GRACE-FO (Gravity Recovery And Climate Experiment - Follow-On) / GFO (GRACE Follow-On)
The GRACE-FO (a.k.a. GFO) mission is heavily focused on maintaining data continuity from GRACE and minimizing any data gap after GRACE. Since 2009, the GRACE Follow-On mission is under definition/negotiation by the US-German GRACE consortium (NASA/JPL, CSR/UTexas, DLR, GFZ Helmholtz Center Potsdam). Many studies have been conducted on national and international levels by agencies, academia, institutions, and by industry to investigate new observation techniques fora long term strategy of observing the gravity field from space. 1) 2) 3) 4) 5) 6) 7)
The GRACE-FO project will be executed in the US under the direction of the NASA Earth Science Division (ESD) within the NASA Science Mission Directorate (SMD) and the Earth Systematic Missions Program Office at GSFC (Goddard Space Flight Center). JPL (Jet Propulsion Laboratory) is assigned responsibility for the GRACE-FO project.
The GRACE-FO mission has significant German participation managed by the German Research Center for Geosciences (GFZ). Funding of the German contributions is jointly secured by the Federal Ministry of Education and Research (BMBF), the Federal Ministry for Economic Affairs and Energy (BMWi), the Helmholtz Association (HGF), the German Aerospace Center (DLR) (LRI in kind contributions) and the German Research Center for Geosciences). The GRACE-FO mission will be operated by DLR/GSOC.
In the fall of 2012, after more than 10 years of very successful operation in orbit, the US-German GRACE-1 mission has demonstrated in a very impressive way its outstanding capability to monitor mass motions in the Earth system with unprecedented accuracy and temporal resolution. These results have stimulated many novel research activities in hydrology, oceanography, glaciology, geophysics and geodesy, which also indicate that long-term monitoring of such mass motions, possibly with improved spatial and temporal resolution, is a must for further understanding of various phenomena.
GRACE-1 gave many breakthroughs in the understanding of (Ref. 16):
- Changes in the terrestrial water cycle
- melting and growing of glaciers and ice sheets
- sea level rise and its causes (ice melt, thermal expansion)
- solid Earth (past glaciation, large Earthquakes).
GRACE-FO mission objectives:
• The primary objective is to continue the high-resolution monthly global models of Earth's gravity field of the GRACE-1 mission for an expected length of 5 years.
• The secondary objectives are:
- to demonstrate the effectiveness of a LRI (Laser Ranging Interferometer) in improving the low-low SST (Satellite-to-Satellite Tracking) measurement performance.
- to continue measurements of GRACE radio occultations for operational provision of e.g. vertical temperature/humidity profiles to weather services.
Satellite gravimetry, that is, measuring spatial and temporal change in the gravity field caused by mass variations from space, provides a unique opportunity to advance mass transport studies and improve our understanding of the Earth system. From the GRACE mission, new fundamental insights into the changing mass distribution have been achieved in the first decade of the 21st century. The results from the GRACE mission have actually revolutionized the field of Earth system research and have established the necessity for future satellite gravity missions. 8)
The large-scale mass distribution in the Earth system is continuously changing. Most of the mass transport is associated with well-monitored atmospheric variability, and with the global water cycle. Through this cycle, the ocean, atmosphere, land, and cryosphere storages of water interact through temporally and spatially variable water mass exchanges. The distribution of water mass in these reservoirs changes at timescales ranging from sub-daily to inter-annual, and decadal, and is strongly related to long-term global change, including sea-level rise, loss of land ice, and extensive droughts and floods. These mass variations may indicate a change in the forcing or the feedback mechanisms that moderate the climate. Water mass variations may therefore be considered a proxy for ongoing climate variations driven by natural and/or anthropogenic causes, which has the potential for impacting society very strongly (Ref. 8).
Gravity is determined by mass. By measuring gravity, GRACE shows how mass is distributed around the planet and how it varies over time. Data from the GRACE satellites is an important tool for studying Earth's ocean, geology, and climate. Mass variations are caused either by redistribution of mass in, on or above the Earth's surface or by geophysical processes in the Earth's interior. The first set of observations of monthly variations of the Earth gravity field was provided by the GRACE-1 twin-satellite mission beginning in March 2002. In 2015, this mission is still providing valuable information to the science community. 9)
• Mapping the global gravity field (Ref. 34)
- Static and dynamic components
- Many applications in geosciences
- Orbit determination and tracking
- SST (Satellite-to-Satellite Tracking)
• Recent gravity field satellites
- GRAIL (lunar gravity).
