channel detectors Assembly metric calibration target for infrared channels Internal Calibration Target Provides radio Active Cooler Provides cooling capability and thermal control of detectors (Cryocoolers) Thermal Control Radiator Provides Sensor Unit thermal control and Heat Pipes within scan cavity Addresses sun’s heat Scanner shrouds Provides video processor to read out detector arrays as well as Sensor Unit Electronics thermal control electronics and control of other mechanisms 3 - 6

35 Optical Port Sunshield Assembly duces the stray light entering the The Optical Port Sunshield Assembly (OPSA) primarily re system. The OPSA has a one time deployable Optical Port Cover (OPC) that is stowed prior to - launch and secured using a non - - puller launch lock. The explosive, shape memory alloy (SMA) pin cover protects ABI against contami nation prior to launch and solar intrusion throughout the launch - raising portions of the mission. Once the launch lock is released, spring - loaded hinges and orbit automatically open the OPC where it is captured on a mechanical stop using a Velcro strip. Optical System After deployment of the OPC on orbit, the SU can begin its primary function of collecting scene - radiance and converting it to digital counts for processing by the EU. The optical collection of the line - - sight at the desired scene. ABI has two radiance begins with the scan mirrors directing of Scan Mirror Assemblies that consist of a scan mirror, a Scan Drive Assembly (SDA), and Support Bearing Assembly. Each SDA mounts to the Optical Bench and supports one side of the scan mirror. It primarily controls the mirror via the motor and monitors the position of the mirror by way of an optical encoder. The SDA motor is controlled via the Scanner Interface & Motor Driver ssembly (CCA) in the EU. The Support Bearing Assembly also mo C C ard A ircuit unts to (SIMD) the Optical Bench and anchors the other side of the mirror. 4 - Figure 3 the two scan mirrors are oriented orthogonally to one another to As shown in , angle is twice the in the NS and EW directions. The optical LOS LOS independently scan the mechani cal angle. T he separation of EW and NS scanning allows for scans parallel to the equator without image rotation and inherently compensates for polarization. Both mirrors can operate LOS simultaneously, permitting scans and slews over a wide range of angles th at allows the to , an to the Blackbody Calibration Target as be pointed anywhere within the f ield - of - r egard as well internal calibration source for the ABI . EW Scanner Z (Earth) NS Scanner Y X (South) (East) Blackbody Calibration Target Figure 3 - 4 . ABI Scanner 3 7 -

36 a telescope and the Aft Optics In addition to the scan mirrors, the ABI optical system consists of Figure 3 5 . The scan mirrors direct the incoming radiance into the telescope. The - as shown in ABI telescope is comprised of four mirrors and forms the image of the scene on each of the three Focal Plane Modules ( . One of the four telescope mirrors can be driven by motor to make FPMs ) The Telescope Assembly consists of the telescope plus the Visible – minor adjustments in focus. Infrared (VIS/IR) beamsplitter and fold mirror. The VIS/IR beamsplitter (BS1) separates the inc oming radiance into VNIR and infrared spectral components. Wavelengths greater than 3 μm are reflected toward the IR focal planes, and those less than 3 μm are transmitted to the fold mirror and then onto the VNIR FPM. The Aft Optics provides additional s pectral separation and holds all the optical components together to provide co - alignment of the FPMs. The separation of the infrared into MWIR and LWIR – Longwa radiance occurs via the Midwave Within each spectral ve (MW/LW) beamsplitter (BS2). band (VNIR, MW IR, LWIR), narrowband spectral selection for each channel is accomplished using filters integrated into the FPMs. The Aft Optics also includes windows and cold stops and provides a controlled cryogenic environment for the FPMs. The LWIR and MWIR optics and FPMs are maintained at approximately 60K. The VNIR optics and FPM are maintained at approximately 170 K (180 K for F PM on GOES - R ). Cooling of the focal planes to these cryogenic temperatures is accomplished using the cryocooler. BS1 Fold SM2 Mirror (E/W) Scan FMA Exit Pupil Mirrors Telescope +X W1 W2 CS - Y W3 Z - BS2 SM1 VNIR (N/S) 0.47, 0.64, 0.865, 1.378, MWIR , MWIR BS2 CS, 1.61, 2.25 60 K 3.9, 6.185, LWIR & LWIR FPMs (170 K) 6.95, 7.34, 8.5 9.61, 10.35, 170 K W3, V/SWIR FPM (180 K for (60 K) 11.2, 12.3, 13.3 (60 K) - GOES ) R Remaining structure 300 K 5 D ock Bl Optics . iagram - 3 Figure 8 - 3

37 Focal Plane Modules The filter the image produced by the telescope into select spectral channels, detect it, and FPMs for the v ideo p convert it into analog signal s rocessor. There are three FPMs corresponding to the ee spectral bands: VNIR, MWIR, and LWIR. Spectral selection is accomplished using thr bandpass filters positioned above the linear detector arrays. The VNIR FPM provides six spectral channels centered on wavelengths ranging from 0.47 μm to 2.25 μm. The MWIR FP M provides five spectral channels centered on wavelengths ranging from 3.9 μm to 8.5 μm. The LWIR FPM provides five spectral channels centered on wavelengths ranging from 9.61 μm to 13.3 μm. The Focal Plane Arrays ( FPAs ) are shown in layout of the VNIR, MWIR, and LWIR Figure 3 - 6 , Figure 3 7 , and Figure 3 - 8 , respectively. The perspective is that of an observer viewing the side of the - FPA illuminated by incident radiance. ABI’s FPA s are the combination of a detector array and its associated Read - Out Integrated Circuit (ROIC) for a single spectral channel. The FPA is the portion of the FPM that detects the incident ral physical properties of each radiance and converts it to an electrical signal. Table 3 - 5 lists seve of the ABI FPAs. Each FPA provides two - fold redundancy. One redundancy consists of separate Side 1 and Side 2 electronics. The second redundancy is the availability of multiple columns of de. The number of redundant columns for each spectral detectors within each electronics si - 5. A single detector channel is shown in the column labeled “Columns per Side” in Table 3 ked during any data collection. This is accomplished using a element from each row is downlin of the selected detector elements for each channel called the Best Detector configurable table Select (BDS) map , as illustrated in Figure 3 9 . - 1 5 8 1 . ° S S S S S S N N N N N N s t s s s s s t t t t t n n n n n n e e e e e e m m m m m m e l e e e e e l l l l l e e e e e e R 0 6 2 6 2 6 I 6 7 7 7 7 7 4 6 3 3 6 6 m m m m m m 1       N 1 8 7 6 4 6 S S S S S 6 8 2 4 6 3 S . . . . . . V N N N N N 9 8 . ° 4 0 0 1 0 0 1 2 N I m m m m m 5 3 2 4 6 1 m      # # # # # #  4 4 4 4 4 B 1 4 2 4 2 2 1 × × × × × × A W W W W W W E E E E E E m m m m m m       4 4 4 4 4 3 5 2 2 5 2 1 N W E 4 - 0 ° 0 7 6 1 . ° 0 1 5 2 0 ° 0 + . 5 . 0 - 2 ° 1 4 ° 0 9 8 ° 5 8 8 9 8 . 0 - 7 + 0 . 8 9 1 + ° 4 8 2 6 5 . 0 S F lane M VNIR . 6 - ocal e 3 Figur P odule L ayout 3 - 9

38 5 5 1 ° 1 . S S S S S N N N N N s s s s s t t t t t n n n n n e e e e e m m m m m e e e e e l l l l l R e e e e e I 2 2 2 2 2 3 3 3 3 3 m m 3 3 3 3 3 m m m   W    0 4 0 9 5 5 3 S S S S S . 9 9 1 . . . . 0 ° 6 0 . 9 M 8 7 N N N N N 6 6 3 0 1 m m m m m I 7 9 8 1 1      # # # # # 0 0 0 0 0 B 5 5 5 5 5 × × × × × A W W W W W E E E E E m m m m m      4 4 4 4 4 5 5 5 5 5 S E W 4 4 7 4 + 0 . 7 4 8 9 5 ° + 0 . 3 7 7 4 7 ° 0 ° - 0 . ° 5 9 8 4 7 . 0 - ° 3 N 3 ocal M odule L ayout P Figure lane - 7 . MWIR F 8 1 . 5 ° 4 S S S S S N N N N N s s s s s t t t t t n n n n n e e e e e m m m m m e e e e e l l l l l e e e e e R 8 2 8 8 8 I 0 0 3 0 0 m m m m 4 3 4 4 4  m     5 3 W 1 2 . . 3 3 . 6 . S S S S 3 S 1 . 0 2 1 1 9 N L N N N N 1 9 ° . 0 6 0 1 6 4 2 m m m m m I 3 5 1 1 1      1 1 # # # # 0 # 0 0 0 0 4 B 4 5 4 4 × × × × × A W W W W W E E E E E m m m m m      6 6 6 4 6 3 3 3 5 3 N W E + 0 . 7 7 1 0 7 ° + 0 . 4 4 5 9 6 ° ° 1 4 4 0 . 0 + 9 ° 4 4 9 5 7 . 0 - 0 ° 6 5 3 . 0 - 4 ° 7 S ocal ayout - Figure 3 L 8 . LWIR F odule P lane M 10 3 -

39 Table 3 - 5 Focal P lane A rray P roperties . Detector Number of Channel Columns per FPM Side NS Rows (μm) Type Silicon 676 0.47 3 Silicon 3 0.64 1460 Silicon 0.86 3 676 HgCdTe 1.378 6 372 VNIR HgCdTe 1.61 6 676 HgCdTe 2.25 372 6 6 HgCdTe 3.90 332 6 HgCdTe 332 6.185 6 HgCdTe 332 6.95 MWIR HgCdTe 6 7.34 332 HgCdTe 6 332 8.50 6 HgCdTe 9.61 332 HgCdTe 10.35 6 408 HgCdTe 11.2 6 408 LWIR HgCdTe 12.3 6 408 HgCdTe 13.3 6 408 3 - 11

40 Side 1 Side 2 2 3 4 5 6 6 1 1 2 3 4 5 N 1 Legend N 1 . 2 . 2 Selected Detector Selected Detectors . 3 3 . . 4 4 . Unselected Detector . 5 5 . Unselected . . . . Detector . . . . . . . . Detector Stack in output . . . . . . . . Video Processor to the . . . . . . . . from North to South for . . Detector stack Detector stack . . Side 1 . . . N . is output to the is output to the . . Video Processor . . Time Sync Time Sync Video Processor W E . . . . from North to from South to . . North for Side 2 . S South for Side 1 . . . . . . . . . . . . . . . . . . . . . . . . . . 5 . 5 4 . 4 . . 3 . 3 2 . 2 . 1 N 1 N Focal Plane Array Stack Stack Focal Plane Array Side 1 Side 2 1 2 1 2 3 4 5 6 6 5 4 3 Legend 1 N 1 N 2 . 2 . Selected Detector 3 . 3 . . 4 . 4 Unselected Detector 5 . . 5 Detector Stack in output . . . . to the Video Processor . . . . . . . . from South to North for . . . . . . . . Side 2 . . . . . . . . . . Detector stack Detector stack . . . . . is output to the . N is output to the . . . Time Sync Time Sync Video Processor . Video Processor E W . . from South to from North to . . . . North for Side 2 South for Side 1 . S . . . . . . . . . . . . . . . . . . . . . . . . . . 5 5 . 4 . 4 . . 3 3 . 2 . 2 . N 1 N 1 Plane Array Stack Stack Focal Plane Array Focal Select Maps (Note: NS is reversed for MWIR FPAs) 3 - 9 Figure . Best Detector 3 12 -

41 Thermal Control ABI utilizes several assemblies and subsystems to maintain thermal control: the radiator, loop heat pipes, scan shroud, cryocoolers, and heaters R adiator/Loop Heat Pipe Assembly r The a ssembly , as shown in Figure 3 - 10 , work in concert to adiator and Loop Heat Pipe (LHP) adiator is a large reflective surface r thermal energy to space. The ensor Unit reject the excess S nstrument and the that radiates energy to space. The LHPs are the interface between the i adiator. r adiator; they transfer excess energy from the rest of the SU to the r 3 Thermal Control Radiator & Heat Pipes Figure - 10 . 3 - 13

42 S can Shroud Assembly 11 , as shown in consists of a series of shields that protect The s can s hroud a ssembly , Figure 3 - the internal instrument structure from direct solar loading through the optical port during the times in the orbit when solar energy enters the internal instrument cavity. The solar he at is collected within the metal shields and transported to the r /LHP ssembly via constant conductance adiator a heat pipes. - Figure 3 11. Scan Shroud Assembly Cryocooler The c ryocooler cools the focal plane arrays to their requisite cryogenic temperatures. It is a two - stage pulse tube active cooler that pumps thermal energy from the focal plane arrays to the r adiator/ Loop Heat Pipe ( LHP ) Assembly. The crycooler is illustrated in Figure 3 - 12 . The ABI has two redundant cryocoolers that can be operated individually or together. Each consists of a Cryocooler Control Electronics ( CCE ) . The TDU consists of an Thermal Dynamic Unit (TDU) and integral cooler , remote cold head, and transfer line. Waste energy is pumped from the FPAs by the TDU to the loop heat pipes where it is transferred to the radiator and rejected to space. 3 - 14

43 Active cryocooler thermal dynamic units mounted to aft optics Figure 3 - 12 . He aters ABI has several types of heaters to maintain temperatures: survival, located within the SU operational, and outgas. All heaters are fully Side1/Side2 redundant. The survival heaters are powered directly by the spacecraft via 70V power and ensure ABI stays at safe temperatures in the absence of operational powe r to ABI. The outgas heaters are used during the outgas phase of the mission to increase the temperature of the SU in order to drive off contaminates from the rs optical surfaces prior to cooling the Aft Optics to cryogenic temperatures. The operational heate are used to control the temperature of the SU during operation. The operational and outgas can have their control points and enable/disable status controlled via ground heaters Power for survival heaters is controlled by spacecraft command, and their set points command. are controlled by fixed thermostatic switches. Electronics ABI’s electronics are dispersed among the three units (Sensor Unit, Electronics Unit, and . The Electronics Unit contains the bulk of the CCAs and Cryocooler Control Electronics) interfaces between the spacecraft and the other ABI units. The contains , located in the Sensor Unit, SUE The CCAs needed to digitize the focal plane data and control SU mechanisms. CCE controls the 13 . Figure 3 cryocoolers. The ABI electronics architecture is shown in - 3 - 15

44 3 - 13 Figure ABI Electronics Block Diagram . Sensor Unit Electronics is comprised of the The SUE and the Peripheral and Thermal Control V ideo P rocessor s (VPs) It is fully Side1/Side2 redundant. The VPs provide the interface between the (P&TC) electronics. focal plane arrays and the EU. The VPs generate timing signals and bias voltages used to read out the FPAs. They also collect the detector samples from the FPAs and format these data for transmission to the Data Processor in the EU. canner) for The P&TC CCA provides thermal and mechanism control (with the exception of the s the Sensor Unit. The P&TC controls the temperature of the Internal Calibration Target (ICT), VNIR FPM, itch between driving the heaters, and outgas heaters. The P&TC motor driver can sw LHP Cover (SCC) and the telescope focus motor. A serial command and telemetry ibration Solar Cal interface is provided for receiving control information from the Telemetry & Timing (TNT) CCA located in the EU. The SUE also controls release of the Optical Port Cover (OPC) and SCC launch locks. Electronics Unit (EU) The Electronics Unit is the primary electrical interface between the spacecraft and the h command upplies and provides the instrument wit s Sensor Unit. The ower EU contains the ABI p and control, data processing, telemetry gathering, and scan control. It consists of a chassis, parent board, and various CCAs . The parent board and CCAs are fully Side1/Side2 redundant. Functional descriptions of the EU CCAs can be found 6. - Table 3 below in 3 - 16

45 3 6 . EU Circuit Card Assemblies Table - Function CCA Converts the +28VDC voltage provided by the Power Supply to voltages required by ABI electronics spacecraft perates the ABI instrument Instrument Controller (IC) Single board computer that o High Speed I/O (HSIO) Communications with the spacecraft via SpaceWire Formats and packetizes detector data provided by the Data Processor V ideo Processor Telemetry and Timing (TNT) s systems clocks and handle s ABI telemetry Generate Interface and Motor Driver Controls the motion of the scan mirrors Scanner (SIMD) Power optical encoders and compute scan mirror EW and NS Encoder Processors (EP) position Cryocooler Control Electronics The two CCE s (one per cryocooler) mount directly to the spacecraft and contain the electronics and software required to operate each cryocooler’s TDU mounted within the Sensor Unit. Waste heat from the CCE is rejected to the spacecraft. The CCE monitors the temperature of the cold PRT ( and adjusts the duty cycle of the power amps ) head via a platinum resis ta nt thermometer to maintain the cold head at its set point temperature. Operation Collection of data by ABI is driven by scenes and timelines. A scene defines the region of interest to be scanned. Each scene is comprised of one or more straight line scans called swaths. A ted scene are collected and the timeline is a schedule that defines when the swaths of each selec - based timeline architecture is that the collection of duration. The primary benefit of ABI’s swath swaths from multiple scenes can be interleaved. Below are a few definitions that are helpful in understanding how ABI collect s data.  Scene: c ommanded area to be observed; constructed from a set of ordered swaths area of scene collected in a single scan defined by start and end  Swath: s ub - coordinates. The scan can be a straight line at any angle but are typically west - to - east, parallel to equator s east)  - Scan: to can maneuver during a swath at constant velocity (nominally 1.4° west -  Stare: s wath with same start and end coordinates typically used for calibration  Slew: s can maneuver between swaths efines what to observe w  Timeline: sequenced set of scene swaths and d hen; it is a time - durations While ABI is flexible in its ability to handle scene definitions, scenes are typically defined in a raster scan pattern as noted in the swath and scan definitions and depicted in Figure 3 - 1 4 . This manner of scanning, made possible by the advanced scanner performance, has several benefits. It allows for a constant time interval across swath boundary, which minimizes temporal distortion. in a similar direction. Image shear is also minimized as all swaths are collected 3 - 17

46 - 1 4 . Figure 3 ABI Raster Scan Fixed Grid Frame All scenes on ABI are defined using the Fixed Grid Frame (FGF) coordinate system, which - of - sight into elevation and azimuth angles . The z - axis parameterizes the line points to n adir, the y axis points south, and the x - axis points east, with the origin of the FGF coordinate system at - 5 - . Figure 3 - 1 the ideal sub offers two perspectives on the definition of angles within satellite point FGF. axis to center of the Earth (and satellite point) -  ideal sub z -  x - axis to east  y - axis to south  Azimuth Coordinate System) - Figure 3 - 1 5 . Fixed Grid Frame (Elevation This coordinate system aligns naturally to the two - mirror ABI scan system as the NS angle (  ) is simply rotation about the x axis and - the EW angle (  ) is rotation about the y’ - axis (the rotated - - of sight unit vector position of the y - axis). The roll and pitch angles are computed from the line components using the equations: 3 - 18

47 For a perfectly aligned scanner located at the Optical angles are twice the mechanical angles. ideal satellite the ideal attitude, the EW and NS Fixed Grid Angles for the instrument position with are simply twice the EW and NS scan mirror sha line of sight , respectively . The ft angles navigation algorithm accounts for the slight deviations from this ideal case using various corrections. The FGF also defines the spacing of pixels. The Nadir spatial pixel resolution of the collected 5 to 1 km in th imagery ranges from 0. visible channels and 1 to 2 km in the infrared channels . e The centers of the 0.5, 1, and 2 - km pixels , as depicted in blue, green, and r ed, respectively, in - 1 are not coincident. This is done so that the nominal area corresponding to a Figure 3 km - 16 , pixel contains within it the nominal areas corresponding to four 0.5 km pixels. Similarly, the - nominal area corresponding to a 2 - km pixel contains within it the nominal area s corresponding to four 1 - km pixels. The spacing of pixels in the FGF is uniform in fixed grid angles. . Pixel spacing in Fixed Grid Frame Figure 3 - 1 6 Standard Earth Scenes ABI operations consist primarily of three types of Earth scenes plus additional scenes necessary for radiometric and geometric calibration. As the calibration scenes are quietly processed in the below background, this section will focus on the standard Earth sce nes shown in Figure 3 - 1 7 . Full m eso scale . The operators can define custom scenes as well, which can be d isk, CONUS, and uploaded anytime during the mission. 3 - 19

48 u l l k F i D s C O N U S E M O S cenes S Full Disk, CONUS, and Meso . 7 1 - 3 Figure Defined as a 17.4 degree diameter circle centered at nadir. It is comprised of 22 d - Full w isk: est . to e ast swaths and is used in Scan Mode 3, 4, and 6 timelines - (Scan Modes are defined in a later paragraph . The flight software automatically extends the defined swaths of the Full Disk scene ) off the Earth to gather both full disk data and a look to space, or as part of a single scan motion , which is used for solar calibration. “spacelook” : Defined as a 3000 km (NS) x 5000 km (EW) rectangle. It is U. Contiguous CONUS ( S.) - to - e ast swaths and is used in the S can Mode 3 and 6 comprised of 6 timeline s . The definition w est - of the CONUS scene is contingent upon the orbital position: GOES - East, GOES - West, or GOES Centr al as shown in Figure 3 - 1 8 . , which can be located Meso: Defined as a square 1000 x 1000 km area (at the satellite sub - point) est ast swaths and is used in of - regard. It is comprised of 2 w - - anywhere within the ABI field to - e Scan Mode 3 There are two Meso scenes available within ABI that can be . s timeline and 6 - adjusted on - the fly by commanding new center locations. This allows an operator to track hurricanes and other storm events by performing rapid revisits of these “mesoscale” phenomena. 20 - 3

49 CONUS Scene for Each Orbital Position Figure 3 - 1 8 . Timelines Timelines are the schedules that dictate when each swath in a selected scene is scanned. Each swath in a timeline is assigned a starting time and duration. The starting time is defined relative to the start of the timeline. The swaths of the various scenes included in the select timeline can be scanned in any order. A helpful analogy for timelines is a musical playlist as laid out in Table The tracks of albums are akin to swaths of scenes, and just as a playlist can mix and reorder . 3 - 7 tracks of multiple albums, so too can a timeline interleave swaths from different scenes. 3 - 7 . Table Timeline / Playlist Analogy ABI MP3 Player Album Scene Track Swath Scan Listen Change Tracks Slew Timeline Playlist 3 - 21

50 To illustrate, a hypothetical timeline that includes two scenes is shown below in Figure 3 - 1 9 . The figure contains depictions of two scenes: Scene A and Scene B. Scene A is comprised of three swaths and Scene B is made up of two swaths. This scanning begins 5.0 seconds after the start of the timeline to allow the scanner to slew to the starting coor dinates and takes 3.0 seconds to scan. The second swath to be scanned is Swath 1 from Scene B, which will begin 10.0 seconds T after the start of the timeline and requires 2.5 seconds. he 2.0 seconds between the completion h in the timeline and the start of the second is added to allow for of the scanning of the first swat the scanner to complete its slew maneuver to the starting coordinates of the second swath. The remainder of the timeline continues with the interleaving of the swaths. 1 – Swath Scene A 2 Swath – Scene A Swath – Scene A 3 1 Scene B – Swath Scene B 2 Swath – Hypothetical Timeline Illustration . 1 - 3 Figure 9 ABI Scan Modes s 3 ABI has primary scan modes: Scan Mode 4 collects just the full disk while Scan Mode three and 6 provide more flexible storm watch capability . Table 3 - 8 p rovides information on the refresh , 20 - - 21 , and Figure 3 rate for the Figure 3 - Figure 3 standard Earth scenes for each scan mode. the standard timelines graphically with ti - ime t me diagrams for the three illustrate 22 scan modes. Each row of these diagrams depicts 30 seconds of timeline activity. - 8 Scene Refresh Rates by Scan Mode 3 . Table Image Collection Revisit Intervals [Minutes] ABI Images Scan Mode 4 Scan Mode 6 Scan Mode 3 (Scenes) (Continuous ( Super Flex Mode) (Flex Mode) Full Disk) 10 5 15 Full Disk 5 5 CONUS --- 0.5 1 0.5 1 Mesoscale #1 --- --- 1 --- 1 --- Mesoscale #2 22 - 3

51 ull - minute timeline that provides one a 15 d isk scene (every 15 minutes), three Scan Mode 3: f CONUS scenes (every 5 minutes), and 30 Meso scenes (one every 30 seconds or two at 1 minute executes the necessary scenes for calibration ( Infrared ( intervals each). It also ) Calibration & IR Spacelook) and INR ) ( Image Navigation and Registration ( tar scenes). s Figure Time Diagram - Scan Mode 3 Time . 3 - 20 isk Scan Mode 4: scene every 5 minutes as well as a 5 - minute timeline that provides one f ull d the necessary scenes for calibration (IR Calibration & Spacelook) and INR (Star scenes). - Time Diagram Figure Scan Mode 4 Time 3 - 21 . 3 - 23

52 : a 10 - minute timeline that provides one f ull disk scene (every 10 Scan Mode 6 minutes), CONUS scenes every 5 minutes, and one Meso scene every 30 seconds ( or two scenes at 1 minute intervals) . It also executes the necessary scenes for calibration ( Infrared ( IR ) Calibration & s Spacelook) and ) ( INR tar scenes). Image Navigation and Registration ( Figure 3 - 22 . Scan Mode 6 Time - Time Diagram Line - of - Sight Pointing Compensation oftware to improve collection of its imagery. (FSW) ABI has several functions within its Flight S ABI receives orbit and attitude data from the spacecraft at 1 H z and angular rate data at 100 Hz. It uses this information while computing scan operations to compensate for differences between measured and ideal orbit and attitude parameters. These functions can be disabled/enabled by ground command. LOS Motion Compe nsation (LMC) : corrects for scanner non - orthogonality, internal misalignments, and instrument - to - spacecraft misalignments. It is applied in real - time. corrects for non ideal attitude using attitude and gyro - Spacecraft Motion Compensation (SMC) : rate data. It is applied in real time. - Orbit Motion Compensation (OMC) : corrects for non - ideal orbital position. It is computed on a - per swath basis and adjusts swath end points prior to collection to remove bias but not swath curvature due to non - ideal orbital position. Bright Object Avoidance While computing its scan operations as defined by the commanded scenes and timelines, ABI un. T he will autonomously adjust its scan pattern to avoid scanning close to the center of the s exclusion zone is a region around the sun that is designed so that swath truncation restricts the 3 - 24

53 field - of - view from reaching the edge of the sun. Depending on the sun location, swaths can be truncated on the e ast side, w est s ide, or skipped altogether. The depth of the “bright object avoidance” outage varies by band due to the East/West position offsets between ABI spectral bands on the focal plane arrays. Each band has a different offset from the instrument line of sight (LO S), which means that the bands are not all viewing the same spot at exactly the same time. The offset between the extreme bands is approximately 1.5 degrees LOS. shows Figure 3 - 23 Figure 3 shows examples of truncations 24 - depictions of e ast - s ide and w est - s ide truncation, and on f ull d isk swaths. Side Truncation - West - Side Truncation East 23 Figure 3 - . Bright Object Avoidance Swath Truncation . Examples of Swath Truncations on Full Disks Swaths Figure 3 - 24 3 - 25

54 Radiometric Calibration ABI conducts two types of radiometric calibration during operations. Calibration of the MWIR and LWIR channels is done with the Internal Calibration Target (ICT). The VNIR channel s are calibrated with the Solar Calibration Target (SCT). Observations of space provide the background measurements for all channels. Infrared Channel Calibration Infrared calibration is conducted via observations of the Internal Calibration Target (ICT) , a high - emissivity, full aperture blackbody calibration source based on a Harris - patented design. The ICT radiance value used for calibration is determined via its temperature, which is maintained at ~302 Calibration occurs at the start of each operatio nal timeline to ensure the MWIR and LWIR K. channels have updated calibration coefficients for data collected during the timeline. This frequent calibration captures the effect of the constantly changing background temperatures of ABI itself. Each timeline b egins with an observation of space to assess background radiance followed by an observation of the ICT. Solar Calibration Solar calibration is conducted via observations of the SCT , which is a partial - aperture, diffuse into the optical system. s part of the SCA within the OPSA. It i white surface that reflects sunlight calibration for the VNIR channels is not conducted frequently but on an as - needed basis. Solar The constant energy of the sun reflecting off the SCT provides a known radiance source for bration. cali ince a specific geometry with the sun must be achieved to ensure significant sunlight S falls upon the SCT, solar calibrations can only be performed during a specific 15 - minute window for any given date. Ground Processing Algorithms downlinked ABI science data packets (Level 0) Ground processing is the method of converting into calibrated, geo - located pixel images (Level 1B). This conversion is accomplished via the ABI compresses the samples from a single detector Ground Processing Algorithms (GPAs). element into a compression block via the onboard lossless Rice compression algorithm during the creation of the CCSDS packets. The process of converting packet data to pixels consists of Figure 3 esampling) as shown in r - four m ajor steps ( d ecompression, calibration, n avig ation, and 2 . 5 3 - 26

55 . Figure 3 - 2 5 Ground Processing Algorithm Flow undoes the onboard lossless Rice compression during the creation of the Decompression : CCSDS packets. converts raw instrument samples (digital counts) into calibrated radiance samples. : Calibration The detector gain and offsets determined from observations of calibration targets (ICT/SCT) and space are applied to each detector sample. It includes a correction for sca n mirror reflectivity and emissivity. determines the location of individual detector samples within the FGF Navigation : . This process reported location. It corrects is a matter of applying a series of corrections to the scan encoder - It also applies corrections from non - ideal for the offset of a si ngle detector to the ABI LOS . (a quadratic estimation algorithm) using data from spacecraft attitude and orbit via a Kalman filter the spacecraft combined with ABI star observations. lue of an FGF pixel as the weighted sum of surrounding calibrated Resampling: estimates the va detector samples, where the weight assigned to each navigated detector sample is based upon its proximity to the selected pixel. 3 - 27

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57 4. Geostationary Lightning Mapper The GOES - R series Geostationary Lightning Mapper (GLM) , manufactured for the NOAA/NASA GOES R project by Lockheed Martin Space Systems’ Advanced Technology Center in Palo Alto, - California, is a nadir - pointed, high - speed video camera that detects the optical signature of lightning illuminating cloud tops at 77 7.4 nm, a wavelength associated with the neutral atomic oxygen emission line of the lightning spectrum. Mounted on the satellite’s EPP , GLM provides hemispherical coverage with its 16 - degree field of view, staring continuously at the cloud tops resolution of 8 (8 km at nadir, increasing to 14 km at the edge of the 14 km - with a near uniform field of view) R flight model of GLM undergoing ground testing can be seen in Figure The GOES - . 4 - 1. The Electronics Unit (EU) can be seen Figure 4 - 1. the Sensor Unit (SU) on on the le ft, and the right. Remote Sensing of Lightnin g Lightning mapping is the process of determining when and where lightning flashes occur. By measuring lightning, forecasters and researchers can monitor important parameters that indicate re thunderstorm development, and predict the formation of tornadoes or the onset of other seve severe weather events. Most of the electrical energy generated by a thunderstorm is dissipated by lightning. The lightning flash rate is quantitatively related to the electrical energy generation in 4 - 1

58 - to - ground and intra - cloud) closely a thunderstorm, and total lightning activity (including both cloud mirrors thunderstorm development. Tracking of convective weather has important applications not just in severe weather “n ow - casting” but also in traffic flow management of air and sea transportation networks, and long - term climatological trending over decadal time scales. A lightning discharge creates and excites atomic oxygen, which decays from its excited state by emitting photons at characteristic wavelengths. To detect a lightning flash, optical lightning mappers typically rely on a prominent oxygen triplet whose emission lines are near 777 nm. The transient optical signature of a lightning pulse diffuses through the surr ounding cloud and 2 illuminates a wide area of the cloud top, typically tens of km . The cloud medium is optically thick but absorbs very little at near - infrared wavelengths, so the resulting multiple scat tering blurs the source geometry, and delays and time - broadens the pulses. Observed on the cloud top, each lightning flash consists of a series of short (less than one millisecond ) strokes separated by . 2 - several milliseconds as shown below in Figure 4 An optical sensor positioned above the cloud top can thus sense the diffuse 777 nm glow from the individual optical pulses generated by the strokes without having a direct view of the lightning plasma channel itself. Peak Optical Pulse Duration 0.9 0.5 400 μs Amplitude 0.1 800 μs Time ulse Typical L ightning O ptical P 2 P rofile - Figure . 4 Detection of light ning is complicated by the presence of bright sun light reflected from the cloud thick cumulonimbus - top. (It is an unfortunate feature of lightning that it usually occurs in optically erference filter and a focal clouds that are particularly reflective). Using a one - wide int nanometer - plane that operates at 500 frames per second, the cloud background created by reflected solar illumination can be subtracted and the transient lightning signal can be detected above the r illumination conditions. Even with the high frame rate - residual noise even under worst case sola and narrow band filter, the background signal can still be orders of magnitude brighter than the lightning signal. Taking advantage of the characteristic temporal signature of lightning, where lash typically consists of a series of strokes separated by several milliseconds and each f 4 - 2

59 generating temporally distinct optical pulses, a lightning mapper system can discriminate lightning even in the presence of various sources of noise and sensor artifacts. A coverage map for GLM is overlaid on a worldwide flash rate map produced from Lightning ( LIS/OTD ) data in Figure 4 - 3 . Imaging Sensor/Optical Transient Detector As can readily be observed, the majority of lightning occurs on land, and there is overlapping coverage of the West. Based on a mean worldwide flash - East and GOES - contiguous United States from GOES rate of approximately 45 flashes/sec, GLM is expected to see ~12 flashes/sec or ~1 million flashes/day on average, with much higher rates during peak periods of convective weather . E - West position - Figure 4 - 3 . ast position (green) and GOES GLM coverage from GOES (cyan), overlaid on a map of flash rate. Instrument Design - arth pointed Sensor Unit mounted on the satellite’s E GLM consists of a nadir - pointing platform , mounted inside the satellite bus. The two units are linked by an and an Electronics Unit (EU) nstrument wiring harness, as shown An electronics block diagram of the . i 4 below in Figure 4 - 5. 1. Table 4 GLM is illustrated in Figure 4 - Key instrument design parameters are list ed in - 4 - 3

60 Sensor Unit (SU) HEATERS, CRITICAL SENSORS) POWER, SPACEWIRE (2) REDUNDANT 59” 19” (48cm ) PRIMARY 150 cm ( ) 27cm 11” ( ) REDUNDANT 14” ( 36cm ) PRIMARY Electronics Unit (EU) Harness 25” (63.5cm) arness and Electronics Unit (EU), with - 4 pproximate A H GLM Sensor Unit (SU), 4 . Figure S izes 4 - 4

