SCIENCE & MISSION REQUIREMENTS DOCUMENT · SCIENCE & MISSION REQUIREMENTS DOCUMENT REVISION:...

GRAVITY RECOVERY ANDCLIMATE EXPERIMENT (GRACE)

SCIENCE & MISSIONREQUIREMENTS DOCUMENT

327 – 200RELEASE 15 MAY 1998

JET PROPULSION LABORATORY

UNIVERSITY OF TEXAS

CENTER FOR SPACE RESEARCH

GEOFORSCHUNGSZENTRUM POTSDAM

SCIENCE & MISSION REQUIREMENTS DOCUMENT REVISION: RELEASEGRACE 327 – 200 MAY 15, 1998 PAGE 2 OF 48

Prepared by:

_______________________________Richard Stanton, JPLGRACE System Engineering Manager

_______________________________Srinivas Bettadpur, UTCSRGRACE Science Team MSET Representative

_______________________________Charley Dunn, JPLGRACE Instrument Manager

_______________________________Klaus-Peter Renner, GSOCGRACE Operations Manager

_______________________________Michael Watkins, JPLGRACE Science Data System Manager

Approved by:

______________________________ _______________________________Ab Davis, JPL Byron Tapley, UTCSRGRACE Project Manger GRACE Principal Investigator

_______________________________ _______________________________Richard Fitzgerald, GSFC Christoph Reigber, GFZ PotsdamGRACE Mission Manager GRACE Co-Principal Investigator.


TABLE OF CONTENTS

1.0 INTRODUCTION

1.1 PURPOSE - - - - - - - - 6

1.2 REQUIREMENTS DEVELOPMENT OVERVIEW - - 6

1.3 DESIGN TO COST - - - - - - - 6

1.4 CHANGE CONTROL - - - - - - 71.4.1 Changes to the Baseline Science Mission - - - - 71.4.2 Changes to the Minimum Science Mission - - - - 71.4.3 Changes to the System Requirements - - - - 71.4.4 Changes to Subsystems and Assemblies - - - - 7

1.5 RELEVANT DOCUMENTS - - - - - 7

2.0 SCIENCE REQUIREMENTS

2.1 OVERVIEW - - - - - - - - 9

2.2 SCIENCE OBJECTIVES - - - - - - 102.2.1 Primary science objectives - - - - - - 102.2.2 Typical science applications - - - - - - 11

2.3 BASELINE SCIENCE MISSION - - - - - 132.3.1 Gravity potential model - - - - - - 132.3.2 Long term gravity model - - - - - - 132.3.3 Temporal variability - - - - - - - 132.3.4 Geoid height error - - - - - - - 132.3.5 Orbit characteristics - - - - - - - 132.3.6 Atmospheric soundings - - - - - - 13

2.4 SCIENCE DATA DELIVERABLES - - - - 132.4.1 Level 0 data - - - - - - - - 142.4.2 Level 1 data - - - - - - - - 142.4.3 Level 2 data - - - - - - - - 15

2.5 MINIMUM SCIENCE MISSION - - - - - 152.5.1 Earth gravity field model - - - - - - 152.5.2 Mean model - - - - - - - 152.5.3 Geoid height error - - - - - - - 162.5.4 Orbit characteristics - - - - - - - 16

2.6 MINIMUM SCIENCE DATA DELIVERABLE - - - 16


3.0 MISSION OVERVIEW

3.1 MISSION SYSTEMS - - - - - - 173.1.1 Launch Vehicle System (LVS) - - - - - 183.1.2 Satellite System (SAT) - - - - - - 183.1.3 Science Instrument System (SIS) - - - - - 183.1.4 Mission Operations System (MOS) - - - - - 183.1.5 Science Data System (SDS) - - - - - - 18

3.2 MISSION PHASES - - - - - - - 203.2.1 Pre-launch phases - - - - - - - 203.2.2 Launch & Early Orbit Phase (LEOP) - - - - - 203.2.3 Assessment Phase (AP) - - - - - - 213.2.4 Observational Phase (OP) - - - - - - 21

3.3 MISSION REQUIREMENTS - - - - - 213.3.1 Injection altitude - - - - - - - 213.3.2 Satellite separation - - - - - - - 213.3.3 Orbit inclination - - - - - - - 213.3.4 Satellite interchangeability - - - - - - 213.3.5 Minimum data capacity - - - - - - 223.3.6 Data gaps - - - - - - - - 223.3.7 Non-science operations - - - - - - 223.3.8 Onboard fuel (GN2) utilization - - - - - 223.3.9 Single point failures - - - - - - - 22

3.4 RADIATION REQUIREMENTS - - - - - 223.4.1 South Atlantic Anomaly (SAA) - - - - - 223.4.2 Single-event upsets & latch-up - - - - - 223.4.3 Radiation Design Margin - - - - - - 22

3.5 DERIVATION OF SYSTEM REQUIREMENTS- - - 23


4.0 SYSTEM REQUIREMENTS

4.1 LAUNCH VEHICLE SYSTEM4.1.1 Launch period - - - - - - - - 274.1.2 Dual launch - - - - - - - - 274.1.3 Orbit altitude - - - - - - - - 274.1.4 Payload mass - - - - - - - - 274.1.5 Eccentricity - - - - - - - - 274.1.6 Inclination - - - - - - - - 274.1.7 Ascending node - - - - - - - 284.1.8 Launch window - - - - - - - 284.1.9 Launch vehicle interference - - - - - - 284.1.10 Tip-off rates - - - - - - - - 284.1.11 Separation velocity - - - - - - - 28

4.2 SATELLITE SYSTEM4.2.1 General Requirements- - - - - - - 294.2.2 Operating Modes - - - - - - - 294.2.3 Coordinate Systems - - - - - - - 324.2.3 Attitude and Orbit control - - - - - - 344.2.4 Critical Alignments & Stabilities - - - - - 344.2.5 Calibrations & Adjustments - - - - - - 364.2.5 Command and Data Handling - - - - - 36

4.3 SCIENCE INSTRUMENT SYSTEM4.3.1 General requirements - - - - - - - 374.3.2 SuperSTAR Accelerometer (ACC) - - - - - 374.3.3 Star Camera Assembly (SCA) - - - - - 394.3.4 Instrument Processor Unit (IPU) - - - - - 394.3.5 K-Band Ranging Assembly (KBR) - - - - - 404.3.6 Ultra-Stable Oscillator (USO) - - - - - 404.3.7 Laser Retro Reflector (LRR) - - - - - - 41

4.4 MISSION OPERATIONS SYSTEM4.4.1 MOS Functions - - - - - - - 424.4.2 Real-time Operations - - - - - - - 424.4.3 Navigation Requirements - - - - - - 43

4.5 SCIENCE DATA SYSTEM4.5.1 General Requirements- - - - - - - 454.5.2 Data Processing and Distribution - - - - - 454.5.3 Verification - - - - - - - - 464.5.4 Precision Orbit Determination - - - - - 46

GLOSSARY - - - - - - - - - - 47


CHAPTER 1

INTRODUCTION

1.1 PURPOSE

This document establishes the science and mission requirements for the Gravity Recovery andClimate Experiment (GRACE) mission. It also provides a mapping of these requirements to thefive Project systems that constitute the top level of technical implementation responsibility.

1.2 REQUIREMENTS DEVELOPMENT OVERVIEW

GRACE requirements arise from a number of distinct sources. The first source is the GRACEMission Definition and Requirements Agreement (NAS5-97213) summarizing top-level sciencerequirements. These requirements, for both the Baseline Science Mission and the MinimumScience Mission are expanded and elucidated in the Chapter 2 of this document. A description ofthe GRACE Mission, including overviews of the systems, operational phases, and environmentalrequirements are presented in Chapter 3. Chapter 3 also provides a breakdown of the GRACEmeasurement error budget used to derive many of the system-level requirements presented inChapter 4. Apart from the environmental requirements, top-level reliability and quality assurancerequirements are documented in the GRACE Mission Assurance Plan (327-120), and are notrepeated here.

The major emphasis Chapter 4 is the detailed specification of the requirements on each of the fiveMission Systems. These requirements establish the basis for subsystem and assemblyimplementations and will be reflected and implemented in lower level documents. These include,but are not limited to the documents listed in Section 1.5, Relevant Documents. In general, theGRACE Project will flow down requirements to lower levels, such as electronic assemblies,through direct inclusion individual specifications, not by reference to “applicable documents”.

1.3 DESIGN TO COST

The requirements listed herein on the Mission Systems are considered necessary and sufficient tomeet the requirements of the Baseline Science Mission defined in Chapter 2. Since GRACE is acost-capped, “design to cost” mission, performance is subject to limits imposed by stringentresource constraints. In the event that requirements cannot be met within the cost cap, somerequirements may be renegotiated under the change control process described below. AMinimum Science Mission has been defined (Section 2.5), characterizing the minimum viableoutput for the GRACE mission. In the event of serious problems, the Principal Investigator hasthe responsibility for determining if the descoped Project can provide this minimum science.


1.4 CHANGE CONTROL

The signed SMRD will be placed under the change control authority of the GRACE Mission andSystem Engineering Team (MSET). Requests for additions, changes, and deletions to the SMRDshould be submitted directly to the GRACE Systems Engineering Manager (SEM), who willcoordinate the appropriate review process. The review process, and the required levels ofapproval, are functions of the category of change requested.

1.4.1 Changes to the Baseline Science Mission

Proposed changes are evaluated by the MSET team in terms of expected benefits, effect on othersystems, and cost, schedule, and risk impacts. Based on this assessment, the Project Managershall recommend either approval or disapproval by the Principal Investigators, who make the finaldecision.

1.4.2 Changes to the Minimum Science Mission

Requirements for the Minimum Science Mission, representing the minimum expected return forthe program investment, can only be changed with approval of the NASA Earth System SciencePathfinder (ESSP) Program Office.