In the framework of a cooperation on future EO technologies and missions, NASA and ESA established an Interagency Gravity Science Working Group of experts tasked to define the overall objectives and the observation requirements for a future constellation with better performance than that of a concept elaborated by a single agency – in fact, a performance better than "the sum of the parts" when these were not jointly defined. 10)
Figure 1 outlines the expected performance (in terms of spatial and temporal resolution) of a next generation gravity constellation, with two cases (three and ten day solutions) separately considered, and also indicates the performance of GRACE and GOCE as a reference. For geophysical signals the targeted geoid height accuracy of 1 mm at scales of 500 km and 150 km, respectively, will give (compared to GRACE) increased temporal and spatial performance and so allow resolving various geophysical processes in far more detail.
Studies have been performed in order to find the constellations of two satellite pairs equipped with laser-based distance metrology (of similar but not necessarily equal performance) capable to deliver the best scientific return. This was supported by closed-loop simulations, with realistic inputs from system design aspects. Constellation candidates have been found, with orbits at different inclinations that optimise the return for the different applications whilst "absorbing" in their gravity solutions most of the disturbing effects of daily and subdaily phenomena with a dedicated parameterisation. This would remove the need for de-aliasing based upon external models, as is current practice for GRACE data analysis, and almost entirely remove the "striping" or "striations" seen in single-pair gravity field solutions.
In November 2012, Airbus DS (formerly EADS Astrium GmbH, Friedrichshafen) has been commissioned to build two new GRACE Follow-On research satellites for NASA/JPL ( Jet Propulsion Laboratory). The goal of the GRACE Follow-on mission is to continue the extremely accurate measurement data collection of the first twin GRACE satellites, which have been in orbit since March 17, 2002. A launch of the mission is planned for August 2017, the minimum mission life is 5 years. 11) 12) 13) 14)
JPL leads the development of the GFO satellite system in partnership (contract) with Astrium GmbH. Astrium provides major elements of two flight satellites based on an existing small satellite design for the CHAMP, GRACE and SWARM missions. The satellite system consists of the following subsystems where most are available with main and redundant units. 15)
TT&C (Telemetry, Tracking & Control): The TT&C activities are carried out using a pyro-deployed S-band receive and transmit antenna, mounted on a nadir-facing deployable boom. Two back-up zenith antennae, one each for transmitting and receiving, along with the appropriate RF electronics assembly, complete the telemetry and telecommand subsystem. The telecommand function of the satellite is designed according to the ESA CCSDS (Consultative Committee for Space Data Systems) packet telecommand standard tailored for G-PUS (GRACE-FO Packet Utilization Standard) with adaptations mutually agreed with DLR/GSOC (German Space Operation Center). The satellites support the command and control capabilities of the MOS (Mission Operations System) by means of:
1) HPC1 (High Priority Commands of priority 1), which are directly handled by the telecommand decoder and by-passes all OBC (On-Board Computer) software
2) Normal telecommands, which will be processed by the OBC on-board software.
After reception of the uplinked command stream via the S-band antenna and the receiver of the RF Electronic Assembly, the telecommands are decoded in above two command categories. The HPC1 are directly executed within the telecommand module of the OBC; i.e. corresponding bi-stable relays are set. The normal telecommands are read from the telecommand handler of the on-board flight S/W via system calls. The telecommand handler further validates and converts the telecommand packets into OBPs (On-Board command Packets). The OBPs are further distributed according to their indicated functionality.
Power subsystem: The power subsystem is responsible for the generation, storage, conditioning and distribution of electrical power in accordance with instrument and satellite bus users needs. Electrical energy is generated using solar arrays of triple junction GaAs (Gallium Arsenide) cells, placed on the top and side exterior surface of the satellites. Excess energy is stored in a battery of Li-Ion cells with a capacity of 66 Ah at mission start. The power bus delivers unregulated power to all users at the respective user interface.
TCS (Thermal Control Subsystem): The TCS consists of 96 independent heater circuits, 128 YSI-type thermistors and 36 PT-type thermistors for in-flight temperature housekeeping, monitoring and heater control, as well as for on-ground verification testing.