61 Electronics Block Diagram . GLM 5 - Figure 4 4 - 5

62 Key GLM I nstrument D esign P arameters 1. - 4 Table Value Unit Design Parameter Lens focal length mm 134 - Lens f number 1.2 +/ 8 deg Lens field of view - pixels CCD imaging area size 1372 x 1300 30 x 30 μm Pixel size (variable, up to) Well depth (variable) 2e6 electrons Ground sample distance 8 – 14 km Frame rate 50 3 fps Filter center wavelength 777.4 nm Filter band pass 1 nm ADC resolution 14 bits 1 - Event rate sec ≥1e5 7.7 Mbps Downlink rate sec <20 Latency 125 Mass (total) kg 67 Mass (Sensor Unit) kg 41 Mass (Electronics Unit) kg Operational power 290 W Flash detection efficiency >80 % (24 hr avg) event % False rate <5 ≥10 years Operating life Sensor Unit element refracting lens with a field of view of 16 degrees, extending almost The SU uses a seven - Earth disk on a Charge Coupled Device (CCD) to the limb of the Earth, to form an image of the iameter, needed to collect focal plane. The lens assembly has an entrance pupil of 110 mm d enoug h lightning photons from the ~40,000 km range of geosynchronous orbit. The stray light design of the lens assembly i s particularly challenging due to the sun coming very close to the field of view during eclipse entry and exit. The SU lens assembly contains three interference filters of increasingly narrow spectral width: a solar rejection filter (SRF) at ~30 nm full - wi dth half - maximum (FWHM) that performs the task of band solar radiation, a solar blocking filter (SBF) at ~3 nm FWHM, and - of rejecting the bulk of out - the key narrow band filter (NBF) at ~1 nm FWHM centered on the lightning triplet. Due to their and stringent spectral requirements, these filters pushed the boundaries of large size manufacturing capabilities. 4 - 6

63 The SU camera electronic assembly is known as the Focal Plane Array Assembly (FPAA) and is . It contains the CCD, 6 associated clock drivers and biasing seen in both Figure 4 - 5, and Figure 4 - circuits, and analog amplifier stages to read out the 56 parallel outputs at a pixel rate of 20 MHz. The FPAA contains 55 separate circuit cards mounted in a cold plate chassis, to which the CCD is bonded. The FPAA is mated to the lens assembly with three shims that set the focus of the - camera to a precision of approximately 10 microns. The FPAA is partly redundant (single string CCD and output amplifiers) and outputs 56 - channel analog video. he GLM Focal Plane Array Assembly (FPAA), mated to Loop Heat Pipe (LHP Figure 4 - 6. T condenser plate at top mates to radiator). The CCD is visible at the center (blue rectangle) with the circular thinned area onto which the Earth is imaged. Red frame is for ground handling. F lex cables carry the analog video to the SU digitizer, known as the Sensor Electronics Box (SEB), which processes the analog FPAA output into digital video for processing by the EU. The SEB contains the Analog to Digital Converters (ADCs) and associated in put circuits, power conditioning circuits, digital logic to support SU command and telemetry, and electronics that assemble the digital video stream. The SEB is fully redundant and outputs 56 - channel digital video with a resolution of 14 bits, serialized i nto 14 parallel SERDES (SERializer - DESerializer) links operating at 1.6 GHz. The SU structure, visible in consists of a titanium/carbon fiber hexapod, with struts 7 Figure 4 - having a zero coefficient of thermal expansion (CTE) for pointing stability. The thermal design of the SU isolates the lens assembly as much as possible, providing a stable thermal environment mechanical distortion. with minimal thermo - A skirt - like carbon fiber honeycomb structure known hown) The entrance optics (not s as the baffle support encapsulates the SU optics and electronics . ring launch and orbit raising, are fitted with a baffle and door assembly, to protect the optics du and to keep direct sun away from the optics to the extent possible. The door is a single - 4 - 7

64 deployment, spring - loaded mechanism released by a redundant High Output Paraffin Actuator (HOPA) some weeks after reaching Geosynchronou s Earth Orbit (GEO), following a period of The digital video signal from the SU outgassing. It is the only mechanism in the GLM instrument. is carried across the 4 - meter instrument harness to the EU, located inside the satellite bus. - of the interior of the GLM Sensor Unit during preparations for thermal Figure 4 7. View vacuum testing, with exterior baffle support removed, revealing lens assembly (top), Metering Tube and hexapod (lower half), FPAA (middle, partially hidden behind ttom). harnesses) and SEB (bo Electronics Unit ), the The EU contains image processors known as Real Time Event Processors (RTEP s SpaceWire communications board, and the power supplies. The EU is a fully redundant unit with 13 electronic modules that plug into a back plane, as of photograph A diagrammed in Figure 4 - 8 . - 9 . the EU is provided in Figure 4 4 - 8

65 A SIDE INPUTS ELECTRONICS UNIT B SIDE INPUTS CHANNELS RTEP BOARD #1 RTEP BOARD #1 CHANNELS (1,2) (2 RTEPS+DF) (2 RTEPS+DF) (1,2) CHANNELS RTEP BOARD #2 RTEP BOARD #2 CHANNELS (3,4) (2 RTEPS+DF) (2 RTEPS+DF) (3,4) CHANNELS RTEP BOARD #3 RTEP BOARD #3 CHANNELS (5,6) (2 RTEPS+DF) (2 RTEPS+DF) (5,6) ADC SENSOR UNIT ADC SENSOR UNIT CHANNELS RTEP BOARD #4 RTEP BOARD #4 BOARDS BOARDS CHANNELS (7,8) (2 RTEPS+DF) (2 RTEPS+DF) (7,8) SCIENCE DATA SCIENCE DATA 1.6Ghz (Ea. Channel) 1.6Ghz (Ea. Channel) CHANNELS RTEP BOARD #5 RTEP BOARD #5 CHANNELS (9,10) (2 RTEPS+DF) (2 RTEPS+DF) (9,10) CHANNELS RTEP BOARD #6 RTEP BOARD #6 CHANNELS BACKPLANE BACKPLANE (11,12) (2 RTEPS+DF) (2 RTEPS+DF) (11,12) CHANNELS RTEP BOARD #7 RTEP BOARD #7 CHANNELS (13,14) (13,14) (2 RTEPS+DF) (2 RTEPS+DF) POWER POWER SC DATA SC DATA CMD/TLM CMD/TLM TELM & CMD TELM & CMD SENSOR UNIT SENSOR UNIT Temp Sensors SPACEWIRE & SPACEWIRE & Temp Sensors Power On Detect HEALTH & STATUS HEALTH & STATUS SPACEWIRE SPACEWIRE PORTS SPACECRAFT PORTS SPACECRAFT +28V INPUT +28V INPUT POWER POWER EU POWER BOARD EU POWER BOARD Operational Operational Heaters Heaters 28V SENSOR UNIT + 28V SENSOR UNIT + , Door Control , Door Control PWR EN PWR EN SU POWER SU POWER BOARD SU POWER BOARD SU POWER PRIMARY REDUNDANT Chassis Chassis Temp Temp Test Connectors . Figure 4 Electronics Unit physical block diagram. Note the Real Time Event Processor - 8 (RTEP) boards are assembled into seven modules with a primary and redundant board each; the total number of physical modules in the EU chassis is 13. 4 The GLM Electronics Unit being readied for shipping. From left to right, the Figure - 9. card modules are: 2x SU power supplies, 2x EU power supplies, 2x SpaceWire modules, and 7x redundant RTEP modules. Red unit in foreground is an environmental recorder used for monitoring conditions during shipping. - 9 4

66 the Seven identical RTEP modules each receive two streams of 1.6 Gbps SERDES data from SU, for a total of eight subarrays handled by each mo dule. The RTEP modules contain the RTEP Field Programmable Gate Arrays ( FPGAs ) event detection logic, implemented in with off - chip memory for storage of the background average. Event data are stored in one FIFO (First In, First Out) queue per subarray, and then formatted in groups of four subarrays by the DF (Data Formatter) logic that sends formatted event data to the SpaceWire board for downlink. All command and telemetry into and out of GLM is controlled by the SpaceWire module. The SpaceWire module gene rates CCSDS packets from the event data, and sends them along to the spacecraft for downlink using the GOES - R Reliable Data Delivery Protocol (GRDDP). GLM does not have a microprocessor; all of the functions related to command and telemetry, time keeping, data formatting, thermal control, and fault management are performed by an FPGA and the SpaceWire Application Specific Integrated Circuit ( ASIC ) , which provides dual redundant data links to the spacecraft clocked at 132 MHz. The SpaceWire module also perfo rms ADC of analog temperature sensors found throughout the instrument. Electrically Erasable Programmable Read board software is stored in - Only Memory - The GLM on ( that can be updated on orbit. The EEPROM also stores configuration tables for the ) EEPROM numerous FPGA registers that control operation of the instrument . In operation, these tables are (RAM) copied into Random Access Memory with Error Detection and Correction ( EDAC) protection to provide resilience against single - event effects from the radiation environment. The EU also contains the power boards for the EU and the SU. These boards take the +28 V DC y regulated voltages used by spacecraft primary power input and generate the various secondar the SU and EU components via DC to DC converters. Event Detection Operation As a digital image processing system, GLM is designed to detect any positive change in the image that exceeds a selected detection threshold. This dete ction process is performed on a pixel - by - pixel basis in the RTEP by comparing each successive value of the pixel (sampled at 500 Hz in the incoming digital video stream) to a stored background value that represents the recent history of that pixel. The bac kground value is computed by an exponential moving average with an adjustable time constant. A longer time constant reduces background noise but increases lag of the background when the pixel illumination changes, such as when the cloud scene evolves. If t he difference between the latest pixel value and its background average value exceeds the detection threshold, an event is generated, as illustrated for just one of ~1.5 million pixels in the - 10 vent occurring at frame 49. . The figure shows a single e example time history of Figure 4 4 - 10

67 Figure 4 - 10 . Time history of a single pixel. A single event occurs at frame 49. Note how background reacts in frame 50. K is the value of the background time constant; higher values of k incorporate more frames into the exponential average. A simplified functional bl ock diagram of the event detection logic is provided in The Figure 4 - 11 . channel basis, among 32 different values - by - detection threshold is determined on a channel nt selected by table lookup based on the brightness of the background as reported with each eve table 4.2 (see ). This allows the detection threshold to be increased in accordance with the brightness (and associated shot noise) of the cloud background signal to maintain a constant rate of false events regardless of illumination. 4 ogic L 11 Figure - etection . Simp lified F unctional B lock D iagram of RTEP E vent D 4 - 11

68 bit data structure describing the identity of the pixel, the camera frame (i.e. time) - An event is a 64 ackground in which it occurred, its intensity with respect to the background, and the value of the b - below self. in Table 4 The data structure for one 2 . Device status and consecutive event is shown it event status are internal status flags relating to internal details of the detection logic of the RTEP. telemetry downlink, and are therefore optimized to Events make up the bulk of the GLM science fit within just 64 bits. - vent Table 2. Data Structure for One E 4 Normal Event Data # Bits Bit Position 0 2 Packet ID 3 Device Status 7 3 9 11 Zero fill 2 10 12 Consecutive Event Status 2 13 23 Frame ID 10 14 24 4 13) - 27 Data Formatter ID (0 28 29 RTEP ID (0 - 3) 2 30 44 Pixel within RTEP (0 - 31849) 15 45 58 Intensity 14 Background Most Significant Bits (background 63 5 59 bits 0 to 4) board image processing Performing on - in the RTEPs and reporting changes in the Earth scene by exception only (when an event is triggered) reduces the downlink data bandwidth of the instrument to a reasonable level, from 14 bits/pixel * (1372 * 1300) pixels/frame * 500 frames/sec = 12.5 Gbps o f raw video data to just ~6 Mbps of processed event data. This is equivalent to a video compression factor of greater than 2,000. The intensity of lightning pulses, like many phenomena in nature, approximately follows a power law. There are relatively fewe r bright and easily detectable events, and a “long tail” of dim events that eventually get drowned out by instrument noise. To achieve high detection efficiency, GLM possible de tection must reach as far into this long tail as possible by operating with the lowest - threshold. The challenge of lightning event detection is then to lower the detection threshold so low that it starts flirting with instrument noise, where random excursions in the value of a pixel can - called “false” event that does not correspond to an optical pulse. The ratio of the trigger a so detection threshold to the standard deviation of the underlying instrument noise is known as the threshold - to - noise ratio, or TNR, and is typically set to about 4.5, meaning that a 4.5 - sigma sigma event - positive deviat ion from the mean value of the pixel will trigger an event. While a 4.5 has only a 1 in 3.4 million chance of happening in any given pixel and any given camera frame, ly be several thousand when scaled up by the frame rate and the number of pixels, there will typical noise false events per second. There are numerous other sources of “false” events, such as the GEO radiation environment (energetic particles that strike the focal plane and generate spurious streaks of light), the sun 4 - 12

69 glinting off t he surface of the ocean, lakes or rivers, and various instrumental effects. The GLM flight hardware has no way of distinguishing these from lightning; indeed, the event telemetry y ground processing, stream contains only a minority of lightning events that must be sifted out b where much more powerful algorithms can be employed than on board the flight hardware. In addition to the event data stream, GLM downlinks a background image every 2.5 minutes (i.e. the downlink. This background image (shown every ~75,000 frames), using only a small portion of 12 in Figure 4 - ) consists of the averaged value of each pixel as stored in RTEP memory, and is not a raw camera frame from the Sensor Unit. Background images are used in the process of event navigation, where dayti me Earth scenes are analyzed for the location of coastlines. Based on the known locations of these coastlines, the geodetic location of each GLM pixel can be derived on the ground. In the near on - infrared band where GLM operates, the contrast between vegetati and water is strong, which facilitates this process of coastline identification. Background images are also used in certain ground processing event filters and help to assess the quality and reliability of GLM data products. GOES-16 GLM Background Image, 2017-02-07 17:31:17 0 200 400 600 y (pixels) 800 1000 1200 1200 1000 0 200 400 600 800 x (pixels) - - Figure 4 Background image from GLM on GOES . 12 16. This image is displayed at fixed pitch, which effectively stretches the edges of scene. Note the strong contrast between land and water . 4 - 13

70 Focal Plane purpose - built solid state CCD with The GLM focal plane, located in the Sensor Unit’s FPAA, is a an overall size of 1372 x 2624 pixels, and an image area of 1372 x 1300 pixels. The CCD is a - transfer device, where each half of the previous image frame is stored in a peripheral area frame of the chip that is not sensit ive to light, allowing readout of the previous image frame in parallel with exposure of the current image frame. The frame transfer architecture enables shutter - less operation at 500 frames per second. The CCD is back fficie ncy side thinned to improve quantum e (QE) over an area matching the image size of the Earth disk on the focal plane. Each pixel of the CCD has a large charge capacity of approximately 2 million electrons (varying proportionally to pixel area), which accommodates the background signal from sunlight reflected by clouds while still providing enough head room to detect lightning. The deep well helps to increase the signal to shot - noise ratio for lightning detected during daytime. pixels tall by 650 pixels wide, The focal plane is sub - divided into 56 physical regions, each 49 known as subarrays. Each subarray is read out in parallel and has an independent signal chain consisting of amplifier, ADC, and RTEP event detection. 0 28 1 29 2 30 31 3 6 4 32 5 33 6 34 35 7 4 8 36 9 37 10 38 2 11 39 40 12 13 41 0 42 14 43 15 Degrees from nadir 16 44 -2 17 45 18 46 47 19 -4 48 20 49 21 50 22 23 51 -6 24 52 53 25 54 26 55 27 2 4 0 -2 -4 -6 6 Degrees from nadir 1 The subarrays and their numbering, overlaid with typical Figure 4 coast lines as - . 3 might be imaged onto the focal plane from the GOES - East position (no yaw flip). Dotted line is the periphery of the backside thinned image area. 14 - 4

71 Figure 4 - 1 3 appear shorter than those at the center The subarrays near the top and bottom of due to the variable pitch of the GLM pixels. The GLM CCD was designed such that the Ground Sample Distance (GSD), the projected area of each pixel on the Earth’s surface, is approximately km matched to the typical size of a storm cell. When following constant with a target value of 8 the development of severe thunderstorms it is important to track the lightning flash rate of individual storm cells; therefore, constant ground sample distance over the Earth is preferred. Near the edge of the field of view, this design (patented under U.S. Pat. 8063968) uses reduced pixel pitch to compensate for the foreshortening as the view shifts away from nadir. This ensures ssociated with the particle that the cloud background signal (and its associated shot noise , a nature of light to - - ) is minimized while lightning signal is maximized, thus preserving a good signal United States is shown The resulting GSD over the contiguous noise ratio near the Earth’s limb . in Figure 4 - 1 4 . The largest pixe ls, near nadir, are sized 30 x 30 μm, with pitch reducing in steps to the smallest size of 20 x 24 μm. The vertical and horizontal discontinuities in the GSD reveal pixel pitch boundaries. The hat every location in the overlapping coverage from GOES - East and GOES - West will ensure t conguous United States is covered with a GSD of 11 km or better. CONUS Ground Sample Distance, GOES-East (km) 0 15 50 14 100 13 150 12 200 11 Y (pixels) 250 10 300 9 350 8 400 7 200 700 600 500 400 300 100 0 800 X (pixels) 1 Figure - 4 - East 4 . Ground sample distance (GSD) over CONUS. In this GOES visualization, CONUS is located in the top left corner of the GLM image; the full image is 1372 x 1300 pixels. Each subarray is read out in parallel with the pixel ordering shown in - 1 5 , which is Figure 4 reflected in the raw science telemetry stream from the instrument. 4 - 15

72 Line readout Line readout North (“Horizontal” clocking) (“Horizontal” clocking) West half of CCD East half of CCD 0 49 49 0 Subarray 0 Subarray 28 31849 31849 48 97 97 48 0 49 49 0 Subarray 1 Subarray 29 31849 48 97 31849 97 48 ... ... 40) (Subarrays 30 – (Subarrays 2 – 12) 0 49 49 0 Subarray 13 Subarray 41 31849 31849 48 97 97 48 97 48 31849 48 97 31849 Subarray 14 Subarray 42 0 49 49 0 ... ... (Subarrays 15 – 53) (Subarrays 43 – 25) 97 48 31849 48 97 31849 Subarray 54 Subarray 26 49 0 0 49 97 48 31849 48 97 31849 Subarray 55 Subarray 27 0 49 49 0 Line (column) transfer Line (column) transfer (“Vertical” clocking) (“Vertical” clocking) South - 1 Pixel N umbering C onvention and R eadout D irection of the F ocal P lane Figure 4 . 5 Thermal Control mounted to a thermally controlled panel inside the satellite bus, and The Electronics Unit is wet - is conductively cooled through its base plate. of the satellite. The Sensor Unit is cooled by an external radiator dedicated to GLM, on the +Y side Waste heat from the FPAA is transported to the radiator by a LHP . This dual redundant LHP is board software and provides variable thermal conductance as - actively controlled by the on plane. Waste heat from the SEB is needed to maintain a constant temperature at the focal transported to the radiator by thermal straps, consisting of flexible stacks of aluminum foil layers. Both the LHP and thermal straps were designed to impart the lowest possible disturbance forces to minimize thermo - mechanical distortions that could cause errors in on the Sensor Unit, so as the navigation of lightning events. All three spectral filters (SBF, SRF and NBF) are temperature - controlled using operational tability becomes important when the heaters, to maintain a stable center wavelength. Spectral s band pass is as narrow as 1 nm; active heater control prevents large temperature drifts that could push the center wavelength away from the xygen triplet and start cutting off the lightning signal. o signed to thermally isolate the lens assembly to the maximum extent The sensor unit is de m ulti - layer i nsulation) thermal possible. From the outside in, this isolation is accomplished by MLI ( emissiv - blankets, the baffle support structure itself (coated on the inside with a low ity surface), and the gold plating on the lens assembly. Survival heat when GLM is not operating is provided by thermostatically controlled survival patch heaters located on the SU lens assembly, and prevent the lens assembly and nearby electronics 16 - 4

73 from cooling below their rated temperature range. These survival heaters are supplied by +70V DC power from the spacecraft. Operational Modes GLM has a very simple concept of operations, with a small set of states and modes as shown in 6 e in the figure is the auto boot path, which occurs during power up without . The bold lin Figure 4 - 1 - commanding and if no errors occur. Normal mode is reached within a few hours of power on, with ter heaters. the timeline being driven by the slow thermal responses involved in LHP start and fil Once in normal mode, GLM generates science telemetry (events and backgrounds) and housekeeping telemetry. Ground Storage and Transportation OFF (no operational or survival power) Launch and On - Orbit Storage (survival power only) SURVIVAL Operational Power Off Detected Error or Watchdog Cold, Warm, or RECOVERY BOOT Manual Boot Recovery reset Load complete Safe Mode switch Reboot/Restart DIAGNOSTIC SAFE to Boot Exit Safe will Instrument power /Safe Diag return to up sequence Mode switch Diagnostic NORMAL GLM odes 6 tates and . 1 - M Figure 4 S While power may be removed at any time, a transition to safe mode is typically used to prepare instrument to lose power and allows for a programmed and ordered shutdown of the the electronics. Ground Processing GLM hardware is designed to detect events, including many events caused by various sources of noise, and sends all these events to the ground for further processing. These raw events are data consists of part of the Level 0 (L0) data stream. The processed data has two leve ls; L1b L2 data consists of the L1b navigated, calibrated events, and events, groups, and flashes . (described later) The first step in the processing is to remove the non lightning events from the data stream. - y reviewing the remaining events. The ground processing algorithms Flashes are then identified b include many filters designed to remove events not caused by lightning, including radiation hits ning un on the ocean. Most of the filters are based on previous work on the Light s and glint from Imaging Sensor (LIS) that flew on NASA’s Tropical Rainfall Measuring Mission (TRMM). The most important filter is the coherency filter. This filter relies on the fact that true lightning events are 4 - 17

74 not. This is the filter that enables GLM to coherent in time and space, whereas noise events are operate near its noise floor, sending many noise events to the ground and detecting fainter lightning events in the process. As viewed from space, any given lightning flash will generate several to several tens o f optical pulses. Flashes can be up to several seconds long, and contain multiple optical pulses detected in the same pixel or adjacent pixels. A noise event will not have of several this coherent behavior. Although many noise events may be triggered over the course seconds, they are unlikely to occur in the same or adjacent pixels. The coherency filter calculates the probability that any given event is a noise event, based on the event intensity, the electronics noise, and the photon noise of the backgroun d. When another event occurs in this same pixel or an adjacent pixel, the filter calculates the probability that both of these events are noise events, based on the new event intensity, the instrument and photon noise, and the time elapsed between the two events. When two events have a sufficiently low probability of both being noise, the events are reported as lightning. This probability threshold is adjustable to allow more or less stringent filtering of the data as desired by the user community. The next step in ground processing is to geo - locate the lightning events, by converting their position on the GLM focal plane (in units of pixels) into a navigated location on the cloud top (as a geodetic longitude and latitude). Navigation takes into account the satellite’s position and attitude, the Earth’s rotation, and must also compensate for small distortions arising from thermal, optical and even relativistic effects. For the purpose of geo - location, lightning is assumed to emanate from a “lightning ellipsoi d,” an imaginary surface several km above the ground at the typical altitude of cloud tops; this minimizes parallax errors at higher latitudes. The navigation process also tags each lightning event with its origination time, which is earlier than its satel lite received time. As observed from GEO, the light travel time from the cloud top to the GLM focal plane is ~0.12 s, during which time GLM has already acquired another 60 image frames. Events are time - tagged to a precision of 1 millisecond. in ground processing is to calibrate the events, a process that converts the L1b The final step intensity of each lightning event from raw units of detector counts to physical units of energy. The L1b data processing consists of all the events labeled as lig htning by the ground output from filters, navigated in latitude and longitude, and calibrated in units of j oules. The L2 algorithms sort the L1b events into groups and flashes. Groups are sets of events that occur in the same frame and are contiguous on the focal plan e. Groups are equivalent to the optical pulses generated by lightning at the top of clouds. Flashes are sets of groups that are associated with each other in time and space. The exact rules for which groups get associated L2 ted by parameters in the algorithm, known as the Lightning together into flashes can b e adjus - Filter Algorithm (LCFA). Cluster The L2 data product from GLM is then used by downstream algorithms to convey information e lightning to its reporting in about lightning to end users. The latency from occ urrence of th L2 data (including propagation time to the satellite, processing by the instrument, downlink to the L2 ) is required to be less than 20 seconds. ground station, and data processing from L0 to L1b to casting” of severe weather based on lightning flash rates This short la - tency is what enables “now tracked within individual storm cells. 4 - 18

75 Space Environment In - Situ Suite 5. - Situ Suite , manufactured by Assurance Technology Corporation, The Space Environment In measures the energetic charged particle environment in geosynchronous orbit, providing real - time dat a to the Space Weather Prediction Center (SWPC), one of NOAA ’s National Centers f or Environmental Information (NCEI). SWPC receives, monitors, and interprets a wide variety of solar terrestrial data, and issue s reports, forecasts, and alerts to the community for the “space weather” conditions. SEISS comprises five individual sensors and a dual redundant data processing unit (DPU) , as (MPS Low Energy Range LO) - shown in Figure 5 - 1. The Magnetosp heric Particle Sensor – measures fluxes of ions and electrons in the 0.03 keV to 30 keV energy range in twelve angular - – High Energy Range (MPS zones. The Magnetospheric Particle Sensor HI) measures protons in the 0.08 MeV to 12 M eV energy range in five angular zones and electrons in the 0.05 MeV to 4 MeV energy range in five angular zones. Two identical Solar a nd Galactic Proton Sensors (SGPS) measure protons and alpha particles in the 1 MeV to >500 MeV energy range. One SGPS or faces east and the other faces west. The Energetic Heavy Ion Sensor (EHIS) measures sens e with single lium ions in the 10 MeV/nucleon to 200 MeV/nucleon energy range for h ydrogen and h n i ckel . The DPU provides the power, telemetry a nd command element resolution through interface to the spacecraft. The DPU also synchronizes the data acquisition of the five SEISS sensors. Figure 5 - 2 shows the location of the sensors on the GOES - R series spacecraf t, and 3 shows the units themselves. Figure 5 - 5 - 1

76 S iagram of the SEISS D uite Figure 5 - 1 . Block 5 Spacecraft 2 Figure - - Series . Location of the SEISS S ensors on the GOES R 5 - 2

77 5 - 3. Figure - Situ Suite Sensors The Space Environment In represents a significant leap forward in technology and capability for The GOES - R series monitoring the energetic charged particle environment. The improvements in SEISS, over the space weather heritage GOES - NOP Space Environment Monitor suite, support the latest NOAA spec ification and prediction requirements. SEISS covers a wider range of particle types, energies and arrival directions than with the previous GOES instruments. Meeting the latest NOAA requirements necessitates addition of two new instruments not previously f lown on GOES – the LO suprathermal plasma analyzer and the EHIS heavy ion cosmic ray detector. The new - MPS MPS HI and SGPS medium and high energy electron and proton instruments have been designed - x range than the previous GOES to make accurate measurements over a much wider dynamic flu MAGPD (Magnetospheric Proton Detector) , MAGED and ( Magnetosphere Electron Detector) instruments. compares graphically 4 below Figure EPEAD (Electron, Proton, Alpha Detector) - 5 trons of the SEISS sensors with the heritage the measurement capabilities for protons and elec GOES - NOP sensors. The MPS - LO, MPS - HI and SGPS sensor units do not contain microprocessors and therefore have no software. However, they do contain firmware in the form of Field Programmable Gate Unit (CPU). The EHIS does contain a Control Processing Arrays (FPGAs). 5 - 3

78 lectron Figur e 5 - 4. Graphical C omparison of the P roton and E apabilities M easurement C of the GOES eritage GOES R SEISS S ensors and the H - - NOP S ensors DPU to the spacecraft for the power, command, control and DPU provides the electrical interface The telemetry output. It controls the operation of the five SEISS sensor units and performs a power from the spacecraft and output - volt (V) power converter function to take the prime 28 regulated voltages of ±15V and ±7.5V to each of the five sensor units. Two fully redundant power supplies provide power to the five SEISS sensor units. A SpaceWire interface with full strap capability is used to communicate - redundant cross y ween the spacecraft and the active digital circuit card assembly command and telemetry data bet . in the DPU DPU processing includes command and telemetry data handling, but no additional sensor unit data manipulation. The sensors each process their own data. The DPU communicates with the sensor units over an RS - 422 synchronous serial interface. Fully redundant is a functional diagram for the DPU Figure 5 - data buses are provided to the five sensor units. 5 and shows the DPU interfaces with the sensor units. 5 - 4

79 5 - 5. Figure DPU Functional Block Diagram MPS - LO Sensor MPS LO measures ions and electrons in the 0.03 keV to 30 keV energy range. The instrument - measures and reports particle fluxes in 15 logarithmically - spaced energy channels and 12 unique angular zones. The - 1 lists the MPS - LO energy band centroids for ions and electrons. Table 5 - 5.8% of the centroid energies. Figure 5 shows the configuration of the 6 energy band widths are angular zones with respect to the instrument and the spacecraft pointing. and Because electrons protons at these energy levels generate very small signals, solid state detector ( SSD ) technology is not feasible for their measurement. The refore, the MPS - LO sensor uses an electrostatic analyzer (ESA) and multipliers. The ESA operates on the princi ple that a particle of a certain energy in a cylindrical or spherical electric field will travel in a specific circular path. An ESA uses a pair of deflection electrodes, with an electric field across the gap between the electrodes to guide charged particl e trajectories. Particles with just the right energy pass between the electrodes without collision. Those with energy too high or low collide with the electrodes and are lost. The deflection electrodes operate as a narrow band energy filter. The MPS - L O has four sets of deflection electrodes, two for electrons and two for ions. The four sets are mounted in two triquadrisphere assemblies. Each triquadrisphere provides a 120 - degree of - o each other to provide a view and the two triquadrispheres are mounted at right angles t - field 5 - 5

80 - degree field - of - total 180 The twelve 15 deg angular zones will enable determination of view. pitch - angle distributions from the orbital data . Each triquadrisphere assembly contains both an . des and an inner set of ion deflection electrodes outer set of electron deflection electro The bias voltage between the deflection electrodes is rapidly stepped to selectively filter particles from 0.03 keV to 30 keV. At a particular voltage step, only particles of a specific energy and charge w ill pass through the plates to reach the MCP detectors. A complete sweep of the energy range takes one second. Particles at the correct energy pass between the electrodes and are accelerated into an he Microchannel Plate (MCP), which electron multiplier, an avalanche device referred to as t generates a signal large enough to measure or count. The 270 deg geometry of the MPS - LO deflection electrodes preserves the angle of incidence information of the incident particles and provides excellent energy resolutio n (  E/E = 0.058). The 270 deg geometry, combined with the LO is solar blind. Another unique feature of - deflection electrode coating, also ensures that MPS the MPS LO is the inclusion of detectors, shielded from the suprathermal plasma, that provide an - pendent measurement of backgrounds due to penetrating radiation. Those background inde measurements are utilized in the ground processing algorithms to correct the particle data during solar particle events and major geomagnetic storms. - MPS LO The consists of t wo triquadrisphere assemblies, the MCP assemblies, sensing and counting electronics, and an RS - 422 interface with the SEISS Data Processing Unit (DPU) implemented in firmware within a single Field Programmable Gate Array (FPGA). ability for space weather measurement on the GOES series - MPS - LO represents new cap R spacecraft. MPS - LO data will reveal the level of charging by low energy electrons that the GOES - Spacecraft charging can cause electrostatic discharge ( R spacecraft is undergoing. and ESD ) arcing be tween two differently charged parts of the spacecraft. This discharge arc can cause serious and permanent damage to the hardware on board a spacecraft, which affects operation, navigation and interferes with measurements being taken. The data will also be used as inputs to develop and validate new models of the space radiation environment. 5 - 6

81 Table 5 - 1. MPS - LO E nergy B and Centroids for I ons and E lectrons. Note that energy band width is 5.8% of the centroid value. Energy Band Centroids (keV) Band Electrons Ions 0.030 E15 0.025 E14 0.049 0.040 E13 0.080 0.066 E12 0.130 0.115 E11 0.212 0.192 0.346 E10 0.316 E9 0.564 0.527 E8 0.926 0.888 E7 1.514 1.502 E6 2.490 2.439 E5 4.094 4.043 E4 6.588 6.732 E3 11.200 11.200 E2 18.590 18.590 E1 30.810 30.810 5 - 7

82 MPS-L O F i e l d -O f-Vi e w Sh o w i n g Se c to r s a n d s e n o Z +X +Y FO V Fa n in t he Y - ne Z P la o o n 1 V FO FO 5 e v e , S V 180 or s P e r S phe r e S e c t X . -Y i- r a hw t r a E d) nt A Z ( - -X ZG R ZG L -Y +Y 0 • 7 Zone s For 3 0 eV t o 3 keV 2 Z1 Z1 B M nd 1 k a c s nt me e ur s a gr ound e A Z2 1 Z1 c ie s . P e r S phe r Zone e e r S pe P Z3 0 Z1 f O ut % O 0 1 t e e o M d T e de e nd N a B Z4 ) R 7 quir e 9 D e O P ( nt me R Z9 Z8 Z5 R Z6 Z7 L Z7 R L Z6 o 30 p in C ov e r a ge la O v e r o or t c T o 1 5 w S e s r r A nt i- E a Z ( t hw a - d) -0 -A0 AM1 4 -0 1 8 0 - LO F ield of axis points – Figure 5 - 6 . MPS - Z V ones. The Z ngular A iew and earthward. The Y - axis points north - south. - anti MPS - HI Sensor MPS - HI measures protons in the 0.08 MeV to 12 MeV energy range and electrons in the 0.05 MeV to 4 MeV energy range. The instrument records and reports proton fluxes in 10 energy spaced electron logarithmically - spaced proton energy channels and 10 logarithmically - 3 - Table 5 channels with an additional >2 MeV integral channel for electrons. list - 2 and Table 5 HI proton and electron telescope channels. The instrument - MPS the energy bands for the solid state detector (SSD) telescopes silicon comprises 5 proton and 5 electron telescopes. Each telescope has a 30 deg ree full - angle field of view and the telescopes are arranged to provide a - 7 170 deg ree shows the configuration of the telescope fields of view with field of regard. Figure 5 respect to the instrument an d the spacecraft pointing. MPS - HI measurements are similar to measurements performed by the heritage GOES space weather instruments; however, MPS - HI provides greater energy range and greatly improved out - of medium and high band rejection compared to previous sensors. The MPS - HI sensor monitor s - energy protons and electrons which can shorten the life of a satellite. High energy electrons are extremely damaging to spacecraft because they can penetrate and pass through objects which owns and result in discharge damage inside of equipment. can cause dielectric breakd , while each electron telescope implanted - Each proton telescope comprises three ion SSDs comprises nine SSDs. In the electron telescopes, three of the SSDs form individual detector remaining six detectors are connected into a single detector channel. The channels, while the thicknesses and dimensions are selected to meet the energy range and field of view requirements for the system. The proton telescopes also include a magnet in the aperture to defle ct low - energy - general, the low energy electron flux exceeds electrons from reaching the first detector. Since in the low - energy proton flux, sometimes by orders of magnitude, the magnetic deflection aids in the particle - type discrimination. 8 5 -

83 MPS - HI al so includes two dosimeter sensors for measuring the Linear Energy Transfer (LET) The energy particles behind two different shielding thicknesses. The dosimeter Field of View ( FOV ) of is approximately 180 degrees. Figure 5 - . Fields of view of the 5 electron telescopes (E1 - E5) and the 5 proton telescope 7 south. (P1 - - P5) for MPS - HI. The – Z - axis points anti - earthward. The Y - axis points north Energy elescopes Table 5 - 2. T B ands for the F ive MPS - HI P roton Energy Bands (keV) Proton Channel Upper Lower P1 80 115 P2 115 165 P3 165 235 P4 235 340 340 P5 500 P6 700 500 P7 700 1,000 P8 1,000 1,900 P9 1,900 3,200 3,200 6,500 P10 12,000 6,500 P11 5 - 9