1.4.3 Changes to the System Requirements

Proposed changes are evaluated by the MSET team in terms of expected benefits, effect on othersystems, and cost, schedule, and risk impacts. Based on the Team recommendations, the SEMshall approve or disapprove the recommended change, subject to review by the Project Managerin the event of a cost and/or risk impact. Changes that affect science shall only be accepted withthe approval of both the Principal Investigators and the Project Manager.

1.4.4 Changes to Subsystems and Assemblies

Review and approval by the MSET is not required for lower-level changes that do not affecteither the Science or System-level requirements

1.5 RELEVANT DOCUMENTS

The SMRD is consistent with, and responsive to, the following documents:

NAS5 – 97213 GRACE Mission Definition & Requirements Agreement (MDRA)327-100 GRACE Cooperative Project Plan327-120 GRACE Mission Assurance Plan (MAP)GR-GFZ-SP-0001 Launch Vehicle Requirement SpecificationNMI 1700.8 Policy to Limit Orbital Debris GenerationNSS 1740.14 NASA Safety Standard Guidelines and Assessment Procedures for

Limiting Orbital Debris


The following documents are consistent with, and responsive to, the SMRD (partial listing):

327-300 Launcher Interface Document327-410 Satellite System Specification327-420 Flight Software Requirements Document327-430 Satellite Assembly, Integration, Verification and Test Plan327-510 K-Band Ranging Specification327-520 Accelerometer Specification327-530 Star Camera Specification327-540 Instrument Processor Unit Specification327-550 Ultra Stable Oscillator Specification327-610 Ground System Requirements Document327-710 Science Data System Requirements


CHAPTER 2

SCIENCE REQUIREMENTS

2.1 OVERVIEW

The primary objective of the GRACE mission is to obtain accurate global models for the meanand the time variable components of the Earth's gravity field. This objective will be achieved bymaking measurements of the inter-satellite range and its derivatives between two co-planar, lowaltitude near-polar orbiting satellites, using a micro-wave tracking system. In addition, eachsatellite will carry geodetic quality Global Positioning System (GPS) receivers and high accuracyaccelerometers to enable accurate orbit determination, spatial registration of gravity data and theestimation of gravity field models.

GRACE is a joint project between the National Aeronautics and Space Administration (NASA)and the Deutsches Zentrum für Luft und Raumfahrt (DLR). Overall responsibility forimplementation of GRACE rests with Dr. Byron Tapley, Principal Investigator, of the Universityof Texas at Austin, Center for Space Research (UTCSR), and Prof. Christoph Reigber of theGeoForschungsZentrum (GFZ) Potsdam, Co-Principal Investigator.

The Earth gravity field estimates obtained from data gathered by the GRACE mission willprovide, with unprecedented accuracy, integral constraints on the global mass distribution and itstemporal variations within the Earth System. The need for a mission to obtain such data withglobal coverage, high resolution and homogeneously high quality, for several applications inEarth System Sciences has been consistently articulated since the Williamstown Conference in1969.

Present models for the mean gravity field of the Earth have been developed using a combinationof various space and ground based data gathered over the past several decades. These datainclude terrestrial and space-based satellite tracking, conventional gravimetry over land, andsatellite altimetry over the oceans. The accuracy of the mean Earth gravity models is currentlylimited by the heterogeneity of the data quality and origin, and lack of global coverage withuniform accuracy. Present techniques capable of providing estimates of global temporalvariability of the Earth’s gravity field are limited to the longest wavelengths (first few sphericalharmonics of the geopotential) and certain selected frequency regimes. Due to its global coverageand high accuracy, a quantum leap in the accuracy of the models for the Earth’s gravity field willbe feasible with GRACE.

The path or the orbit of any satellite in Earth orbit is dependent on the integrated effect of themass distributions and movements in the Earth system. Orbits of the two GRACE satellites,sensing these effects at slightly different phases, will be perturbed differentially. This difference


in perturbations is manifested in the inter-satellite range changes. The GRACE microwaveranging instrument will provide very accurate measurements of these range changes. By makingthese differential measurements, the higher frequency content of the gravitational signals will bemagnified, thus enabling significantly improved estimates of the higher resolution features of theEarth’s gravity field. Since the two satellites are separated in orbit by only a few hundredkilometers along track, the errors due to media effects are minimized as compared to space-basedor ground-based tracking. This also ensures the homogeneity of data quality over the missionlifetime of GRACE. By choosing to fly in nearly polar orbits, the coverage of the globe is alsomaximized. The presence of high accuracy accelerometers on board each satellite ensures that thedifferential range changes due to the non-gravitational accelerations can be accurately modeledand removed in the processing of the GRACE inter-satellite tracking data. Each satellite alsocarries a geodetic quality GPS receiver to ensure that the orbits for the satellites can becontinuously and accurately determined and that the gravity field estimates can be correctlyregistered in a terrestrial reference frame. A dedicated Laser Retro Reflector (LRR) on eachsatellite allows an external calibration of the onboard GPS system for orbit determination. Laserranging from the ground can also be used in conjunction with the GPS to support precise orbitdetermination and gravity field recovery.

The improved estimates of mean and time variable components of Earth’s gravity field fromGRACE, in conjunction with other in-situ and satellite data, as well as geophysical models, willprovide impetus for advances in a wide variety of Earth System Science disciplines. In theoceanographic community, the knowledge of the mean geoid, in conjunction with satellitealtimetry data, will allow significant advances in the studies of oceanic heat flux, long term sealevel change, upper oceanic heat content, and the absolute surface geostrophic ocean currents.Further, the estimates of time variations in the geoid obtained from GRACE will help the sciencecommunity unravel complex time variable processes in oceanography (e.g. deep ocean currentchanges), hydrology (e.g. large scale evapo-transpiration and soil moisture changes), glaciology(e.g. polar and Greenland ice sheet changes), and the solid Earth sciences (e.g. postglacialrebound, solid Earth’s isostatic response)

2.2 SCIENCE OBJECTIVES

2.2.1 Primary science objectives

The primary objective of the GRACE mission is to provide, with unprecedented accuracy,estimates of the global and high-resolution models of the Earth’s gravity field for a period of up tofive years. The temporal sequence of gravity field estimates will yield the mean Earth gravityfield, as well as a time history of its variability. An additional science objective of the GRACEmission is to provide several hundred globally distributed profiles each day of the excess delay, orbending angle of the GPS measurements due to the ionosphere and the atmosphere, using GPSlimb-sounding. These can be converted to total electron content and/or refractivity in theionosphere and troposphere, respectively.


The science data required to realize these objectives are defined in Table 2-1, which establishesthe primary measurement objectives for GRACE.

Table 2-1 Science Measurement Objectives

Science Measurement Instrument Spacecraft

Inter-Satellite RangeChange

K-Band Rangingµ-wave link

2

Earth GravityField Non-Grav.

AccelerationsAccelerometer 2

Satellite Orbits GPS Turbo-Rogue Receiver 2

AtmosphericOccultation

GPS-to-GRACE phasechange

GPS Turbo-Rogue Receiver 1

This science data, in conjunction with ancillary data, will then be used to obtain estimates ofspherical harmonic coefficients of the Earth’s gravitational potential. A typical 30 day span ofdata, collected in a 475 km altitude polar orbit, shall have a global root mean square (rms) geoidheight error due to the measurement system errors as specified in Table 2-2.

Table 2-2 Geoid Height Error over 30 days

Harmonic DegreeGeoid Height Error

Per Degree(mm)

Geoid Height ErrorCumulative (from n=3)

(mm)

n = 2 < 0.10 -

3 ≤ n ≤ 10 < 0.01 < 0.02

10 ≤ n ≤ 70 < 0.15 < 0.40

70 ≤ n ≤ 100 < 1.50 < 3.50

100 ≤ n ≤ 150 < 65.0 < 200

2.2.2 Typical science applications

The estimates of the Earth gravity field from GRACE, in conjunction with other space-basedmeasurements, in-situ data and geophysical models, will be used to discriminate time varyingchanges in the mass of the Earth’s dynamical system due to different geophysical processes.


Examples include the discrimination of effects due to sea level rise, continental water storage, ice-sheet changes and other geophysical phenomena. Additionally, atmospheric model studies willbenefit by the recovery of refractivity (and the derived quantities of temperature and water vapor)from the use of GPS limb sounding. Furthermore, limited GPS sounding of the ionospherebeginning at altitudes in the region of 100 km will be available for studying fine ionosphericstructure.

As an example, Table 2-3 summarizes the advances in science applications achievable from theGRACE mission data obtained in a 475 km altitude polar orbit. For each scientific application,the principal spatial, and where appropriate, the temporal scales of associated geoid variability aregiven, along with the improvement obtainable from such a mission scenario.

Table 2 - 3 Science Applications Summary

APPLICATION RESOLUTION TIME SCALE ACCURACY COMMENTS

STATIC GRAVITY FIELD

Oceanic HeatFlux

> 1000 km > 40 percentimprovements

Ocean Currents > 1000 km < 1 mm geoid error Improves to <0.1 mmfor longest scales

Solid EarthSciences

> 300 km Approx. 1 cmgeoid error

TIME VARIABLE GRAVITY FIELD

Ocean BottomPressure

> 500 km Seasonal 0.10 mBar pressure 30 day estimate

Deep OceanCurrents

> 500 km Seasonal 1 mm/sec currentvelocity

30 day estimate

Sea Level Rise > 700 km Secular 0.1 mm/yr. Waterlevel

5 year estimate

Evapo –Transpiration

> 300 km Seasonal < 1 cm waterequivalent

30 day estimate

Greenland /Antarctic Ice

Secular

Seasonal

0.4 - 0.8 mm/yr.Ice thickness

3 – 10 mm change

5 year estimate

yearly estimate


2.3 BASELINE SCIENCE MISSION

The science data required to meet the primary science objectives has been defined in Table 2-1.The Baseline Mission Definition, as stated below, shall be the foundation underpinning theMission design and the lower level requirements described in the remainder of this document.