OBC (On-Board Computer) subsystem: The OBC subsystem provides processor and software resources, as well as necessary I/O capabilities for AOCS (Attitude and Orbit Control System), power subsystem and TCS operations, including necessary fault detection, isolation and recovery operations.
AOCS (Attitude and Orbit Control Subsystem): The AOCS consists of sensors, actuators and software to:
• Provide adequate knowledge of satellite attitude during all phases of the mission
• Generate on-board error signals to accurately maintain satellite attitude
• Provide necessary orbital control to satisfy the GRACE-FO mission requirements.
The sensors include a CESS (Coarse Earth Sun Sensor), an IMU (Inertial Measurement Unit) and a fluxgate magnetometer, as well as the STR (Star Tracker Assembly) and GNSS receiver.
The CESS provides for omnidirectional, coarse attitude measurement in the initial acquisition, survival and stand-by modes of the satellite. It comprises of six thermistors orthogonally mounted on the satellite. By assuming that the Sun is the hottest object in the field of view and the Earth is the second hottest object in the field of view, the CESS provides the Sun and Earth vectors relative to the body frame at a rate of 1 Hz with an accuracy of ~5º.
The IMU is used in survival modes and provides 4-axis rate information. The unit comprises of three solid-state fiber optic gyros, and three solid-state silicon accelerometers that measure velocity and angle changes in a coordinate system fixed relative to its case. A fluxgate magnetometer provides additional rate information.
The actuators for the AOCS include a CGA (Cold Gas Assembly) and a magnetorquer system. The GN2 (Gaseous Nitrogen) reaction control system includes two pressure vessels, valves, regulators and filters, along with 12 attitude control thrusters and two orbit control thrusters. Three magnetorquers with linear dipole moments of 27.5 Am2 complete the set of AOCS actuators.
µSTR (micro Star Tracker Assembly): The µSTR determines the orientation of the satellite by tracking it relative to the position of the stars. These measurements are used for fine-pointing and on-board control of the satellite. Additionally they are required for the interpretation of measurements made in the satellite reference frame, such as those from the SuperSTAR accelerometer.
The SCA (Star Camera Assembly) consists of three temperature controlled CCD star cameras mounted to the accelerometer, along with the respective baffle assemblies. The STR delivers its video frames to the OBC, which then computes the attitude quaternions. The OBC also acts as the power and command/control interface to the STR. Once switched on and initialized, the STR proceeds with automatic coarse attitude acquisition and then on to fine attitude derivation.
CMT (Center of Mass Trim) assembly: The CMT assembly consists of six (two per axis) MTMs (Mass Trim Mechanisms), associated electronics, and the power and signal harness. Each MTM consists of a trim mass driven on a nut rotor with a stepper motor. The CMT assembly is used to center the CG (Center of Gravity) of the satellite at the center of the proof-mass of the accelerometer after CG calibration maneuvers.
Figure 2: Artist's view of the GRACE Follow-on mission (image credit: Astrium)
Figure 3: General accommodation of the GRACE-FO instruments on the Flexbus platform (image credit: GRACE-FO consortium)
• MDR (Mission Definition Review) and SSR (System Requirements Review) passed in July 2012.
• Formal transition to Phase A occurred in January 2012 16)
The two GRACE-FO satellites with a size of around 3 m x 3 m x 0.8 m and a mass of ~600 kg, will orbit Earth in a co-planar polar orbit, spaced 220 km apart at an altitude of ~490 km.
Figure 4: In 2015, the build phase of the GRACE-FO climate satellites has begun with the delivery of the highly stable carbon fiber composite structure with a mass of ~200 kg in Friedrichshafen, Germany (image credit: Airbus DS, Mathias Pikelj) 17)
Figure 5: Photo of the first of the GRACE-FO research satellites, being prepared for transport in an Airbus cleanroom in Friedrichshafen (image credit: Airbus DS GmbH, A. Ruttloff)
Some project development status dates:
• May 2, 2017: The two GRACE-FO satellites, developed at Airbus DS Friedrichshafen (Germany) for NASA/JPL, were 'given an earful' during recent acoustic tests. The sound impact that builds up during rocket launches was simulated in an echo chamber with a volume of around 1,400 m3 at IABG in Ottobrunn near Munich. 18)
- In four test cycles, the satellites in their flight position were subjected to a sound impact of nearly 140 decibels (dB). In comparison, a pneumatic drill produces 100 dB and the human pain threshold is 130 dB. Both satellites passed the tests with flying colors.