84 3. Energy B ands for the F ive MPS - HI E lectron T - Table 5 elescopes Energy Bands Energy Bands (keV) (keV) ETel 3, 5 Channel ETel 1, 2, 4 Lower Upper Lower Upper E2 90 145 95 140 E3+E3A 145 230 140 275 E4 325 275 405 230 E5 325 460 405 609 E6 460 705 609 794 E7 705 1360 794 1364 1903 1364 E8 1360 1785 E9 2685 1903 2842 1785 2685 4345 2842 4515 E10 4345 5660 E10A 4515 5899 E11 2000 2000 SGPS Sensors The SGPS instruments measure protons in the 1 MeV to >500 MeV energy range. The two identical instruments measure and report the flux in 10 logarithmically - spaced energy channels channel for protons >500 MeV. The energy range is covered using three separate and 1 integral SSD telescopes. angle - Telescope 1 measures 1 MeV to 25 MeV protons and has a 60 degree full - cone field of view. Telescope 1 comprises two SSDs. Telescope - 2 measures 25 MeV to 80 MeV - protons and also has a 60 degree full - angle conic field of view. Table 5 - 4 lists the energy bands for the two SGPS sensors. Telescope - 2 comprises three SSDs and aluminum energy degraders. 3 measures 80 MeV to >500 MeV protons, has a 90 degre Telescope - angle conic field of e full - view and comprises three SSDs and copper energy degraders. he fields of view of each of the T telescopes are co - aligned. The fields of view of one SGPS unit point east, the other points west. Figure 5 - 8 shows the SGPS configurati on and fields of view. SGPS measurements are similar to measurements performed by the heritage GOES space weather instruments; however, SGPS provides greater energy range and greatly improved out - NOAA sist of - band rejection compared to previous sensors. The data pro vided by SGPS will as ’s in providing solar radiation storm warnings. These particular measurements are crucial to NCEI the health of astronauts on space missions, though passengers on certain airline routes may 5 - 10

85 experience increased radiation expos ure as well. In addition, these protons can cause blackouts of radio communication near the Earth's poles and can disrupt commercial air transportation flying polar routes. The warning system allows airlines to reroute planes that would normally fly over E arth’s poles. energy protons (>80 MeV) - A significant design challenge for SGPS is the differentiation of high that enter a telescope from the front or the back. To facilitate that differentiation, we have included front circuitry that compares the signals in the most SSD and the rear - most SSD. Protons entering - from the front of the sensor, will deposit slightly less energy in SSD 1 than in SSD - 3. For protons - entering from the rear of the sensor, the opposite will be true. The design uses a precision tor to compare the signal in SSD - 1 and in SSD - 3. The results of that comparison aid the compara entry particles. - entry from rear - on - board logic in differentiating front Figure 5 The fields of - 8 . Fields of view of the three SGPS telescopes. view of the three telescopes are co 2) . - - aligned and point east (SGPS - 1) and west (SGPS 5 - 11

86 5 Energy B ands for SGPS+X and SGSP - X (50% points) - 4. Table SGPS+X SGPS - X Channel High (MeV) Low (MeV) Low (MeV) High (MeV) 1.86 1.02 1.86 P1 1.02 1.90 2.30 1.90 2.30 P2A 3.34 2.31 2.31 3.34 P2B 3.40 6.48 3.40 6.48 P3 11.0 5.84 11.0 P4 5.84 11.64 23.27 P5 23.27 11.64 25.9 39.1 24.9 38.1 P6 41.2 74.3 40.3 73.4 P7 98.5 83.7 P8A 82.9 99.8 121 96.4 118 99.9 P8B 114 148 115 143 P8C 242 160 160 242 P9 P10 276 404 276 404 P11 540 540 EHIS S ensor a The heavy ion measurements made by EHIS are series space new capability for the GOES - R weather instruments. The EHIS is responsible for measuring heavy ion fluxes in the magnetosphere to provide a complete picture of the energetic particles surrounding Earth. This includes particles trapped within Earth’s magnetosphere and particles arriving directly from the c rays which have been accelerated by electromagnetic fields in space. This sun and cosmi information will be used to help scientists protect astronauts and high altitude aircraft from high levels of harmful ionizing radiation. EHIS measures fluxes of ions from protons through nickel. The energy range of measurement is 10 – 200 MeV for protons. For heavier ions it is the energy range for which that ion penetrates the same amount of material, in areal density, as do the 10 – 200 MeV protons. The data are binned in five lo ydrogen garithmically spaced intervals in this energy range for five mass bands, H , hosphorus He , Carbon through O xygen , Ne on through P lium , and C h l orine through Ni ck el . Table . The 5 - 5 lists the energy bands for representative ions that span the EHIS detection range counting rate over which the instrument functions ranges from the galactic cosmic ray rate of ~0.3 . events/sec through the solar particle event rate of 30,000 events/sec 5 - 12

87 angle conic field of view. The - EHIS contains a single SSD telescope with a 28 degree full , with a pointing accuracy of 2 degrees. earthward - 9 shows the telescope is pointed anti Figure 5 - telescope. - EHIS sensor, its field of view, and a cross sectional view of the EHIS uses commonly accepted techniques and a unique trajectory system to achieve these goals. The technique EHIS employs is an energy loss (dE/dx) vs. residual energy (E) measurement co taken with a stack of axially mounted SSDs. A pl astic scintillator acts as a veto for thirtee n - entry particles. side - Ions travel through detectors depositing energy in the silicon SSDs that is read out as a current t The firs pulse. Each SSD outputs a charge (current) pulse proportional to the energy deposited. three SSDs are used in the Angle Detecting Inclined Sensor (ADIS) system. Accurate incident particle angle information allows for the individual elemental separation. I n order to identify heavy by - event basis us ing a dE/dx vs. E method, it is ions with good mass resolution on an event - necessary to account for the angle of incidence of the particle. Identical ions entering an instrument at different angles will deposit different amounts of energy in the various detectors. (D2 and D3) t with two detectors mounted at known angles ADIS takes advantage of this effec with respect to the plane normal to the telescope axis. The polar angle is 30 deg for both D2 and D3 but the azimuthal angles are different. The energy deposition in these tilted detectors is the energy deposition in the one mounted normal to the instrument axis. From the compared to energy deposited in these inclined detectors, the on - board processing determines the differen t angle of incidence of each particle - With these angles and the measu red energy deposits in the first six SSDs, EHIS uses an on The ZCAL method, board processor to calculate the charge of a particle using the ZCAL method. developed California Institute of Technology , is an approximation that enables the originally at determ The on - board ination of the ion charge based on the energy deposited in the detectors. processor can analyze between 1,800 and 2,700 events/second. Results are stored in charge - - energy histograms for telemetry to the ground on a 1 minute cadence. 5 - 5. EHIS E nergy B Tabl R epresentative I ons. e ands for ions from H (Z=1) to Ga (Z=31) EHIS reports abundances of all E4 E1 E5 E3 E2 (MeV/nucl) (MeV/nucl) (MeV/nucl) (MeV/nucl) (MeV/nucl) t A omic High Element Low High Low High Low High Low High Low Number 1 H 179.75 109.50 13.0 0 31.50 31.00 44.50 43.50 56.25 54.00 92.25 2 He 31.75 31.00 43.75 44.00 52.75 54.50 91.75 110.00 194.00 10.00 C 6 81.50 59.50 57.00 82.25 18.50 98.75 101.00 171.75 210.00 335.25 7 N 19.75 65.00 61.25 89.75 107.75 108.50 184.75 229.00 367.00 87.25 8 O 67.00 400.75 97.00 22.00 70.75 249.50 96.50 117.00 119.00 200.00 12 Mg 87.00 82.75 118.00 118.25 143.00 26.25 147.00 247.50 312.25 493.75 14 Si 92.50 29.00 97.25 132.50 132.00 160.25 164.50 279.00 352.50 567.25 26 Fe 825.50 37.50 131.50 124.50 180.50 180.50 219.75 226.50 393.25 501.00 5 - 13

88 . 5 sectional view of the - The EHIS sensor, 28 degree field of view and a cross Figure - 9 Detecting Inclined EHIS telescope showing the inclined detectors used in the Angle . Sensor 5 - 14

89 Magnetometer 6. - R series Magnetometer assembly (MAG) consists of two m agnetometer instruments The GOES operating simultaneously . The instruments are mounted on a deployable to allow for gradiometry so as to minimize the spacecraft influence on boom to keep them away from the spacecraft . instrument measurements assembly is provided by ATK, while the individual The MAG magnetometer sensors are provided by Macintyre Ele ctronic Design Associates, Inc (MEDA) . Each M - axis sensor and an electronics unit. Each agnetometer instrument consists of a three agnetometer measures three orthogonal vector components of the magnetic field in m instrument M agnetometer axes are orthogo nal to within +0.5 degrees the vicinity of the spacecraft. The three and calibrated to within less than 0.1 degrees. Each sensor has a linear range of +512 nanoTesla (nT). This includes a measurement resolution of 0.016 nT and measurement bandwidth 2.5 Hz. The determination of the ambient magneti c field within the vicinity of the spacecraft is simultaneous and continuous. The MAG supports the following mission objectives:  Map the space environment that controls charged particle dynamics in the outer region of the magnetosphere  Measure the magnitu de and direction of the Earth’s ambient magnetic field in three orthogonal directions in the geosynchronous equatorial orbit  Determine general level of geomagnetic activity - Detect magnetopause crossings, storm sudden commencements, and sub storms  measure s and map s MAG the space environment magnetic field that controls charged The particle dynamics in the outer region of the magnetosphere. These particular particles can be dangerous to spacecraft and astronauts. T hese geomagnetic field readings are importan t for providing alerts and warning s to many customers including satellite operators and power utilities. The MAG also d etermine s the level of geomagnetic activity as well as detect s magnetopause crossings and storm sudden commencements. provide the The GOES - R series M AG to s vital information to both the satellite itself, as well as individuals monitoring weather patterns on the ground. This system serve s space as an early warning system for large scale magnetic storms. These measurements are used to validat e large 1 scale space models that are used in operation. Figure 6 - and 6 - 2 show notional depiction s of are mounted on a deployable boom as shown with magnetometers The two the MAG instrument . ponents are thermally supporting electronics mounted onboard the spacecraft. These com controlled to maintain operational temperatures. 6 - 1

90 Figure 6 - 1. Magnetometer Instrument Locations +Z +X + +Z Y - + Unstowed Stowed Configuration Configuration Figure /Unstowed Stowed . Magnetometer 2 6 Configuration - 2 - 6

91 MAG is used for the estimation of the Earth’s magnetic field. A gradiometric The data from the algorithm is used in this estimation, in conjunction with a priori knowledge of the satellite magnetic ambient magnetic field. characteristics to accurately estimate the - axis m agnetometers utilize a gradiometer effect, allowing for continuous monitoring The two three of the spacecraft magnetic field. The gradiometer uses a mathematical model to represent the spacecraft’s field characteristics as a s ingle dipole. The inboard m agnetometer is deployed 6.35 meters from the spacecraft on the boom, and the outboard magnetometer is deployed 8.56 meters away from the spacecraft. These large distances from the spacecraft significantly reduce from the spacecraft’s body. The boom employs a light - weight , composite, coiled magnetic effects design to permit compact stowage. Its release mec hanism consists of a frangibolt with an actuated release pin. The sensors themselves are fully blanketed with thermal blankets t o protect them from the thermal environment. electronics for the magnetometer assembly are located on the , the 3 - Figure 6 As can be seen in - 4, t he sensor data passes As shown in the block diagram in Figure 6 spacecraft on the +Y panel. through a Remote In terface Unit (RIU) to the Command and Data Handling (C&DH) subsystem. ) FSW pass The flight software ( es the 10Hz magnetometer data to the ground , and the ground processing algorithms take each magnetometer measurement and generate an estimate of the . ambien t field based on the gradiometer algorithm, calibration data, and transient compensation Mission operations maintains system performance parameters via periodic in - flight calibrations. spacecraft static magnetic situ calibration takes into account the characterization of the The In - haracterization of the spacecraft AC magnetic field at the field at the magnetometer sensors, the c haracterization of sensors, and the c the magnetic sensor alignment with respect to the attitude reference frame . +Zsc +Xsc Magnetometer Electronics Units Figure - 3 . Location of Magnetometer Electronics Units 6 6 - 3

92 Figure 6 - 4 . Magnetometer Block Diagram - has four modes of operation. Normal mode provides 3 MAG axis science data at 10Hz, time The stamps and status data. Maintenance Mode is used to support ground calibration and testing along with supporting upload/download of calibration coefficients and firmware images. Health Test Mode executes a set of step commands and reports telemetry exactly as it does in normal . mode T his generates a predefined magnetic field step sequence to confirm the health of the unit. ode. M is Off The last mode is Safe Mode , which 6 - 4

93 7. The Extreme Ultraviolet and X - ray Irradiance Sensors Instrument Space Weather Monitoring : The Sun/Earth Weather Connection Measuring solar radiation variability is an important component of space weather monitoring as s un, in the form of solar flares and associated coronal mass huge eruptions of energy on the , can have severe impacts on the Earth’s atmosphere and human endeavors. This solar ejections radiation variability drives the heating, ionization, chemistry and dynamics in the Earth’s atmosphere which in turn can create hazards for astronauts in the form of an increase d exposure to radiation and for orbiting satellites where atmospheric density changes can affect orbit integrity and tracking. Earth terrestrial level affects can include communications blackouts, disruptions to Global Positionin power grids and errors in g System ( GPS ) navigation. The GOES program at its inception in 1975 had a terrestrial focus that looked to provide continuous Earth imaging and sounding data. Monitoring solar irradiance in the soft X - rays was part of the GOES program rays - starting in 1986, an d on the NOAA SMS series prior to that, leading to the use of the soft X measurements for the classification of flare magnitude. Starting with the GOES N satellite series - providing solar in 2001, capability was added to monitor and study effects of solar dynamics by imaging and extreme ultraviolet irradiance measurements. instrument by the University of Colorado’s The GOES - R satellite series development of the EXIS Laboratory for Atmospheric and Space Physics (LASP) continues these important space wea ther measurements. EXIS focuses on measuring the brightness of the s un at several different wavelengths of light that have been shown to affect the Earth’s atmosphere and terrestrial environment. A detailed instrument description follows. Sun Pointing The Extreme Ultra violet and X - ray Irradiance Sensors instrument resides on the SPP which is mounted to the yoke of the spacecraft solar array. EXIS instrument ) Platform ( s un looking for channels continuously measure the absolute brightness of the full disk of the changes that provide an early warning of an impending solar storm. On a quarterly basis, for a point in order to monitor any changes in short period of time, the SPP performs a 16° off - signals and performing instrument performance by comparing to mission start baseline dark electronic calibrations of the instrument detector systems. Figure 7 - 1 below shows how EXIS into the SPP and how the SPP is integrated onto the . Key physical spacecraft integrates 1. - parameters for the EXIS are shown in Table 7 7 - 1

94 pacecraft S ntegrated onto the I Figure 7 - 1. E XIS Integrated into the SPP and the SPP - Table 7 R EXIS FM1) 1. EXIS Physical Resource Summary (GOES - 7 - 2

95 Figure 7 - 2. Overview of EXIS Subassemblies and Key Metrics 2 ay Sensor (XRS), an r - : an X EXIS is divided into four subassemblies as depicted in Figure 7 - Extreme Ultraviolet Sensor (EUVS), an EXIS Electrical Box (EXEB) and a Front Aperture Assembly (FAA) which includes a door mechanism, a filter mechanism, and a Solar Position - Sensor (SPS). Attending flight software facilitates all intra instrument communication and external communication with the spacecraft. - A block diagram of the EXIS is shown in Figure 7 described in the following paragraphs. 3, w it h each subassembly 7 - 3

96 NOTES: - parentheses, such as (A), refer to an operational side. Letters in The symbol (A, B) indicates device is accessible by both instrument side A and B. - A capital letter not in parentheses refers to an optical channel. - EXIS Block Diagram 3. - Figure 7 XRS s un and use filters to measure the X ray Consists of six photometers, four of which look at the - bands of interest. Two photometers are covered and provide “dark” photometer background 0.4 - A1/A2) covers 0.05 information for subtraction from the prime viewing channels. Channel A ( - nm and channel B (B1/B2) covers 0.1 0.8 nm. All active channels view the sun through two beryllium ( Be ) filters with the thickness of the filters determining the bandpass. Each active XRS - channel consists of a low - sensitivity (A2/B2) and a high sensitivity (A1/B1) detector whose - nsitivity detectors se responses overlap in order to span the required total dynamic range. The low are quadrant photodiodes, which view the sun through a small aperture, allowing X and Y position information to be extracted for bright, impulsive events such as solar flares. The high - sensitivity detectors are single element photodiod es with larger apertures. The aperture assembly of the 4 - 7

97 - house designed electron deflection system (Ramatron) that deflects XRS incorporates an in incoming electrons from the XRS detectors so only X - rays are measured. The Ramatron also provides magnetic shie lding to the outside environment to minimize a magnetic signature that might affect other satellite measuring systems. EUVS Consists of three spectrographs which measure sunlight in select wavelengths of interest. The three spectrographs, denoted A, B and C, give coverage in the bands of 25 - 32 nm (0.6nm - - resolution), 115 285nm (0.1nm resolution) respectively. From 141 nm (0.6nm resolution) and 275 these, a reconstruction of the full solar spectrum between 5 nm and 127 nm is generated. The s use gratings, filters and solid state detectors to make the measurements in three spectrograph their respective wavelength bands. Post - dispersion photon detection is done via custom arrays of discrete silicon photodiodes for the A and B spectrographs, and a linear 512 - elem ent photodiode array for the C spectrograph which is made up of two redundant units, C1 and C2. The A spectrograph makes measurements in the Extreme Ultraviolet (EUV), the B spectrograph in the Far Ultraviolet (FUV), and the C in the Middle Ultraviolet (M UV) portions of the spectrum. The C channel also provides a calibration standard for the A and B spectrographs. EXEB Contains the instrument low voltage power supplies and a FPGA with imbedded microprocessor facilitate housekeeping data collection from and instrument internal/external interfaces that voltage, current and thermal monitors, instrument science data collection from XRS A/B, EUVS - A/B, EUVS C1/C2 and SPS, control of door motor, filter wheel motor, stimulus lamps, instrument and opera tional heaters. All communication to the s pacecraft is via redundant calibration, SpaceWire links using the GOES - R Reliable Data Delivery Protocol ( GRDDP ) . FAA and SPS Are made up of a door mechanism, a filter mechanism, a baffle assembly and a solar position sensor as shown in Figure 7 There are two mechanisms on EXIS, one being the EUVS filter - 4 . mechanism which is used for selecting from redundant filters for EUVS A measurements, and the contamination second one being the EUVS door mechanism whose purpose is to limit the rate that can enter the EUVS optical apertures before going into on - orbit operations (not a hermetic seal). Both mechanisms consist of an aluminum disk directly connected to the output shaft of a stepper motor assembly. Each disk is mounted perpendic ular to the axis of rotation of its shaft. The door disk has open and closed aperture positions that are placed at the locations needed for them to either block or open the lines of sight to the associated spectrograph. Although the door mechanism has the capability to open/close as desired, the on orbit intent is to open and leave open for the mission duration. The filter disk is made up of 24 redundant filters that can be cycled as needed to support the normal solar measurements and additionally to provi de degradation checks of filters that sustain heavy usage. The Solar Position Sensor is made up of a quadrant photodiode, aperture and control electronics whose purpose is to accurately and at high cadence report the position of the solar disk within the nstrument field of view. i 7 - 5

98 Door Mechanism Filter Mechanism Solar Position Sensor Figure 7 - The Front Aperture Assembly (FAA) which includes a door mechanism, a filter 4. mechanism, and a Solar Position Sensor (SPS). Flight Software up code stored in programmable read - only memory ( PROM ) and a Consists of start - reconfigurable/uploadable code stored in EEPROM that executes from RAM. This excludes all FPGA instruction sets. Thermal F 7 - 5 ). EXIS thermal es 1 and 2 (see EXIS has two thermal zones, simply referred to as Zon igure channel detector package and radiator. Zone 1 is the remainder Zone 2 consists of the EUVS C - of the instrument. Each zone has its own redundant set of both operational and survival heaters. Operational heaters are controlled internally by a linear proportional control system located on the operational power boards. Survival heaters receive +70V power from the spacecraft and are switched by thermostatic switches located in each thermal zone. 6 - 7

99 Thermal Zone Operational Range Survival Range Zone 1 C  C to +50 - 5  C to +20  C - 35  Zone 2 30 - 15  C to 0  C C -  C to +50  EXIS Thermal Zones Figure 7 - 5. 7 - 7

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101 Solar Ultraviolet Imager 8. The Solar Ultraviolet Imager, manufactured by Lockheed Martin, is used to determine when to issue forecasts and alerts of “space weather” conditions that may interfere with ground and space systems. These conditions include ionospheric changes that affect radio communication ( both to ground - to - ground and satellite - ground) and magnetospheric variations that induce currents in - electric power grids and long distance pipelines. These conditions can cause navigational errors aft charging, produce high energy in magnetic guidance systems, introduce changes in spacecr particles that can cause single event upsets in satellite circuitry, and expose astronauts to increased radiation. SUVI is designed to provide a view of the solar corona by taking full - disk solar images at high cadence aro und the clock, except for brief periods during eclipse or instrument calibration, in the extreme ultraviolet (EUV) wavelength range. Available combinations of exposures and filters allow coverage of a range of solar features, including coronal holes, X - cla ss flares, and estimates of temperature and emission measurements. Images from SUVI will be used by NOAA and U.S. Air Force forecasters to monitor solar conditions that affect space weather conditions, including the dynamic environment of energetic particl es, solar wind streams, and coronal mass ejections emanating from the sun. These data can be used to issue forecasts of solar phenomena. from a geosynchronous orbit located at SUVI is tailored specifically to observe solar phenomena longitude. SUVI’s primary science objectives include: either 75° or 137° west  - speed solar wind streams causing Locating coronal holes for the prediction of high recurrent geomagnetic storms. These weakly emitting features are good predictors ear solar activity minimum when long - lived of geomagnetic storms for the years n holes are present on the s un.  Locating the position of solar flares, both on the disk and beyond the west limb, to predict the magnitude of particle events. the east limb. Enhanced emission  Identifying solar activity rotating onto the disk from above the limb provides information about solar activity occulted by the solar disk at other wavelengths.  Monitoring the s un for evidence of coronal mass ejection (CME), which is associated with geomagnetic storms. Reliab le indicators of the CME occurrence include separating flare ribbons, post flare loops between them (in long duration events), and large scale coronal dimming.  Observing the size, temperature, morphology, and complexity of solar active regions. Changes i n these properties will be used to predict the rate of growth of solar active regions and the probability that the regions may flare. To meet these objectives, the SUVI images the solar corona in the EUV region of the disk s olar images are provided with a 1280 X 1280 array with 2.5 electromagnetic spectrum. Full - arc second pixels in six wavelength bands from 94 to 304 Å (9.4 to 30.4 nm). The SUVI optical - 8 - 1

102 system employs a Ritchey - Chretien telescope consisting of multilayer coated optics, and a CCD detector at its focus to record images of the solar disk and its atmosphere. Using an aperture selector, SUVI operates at any one of the six EUV narrow spectral passbands via combination of thin film filters and multilayer coated optics. Each optic (mirror) of the telescope has six distinct multilayer coatings that are fine - tuned to reflect at a well - defined EUV wavelength that corresponds to a particular temperature region of the observed solar atmosphere, as shown in Figure 8 - 1. A second intervals is used r egular sequence of exposures tha t are downlinked at ten - to cover the full dynamic range needed to monitor solar activity. presents a sample set 2 - Figure 8 s un in the six chosen wavelengths , of images of the bandpasses , representing a discrete range of plasma temperatures of the sun’s atmosphere, from 0.3 to 3 million Kelvin. Hotter temperatures greater than 3 million K are reached during transient events such as flares and coronal mass . ejections The SUVI 8 - 1. Figure EUV Wavelength B and s used to observe the Range of Solar pa orecasting . Phenomena important for S ce Weather F 8 - 2

103 minute averages of 1 - 2. 60 8 - second exposures, taken on January 23, 2017 of all Figure - six of EUV p ass b ands SUVI’s of a telescope assembly, an electronic box, and the cables that run The SUVI instrument consists between the telescope assembly and the electronics box, as shown in Figure 8 - 3. The top level - Figure 8 block diagram is presented in All SUVI components are mounted on the SPP as shown 4. The SUVI Electronics Box (SEB) provides the instrument control, data Figure 8 - 5 . in management, conditioned power to the rest of the instrument, and the spacecraft interface. T he SUVI Telescope Subsystem (STS) consists of the EUV Telescope Assembly (ETA), the Guide Telescope Assembly (GTA), and the Camera Electronics Box (CEB). The GTA includes the Guide Telescope (GT) and the GT Pre Amp box. The GTA is mounted on the ETA along with the CEB. - The CEB provides CCD sensor control and image processing manageme nt. The GTA provides of solar pointing data with respect to the instrument line - - sight to the s pacecraft during normal - 1. 8 operation. A summary of SUVI’s characteristics is given in Table 8 - 3

104 - The SUVI Instrument Figure 8 . 3 I ument D lock Figure 8 - 4. The SUVI B nstr T op L evel iagram 8 - 4

105 able 8 - 1. Solar Ultraviolet Imager Characte ristics T Mirrors Multi - layer - coated Zerodur Number of coating segments per mirror 6 20 cm Primary diameter Effective focal length 173.04 cm Field of view 45 × 45 arcmin or better Pixel size/Resolution 21 μm/2.5 arcsec CCD detector 1280 × 1280 pixels Detector full well 450 000 electrons 1 per 10 seconds Full image frame rate 0.01 to 1 second Typical exposure times Flight computer BAe RAD750 Mass: Telescope subsystem 39 kg Electronics box 25 kg - instrument harness 8 kg Intra 225 W (peak) Instrument Power Science telemetry Interface to spacecraft 3.5 Mbps orbit 10 years (after 2 years of on Design life - storage) 8 - 5

106 8 Figure 5. Solar Ultraviolet Imager and Solar Array in Deployed Configuration - Instrument System SUVI is mounted on the SPP and co aligned with the EXIS. The SPP assembly is mounted on - the solar array yoke to continuously face the sun. The SPP is actuated using the SPP Elevation Gimbal Assembly (SEGA) in the north - south direction, tracking the sun in solar declination. The Solar Array Drive Assembly (SADA) controls the east - west pointing of the yoke, tracking the S pointing and the yoke E - n. Both the SPP N W pointing are controlled - diurnal motion of the s u loop control systems that utilize two s pacecraft - provided closed - during normal operations by - axis position error data from the SUVI Guide Telescope (GT) . pacecraft provided Sensor (interchangeably called the Sun Pointing The s Interface Unit (SIU) - Platform Interface Unit) and Fine Sun Sensor (FSS) are also mounted on the SPP, as shown in Figure 8 - 6 . The SIU is located between SUVI and EXIS and provides the command and telemetry interface between the instruments and the axis s - SpaceWire pacecraft. The FSS provides two position error data when SUVI GT data is unavailable. 8 - 6

107 M - 6. 8 ounted on the SPP ( h arness and blankets not shown) Figure SUVI and EXIS , electronics, harness) is 72.0 kg, of which 39 kg is the The total mass of the SUVI (telescope telescope assembly. Electrical connections to the GOES spacecraft cross the SADA interface through slip rings. Sixteen power rings have been allocated to SUVI. pacecraft data All SUVI s interfaces ar e carried via SpaceWire to the GOES spacecraft, and sixteen signal rings have been allocated to SpaceWire. The SpaceWire network is capable of handling a SUVI data rate of 10 Mbps. EUV image data are transmitted from the Spacecraft at a rate of approximate ly 6 images per minute via the high - rate s pacecraft Raw Data Link (RDL). SUVI health and safety data is also ealth data link. h bservatory o transmitted from the s pacecraft as part of a separate low - rate Operation Operation of SUVI is controlled through the SEB. The SEB primarily consists of three parts : power distribution, control processing, and data handling or storage. Control is performed via a RAD750 microprocessor. The SEB receives, interprets, validates, and executes both stored sequence commands and immediate execution commands. Through these commands, the SEB coordinates and controls the activities and operations of the SUVI instrument. In addition, the SEB receives, collects, and multiplexes the science, engineering, and housekeeping data from the pacecraft communication subsystem for SUVI components and provides the data stream to the s transmission to ground facilities. SUVI transmits this data as SpaceWire data packets formatted Figure 8 per the GRDD P . - 7 gives a functional flow definition for the SEB command and data Figure 8 - 8. handling. The SEB electronics block diagram is presented in 8 - 7

108 8 - 7. The SEB C ommand/ D Figure F low B lock D iagram ata - iagram D lock Figure 8 B 8. The SUVI E lectronics F unctional - bit resolution pixel/s through either one of two ports at 14 The CEB reads out the CCD at 2 M ega with 40 electrons rms noise (including CCD contribution), stores the image data as the CCD image rate IEEE 1355 SpaceWire ) LVDS is being read out, uses a high - voltage differential signaling ( low - interface with the SEB for image and housekeeping telemetry, and conditions and converts the 28V input power. 8 - 8

109 The Bridge/SpaceWire Board contains both the SpaceWire ASIC and the Bridge FPGA. The ekeeping telemetry, mechanism Bridge FPGA provides interfaces for the power system, hous control, the Guide Telescope, and provides a lockout function to prevent conflicts if both RAD750 processors in the SEB are powered at the same time. The SpaceWire ASIC provides redundant s pacecra 10Mbps SpaceWire links to the ft, a 10Mbps SpaceWire command link to the CEB, and a 50 Mbps SpaceWire data link to the CEB. The Bridge/SpaceWire board combines the CCD image data, GT pointing data, housekeeping, and memory dump data, and passes them to the the SIU s the SPP. The SIU converts the SpaceWire data into RS - 422 for pacecraft through on - ring. The s pacecraft onboard computer transmission over the solar array drive assembly slip extracts and processes the housekeeping and GT pointing data from the combined SUVI data to ult management and SPP pointing functions. support fa SUVI data are transmitted to the ground by two paths, low rate telemetry for housekeeping data - image and telemetry only, and via the Raw Data Link for all SUVI data. The data can be received M , and the NOAA Space Weather Prediction Center (SWPC) in at SOCC in Suitland, aryland olorado . In general, image data are downlinked as rapidly as possible after they are Boulder, C acquired. Power Elec trical power is provided to SUVI from the spacecraft electrical power subsystem. The s pacecraft is at the SEB connector panel. The electrical interface between SUVI and the s pacecraft delivers both +28V and +70V dc power to SUVI, and these power inputs are protected pacecraft The SUVI electrical harnesses are routed from the - to with fuses within the s pacecraft. - s s pacecraft bus via Solar Array Drive Assembly slip rings and across the SPP Elevation Gimbal Assembly. The s pacecraft provides operational power to the SUVI instrument from the power bus that is regulated at 28.0 ± 2.0 V dc during sunlight operation. SUVI uses two separate and redundant T he first redundant pair serves as the main power inputs and is sets of 28V dc input circuits. by the SEB for the use of various SUVI instrument converted into a number of different voltages components and subsystems. The second redundant pair provides power to the CCD decontamination heaters which are used to keep the CCD sensor warmer than the rest of the instrument while the Instrument is tu rned off, so as to not serve as a contamination “getter.” During eclipse, this primary power bus is controlled by battery voltage and a voltage regulator, which maintains the +28 volt bus. The s pacecraft primary bus (70.0 ± 2.0 V dc) provides power to the survival heater power located on the Telescope Subsystem. The input operational power atts maximum during eclipse. consumption by the SUVI is 90 w atts maximum in sunlight and 162 w Telescope Subsystem consists of the ETA, the GTA, and the CEB. The GTA includes the GT and the GT Pre - The STS . Figure 8 - 3 Amp box. Both the GTA and CEB are mounted on the ETA, as shown in EUV Telescope Assembly (ETA) The ETA is a 20 - cm (8 - inch) Ritchey - Chretien telescope that includes a number of mechanisms - 10 . The SUVI metering tube and sensor packages. The layout of the ETA is presented in Figure 8 is the ETA’s primary structure that integrates the main telescope, GTA flexure mounts, aperture s are door, CEB, camera, radiators, associated housing/adapter and flexure mounts. The strut 8 - 9

110 mounted on the metering tube located at the front and aft of the telescope carbon fiber metering tube and interfaces with the SPP. Each of the six mounting struts is identical with adjustment capability and provides the primary load path to the SPP. The GTA is mounted to flexures which are themselves mounted to the ETA metering tube assembly. The Spider Assembly, Front Aperture Housing and Aperture Door are mounted at the front end of the metering tube while the Shutter housing, Filter Wheel Housing, Isolator tube and Detector Housing including radiator are mounted at the aft of the metering tube. The CEB is mounted on the +Z side of the telescope next to the Filter Wheel housing. The Spider Assembly contains the Focus Mechanism, secondary optics, and the structure to mount to the metering tube assembly. One spoke of the Spider Assembly houses the ight e mitting l iode (LED) Assembly that supports aliveness testing of the ETA focal plane array and camera d Front Door Assembly is closed. electronics while the L ETA - Figure 8 ayout 9 . Optics The SUVI optical design is based on the Solar Dynamics Observatory (SDO) Atmospheric Imaging Assembly (AIA) instrument design modified to meet SUVI’s specific performance requirements and co mply with SUVI’s allocated mechanical envelope. The general optical layout . is shown in Figure 8 - 9 To meet mission requirements for spectral sensitivity and response, the design combines a eve the required narrow normal incidence telescope with multilayer mirror coatings to achi bandwidths. A system of aperture masks and internal baffles are also used to suppress out - of - band radiation and eliminate direct paths for non - solar particle radiation that might otherwise reach the CCD sensor. Each wavelength is ac complished by dividing the aperture into six sectors, in Table 2, the 94Å sector is made of below - 8 each with a different multilayer. As shown 8 - 10

111 - Yttrium (Mo/Y) multilayers while the other five utilize a Molybdenum/Silicon (Mo/Si) Molybdenum multilayers. le 8 Tab 2. SUVI Optical Multi - L ayer Prescription - Capping Layer Total Film Channel (Å) Multilayer Number of Multilayers Thickness (Å) 93.9 Mo/Y 120 33 Å Mo 5748 131.2 35 Å Si Mo/Si 50 3357.5 171.1 Mo/Si 40 3534 35 Å Si 195.1 Mo/Si 40 4080 35 Å Si 284.2 3049 Mo/Si 20 30 Å Si 303.8 30 Å Si Mo/Si 20 3300 Mechanisms The SUVI Telescope is equipped with a set of mechanisms that support the imaging operations or keep the instrument safe and clean of contamination during ground transportation prior to The mechanisms are listed in Table 3, along with a brief description 8 launch, and during launch. - of their function. 8 - 11