2.3.1 Earth Gravity Field Model

The Earth’s gravitational potential model, characterized by the coefficients of a sphericalharmonic expansion, shall be estimated from the GRACE science data.

2.3.2 Long Term Mean Model

The long term mean geopotential model shall be characterized by a spherical harmonic model toapproximately degree and order 160 or more.

2.3.3 Temporal Variability

The temporal variability shall be characterized by sets of average values of the geopotentialcoefficients over 30 days or longer, to spherical harmonic degrees 100 or less.

2.3.4 Geoid Height Error

A typical set of 30 day average geopotential coefficients shall have an equivalent root meansquare (rms) geoid height error measurement system errors consistent with Table 2-2.

2.3.5 Orbit Characteristics

These models shall be numerically estimated from measurements made between two satellites,co-orbiting the Earth at approximately 300-500 km altitude, at nearly polar inclinations, separatedalong track by 100-500 km, over a lifetime of 5 years.

2.3.6 Atmospheric Soundings

Approximately 200 GPS atmospheric profile soundings per day shall be acquired, subject to datasystem limitations.

2.4 SCIENCE DATA DELIVERABLES

The science data shall be made available during the Observational Phase, after appropriatecalibration and validation, as defined below. All data shall be time-tagged using an on board GPStime reference.


2.4.1 Level 0 data

The Level 0 data records contain raw, unprocessed telemetry data that has been decommutated byDLR/DFD. The following products shall be made available to the JPL SDS and the GFZ DataCentre within 24 hours of receipt by ground stations.

Table 2 – 4 Level 0 Data

Measurement Sample Rate Time Interval

K-Band Phase Data 10 Hz Continuous

GPS Data (Orbit Det.) 1 Hz Continuous

GPS Data (Occultation) 50 Hz Brief intervals; < 200 /day

Accelerometer 10 Hz Continuous

SCA Quaternions 1 Hz Continuous

GPS Data (Ionosphere) 50 Hz (x2 GPS satellites) 3 min

Housekeeping Data All Continuous

Ground based GPS Data 0.033 Hz (most sites for POD)1 Hz (select sites, for occultation)

Continuous

2.4.2 Level 1 data

The Level-1 data products include the calibrated and verified satellite-to-satellite line of sightbiased range and its derivatives, along with the GPS tracking data and initial ephemerides for theGRACE satellites. These data shall be made available to the scientific community no later than 5days following receipt of the Level 0 data.

Table 2 – 5 Level 1 Data

Measurement Product Sample Rate

K-Band Ranging Biased Range & Derivatives 5 s

GPS Occultation SNR + Phases 50 Hz

Laser Ranges to satellite Ground to GRACE distance (2-way) 5 s

Satellite Position & Velocity 5 s

Acceleration 5 s

SCA Quaternions 5 s

GPS Ionosphere Occultation SNR + Phases variable


2.4.3 Level 2 data

The Level-2 data include the orbits for the GRACE spacecraft, estimates of spherical harmoniccoefficients for the Earth gravitational potential, and excess path delay/refractivities from theoccultation data. The Level 2 data shall be distributed within 60 days of completion of dataacquisition, to both the oceanographic community and the wider scientific community.

Table 2 - 6 Level 2 Data

Measurement Product Sample Rate

Geopotential Field Spherical HarmonicCoefficients

30 days or longer

Satellite Position & Velocity Cartesian 5 s

Excess Delay/Refractivity Profiles < 50 Hz

Temperature/Water vapor Profiles < 50 Hz

2.5 MINIMUM SCIENCE MISSION

In the event of major cost or technical problems, it may be necessary to descope the sciencereturns targeted for the GRACE mission. This section defines the minimum level of sciencereturn that will constitute a successful mission. As a minimum objective for a successful mission,the GRACE measurements shall provide for at least an order of magnitude improvement in themean global geoid. The improvement in the marine geoid will enhance dramatically the recoveryof the general ocean circulation and ocean heat flux from satellite altimetry. This improvement isa current requirement of both Earth Observation System (EOS) and World Ocean CirculationExperiment (WOCE).

2.5.1 Earth Gravity Field Model

The Earth’s gravitational potential model, characterized by the coefficients of a sphericalharmonic expansion, shall be estimated from the GRACE science data.

2.5.2 Mean Model

The mean geopotential model shall be characterized by a spherical harmonic model to at leastdegree and order 100.


2.5.3 Geoid Height Error

The cumulative contribution to global geoid height error from harmonic coefficients to degree andorder 70 shall not exceed 1 cm rms.

2.5.4 Orbit Characteristics

This model shall be numerically estimated from measurements made between two satellites, co-orbiting the Earth at approximately 300-500 km altitude, at nearly polar inclinations, separatedalong track by 100-500 km.

2.6 MINIMUM SCIENCE DATA DELIVERABLE

The data deliverables from the minimum mission are identical with those for the baseline mission(Tables 2-4, 2-5, and 2-6), with the exception of the complete deletion of the atmospheric andionospheric occultation products.


CHAPTER 3 MISSION OVERVIEW

This chapter describes the mission elements and top-level mission requirements that respond tothe science requirements of the previous chapter. An implementation drawing from the strengthsand heritage of international partners is briefly outlined in the definition of the Mission Systems.The mission phases (Section 3.2) and top-level requirements (3.3) reflect a design that hasevolved during extensive analysis and simulation of mapping measurement errors and satellitemisalignments into gravity field errors. Section 3.4 lists the environmental requirements for flightelements. Finally, in Section 3.5 a breakdown of error sources associated with the gravity fieldmeasurement is provided as background for the requirements on each of the Mission Systems arestated in Chapter 4.

3.1 MISSION SYSTEMS

The GRACE Project is divided into five systems. Implementation responsibility for thesesystems, described in the Cooperative Project Plan, is schematically outlined in Figure 3-1. TheScience Data System is a distributed System, with some functions occurring in the United Statesand some in Germany. The Laser Retro Reflector instrument shown separately in Figure 3-1, is adirect German contribution to the GRACE Science Instrument System. Each of the five systemsis described in the following paragraphs. Figure 3-2 shows the integration of all systems, flightand ground elements, and interfaces with the Mission Operations System.

NASA

GSFC

UTCSRJPL

DLRGFZ

MissionOperations

System

LaunchVehicleSystem

LRRInstrument

ScienceInstrument

System

Satellite

System

US ScienceData

System

Ger-ScienceData

System

German GRACEProject Office

Figure 3-1 GRACE Systems


3.1.1 Launch Vehicle System (LVS)

The LVS includes the COSMOS (or equivalent) launch vehicle and the personnel, test equipmentand facilities for preparation, integration and launch of the satellites. The LVS is managed by theDLR Launch Vehicle System Manager and supported by the JPL Project and its contractors.

3.1.2 Satellite System (SAT)

JPL administers a contract with Dornier Satellitensysteme to design and fabricate two flightsatellites with selected spares, based on an existing small satellite designed for the CHAMPmission.

3.1.3 Science Instrument System (SIS)

The SIS includes all elements of the inter-satellite ranging system, the GPS receivers required forprecision orbit determination and occultation experiments, and associated sensors such as the StarCameras. This System also coordinates the integration activities of all sensors, assuring theircompatibility with each other and the satellite.

3.1.4 Mission Operations System (MOS)

The MOS consists of facilities and resources of the German Space Operations Center, trackingantennas at Weilheim and Neustrelitz, and other stations and facilities if needed for supportingLEOP and contingency operations. These facilities are used to monitor and control the satellite,perform initial processing of the telemetry data, and deliver all data to the SDS for furtherprocessing and generating science products. In addition to real-time operations, the MOS functionprovides the Central Checkout System for ground testing using command and data interfaces.The operations team also monitors satellite performance and health throughout the duration of themission. Mission operations are conducted at the GSOC control center in OberpfaffenhofenGermany.

3.1.5 Science Data System (SDS)

The SDS functions include science data processing, distribution, archiving and productverification. The SDS is a distributed entity and managed in a cooperative approach by JPL andUTCSR in the US and GFZ in Germany. The cooperative approach includes sharing of processingtasks, harmonization of product archives and validation/comparison of products. All data andproducts for the different processing levels are specified in a product specification document.Data and products to be processed and archived by the SDS include corrected inter-satellite rangeand accelerometer measurements, GPS orbit and occultation data, orbit, gravity field and GPSoccultation products. The SDS also receives, processes and archives ancillary data (e.g.meteorological fields) necessary for data processing and verification.

DLRTracking StationNeustrelitz (NST)53.32N13.06Ed=7.2 m

S-BandDownlink:BPSK1 Mbps / 32 Kbps

S-BandUplink:PCM/PSK4 KBps

S-BandDownlink:BPSK1 Mbps

Mission Control Center (DLR)o Telemetry Processing/Monitoringo Timeline Generationo Commandingo Orbit Determination/Predictiono Orbital Elements Uplinko Orbit Maintenance Maneuvero Attitude Monitoringo Contingency Analysis/Operations

Raw Data Center (DLR)o Telemetry Decommutationo Long-term Raw Data Archiveo Rolling Level-0 Data Archive

Science Data System (SDS)

DLRTracking StationWeilheim (WHM)47.88N11.08Ed=15 m

CommandRealtime & DumpTelemetry

Realtime & DumpTelemetry

TimelineTwo-Line-ElementsDump Telemetry

Housekeeping DataReception Report

Level-0Data Products

ArchivingInformation

TT&CVoice

GRACE B/AGRACE A/B

LEOP and Contingency Support(Telemetry, Tracking & Command)

WallopsFlightFacility

GoddardSpaceFlightCenter

McMurdo(Antarctica)77.84S103.33Wd=10 m

Poker Flat(Alaska)65.11N147.46Wd=11 m

Svalbard(Norway)78.23N15.39Ed=11 m

Mission Operations System(MOS)

Central Checkout System(DLR)o Pre-Launch Test Supporto SCOE Monitoring & Controlo Telemetry Processingo Command Generationo Test Sequencing

On-Ground Tests

Satellite System (SAT)Science Instrument System (SIS)

Launch Vehicle System (LVS)

Injection Orbit Data TimelineReception Reports

OperationsRequests

Figure 3-2 GRACE System Interfaces


3.2 MISSION PHASES

3.2.1 Pre-launch phases

Plans for the design integration and testing of the GRACE satellites, instruments, and groundsystem are described in the GRACE Cooperative Project Plan (327-100). Launch site activitiesprior to liftoff are described in the Launcher Interface Document (327-300).