- NASA/JPL in partnership with the GFZ (German Research Center for Geosciences) Potsdam , will send both GRACE-FO research satellites into a polar orbit at an altitude of around 500 km and at a distance of 220 km apart.
Figure 6: Preparations for the acoustic noise test for the two GRACE-FO satellites, which Airbus constructed for NASA/JPL (image credit: Airbus DS GmbH, Mathias Pikelj)
• October 28, 2016: Airbus DS is reporting that the company has finished construction of the first of the two GRACE-FO satellites in Friedrichshafen, Germany (Figure 5). The satellite will now be transferred to Ottobrunn near Munich for several months of operational testing in the IABG (Industrieanlagen-Betriebsgesellschaft mbH) test center. — The second GRACE-FO satellite will be ready for testing in about four weeks. 19)
• July 2, 2015: BATC (Ball Aerospace & Technologies Corporation) delivered the "laser frequency stabilization reference flight units" for GRACE-FO to NASA/JPL. Ball's flight units constitute a subsystem that is part of the high-precision LRI (Laser Ranging Interferometer), a secondary payload aboard the spacecraft. 20)
• The project PDR (Preliminary Design Review) was conducted on Jan. 20-24, 2014. 21)
• Feb. 11, 2013: GFZ signs MoU and provides launcher - Today the Board of the GFZ (German Research Center for Geosciences) has signed the Memorandum of Understanding with NASA Administrator Charles Bolden Jr. to mutually realize the follow-on mission of GRACE (Gravity Recovery and Climate Experiment). This agreement solidifies the joint long-term planning and realization of this unique tandem satellite mission, which will, after having successfully passed the PDR (Preliminary Design Review) in January, now enter it's building phase to be launched in August 2017. 22)
Germany does not only provide the launcher but also optical elements of the LRI (Laser Ranging Interferometer) assembly. The distance measurement between the current dual GRACE satellites which has been performed so far with an accuracy of about two thousands of a millimeter using microwave signals. This shall be improved for GRACE-FO by this new technology by a factor of up to 20. This is important as the prime signal to monitor gravity field variations is the change in separation between both satellites. The LRI design is provided by the Max-Planck-Institute for Gravitational Physics (Albert-Einstein-Institute), the flight hardware will be delivered by STI (SpaceTech GmbH), the GFZ project partner).
Besides NASA/JPL (Jet Propulsion Laboratory) and the University of Texas, GFZ is again responsible for science data processing. Funding of the German contributions is jointly provided by the German Ministry of Education and Research, the German Ministry of Economics and Technology and the Helmholtz Foundation. The German Aerospace Center (DLR) provides in-kind contributions for the LRI. Besides this, GFZ provides Laser Retro-Reflectors for both satellites and funds the mission operations for 5 years. Mission operations will be conducted, as for GRACE, by GSOC (German Space Operation Center) of DLR in Oberpfaffenhofen.
• Formal transition to Phase B occurred on September 3, 2012.
Germany will contribute to the GRACE Follow-on mission the following services/developments:
- Launch provision (a Rockot launch is planned as for GRACE)
- Major contribution to the German/US LRI instrument
- Mission operations and ground system
- Science analysis.
Astrium uses a 3rd generation Flexbus for the GRACE-FO mission (Ref. 16). Each of the GRACE-FO satellites measures approximately 3 m x 2 m x 0.8 m and has a mass of around 580 kg.
Throughout the five-year mission, these measurements will be used to generate an updated model of the Earth's gravitational field every 30 days. In addition, every day each satellite will create up to 200 profiles of temperature distribution and water vapor content in the atmosphere and ionosphere.
Launch: A launch of the twin spacecraft GRACE-FO mission is scheduled for early 2018 from VAFB, CA. GRACE-FO is officially confirmed on a SpaceX-Falcon-9 flight as a rideshare mission with 5 Iridium NEXT satellites, according to a joint announcement of Iridium, NASA and the GFZ (German Research for Geosciences). 23) 24)
On January 31, 2017, Iridium Communications Inc. announced that it has contracted with SpaceX for an eighth Falcon 9 launch. Along for the ride are the twin-satellites of the NASA/GFZ GRACE-FO (Gravity Recovery and Climate Experiment Follow-On) mission, which will be deployed into a separate low-Earth orbit, marking the first rideshare deal for Iridium. An agreement of this kind is economical for all parties, and affords Iridium the ability to launch five additional satellites for its next-generation global satellite network. The rideshare is anticipated to launch out of VAFB (Vandenberg Air Force Base) in California by early 2018.