112 SUVI - 3 Table 8 Telescope Mechanisms . Mechanism Description Front Aperture loaded hing e, latch mechanism, and -  Consists of a door, a spring gearbox mechanism for driving the door open. and Door  The latch is operated using redundant paraffin linear actuators. Assembly The drive motors are fully redundant and independent.   The door is designed to sweep 245° to fully open. Aperture  Located in front of the spider assembly Selector  Used to select the desired bandpass out of the six wavelength bands. Focus  Actuated by a DC torque motor that moves the secondary mirror over a ±800 μm range in 3 - μm steps Mechanism  Trims the ETA’s initial focus on ce on orbit  Adjusts telescope focus for slow thermal drifts throughout mission Focal Plane Consists of a circular blade turned by a brushless DC motor  Shutter Blade has two openings: a narrow slit and a wider opening to  support either sweeping t he solar image across the CCD or Mechanism expose the entire CCD all at once  The blade can also be commanded to dwell to support long exposures. Filterwheel  Two filterwheels present in the ETA. Mechanism Selects the proper filter combination to ensure that the desired  channel (wavelength) reaches the detector.  Each forward/aft filterwheel contains 5 positions  The mechanism has a positional accuracy of ±30 arc minutes with a move time less than 1 second be tween adjacent positions.  Each filterwheel is designed to have redundant thin and thick Al and Zr filters so the telescope can continue to image properly despite the presence of pinholes. Each wheel has one open position and the aft filter wheel has an  additional clear glass filter so that the light leak performance of orbit. - the entrance filter can be monitored on Camera System The SUVI camera system primarily consists of a focal plane detector using a CCD, CEB and the ed flex cables and head associat board. The CEB receives raw image data from the CCD, - processes the data, and forwards it to the SEB, via a high speed IEEE 1355 SpaceWire low - 8. voltage differential signaling ( LVDS ) interface, as shown in Figure 8 The SUVI instrument focal plane contains a 1280×1280 pixel CCD detector. The CCD is a design and S similar in construction to the proven Solar - B/Hinode Focal Plane Package (FPP) olar - ial output register to D ynamics O bservatory designs. They feature low voltage clocking of the ser minimize power dissipation in the clock driver electronics. The CCD is back thinned and back - illuminated with 21 μm pixels and operate non inverted to ensure good full well capacity. The - 12 8

113 SUVI thermal control system maintains the CCD at - 30C or less. The CCD operating temperature architecture and readout map is shown below in Figure 8 - 1 0 . Figure 8 - 1 0 . CCD Architecture Guide Telescope Assembly The SUVI Guide Telescope Assembly (GTA) was designed and built to support the o bservatory’s pointing functions. The GTA, via the SEB, provides solar pointing data to the - pacecraft s sun during normal SUVI operation by determining the position of the solar limb relative to the GTA centerline. The GTA is approximately 81 cm (32 inches) in length with a glint - free field of v iew ( 1 along with a short description of key FOV ) of 10°. The GTA layout is shown below in Figure 8 - 1 components. The GTA has a linear range of approximately ±110 arcseconds, and is capable of V, of approximately ±26 arcminutes. s un within a cone, defined as the a cquisition F O acquiring the The GT A is designed to operate between 5 °C to 35 °C. - 1 . GTA Layout Figure 8 1 8 - 13

114 - 4. GTA Components Table 8 GTA Description Component Optics  Design derived from S olar Dynamic Observatory (SDO) Atmospheric Imaging Assembly (AIA) and Solar Terrestrial / Solar Dynamic Observatory Atmospheric Relations Observatory Imaging Assembly (STERO/SECCHI) ground telescope (GT)  Galilean telescope design with a bandpass entrance filter, an objective lens “semi - cemented” doublet and Barlow lens  Lenses are manufactured from radiation hardene d glass  Entrance filter consists of a filter plane sandwiched between two pieces of radiation hardened glass Sensors Four redundant pairs of photodiodes arranged in a cruciform  pattern  Cruciform is located behind an occulter cone to measure the solar limb position Pre - Amplifier  Amplifies the photodiode signals  Built with redundant connectors and ¼” thick a luminum enclosure Operation Modes SUVI has five modes which may be used over the course of the GOES mission. They are distinguished by the telemetry generated in each mode. Each of the five modes use event flags to signal specific events that could affect the modes or the telemetry output. SUVI modes are illustrated below in Figure 8 - 1 2 . Figure D iagram low 8 - 1 2 . The SUVI I nstrument M ode and F 8 - 14

115 SUVI OFF (Survival) Main SUVI instrument power is off. No instrument telemetry is generated. Power is provided for thermostatically controlled survival heaters from 70V dc bus, and the CCD decontamination s pacecraft if they have been heaters are drawing +28V dc power from the turned on by the operator and the s pacecraft is power - positive. The temperatures are monitored by the s pacecraft - pacecraft s telemetry system during this mode using SUVI - provided calibrated thermistors and provided conditioning circuitry. Temperatures moni tored in this mode include the mirror assemblies (both primary and secondary mirrors), the CCD assembly, the door mechanisms, the guide telescope (both forward and aft ends), the camera electronics box, and SUVI electronics box. Restricted Mode In this mod e, SUVI is drawing +28V dc power. SUVI housekeeping and engineering telemetry is generated. Diagnostic and event message te lemetry may also be generated. This mode is designed for software maintenance; i.e. debugging and updating. The instrument enters thi s mode Computer Software d from one of the other modes. either after the initial boot - up or by a comman Configuration (CSC) items may be loaded and unloaded in this mode. The software will not be None of the SUVI functional processes fully operational until all flight software CSCs are loaded. are active in this mode. Standby (Idle) Mode In this mode, SUVI is drawing +28V dc power and all flight software CSC’s have been loaded and activated. No updates to the software can be performed in this mode. SUVI housekeeping and engineering telemetry is generated. Diagnostic and event message telemetry may also be - generated. One or more sub systems may be powered off, as indicated by the sub - mode. Thermal control may be inactive. Operational constraints on sequencing and image sizes are disabled. The SUVI i nstrument can be in this mode while the telescope door is closed. Operations (Sequenci ng) Mode This is the mission mode. This mode may be entered only by ground command. All subsystems level - are powered on, the telescope door must be open and thermal control is enabled. Low ty of inadvertent interruption device control has been disabled. This mode minimizes the possibili of the mission observational program. Housekeeping, engineering telemetry and science data telemetry are generated. Diagnostic and event message telemetry may also be generated as needed. 3 The nominal imaging sequence is shown in Figure 8 - 1 . The full sequence takes four minutes to complete and is divided into twenty - four 10 - second slots during which an image is acquired and processed. Twenty - two images are allocated across the six wavebands to maximize science minute sequence. During - efficacy. Two ca libration images are also collected during every 4 eclipse, SUVI can be commanded to perform mechanism characterization tests. 8 - 15

116 equence S Figure 8 - 1 3 . The N ominal SUVI I mage Safe Mode n SUVI enters this mode upon receiving a “Enter Safe Mode” command or upon detecting an internal fault. The safe mode is used to prepare the instrument to lose power and allows for a programm ed and ordered shutdown of the SUVI subsystems: all EEPROM operations are Thermal control is maintained pped; and the shutter is closed. stopped; image sequencing is sto by the instrument. Only a limited set of commands are accepted in this mode. All instrument telemetry is generated. Event Flags Event flags are used to signal specific events that could affect the modes or the telemetry output. sub modes ” and are not mutually exclusive so SUVI may be in a particular They essentially act as “ mode with two or more event flags. Ground Processing ix SUVI ultraviolet The Level 1b SUVI data product is an image of the sun in one of the s - time resolution, with radiometric and geometric corrections applied. The wavelengths in full space image has been converted to physical units, and supplemental information for further processing is appended as metadata. The image is o riented in the same configuration as if the user was – solar north will be in the top half of the image and viewing the sun from aboard the spacecraft solar east will be in the left half of the image. Ground processing of the downl inked SUVI images is carried out by the SUVI - designed Ground Processing Algorithm (GPA). The GPA reduces and processes the information from the raw image data received from the SUVI CCD readout onboard the S/C into usable level 1b level data products. Re duction of raw image data follows a series of steps to manipulate the data stored in each pixel of the image in order to produce a numerical value in each pixel that corresponds accurately to the number of photons that were incident on that pixel during th e image exposure. Through reduction of the raw image data, by removing error sources, an accurate representation of the field of view (FOV) at the time the image was taken can be reproduced. The SUVI raw image data uses a number of calibration factors in cluding fixed properties of the instrument - (primarily the CCD, camera, and electronics) measured on the ground prior to launch, and on orbit Instrument and Spacecraft factors that are monitored in real time because of variations and other environmental factors. based on time, temperature 8 - 16

117 GOES - R Communications Subsystem 9. - R System is to acquire and disseminate environmental data from a The mission of the GOES - equatorial Earth orbit at geostationary altitude. near The Communications Subsystem provides the following functions: on of instrument data to the Command and Data Acquisition (CDA)  Transmissi Stations  Transmission of spacecraft telemetry to the CDA Stations  Reception of spacecraft commands from the CDA Stations  - way ranging and Doppler from CDA Stations and DSN Stations Two  Relay of Unique Payload Services signals in support of:  GOES Rebroadcast  Emergency Managers Weather Inform ation Network broadcasts  Search and Rescue  D ata Collection Platforms Antenna  reception of GPS Navigation signals Figure 9 1 outlines the GOES - R RF interfaces between the Communications Services to - accomplish these functions . The GOES - R Communications Subsystem consists of a raw data link and six (6) bent p as depicted in Figure 9 - 2. ipe services A suite of transponder payloads provid e communications relay services and GOES mission data transmission. The suite consists DCS), the High Rate information Transmission/Emergency of the Data Collection System ( Managers Weather Information N etwork (HRIT/EMWIN), GOES Rebroadcast (GRB), GOES Raw Aided Tracking (SARSAT) system. - Data Link (RDL) system, and the Search and Rescue Satellite band earth The GRB transponders consist of dual polarized X - band uplink and dual polarized L - coverage downlink . The HRIT/EMWIN transponder is an S - band uplink to L - band earth coverage downlink narrow bandwidth transponder. The SAR and Data Collection Platform Report ( DCPR ) services are UHF uplink to L - band earth coverage downlink narrow bandwidth transponders. The band uplink to UHF earth coverage - Data Collection Platform Command ( DCPC ) service is an S downlink narrow band transponder. 9 - 1

118 Communication Links Overview Figure 9 - 1. Spacecraft - Raw Data, Bent Pipe Transponders DCPR UHF to L UHF to L SAR HRIT/EMWIN S to L Uplinks Raw Data DCPC S to UHF Source X to L GRB GRB X to L KEY UHF L-Band HRIT/ SAR GRB DCPC DCPR S-Band GRB EMRIT Raw Data X-band Downlinks 2 . GOES Bent Pipe Transponders - Figure 9 C) Subsystem provides telemetry, tracking, and The Tracking, Telemetry, and Control (TT& - raising and normal on - command functions through orbit station operations. 2 - 9

119 service functions are: The Comm unications Raising Tracking, Telemetry and Control ( Provide S and Orbit - - ORTT&C ) • b Launch and Orbit Raising ( communications for ) command, telemetry and tracking LOR • Provide S - b and ORTT & - Band Command and Data Acquisition ( CDA ) C and L (housekeeping) TT & C communications for on - orbit command, telemetry and tracking Support terrestrial and oceanographic Data Collection Platforms (DCPs) via the Data • ( DCPC ) & Data Collection Platform Report ( DCPR ) links Collection Platform Command • Relay High Rate Information Transmission (HRIT) and imaging data between Earth terminal s and relay the EMWIN broadcast on the HRIT/EM WIN • Provide rapid detection of distress messages from the Search and Rescue (SAR) Emergency Locator Transmitters (ELTs) and Emergency Position Indicating Radio Beacons (EPIRBs) • Rebroadcast processed GOES sensor data via the GRB data link R operations will be conducted from the NOAA NSOF - on phase, GOES During LOR Missi supported by a global station network. This network will support all command, telemetry and separation acquisition through all apogee thruster firings and - tracking requirements from post depl oyments. - - W and post Post launch test operations will be supported Launch Checkout position at 89.5° primarily by the NOAA ground station. Ground support stations provid e backup, tracking and emergency support during all phases of the GOES missions for the life of the spacecraft series. GOES Rebroadcast (GRB) GOES Rebroadcast is the primary space relay of L1b products and will replace the GVAR service. GRB will provide full resolution, calibrated, navigated, near - real - time direct broadcast data. L1b full set of products from all The content of the data distributed via GRB service includes the instruments onboard the GOES R series spacecraft. This concept for GRB is based on analysis - that a dual - pole circularly polarized L - band link of 12 MHz bandwidth may su pport up to a 31 - Mbps data rate – enough to include all ABI channels in a lossless compressed format as well as data from GLM, SUVI, EXIS, SEISS, and MAG. Data Collection System (DCS) The DCS is a satellite relay system used to collect information from E based data collection arth - - situ environmental sensor data, such as stream or river flow, tide - levels, platforms that transmit in weather conditions, etc. The transmissions can occur on predefined frequencies and schedules, ed conditions, or in response to interrogation signals. Th in response to thresholds in sens e this signal and then rebroadcasts it transponder on board the GOES - R series satellite s detect s - so that it can be picked up by other ground based equipment. Federal, state and local agencies based then m - onitor the environment through the transmission of observations from these surface data collection platforms. The platforms can be placed in remote locations and left to operate with 9 - 3

120 tion. The DCS thus allows for more frequent a nd more geographically minimal human interven The DCS data flow is depicted in Figure 9 - 3. complete environmental monitoring. In the GOES - R era, the number of user - platform channels were expand ed 266 to 433. There from 696 MHz to 1679 MHz, which required the replacement of a frequency change from 1 was also users’ Low Noise Block (LNB) feed s . Direct Readout Ground St ation (DRGS) manufacturers were informed of this change. Data transmission rates in the GOES - R era are 300 bps and 1200 bps. There was no cha nge to the data access policy. Figure 9 3. DCS Data Flows - High Rate Information Transmission (HRIT)/ Emergency Managers Weather Information Network (EMWIN) EMWIN i s a direct service that provides users with weather forecasts, warnings, graphics, and othe m the NWS in near real time. The GOES EMWIN relay service is one r information directly fro of a suite of methods to obtain these data and display the products on the user’s personal - resolution GOES satel lite imagery data computer. The HRIT service provides broadcast of low and selected products to remotely located user HRIT Terminals. GOES Raw Data Link (RDL) The RDL channel broadcasts raw data coming from the GOES instruments directly down to the Ground Stations. These stations are the NSOF in Suitland, M aryland, and the WCDAS at Wallops, Virginia. Search and Rescue Satellite Aided Tracking (SARSAT) As an integral part of the COSPAS - SARSAT international search and rescue satellite program , NOAA operates the SARSAT system to detect and locate mariners, avia tors, and other recreational users in distress almost anywhere in the world at any time and in almost any condition. This system uses a network of satellites to quickly detect and locate distress signals from emergency beacons onboard aircraft, vessels, an d from handheld personal locator beacons s series satellite ( PLBs ) . The SARSAT transponder that will be carried onboard the GOES - R 9 - 4

121 provide the capability to immediately detect distress signals from emergency beacons and relay - calle d local user terminals. In turn, this signal is routed to a SARSAT them to ground stations mission control center and then sent to a rescue coordination center which dispatches a search and rescue team to the location of the distress. GOES - R continues the legacy Geostationary SA R (GEOSAR) function of the SARSAT system onboard NOAA’s GOES satellites which has contributed to the rescue of thousands of individuals era R - by being able to in distress. The SARSAT transponder was modified slightly for the GOES operate with a lower uplin k power (32 dBm), enabling GOES - R series satellites to detect weaker - SARSAT System is shown below in Figur signal beacons. An overview of the COSPAS 4. - e 9 - SARSAT System Overview Figure 9 - 4. COSPAS 9 - 5

122 Tracking, Telemetry, and Command (TT&C) The C subsystem provides telemetry, tracking, and commands through orbit raising, orbit - TT& station operations, and on - station contingency. The TT&C raising contingency, normal on - subsystem can be configured differently depending on the particular phase of the mis sion. Flexibility has been designed into the architecture in order to ensure maximum capability and functionality throughout the GOES - R series mission. The on - board TT&C assemblies consist of each a series of antennas , as shown in Figure 9 - further and electronics b oxes, which are 5, described below. FORWARD OMNI ANTENNA FORWARD OMNI ANTENNA ANTENNA SAR ANTENNA SAR ANTENNA GRB ANTENNA GRB UHF ANTENNA UHF ANTENNA BAND X - BAND - X ANTENNA ANTENNA ANTENNA WING ANTENNA WING ASSEMBLY (AWA) ASSEMBLY (AWA) +Z +Z +Z BAND L/S L/S - BAND - ANTENNA ANTENNA +X +X +X - Y Y - Y - AFT OMNI AFT OMNI GPS GOES GPS - - TO - - TO GOES ANTENNA ANTENNA ANTENNA ANTENNA 5 Communication Subsystem Antenna Locations . - Figure 9 GRB Antenna The GRB antenna, which is mounted on the Antenna Wing Assembly (AWA), transmi ts a dual a multimode horn, polarized - circular The antenna is composed of global coverage beam. polarizer and waveguide to coax transition. The GRB antenna receives processed, reformatted sensor data from WCDAS/RBU stations and rebroadcasts it to a large number of outlying ground This antenna has a ±20 degree field of view cone GRB User Ter minals (GRBT). 9 - 6

123 L/S - Band Antenna The L/S - Band antenna is located on the AWA. This a ntenna transmits and receives a linearly pola rized global coverage beam. It is composed of a high efficiency horn, diplexer and waveguide - to - coax transition. The L/S - Band antenna is used for the DCPR comm service. UHF Antenna The UHF antenna transmits and receives a right hand circularly polarized earth coverage link between the GOES - a 4 - Element UHF R spacecraft and the ground users. The design consists of Helix Array mounted on an aluminum ground pl ane that is excited by a beam forming network. GPS to GOES - R Antenna The GPS Antenna receives a Right Hand Circularly Polarized (RHCP) L - Band signal from the GPS satellite constellation. The antenna consists of a ground plane mounted helix radiating element. BAND ANTENNA - X The X Band Antenna transmits and receives a linearly polarized communication link to and from - the primary GOES - R dedicated ground station at Wallops Command and Data Acquisition Station (WCDAS) in Virginia as well as the remote backup (RBU) facility at Fai rmount, West Virginia. The Antenna consists of a single surface parabolic reflector illuminated by a feed network consisting of horn, orthomode transducer, and two diplexers. The reflector is deployed and pointed by a 2 - axis gimbal allowing orbital slot f lexibility. SAR Antenna The SAR Antenna is required to transmit right hand circular polarized global coverage beams. The Antenna is composed of a ground plane mounted dual tapered helix radiating element. - TT&C Antenna – Forward Omni The Forward Omni Ante nna provides independent hemispherical and toroidal beams. The - Forward TT&C Omni Antenna is a combined Bicone/Crossed Dipole configuration. IT is mounted on a mast extending from the spacecraft’s earth panel. TT&C Antenna – Aft Omni The Aft Omni provides an independent hemispherical beam. The Aft TT&C Omni Antenna is a Crossed Dipole configuration. The Aft Omni Antenna is mounted on a long mast extending from - the spacecraft base panel. 9 - 7

124 TT&C Electronic Assemblies and their functions are as follows: assemblies The communications subsystem electronics Traveling Wave Tube  : amplify and linearize the RF input signals in the Amplifiers downlink band  - B and modulator : provides continuous phase - shift key modulation X X - Band receiver : receives and process two X - Band uplink signals and down converts the  signals as inputs to the GRB L - Band transmitter . The two independent X - Band channels, designated CH1 and CH2 operate in fixed gain mode and are narrowband filtered. UHF receiver : recei ve  and process es two uplink signals at UHF. The two channels, s designated CH1 (DCPR) and CH2 (SAR), are amplified, narrowband filtered, and up converted to L - Band. ransponder  S - Band t : p rovide s command, telemetry and ranging functions during launch he GOES - primarily used during orbit R satellite. It is - and Orb it raising operations of t raising, but is available during normal on , the S - band station operations. In addition - Transponder is used on - station to determine the range to the satellite, to provide dary satellite telemetry on the S - Band downlink and to receive satellite commands secon in the event the CDA command link is unavailable.  S - Band receiver : r eceive s and process es two uplink signals at S - Band. The two narrowband channels, designated CH1 (EMWIN - HRIT) and CH2 (DCPC), are amplified , and UHF band . filtered, and down converted to L  CDAS transceiver : provide s command (CMD) and telemetry (TLM) functions during On Station operations of the GOES - R satellite. The CDAS t ransceiver is comprised of a receiver and Telemetry Transmitter constructed as a single package Command The . CDAS Command Receiver receives and demodulates a - Shift Key ( BPSK ) Binary Phase direct - modulated RF command signal and provides the demodulated digital data and clock for processing by C&DH. In addition, the CDAS Telemetry Transmitter receives digital data which is sued to modulate and transmit a BPSK direct - modulated RF output. provide a reference signal to the communications : 10 Mhz reference oscillator  e a highly stable frequency source. The 10 MHz subsystem components that requir Reference Oscillator is a standalone device which generates a stable 10 MHz output. 9 - 8

125 Command and Data Handling Subsystem 10. The Command and Data Handling (C&DH) Subsystem is responsible for gathering, formatting, s oftware data throughout the spacecraft. It provides several platforms for ing the f light and deliver to execute and serves as the validator and formatter for all ground communication. The C&DH 1 . The C&DH is comprised of the following - Figure 10 subsystem block diagram is shown in components:  Command and Telemetry Processor (CTP)  On Board Computer (OBC) Remote Interface Units (RIU), quantity 4   Sun Pointing Platform Interface Unit (SIU)  Command Decryption Unit Assembly (CDUA) nt Sensor Unit (CSU)  Curre Transient Suppression Unit (TSU), quantity 7  I B A L G M G G N & C & N C / I F I / F S D V L L S V D 5 5 1 - D T S - L I B 3 M p S W p S W S A D E o o t t ) a b t u O ( d r R L D L R D S l i p R g n i s a t C o O n B o r d m p u e r S p a c e S U V I C O B B - A O C - B W i r e S E I X f r c e c a p S t a n d a m n t I u s e n t r p S W p S W p S S p W W r y t T e l e m e g g n n l l i i r r o o r r o o t t t t i i W p S W S p n n n n o o o o S u n C C M M e R e m o t d d n n C C i n i t n o P g a a B B r e t n I e c a f O O f r o m t P a l t i n U S U T r I a c e f t n e ) U I R ( i n U t ( 4 ) ) U I S ( C B - P T C T P - A a e a p S c r c f t W S p P r o c e s s o r C o m m a n d a n d T e l e m e t r y s C m m a n d o a n d n C o l t r o u t n e r C r l l 2 2 o S s r e n e e 2 2 v v e m e l e T y r t 4 4 e e l l S U ) U n i t ( C i S i S A U D C B B R R R R y a l e y a l e e w P o r m n d C o m a r D e e v i r D v i l o r y P a y e R p o t n i D e y r c u n d a n t R e d y P o r s C m d m C s d y b m e s s A l P H S K P H S K t e s m b l n U y s A i F I / r S p a c e e i W T R O O C & T T R C & T A R ) 2 ( ( P ) I F / I / F S W ( ) p S I S S E R C&DH - Block Diagram of the GOES Figure 10 - 1. 10 - 1

126 Command and Telemetry Processor (CTP) T he CTP is the primary gateway for all uplink commanding and downlink state of health telemetry communications in the GOES system. The CTP receives digital command data from the CDAS or ORTT&C uplinks and validates the stream using CCSDS standa rds subsystem via the before the command is allowed to be executed. The CTP is a dual sided unit and operates in a hot - hot configuration. Each side of the CTP continually monitors all four uplink paths from the communications subsystem and is always command receptive on all channels. The CTP also contains discrete commanding capabilities to turn various components on or off via the Relay Drive Card (RDC). telemetry As well as generating C&DH State of Health telemetry packets, the CTP receives FSW (a spacecraft communication network data from the On Board Computer (OBC) via SpaceWire . The CTP formats all state of health based in part on the IEEE 1355 standard of communications) downlink telemetry in the proper CCSDS format and then delive rs the CCSDS transfer frames to CDAS the and ORTT&C for downlink. T he CTP’s downlink is selectable at 1, 4, and 40 kilo - symbols per second (ksps). also and can reboot or reconfigure OBC The CTP is responsible for monitoring the health of the fault is detected. This function is handled by the Redundancy Management Card the OBC if a (RMC). The RMC receives two types of recurring heartbeat signals from the OBC, discrete and configurable critical bus heartbeats. If either of the heartbeats are not received by the RMC in a amount of time, the RMC will first attempt to reboot the ailing OBC and if this fails, the RMC will command the standby OBC to operational and FSW will boot on the new processor . 2. The Command and Telemetry Processor (CTP) Figure 10 - 10 - 2

127 On Board Computer The OBC is an internally redundant component that provides processing resources necessary for FSW to gather and route spacecraft component and instrument commands and data. In addition, and GLM instruments and functions as the OBC provides direct SpaceWire interfaces to the ABI ion terminal for all instrument data for downlink to the ground through the Raw the data collect Data Link (RDL). The OBC acts as the 1553 Bus Controller for the spacecraft, communicating to all of the Remote rackers among other Interface Units (RIU), Solar Array Drive Electronics (SADE), and s tar t components. There is a SpaceWire link between the OBC and CTP where receives FSW to the validated uplink commands and data files as well sending spacecraft telemetry packets C TP for downlink. The OBC also has direct low latency connections to the Inertial Measurement Units (IMU) and Global Positioning System Receiver (GPSR). This data is relayed to the FSW for constant attitude and pointing calculations to be maintained. The Sp aceWire Router Card (SWRC) within the OBC receives all of the instrument science data and formats it for downlink through the RDL. It is also the source for the entire SpaceWire network he data rates a re either 132 for all of the instruments on the spacecraft. Depending on the link, t Mbps or 10 Mbps. All instrument science data is delivered to the SWRC where it is CCSDS formatted, Low Density Parity Check encoded and delivered to the RDL for downlink at a rate of 120 Mbps. The On Board Computer (OBC) Figure 10 - 3. 10 - 3

128 Remote Interface Unit (RIU) and Sun Pointing Platform Interface Unit (SIU) - There are four RIU s (shown below in Figure 10 4) and one SIU on the spacecraft. The RIU/SIU us data for distribution to the component provides for reception of comm ands over the 1553 b The RIU/SIU also, upon receiving a telemetry request over other subsystems of the spacecraft. b us, collects, processes, and transmits the data via the 1553 b us to the b us c ontroller, the 1553 the OBC. The interfaces are ill ustrated in Figure 10 - 5 . Figure 10 4. - Remote Interface Unit (RIU) 10 - 4

129 Figure 10 - 5 . RIU/SIU Block Diagram The RIU/SIU is of modular design. Three boards provide the core circuitry necessary for each RIU/SIU: an Electronic Power Converter (EPC) board, a Control & 1553 Board, and a Harness Board/Backplane. The modular design is accomplished by providing a standa rdized Harness Board interface which allows for the necessary combination of boards tailored for mission requirements. The RIU/SIU common major functions and operational features are: - STD - 1553B remote terminal function  MIL Collecting and processing satel lite telemetry in response to a request received via 1553  us b  Providing requested telemetry consisting of analog, passive, digital logic level, and digital relay status back to the bus controller  Issuing relay drive commands distribution to onents pacecraft comp s Providing heater control circuits to switch the  s pacecraft 70V bus to the heaters  Providing SpaceWire router functionality, interfacing to the SEISS, SUVI, and EXIS instruments 10 - 5

130 The RIU’s additional functions and operational features are:  Thruster [ Rocket Engine Assembly (REA), and Liquid Apogee Engine (LAE) ] control pacecraft 70V bus to the thruster solenoids and thruster heaters circuits to switch the s Motor Drive control function to drive antenna gimbals to the desired pointing direction  Provide Reaction Wheel Assemblies (RWA)  command and telemetry interfaces to control controlling spacecraft attitude and orientation for  Provide command and telemetry interfaces to the Magnetometer elerometer assemblies, used  Provide excitation and telemetry interfaces to measure Acc for determining spacecraft stability  Provide telemetry interfaces to monitor the Coarse Sun Sensor Assemblies, used for determining general spacecraft orientation Provide command and telemetry interfaces to interface with the Fi ne Sun Sensor  Assembly, used for refined spacecraft orientation  Provide AC c urrent to Magnetometer heaters, needed to power special heater assemblies for the Magnetometer instrument - ithium l attery b  Collect l ithium - ion cell bank voltage measurements and provide ion b alancing functionality, used for power subsystem control and conditioning Command Decryption Unit Assembly (CDUA) The CDUA is a standalone component that exclusively interfaces with the CTP. This assembly houses the decryption ASIC that meets Committee for National Security Systems Policy ( CNSSP - 12 ) requirements. Encrypted data is passed to the CDUA by the CTP in 128 bit CCSDS code blocks. Upon successful decryption, the CDUA passes the 64 bit unencrypted message back to ransfer f t rame. Only after successful decryption the CTP where it is reassembled into the CCSDS R s eries of satellites - can any CCSDS processing be performed on uplink transactions. The GOES utilizes decryption (when enabled) only on the uplink. Downlinked telemetry is alw ays unencrypted. There are 16 unique keys per side of the CTP (total of 32 different keys). Each side of the CDUA only interfaces with a single side of the CTP, therefore to send commands through the B side CTP, a different key must be used. The reason for unique keys is twofold. First, by having different keys interfacing with each side of the CTP ensures that both sides of the CTP will not validate an uplink command and send two copies of the same command to FSW. Second, are only valid for a certain period of time. M ultiple keys are per decryption requirements, keys needed to ensure the 15 year mission is satisfied. Current Sensor Unit (CSU) - 6, provides increased perceptibility into faults The Current Sensor Unit , shown below in Figure 10 during the Integration and Test phase (I&T) and Operations. The CSU measures 12 different 70V and 28V power lines which are comprised of 12 primary and 12 secondary channels (24 total). All e independent from the secondary channels so should a failure occur on one primary channels ar channel it would not affect the redundant component. Current passes through the CSU from the 10 - 6

131 Electrical Power Subsystem (EPS) and then is connected to the individual user component. The CSU measure s the amount of current passing through the unit, convert s it to an analog voltage that analog value to an RIU where it is converted to a digital es representation and then pass king to the ground. representation and passed back to FSW for packaging and downlin Figure 10 - 6. Current Sensor Unit (CSU) Transient Suppression Unit (TSU) The TSU - 7, interfaces directly with the harness and spacecraft , shown below in Figure 10 structure. The TSU is not designed to be internally redundant. The spacecraft system has the responsibility to assign the TSU channels to meet system fault tolerance requirements. The TSU is a passive ex tension to the harness. The TSU cannot be commanded, does not provide telemetry, and does not draw nor dissipate electrical power. There are four different spacecraft channel types that the TSU can interface with: analog/digital telemetry, low voltage ser 32V discrete relay drives. Each circuit ial telemetry, 70V power, and - - sensitive hardware circuitry on both sides of the TSU interface. type is designed to protect ESD In the event that the spacecraft harness becomes charged, potentially harmful energy wil l attempt to discharge to ground through whatever electrical path it can find. Often, this is through sensitive electronics. The TSU is designed to provide a discharge path that is safe for the spacecraft components. Seven of these TSUs are installed on th e spacecraft and are located throughout the risk hardware. bus so as to provide protection to all at - 10 - 7

132 7. - Figure 10 Transient Suppression Unit (TSU) 8 - 10

133 11. Electrical Power Subsystem erant 70V and 28V T he Electrical Power Subsystem (EPS) provides tightly regulated, fault tol nstruments. The power is always on. It also provides redundant power to the spacecraft loads and i unregulated 28V power to the deployment devices. The architecture is shown i n Figure 11 - 1. The major components are:  provides p rimary power to the s pacecraft Solar Array : Batteries : provide power when solar array power is less than the total s pacecraft load and  i nstrument demand, e.g. during eclipses Power Regulation Unit (PRU) regulates the flow of power from the Solar Array and b atteries  : igure below the nstruments (green boxes in the f pacecraft loads and i ) s to Fuse Board Assemblies (FBAs) : provide over -  current protection to prevent power fault propagation enable and fire relays to the deployment devices Pyro Relay Assemblies (PRAs) house ,  : F S W t t e r y B a a r g e h C C o n t r o l a r r A r a l o S y o A r l a r a y r S S h u n t s 7 0 V o i t a l u g e R n o t c i n s r c e l E y r t e B a 1 t e r y t a B t t l p e D o y m e n / C h a r g e r R s y a l e s r a h c s g i D r e a r y 2 t t e B s u B V w o L l g a e o v t 0 7 n r e d e t a l g u U o P w e r t s e s l u P V 8 2 o M u l e s o d 8 V 2 e p l o y n e D m t s e g u l a t i o n R D v i c e s e u B c E l e c i t r o n s V 8 2 P o w r e i F u s e s F u b e s u t i o n s D i s t r M s e l u d o i w s n U d e h c t i w S e c t i w S d e h c t i w s n U d d e h c t h V 8 2 w r e w o P V r e w o P V 8 2 0 7 r P V 0 7 o e r e w o P Items in green reside in the Power Regulation Unit 1. Electrical Power Subsystem Architecture - Figure 11 11 - 9

134 The EPS uses a direct energy transfer to distribute power efficiently. The 70V bus is regulated to 70V ± 0.6 V. Shunts in the PRU control Solar Array power. Shunts are turned on/off when less/more power is needed to maintain the bus at 70V. When a shunt is off, power from that solar he Battery Charger/Dischargers (BCDs) use T array circuit flows directly onto the 70V bus. b atteries. When more power is needed than buck/boost converters to regulate power to/from the b atteries. The FSW the Solar Array can provide, the BCDs are c ommanded to discharge the attery telemetry. When the battery state monitors of - charge is low, the FSW commands a charge - b rate to the PRU. When excess Solar Array power is available, the PRU automatically charges the atteries up to the c harge rate commanded by the FSW. Regulated 28V power is developed from b the 70V bus using buck converters. The 28V bus is regulated to 29.3 ± 0.6V. With harness V drops, the voltage at the loads is guaranteed to be 28V ± 2 V. Switched or un - switched power i s - b provided to the loads as needed. Unregulated 28V busses are tapped from the atteries to provide pulses to deployment devices via relays in the PRA. During North South Station Keeping Total load power is typically a little more than 4000 W. - ers it peaks at 8000 W. The Solar Array provides 5000 W to 5500 W at end - of (NSSK) maneuv life with no failures. The batteries can support 4750 W for 1.2 hours using only half their capacity. The worst case power margin throughout the mission is 17% with a Solar Array c ircuit failed and 2 ( - The power budget is shown below in Figure 11 a nstruments include i b attery cell bank failed. ABI, GLM, SEISS, SUVI, EXIS, and unication Services include SAR, GRB, ; Comm MAG ; Band Downlink, and TTC RF HRIT/EMWIN, and DCPR, X Spacecraft Support includes EPS - Electronics, Power Distribution Losses, GNC, C&DH, and h eaters). 11000 Summer Solstice NSSK 10000 9000 8000 7000 Vernal Equinox Sunlight 6000 Eclipse 5000 Power (W) 4000 3000 2000 1000 0 Loads Battery Loads Loads Solar Array Solar Array Capability & Battery Capability Capability Spacecraft Support Comm Services, Etc. Instruments Arcjets Battery Charge Power 2. - 11 Figure Spacecraft Power Budget 11 - 10