3.2.2 Launch & Early Operations Phase (LEOP)

The GRACE satellites will be launched together on a single vehicle. This phase, starting withlift-off, provides frequent opportunities for monitoring the satellite status to increase theopportunity for ground-based intervention. In addition to the GSOC ground stations at Weilheimand Neustrelitz, NASA tracking stations at McMurdo, Spitzbergen, Poker Flat, and Kaena Pointwill be used for this purpose.

The LEOP phase nominally ends when the following conditions have been met:(1) Both satellites are in safe, stable orbits with no danger of collision with each other,

with the launch vehicle, or with co-passenger satellites,(2) Both satellites have attained nominal attitude control including successful star-camera

acquisition(3) Nominal uplink and downlink communications are achieved with GSOC stations(4) No anomalies exist that pose a near-term threat to the mission.

In the absence of major unexpected events, the LEOP phase should be completed within the firstfive to ten days after launch.

The sequence of events can be briefly described as follows. Separation from the launcher occursapproximately 30 minutes after liftoff. The satellites are activated by separation from the launchvehicle, and autonomously begin to measure and remove angular rates imparted by the separationdynamics. A low rate S-band transmitter is activated throughout the LEOP period, until turnedoff by ground command. Roughly 20 minutes after separation, the GRACE satellites will enterthe visibility range of the NASA McMurdo station in Antarctica and some status data may bereceived if attitude rates are sufficiently low.

Assuming a separation with the maximum specified angular rate of 3°/s, the satellites will requireup to 90 minutes to damp the rotation and to reach its safe, earth oriented attitude. About 60minutes after separation, both satellites will enter the coverage area of the Weilheim station, butfull attitude stabilization may not yet be established. Therefore, no telecommanding is plannedfor this first pass. As a minimum, telemetry data (at 32Kbits/s) will allow assessment of bothsatellites’ health status.

Following the first pass over Germany, GRACE will reach the coverage area of Spitzbergen andsubsequently that of Poker Flat. These two NASA polar network ground stations can be used for


reception of telemetry and for relaying telecommands issued by GSOC. The following contact withMcMurdo, beginning 144 minutes after lift-off, will serve for verification, that the satellite hasstabilized its attitude and that the angular rates are below 0.5°/s. Subsequent orbits will be used toverify performance of both satellites and issue commands as needed to complete this Phase.

3.2.3 Assessment Phase

Following LEOP, there will be a six month Assessment Phase in which initial instrument check-outis performed, and GPS, KBR, ACC, and SCA data are evaluated. Preliminary calibrations for thesatellite center-of-mass offset, KBR-star tracker axes, and accelerometer bias and scale factor arecomputed. Using these calibrations, preliminary gravity field solutions will be computed, andverified through a combination of internal consistency checks and comparison with limited in-situgravimetry and ocean bottom pressure data.

3.2.4 Observational Phase

Following the 6 month Assessment Phase, the mission will enter the Observational Phase, in whichscience data is routinely gathered from the science payload. This phase will continue until the endof the mission, with the exception of brief (few day or less) periods for orbit maintenance andrecalibrations.

3.3 MISSION REQUIREMENTS

The capabilities are required to achieve the Baseline Science Mission:

3.3.1 Injection altitude

The satellites shall be launched into the lowest altitude orbit that provides a greater than 90 %probability of not re-entering within 5 years. The final injection altitude shall be chosen at least 3months before the scheduled launch.

3.3.2 Satellite separation

Inter-satellite separation shall be maintained within ± 50 km of the nominal value of 220 km. Itshall be possible to acquire valid K-Band ranging data at inter-satellite separation distances between100 and 500 km.

3.3.3 Orbit inclination

The satellites shall be launched into an 87° - 92° degree inclination orbit.

3.3.4 Satellite interchangeability

The leading and trailing satellites shall be capable of being interchanged during on-orbit operations,if necessary.

SCIENCE & MISSION REQUIREMENTS DOCUMENT REVISION: RELEASE

GRACE 327 – 200 MAY 15, 1998 PAGE 22 OF 48

3.3.5 Minimum data capacity

The onboard data system shall be capable of generating and storing a minimum of 100 Mbytes/day,including housekeeping data.

3.3.6 Data gaps

3.3.6.1 Science data shall be generally collected in continuous streams lasting many days.

3.3.6.2 Nothing in the flight or ground system design shall preclude taking continuous records ofmany days duration.

3.3.6.3 Geographically correlated data gaps shall be avoided.

3.3.6.4 Ground system outages (planned or unplanned) shall result in a loss of less than 10% ofthe total planned data acquisition.

3.3.7 Non-science operations

The Observational Phase shall be planned to require no more than 5% of available time during themission for calibrations, orbit maneuvers or other operations which preclude the collection of validscience data.

3.3.8 Fuel utilization

The onboard gaseous nitrogen fuel shall be allocated as follows:40% for attitude control40% for orbit maintenance20% reserve

3.3.9 Single point failures

The satellites and instruments shall be designed such that no single credible failure can preventaccomplishment of the Baseline Science Mission.

3.4 RADIATION REQUIREMENTS

3.4.1 South Atlantic Anomaly (SAA)

All flight hardware shall be designed to maintain nominal operations while within the SAA.

3.4.2 Single-event upsets & latch-up

The satellites and science instruments shall be designed to be immune to latch-up and tolerant tosingle-event upsets.

3.4.3 Radiation Design Margin

The satellites shall be designed for the radiation dose environment specified in the Satellite SystemSpecification (GRACE 327-410) with a Radiation Design Margin (RDM) factor of 2.0.


GRACE 327 – 200 MAY 15, 1998 PAGE 23 OF 48

3.5 DERIVATION OF SYSTEM REQUIREMENTS

Many of the requirements listed in Chapter 4 concern errors associated with the basic measurementmade by the GRACE satellites. This is fundamentally the variation in the separation betweenfloated accelerometer proof masses in each satellite over periods of minutes to many orbits induration. The motion of these masses is influenced both by gravity and by non-gravitational forcesacting on the satellites. Non-gravitational effects due to atmospheric drag, solar and Earth radiationpressure, thermal effects, thruster firings, etc., must be accurately measured by the SuperSTARaccelerometers in order to derive the gravity field information.

Calculation of the Earth’s gravity field from these measurements entails a complex estimationprocess, which to a large extent obscures the relationship between errors of various types. To dealwith this complexity, extensive simulations are being run to characterize the mapping of expectederrors into gravity-field errors. Many of the requirements listed in Chapter 4 are derived from thesesimulations, taking into account the probable costs associated with different levels of performance.

The total gravity measurement error is broken down schematically into three major components inFigure 3-3. Below each of these components are numbers summarizing the current allocation forthis category of error.

Gravity Measurement Error

(Specified in Table 2-2)

III. Precision OrbitMeasurement

< 1cm in plane, absolute< 1mm in plane, relative

Few cm out-of-plane, abs<1cm out of plane, rel

I. Inter-satellite RangeMeasurement

< 4 microns @ twice/rev< 10 microns total @

0.1 to 0.0001 Hz

II. Non-GavitationalAccelerationMeasurement

< 10-10 m/s2 total errorover 0.1 to 0.00005 Hz

Figure 3-3 Gravity Measurement Error Tree

An important aspect of estimating the gravity field is the distinction between measurement noiseand errors that occur at distinct harmonics of the orbital period (“tones”). The effects of broadband


GRACE 327 – 200 MAY 15, 1998 PAGE 24 OF 48

noise can generally reduced by filtering or averaging over many orbital periods. However, tonesthat occur at orbital harmonics can easily map into gravity terms, thereby limiting the accuracy ofthe resulting field measurement. Tones at low orbital harmonics can arise from a number of sources,such as temperature variations around the orbit, interactions between the magnetic field and theaccelerometer, and variations in the level and direction of the atmospheric drag. Regular firing ofattitude control thrusters is also capable of producing tones, most likely at frequenciescorresponding to the highest spatial resolutions achievable by GRACE.

The requirements in Chapter 4 contain a number of specifications relating directly to tone errors.For example, although the accelerometer output noise is consistent with measuring accelerations of10-10 m/s2, tones at orbital frequency harmonics as small as 4 x 10-12 m/s2 can be significant (Table4-3). Achieving this level of performance places a significant burden on the satellite systems thatcan affect the accelerometer output, such as thermal control. It is important that all such effectsarising for orbital motion be understood and controlled. Only in this way can we expect to achievethe full potential of the GRACE measurement.