This rideshare represents a material savings from other supplemental launch options due to the efficiency of sharing the rocket with GRACE-FO, and the incremental cost during the Iridium NEXT construction period is immaterial when considering the avoidance of unspent amounts contemplated under the Kosmotras program. It also affords Iridium the opportunity to rearrange its launch and satellite drifting plan and launch these five satellites directly into their operational orbital plane while increasing the number of planned in-orbit spares by three satellites. Further, this development allows Iridium to complete the whole operational constellation at a faster rate than it would have with seven launches. Iridium will still consider launching satellites with Kosmotras once approvals are available.
"This is a very smart way to get additional Iridium NEXT satellites into orbit," said Matt Desch, chief executive officer at Iridium. "This launch provides added resiliency to our network for not much more than we had planned originally to launch 72 satellites, including two with Kosmotras."
Table 1: Some background of the GRACE-FO launch provider switch 25)
Orbit: A circular co-planar orbit (non-repeat ground track), initial altitude of 490 ±10 km, inclination = 89.0±0.06º, eccentricity < 0.0025, spacecraft separation: ~220 km.
After separation from the launch vehicle, the two satellites are in slightly differently sized orbits that cause a drift apart with a nominal drift rate of about 200 km/day for several days. This drift has to be stopped and reversed to bring them into the required formation of 220 km apart. The operations of the intersatellite ranging instruments — MWI and LRI — require pointing information based on TLEs (Two-Line Elements) of both spacecraft, which are provided via command from the ground. To improve this TLE-based pointing information onboard each satellite a GPS-based frame correction is applied for K-band pointing. For LRI operations this corrected pointing still is not accurate enough. Thus, the offset between GPS-based and TLE-based pointing is computed and modelled on the ground using a parametric fit method. The computed fit parameters are uploaded to the satellites by ground command together with each TLE set to compute a more accurate pointing using the TLEs and fit data. 26)
Sensor complement: (KBR, LRI, ACC, SCA, TriG)
The GRACE -FO mission will supplement the original GRACE -1microwave measurement (KBR) with a new LRI (Laser Ranging Interferometer) system. The laser's shorter wavelength and excellent wavelength stability will improve measurement accuracy by a factor of 25. Flying a laser system in addition to the existing microwave system provides this improved measurement capability, yet retains continuity with original GRACE measurements. The microwave system will remain the primary instrument, with the laser system to act as a technology demonstrator instrument.
GRACE-FO consists of two identical satellites in circular polar orbits. Although the orbit is nominally freely decaying and therefore does not have a constant repeating ground-track pattern, the satellite separation is maintained between 170 and 270 km by occasional orbit maintenance maneuvers. The intersatellite distance variations induced by the differential accelerations on the two satellites in orbit are measured using a two-way microwave link. Two frequencies 24 GHz (K-band) and 32 GHz (Ka-band) are used to enable the removal of variable ionospheric delays. It is important to note that the two satellites effectively comprise a single measurement system (Ref. 33).
The success of the recent satellite gravity missions (CHAMP, GRACE, GOCE)) has demonstrated that the measured Earth's gravity field provides the unique integral measurements of the mass distribution and redistribution in the complete Earth system that are indicators of phenomena such as ice sheet mass balances and isostatic adjustments, terrestrial water storage changes, sea level rises and ocean circulations, tectonic and seismologic activities. As of 2014, the continuity of these measurements is of high importance in many scientific and societal fields. 27)
KBR (K-Band Ranging)
For GRACE-FO, KBR is the key science (and heritage) instrument of the GRACE-1 mission (it is also referred to as the MWI (Microwave Instrument). The objective is ultra-precise satellite-to-satellite tracking (SST) in low-low orbit. The measurement method employed is referred to as DOWR (Dual One Way Ranging). In this approach, each of the two satellites transmits a carrier signal and measures the phase of the carrier generated by the other satellite relative to the signal it is transmitting. The sum of the phases generated is proportional to the range change between the satellites, while the phase variation due to long-term instability in each clock cancels out. 28)
K-band has a radio frequency of about 24 GHz and Ka-band is near 32 GHz. The GRACE K- and Ka-band frequencies are in an exact 3-to-4 ratio on each satellite. The KBR system can measure the range (with a bias) to the µm level.