135 - launch operations , power is applied to the Solar Array circuits by the Electrical Ground During pre , the EGSE power is removed and the Support Equipment (EGSE). A few minutes before launch half hours after launch, the Solar Array is b atteries provide the s pacecraft power. Three - and - a - deployed to provide power during orbit raising. During orbit raising and geosynchronous b atteries supply secondary operations, the Solar Array supplies the primary power while the power during eclipses and peak power events. Generally, the EPS components are hot redundant at the module level within the components. For example , there are three parallel BCDs so that if one fails, the remaining two provide enough capability to adequately charge/discharge the batteries. The b attery charge control in FSW uses redundant charge control methods to prevent over - charge. The EPS is designed to fly through any fault without affecting the 70V or 28V bus regulation. For those faults that require timely action needed to restore corrective action, the FSW monitors telemetry and takes the minimum long term operability. For example if the battery voltage telemetry circuit fails, charge control , based on battery voltage is disabled and control based on amp - of - charge and cell bank hour state - voltages remains active. Solar Ar ray The Solar Array, shown in Figure 11 - 3 , provides the primary power to the spacecraft and was - manufactured by Lockheed Martin in Sunnyvale, California. It is comprised of 6720 ultra triple junction (UTJ) photovoltai The cel ls are wired into 16 separate c cells supplied by Spectrolab. circuits each connected separately to the 70V bus in the PRU. Each circuit has 10 parallel strings of 42 cells wired in series. String isolation diodes prevent a string short from affecting the rest of the circuit. Circuit iso lation diodes in the PRU prevent a circuit short from pulling down the 70V bus. At the end of the 15 year mission the Solar Array produces 4960 W at summer solstice and 5600 W at vernal equinox. - installed on GOES 16 Figure 11 - 3. The Solar Array 11 - 11

136 he Solar A rray is stowed against the s pacecraft for launch. First stage deployment occurs about T s four hours after the l aunch v ehicle in order to provide the required pacecraft is separated from the power for orbit raising. Once in geosynchronous orbit, the fi nal deployment is performed. Batteries 5, Two atteries , as depicted in Figure 11 - b provide power when the load demand solar array exceeds the solar array power, e.g. during eclipses. Each b attery is comprised of 36 Saft VL48E lithium ion cells. Three cells are connected in parallel to form a cell bank, and 12 cell banks are connected in series to form a b attery. Balancing circuits under FSW control apply current to individual cell banks to balancer their voltages. Bypass switches are used to remove a failed cell bank from the electrical path. The nominal capacity of each cell is 48 amp - hour and full charge - ach battery is 6120 watt voltage is 4.1 volt. Total energy storage of e The battery also hour. . A radiator covered with optical contains temperature sensors and heaters for thermal control s olar reflectors (OSRs) ejects excess heat generated during discharge. Figure 11 - 5 : Solar Array Battery - The FSW performs the b attery charge control using redundant methods to prevent over charge in the face of any b attery system failure. The primary method controls the maximum cell bank voltage. When the voltage is low a constant charge current is commanded until the voltage reaches the end of charge set point. The current is then tapered to fully charge the battery without exceeding the set point. Finally, when the taper is done, the FSW commands the balancer circuits to apply a small current to remaining cell banks until each is fully charged. Backup charge control b methods use - attery voltage and amp hour state of c harge (integrated b attery current). The batteries were manufactured by battery charge control functi on is depicted in Figure 11 - 6. The Saft in Cockeysville, Maryland. 11 - 12

137 Charge Current Max Cell Bank Voltage Taper Charge Constant Current Balancing 6 Battery Charge Control . Figure 11 - Power Regulation Unit , shown in Figure 11 atteries to - The regulates the flow of power from the Solar Array and 7, b PRU . s manufactured by Lockheed Martin in Littleton, Colorado nstruments. i pacecraft loads and s It wa The PRU is comprised of the following modules:  Central Distribution Assembly (CDA) : 70V regulation electronics, command/telemetry interface via 1553 bus to the , 70V load ports and current sensors OBC shunts that control the flow of solar array power onto : Solar Array Shunt (SAS) modules  the 70V bus in response to a control signal from the 70V regulation electronics buck/boost converters that control the flow of power to/from the b atteries  BCD modules : in response to a control signal from the 70V regulation electronics 28V regulation electronics, command/telemetry Low voltage Control Module (LCM)  : interface to the CDA, 28V load ports and current sensors :  Low voltage Power Modules (LPMs) buck converters that control the flow of 70V power onto the 28V bus in response to a control signal from the 28V regulation electronics : nstruments and some i Power Distribution Modules (PDMs) power feed switches for the  i pacecraft loads, s nstrument power feed current sensors The PRU communicates with the OBC us. The command/telemetry interface via a 1553 d ata b attery polls telemetry, including module temperatures, module on/off status, s b olar array currents, b attery charge rate commands, on/off commands currents, and load currents. Commands include to each module, and on/off c ommands to the load switches in the PDMs. 13 - 11

138 - 7 . Figure 11 The Power Regulation Unit Fuse Board Assemblies (FBA) 70V 8. The - An FBA is depicted in Figure 11 current fault protection. Two FBAs provide over - and 28V fuses are segregated to eliminate the risk that the two busses are shorted together. . Figure 11 Fuse Board Assembly - 8 A 11 - 14

139 (PRA) Pyro Relay Assemblies , shown in Figure 11 - 9, Two PRAs provide fault tolerant pulses to deployment devices in response to commands from the Command & Telemetry Subsystem. Unregulated 28V power tapped from the b atteries is applied to the input of each PRA. For each deployment device, a latching enable relay con nects the 28V power to a non - latching fire relay that is closed by a command pulse from the C&DH. When the fire relay is closed, the 28V power is applied to the deployment device to actuate it. Figure 11 - 9 . The Pyro Relay Assembly Energy Balance The Solar Array and b attery sizes were chosen to insure that the worst case b attery depth - of - discharge is less than 50% of capacity and that the batteries are fully recharged at the end of any io starts with a fully charged b . The scenar . ry during sunlight operation. About atte 24 hour period pacecraft midnight, the s pacecraft enters eclipse and the b atteries discharge s 36 minutes before to support the s pacecraft load (negative b attery current represent s discharge, positive represents charge is at a minimum. When charge) . At the end of the 72 minute eclipse the b attery state - of - s s attery charging. the b pacecraft exits eclipse, pacecraft loads and olar a rray power supports the s rcjet burn star ts. The s pacecraft local time a n NSSK a At approximately 05:25 b atteries discharge to support the resulting peak load which is greater than the s olar a rray capability. After 40 minutes atteries recharge. b the a rcjet burn ends and the 11 - 15

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141 12. Guidance Navigation & Contro l The GN&C subsystem provides guidance, navigation, and attitude & articulation control for the - GOES R spacecraft. GN&C activities include determining the attitude of the spacecraft, determining the position of the spacecraft, determining the location of desired targets such as the Sun and Earth nadir, providing attitude control for rate damping from the launch vehicle separation residuals, orienting the spacecraft during cruise, communication, and science operation periods, - providing for and controlling t ranslational delta V maneuvers and station - keeping, and managing 1. - momentum of the spacecraft. A block diagram of the GN&C Subsystem is shown in Figure 12 Attitude determination is nominally accomplished using one Inertial Measurement Unit (IMU) and two Star Trackers. Two IMU’s and three star trackers are included in the subsystem, which yros (HRGs) and two provides redundancy. Each IMU consists of four hemispheric resonating g view star tracker, which provides accelerometers. The Star Tracker design is a wide field - of - attitude acquisition from unknown initial conditions and provides attitude updates at up to 20 Hz in track mode. Orbit determination is provide d by a global positioning system (GPS) receiver during the operational orbit. The GN&C subsystem also contains analog sun sensor assemblies, which are used for sun acquisition and contingency operations. and the propulsion system. There are six Attitude control is provided using reaction wheels reaction wheels, and all are nominally operated simultaneously. Each wheel has a momentum - s. The reaction wheels serve as the primary actuators for attitude - m storage capacity of up to 75 N control. The GN&C subsys tem also has responsibility for control of the spacecraft’s gimbals. The - spacecraft design includes 2 axis gimbals for the X - band antenna, a single - axis gimbal for the axis gimbal for the sun - pointing platform. solar panel, and a single - 12 - 1

142 Figure 12 - 1. GN&C Subsystem Block Diagram Sun Sensors - S includes four analog sensor heads R CSSA Coarse Sun Sensor Assembly (CSSA) : The GOE integrated into a single pyramid shaped mounting bracket. The CSSA provides current outputs that can be used to obtain a coarse knowledge of the sun’s position. There are six Coarse Sun sensor Assemblies mounted on the GOES spacecraft and identified as A1, A2, B1, B2, C1, and . The six Coarse Sun Sensors Assemblies support the mission 2 C2 , as shown in Figure 12 - objectives of deter mining the position of the Sun with respect to the spacecraft and to determine the position of the Sun at all times. Four sensors are mounted on the solar array; two on the active side of the solar array, and two mounted on the backside. The other two sens ors are located on a bracket mounted to the spacecraft base panel. The six CSS units are configured as two redundant systems, each system consisting of three CSS units. During eclipse, the Coarse Sun not being in view. Sensors do not converge to a so lution because of the sun 12 - 2

143 Figure 12 - 2 . Coarse Sun Sensor Assembly Orientation R FSSA includes the Fine Sun Sensor Head - The GOES : Fine Sun Sensor Assembly (FSSA) and the Fine Sun Sensor Electronics. The FSSA supports the mission objective of determining the position of the sun with respect to the SPP, and determining the position of the sun with high SSA is a backup sensor to the SUVI Guide Telescope. It is mounted on accuracy. The F the SPP, aligned with the electronics box on the side of - as shown in Figure 12 - 3 . The optical head is co the SPP. Figure 12 Fine Sun Sensor Assembly Mounting Location . - 3 12 - 3

144 R GOES GN&C Performance Requirements - R series Earth - observing The increased spatial, spectral and temporal resolution of the GOES - instruments impose extremely demanding performance requirements on the spacecraft Guidance luding attitude knowledge, Integrated Rate Error (IRE), Navigation and Control (GN&C) design, inc orbit knowledge, pointing, pointing stability, and jitter. The GOES - series attitude knowledge requirements are primarily driven by the instrument R . The spacecraft GN&C subsystem is requ ired to provide inertial attitude knowledge requirements - to the instruments as a time ed attitude quaternion at 1 Hz. The spacecraft is also required tagg - - axis attitude rate data to the ABI at 100 Hz. The ABI uses this knowledge to provide low latency 3 - time control of its LOS. The ABI propagates its own attitude knowledge to achiev e to provide real the GOES - . For other instruments, the attitude knowledge information is used R INR performance during the post processing on the ground. GOES - R derives attitude and attitude r ate estimates - from data from the IMU and The stringent attitude using attitude rate data the star tracker. attitude - knowledge requirements shown in Table 1 drive the spacecraft design to co - locate the Earth - 12 observing instruments with the IMUs and star trackers on the EPP. The spacecraft GN& C is also required to provide orbit position and velocity to the instruments at a 1 Hz rate. Position accuracy requirements are driven by pixel navigation performance , and r ate en the 1 second accuracy requirements are driven by the need to propagate orbit position betwe - orbit board Global Positioning System (GPS) receiver to provide updates. GOES - R has an on data with the specified accuracy. Because GOES - R is a geostationary satellite, this involves tracking extremely low level signals while operating ab ove the GPS constellation. The GOES - R series pointing and pointing stability requirements are comparable to other - pointing missions. However, the GOES precision R requirements apply during spacecraft - maintenance events, such as momentum unloads and station - keeping maneuvers. This “operate - through” capability is unique to this mission. Within the control design, there are a number of feed - - through capability. For example, the ABI instrument forward paths to facilitate the operate edictions of the disturbance forces and torques created by ABI mirror provides to the spacecraft pr motion for use in feed forward co mpensation. 12 - 4

145 - 12 - Observing Instruments 1: Table Summary of GN&C Requirements for Earth Value Requirement Attitude Knowledge Static 1200 μrad 3σ per axis (prior to on - orbit calibration) Slow Dynamic 45 μrad 3σ per axis 30 μrad 3σ per axis Dynamic Integrated Rate Error Sec 1 1 μrad 3σ X/Y axis; 1.5 μrad 3σ Z axis 30 Sec 2 μrad 3σ X/Y axis 2.5 μrad 3σ Z axis 300 Sec 7 μrad 3σ per axis 900 Sec 18.5 μrad 3σ per axis 1 Latency Latency requirement curve as shown in Figure 12 - Orbit Knowledge 3σ 75 m In - Track Position Cross - Track 75 m 3σ Position Radial Position 100 m 3σ Velocity 6 cm/sec 3σ per axis Pointing Accuracy 270 μrad 3σ per axis Pointing Stability, 60 224 μrad 3σ per axis sec Attitude Rate Error 58.7 μrad/s 3σ per axis, based upon 15 ms latency < 120 minutes per year of lost observation time Availability discussed in the previous section, the stringent GOES - R As attitude series spacecraft determination requirements dictate that the IMUs and star trackers be co - located with the Earth - observing instruments ( ABI and GLM ) . The resulting configuration is shown in Figure 12 - 4 , where the placement was driven by the instruments’ field of regard, and the star tracker keep out zones. - GLM is a static staring instrument with no capability to compensate for alignment biases or shifts. As the GOES - R series satellites come on - orbit in the operational configura tion, the GLM line - of - sight is pointed at nadir. With its scanning mirrors, ABI has the capability to compensate for any - - of sight. offset between the ABI and GLM lines 12 - 5

146 - Figure 12 - 4 . Earth Point ing Platform Configuration for the GOES R Series Spacecraft The “operate through” requirement for GOES - R drove the development of Aerojet Rocketdyne’s (LTR) for use during momentum a miniature 0.08 N Low Thrust REA djust (MA) maneuvers and EWSK maneuvers. Th e design trade for these small thrusters involved many factors, including constant and predictable low thrust, high throughput, long life, and design simplicity (the GOES - rom the R design uses 16 of them). The small thrust from the LTRs can be balanced by torque f reaction wheels, which allows continuous firing of the LTRs with minimal spacecraft attitude disturbance. The 0.2 N arcjet thrusters (also built by Aerojet Rocketdyne) are used for ewton the small thrust from the arcjets NSSK because of their high Isp of ~570 sec. As with the LTRs, can be balanced with torque from the reaction wheels, and the attitude excursions during NSSK maneuvers remain within pointing requirements. T he gimbal design used for the azimuth and elevation control of the solar array and SPP incorporates the proven low - disturbance design first implemented on the Mars Reconnaissance backlash harmonic drive with a relatively high gear Orbiter. The design is based upon a zero - - phase brushless motor reduction of 200:1. by a 2 The low disturbance capability is provided driven by a sine drive commutation, which effectively eliminates motor cogging. Additionally, a high bandwidth rate - loop is implemented on the motor rate, which essentially eliminates most of the harmonic drive friction and no nlinear effects. Attitude Determination P erformance utilizes the Northrop Grumman Scalable Space R - GOES For attitude determination, the series Inertial Reference Unit (SSIRU) for the IMU, and the SODERN Hydra with three optical heads for . The design includes 2 SSIRUs with 4 gyros each, but only one SSIRU is powered the star tracker on at a time. The SSIRU’s 4 gyros are sampled at 200 Hz, and the star tracker optical heads are 12 - 6

147 gyro data are collected, filtered, bias - corre cted, and - sampled at 20 Hz. Two samples of 4 axis rate data before sending to the ABI at 100 Hz. Attitude estimation is performed - converted to 3 - using a kinematic 6 tate extended Kalman filter , which combines quaternion outputs from the s star tracker with angular rate measurements from the - state attitude error SSIRU to produce a 3 estimate and 3 state gyro bias error estimate. - As with previous GOES satellites, accurate attitude and rate estimates are critical to INR R - GOES the requirements as they are used in the ground - based motion compensation. For series time ABI mirror control to steer out jitter up to the first - , rate estimates are also used for real . In the GOES - attitude determination implementation, star tracker series , instrument mode R measurements and SSIRU measurements are synchronized with the spacecraft control frame to provide the most accurate attitude estimate possible. The SSIRU plays a key role in meeting the GOES - R series INR requirements. High bandwidth, time mirror low - latency rate measurements are critical for accurate motio n compensation and real - control for the ABI instrument. The IRE requirements specify how much error can be accumulated 12 when integrating measured gyro rates. As shown in Table 1, IRE requirements are specified - over different time windows from 1 to 900 seconds. The 1 second window is completely driven by gyro performance, particularly angle white noise. The other windows are driven by a combination of gyro performance, Kalman filter bias estimation, and stability of the mounting interface. The SSIRU was selected for GOES - R series because of the high bandwidth and low latency of the - 422 gyro data output, as well as the low - noise characteristics of the four hemispherical the RS resonator gyros. The three - head SODERN Hydra Star Tracker is used for attitude measurements, with two heads operating continuously and one serving as a cold spare. The Hydra design provides the capability to synchronize each star measurement with an externally provided 20 Hz reference signal. Star measurements from the mult iple heads are combined within the star tracker software . Each head can track up to 15 stars at 20 Hz . The IRE requirements are the most unique AD performance requirements for GOES - R. To f 200 Hz data from the gyros establish the performance characteristics of the gyros, 24 hours o were collected with the GOES SSIRU mounted to a granite block. Analysis of the gyro - R series data was performed to estimate the angle white noise, angle random walk, and rate random walk parameters for each gyro. The SSIRU mode l in the AD simulation was configured with these model parameters, and the attitude and attitude rate profiles discussed above were used as inputs to the AD simulation. IRE performance requirements were met with considerable margin for the GOES - series SS IRU. R Orbit D etermination Performance A key part of the GN&C component suite for the GOES - R series is the upgraded Viceroy GPS Receiver (GPSR) from General Dynamics coupled with a GPS antenn a designed by Lockheed series program. The antenna Martin. The Viceroy - 4 was developed specif ically for the GOES - R design is tailored for operations in a G . The new GPSR design and custom GEO antenna EO 12 - 7

148 design enable onboard autonomous navigation, which is a critical enabling technology for this mission. A GPS receive r at GEO altitude (~35,786 km) is 15,000 km farther away from Earth than the GPS constellation. Satellites comprising the GPS constellation are designed to transmit signals series have its towards Earth. Therefore, a GEO spacecraft such as those in the GOES - R must GPS receive antenna nadir pointing in order to receive the GPS signals that leak around the Figure 12 Earth, as shown below in - 5 . GPS Signal as Seen by a Geostationary Satellite . Figur e 12 - 5 s atellites. Analysis shows that A GPS receiver performs optimally when it tracks 4 or more GPS this cannot be achieved at GEO when only the GPS mai n lobe signals are used . Because GPS side lobe signals are inherently weak, and because free space path loss at GEO is up to 10 dB compared with low - Earth orbit, tracking GPS side lobes is extremely challenging. - lobe signal power is specified in the GPS system specification. The antenna designers O nly main for the various GPS vehicles (Block II, IIA, IIR, IIRM, and IIF) have chosen slightly different - lobe power requirements. As a result, the side - lobe characteristics vary methods of meeting main depending upon the specific GPS satellite in view, as illustrated in 12 Figure - 6 . A GPS receiver attempting to exploit side lobe information must have the dynamic range to distinguish a low power igh power signal is present . This characteristic of signal from noise while not saturating when a h the GPS constellation makes analysis of GEO receiver availability particularly difficult. 12 - 8

149 Figure 12 - 6 . GPS T ransmit A ntenna P attern in 3D I llustrating S ide L obe S tructure D etail The time to initialize upon power up is a good indicator of the performance capability of a GPSR. Th e GOES - R GPSR acquires a position fix within eight minutes for more than 95% of the cases. Once acquisition is attained, the Viceroy - 4 outputs the spacecraft position and velocity in the - Fixed (ECEF) reference frame at 1 Hz. The GPSR - Centered Earth provided ECEF position Earth - and velocity are converte - d to the International Celestial Reference Frame (ICRF) by the GOES R series onboard software. The ICRF position and velocity are converted to equinoctial elements, which are used to propagate the orbit at 20 Hz and to provide the nadir and orbit normal ve ctors needed by the attitude control system. Pointing Control and Stability P erformance The disturbances affecting low frequency pointing performance include solar array articulation, ABI scan mirror disturbances, and RWA friction, gyroscopic, and zero c rossing disturbances. In addition to these common disturbances, momentum adjust cases also include LTR thruster disturbances (torque and thrust variation). Because many of these disturbances are deterministic and predictable, GOES rward capabilities to improve pointing stability. FSW uses feed fo series - R The gimbal articulation controller uses spacecraft body rate estimates as a feedforward term to steer out spacecraft body motion for the sun pointed instruments. - observing instruments, ABI s For the Earth can mirror disturbances are mitigated through feedforward of the ABI PIFT data from the instrument. Torques produced by the LTRs and arcjets are predicted and fed forward as compensation for those disturbances. Gyroscopic torques due Along with feedforward to the spinning RWAs a re also fed forward through the attitude controller . prediction algorithms, RWA friction and zero crossing disturbances are mitigated by the implementation of a wheel speed controller, which acts on the error between the commanded eed and the actual wheel speed estimate. sp 12 - 9

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151 Propulsion Subsystem 13. - propulsion system provides the means for reaction wheel momentum R The GOES series keeping, relocation, - management, attitude control, station decommissioning, and the velocity change at apogee required for final injection into geostationary orbit. The propuls ion system 1. schematic is shown below in Figure 13 - 1 P 1 V S P H P P 3 V P 7 0 1 V V i L i N 0 4 u q d e n i g n E e e g o p p A 4 e G H e G H P 9 V e n i z a r i H N 2 2 y - o r p B d P k n a t t n s r u s s e r a l r e t s u r h p e l T a n t H A E R e n i z 2 2 N a y d r V 6 P S V 2 F 1 E R t s w A r L o u T h y V v l a N o r m a l l e C l o s e d P y r o 1 R V 1 3 S 9 V S P a l l y O p e n N y r o V a l v e o r m 1 V 1 S P P 5 C h e c k V a l v e P a r a l l e l S V 7 e a v e S e r v i c l V P V 8 a & b 4 V S P P 4 V l v t c h i n g B ( k c e a a a L S 2 1 V L d i r e c - s s e r P u e R e l i e f r X + - X F u e l n s t ) n h w i o o H N N N O O 2 4 i o i o x i d r d i z e r z e x 2 2 4 4 k n a t n a t F i l k r n t k t a e 8 V S r e s L e h g i o s w P H d n a r u S V 6 F 7 P F 6 P H S V 3 t s n e r u s s e r P r e c u d a r O R 1 R 2 O P 3 P i e s R e d u n d a n t R e g u l a t o r S e r P 2 P 2 V P a V P b 2 1 P V b P V a 1 0 S V 1 c i r O f i e L 7 V F 4 S 5 V r h A r c j T e t u s t e r 3 F F 2 o c e d n w E h e r a n p P E U C U E P C L L V L L 1 0 9 V 8 V L c o n d i V L 1 L t i V n t i o n i n g U 2 L L V S 4 1 V 1 5 S P 4 V P 5 V R 3 O 4 O R 0 2 8 2 1 3 4 2 3 9 1 2 3 2 2 7 2 3 2 2 8 6 7 1 9 2 1 2 2 1 5 0 1 5 1 4 1 6 3 1 3 2 2 8 1 0 1 1 9 7 5 6 1 1 - 1. Figure GOES - R Series 13 Propulsion System Schematic It is a dual mode storable p ropellant propulsion system derived from the Lockheed Martin Space Systems Company A2100 AX - class design. The delta - v at apogee is provided by a high performance 450 N (101 lbf) hypergolic Liquid Apogee Engine (LAE) using hydrazine as a fuel and MON 3 as a n oxidizer. Two 22 N (5 lbf) hypergolic hydrazine bipropellant thrusters (HBTs) - are used for relocation, decommissioning, and as a backup to the LAE and use the same propellants as the LAE. Eight 22 N (5 lbf) monopropellant hydrazine reaction engine assemb lies (REAs), configured in half systems, are used for settling burns prior to LAE ignition, attitude control during LAE firings, and relocations. Sixteen 90 mN (20 mlbf) monopropellant hydrazine LTRs, configured in half systems, are used for momentum manag ement and station - keeping. Four - 225 keeping. The low thrust, highly predictable, stable mN (50 mlbf) arcjets are used for station 13 - 1

152 performance of the LTRs and arcjets allows the payload to operate through thruster use. Pictures 2. - Figure 13 of the thrusters are shown in LAE REA HBT LTR Arcjets 13 - 2. Figure - R Series Spacecraft Thrusters GOES The propellants are stored in two cylindrical titanium alloy oxidizer tanks and one cylindrical graphite reinforced titanium alloy fuel tank, shown in Figure 13 - 3. Oxidizer Tank Fuel Tank GHe Tank 3. R Series Spacecraft Propellant and Pressurant Tanks - Figure GOES 13 - All three propellant tanks include internal propellant management devices (PMDs) to control the location of propellant in the zero - gravity space environment and to ensure gas - free propellants o all thrusters over the operational life of the spacecraft. The propellant tanks are are supplied t 2 - 13

153 pressurized by gaseous helium (GHe), supplied from two cylindrical lightweight titanium alloy graphite over wrapped pressurant tanks. Check valves upstream of the propell ant tank prevent - migration of propellant vapors into the pressura nt system. A pressure regulator maintains constant propellant tank pressure throughout transfer orbi Once t for consistent LAE operation. the spacecraft has achieved geostationary orbit, the L AE, oxidizer tank, and GHe tank are isolated by firing pyrovalves closed and the system operates in blowdown for the rema inder of its operational life. One additional set of a normally open and a normally closed pyrovalves allows for a mid - life against propellant release on repressuri zation of the fuel tank. Latch valves provide an inhibit or the ground and allow for isolation of a half system of thrusters on orbit. In line filters in both the - ng clea n propellant to the GHe and propellant systems protect against contamination, ensuri thrusters. Pressure transducers provide pressure telemetry at several points in the system throughout operational life. Fill and drain service valves provide the means for loading propellants and pressura nt into the propulsion sy stem. They are also used as test ports and can be used for offloa ding propellant, if necessary. Various propulsion system components are shown in Figure 13 4. - Pressure Regulator Pyrovalve Filter 13 - 4. GOES R Propulsion System Components Figure - 3 - 13

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155 14. Thermal Control Subsystem The GOES - R series spacecraft thermal control subsystem is designed to ensure that thermal requirements are met for all mission phases from launch to end of life. The GOES - R spacecraft uses Lockheed Martin A2100 heritage techniques such as heat pipes, mirrors, MLI blankets, and heaters to accommodate variations in spacecraft configuration, environmental heat loads, and - degradation of materials to meet these requirements. The thermal features of GOES own R are sh for the deployed operational configuration and - Figure 14 - 1 Figure 14 in 2 for the stowed orbit raising configuration. The +Z axis is Earth facing and the solar array is pointed south and tracks the sun during the operational mission phase. R Thermal Control Features (Deployed Configuration) Series - Figure 14 - 1. GOES 14 - 1

156 - 2. Figure 14 R Thermal Control Features - GOES Passive thermal control features include external and internal thermal materials and hardware. tors the design uses heat pipes, wet mounting of To aid the transfer of waste heat to radia components, and high emissivity coatings. To reduce the flow of heat where needed, the design uses MLI blankets, low emissivity coatings, and low conductivity stand offs. Radiation panels are - mmonia filled heat pipes to enhance the heat spreading throughout the panel embedded with a - and serve to lower hot spots and improve heat rejection into space. Radiators are covered with ct sunlight and optical solar reflectors (OSRs), a high emissivity/low absorptivity material, to reje emit heat to space. Portions of the panels are covered in MLI to reduce heater power consumption while maintaining internal temperatures on - orbit. cal : (1) mechani series the There are two types of regulated heater circuits designed for - GOES - R thermostat controlled circuits; and (2) OBC - controlled circuits. Mechanical - thermostats are used for controlling heaters intended for short durations during the early part of the mission. For example, on deployment mechanism rate dampers for the solar array and antenna shelf. These set points and are usually enabled just prior to deployment. The - circuits have fixed - temperature - - thermostats for survival heater control. OBC - controlled circuits instruments also use mechanical ses where heater control is required; storage, transfer orbit, and are used during all mission pha on - orbit mission phases. These circuits have software defined control characteristics and use feedback sensing from thermistors for control. Control algorithms include MAX/MIN logic, or the - heater can be controlled using a straight duty cycle. All heater control parameters may be adjusted at any point during the mission as long as a command link is available. All heaters are fully redundant, except for the Magnetometer heaters. 14 - 2

157 Finally, bot h types of heater circuits can be manually overridden to force them on or off as desired and are also protected by a fault management system. Most of the spacecraft heaters are magnetically compensated heaters, which are designed to meet the low magnetic d ipole requirements of the M agnetometer sensors. outh” panels) sides to reduce s orth” and “ Equipment panel radiators are oriented to face +Y/ - Y (“ n direct solar heating and be better thermally controlled. Many of the instruments have a SU that is mounted ext ernal to the spacecraft and an EU that is mounted to the inside of the spacecraft on the radiator panels. Both EUs and other electronic boxes take advantage of the thermally nd a controlled panels. Other components mounted to the panels include RWA s , the OBC, a variety of communication hardware like traveling wave tube assemblies (TWTAs). When higher conductivity between a box and the panel is required, components are wet mounted with a high thermally conductive adhesive bond. When a conduction enhancement i s not necessary, components are dry mounted to the panels. Base plates may be treated with irridite for electrical conductivity. Most units and some internal panels are painted black to maximize radiation heat transfer internally to the radiator panels as well as help create an isothermal environment inside. X sides of the spacecraft are blanketed with MLI. MLI blankets help minimize the – The +X and diurnal temperature swing experienced from eclipse and maintain the spacecraft cavity within acceptable tempe ratures, where many of the propulsion components are mounted. The LAE is mounted to the base ( Z) side of the spacecraft and has a heat shield to protect surrounding - hardware from extreme temperatures during firings for transfer orbit maneuvers. The antenna wing assembly (AWA) is mounted on the – X side and faces in the +Z direction after deployment. An array of horns are mounted to the antenna wing and all are covered with RF transparent s unshield blankets to protect the horns from direct solar light. There is also a gimbal controlled X - band reflector, mounted on the +X side of the spacecraft which is blanketed as well. Magnetometer sensors are mounted on a deployable boom mounted to the +X side of the spacecraft that is deployed once in GEO orbit. Each sensor unit has a heater and is covered in MLI blankets to maintain temperatures. GOES - R series batteries take advantage of many thermal control strategies to stay within thermal requirement temperatures. To provide redundant thermistors the set up en sures two thermistors in each of the six heater zones on each battery. Each zone is heater controlled separately to evenly heat the battery as needed. The batteries are mounted to the bottom deck on the – Y side and are completely isolated from the rest of the spacecraft. Dedicated radiators facing – Y direction are required in order to dissipate heat from the batteries. Areas not covered with MLI are shown 3. - Figure 14 in gray in 14 - 3

158 Figure 14 - 3. Electrical Power Subsystem Architecture SEISS Cabinet - alone thermal - X side of the spacecraft as a stand structural The SEIS S cabinet is mounted to the – – assembly with embedded heat pipes. All instruments are coupled to the two radiators on the +/ Y sides of the cabinet via a thermal wet mount. MLI is used to cover the rest of the cabinet as well as the instruments themselves. Heaters and control thermistors are used on the cabinet heat pipes to maintain minimum temperatures. Both MLI and radiator panels are shown below Figure in 14 - 4 . SEISS Cabinet and Sensors Figure 14 - 4. 14 - 4

159 Earth Pointing Platform The E arth Pointing Platform is mounted on top of the nadir deck. The EPP is attached to the nadir deck via four launch lock assemblies that mechanically isolate the EPP from the rest of the re spacecraft. The EPP is the mounting location for two of the GOES - R instruments, shown in Figu 5. - 14 GLM and ABI are both conductively isolated from the EPP via titanium mounting feet. Both instrument SUs are covered with thermal blankets to insulate the instrument from the EPP, other al distortion in the form of diurnal instruments, and space. MLI blankets are used to reduce therm swings and temperature gradients. ABI has a dedicated radiator provided with the instrument which it is attached via LHPs. To accommodate the GLM an isolated, free standing aluminum honeycomb radiator panel with embedde d heat pipes was designed with heater control. GLM is attached to the radiator by both LHPs and thermal straps. Both ABI and GLM have associated EUs that are mounted inside the spacecraft on the – Y equipment panel. To meet instrument he star tracker is thermally controlled with a separate radiator, heat straps, pointing requirements t and heaters. To keep the scalable SSIRUs within temperature requirements they are mounted to a dedicated heat pipe cold plate radiator assembly. - In order to prevent excessive en vironmental heating due to solar entrapment, the volume of space between ABI and GLM is closed out with a sunshield membrane. Figure 14 - 5. EPP and Instruments Wing Assembly Array Solar series , holds SUVI The S A WA is shown below - in Figure 14 - 6. The SPP, unique to the GOES R and EXIS on the solar wing. Both instruments must face the sun at all times which is done via two motor driven gimbals, the SEGA and SADA. In addition, the SEB and the SIU of the sun pointing 14 - 5

160 back of the SPP is a black painted radiator to help emit subsystem are mounted to the SPP. The heat dissipation from the SEB and SIU. MLI blankets wrap around the individual instruments as oke is well as the spacecraft components mounted on the panel between them. The solar array y facing side with MLI. Both gimbals are also covered with MLI except at - ed on the sun also blanket the rotation interfaces and are equipped with heaters. Array 14 Figure Wing Assembly and Sun Pointing Platform - 6. Solar 14 - 6

161 15. Mechanisms - R s eries s atellite s ha ve five appendages that The GOES be stowed and restrained for must launch and later deployed at different stages of the mission. These appendages are listed below: W ing Sub system  Solar Pointing Platform (SPP) (SWS) and Sun  Antenna Wing Assembly (AWA) - and Reflector Antenna (X b X  Band) -  Magnetometer Boom (Mag Boom)  Earth Pointing Platform Deployable Structures - Figure 15 - 1. R Series GOES Various mechanisms are needed in order to restrain, deploy and position the GOES - R s eries appendages. These mechanisms are discussed briefly below. Restraint Mechanisms A shear tie is a mechanism that restraints a deployable structure while it also reacts to external loads resulting from transportation of the spacecraft and loads resulti ng from the launch A shear tie assembly is composed of two main sub - assemblies; the actuator and environment. the retraction mechanism. The actuator secures the bolt or cable that keeps the deployable in a 15 - 1

162 preloaded and secured state. Once the actuator is commanded to release, the retraction mechanism ensures that the cable or bolt securing the deployable is retracted out of the way to allow the deployment. Figure 15.2 is representative of the shear ties used to secure the GOES - R s eries a ppendages. Fig 15 - 2. Example of Solar Array Shear T ie ure The GOES - R series use different kind of shear tie actuators depending on the application and preload required to secure the deployable assembly. All shear tie actuators are fully redundant as they contain independent primary and secondary circuits for release. The differen t kind of restraint and release actuators are: Frangibolts: made by TiNi Aerospace Inc.  a o Used on the SWS f rame s hear t ies, X - b and ntenna and AWA shear ties  Separation Nuts (Sep Nuts): made by Eaton o Used on the Solar Array Panel shear ties  Split Spool Re lease Devi c e (SSRD): made by NEA Electronics o Used on the SPP shear ties Deployment and Positioning Mechanisms The deployment and movement of the GOES - R series appendages is achieved by the use of passive hinges, gimbal actuators and stepper motor drives. Hinges The hinges are used during one time deployments and their rotation is controlled using thermally controlled viscous dampers. These viscous dampers can be integral to the hinge as is the case of the Root Hinge (used in the SWS) or the HA , as shown in Figure ed on the AWA) 90 Hinge (us - 15 - 2