The first category of error shown if Figure 3-3, inter-satellite range errors, can be further subdivided(Figure 3-4). Characterization of this measurement must include allocations for many

I. Inter-Satellite Range Error

B. SatelliteGeometric

Stability

A. MicrowaveRange

Measurement

D. Ultra-Stable

Oscillator

C. Attitude& Pointing

Control

- Multipath errors

- Electronic drifts

- Ionosphere

- Temperature drifts

- Thermal gradients

- Day/night flexure

- CM motion

- Attitude knowledge

- Pointing control

- K-Band Boresight Knowledge

- Satellite-satellite vector prediction

- Drifts

- Thermal effects

- Magnetic field

Figure 3-4 Inter-Satellite Range Error

effects ranging from the microwave measurement itself (both noise and tone errors) to attitudeerrors and thermal variations affecting the stability of the proof-mass to phase-center separation. As


GRACE 327 – 200 MAY 15, 1998 PAGE 25 OF 48

an example, errors in pointing the K-Band Boresight (KBB) at the other satellite introduce rangeerrors, both through multipath effects on measuring microwave phase, and through purelygeometric projection effects. These pointing errors arise from several effects, including (1) attitudecontrol errors, (2) uncertainty in onboard predictions of the other satellite location, and (3)imperfect calibration of the actual KBB relative to the satellite coordinates. Figure 3-5 provides athird-level allocation for these error sources, as an example of the next level of the error tree.

While pointing inaccuracy can introduce significant errors, temperature effects, particularly thosecorrelated with the orbital period, can be even more important. Thermal expansion and warping asthe satellite passes from sunlight to night and back every orbit result in geometric shifts that maymap directly into the gravity measurement. For this reason the satellite equipment panel and otherstructural elements are fabricated using low expansion composite material (CFRP). In addition, thetemperature of key elements is stabilized to better than 1K (ACC to 0.1K) and the gradient acrossthe accelerometer to the order of 0.1K. Heaters to achieve this level of stabilization represent asignificant portion of the total satellite power budget.

C. Attitude & Pointing Control

1. Alignment StabilityKBB < 0.3 mr

ACC < 0.3 mr

4. On-Board OrbitPrediction Error

< 0.5 mr

2. Attitude ControlAccuracy (rms)

0.5 mr (p/y)10 mr (roll)

3. K-Band BoresightCalibration

< 0.3 mr

Figure 3-5 Attitude and Pointing Control Errors

Figure 3-6 presents the second-level breakdown of errors associated with measuring the non-gravitational forces affecting the motion of each proof mass. Since it is essential to separate outthese forces when calculating the effects of gravity, the performance of the precision SuperSTARaccelerometer is vital to success of GRACE. The Figure lists the key performance parametersneeded in for achieving the required performance of 10-10 ms-2 over a measurement bandwidth of10-4 to 0.1 Hz. Requirements for these accelerometer parameters, as well as those for other scienceinstruments and Systems, are presented in Chapter 4.


GRACE 327 – 200 MAY 15, 1998 PAGE 26 OF 48

II. Non-Gravitational AccelerationMeasurement

CalibrationErrors

EnvironmentalEffects

Dynamic RangeNoise & Tone

Errors

- Temperature drifts

- Temperature gradients

- Magnetic field

- Scale factor

- Axis orientation

- Non-linear terms

- Read noise

- Corner frequency

- Error tones

- Intermodulation

- Proofmass control range

- Measurement Range

Figure 3-6 Non-Gravitational Acceleration Measurement Errors


GRACE 327 – 200 MAY 15, 1998 PAGE 27 OF 48

CHAPTER 4

SYSTEM REQUIREMENTS

This chapter contains the top-level requirements on the Mission Systems listed in Chapter 3.Requirements in this section derive from the Baseline Science Mission requirements (Chapter 2)either directly or through the allocation and analysis process described in Section 3.5.Requirements for the Minimum Science Mission (Section 2.3) are not explicitly addressed.

All requirements are arranged by System, and are comprehensive within each subsection. EachSystem shall respond to these requirements by reflecting them in lower level documentation andassuring their successful implementation. After the first release of this document, all changes shallbe formally documented on Engineering Change Requests (ECR), reviewed by the Mission &System Engineering Team (MSET), and approved by the System Engineering Manager, or theProject manager and Principal Investigators for Class 1 changes.

4.1 THE LAUNCH VEHICLE SYSTEM (LVS)

The GRACE LVS shall provide the following capabilities.

4.1.1 Launch period

The GRACE satellites shall be launched in the summer of 2001

4.1.2 Dual launch

Both GRACE satellites shall be launched as a dual payload on a single launch vehicle.

4.1.3 Orbit altitude

The maximum allowable tolerance for the injection orbit semi-major axis shall be ±16 km (1σ).

4.1.4 Payload mass

The LVS shall be capable of inserting a payload mass of 1050 kg (two satellites plus adapter) into a500 km or lower circular orbit of 87 degree inclination.

4.1.5 Eccentricity

The eccentricity of the insertion orbit shall be e < 0.005.

4.1.6 Inclination

Inclination shall be in the range 878 to 928 (goal = 908).


GRACE 327 – 200 MAY 15, 1998 PAGE 28 OF 48

4.1.7 Ascending node

4.1.7.1 The satellite shall be launched into a noon orbit.

4.1.7.2 The launcher shall meet this requirement with an accuracy of ±138.

4.1.8 Launch window

The launch window shall be at least one hour each day, within a period of at least 30 days.

4.1.9 Launch vehicle interference

The launch vehicle, or co-passenger satellites, shall not impact or contaminate the satellites afterseparation.

4.1.10 Tip-off angular rates

After separation the angular velocities of the satellites shall be less than 38/s around each axis.

4.1.11 Separation velocity

4.1.11.1 The along track velocity difference of the two satellites immediately after separationshall be greater than 0.25 m/s and less than 1 m/s.

4.1.11.1 The along track velocity difference of the satellites and the upper stage immediatelyafter separation shall be greater than 0.5 m/s.


GRACE 327 – 200 MAY 15, 1998 PAGE 29 OF 48

4.2 SATELLITE SYSTEM (SAT)

4.2.1 General Requirements

4.2.1.1 The two GRACE satellites shall differ only in the S-band radio frequencies used forcommunication with the ground, and in the K-Band frequencies used for the inter-satellitelink.

4.2.1.2 Both satellites shall be capable of flying either in the lead or trailing positions, “forward”or “backward” into the residual atmospheric wind.

4.2.1.3 Both satellites shall be designed to an operating lifetime of five years.

4.2.1.4 Expendable or degrading resources (solar array power, cold gas propellant and battery)shall be compatible with the full five-year mission.

4.2.1.5 The satellites and payload adapter shall be compatible with the COSMOS, or equivalent,launch vehicle system, and associated environments, both on the ground and duringlaunch.

4.2.1.6 The launch mass, including two satellites and payload adapter, shall not exceed the LVcapability specified in 4.1.4.

4.2.1.7 The satellites shall be capable of full Science Mode operation at any altitude between 300and 500 km.

4.2.1.8 The ballistic coefficients of the two satellites shall differ by < 0.5% irrespective of whichsatellite is flying “forward” (xs-axis in the direction of orbital motion) and “backward”(xs-axis opposite the direction of orbital motion).

4.2.1.9 All “fine pointing modes” (Figure 4-1) shall have all science instrumentation poweredcontinuously.

4.2.1.10 The operating set-point temperature of the satellite equipment panel shall be settable overa range of 0 to 30°C with increment no larger than 1°C.

4.2.2 Operating Modes

The GRACE satellites can be operated in eight different modes. These modes, and transitionsbetween them, are illustrated in Figure 4-1.

4.2.2.1 Off-Mode (Pre-launch only)In this mode the satellite is completely switched off.

4.2.2.2 Test Mode (TM)This mode shall provide special commands for testing, failure isolation and detection, on the ground.


GRACE 327 – 200 MAY 15, 1998 PAGE 30 OF 48

LaunchMode

OffMode

TestMode

On-Orbit Modes

Fin

e P

oin

tin

g M

od

es

LaunchMode

OffMode

Autonomous, triggered by Separation

(i.e. Power-on-reset)

TestMode

On-Orbit Modes

No

min

al M

od

es

Attitude Hold Mode

CalibrationMode 1(CoM)

ScienceMode

CalibrationMode 2

Sin

gle

Sat

ellit

e M

od

esD

ual

Sat

ellit

e M

od

es

Loss of Attitude orLoss of Fine Attitude

Measurement orCommand

SafeMode

OBDH Reconfiguration orMain Power BusUnder-Voltage or

Command

ShortSurvivalMode

Autonomous

On Command

Sat

ellit

e In

itia

lizat

ion

or

Off

-No

min

al C

on

dit

ion

s

Cmd.Cmd.

Cmd.

End of Cal.

End of Cal.

Cmd.

Figure 4 - 1 Satellite Operating Modes

4.2.2.3 Launch Mode This mode shall be active during the last 2 hours before lift-off and during launch. ThePower Conditioning Unit (PCU) is active and powered via the satellite’s battery, all otherequipment is switched off.

4.2.2.4 Short Survival Mode The Short Survival Mode shall implement the following capabilities:

4.2.2.4.1 Damp angular rates and maintain them below 10 mr/s

4.2.2.4.2 Establish & maintain pointing of the satellite z-axis to the nadir within 158 (1σ)

4.2.2.4.3 Switch off non-essential loads if an undervoltage condition is detected

4.2.2.4.4 Establish & maintain the satellite’s y-z-plane pointed toward the sun with an accuracyof 158 (1σ) if needed to maintain battery charge.


GRACE 327 – 200 MAY 15, 1998 PAGE 31 OF 48

4.2.2.4.5 Maintain on-board data recording.

4.2.2.4.6 Activate and maintain data transmission to ground

4.2.2.5 Safe Mode This mode shall be entered either due to loss of star references, or by ground command.The Safe Mode shall have the following characteristics:

4.2.2.5.1 Switch on all equipment (except cold redundancies and components which areintentionally switched off) and maintain nominal thermal control.

4.2.2.5.2 Record data in the mass memory unit and transmit H/K data at low data rate (32 kbps).

4.2.2.5.3 Establish and maintain attitude control as in the Short Survival Mode.

4.2.2.5.4 The Safe Mode shall be terminated only via ground command.

4.2.2.6 Attitude Hold Mode

The satellites shall be capable of acquiring star references and maintaining a referenceattitude with a pointing accuracy of better than 58 for all axes, even while firing the orbitcontrol thruster(s).