Variations in the gravity field cause the range between the two satellites to vary. The relative range is measured by KBR (a microwave link which is integrated with a GPS receiver). The measured range variations are corrected for non-gravitational effects by an accelerometer called SuperSTAR. KBR consists of the following elements: USO (Ultra Stable Oscillator), the MWA (Microwave Assembly), the horn, and IPU (Instrument Processing Unit). The IPU and the SPU (Signal Processing Unit) constitute the heart of the instrument system. 29) 30)
Figure 7: A schematic drawing of the GRACE instrument system (image credit: NASA/JPL)
Legend to Figure 7: The IPU, SPU, KBR and ACC are internally redundant, and the ultra-stable oscillator (USO) is redundant.
USO (a heritage instrument of GRACE-1/GRAIL) serves as the frequency reference. The microwave assembly, or sampler, is used for up-converting the reference frequency to 24 and 32 GHz; down-converting the received phase from the other satellite; and for amplifying and mixing the received and the reference carrier phase. The horn is used to transmit and receive the carrier phase between the satellites. - The IPU is used for sampling and digital signal processing of not only the K-Band carrier phase signal, but also the signals received by the GPS antenna and the star cameras. Each satellite transmits carrier phase to the other at two frequencies, allowing for ionospheric corrections. The transmit and receive frequencies are offset from each other by 0.5 MHz in the 24 GHz channel, and by 0.67 MHz in the 32 GHz channel. This shifts the down-converted signal away from DC, enabling more accurate measurements of the phase. The 10 Hz samples of phase change at the two frequencies are downlinked from each satellite, where the appropriately decimated linear combination of the sum of the phase measurements at each frequency gives an ionosphere-corrected measurement of the range change between the satellites.
Figure 8: Block diagram of the dual one-way ranging system (image credit: NASA, Korea Aerospace University) 31)
LRI (Laser Ranging Interferometer)
The most significant difference to the GRACE-1 mission is the additional inclusion of an experimental LRI as a technology demonstrator. The objective of LRI is to measure the same range fluctuations as the KBR (K-Band Ranging) instrument, but with less noise, and in addition provides precise measurements of the relative pointing of the satellites to each other. As the LRI is an experimental complimentary instrument, it has less stringent lifetime and reliability requirements than the primary KBR instrument. It is hoped, however, that both instruments can be operated in parallel for a large part of the mission life and thus produce better final results by new combinations of their data, mutual calibration and consistency checks etc. If the LRI performs better than the KBR, as hoped, it will probably be used in future GRACE-type missions as primary instrument. 32) 33) 34) 35) 36) 37)
In addition, LRI will be the first spaceborne laser interferometer to measure distance variations between remote spacecraft, with significant implications for other missions using the same basic measurement such as LISA (Laser Interferometer Space Antenna).
The LRI instrument is a joint US-German development with additional contributions from Australia. NASA will provide the laser, cavity assembly, and ranging processor; the German Space Program (DLR) will provide the measurement optics and steering mirror assembly along with instrument integration and testing. The addition of the LRI instrument is complicated by the fact that its design must be adapted to the existing spacecraft platform design.
The various collaborative institutions of the RFI (Radio Frequency Interference) project are simply referred to as the "RFI consortium", the main participants are: NASA/JPL (Jet Propulsion Laboratory), Pasadena, CA, USA; Max Planck Institut für Gravitationsphysik, Hannover, Germany; STI (SpaceTech International) Immenstaad, EADS Astrium Immenstaad, Germany; DLR (German Aerospce Center), GFZ (GeoForschungsZentrum), Potsdam, Germany; Tesat Spacecom, Backnang, Germany; ANU (Australian National University), Acton, Australia.