163 4 - assembly as is the case of the Solar Array inter - 15 . Also, the dampers can be a separate sub - - 3 . panel hinges and the SPS frame hinges , as shown in Figure 15 GOES - R series contain redundant the springs for uniform torque The hinge assemblies for application throughout the deployment, hard stops and latches for stiff lockout of the deployable assembly once it reaches its final position. and Damper Solar Array Inter Panel and SPP Frame Hinge 3. - Figure 15 15 Figure (right) - 4. SWS Root Hinge (left), AWA HA - 90 Hinge Gimbal eflector to be The gimbal (shown in Figure 15 - 5) r is a dual axis actuator that allows the X - b and unted orthogonally repositioned on orbit as required. The gimbal contains two stepper motors mo .003°/step resolution for each axis. The gimbal contains redundant course and fine t provide tha potentiometers for reflector position telemetry. 15 - 3

164 Dual Axis Gimbal Assembly . Figure 15 - 5 Solar Array Drive/ Slip Ring Assembly (SADA/SRA) the Slip is composed of the Solar Array Drive Assembly The SADA/SRA (shown in Figure 15 - 6) Ring Assembly. The SADA allows the Solar Array Wing (SWS) to rotate 360° (in both positive and negative directions) while tracking the sun. The Slip Ring Assembly (SRA) allows power to hru a rotating interface. The SADA be transferred from the Solar Array back into the spacecraft t contains redundant motors and resolver circuits. The SADA is driven via software and by the SADE (Solar Array Drive Electronics Box). Figure 15 - 6. Solar Array Drive/Slip Ring Harness Assembly Pointing Platform El evation Gimbal Assembly (SEGA) and Trailer - Sun Bearing Assembly (TBA) The SEGA gimbal works in conjunction with the Trailer Bearing Assembly (TBA) to point the SPP olar i which contains the GOES s 7 - nstruments , as depicted in Figure 15 . This pointing ability pro vides seasonal position/offsets for the instruments on the SPP as well as the ability to perform calibration scans which are driven by the instruments. The SEGA is a single axis motor with an attached drive train that provides enough torque to move the SPP . The TBA which is located at the opposite end of the SPP rotation line provides a low friction support to the SPP and allows the SEGA to rotate the SPP with high precision. 15 - 4

165 15 - Figure SEGA (top & bottom left) and TBA (bottom right) 7. Solar Array Dep loyment The deployment is completed in two stages. The first stage occurs autonomously within SWS hour s after launch. At this point (r ), only the solar panels are deployed by Figure 15 - 8 eference panel hinges to deploy the solar array - er shear ties and allowing the int firing the six separation nut s y oke and f irst portion of the wing. The remaining SPP, tage shear ties are not fired at this time. Once the satellite has reached final orbit, the remaining Solar Array shear ties (the two Frangibolt ties, the six SSRD Frame shear ties and the two (not visible as they are on the far Frame shear Figure 15 8 are independently commanded to fire in order to side of ) - inal Frangibolt shear ties ) f f the second stage 9 deploy the complete wing assembly. - Figure 15 below shows the locations o deployment shear ties. Once these shear ties are fired the wing separates from the spacecraft. The Root Hinge and frame hinges deploy the wing into its final position. 15 - 5

166 Figure 15 - 8. First Stage Solar Array Deployment Second Stage Solar Array Deployment 9. - 15 Figure After the Solar Wing is deployed, the SADA and SEGA are able to rotate the Solar Wing Assembly 10). - The and SPP via commanded instructions to the desired position (reference Figure 15 s primary function of the SADA and SE GA is to precisely maintain the pointing of the olar un for s nstruments at the s i un. This action by default then also points the Solar Array cells at the spacecraft power generation. 15 - 6

167 15 - Figure SADA and Root Hinge Interface with Solar Wing Assembly 10. X - b and Reflector Antenna Deployment The X - b and Reflector Antenna is held in place using three Frangibolt shear ties. Once the command is given to fire these shear ties , the X - band Gimbal is commanded to deploy the - Figure 15 11. reflector into its final position as show n in ie and imbal ocations Figure 15 - 11. X - b and Reflector Shear T L G Antenna Wing Assembly (AWA) Deployment is restrained using four Similar to the X - b and Reflector, the AWA (shown in Figure 15 - 12) 90 Frangibolt shear ties. - The AWA is deployed into its final position via the deployment of the HA hinge. The AWA final deployed position is 90 degrees from the stowed position. 15 - 7

168 AWA 12. Figure 15 - Magnetometer Boom Deployment is deployed via a single commanded Frangibolt 13) The Magnetometer Boom (shown in Figure 15 - shear tie release. The boom is deployed by its own stored strain energy in its structural elements. 15 - 8

169 Figure 15 - 13. Magnetometer Boom Stowed and Deployed Configuration Earth Pointing Platform Uncaging The EP P contains four identical Launch Lock Assemblies which retain the EPP during s atellite ock assemblies are shear tie type L . The Launch 14) - transportation and launch (see Figure 15 devices that work wi th a SSRD. The Launch Lock Assemblies allow the uncaging of the EPP and allow it to become suspended via a dampened strut/isolation system. Once released, the Launch Lock upper and lower housing separate providing the clearance required to accommodate the EPP isolation system. Figure EPP and Launch L ock Assembly 15 - 14. Structures R series spacecraft structure is based on the Lockheed Martin A2100.The core The GOES - structure consists of honeycomb structural panels, which form a box to support the propulsion 15 - 9

170 The core has an integral adapter ring that provides a mating 16) - system (see Figures 15 - 15 and 15 - pipe embedded panels are interface to the launch vehicle. In parallel, three thermal heat assembled to support the system module integration. Once these two structures are mated, tures are added to support the primary instrument payloads. A honeycomb EPP additional struc is mounted on top of the primary structure and also carries star trackers and the inertial measurement units. A honeycomb cabinet is mounted to the base panel for instrument Finally, an articulating honeycomb SPP is mounted to the bus structure just accomm odation. - structures have a mass of 608 kgs. series inboard of the solar array. The GOES R 15. Spacecraft Primary and Secondary Structures Figure 15 - 15 - 10

171 - Spacecraft Primary and Secondary Structures Figure 15 16. 11 15 -

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173 Ground System Architecture 16. - R The GOES eries g round s ystem (GS) consists of the following: s  GOES - R s eries c ore GS  s ystem Antenna Tools and simulators hosted or integrated in the GS   GOES Rebroadcast (GRB) s imulators  Product Distribution and Access (PDA) components to satisfy GOES - R Access Subsystem requirements The GOES eries GS operates from three sites. The NOAA Satellite Operations Facility s R - (NSOF) in Suitla houses the primary Mission Management (MM), and selected aryland, nd, M Enterprise Management (EM), Product Generation (PG), and Product Distribution (PD) functions, including the Environmental Satellite Processing Center (ESPC) PDA capability. The Wallops (WCDAS), located in Wallops, Virginia Command and Data Acquisition Station , provides the primary Radio Frequency (RF) communications services, EM and MM functions, and selected PG and PD functions. The third site is a geographically diverse Consolidated Ba ckup facility (CBU), located at Fairmont, W est V irginia . It functions as a completely independent backup for the MM and selected PG and PD functions for the production of Key Performance Parameter end products capable of concurrent and remote operations (KPPs) and GOES Rebroadcast (GRB) data, and is - from NSOF and WCDAS. The CBU has visibility to all operational and on orbit spare satellites. The KPPs consist of the Level 2+ (L2+) Cloud and Moisture Imagery (CONUS, ull d isk, and f provides an overview of associated sectorized products. Figure 16 - 1 Mesoscale) product and the the GOES - R System and GS. In addition to the operational sites, two Operational Support Locations (OSLs) have been implemented to support GS sustainment and maintenance activities and reso lve anomalies. The first, known as OSL1, is located at the development cont ractor facility in Melbourne, Florida . From OSL1, support personnel can access the system at all 3 sites. From the second site (OSL2), Climate Prediction (NCWCP) in College Park, located at the National Center for Weather and , support staff and government personnel are able to access the NSOF L2+ product M aryland generation and distribution capabilities. 16 - 1

174 Figure 16 - 1. GOES - R System and GS Overview The satellites are commanded throughout their mission lifetime from the NOAA Satellite Operations Control Center (SOCC) located at NSOF with the ground station RF interface located nd CBU, and at WCDAS and CBU. The engineering telemetry streams are received by WCDAS a then ground relayed to the SOCC for processing and monitoring at all locations. In nominal operations, the raw sensor data is received by WCDAS, processed by the PG function at WCDAS to create Level 1b (L1b) and L2+ GLM products. These L1b a nd L2+ GLM products are then rebroadcast through the spacecraft GRB transponder. Additionally, sectorized L2+ cloud and moisture imagery products are distributed directly from WCDAS to the NWS AWIPS, and from there onto NWS Weather Forecast Offices (WFO s ) and other AWIPS users. The GRB data is received at NSOF where the rest of the L2+ products are created. Ancillary data used in generating the L2+ products are ingested from the Ancillary Data Relay System (ADRS). Applicable products are directly distribu the ted to the PDA component of ESPC, which provides GOES NESDIS offices, NWS, R Access Subsystem (GAS) functionality, and provides data to - CLASS for long term archive and access supporting retrospective users of GOES data and other GOES data users. e CBU, the raw sensor data, as well as GRB, is received through its RF interface and At th processed by the PG function. The CBU is limited to the production of data to support L0, L1b, and L2+ GLM generation in support of the creation and distribution of GRB, a nd the production of sectorized KPPs for distribution to AWIPS. The CBU is an always - on “hot” backup in order to minute failover in support of high availability mission allow the GS to meet its requirement for 5 - operations and KPP generation and distributi on. The GS includes separate development and integration and test (I&T) environments for the purposes of ongoing development and I&T throughout the GOES - R s eries mission. Portions of these environments are located at both NSOF and WCDAS to support local si te development and shows the primary data flows through the system. I&T activities. Figure 16 - 2 16 - 2

175 GS Primary Data Flows 2. - 16 Figure 16 - 3

176 Overview GOES - R series core g round s ystem provides the following high level functions: The  Mission Management (MM), which includes o Space - Ground Communications (SGC) o Telemetry and Commanding o Spacecraft Navigation Mission Planning and Scheduling o Data Operations (DO), including  L0, L1b, and L2+ product generation o o Product distribution to PDA, and via the AWIPS interface o Product monitoring o Product performance monitoring  Enterprise Infrastructure (EI), including Enterprise supervision o o Configuration Management (CM) Data Storage, including o Mission life data storage  5 - o day (revolving temporary) data storage for CCSDS t ransfer f rames o 2 - day (revolving temporary) data storage for other data products, intermediate products, and related data o Life of mission storage for command and telemetry data Network Management  Mission Management (MM) Element The Mission Management (MM) element provides the capabilities needed for satellite operations. shelf (OTS) products, The MM element architecture uses both custom - developed and off - the - including the OS/COMET® stems. MM can operate roduct for satellite ground control sy s oftware p without any dependence on the other GS elements. Control and status for ground equipment is provided through device - specific drivers that interface with the equipment and integrate the use s ystem is of those drivers into the OS/COMET toolset. C ontrol and status for the a ntenna integrated with OS/COMET. 16 - 4

177 Mission Management Element Functions The Mission Management element provides the following functions that are directly related to user and operator mission operations:  Mission Operations : Handles satellite command and command verification, flight software maintenance, ground directives, and anomaly responses. It also handles satellite telemetry monitoring and processing and provides control and status of ground equipment. OS/COM ET is used to accomplish the mission operations capabilities.  Spacecraft Navigation (SCN) : Performs Orbit and Attitude (O&A) determination, compares O&A solution to that generated by the spacecraft, generates orbital event spacecraft maneuvers. Spacecraft O&A and orbital events are used times, and plans to support instrument planning and operations. SCN is comprised of the OTS product - R series Focus Suite, which has been tailored and configured to support the GOES mission.  Mission Plannin g and Scheduling (MPS) : Handles mission scheduling and planning for all satellite activities and handles integration with ground maintenance schedule. - MPS is composed of custom developed software that integrates with both OS/COMET and Focus Suite .  Space - end round Communication (SGC) : Using a set of modems and front - G - processes satellite telemetry data used for processors, the SGC ingests and pre satellite health and status monitoring. MM also ingests the X band raw data and - processes it to remove higher - el Consultative Committee for Space Data Systems lev (CCSDS) protocols; then outputs Command and Data Acquisition (CDA) telemetry, supplemental instrument data, and science data packets to PG. MM Storage :  Coordinates logging and storage of MM operational data, including storage that is available for secure remote access. MM Storage makes data available to the CASSIE (Contextual Analysis for Spectral and Spatial Information) engineering analysis tool .  GRB Data Routing : Receives GRB data from PG and routes it t o the a ntenna sy stem for RF distribution. R  : Provides the monitoring and controls of the GOES - Antenna Monitor and Control antennas. Data Operations Description The core GS Data Operations (DO) func tions are comprised of the PG and Product Distribution (P D elements. These functions include:  L0, L1b, and L2+ Processing  PG Infrastructure and Service Management  Product Monitoring and Product Performance End Product sectorization and re projection, and formatting -  16 - 5

178  Distributing end p roducts (per current PD configuration) and non - products (e.g. ancillary data, algorithm packages)  Formatting Intermediate data files for PG and storing them in the 2 - - Store (2DS) Day Processing retransmission requests for PDA  Product Generation Element Functions The PG element generates L0, L1b and L2+ products from each GOES - R series operational satellite on a continuous basis, meeting the applicable product latency requirements. PG s received at all three GOES continually processes data as it i - gh the PG R GS sites. Althou functions are fully automated, the Ground System Product Operator can monitor generated product processing and quality. The PG element functions fully support a satellite’s ABI, GLM, SEISS, EXIS, SUVI and Magnetometer instruments when in its oper ational slot, and alternately supports a satellite’s SEISS and Magnetometer instruments when in its on - orbit storage slot. To enable the high throughput, low latency required for DO, a solid state distributed memory cache ata fabric”). This distributed memory cache provides high is utilized (referred to as the “d throughput, low latency, flow control, fault tolerance, and linear scalability with direct access to its contents from computer servers across the GS. All of these features are key to satisfying pro duct latency, operational availability, data delivery, and scalability requirements. Product Monitoring monitors the GRB data downlink products with respect to radiometric and ional insight geometric performance and reports status. Product performance provides for addit into the performance of L1b and L2+ data operations. Product Distribution Element Functions The P D element provides near - real - time continuous network distribution of GS products and data. WCDAS, and the CBU. The CBU and WCDAS PD functionali locations: NSOF, ty is at all three GS provide selected PD functions for the generation, formatting and distribution of sectorized products to the NWS via the AWIPS interface and via GRB. In addition, the CBU is capable of standalone PD operations for a li mited set of products to enable generating, formatting, and distributing products when one or both of the other two facilities (WCDAS and NSOF) are inoperable. PD provides configuration displays through a i nterface (GUI) that allows an ser g raphical u rator to monitor delivery status, configure message filtering, configure PD products for AWIPS, ope or configure which products are being sent to PDA. Any authorized operator’s console can be configured to show the PD displays. PD also provides key storage fo r the core GS in the form of the 2DS and the mission life store (MLS). The 2DS provides for retransmission to PDA if an error in transmission occurs. Instrument c alibration data sent to PDA comes from the PD MLS . The 2DS can also be accessed by PG to extra ct, if needed, operator - selected products for use in support of analyzing anomalous conditions. PD moves data into the PD MLS automatically as well as manually. Selected e nd along with histories of ground dir ectives and p roducts are automatically stored in the PD MLS security events. 16 - 6

179 Enterprise Infrastructure The Enterprise Infrastructure (EI) capabilities of the GS are comprised of the Enterprise Management (EM) and Infrastructure (IS) elements. Together these elements provide the core GS monitoring and c ontrol capabilities. The IS element contains the majority of the core GS hardware. EI also provides the GS with its security monitoring, access control/authentication, and public key infrastructure (PKI) functionality. - site Network Communication Inter NOAA Science Network (N - WAVE) provides inter The site communication services for GOES - R - data transfers between NSOF, WCDAS, and CBU, as well as providing communications between ci WCDAS/CBU and the AWIPS Network Control Facility (ANCF) and Test Network Control Fa lity (TNCF) in Silver Spring, Maryland , and the AWIPS Backup Network Control Facility (BNCF) in est V Fairmont, W . In addition, N - WAVE provides communication support between the irginia system and the GS factory, and between the system and NCWCP. These ser vices are provided R services are - etwork (WAN). GOES n rea through a m ulti - p rotocol l abel s witching (MPLS) w ide a NOAA Office of Satellite and Product Operations ( OSPO ) MPLS WAN. Path a part of the - availability between the GS and AWIPS is specified at 99.9%. A high level overview is shown below in Figure 16 - 3. - 3) ( carrier (OC 155.52 Mbps) links are in place for direct In addition to N - WAVE, a set of optical - between sites to support mission operations. These circuits provide a - to - point connectivity point parallel path via a separate, independent network service provider to help increase reliability for - critical MM inter site network traffic. R Series GS N etworks - Figure 16 - 3. Inter - site GOES 16 - 7

180 GOES - R Series GS Environments In order to support concurrent operations, integration and test, and development activities, the core GS is segregated into three environments. Each of these environments spans sites and security boundaries to meet GS functional requirements. In addition, the core GS edge provides R data, and allows for authorized series the interfaces to external systems that receive GOES - n users to remotely access GS edge resources via the NOAA v irtual p rivate etwork (VPN) the GS environments across sites, security an overview of capabilities. Figure 16 - 4 provides zones, and functional elements. 4. Figure 16 - GS Environments by Site and Element Operational Environment GOES The GOES - R series R series Operational Environment (OE) supports the operational - mission. OE functionality spans all sites. In the Satellite Operations Zone (SOZ) (within the NOAA 5050 security boundary), all three sites provide operational mission management functionality. WCDAS and CBU provide data operations for L0, L1b, and select ed L2+ processing for data distribution via GRB and to the NWS via the AWIPS dedicated interface. At NSOF, the Product Processing Zone (PPZ) OE receives L1b data via the GRB stream and processes it further into L2+ end products, which are distributed to th e user community via the ESPC PDA system. The OE is tightly configuration managed and changes to the OE are developed, deployed, and tested in non - operational environments before moving to the OE. y zones via the GOES R edge. - OSL1 accesses all three sites and both SOZ and PPZ securit OSL2 only has access to the PPZ in support of NSOF L2+ data operations. 16 - 8

181 Integration and Test Environment ITE c onsists of capabilities to verify the element functionality prior to deployment in the operational ITE capabilities are implemen ted at the NSOF and WCDAS sites. environment. The The ITE only has one PG/PD (DO) data processing string. As a means to test the distribution of products to ty to the PDA ITE and the TNCF with the full operational load, the ITE includes the capabili simulate the distribution of products from two satellites concurrently making use of the data generated from the one available PG/PD data processing string. The Core GS ITE supports realistic test activities by mimicking the OE. This realistic testin g extends beyond the c ore GS to the external interfaces (AWIPS, Level Zero Storage Service ( ) , and PDA) integration and test environments. Thus, GS software and hardware LZSS modifications can be tested across the end - to - end system prior to being deployed t o the operational environments. Development Environment The GOES - R Development Environment (DE) consists of two physically isolated sets of hardware that do not directly communicate electronically with each other. The two DEs are separated by site and are in different security domains. The SOZ DE resides at WCDAS with additional SOZ workstations at NSOF. Note there is not a DE at the CBU. The PPZ DE resides at NSOF. The PPZ DE may also be accessed by authorized users remotely through the NOAA Network VPN to employ read/write/execute transactions, but not for data upload ) NOC ( Operat ions Center or download. Remote Access GOES - R series GS provides a set of resources that are accessible to authorized users who The - R Edge. This remote access capability allows are not physically located at a GS site via the GOES users with accounts and permissions to utilize a limited set of GS resources, whil e not within the physical boundaries of the GS facilities. Using remote access, users have the ability to download data to their own computing resources, view data and manipulate data using tools provided by the core GS. Capabilities available via remote a ccess include:  Access to Level 0 data via the LZSS using the secure file transfer protocol (sFTP)  Access to telemetry data and the CASSIE engineering analysis tool via Mission Management servers located in the edge at all three GS facilities  r MM data such as ground directives, command procedures, schedules, Access to othe and INR reports  Access to the ABI PLEIADES (Post - Launch Enhanced Image and Data Evaluation PLT System) tool to support ABI post - launch test ( ) and anomaly resolution activities (limited acc ess to ABI vendor only)  Access to the PPZ DE in support of the implementation and assessment of potential changes and upgrades to the GS L2+ DO capabilities Remote access is accomplished via obtaining access to the NOAA NOC VPN, and by obtaining system account with remote access permission. Remote access is constrained by the a GOES R - available bandwidth between the GS and external partners, as well as by the number of concurrent remote user sessions permitted by the NOAA NOC. 16 - 9

182 Level 0 Data Products - R L0 data is a collection of CCSDS packets for each instrument collected over a period GOES of time unique for that instrument. This consists of reconstructed unprocessed instrument science data and instrument engineering data packets at full resolution, as sent by the instrument, with all communications artifacts (e.g. synchronization frames, communications headers) removed. These packets are extracted from the multiplexed packets within the transfer frame. looking instruments (SUVI, - This data comes from E arth - looking (ABI pace and GLM ) and s EXIS, MAG and SEISS) at full resolution and includes science, engineering and diagnostic data along with their instrument calibration parameters. L0 data also includes Orbit and Attitude (O&A)/Angular Rate (OAR) telemetry da ta, containing orbit ephemeris and satellite position extracted from selected telemetry packets. Each instrument receives OAR data as a part of its L0 data stream and includes this data as a part of its L0 product files. Level 1b Data Products The L1b pro duct is composed of GOES - R Level 0 (L0) data with radiometric and geometric corrections applied to produce parameters in physical units. It includes calibration tables and addition associated metadata as developed by the GOES - R Product Generation (PG) software. In to being a standalone product, the L1b product is an input into Level 2+ (L2+) product processing. GRB Content GRB is a 31 Mbps direct readout broadcast that replaces the 2.1 Mbps GVAR legacy format. It b products from all instruments, L2+ GLM and contains a set of products consisting of L1 associated metadata, and GRB Information (INFO) Packets containing satellite operations schedules, status information and orbit state vectors. GRB INFO packets also include semi - static L1b algorithm calibration parameter tables which are transmitted after an update. GRB is sent to - R satellites from WCDAS for rebroadcast to the GOES the GOES - R sites and GRB users. Back - up GRB transmission capability is available CBU site. The GRB is received at the NSOF, from whic h L1b, L2+ GLM and GRB INFO Packets products (including L1b algorithm calibration parameter tables) are recovered, and the remainder of the L2+ products are created, as well as sectorized Cloud and Moisture Imagery (CMI) products GRB. A dual circular polarization is used to accommodate the 31 Mbps A simplex link is used for data rate within a frequency bandwidth constraint of 12 MHz using a standard downlink modulation at 1686.600 MHz (L band). The GRB processed instrument data source is packetized - 1), complian - B - t with CCSDS Space Packet Protocol standard (CCSDS recommendation 133.0 and utilizes lossless data compression to fit within the allocated bandwidth. Level 2+ Data Products GOES - R L2+ products include all Level 2 and higher products. Level 2 refers t o derived environmental variables (e.g., sea surface temperature) at a comparable temporal and spatial resolution to the Level 1 source. L2+ includes data or retrieved environmental variables which have been spatially and/or temporally resampled (i.e. der ived from Level 1 or 2). Such resampling may include averaging and/or compositing. L2+ can also include model output or results from analyses of lower level data (i.e., data that are not directly measured by the instruments, but are urements). derived from these meas 16 - 10

183 GOES - R L2+ products may be distributed to end users either via the ESPC PDA interface or via the direct interface to the NWS (for sectorized cloud and moisture imagery only). GOES - Overview R Series Antenna System - The GOES s eries antenna system is part of the comprehensive GS and supports the mission R management element of the core GS. The antenna system includes all components across all three sites required to receive or transmit RF signals to/from the satellites through the Intermediate Freque ncy Distribution System (IFDS) interface demarcation point with the GS. The ground station resources consist of three new 16.4m hurricane - rated (HR) antenna stations at WCDAS, three new 16.4m HR antenna stations at CBU, and upgrades to four existing 9.1m r eceive - only systems at NSOF. Figure 16 - 5 provides a notional view of a 16.4m antenna station at site. N e t M A C , E I T k n i L , B F I S D M n i t o r i n g o U n p w - D d n a b - X d n a b S o n i L k n i A P S S L k A P S S d S v . 1 R - N T S C S M - 5 0 0 - T I E _ T N A s - 16 - GOES Figure 5. Antenna Station R Series The architecture is divided into three major functional subsystems: the Antenna Subsystem, the Control (M&C) Subsystem and the Site Preparation and Construction Subsystem. Monitor and The Antenna Subsystem consists of the components for the 16.4m Antenna (including Antenna band), reflectors, trusses, drive trains, an - , and L - , S - band feed design (X - Control, Tri d pedestal bases), the RF Uplink and Downlink functionality, Data Collection System (DCS), Intermediate Frequency Distribution Switch (IFDS) and Timing and Frequency Reference System (TFRS). The Antenna Subsystem also upgrades existing 9.1m antenna feed as semblies to support the receipt of the circularly polarized GRB signal from GOES - R series spacecraft. The M&C Subsystem 11 - 16

184 In ink m onitoring, Built - In - Test/ Built - l - Test - includes the functionality for Antenna Station c ontrol, System Interface Simulators (ASIS). The Site Preparation Equipment (BIT/BITE), and Antenna and Construction Subsystem consists of the foundation design, power interfaces, HVAC, physical Figure 16 - 6 shows the a ntenna s ystem architecture components security and safety components. facility. at each e G O E S S t n e m g p S e c a E P O N - S G O E S - R G O 5 1 - 3 1 ( ) s h t P l a n g i S a t e l l i t S s e a S P G r o d u c t D h t a P a t P a s e t i l l e t a S T e s e h t t a P l a n g i S t s S a i l l e t d C n a a u t S t o n t r o l s n k i X n d a n d S - - B B a n d u p l a t C S P i l o D S i g n a l G P S D o k w n l i n S P G d n a L - B k n i l n d - B w o d n a B - L d n S a n d , d n a B - X , a n - X - k n i l n w o d d n a B L d n a , d B a B - S , d n a l D o w n i n k o k n D w n l i X a n d u B a - d a n d p l i n k S - B n r ( G R B o ) G V A R N - T N E M G E S D S E O G U O R G R R - S E O G A T N A S G S Y S T E M N N E i l f o o R g n d i u B s p O E M E N T M A N A G E N S I R P R E T E ) M E ( N E T N ) A 3 x ( F R / A N A ) E T F R / N N ( x N A 4 ( x 3 ) N A / R F N E T N A T N A S G R - E A N S E N O G S A N N E T N A G G R - O E S O N A T N A S G R - G N E S E U B S Y S T E S M U Y S S B S T E M E S U B S Y S T M g n i d l i u B s p O i n g u g n i d l i u B s p O p s B O i l d R F T F R T T F R G ( ) P F ) P F G ( o M r o t i n o M t r i n o M o n i t o r C S D & & & l o l o r t n o C C r t l o r o n t n o C D S I F F D S I I S F D U p n l i / k k n i l n w o D / k n U p l i D o n l i n k w E t n e m e g e s i r p r a e n t e r p r i s e M a n a g e m e n t n a M E n t I F t t e r p e n r i E n e m e g a n a M s w n l i n k I F D o M & C F I C & M ) ) E M E ( M ( E M ) ( & C M i s s i o n M a M a g e m e n t n M i s s i o n M a n a g e m e n t n n e m e g a n a M o i s s i M t M ) M ( M ( ) M ) M ( M i D o m 5 n a G O E S G S S - N S S u i t l a n d , M D m V , t n o W r i a F , A d n a l s I s p o l l a W V A s o i t a e p O e t i l l e t a S n A O N F a c i l i t y ( N S O F ) r i c a f R e m o t e B a c k u p ) U B R ( y t i l c a d n a m m o C s p o l l a W D a t a A n q u i s i t i o n S t a t i o n ) S A D C d W ( O C R A - 6 0 0 - T I E S _ T N A d s v . 1 R - W V - 16 - 6. GOES Figure R Antenna System Architecture ground communication functions and sensor data - WCDAS currently provides all primary space the orbit GOES constellation and will perform the same role for - processing for the on R - GOES series . WCDAS house s the antenna suite required for dedicated links to each operational and end equipment to acquire data and to uplink commands and data stored spacecraft, the front - - system. ies ser R services, and the associated network interfaces to provide data to the GOES WCDAS also interface s with and provide s uplinks to the Unique Payload Services for broadcast. Unique Payload Services provides communications support to the Data Collection System (DCS), the High Rate Information Transmission/Emergency Managers Weather Information Network More i Rescue Satellite Aided Tracking (SARSAT). (HRIT/EMWIN), and Search - and - nformation - R Unique Payload Services can be found in the Communications Subsystem on the GOES R and section. The 16.4m HR antennas are fully backward compatible to support both the GOES - legacy GOES missions. - At NSOF, GRB data is received directly from GOES - R Series spacecraft via four 9.1m receive only antennas located on the roof of the NSOF facility. These antennas have been upgraded to 16 12 -

185 receive GRB from GOES - R in addition to GVAR data from legacy GOES spacecraft. Also, the capabilities to perform remote operation of WCDAS functions. NSOF has - R s eries to include a remote site Continuity of Operations (COOP) requirements drive the GOES that provides the critical functions of WCDAS and NSOF through the production and distribution of GRB and key product data. Operation of the CBU is the primary contributor to fulfilling COOP requirements and may also be used to enhance system availability. Although the new GOES - R are antennas at CBU series compatible with legacy GOES, the CBU does not include the ground proces sing equipment to provide backup for legacy GOES satellites. Unless the antennas are being used for testing, training, or are in maintenance, the CBU antennas serve as back up antennas, positioned at the nominal satellite look angles to minimize switchov er - time. The only exception to this would be the need to stow the antennas under high wind conditions. Under high wind conditions, the 16.4m and 9.1m antennas must be driven to the stow position (90º elevation). The 16.4m antennas are specified to operate in up to 110 MPH winds and must be driven to stow in 135 MPH winds. Periodic performance testing can be performed remotely on the RF uplinks and downlinks. The antennas and control systems at CBU will be exercised remotely to verify proper tracking perfo rmance on satellite. These steps ensure that these systems will perform as expected upon switchover to CBU operations. R 16.4m Antenna - Figure GOES 16.7 ools and T S imulators - R series GS. Integrated tools reside in the GOES Three classes of tools interact with the GOES - R GS OE, ITE, and/or DE workstations or servers. These tools are allocated GS requirements 16 - 13

186 and are necessary for system operations. Hosted tools reside in the core GS for mission operations support, post - launch test ing, or calibration/validation activities. These tools are not - allocated requirements, but they support the GOES series mission. Finally, a third set of offline R tools receive GS data but are not hosted within the core GS. These tools may reside at vendor factory sites or on offline workstations at NSOF. Offline tools are not considered part of the GOES - R series GS. Integrated Tools Examples of integrated tools include CASSIE, the GOES - R Parameter Database (PDB) tool, the Custom Object Dump Tool (CODT), an d the Level - 0 Storage Solution ( LZSS ) . These tools fulfil specific mission requirements for mission or data operations. CASSIE provides engineering ) MOST ( and accesses data directly analysis capabilities to the Mission Operations Support Team from the core GS mission life store. The PDB and CODT are used for spacecraft and MAG memory management and interact directly with MM software used to uplink commands and memory loads, and to dump memory files. The LZSS receives L0 data directly from the core GS as net CDF (Network Common Data Form) files. These files are stored permanently for PLT data and for two years for non - PLT data, and they are made available to LZSS users via the NOAA remote access interface. Hosted Tools Hosted tools have been developed by multi ple parties. These tools are not necessary to meet GOES R requirements, but they are useful to mission and data operations teams during various - phases of the GOES R mission. Hosted tools include the GOES - R ABI Trending and Data - Analysis Toolkit (GRATDAT) a nd PLEIADES. These tools are used for post - launch data assessments as well as for long - term or infrequent calibration activities. Hosted tools reside within the GOES - R GS on dedicated servers with operator access from both WCDAS and NSOF. (In the case of P LEIADES, a separate set of hardware has been installed to host the tool). . 8 - Figure 16 Environments for hosting of external tools are shown in 1 6 - 14

187 System Storage and Hosting of Tools - 8 . 16 Figure Offline Tools - line tools reside outside of the GOES - Other off GS at locations such as instrument R series vendor sites or NESDIS’ Center for Satellite Applications and Research (STAR). These tools are - not a part of the GOES R GS, but they may be used by mission operations staff in support of the - R series mi ssion. These offline tools may access L0 data via the LZSS; L1b, L2+, and GOES - term (greater than seven days’ instrument calibration data via the ESPC PDA system; and long worth) data via the CLASS. Externally Provided Simulators tbed (SAST) provides an all Software Tes - The Spacecraft All software simulation of the satellite - f light p roject for use by g round for telemetry and command functions. SAST is delivered to the s ystem contractor to support development of the Mission Management System and to support launch updates. The SAST mission operations rehearsals, pre - launch test activities, and post - - and and S and simulates spacecraft command and telemetry packets associated with the L b b - links. The SAST does not model the encoding, randomization, encryption, or RF tra nsfer associated with spacecraft communications subsystem. light CTP and The f sided OBC/ - provided h ardware in the loop (HWIL) test bed contains a single - can accommodate any combination of the following instrument emulators: ABI, GLM, SEISS, When the instrument emulators are not present, models of the instrument are SUVI and EXIS. provided in the Simulation SW Subsystem. 16 - 15

188 In its role as a Flight Software Development Environment (FSDE) or Flight Software Maintenance Environment (FSME), the HWIL suppor ts the development teams in their integration, test bed test, and verification activities. When used in the SatSim capacity, the HWIL test bed supports spacecr aft integration and test (I&T), operations and maintenance (O&M) command procedure verification, crew trai ning, and on - orbit anomaly resolution. The test bed interfaces with the GOES - R s eries GS at all three sites, and at the spacecraft factory. GOES Rebroadcast Simulators gest and data handling The purpose of the GRB imulator is to allow for on - site testing of user in s systems at GRB field terminal sites. The unit simulates GRB downlink functionality by generating CCSDS formatted GRB output data based on user - defined scenarios, test patterns, and proxy data files. GRB signals in the GOES - R era will replace the current legacy GVAR signal and will have substantially different characteristics, including a change in data rate from a single 2.1 Mbps stream to two digital streams of 15.5 Mbps each. The GRB s imulator is a portable system that - fidelity stream of CCSDS formatted GRB packet data equivalent to live GRB data. outputs a hig h The data is used for on site testing of user ingest and data handling systems known as field - terminal sites. s imulator is a fully self - The GRB s all the hardware and software contained system that include required for operation. The operator manages configurations to edit preferences, define individual test scenarios, and manage event logs and reports. Simulations are controlled by test scenarios, s fy the test data and provide a series of actions for the GRB imulator which are scripts that speci to perform when generating GRB output. Scenarios allow for the insertion of errors or modification of GRB packet headers for testing purposes. The GRB s imulator provides a built - in edit or for managing scenarios. Data output by the simulator is derived from either proxy data files containing L1b or GLM L2+ image test pattern generation commands specified from within data, test pattern images, or non - s imulator outputs a scenario. The GRB packets containing both instrument and GRB Information data. Instrument packets contain data simulated from any of the six GOES R instruments: ABI, - SUVI, SEISS, EXIS, GLM, and the Magnetometer. The INFO packets contain information such as satellite operati ons schedules, status information, orbit state vectors, static unit conversion tables, and static calibration tables. s The GRB imulator provides GRB data as either baseband (digital) or Intermediate Frequency (IF) output to the user ingest and data handl ing systems. GRB packet data is sent in the same two output streams as used in the operational system: one for Left Hand Circular Polarization (LHCP) and one for Right Hand Circular Polarization (RHCP). Use of circular polarization in the operational syste m allows the transmitting antenna to multiplex the two digital streams into the same signal, thereby doubling the available bandwidth. The GRB simulator is designed to be used at any site that receives GRB downlink. ream by generating CCSDS formatted GRB packets. imulator produces a GRB data st The GRB s The operator can configure the setup and runtime parameters and create scripts for the runtime simulations. The GRB Simulator normal operations include: configuration, scenario, test patterns, and proxy files . 16 - 16