4.2.2.7 Science Mode This mode shall support full science operation including the following capabilities:

4.2.2.7.1 Acquire and maintain a reference attitude with a pointing accuracy of 0.5 mr (1σ) inpitch and yaw, and 10 mr (1σ) in roll.

4.2.2.7.2 Acquire and store housekeeping data and data from all science instrumentationconcurrently.

4.2.2.7.3 Transmit stored and real-time data to the ground without interrupting data acquisitionby any instrument.

4.2.2.8 Calibration Modes The satellite shall execute a number of pre-defined sequences (attitude scans or othermaneuvers) to calibrate elements such as the accelerometer scale factor, center-of-massposition and the K-Band boresight.


GRACE 327 – 200 MAY 15, 1998 PAGE 32 OF 48

4.2.3 Coordinate Systems

The following coordinate systems, illustrated in Figure 4-2, are defined for the satellite system:

(1) xs, ys, zs = satellite coordinates defining locations and orientations of satellite hardware.This coordinate system’s origin is a reference insert on the satellite equipment panel, 125 mm(TBC) below the CM of the SuperSTAR accelerometer proof mass, oriented as follows:

xs = long axis of symmetry of the satellite, positive in direction of Ku/Ka hornzs = vertical axis of symmetry, positive toward radiatorys = completes right-handed coordinate system

(2) xorb, yorb, zorb = orbital reference coordinatesSame origin as satellite coordinates, oriented as follows:

xorb = positive in direction of satellite velocityzorb = positive toward nadiryorb = completes right-handed coordinate system

(3) xa, ya, za = SuperSTAR accelerometer coordinatesThese axes are not aligned with the corresponding satellite letter coordinates due to the definition ofxa as being in the direction of the insensitive axis of SuperSTAR:

xa = insensitive axisza = vertical axis of symmetry, positive toward nadirya = completes right-handed coordinate system

+xa is aligned along +ys, +ya is aligned along -xs, and + za along + zs (nadir pointing).

(4) xcn, ycn, zcn = Star Camera #n coordinates

(5) In addition to these systems, the following frames are defined to elucidate some of thealignment and stability requirements:

SF = Ideal satellite frame defined by a target center-of-mass (CM) at xyz = (0, 0, 0), atarget antenna phase center (PC) point at xyz = (1.4m, 0, 0), and a z-axis constructedperpendicular to the CM/PC vector in the xszs plane. (The actual x separationattained during fabrication can differ from 1.4m, but once the CM and PC points arephysically established, placements and alignments with respect to these preciseactual points are unambiguous.)

KBB= K-Band Boresight defined as the actual vector between the actual CM of theaccelerometer proof mass and the physical phase center.


GRACE 327 – 200 MAY 15, 1998 PAGE 33 OF 48

Figure 4 – 2 Satellite Coordinate Systems


GRACE 327 – 200 MAY 15, 1998 PAGE 34 OF 48

4.2.3 Attitude & Orbit Control

4.2.3.1 The satellite’s attitude shall be controlled using cold nitrogen thrusters augmented bymagnetic torquers (to reduce gas utilization during periods of favorable orientation withthe Earth’s field).

4.2.3.1.1 The thrusters shall be capable of being operated individually, or in coupled pairs

4.2.3.1.2 The roll and yaw thrusters shall be oriented to thrust only in the Ky direction.

4.2.3.2 The satellites shall meet the pointing requirements in Table 4-1 after failure of onethruster or thruster string.

4.2.3.3 The size of the attitude deadband shall be adjustable by ground command over a range ofapproximately 0.5 to 100 mr.

4.2.3.4 The satellite shall be capable of orienting and maintaining the orbit trim thrust to anydirection in the plane perpendicular to the radius vector during the maneuvers, subject topower constraints.

4.2.3.5 Orbit adjustment maneuvers shall be accomplished through extended activation (up to 18hours) of one or both 40 mN orbit trim thrusters.

4.2.3.6 Propellant pressure and temperature measurements shall be provided for in-flightcalibration and performance prediction of the propulsion system

4.2.3.7 The satellite shall be capable of autonomously recovering from angular rates of 100 mr/sabout any axis and from any attitude.

4.2.3.8 The satellite principal axes of inertia shall be aligned with the satellite axes (xyzs) withinthree degrees.

4.2.3.9 The satellites shall achieve the attitude and pointing accuracies in Table 4-1 in the ScienceMode and Calibration Modes.

Table 4-1 Attitude and Pointing Requirements

Component Accuracy (1σσ)

Attitude Control relative toinertial coordinates

Pitch, Yaw: 0.5 mr

Roll: 10 mr

K-Band Boresight (KBB) Pointing Controlrelative to other satellite

Pitch, Yaw: 1.0 mr

Roll: 10 mr


GRACE 327 – 200 MAY 15, 1998 PAGE 35 OF 48

4.2.4 Critical Alignments & Stabilities

The satellites shall meet the alignment accuracies and stabilities of critical dimensions enumeratedin Table 4 – 2. The first column “Allowable Error” specifies is the requirements for the ScienceMode. In some cases, these requirements can be met by ground measurement, while others requirecalibration on orbit. The final column addresses the on-orbit stability requirements for keyparameters.

Table 4-2 Alignments & Stabilities

Component Ground Error b) In-flight CalibrationAccuracy

In-flight Stability

Star Camera wrt SFa) P/Y: < 0.8 mrRoll: < 10 mr

< 0.3 mr c)

--< 0.3 mr< 1 mr

Accelerometerorientation wrt starcamera

< 0.3 mr -- < 0.1 mr

Accelerometer wrt SF a) xyz: < 300 µmangles: <1.2 mr

----

< 100 µm< 0.1 mr

Satellite CM offset wrtaccelerometer

xyz: < 500 µm < 100 µm < 100 µmper 6 mos.

K-Band horn wrt SF a) Angles: < 5 mrPC x:: < 2 mm

PC yz: < 500 µm

------

< 1 mr<3 µm(1/rev)d)

< 400 µm

GPS antenna wrt SF a)

: Primary:

Secondary:

PC xyz: 1 mmAngles: < 10 mrPC xyz: 2 mm

Angles: < 15 mr

< 300 µm

< 300 µm

< 300 µm< 1 mr< 1 mm< 10 mr

Front Panel surface wrtSF

2 mm< 10 mr

----

< 200 µm< 2 mr

a) “SF” refers to the Satellite Frame, defined by a target center-of-mass point at xyz = (0, 0, 0), by a target antennaphase center (PC) point at xyz = (1.4m, 0, 0), and a z-axis constructed perpendicular to the CM/PC vector in thexszs plane. (The actual x separation attained during fabrication can differ from 1.4m, but once the CM and PCpoints are physically established, placements and alignments with respect to these precise actual points are asspecified in the table.)

b) Maximum on-orbit error before flight calibration, if any. Includes ground to orbit changes.c) Knowledge error, after flight calibration, of the actual K-Band Boresight vector between the proof mass CM and

the antenna phase center.d) This specification is 1 µm for a 2/rev error or other signatures.


GRACE 327 – 200 MAY 15, 1998 PAGE 36 OF 48

4.2.5 Calibrations & Adjustments

The following calibrations and adjustments shall be performed during orbital operations. Theaccuracy needed for these calibrations is specified in Table 4-2.

4.2.5.1 In-flight calibrations shall not be required more frequently than every 60 days in order tomeet full performance.

4.2.5.2 Offset between the satellite CM and the accelerometer proof-mass CM shall be measuredto an accuracy of less than 50 micrometers in all three axes.

4.2.5.3 The satellite CM shall be adjustable with a 10 µm step size over a total adjustment rangeof ±2 mm in each axis.

4.2.5.4 The accelerometer scale factor and non-linearity coefficients shall be measurable in flightto a level consistent with Table 4-3.

4.2.5.5 Alignment of the K-Band Boresight relative to star camera and accelerometer coordinatesshall be measured based on correlation between satellite attitude and range deviations.

4.2.6 Command & Data Handling

4.2.6.1 The satellites shall be designed to support ESA CCSDS standard for command andtelemetry, established in coordination with the German Space Operations Center.

4.2.6.2 The satellites shall maximize design compatibility with the CHAMP Onboard DataHandling System (OBDH) for compatibility with GSOC ground elements.

4.2.6.3 The satellites shall transmit downlink data at 1 Mbit/s.

4.2.6.4 An additional rate of 32 Kbit/s shall be provided for transmitting real-time housekeepingdata in safe mode and during initial acquisition.

4.2.6.5 The satellites shall provide the capability to return telemetry data via direct real-timetransmission and by playback of recorded data.

4.2.6.6 The satellites shall support a downlink bit error rate of <10-6 with a 3 dB margin.

4.2.6.7 The satellites shall receive uplink data at 4 Kbits/s.

4.2.6.8 The uplink bit error rate shall be less than 10-8 for commanding with a 3 dB margin.

4.2.6.9 The Mass Memory, used for storing science and telemetry data, shall have a capacity of 1Gbit.

4.2.6.10 The standard data interface between the OBDH computer and payload instruments shallbe the EIA RS-422 interface operating at 19.2 Kbaud (maximum 9560 bits/s).


GRACE 327 – 200 MAY 15, 1998 PAGE 37 OF 48

4.3 SCIENCE INSTRUMENT SYSTEM


4.3.1.1 The instrument system, including thermal expansion of the satellite bus, shall notintroduce errors in the range measurement at a twice-per-rev frequency having magnitudelarger than four (4) microns.

4.3.1.2 At frequencies other than twice-per-rev and between 10 Hz and once-per-rev, theinstrument system, including thermal expansion of the satellite bus, shall not introduceerrors into the range measurement larger than 10 microns.

4.3.2 SuperSTAR Accelerometer (ACC)

The SuperSTAR accelerometer, a modified version of the previously flown ASTRE high precisionaccelerometer, shall provide the measurement performance specified in Table 4-3.