The instrument hardware on both spacecraft is identical. It consists of a single frequency laser, a Fabry Perot reference cavity, a triple mirror assembly (TMA) and an[OBS (Optical Bench Subsystem), consisting of the optical bench assembly (OBA) and associated electronics (OBE)] and the instrument electronics, the so called laser ranging processor (LRP). Figure 9 shows an overview sketch of the whole instrument configuration. 38)
Figure 9: Overview of the LRI architecture (image credit: LRI consortium)
On one spacecraft (the master S/C), the laser is frequency locked to the reference cavity via a fiber link and the beam is launched via the fiber injector onto the OBA. On the OBA it is directed to the quadrant photo receiver as well as to the other spacecraft by means of a beam splitter.
The top level instrument requirement is the ranging noise requirement of:
From this requirement several key and driving design requirements for the free space optics part of the LRI (which is the OBA, the TMA and the instrument baffles) and the OBA are derived and listed in Table 1.
Table 2: Key and driving LRI optical design requirements
The beam diameter requirement is derived from the overall power budget, taking required power at the distant spacecraft and heterodyne signal loss under the influence of co-alignment errors into account. The same is true for M2, the required wavefront planarity and the beam co-alignment. The field of regard defines the required laser beam steering range relative to the S/C because of pointing misalignments after launch (caused by uncertainty of the S/C pointing due to star camera errors, launch settling and optimized pointing for the microwave ranging system) and spacecraft AOCS pointing performance of about ±300 µrad.
The straightforward approach of routing the laser beam back and forth along the connecting line between the two spacecraft's centers of mass is not possible in GRACE-FO since that path is blocked by existing components such as the KBR horn antennas. Hence, an alternative design is used, the so-called 'racetrack' configuration shown in Figure 10 (Ref. 32).
The key geometric element of this design is the system of three mirrors on each spacecraft used to route the beam through each spacecraft. These three mirrors form a corner cube configuration, i.e. their mirror planes are perpendicular to each other. For small rotations of the device, the beam spots are nearly stationary and the corner cube doesn't need to be complete — only the actually reflecting segments need to be present. This arrangement has a number of useful special properties: The intersection point of the three mirror planes (the vertex of the retro-reflector), can be located outside the mirror device, allowing the effective fiducial measurement point to be placed inside the accelerometer housing. Additionally, two important parameters are invariant under rotation around the vertex, namely:
• the round-trip path length, which is twice the distance between the beam starting point and a plane normal to the beam direction and intersecting the vertex
• the propagation direction of the reflected beam which is always anti-parallel to the incident beam.
These are essential elements to construct a system with high immunity to spacecraft attitude jitter.
The second fundamental ingredient is the frequency domain scheme of offset phase locking. The laser in one of the spacecraft (called S/C 1) in the otherwise perfectly symmetrical setup is designated as master laser and locked in frequency to an on-board reference cavity, to minimize noise originating from laser frequency fluctuations. When arriving at the other spacecraft (S/C 2), the light has picked up a Doppler shift of a few MHz, either positive or negative. Due to the beam divergence and the small apertures necessary to cope with misalignments, less than 1 nW out of 25 mW laser power, are received at S/C 2, such that direct retro-reflection is infeasible. Instead, a local laser is phase-locked to the incoming light with a frequency offset chosen such that after picking up once more the Doppler shift (with the same sign) on the way back to S/C 1, the beat note between the light arriving back at S/C 1 and the local, stabilized, laser at S/C 1 is in the sensitive frequency range of the phasemeter (approximately 2 - 15 MHz).
Figure 10: Optical layout for the laser ranging interferometer (Laser frequency stabilization subsystem not shown), image credit: LRI consortium
Legend to Figure 10: The microwave ranging system is labelled K/Ka band ranging is centered on axis.
Optical bench: The optical bench features the ultra-stable fiber injector (FIA), the beam fine steering mirror (FSM), superposition of the laser signal, imaging optics as well as the photoreceiver frontends (PRFs) with the quadrant photodiodes. Its key functions are to:
- launch the laser beam to other S/C (via TMA)
- superimpose the received beam and the local oscillator on the PRFs
- convert the optical heterodyne signal to electrical signal and provide to LRP
- enable the beam steering (by LRP command) to cover the field of regard
- enable closed loop beam steering (by LRP) to keep the received beam and the transmitted beam co-aligned
- image the OBA entrance aperture and the FSM onto the PRFs (for beam size adaption, avoidance of beam walk, avoidance of diffraction rings).