189 imulator operates in two modes: online and offline. The online mode is for the actual s The GRB execution of the GRB simulation whereas the offline mode is for editing and configuration activities the GRB s performed by the operator. When a simulation is started, imulator is placed in the online mode. The actual generation and output of the CCSDS formatted GRB packets occur while in online mode. During a simulation, GRB packets are written to a port, making them available to imulator hardware. A new event log is generated for the currently running simulation the GRB s imulator is placed in online mode. In offline mode, event log reports may be each time the GRB s generated and tasks such as maintaining configurations and scenarios may be performed. Offline unctions are not available while the GRB s imulator is in online mode. The GRB imulator is f s placed into offline mode upon termination of a simulation or a user requested stop simulation. s imulator is packaged in two transit cases. The key component s of the GRB s The GRB imulator include: : hosts the GRB s imulator software that produces GRB data and  Simulator Processor provides an interface for the users Front End Processors (FEP) : c  reates Channel Access Data Units (CADU) and sends the transfer frames to the m odem Modem :  outputs the modulated IF signal via Digital Video Broadcasting Satellite Second Generation (DVB - S2) streaming provides system time to the simulator :  Time Code Generator :  KVM ) Switch Keyboard, Video and Mouse ( provides the primary operator interface ESPC Product D istribution and A ccess NESDIS has moved to an enterprise - wide satellite ground processing solution for all NOAA missions. Under the Environmental Satellite Processing and Distribution System (ESPDS) development effort, NESDIS has developed and deployed the ESPC PDA which fulfils GOES - R GS requirements for L1b and L2+ data distribution to authorized ad hoc and subscribed users. access capabilities for PDA provides an integrated solution that includes product distribution and National Polar orbiting Partnership Joint Polar Satellite System , GOES - R, Suomi ) - NPP (S uomi JPSS ) , and Legacy operations, as well as future NOAA and non - NOAA satellite systems. Figure ( 9 16 - shows the GOES - R and PDA systems in the c ontext of the broader ESPC architecture. While the PDA system is not physically part of the GOES - R series GS, it is tightly integrated with delivery requirements to users. it to meet GOES - R series 16 - 17

190 PDA’s Role in the ESPC Enterprise at NSOF 16 - 9 . Figure The new ESPC PDA system serves as a unified provider of NOAA’s satellite data and product offerings. A single system intakes and distributes products for real time users, receiving inbound - product files from multiple product generat ion systems, and enabling distribution to all registered - time ESPC users. The PDA development provides a web portal to end - users of NOAA’s real satellite product offerings, including GOES - R. End users are able to search for and order satellite products, vi a ad hoc requests or subscriptions. The future system will also provide a single interface for product generation system operators to subscribe to and receive ancillary data files. om 10 different data PDA is estimated to make available 30 TB of daily product volume by 2020 fr sources, including GOES R. - 16 - 18

191 Spacecraft Mission Phases 17. - station location in geostationary orbit (station acquisition), the GOES - R To reach the required on series spacecraft undergo four distinct mission phases: off to satellite separation Launch/Ascent  — From Atla s V 541 lift - burns  Liquid Apogee Engine (LAE) : Series of 5 LAE burns to raise perigee and reduce inclination to near geosynchronous orbit drift stop maneuvers with Propellant Thruster (HBT)  Station Acquisition: H ydrazine Bi - ion acquisition at PLT longitude stat Post Launch Test and on station p erformance testing of the payloads to initial operations  Ground Stations Various ground centers and tracking stations are involved throughout the mission phases: • Universal Space Network (USN) stations at Dongara, South Point, and Hartebeesthoek • Diego Garcia Station (DGS), an Air Force remote tracking station WCDA station located at Wallops, VA • CBU in Fairmont WV  Launch/Ascent Phase The GOES - R series spacecraft are launched from Cape Cana veral Air Force Station Space Launch Complex 41 by a United Launch Alliance (ULA) Atlas V 541 rocket. aboard a ULA R (now GOES Atlas V 541 rocket at Cape Canaveral - Figure 17 - 1. GOES - 16) Air Force Station’s Space Launch Complex 41. 17 - 1

192 The Atlas V 541 ascent trajectory utilizes a three Centaur burn extended coast mission profile. The first Centaur burn achieves a parking orbit. The second Centaur burn achieves an interim orbit. There is then an extended coast (current baseline is 2.75 hours from Main En gine Cutoff (MECO) 2 to Main Engine Start (MES) 3) with the third Centaur burn achieving the Transfer Orbit (GTO). Figure 17 - 3 Geosynchronous provides an overview of the ascent profile. Atlas/Centaur performance increases as the extended coast duration inc reases. There is a requirement for upper stage disposal 500 km below GEO. Normally, the longer the extended coast and the closer separation is to apogee, the harder it is for Centaur to maneuver to the disposal - orbit. For the GOES n, the Atlas/Centaur is targeting satellite separation R series trajectory desig for 500 km below GEO. This provides for the ability for the launch vehicle (LV) to maximize the extended coast performance improvement. The satellite raises the 500 km low orbit during the acquisition phase. HBT station 2. 16) lifted off at 6:42 p.m. EST on November 19, 2016 - R (now GOES Figure 17 - - GOES from Cape Canaveral Air Force Station's Space Launch Complex 41, aboard a United Launch Alliance Atlas V 541 Rocket 17 - 2

193 scent Phase Profile . Overview of the A 3 Figure 17 - LAE Burn Targeting - A robust LAE burn plan has been developed for the GOES R series that meets the constraints R spacecraft. Given the separated mass of the and addresses the unique LV targeting on GOES - GOES E burns are required for orbit raising. The driving constraint for - R series spacecraft, five LA the LAE burn plan is the PLT longitude at 89.5° W. The secondary driver for the LAE burn plan is the LV GTO. This orbit drives the target for the last LAE burn. drift stop maneuver plan and also the targeting of the last LAE burn. The The LV GTO drives the injection apogee altitude of 500 km below GEO stays relatively constant during orbit raising. The A) post - LAE 5 perigee altitude is targeted 60 km below GEO for collision avoidance (COL mitigation. The resulting post - LAE 5 drift rate is 3.6° E/rev. This is a relatively large drift rate. 2° W of the PLT longitude with a post Typically LAE burn plans are targeted approximately 1.5 - - - R series spacecraft, with a larger post - last LAE drift rate of approximately 0.6°/rev. For the GO ES last LAE drift rate of 3.6°/rev, the last LAE longitude offset is increased. The drift stop maneuver plan was actually developed first in order to define the longitude offset required for the last LAE burn. There is a four HBT burn drift stop maneuver plan which requires a longitude offset for the last LAE burn of 10.3°. Therefore, the last LAE is targeted at 260.2°E/99.8°W (10.3°W of 89.5°W PLT longitude). s can be targeted. Apogee 3 is chosen With the last LAE burn determined, the first four LAE burn as the first LAE burn apogee to provide the necessary two revs from satellite separation for Orbit Determination (OD) and maneuver planning. The resultant LAE 1 longitude of 99.0°E provides rom Diego Garcia and Dongara. A relatively large 40.8 minute good dual station coverage f maneuver is planned, in part to maintain LAE burns 2 and 3 below the 50 minute burn limit. Any 17 - 3

194 performance dispersions in this first LAE can be readily accommodated in planning the next four LAE burns. The LAE 2 longitude of 61.2°E at apogee 5 provides two revs between maneuvers and provides solid tracking station coverage from Diego Garcia, Dongara, and Hartebeethsoek. The 47.8 minute LAE 2 burn duration is comfortably under the 50 min ute burn co nstraint imposed by p ropulsion. The resultant LAE 3 burn longitude of 331.4°E at apogee 7 again provides two revs between maneuvers and dual tracking station coverage from Santiago and Hartebeethsoek. The y under the 50 minute limit. The resultant 48.5 minute LAE 3 burn duration is again comfortabl LAE 4 longitude of 157.4° E at apogee 9 again provides two revs between maneuvers and provides dual tracking station coverage from Dongara and South Point, Hawaii. A lower duration maneuver time of 30.7 minutes pr ovides a lower 206.5 m/sec ΔV which lowers the burn dispersion magnitude which in turn provides more accuracy in targeting the final LAE 5 burn longitude. A drift rate target of 51.4° is chosen as a divisor of 360° such that for a missed LAE 4 contingency, the spacecraft will return to the LAE 4 burn longitude in 7 rev olution s. Finally, the final LAE 5 burn longitude of 260.2°E at apogee 11 is targeted based on the station acquisition plan. The 24.3 minute maneuver duration imparts a ΔV of 173.7 m/sec whic h provides higher maneuver accuracy for meeting the drift rate target of 3.6 °E/rev. There is a 6.8 day mission duration from satellite separation to the last LAE 5 burn. Additionally, the LAE burn longitudes are targeted at gaps in the GEO belt per stand ard practice to mitigate potential collision avoidance conjunctions with GEO satellites. Given the 500 km below - R. Conjunction analysis will be performed by GEO apogee, COLA at GEO is mitigated on GOES GSFC Flight Dynamics during the LOR period. Targeting the LAE burns for gaps in the GEO belt also minimizes Radio Frequency Interface (RFI) issues. Table 17 - 1 and Table 17 - 2 provides a summary of the nominal LAE burn plan with information on targeting and optimization. The complete GOES 2. R nominal LAE burn pl - - 1 and Table 17 an is shown below in Table 17 - - Table 17 LAE Burn Plan 1 17 - 4

195 LAE Burn Plan Table 17 - 2 HBT Station Acquisition Targeting The HBTs are used after the LAE burns to raise the orbit to GEO and stop the drift rate at the 89.5°W PLT longitude. Station acquisition is also used to describe the post - LAE drift stop d on the 500 km low maneuver sequence. The LV GTO drives the drift stop maneuver plan base apogee altitude. A one and a half day duration between the last LAE 5 and HBT 1 is chosen for time for a good OD and for satellite deployments. HBT maneuvers are performed in the deployed configuration. Four HBT maneuvers provide a robu st plan with good flexibility and orbit/drift control. The LAE maneuvers take out all the inclination and therefore the HBT drift stop track maneuvers. Approximately five days are required to stop the - maneuvers are all positive in drift at the PLT longitud e. Satellite Separation Attitude The satellite separation attitude is optimized to provide both adequate sun angle on the outboard panel of the stowed solar array and adequate coverage from the +Z hemispheric (hemi) antenna. satellite separation. After solar array stage 1 deployment the satellite The hemi antenna is used at goes into Sun Search Mode and is oriented in a sun coning attitude nominally at satellite separation + 45 min, and the hemi antenna is still used. Subsequently, a ground command sets the satellite to the cruise attitude where the bi - cone antenna is used. The satellite separation attitude needs to be rotated from normal to the sun for the outboard solar array due to hemi antenna pattern interference/nulling in the satellite XY plane. L a unch and Orbit Raising Cruise / Orbit Raising Attitude The bi cone antenna is used for launch and orbit raising ( operations after satellite ) - LOR separation, Solar Array stage 1 deployment, and sun coning, all of which use the hemi antenna. cone antenna occurs after the satellite attains sun The transition f rom the hemi antenna to the bi - coning attitude nominally at satellite separation + 45 min and before commanding the satellite to slew to the LOR cruise attitude. The LOR cruise / orbit raising attitude is optimized based on blockage of the bi cone antenna by both the ABI and GLM instruments and for SUVI sun angle - . constraints 17 - 5

196 LOR Tracking Station Network and Contacts For LOR activities, GOES - R has the following tracking station options: • South Point, Hawaii (USN) 204.3°E (155.7°W) • White Sands 253.4°E (106.6°W) 284.5°E (75.5°W) • Wallops (G round Network) • Santiago, Chile (GN) 289.3°E (70.7°W) • Hartebeesthoek, South Africa (USN) 27.7°E ) ir Force Satellite Control Network (AFSCN) 72.4°E • Diego Garcia (A • Dongara, Australia (USN) 115.3°E ) 103.0°E • Singapore ( Kongsberg ( KSAT ) • Hanger AE/KSC (ORTT&C 1K data will originate from here until separation) Post Launch Test spacecraft is After the engineering handover is complete, post launch test begins, and the checked for proper performance before entering service at either of two assigned locations. At the 89.5 W checkout station, the orbit apogee and perigee radii will be at the geosynchronous radius of 42,164 km. By international agreement for t he GOES system, two spacecraft orbital (t positions have been assigned: 75° and 137° West longitudes he latter is a shift from previous ⁰ GOES at 135 W in order to eliminate conflicts with other satellite systems) . From these two - R vantage points, roughly over Ecuador and the Marquesas Islands, respectively, the GOES series instruments cover both the Atlantic and Pacific oceans. The major operations performed upon station acquisition are: • Outgas instrument contaminants • Activate and checkout communicati ons payload data services • Deploy i nstrument cooler covers • Activate space environment monitor equipment • Characterize and optimize i nstrument radiometric performance • Activate and evaluate image navigation and r egistration begin on - station operations • Enter storage mode or 17 - 6

197 18. On - Orbit Mission Operations Each spacecraft in t GOES - R series he is designed for 10 years of on - orbit operation preceded by the active science data orbit storage. This section concerns itself only with - up to five years of on collection aspect of the on orbit mission. On - orbit operations consist of daily (routine) and periodic - operations, both of which are planned in advance and executed as per the operations schedule. Routine operations driven by the on - board schedule include instrument commanding and one housekeeping period (for clock adjustments and momentum dumping) with the spacecraft on - board systems controlling the spacecraft attitude, systems monitoring, and maintaining general th monitoring. operations and heal Mission Operations system is a critical national resource that requires the highest level of mission The GOES - R series practices, rigorous engineering configuration operations support, utilizing mission operations best - management, and extensive development and testing of normal and contingenc y operations procedures. Console operations are continuously staffed at the NSOF. Operators ensure proper execution of all satellite commanding, monitor the performance of the satellite and ground segment, and respond to any real time request or anomaly. Operators can also remotely monitor the status of WCDAS elements and CBU functions, and configure those resources as required. Spacecraft engineering ensures spacecraft health and safety and maintain a continuous flow of - term trending of high quality mission data. This su pport includes performance analysis and long all spacecraft subsystems, INR analysis and operations, anomaly investigation and resolution, maneuver planning and execution, and engineering procedure and database development and Satellite and operations pro maintenance. cedure development may utilize OTS systems for development and configuration management. Instrument engineering monitors instrument performance and detect, diagnose and resolve instrument anomalies. Instrument performance anal ysis evaluates significant instrument performance parameters, analyzing short and long term trends, archiving all pertinent data for future use, and performing statistical analysis of data pertaining to instrument radiometric ctivities include assistance in resolving product data anomalies, calibration and performance. A quantitative monitoring of product data at Level 0 and Level 1, and providing information to support data calibration activities to maintain the highest quality products on a continuous basi s. Norma l Operations : Day in the Life GOES - A typical day in the life of satellite operations includes a maneuver, or some R series combination of maneuvers, plus interactions with the on - board file system. Each day’s maneuver MA and an EWSK maneuver, or a NSSK maneuver plan consists of either a stand - alone MA, a n immediately preceded and followed by MA maneuvers. In addition to the planned maneuver, each operations day includes uplink and downlink of files to/from the on board file system. - 18 - 1

198 These interactions include: • Loading and activating a new 7 - Day Absolute Time Sequence (ATS) – this file is updated daily and includes all planned commands for the next seven days (in compliance with requirement of seven days of autonomous operations). The new file can be uploaded at any time of the day and is activated at the start of each day. • Lo ading and dumping the 7 - Day Target Star Table for ABI • In addition, some other files are updated and uploaded on a weekly basis, including: o the backup ephemeris file, which is used if contact is lost with the GPS constellation o ion Parameters (EOPP) file, needed for the UT1 time the Earth Orientation Predict updates used in some GN&C algorithms. Instrument operations are integrated with spacecraft operations in the 7 - Day ATS. There are a ng on a daily, weekly, variety of typically planned instrument operations that require commandi and less frequent basis. Figure 18 - 1 presents an example of a typical operations day in the life – Figure 18 showing a day with planned MA and EWSK maneuvers, while - 2 shows activities on a typical NSSK day. Operations Activities (EWSK Day) – Day in the Life Figure 1 8 - 1. Typical 18 - 2

199 18 - 2. Figure Typical Day in the Life – Operations Activities (NSSK Day) Normal Operations : Week in the Life Normal Operations of the - R series satellites are typically planned over a week long GOES time frame and implemented via a 7 - d ay ATS that contains all planned commanding for the next 7 days, with a new ATS uploaded and initiated every day. Spacecraft and instrument commanding is combined into a single ATS. A constraint checker within the ystem Mission and s round g - Planning System is used to ensure that there are no conflicts in the uploaded sequences. The 7 day sequence is stored within the Stored Command Processing (SCP) Computer Software Component (CSC). The SCP CSC resides in the OBC FSW that ex ecutes on the RAD750. In addition to the typical da ily operations, the station keeping and momentum a djust maneuvers day repeating cycle. are planned via the 7 - day ATS according to a 4 - Figure 18 - 3 shows a typical schedule for satellite maneuvers over a we ekly period. Typical Seven Day Maneuver Schedule - 3. Figure 18 18 - 3

200 Operations Year in the Life : Normal In addition to the tasks that are repeated on a daily or weekly timeframe, there are also recurring tasks that need to be performed on a less frequent basis ranging from monthly to quarterly to on frequent basis include a monthly an annual basis. Spacecraft bus tasks required on this less enter of g ravity (CG) location and an annual repeat of the IMU to s update of predicted vehicle c tar racker calibration (with maneuvers about each vehicle axis) which keeps the IMU scale factor t within required limits. Flus hing of the HBT thrusters is also required infrequently (~ every 239 days) via a 1 - up of ferric nitrate in the HBT 2 sec thruster burn. This flushing burn limits the build - Charge valves. Battery charging parameters are also modified on an annual basis. The End of Voltage (EOCV) is increased by 0.01V each year to account for fade and increasing depth of discharge (due to less solar array power available to support NSSK power needs). ument calibrations that In addition to the needs of the spacecraft bus, there are a number of instr are required on a less frequent basis. These calibrations can be grouped into those that do not require any spacecraft bus operational changes, and those that do require bus operational - pointing o support (specifically temporary off f the SAWA and SPP from the sun). Instrument calibrations without spacecraft bus operations include: ABI Scan Encoder Calibration (every 1 - 2 months) • • ABI Blackbody and Spacelook Calibrations (as needed) ABI Star Catalog Update (~ twice over mission life • ) • EXIS Calibration Sequence ( q uarterly): EUVS: A/B/C Dark, A/B/C Flat Fields, A/B Gain ilter f Calibrations ( omparisons) c Magnetometer Electronic Calibrations (as needed) • SUVI and EXIS calibrations that require off - pointing of both the SAWA and the SPP from the sun are: uarterly) • EXIS Calibration Sequence ( q – includes XRS, EUVS, and SPS calibrations Fielding/Vignetting, Guide Telescope to - • SUVI Off s un Calibrations ( q uarterly) – Flat Science, Telescope Cross Calibration, Light Transfer Curves re also a number of operations that occur very infrequently (once or twice over mission There a p life) such as tar t racker s tar catalog updates, s ropulsion s ystem repressurization, and EEPROM refreshes. Instrument Operations ABI and SUVI can operate autonomously u sing programmable internal schedules, or, interactively in response to a command sequence. Autonomous operations may be enabled, disabled, or interrupted by command. GLM, SEISS, EXIS and MAG operate autonomously without the need for frequent uploads or cal ibration commands. In addition to their Normal Operating Modes, all of the GOES - R series instruments support “Instrument Diagnostic” and “Health and Safety” modes, including an autonomous “SAFE” mode. s eries instruments are designed to execute transitions between modes in such a - R The GOES manner as to prevent damage to the instrument, and will report the present operating mode for 18 - 4

201 metry is each instrument in the housekeeping telemetry for that instrument. Housekeeping tele transmitted in all powered instrument modes. Flight Software for the instruments is reprogrammable on orbit, and Computer Software Units (object code modules) are usable immediately after upload, without restart of the internal computer, or requ - board memory iring completion of the entire software package upload. All on may be dumped to the ground system on command without disturbing normal operations of instrument data processing. ABI Operations n an The ABI collects Earth scene data swath by swath i direction and builds the image e ast/west n orth/ s outh direction. ABI is able to scan across the s un at its normal from successive swaths in a scan rate two times within 30 seconds or less without interrupting normal imaging operations or sustaining dam age, although performance may be degraded. While in Normal Operating Mode, the ABI concurrently acquires all secondary observations required to meet radiometric and INR requirements within the scan period allocated for primary imaging. When star - sensing is active, the scan pattern is autonomously adjusted to perform the necessary acquisition. Integral parts of each scan mode are space and blackbody calibrations needed to meet cated time for radiometric performance requirements. These calibrations are included in the allo each. It is planned that all instruments will operate concurrently and continuously with minimal downtime for housekeeping operations. ABI exploits the “operate through” capability of the - spacecraft bus for continuous imaging within specific ation during housekeeping activities and orbit maneuvers. No special “keep - out - zone” commanding is required for s un or m oon avoidance in normal operations. ABI is capable of scanning across the Earth limb with the s un present in the FOV at the normal scan rate without damage, but onboard software will prevent direct s un impingement during normal imaging operations with minimal loss of image data. Solar and lunar exclusion - look calibrations are automatically computed by the AB I flight zones for star looks and space software using on board spacecraft ephemeris data. GLM Operations During nominal operations, the GLM requires no commanding from the ground. When the instrument is in NORMAL mode, valid science data is linked to the ground, and processed through he ground processing algorithms. The resulting navigated lighting events are provided to higher t level processing to produce the lightning weather data products. So a typical day includes no commands to the instrument; housekeeping and engineering telemetry are continuously generated; and science telemetry is autonomously generated according to the lightning activity on the earth. For the GLM, much of the raw data processing occurs on the ground (GLM raw data downlink orb it operational requirements are very limited for the GLM. rate is approximately 7.5 Mbps). On - Detector navigation is performed on the ground using spacecraft bus attitude solutions. No routine on - orbit calibrations are required. A large portion of the raw data processing involves the discrim ination of true lightning events from detector stimulation produced by charged particles, 18 - 5

202 - induced events. The flash false alarm probability is less than 5% surface glint, or electronic noise after processing. Some operational characteristics of the GLM ar e: • Continuous operation through eclipse periods • Withstands sun in the field - of - view indefinitely without damage • Autonomous background imaging (intensity of every detector element) once every 5 minutes, or upon ground command GLM data reported for each lightning event will include geolocation of the event to 5 km accuracy, intensity of the detected event, time of the event to an accuracy of 500 microseconds, and the identification of the imager pixel that detected the event. Li ghtning events can be overlaid onto ABI imagery (via ground processing) and the GLM instrument takes background images every 150 seconds. Space Weath er and Solar Imaging Operations The SUVI, EXIS, SEISS, and MAG operate and transmit data during eclipses a nd stationkeeping maneuvers. Each operates independently of the other instruments on the spacecraft bus. All instruments observe simultaneously and do not invoke different observing modes. It is possible that the SEISS and MAG instruments may be operationa - orbit storage to collect space l during on environmental data from the storage location. This function will depend on the storage mode attitude control mode and the downlink antenna geometry. SEISS, SUVI, and EXIS calibrations vary by instrument. The sola r - pointing instruments require s to measure rees un by up to 15 deg periodic (no more than 4 times per year) off - pointing from the background. Sequential orthogonal slews across the solar disk (cruciform slews) are required for but these activities can be combined into a unified operation for the SUVI and EXIS instruments, the SPP suite and are required no more than four times per year. Initial on - orbit calibration of the magnetometer instrument offset bias (instrument plus spacecraft) required successive lar - ge angle (multi - rev) spacecraft rotation maneuvers. The M agnetometer offset determination was a one - time calibration maneuver involving large angle attitude slews performed during the spacecraft post - launch test period in the vicinity of local noon. Like the GLM, the SUVI, EXIS and SEISS instruments require no commanding from the ground but save for the infrequent solar off pointing calibration maneuvers. during nominal operations, - The instruments should require minimal operational resources. Operations Housekeeping Housekeeping operations are activities occurring on a regular basis for maintenance of satellite functions. Examples of routine housekeeping activities include momentum management, clock maintenance, memory dumps or other onboard processor manag ement, or subsystem reconfigurations not covered by onboard autonomy. Any periodic instrument calibration, such as - pointing or MAG calibration sequence commanding, is scheduled as a SUVI/EXIS platform off housekeeping activity. Station keeping management r equirements are met using frequent incremental delta - v maneuvers. Any housekeeping activity not controlled autonomously onboard is planned by the scheduling function. Daily “outage” periods to accomplish housekeeping R. - functions are not specified for GOES Stringent total yearly outage requirements drive all 18 - 6

203 routine housekeeping activities to be accomplished without interruption in instrument data collection or relaxation of performance specifications. Special Operations Special operations are activities not occurring during the course of daily routine operations and are associated with a higher level of risk than routine operations. Typically, special operations activities are supported by engineering staff and managed us ing prescribed operational procedures. Activities with a high degree of complexity and risk, such as non - routine attitude or orbit maneuvers, also require significant preparation. This preparation includes all associated s planning and scheduling, detailed review, contingency planning, and equence of events ( SOE ) fidelity spacecraft simulation. Special operations - SOE validation and crew rehearsals via high may include station keeping or station change maneuvers, spacecraft subsystem or instrument on changes, transition to storage mode, or special instrument calibrations or configurati diagnostics. Anomaly Operations Satellite anomaly operations will occur when the spacecraft bus or instruments experience a t affects normal data collection, or otherwise failure or degradation in function or performance tha compromises the health and safety of the satellite system. Anomalies could be sudden, discrete events, such as the failure of a critical component, or could be a gradual degradation in engineering trending that permits action prior to the occurrence of a performance detected by - mission threatening situation. Onboard failure detection and correction will respond autonomously to spacecraft and instrument anomalies in many cases, but it is the responsibility al time operators to respond to any contingency situation in accordance with pre - defined of re procedures. Mission operations engineering will receive notification of any actual or suspected satellite anomaly through either operator contact or automated ground s ystem functionality. Engineering support will respond in near real time when required. The remote access system will enhance anomaly response by allowing engineers who may be off - site to acquire and analyze satellite telemetry expediently. All anomaly inve stigations and corrective actions will be thoroughly documented in reports and managed under document configuration control to ensure that the knowledge base is maintained throughout the program lifetime. ubsystem anomalies, a key feature of maintaining For serious spacecraft attitude control or other s health and safety is the use of the Safe Hold Mode (SHM), which permits automated acquisition positive and thermally safe condition while - and long term attitude control of the satellite in a power g ground communication. Entry into SHM may be triggered by an event or condition maintainin detected autonomously onboard, or it may be commanded manually based on engineering assessment of a failure or degraded condition. Recovery to normal Earth - pointing or storage mode attitude from SHM would involve significant planning and engineering preparation. Although flight system anomalies present the greatest threat to mission health and safety, ground segment anomalies traditionally comprise the great majority of interr uptions in product data flow. Ground anomalies are analogous to satellite anomalies in that autonomous failure detection and isolation is performed by the Enterprise Management function, but operators are still responsible system engineers will respond similarly for ensuring proper correction of any system fault. Ground to spacecraft engineers in the event of a significant problem. 18 - 7

204 autonomous capabilities Spacecraft The spacecraft bus has autonomous fault detection and correction capability, enabling it to survive the occurrence of any credible single component failure or processor upset. Onboard autonomy t is capable of executing stored drive many aspects of the operational procedures. The spacecraf command sequences and table loads that permit up to seven days of autonomous operation without ground interaction. The spacecraft bus performs uninterrupted image data collection aft bus has sun - during stationkeeping maneuvers. The spacecr positive safehold mode. The spacecraft flight software has telemetry points modifiable on - orbit. The flight software is that are able to be uploaded without disrupting normal processor or spacecraft operations. Image Na vigation and Registra tion (INR) Image Navigation and Registration is a set of image quality metrics pertaining to the location referenced instrument pixels in Level - 1b data. Navigation is absolute pixel location errors of Earth - accuracy, and the various registration requiremen ts specify relative pixel location accuracy. Within frame registration and line - to - line registration are relative pixel - to - pixel location errors within a to single frame. These errors result in image distortion and shear within a single image. Frame - - frame registration is the relative motion of a given pixel in sequential frames. This error produces jumps when successive images are looped. Channel - to - channel registration is the offset between spectral channels for a given pixel location. These errors affect multi - spectral products derived from raw imagery. - end system; the - INR requirements are met through a coordination of all elements of the end to instruments, spacecraft, and ground processing system. INR processing will utilize precision onboard orbit solu tions, star measurements made by the instrument, and spacecraft attitude and angular rate measurements together with ground - based resampling techniques to locate each - grid reference. Responsibility for meeting INR requirements, from photon pixel in a fixed collection through generation of Level 1b data, is placed on the instrument contractors (working to spacecraft - to - instrument interface pointing requirements met by the spacecraft manufacturer). This represents a departure from previous GOES series, in whi ch INR was performed by the prime contractor (GOES I – M) and the spacecraft contractor (GOES N – P ). Image navigation for the SUVI involves all the spacecraft bus pointing considerations of the ABI, to but with the additional complications of solar array platfo rm - body dynamic interactions. Body - - fixed instruments such as the SEISS and MAG are navigated via simple coordinate transformations using the spacecraft bus attitude estimate. Yaw Flip The GOES - R series was designed to not require semi - annual yaw - flip (1 80 degree rotation about the nadir axis) maneuvers, although the capability to perform such maneuvers exist. The yaw flip maneuver may increase seasonal radiometric performance. The instrument designs are not dependent on a semi - annual yaw flip maneuver. I f the need to perform a yaw flip arises, the GOES - R series will not perform imaging during Yaw flip maneuvers and will recover and commence imaging within a prescribed period of time. The cumulative time that imaging is interrupted due to all momentum mana gement, stationkeeping, and yaw flip maneuvers will be under 120 minutes/year. This is compared to 3650 minutes/year for momentum management eries. - s N alone on the GOES 18 - 8

205 Station Relocation The longitude station of a satellite may be changed several times ov er the duration of the mission. Station relocation will occur, for instance, when a satellite is “drifted” from the 89.5 W check - out location to the 105 W storage location or when a satellite is moved from the storage location into tations. A satellite may also be relocated from an operational station at one of the operational s the end of its operational mission for other use before it is decommissioned. Emergency station relocation may be required to replace a failed operational satellite and meet availabi lity east/west delta requirements. Station relocation maneuvers are initiated by applying an v at an - apsis to raise or lower the semi - major axis and induce a “drift” rate in geosynchronous longitude. When the desired station is approached, a roughly equal and opposite delta - v at the same apsis re circularizes the orbit at the new location and stops the drift. The GOES - R series spacecraft will - be capable of up to 2 emergency station relocation maneuvers at a longitude drift rate of 3 deg rees /day, and 3 drift maneuvers at a drift rate of 1 deg ree /day. INR specifications will be maintained at the 1 deg /day drift rate. However, downlink of X - band frequencies during the ree station drift may be restricted, so that imaging may not be possible during station relocat ion and no INR specifications may be applicable. Station relocation events include the following: - checkout location to an on degree/day orbit storage location at a minimum of 1 • From shift From the on - orbit storage location to the operational station location at a minimum of 1 • degree/day shift • Three changes of operational station location while meeting Attitude Control System pointing performance specifications at a minimum of 1 degree/day shift • Two emergency relocations at a minimum of 3 degrees /day shift life longitude at a minimum of 1 From the operational station location to end - of - degree/day shift . Eclipse Operations The GOES - R series spacecraft are designed to support full operations through the maximum geosynchronous eclipse duration of 72 minutes. All instruments are capable of continuous operation through eclipse. Consequently, no special operations should be requi red to accomplish the daily eclipse entry and exits, with the possible exception of commanding to accomplish battery charge management. Seasonal reconfigurations such as for the thermal and electrical power subsystems may be required, but should not signif icantly affect operations. Leap Second Adjustments In the event that the UTC ) is adjusted for a leap second, both GOES Coordinated Universal Time ( spacecraft will be placed into a special housekeeping period at 0000 GMT on the day of the leap second. A cl ock adjustment will be performed to compensate for this change in UTC over the duration of the housekeeping period. Upon exit of the housekeeping period, the onboard clock will be synchronized with UTC within specifications. Orbit - De The deactivation phase occurs when a satellite is declared to be incapable of providing useful mission data or other services and requires disposal to meet international guidelines for the 18 - 9

206 ssion planning and stewardship of geosynchronous resources. This phase includes all mi execution to boost the satellite to a supersynchronous orbit with a perigee no less than 300 km above geosynchronous altitude. This operation is also referred to as “de - orbit”. Following orbit t extent possible and all systems are deactivated so boost, propellant is depleted to the greates up is minimized. All that no spurious RF is radiated and the probability of vehicle break - deactivation activities are accomplished by NOAA o perations, with planning support from the . rogram GOES - R series p 18 - 10

207 Technical Performance Summary 19. nal capabilities of the GOES - This section summarizes the typical and nomi system unless R series otherwise stated. The numbers quoted do not necessarily represent worst case parameter values for all extreme conditions in special modes . Spacecraft Dimensions Height 6129.27 mm (241.31 in) Width 3879.60 mm (152.74 in) Depth 2687.57 mm (105.81 in) Mass GOES - R Subsystem (kg) Spacecraft Structure 607.82 135.85 Thermal Control Mechanisms 108.67 161.45 GN&C C&DH 84.40 TT&C RF 17.36 Propulsion 206.22 Power 335.78 Harness 302.00 Bus Harness 163.82 138.18 Payload Harness Comm 134.69 Antenna 84.42 627.54 Instruments 19 - 1