Table 4-3 SuperSTAR Accelerometer

Quantity SpecificationPrelaunch

VerificationMaximum Instability

Temp. variationa) < 0.1o C < 0.1o C —

Magnetic environment < 0.3 Gauss< 10 Am2 @ 0.4m

< 0.3 Gauss —

Transfer functionb)

Bandwidth

Amplitudec)

Group delay errorc)

Dispersion d)

3 Hzflat, < -90 dB

< 0.2 msec

< 0.002o

3 Hz< -90 dB

< 0.2 msec

< 0.002o

—< -90 dB

< 0.1 msec

< 0.002o

Intermodse) -90 dB -90 dB

Output samplesSample rate, S/sBits per sample

Time-tag errorf)

1020

< 0.1 msec

——

< 0.1 msec

——

< 0.1 msec

a) Baseplateb) From input acceleration to output acceleration at 10 S/sc) Calibration error for group delay across signal band, 5 x10-5 to 0.04 Hzd) wrt the linear phase response due to group delay, only across signal band

e) Signal-band intermods from 50 µm/s2 tones at any frequenciesf) wrt satellite clock


GRACE 327 – 200 MAY 15, 1998 PAGE 38 OF 48

Table 4-3 SuperSTAR Accelerometer (cont)

Quantity c) SpecificationPrelaunch

Verification

In-flightCalibrationAccuracy

MaximumInstability

x-axis (LOS)

S(f)..10-20m2 s-4Hz-1

c0.................ms-2

c1......................

c2................m-1s2

c3................m-2s4

Error tones a).....ms-2

Control range... ms-2

Meas. Range.....ms-2

< (1+0.005/f)

< 2 x 10-6

1.0< 10

< 104

< 4x10-12/tone

2.5x10-4

0.5x10-4

< (1+0.005/f)

< 10-6

< 2 %< 10

< 3x105

< 10-12

2.5x10-4

0.5x10-4

——

< 0.01 %—————

< +20 %

< 4x10-12/tone< 0.01% / yr

< +20 %< +20 %< +20 %

——

z-axis (“radial”)

S(f), 10-20m2 s-4Hz-1

c0.................ms-2

c1......................

c2................m-1s2

c3................m-2s4




< (1+0.005/f)

< 2 x 10-6

1.0< 20

< 105

< 4x10-12/tone

2.5x10-4

0.5x10-4

< (1+0.005/f)

< 10-6

< 2 %< 10

< 3x105

< 10-12

2.5x10-4

0.5x10-4

——

< 0.2 %—————

< +20 %

< 4x10-12/tone< 0.2 % / yr

< +20 %< +20 %< +20 %

——

y-axis (“cross-track”)

S(f)..10-18m2 s-4Hz-1

c0.................ms-2

c1......................

c2................m-1s2

c3................m-2s4




< (1+0.1/f)

< 10-6

1.0< 50

< 2x105

< 4x10-12/tone

2.5x10-4

0.5x10-4

< (1+0.1/f)

< 10-6

< 2 %< 50

< 2x105

< 10-12

2.5x10-4

0.5x10-4

——

< 1 %—————

< +20 %

< 4x10-12/tone< 1 % / yr< +20 %< +20 %< +20 %

——

Axis directions….mr b) 0.1 0.1 — 0.1

a) Within signal band, f = 5 x10-5 to 0.04 Hz b) Axis orthogonality and orientationc) The actual acceleration, ad , is related to the ACC output, G, by the relation:

G = c0 + c1 ad + c2 ad2 + c3 ad

3 + S(f)


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4.3.3 Star Camera Assembly (SCA)

The Star Camera Assemblies shall meet the requirements specified in Table 4-4.

Table 4-4 SCA Requirements

Parameter Measurement error

Single axis accuracy relative to LOS (P/Y) 30 µr

Single axis accuracy (Roll around LOS) 240 µr

Noise Equivalent Angle (P/Y) 25 µr

LOS angular rate (P/Y) < 10 mr/s

Nominal Update Rate 1 / s

4.3.4 Instrument Processor Unit (IPU)

4.3.4.1 The GPS function shall provide computed position to an accuracy better than 50m (1σ) intelemetry and for real time use by the AOCS.

4.3.4.2 The IPU shall provide a method of time calibration to an accuracy better than 100 µs

4.3.4.3 The IPU shall provide a dual-frequency GPS range and integrated GPS carrier phase forPrecise Orbit Determination at 10s intervals, of the following precision:

Phase (Ionosphere-free) < 1.0 cmRange (Ionosphere-free) < 100 cm.

4.3.4.4 The IPU shall provide 50Hz, dual frequency integrated GPS carrier phase and amplitudefor atmospheric occultation with the following precision:

Phase (Ionosphere-free) < 0.1 cm Range (Ionosphere-free) < 100 cm.

4.3.4.5 The IPU shall provide 10 Hz measurements of the K and Ka band carrier phases of theranging signal received from the other satellite with a precision of 10-4 cycle.

4.3.4.6 The IPU shall output attitude measurements in the form of quaternions for use by the AOCS at a 1 Hz rate.

4.3.4.7 The IPU shall output attitude solutions no more than 200 ms after receipt of a star camera image frame.


GRACE 327 – 200 MAY 15, 1998 PAGE 40 OF 48

4.3.4.8 The IPU shall report attitude in local coordinates or ECEF coordinates, selectable bysoftware.

4.3.4.9 All IPU data shall be timetagged in seconds since January 6, 1980.

4.3.4.10 The IPU shall output ECEF measurements in WGS-84 coordinates.

4.3.4.11 The phase center of the GPS zenith antenna and the GPS backup antenna shall bedetermined to the accuracies given in Table 4-2 in all dimensions relative to its mountingsurfaces.

4.3.4.12 The mean time between IPU outage (due to SEUs) shall be greater than 30 days. Theduration of each such outage shall be less than 2 minutes.

4.3.5 K-Band Ranging Assembly (KBR)

4.3.5.1 The KBR shall incorporate a dual frequency system selected to enable calibration of theeffects of the ionospheric delay.

4.3.5.2 The KBR shall generate carriers each radiated at greater than 46 dBm EIRP.

4.3.5.3 The KBR shall process the received microwave signal to produce an appropriate basebanddigital signal having C/N0 greater than 60 dB Hz when the satellites are separated by 500km.

4.3.5.4 The KBR shall be designed to reject multipath from the satellite to better than -56 dBrelative to the direct signal.

4.3.5.5 The location of the Ka and K band phase centers shall be known and referenced to themechanical interface of the KBR as specified in Table 4-2.

4.3.5.6 The measured phase of the K or Ka band signal shall vary with temperature of the RFelectronics by no more than 10-4 cycle per Kelvin.

4.3.5.7 The measured phase of the K or Ka band signal shall vary with temperature of theAntenna by no more than 10-4 cycle per Kelvin.

4.3.6 Ultra Stable Oscillator (USO)

4.3.6.1 The USO shall have a long term stability better than 1 x 10-10 per day after 30 days.

4.3.6.2 The Allan Variance of the USO shall be less than the following:0.2 Second 2 x 10-13

2 Second 2 x 10-13

10 Second 1 x 10-13

102 Second 1 x 10-13

103 Second 2 x 10-13

104 Second 6 x 10-13


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4.3.6.3 The USO phase noise, measured at 5 MHz, shall be less than -110 dBc integrated overf > 0.1 Hz away from the carrier

4.3.6.4 Spurious tones shall have a power less than –110 dBc within 0.1 Hz of the carrier.

4.3.6.5 The total power of spurious tones outside of 0.1 Hz from the carrier shall be less than–110 dBc.

4.3.6.6 The USO stability with respect to radiation shall be better than 10-12 per rad.

4.3.6.7 The USO stability with respect to magnetic field shall be better than 10-12 / gauss.

4.3.6.8 The USO shall be shielded to reduce local magnetic field strength to < 0.05 gauss.

4.3.6.9 The radiation sensitive components in the USO shall be shielded so as to receive no morethan 0.4 rad-si due to passage through the South Atlantic Anomaly at solar minimum and500 km altitude.

4.3.7 Laser Retro Reflector (LRR)

The dedicated Laser Retro Reflector (LRR) on-board of both satellites serves for the externalcalibration of the on-board microwave orbit determination system (GPS). Laser ranging data can beused to support the precise orbit determination in connection with GPS data for gravity fieldrecovery as well.

4.3.7.1 The LRR shall deliver sufficiently high return signal to enable both night and daylightlaser ranging to the GRACE satellites over their entire range of altitudes. This shall bewarranted by sufficiently large aperture and a dedicated far field reflection pattern tocorrect for the effect of velocity aberration to a high degree.

4.3.7.2 The optical design of the LRR shall minimize the signature of the returning signal byusing a minimum number of prisms.

4.3.7.3 The single shot accuracy (assuming a ground station of 3rd

generation performance)shall be better than 10 mm. The range correction versus elevation shall be less than4 mm for culmination heights of the GRACE satellites up to 85° over the rangingground station.

4.3.7.4 The tolerance of the effective reflection plane location shall be less than 0.5 mm.


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4.4 MISSION OPERATIONS SYSTEM

A mission control system, dedicated to the GRACE Mission, shall be in place at GSOC (DLR),complete, documented & tested, two months before launch, with the following characteristics:

4.4.1 MOS Functions

4.4.1.1 The MOS shall provide all normal monitoring and control functions for operation of thesatellite, including mission planning, command generation and execution, scheduling,navigation, and satellite performance analysis.

4.4.1.2 The navigation function, including orbit determination and correction, shall be providedby the Mission Operations System.

4.4.1.3 The MOS shall provide data system-processing functions necessary to capture and savethe satellite telemetry, including the data processing system interfaces to JPL and sciencedata processing center(s).

4.4.1.4 The MOS shall supply the Central Checkout System (CCS) prior to satellite electricalintegration and support check-out operations.