208 - M^2 kg mm M^2 kg - Mass X - Y Pxz Pyz - cg Z - cg Ixx Iyy Izz Pxy cg ) (kg - 5191.64 - 54.5 23.4 1886.9 8345.7 8557.5 120.2 3987.7 - 94.4 LIFTOFF 26.6 0 0 1786.1 686.5 1627.7 686.5 31.8 0 0 0 Hydrazine 0 0 1358.5 55.4 699.7 468.5 418 0 0 0 Oxidizer 0 0 0 768.4 349.85 0 1358.5 27.7 27.7 2.5 East Ox - 349.85 0 1358.5 27.7 2.5 0 0 0 27.7 West Ox 768.4 6.9 - 0.7 0 263 0.9 0.8 0 0 0 1.4 Pressurant 113.7 310.7 5191.53 7.7 11.5 20.9 1933.4 8311.8 10604.7 6002.2 PRELAE 0 0 0 1627.58 0 0 1851.9 781 781 32.9 Hydrazine 0 0 0 699.7 0 0 1549.7 103.9 516.9 419.1 Oxidizer 768.4 0 1549.7 51.9 51.9 349.85 3 0 0 0 East Ox - 349.85 0 0 1549.7 51.9 51.9 3 0 0 West Ox 768.4 6.9 - 0.7 0 329.4 1.9 1.6 0.8 0 0 0 Pressurant 7834.1 107.1 3765.94 15.9 28.8 1994.5 9736.9 5595.3 310.2 4 STLAE PO 466.1 466.1 0 865.35 0 0 1776.6 0 24.2 0 Hydrazine 0 13.5 36.35 0 0 827.5 35 21.6 0 0 Oxidizer 18.17 768.3 0 0 827.5 6.8 6.8 0.1 0 0 East Ox - 18.17 0 0 0 827.5 6.8 6.8 0.1 0 We st Ox 768.3 0 6.9 - 0.3 0 0 4.5 5.1 1.2 0 1024.9 Pressurant 19 - 2

209 Mass Properties at End Of : M^2 kg kg - - mm M^2 X - cg Mass (kg Y - cg Z - cg Ixx Iyy Izz Pxy Pxz Pyz ) In Orbit 3751.61 - 71.6 322.3 2015.3 8641.4 12694 58.8 - 519.2 320.4 16585.6 Test 12681.2 3692.42 72.8 327.5 2019.1 16541.1 8597.6 - 60.2 - 518.2 315.8 Year 1 3650.28 - 73.6 331.3 2019.8 16517.3 8578.2 12675.5 61.2 - 518 315 Year 2 315 3617.14 - 74.3 334.3 16499.6 8564 12670.9 62 - 518 2019.8 Year 3 3582.98 75 337.5 2020 16480.8 8548.9 - 12666.1 62.8 - 518 314.7 Year 4 3547.97 - 75.7 340.8 2020.2 16461.7 8533.6 12661 63.7 - 517.9 314.4 Year 5 310.6 3477.28 - 77.3 347.8 2023.4 16412.6 8 12647.3 65.6 - 517.1 489.2 Year 6 3430.27 78.3 352.5 2024.8 16384.5 8466.6 - 12640.1 66.9 - 516.7 308.9 Year 7 8 3383.54 - 79.4 357.4 2026 16357.4 445 12632.7 68.2 - 516.4 307.4 Year 8 516.3 - 8426.7 69.5 3337.32 - 80.5 362.4 2026.3 16333.4 307.1 12625.1 Year 9 3282.12 81.9 368.5 2028.2 16301.3 8401.5 - 12615.6 71.2 - 515.8 304.8 Year 10 3238.21 - 83 373.4 2030.6 16273.8 8379.7 12608 72.5 - 515.1 301.9 Year 11 294.1 3195.9 - 84.1 378.4 2037.1 16234.3 12600.4 73.8 - 513.4 8345.9 Year 12 - 85.2 383.3 2044.6 16192.2 155.35 8309.3 12593 75.1 - 511.4 285 3 Year 13 3116.67 8274.7 - 86.2 388 2051.6 16152.1 12585.8 76.4 - 509.5 276.5 Year 14 266.4 507.2 - 77.9 3071.52 - 87.5 393.7 2060 16104.5 8233.7 12577.1 Year 15 262.6 3059.9 - 87.8 395.2 2063.1 16087.8 8216 12572.2 78.4 - 506.4 RESIDUAL 0 183.88 0 0 1559.8 106.9 6.3 0 0 106.9 Hydrazine 0 0 0 11.77 0 0 788.7 4.8 11.8 7 Oxidizer 2.4 768.3 0 788.7 2.4 5.88 0 0 0 0 East Ox - 0 0 5.88 0 0 0 788.7 2.4 2.4 West Ox 768.3 0 0 0 6.9 - 0.2 0 1227.2 4.1 4.8 1.2 Pressurant 19 3 -

210 Electrical Power Subsystem Solar Array Single axis sun tracking Spectrolab Ultra triple junction (GaInP2/GaAs/Ge) Cell type 5 panels, 135.7 cm x 392.3 cm each Panels Redundancy 16 for 15 circuits Power Solar Array Output Satellite Load 5177 W 3748 W BOL summer solstice BOL autumnal equinox 5956 W 4605 W 4830 W 3535 W EOL summer solstice EOL autumnal equinox 4530 W 5489 W Batteries 2 lithium ion batteries, 36 cells each Cell configuration 3 parallel cells per bank, 12 series banks Redundancy 23 for 24 cell banks Capacity 170 Wh/cell Depth of Discharge < 60% of measured capacity Eclipse load 4650 W BOL, 4410 W EOL, max 72 min eclipse 70 V ± 0.6 V at regulation point 70V Bus Voltage 70 V +0.6 V/ Used for housek eeping and 2.0 V at source – Auxiliary Communications 28V Bus Voltage 29.3 ± 0.6 V at source – 3.0 V at load Used for Instruments 29.3 +0.6 V/ 19 - 4

211 Propulsion Design Bipropellant Propellant Tank volumes/ capacity – Fuel (60.2 ft3) / 1637 kg (3609 lbm) Hydrazine 1704.7 L 657.0 L (23.2 ft3) / 900 kg (1984 lbm) Oxidizer – Nitrogen Tetroxide 167.1 L (5.9 ft3) Pressurant - Helium Total Propellant Mass Loaded 1626 kg (3584 lbm) Fuel (1) 700 kg (1543 lbm) Oxidizer (2) Helium (2) 7.3 kg (16.0 lbm) Thrusters LAE (1) 445 N (100 lbf) HBT (2) 22 N (5 lbf) REA (8) 22 N (5 lbf) LTR (16) 90 mN (20 mlbf) 225 mN (50 mlbf) Arcjets (4) 19 - 5

212 GOES - R Performance – Guidance Navigation & Control Attitude Knowledge 569.6 μ Static σ per axis rad 3 Slow Dynamic 26 μrad 3σ per axis 11.8 μ rad 3 σ per axis Dynamic Integrated Rate Error 1 Sec 0.9 μrad 3σ X/Y axis; 0.89 μrad 3σ Z axis 1.16 μrad 3σ X/Y axis; 1.23 μrad 3σ Z axis 30 Sec 300 Sec 4.9 μrad 3σ per axis 900 Sec 12.0 μrad 3σ per axis Knowledge Orbit Track Position 10.1 m 3σ - In Cross - Track Position 11.6 m 3σ Radial Position 51.3 m 3σ Velocity 2.4 cm/sec 3σ per axis Pointing Accuracy 184.5 μrad 3σ per axis Pointing Stability, 60 sec 215.3 μrad 3σ per axis +/ Attitude Rate Error 3σ per axis, based upon 15 msec - 37 μrad/s Thermal Control Subsystem Nominal Spacecraft Internal Dissipation Thermal Load BOL ~1400 W Nominal Spacecraft Internal Dissipation Thermal Load EOL ~1526 W Primary heat rejection Primary heat rejection Aluminum h oneycomb panels with embedded heat pipes, covered with OSRs 247 Spacecraft heaters FSW controlled by OBC Heater Control 19 - 6

213 Command (Command Data Acquisition (CDA)) Receive Characteristics Uplink frequency 2034.200 MHz - Station) - 33 dB/K Minimum G/T (On Minimum G/T (Safehold Mode) - 59 dB/K over 95% spherical coverage - 120 dBm to - Referenced to Command 50 dBm ( Dynamic ran ge (4 ksps) ) Receiver input - Dynamic range (64 ksps) 108 dBm to - 50 dBm ( Referenced to Command Receiver input ) Modulation and Data Ra te Command modulation Direct BPSK Command data rate (uncoded) 3.5 ksps or 56 ksps Command data rate (coded) 4 ksps or 64 ksps - 05 Bit error rate (after decoding) ≤ 1E 19 - 7

214 - Raising Telemetry Tracking and Command Command (Orbit (ORTT&C)) Characteristics Receive Uplink frequency 2036.000 MHz Polarization (Orbit - Raising/On - Station using Hemis) RHCP Polarization (Safehold Mode using Hemis) RHCP Polarization (Orbit Raising using Bicone) Linear - Minimum G/T (On - Station using Hemis) - 42 dB/K Min imum G/T (Safehold Mode using Hemis) - 52 dB/K over 75% spherical coverage Raising using Bicone) 48 dB/K - Minimum G/T (Orbit - - 121 dBm to - 50 dBm Dynamic range (1 ksps) Referenced to Command Receiver input ) ( Dynamic range (4 ksps) - 115 dBm to - 50 dBm ( R eferenced to Command Receiver input ) Modulation and Data Rate BPSK on Subcarrier Command modulation Subcarrier modulation 16 kHz subcarrier phase modulated on carrier Command data rate (uncoded) 875 sps or 3.5 ksps Command data rate (coded) 1 ksp s or 4 ksps 05 - Bit error rate (after decoding) ≤ 1E 19 - 8

215 (Command Data Acquisition (CDA) Telemetry Transmit Characteristics Downlink frequency 1693.000 MHz Polarization (On Station using L/S - Band EC Antenna) RHCP - Polarization (On - Station using Hemis) RHCP Polarization (Safehold Mode using Hemis) RHCP Minimum EIRP (4 ksps) 25 dBmi Minimum EIRP (40 ksps) 33 dBmi Modulation and Data Rate Telemetry modulation Direct BPSK 3.4375 ksps or 34.375 ksps Telemetry data rate (uncoded) Telemetry data rate (coded) 4 ksps or 40 ksps - ≤ 1E 07 Bit error rate (after decoding) Raising Telemetry Tracking and Command - Telemetry (Orbit (ORTT&C)) Transmit Characteristics Downlink frequency 2211.040 MHz Polarization (Orbit - Raising/On - Station using Hem is) RHCP Polarization (Safehold Mode using Hemis) RHCP Polarization (Orbit - Raising using Bicone) Linear - Station using Hemis) Minimum EIRP (On 26 dBmi 24 dBmi over 75% spherical Minimum EIRP (Safehold Mode using Hemis) coverage Modulation and Data Rate Telemetry modulation BPSK on Subcarrier Subcarrier modulation 1.024 MHz Subcarrier Phase Modulated on Carrier Telemetry data rate (uncoded) Telemetry data rate (coded) 1 ksps or 4 ksps 05 - Bit error rate (after decoding) ≤ 1E 19 - 9

216 Tracking Tracki ng Characteristics Turnaround tone ranging Method Turnaround frequency ratio 240/221 8000 nsec Maximum spacecraft ranging signal delay Spacecraft ranging signal delay uncertainty ±40 nsec Receive Characteristics Uplink frequency 2036.000 MHz - RHCP Station using Hemis) Polarization (Orbit - Raising/On RHCP Polarization (Safehold Mode using Hemis) Polarization (Orbit - Raising using Bicone) Linear Transmit Characteristics 2211.040 MHz Downlink frequency sing Hemis) RHCP Polarization (Orbit - Raising/On - Station u Polarization (Safehold Mode using Hemis) RHCP Linear Polarization (Orbit - Raising using Bicone) 19 - 10

217 Raw Data Link (RDL) Transmit Characteristics Downlink frequency 8220.000 MHz H - Polarization Linear - V or Linear 33.65 dBi Minimum antenna ga in (includes pointing error) Maximum antenna gain 37.0 dBi Gimbaled to provide Antenna coverage coverage to WCDAS and RBU (to WCDAS/RBU from 137 deg W) 70.3 dBmi Minimum EIRP Minimum EIR P (to WCDAS/RBU from 75 deg W) 69.6 dBmi Modulation an d Data Rate OQPSK Transmit modulation Transmit data rate (uncoded) 105 Mbps Transmit data rate (coded) 120 Mbps Transmit bandwidth ≤ 130 MHz - 12 Bit error rate (after decoding) ≤ 1E 19 - 11

218 Broadcast (GRB) - GOES Re Receive Characteristics Uplink frequency 7216.600 MHz Linear Polarization V and/or Linear - H - Minimum antenna gain (includes pointing error) 33.2 dBi Maximum antenna gain N/A Gimbaled to provide Antenna coverage coverage to WCDAS and RBU Minimum G/T 3.6 dB/K - - 103 .4 dBmi to Dynamic range 92.1 dBmi Transmit Characteristics 1686.600 MHz Downlink frequency RHCP and/or LHCP Polarization Minimum antenna gain (includes pointing error) 14.8 dBi Maximum antenna gain 17.7 dBi Antenna coverage Earth coverage Minimum EIRP 60. 3 dBmi at EOC Modulation and Data Rate 8 - PSK or QPSK Modulation Data rate (uncoded) per polarization 15.5 Mpbs Data rate (coded) per polarization 23.25 Mbps (8 - PSK) Channel bandwidth ≥ 11.6 MHz - 10 Bit error rate (after decoding) ≤ 1E 19 - 12

219 Report (DCPR) Data Collection Platform Receive Characteristics Uplink frequency 401.900 MHz (Dom) / 402.200 MHz (Int'l) Polarization RHCP Minimum antenna gain (includes pointing error) 12.9 dBi Maximum antenna gain N/A Earth coverage Antenna coverage Minimum G/T - 15.5 dB/K at EOC - - 133.5 dBmi to Dynamic range 110.6 dBmi Transmit Characteristics Downlink frequency 1679.900 MHz (Dom) / 1680.200 MHz (Int'l) Polarization Linear Minimum antenna gain (includes pointing error) 14.2 dBi Ma ximum antenna gain 16.7 dBi Antenna coverage Earth coverage Minimum EIRP 51.3 dBmi at EOC Modulation and Data Rate 8 - PSK Modulation Data rate (uncoded) per carrier 300 bps or 1.2 kbps Data rate (coded) per carrier 450 bps or 1.8 kbps Channel bandwidth ≥ 400 kHz 06 - Bit error rate (after decoding) ≤ 1E 19 - 13

220 Data Collection Platform Command (DCPC) Receive Characteristics Uplink frequency 2032.775 MHz (East) / 2032.825 MHz (West) Polarization Linear Minimum antenna gain (includes pointing error) 14.5 dBi Maximum antenna gain N/A Earth coverage Antenna coverage Minimum G/T - 16.7 dB/K at EOC - - 120.2 dBmi to Dynamic range 107.1 dBmi Transmit Characteristics 468.775 MHz (East) / Downlink frequency 468.825 MHz (West) Polarization RHCP Minimum antenna gain (includes pointing error) 13.0 dBi Maximum antenna peak to - - edge gain ratio 1.7 dB Antenna coverage Earth coverage Minimum EIRP 47.2 dBmi at EOC Modulation and Data Rate BPSK with CDMA Modulation Data rate (uncoded) 306 .1 bps Data rate (coded and chipped) 22.225 kbps Channel bandwidth ≥ 90 kHz ≤ 1E 05 Bit error rate (after decoding) - 19 - 14

221 Rate Information Transmission/Emergency Management High - formation Network (HRIT/EMWIN) Weather In Receive Characteristics Uplink frequency 2027.100 MHz Linear Polarization Minimum antenna gain (includes pointing error) 14.5 dBi Maximum antenna gain N/A Antenna coverage Earth coverage - 16.7 dB/K at EOC Minimum G/T Dynamic range - 91.5 dBm to - 76.7 dBm Transmit Characterist ics 1694.100 MHz Downlink frequency Linear Polarization Minimum antenna gain (includes pointing error) 14.2 dBi - Maximum antenna peak - to edge gain ratio 1.4 dB Antenna coverage Earth coverage Minimum EIRP 56.8 dBmi at EOC Rate Modulation and Data BPSK Modulation Data rate (uncoded) 400 kbps Data rate (coded) 927 kbps Channel bandwidth ≥ 1.2 MHz - 08 Bit error rate (after decoding) ≤ 1E 19 - 15

222 Search and Rescue (SAR) Receive Characteristics Uplink frequency 406.050 MHz tion Polariza RHCP Minimum antenna gain (includes pointing error) 12.9 dBi Maximum antenna gain N/A Earth coverage Antenna coverage Minimum G/T - 15.5 dB/K at EOC 149.7 dBmi to - Dynamic range - 124.7 dBmi Transmit Characteristics 1544.550 MHz Downlin k frequency RHCP Polarization Minimum antenna gain (includes pointing error) 11.5 dBi Maximum antenna gain 14.0 dBi Antenna coverage Earth coverage Minimum EIRP 44.5 dBmi at EOC Modulation and Data Rate PM/BPSK Modulation Data rate (uncoded) 400 bps Data rate (coded) N/A Channel bandwidth ≥ 80 kHz - 05 Bit error rate (uncoded) ≤ 5E 19 - 16

223 Advanced Baseline Imager - View Defining Element Detector - Field of View - of Focal Plane Module Field - - IR 0.88° (NS) x 1.85° (EW) ear Visible & N Midwave IR 0.91° (NS) x 1.55° (EW) 0.89° (NS) x 1.57° (EW) Longwave IR of - Ellipsoid – 20.5° (NS) x 22.7° (EW) Field Regard - Simultaneously ing 16 channel Imag - Scenes (Refresh Rate in Minutes) Scan Capability Full Disk (15), CONUS (5), Mesoscale (0.5) Mode 3 Full Disk (5) Mode 4 User - Custom defined scenes and refresh rates Pixel Spacing Channel (Wavelength) / Detectors - IR Visible & Near 1 km Band 1 (0.47 um) / Silicon Band 2 (0.64 um) / Silicon 0.5 km 1 km Band 3 (0.86 um) / Silicon 2 km Band 4 (1.38 um) / HgCdTe HgCdTe 1 km Band 5 (0.86 um) / Band 6 (1.38 um) / HgCdTe 2 km Midwave IR Band 7 (3.90 um) / HgCdTe 2 km Band 8 (6.185 um) / HgCdTe 2 km B and 9 (6.95 u m) / HgCdTe 2 km Band 10 (7.34 um) / HgCdTe 2 km Band 11 (8.50 um) / HgCdTe 2 km Longwave IR Band 12 (9.61 um) / HgCdTe 2 km Band 13 (10.35 um) / HgCdTe 2 km Band 14 (11.20 um) / HgCdTe 2 km B and 15 (12.30 um) / HgCdTe 2 km Band 16 (13.30 um) / HgCdTe 2 km Space, internal blackbody, solar diffuser Radiometric Calibration Frequency of Calibration Space Look ≤ 30 seconds Infrared Blackbody 15 minutes (Mode 3) / 5 minutes (Mode 4) Solar Diffuser On Demand System Absolute Accuracy Bands 1 – 3, 5 – 6 ± 3% @ 100 % Albedo Band 4 ± 4% @ 100% Albedo ± 1K @ 300 K Bands 7 – 16 19 - 17

224 System Relative Accuracy Pixel - Pixel

225 GLM Design Lens focal length 134 mm Lens f number 1.2 - Lens field of view +/ 8 deg 1372 x 1300 pixels CCD imaging area size 30 x 30 μm Pixel size (variable, up to) Well depth (variable) 2e6 electrons Ground sample distance 8 – 14 km 500 fps Frame rate 777.4 nm Filter center wavelength 1 nm Filter band pass ADC resolution 14 bits Event rate ≥1e5 sec - 1 7.7 Mbps Downlink rate 125 kg Mass (Total) Mass (Sensor Unit) 67 kg 41 kg Mass (Electronics Unit) Operational power 290 W Flash detection efficiency >80% ≥10 years Operating life 19 - 19

226 EXIS Design________________________________________ Parameter Design XRS  Range 0.05 – 0.40 nm 0.80 nm – 0.10 - 10 9 W/m2 XRS Dynamic Range – 10 - 3 W/m2 XRS SNR >30:1 over 10 min. average XRS Data Product Accuracy 14% over mission life 3 sec XRS Cadence EUVS  Range 5 - 127 nm (data product) EUVS  Resolution 115 nm; 5 nm bins 5 – - 127; 10nm bin 117 >20:1 over 10 min. average EUVS SNR EUVS Data Product Accuracy 18% o ver mission life 27 sec EUVS Cadence 19 - 20

227 SUVI Design Mirrors Multi - layer - coated Zerodur Number of coating segments per mirror 6 20 cm Primary diameter Effective focal length 173.04 cm 45 × 45 arcmin or better Field of view Pixel size/Resolution 21 μm/2.5 arcsec CCD detector 1280 × 1280 pixels 450 000 electrons Detector full well 1 per 10 seconds Full image frame rate Typical exposure times 0.01 to 1 second Mass: Telescope subsystem 39 kg Electronics box 25 kg 8 kg Intra - instrument harness Power 225 W (peak) Instrument Science telemetry Interface to spacecraft 3.5 Mbps - orbit storage) Design life 10 years (after 5 years of on 19 - 21

228 Magnetometer Performance Summary Dynamic Range +/ - 512 nT Resolution 0.0016 nT Accuracy <1.7 nT Noise <0.1 nT RMS Bandwidth 2.5 Hz 0.1 deg post calibration Sensor axes orthogonality Within +/ - 19 - 22

229 - Situ Suite (SEISS) Performance Summary Space Environment In - High Energy (MPS - Magnetospheric Particle Sensor HI) Function Measure flux of protons from 80 keV – 12 MeV and electrons from 12MeV 50 keV – 10 solid state silicon detector telescopes: 5 electron and 5 proton Sensor assembly telescopes 2 dosimeters Energy bands Protons 11 energy bands from 80 keV – 12000 keV 400 keV Electrons 12 energy bands from 50 keV – 2 integral channels: >2000 keV and >4000 keV Once every 1 second Sampling rate Field of view 30° cone per telescope, total 170° YZ plane Z direction - 180° per dosimeter, Magnetospheric Particle Sensor – Low Energy (MPS - LO) – 30 keV Function Measure flux of ions and electrons from 30 eV 2 sensor heads, each with 1 electron and 1 ion microchannel plate Sensor assembly Energy bands Ions 15 logarithmically spaced 15 logarithmically spaced Electrons Once every 1 second Sampling rate 180° YZ plane Field of view 19 - 23

230 Solar and Galactic Proton Sensor (SGPS) Function Measure flux of protons from 1MeV – 500 MeV and alpha particles from 4 MeV – 500 MeV 2 units, each with 3 solid state silicon detector telescopes Sensor assembly Energy bands Protons 14 front entry particles, 4 rear entry particles only Alpha particles 12 front entry particles, 3 rear entry particles only Sampling rate Once every 1 second Field of view 1 unit pointed +X, 1 unit pointed – X 3 telescopes / unit: 2 at 60° cone, 1 at 90° cone XY plane Energetic Heavy Ion Sensor (EHIS) - 200 MeV / nucleon Function Measure proton and heavy ion flux from 1 5 nickel in 30 mass Measure individual elements from hydrogen to bands Sensor assembly 1 solid state silicon detector telescope Energy bands 5 logarithmically spaced energy bands at each of 30 mass bands Sampling rate Full data set once per 60 seconds 56° cone YZ plane Field of view 19 - 24

231 Acronyms 20. PPS One Pulse per Second - 1 - Day Store 2DS 2 A Angstrom ABI Advanced Baseline Imager ACS Antenna Controller System Antenna Control Unit ACU AD Attitude Determination ADC Analog to Digital Converter ADC Digital Conversion Analog to ADIS Angle Detecting Inclined Sensor Ancillary Data Relay System ADRS Atmospheric Imaging Assembly AIA AWIPS Network Control Facility ANCF AS Application Server ASD Acceleration Spectral Density ASIC Application Specific Integrated Circuit ASIS Antenna System Interface Simulators ATS Absolute Time Sequence AVD Active Vibration Damping AWA Antenna Wing Assembly Advanced Weather Interactive Processing System AWIPS Battery Charger/Discharger BCD Best Detector Select BDS Built - In - Test/ Built - In - Test - Equipment BIT/BITE AWIPS Backup Network Control Facility BNCF Binary Phase - Shift Key BPSK Body Reference Frame BRF BS Beamsplitter C&DH Command and Data Handling Channel Access Data Units CADU CASSIE Contextual Analysis for Spectral and Spatial Information CBU Consolidated Backup CCA Circuit Card Assembly CCD Charge Coupled Device CCE Cryocooler Control Electronics CCTV Closed Circuit Television CCSDS Consultative Committee for Space Data Systems Central Distribution Assembly CDA CDA Command and Data Acquisition 20 - 1

232 Contract Data Requirements List CDRL Command Decryption Unit Assembly CDUA CEB Camera Electronics Box Center of G ravity CG CI Configuration Item - data Stewardship System CLASS Comprehensive Large Array CM Configuration Management CMDB Configuration Management Database CME Coronal Mass E jection CMI Cloud and Moisture Imagery CODT Custom Object Dump Tool Collision A voidance COLA CONUS Contiguous United States COOP Continuity of Operations COSI Common Operating System Image CPE Certified Principle Engineer CS Consolidated Storage CSC Computer Software Component CSC Computer Software Configuration CSSA Sensor Assembly Coarse Sun CSU Current Sensor Unit Coeficient of Thermal Expansion CTE CTP Command and Telemetry Processor DB Database Direct Current DC Data Collection Platform DCP DCPC Data Collection Platform Command DCPR Data Collection Platform Report Data Collection System DCS DE Development Environments DF Data Formater DGS Diego Garcia Station DN Digital Number DO Data Operations Data Processor DP DPU Data Processing Unit DRGS Direct Readout Ground Station DVB S2 Digital Video Broadcasting Satellite Second Generation - ECEF Earth - Centered Earth - Fixed EDAC Error Detection and Correction - Electronically Erasable Programmable Read Only Memory EEPROM 20 - 2

233 Electrical Ground Support Equipment EGSE Energetic Heavy Ion Sensor EHIS En terprise Infrastructure EI Emergency Locator Transmitter ELT EM Enterprise Management EMWIN Emergency Managers Weather Information Network EOCV End of Charge Voltage Earth Orientation Prediction Parameters EOPP EP Encoder Processor EPEAT Electron, Proton, Alpha Detector EPC Electronic Power Converter EPIRB Emergency Position Indicating Radio Beacons Earth Pointing Platform EPP Electrical Power Subsystem EPS ESB Enterprise Service Bus ESD Electro - static discharge ESPC Environmental Satellite Processing Center ESPDS Environmental Satellite Processing and Distribution System ETA EUV Telescope Assembly EU Electronics Unit Extreme Ultraviolet EUV EUVS Extreme Ultraviolet Sensor - W est EW East EWSK - West Station K eeping East EXEB EXIS Electrical Box EXIS Extreme Ultraviolet and X - ray Irradiance Sensors FAA Front Aperture Assembly FBA Fuse Board Assembly FEP Front End processors FGF Fixed Grid Frame FIFO First In, First Out FM Flight Model Field of FOV View FPA Focal Plane Arrays FPAA Focal Plane Array Assembly FPGA Field Programmable Gate Array FPM Focal Plane Module FPP Focal Plane Package FSDE Flight Software Development Environment FSME Flight Software Maintenance Environment Fine Sun FSS Sensor 20 - 3

234 Fine Sun Sensor Assembly FSSA FSW Flight Software FUV Far Ultraviolet Full - Width Half - FWHM Maximum GAS GOES - R Access Subsystem t GEO Geosynchronous Earth Orbi GEOSAR Geostationary Search and Rescue GHe Gaseous Helium GLM Geostationary Lightning Mapper GN&C Guidance Navigation and Control GOES Geostationary Operational Environmental Satellite Ground Processing Algorithm GPA GPS Global Positioning System GPSR Global Positioning System Receiver GRATDAT GOES - R ABI Trending and Data Analysis Toolkit GRB GOES Rebroadcast service GRBT GRB User Ter minals GRDDP GOES - R Reliable Data Delivery Protocol round G ystem S GS GSD Ground Sample Distance Guide Telescope Assembly GTA GTO Geosynchronous Transfer Orbit GUI Graphical User Interface GOES VARiable GVAR Hydrazine Bi - Propellant Thruster HBT High Output Paraffin Actuator HOPA HR Hurricane - Rated Hemispheric Resonating G yros HRG HRIT High Rate Information Transmission HSIO High Speed I/O HWIL Hardware In the Loop Hz Hertz Integration and Test I&T IC Instrument Controller ICRF International Celestial Reference Frame ICT Internal Calibration Target IEEE Institute of Electrical and Electronics Engineers IETF Internet Engineering Task Force IFDS Intermediate Frequency Distribution System Facility Link IFL Inter - IFL Intermediate Frequency 20 - 4

235 Inertial Measurement Unit IMU Information INFO Image Navigation and Registration INR Integrated Rate Error IRE IR Infrared IS Infrastructure Isp Specific Impulse Integration and Test Environment ITE JPSS Joint Polar Satellite System keV Kilo Electron Volt km Kilometer KPP Key Performance Parameter Kilo - Symbols per Second ksps Keyboard, Video and Mouse KVM Liquid Apogee Engine LAE LAN Local Area Network LASP Laboratory for Atmospheric and Space Physics LCFA Lightning Cluster - Filter Algorithm LCM Low voltage Control Module LED Light Emitting Diode Lower Equipment Room LER LHCP Left Hand Circular Polarization Loop Heat Pipe LHP Lightning Imaging Sensor LIS Lightning Imaging Sensor / Optical Transient Detector LIS/OTD LOS Motion Compensation LMC Lockheed Martin Space Systems Company LMSSC L ow Noise A mplifiers LNA Launch and Orbit Raising LOR - L ine - o f ight S LOS LPM Low voltage Power Module LTR Low Thrust REA Launch Vehicle LV Low - Voltage Differential Signal LVDS LWIR Longwave infrared LZSS Level Zero Storage Service M&C Monitor & Control MA Momentum Adjust MAG Magnetometer Electron Detector MAGED Magnetosphere MAGPD Magnetospheric Proton Detector 20 - 5

236 Megabits per second Mbps Microchannel Plates MCP MECO Main Engine Cutoff Macintyre Ele ctronic Design Associates, Inc MEDA MES Main Engine Start MeV Mega Electron Volt MegaHertz MHz μrad Mic roradian MLI M ulti - L ayer I nsulation MLS Mission Life Store MM Mission Management Mo/Si Molybdenum/Silicon MOST Mission Operations Support Team Mo/Y Molybdenum - Yttrium MPLS Multi - Protocol Label Switching MPS Mission Planning and Scheduling MPS - HI Magnetospheric Particle Sensor - High energy range MPS - LO Magnetospheric Particle Sensor - Low energy range - Momentum and Station Keeping Simulation MSKSim MUV Middle Ultraviolet Multiplexer MUX Midwave – Longwave MW/LW Midwave infrared MWIR MY Minus Y National Aeronautics and Space Administration NASA Narrow Band Filter NBF NCEI National Centers for Environmental Information NOAA Center for Weather and Climate Prediction NCWCP Noise Equivalent delta Temperature NEdT NESDIS National Environmental Satellite, Data and Information Service NF Network Fabric NOAA National Oceanic and Atmospheric Administration Network Operations Center NOC National Polar - orbiting Partnership NPP NS N orth - S outh NSOF NOAA Satellite Operations Facility NSSK North South Station Keeping nT nanoTesla N - WAVE NOAA Science Network National Weather Service NWS Orbit and Attitude O&A 20 - 6

237 Operations & Maintenance O&M O&A Angular Rate OAR On Board Computer OBC Operations Configuration Change Request OCCR OD Orbit Determination OE Operational Environment OMC Orbit Motion Compensation Orthomode Transducer OMT OPC Optical Port Cover OPSA Optical Port Sunshield Assembly ORTT&C - Raising Tracking, Telemetry and Control Orbit Operational Support Location OSL Office of Satellite and Product Operations OSPO Optical Solar Reflector OSR OTS Off - the - Shelf OTD Optical Transient Detector P&TC Peripheral and Thermal Control PD Product Distribution Product Distribution and Access PDA PDB Parameter Database PDM Power Distribution Module Power Drive Unit PDU PG Product Generation PIFT Predicted Interface Force and Torque PKI Public Key Infrastructure PLEIADES Post Launch Enhanced Image and Data Evaluation System PLT Post - Launch Test PMD Propellant Management Device PMU Personal Maintenance Unit PPS Pulse per second PPZ Product Processing Zone PRA Pyro Relay Assembly Platinum Resistan t Thermometer PRT Power Regulation Unit PRU PS Power Supplies PSU Personal Safety Unit PTR Program Tracking Reports PY Plus Y QE Quantum Efficiency QJ Quad Junction Restraint and Release R&R 20 - 7

238 Random A ccess M emory RAM Relay Drive Card RDC RDL Raw Data Link Rocket Engine Assembly REA RF Radio Frequency RFI Radio Frequency Interface Right Hand Circular Polarization RHCP RIU Remote Interface Units RMC Redundancy Management Card RMS Root Mean Square ROIC - Out Integrated Circuit Read RTEP Real Time Event Processors Reaction Wheel Assembly RWA Solar Array Drive Assembly SADA SADE Solar Array Drive Electronics Box SAR Search and Rescue SARSAT Search and Rescue Satellite - Aided Tracking SAS Solar Array Shunt SAST Spacecraft All - Software Testbed Solar Array Wing Assembly SAWA SBF Solar Blocking Filter Solar Calibration Assembly SCA SCC Solar Calibration Cover SCN Spacecraft Navigation Stored Command Processing SCP Solar Calibration Target SCT Drive Assembly Scan SDA SDO Solar Dynamics Observatory Sensor Electronics Box SEB SUVI Electronics Box SEB SEGA SPP Elevation Gimbal Assembly SEISS - Situ Suite Space Environment In Separation Nuts Sep Nuts SERDES SERializer - DESerializer sFTP secure File Transfer Protocol SGC Space Ground Communications SGPS Solar and Galactic Proton Sensors SHM Safe Hold Mode SIMD Scanner Interface & Motor Driver ointing Platform Interface Unit, Sensor Interface Unit SIU Sun P SMA Shape Memory Alloy 20 - 8

239 SMC Spacecraft Motion Compensation Single Mode Fiber SMF SMS - 1 first Synchronous Meteorological Satellite SNR Signal to Noise Ratio SNMP Simple Network Management Protocol SOCC Satellite Operations Control Center SOZ Satellite Operations Zone SPP Sun Pointing Platform SPS Solar Position Sensor SPS Sun Pointing Subsystem SRA Slip Ring Assembly Static Random Access Memory SRAM Solar Rejection Filter SRF SRS Shock Response Spectra SSD Solid State Detector SSIRU Scalable Space Inertial Reference Units SSPA Solid State Power Amplifier SSRD Split Spool Release Device STAR Satellite Applications and Research STS SUVI Telescope Subsystem SU Sensor Unit SUE Sensor Unit Electronics SUVI Solar Ultraviolet Imager Solar Wing Assembly SWA SWPC Space Weather Prediction Center SWRC SpaceWire Router Card TBA Trailer Bearing Assembly TCP/IP Transmission Control Protocol/Internet Protocol TDU Thermal Dynamic Unit TFRS Timing and Frequency Reference System TNCF Test Network Control Facility TNR Threshold - to - Noise R atio TNT Telemetry & Timing TRMM Tropical Rainfall Measuring Mission Transient Suppression Unit TSU Tracking, Telemetry, and Control TT&C TWTA Traveling Wave Tube A ssembly ULA United Launch Alliance USN Universal Space Network UTC Coordinated Universal Time Ultra Triple Junction UTJ 20 - 9

240 VEM Visco - Elastic Material VIS/IR Visible – Infrared Visible and Near Infrared VNIR Video Processors VP VPN Virtual Private N etwork WAN Wide Area Network Wallops Command and Data Acquisition Center WCDAS Weather Forecast Office WFO XRS X - Ray Sensor LZSS 0 Storage Solution Level - 20 - 10

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