4.4.2 Real-time Operations

4.4.2.1 The MOS shall utilize the satellite telemetry, tracking, and command data links of theDLR tracking station at Weilheim and telemetry and tracking at Neustrelitz. It shall becapable of downlinking data from selected NASA stations ( McMurdo, Poker Flat, KaenaPoint and Spitzbergen) in support of LEOP and contingency operations.

4.4.2.2 The MOS shall provide scheduling of the forward (up), return (down), and ground linkservices for satellite communications.

4.4.2.3 The MOS shall monitor the health of the satellites for current status in real-time.

4.4.2.4 The MOS shall capture and save all command and satellite telemetry data

4.4.2.5 The MOS shall command the satellites to perform nominal onboard sequences of eventsand to recover from anomalous behavior defined in the Satellite Operations Handbook.

4.4.2.6 The MOS shall be capable of uploading system software for both the OBDH and IPU.

4.4.2.7 The MOS shall be capable of downloading data from both satellites during the samepass.

4.4.2.8 During LEOP, the MOS shall be capable of uploading commands to both satellites duringthe same pass.


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4.4.2.9 All operations events and data time tagging shall be relative to Coordinated UniversalTime (UTC).

4.4.2.10 Operational orbit determination, and the orbit update parameters to be sent to the twosatellites, shall be provided by GSOC.

4.4.2.11 The MOS shall be capable of acquiring and processing 100 Mbytes of science andhousekeeping data per day for each satellite. (Actual data return depends on manyfactors, including orbit geometry and station availability.)

4.4.2.12 The MOS shall successfully acquire and process an average of at least 50 Mbytes/day ofdata per satellite.

4.4.2.13 The MOS shall provide Level 0 satellite data to the SDS within 24 hours of acquisition.

4.4.3 Navigation

Operational orbit determination, performed by GSOC, fits the tracking data over a chosen timespan; the orbit determined for that time span is called the definitive orbit. The extension of thedefinitive orbit beyond the data span is called the predicted ephemeris. Operational requirementsfor satellite predicted ephemeris data are listed in Table 4-5.

The MOS must produce ephemeris information for mission planning, command generation,navigation planning, maneuver development and evaluation, and tracking pass prediction. Inaddition, it must provide orbital parameters from which each satellite’s onboard computer calculatesrelative positions of both satellites. This information is used to predict the desired K-BandBoresight vector that is relayed to the AOCS function. The relative radial and cross-track positionerrors for each satellite include both the ground based orbit determination and prediction errors(given in Table 4-5) and errors due to the onboard orbit propagator (OOP).

4.4.3.1 The MOS shall determine the GRACE orbit based on GPS position measurements.

4.4.3.2 The predicted ephemerides shall comply with the absolute and relative accuraciesspecified in Table 4-5, even in high solar flux and low altitude conditions.

4.4.3.3 The MOS shall provide operational orbit products fora. Schedulingb. S/C Operationsc. Antenna Pointingd. Maneuver planning


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4.4.3.4 The MOS shall provide orbit control to establish the operational orbit within 30 days afterinjection.

4.4.3.5 The MOS shall provide information about planned maneuver magnitudes and associatedturns and after-the-fact information about maneuver event times, and measured errors inmaneuver performance.

4.4.3.6 The MOS shall perform in-flight calibration of the trajectory correction maneuvers andpredict maneuver performance based on in-flight telemetry and operational orbitdetermination.

Table 4-5 Predicted Ephemeris Data

Source Purpose Parameter Requirement(3 sigma)

TimeSpan

MOS Mission Planning andScheduling

Absolute Along-TrackPosition

± 200 km1 7 days2

MOS Command Generation,Ground Station Acquisition

Absolute Along-TrackPosition

± 10 km 2 days3

MOS Satellite Spacing Control Relative Along-TrackPosition

± 50 km 30 days4

SIS Satellite Pointing Relative Radial / Cross-Track Position

± 50 m5 1 day

1 Equivalent to 30 s transit time error2 14 days adequate for high altitudes/low drag; 7 days needed for low altitude/ high drag3 Expected navigation operations schedule4 Desired minimum time between maneuvers5 MOS OD/OP error budget excluding contributions by the onboard orbit propagator (OOP); equivalent to 0.25 mr at200 km separation


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4.5 SCIENCE DATA SYSTEM

The Science Data System (SDS) will generate science data products, including data processing,verification, distribution, and archiving of science data. The SDS is a distributed entity, with somefunctions occurring at each of JPL, GFZ, and UTCSR. Specific processing tasks are performedonly at one specific entity, while other tasks occur at each of JPL, GFZ, and UTCSR. Unification ofdata and product archives is planned as well as comparison and validation activities. The GFZ andJPL SDS subsystems include elements derived from other projects, including CHAMP, GENESIS,and the PODAAC.


4.5.1.1 The SDS shall generate science data products, including data processing, verification,distribution, and archiving of science data.

4.5.1.2 The SDS shall provide for Precision Orbit Determination, including acquisition oftracking data, development of force and measurement models, and incorporation of PODdata into the science data product.

4.5.1.3 A complete, documented, science data processing system shall be in place 3 monthsbefore launch of the satellites.

4.5.1.4 The SDS shall begin science data processing at the initiation of the collection of sciencedata by the GRACE satellites and continue until 3 months after the end of the mission.

4.5.1.5 The SDS shall receive satellite science and housekeeping data from the MOS.

4.5.1.6 The SDS shall receive and process ancillary data (e.g. meteorological) necessary for dataprocessing and verification.

4.5.2 Data Processing and Distribution

4.5.2.1 The SDS shall process all science data collected by the GRACE satellites

4.5.2.2 After completion of the verification phase, the SDS shall make level 0 and 1 dataavailable to the science investigators within 5 days of collection. Level 2 data shall beavailable within 60 days of acquisition.

4.5.2.3 The SDS shall archive, or arrange for archiving through existing data centers, all sciencedata from the GRACE mission. The JPL archive shall be the Physical OceanographyDistributed Active Archive (PODAAC) of EOSDIS. The German archive shall be at GFZPotsdam.

4.5.2.4 The SDS shall make available all measurement corrections required to make full scienceuse of the GRACE data. These include:

(a) propagation media delay(b) phase center location(c) instrumental alignments


GRACE 327 – 200 MAY 15, 1998 PAGE 46 OF 48

(d) satellite CM location.(e) other corrections determined necessary through further study

4.5.3 Verification

4.5.3.1 The SDS shall have in place, prior to launch, a capability to infer the accuracy of theGRACE orbital ephemerides, the GRACE inter-satellite ranging system, and the gravityfield models derived from that data.

4.5.3.2 Initial verification of the science data shall take place during the Assessment Phase.

4.5.3.3 Periodic verification of science data and products shall continue throughout the entiremission duration.

4.5.4 Precision Orbit Determination

4.5.4.1 POD shall produce an ephemeris for each GRACE satellite with the followingcharacteristics:

(a) Absolute accuracy (with respect to mass center of the Earth, small rotations about adefined frame are allowable) for all science phases of the mission as specified in theScience Data System Requirements (327-710).

(b) Relative accuracy of the ephemeris of one GRACE satellite to the other shall be asspecified in 327-710.

4.5.4.2 POD shall determine and apply necessary calibrations and corrections to measurementdata including those for

(a) propagation media delay(b) phase center location(c) instrumental alignments(d) satellite CM location.(e) other corrections determined necessary through further study


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GLOSSARY

ACC SuperSTAR Accelerometer SubsystemAOCS Attitude and Orbit Control SubsystemAP Assessment PhaseC/N0 Ratio of carrier power to Noise Power Density (bandwidth independent SNR)CCS Central Checkout SystemCCSDS Consultative Committee for Space Data SystemCFRP Carbon Fiber Reinforced PlasticCHAMP A Challenging Microsatellite Payload for Geophysical Research and ApplicationCM Center of Mass = Center of GravityDLR Deutsches Zentrum für Luft und RaumfahrtDFD German Raw Data CenterECEF "Earth Centered, Earth Fixed"EIRP Emitted Isotropic Radiated PowerEOS Earth Observation SystemERC Engineering Change RequestESA European Space AgencyESSP Earth System Science PathfinderGFZ GeoForschungsZentrumGPS Global Positioning SystemGRACE Gravity Recovery and Climate ExperimentGSFC Goddard Space Flight CenterGSOC German Space Operations CenterH/K Housekeeping (data)IPU Instrument Processor UnitJPL Jet Propulsion LaboratoryKBB K-Band Boresight = Actual vector between proof mass CM and phase centerKBR K-Band Ranging AssemblyLEOP Launch and Early Orbit PhaseLOS Line-of-sightLRR Laser Retro-Reflector SystemLV Launch VehicleLVS Launch Vehicle SystemMDRA Mission Definition & Requirements AgreementMOS Mission Operations SystemMSET Mission & System Engineering TeamNASA National Aeronautics & Space AdministrationNMI NASA Management InstructionOBDH On Board Data Handling SubsystemOOP Onboard Orbit PropagatorOP Observational PhasePC Phase Center of the K-Band horn


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POD Precision Orbit DeterminationPODAAC Physical Oceanography Distributed Active ArchiveRDM Radiation Design MarginRF Radio FrequencyRINEX Receiver Independent Exchange FormatS(f) Noise spectrum of ACC outputSAA South Atlantic AnomalyS/A Solar ArraySAT Satellite SystemSCA Star Camera AssemblySDS Science Data SystemSEM System Engineering ManagerSF Satellite FrameSIS Science Instrument SystemSMRD Science and Mission Requirements DocumentSNR Signal-to-Noise RatioTM Test ModeUSO Ultra Stable OscillatorUTC Coordinated Universal TimeUTCSR University of Texas at Austin, Center for Space ResearchWGS-84 World Geodetic SystemWOCE World Ocean Circulation Experiment

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