Figure 11 shows the setup of the optical bench hardware during the assembly and test campaign at STI. The individual components are either placed in or attached to the optical bench body out of titanium. This approach has been selected to achieve a low thermal noise (due to the high thermal mass enclosing the components) and high beam pointing stability (due to the symmetry of the titanium body with respect to the plane containing the beam). The optical components are out of BK7G18, which is well CTE-matched to titanium, further improving the thermal stability of the OBA.
Figure 11: Optical bench illustration, left: top down view, showing the optical path; right: with graphical representation of the titanium body (image credit: LRI consortium)
The first function of the lens system is to simultaneously image both the receive aperture and the steering mirror pivot plane onto the photodetector such that a beam tilt at either the aperture or steering mirror leads to a pure tilt (i.e. without beam walk) on the photodetector. To achieve the correct imaging for both beams, the steering mirror and the aperture on the optical bench must have the same effective distance from the lens system. In addition the imaging system is used to match the beam size to the (smaller) quadrant photodetector.
A high-fidelity optical imaging in the classical sense is not necessary here, since the quadrant photodetector has only four 'pixels'. The main requirement is a small coupling of tilt into the length measurement. Aberrations of the imaging system are largely common to both beams. The remaining residual coupling is mainly caused by the different spatial profile of the two beams and residual beam misalignments.
The path length through a single beamsplitter is angle dependent which leads to a coupling of the spacecraft attitude jitter into the round-trip length measurement. The coupling factor for yaw is dominant and amounts to 2.2 mm/rad for a 7 mm thick fused silica beamsplitter. For pitch, the coupling factor is quadratic and nominally at a turning point. These coupling factors as a function of spacecraft rotation are shown in Figure 12.
The yaw coupling factor of 2.2 mm/rad is too large by about 20 - 50 times for typical assumed spacecraft pointing jitter and noise budgets. It can be almost completely removed by adding a compensation plate made from the same material as the main beamsplitter but rotated by 90 degrees. This shifts the coupling factor for yaw also to a nominal minimum working point. The coupling factor with a 1 mrad error from the nominal minimum amounts to only 10 µm/rad with the compensation plate. These coupling factors were estimated using simple raytracing along the sensitive path of the interferometer and for nominally perfect geometry.
Figure 12: Coupling factors for pitch, yaw and roll into the sensitive path length for only the beamsplitter (shown on the left) and for both the beamsplitter and the compensation plate (shown on the right), (image credit: LRI consortium)
Figure 13: Illustration of the Laser subsystem (image credit: Tesat Spacecom)
Satellite pointing: The orientation of the GRACE satellites to each other is controlled to a level of a few mrad based on the orbits of the satellites and the star cameras. The actuators are magnetic torque rods and cold gas thrusters in case the Earth's magnetic field lines are unfavorably aligned or if the disturbances are such that the magnetic torque rods have insufficient control authority. GRACE-FO will use the same basic attitude control scheme. While this level of attitude control is sufficient for the microwave ranging system due its wide beam and receive field of view, these misalignments are too large for the laser interferometer, which requires about 100 µrad pointing accuracy. Therefore, active pointing control is required.
For the LRI instrument of the GRACE Follow-on mission, the pointing control must be implemented internally in the instrument itself. The DWS (Differential Wavefront Sensing) function of the interferometer is used as sensor, and a single steering mirror per spacecraft as actuator, that simultaneously optimizes the interferometer contrast in the receive path as well as the transmit beam pointing.
Beam pointing control does not eliminate the coupling of spacecraft jitter completely from the measurement, since the spacecraft still has a variable physical misalignment which leads to varying beam paths. For example, an offset of the triple mirror vertex from the accelerometer reference point (the effective rotation point after using the accelerometer output in data processing) leads to a coupling of satellite attitude jitter into the measured round-trip length variations. Typical numbers lead to a static alignment requirement of the order of 100 µm for the vertex location in the two directions orthogonal to the beam axis.
Other effects that also lead to such coupling include effects on the optical bench such as incomplete compensation of the beamsplitter path length error by the compensation plate, differential-mode aberrations in the beam compressor lens system, stray light and diffraction effects in the triple mirrors and baffles, etc. It is expected that these effects will be mostly a systematical and reproducible function of the 2D pointing error of the spacecraft, such that they can be at least partially removed in data post-processing if an accurate measurement of the instantaneous pointing error is available. In the LRI, the feedback signal fed to the steering mirror will serve this purpose and will therefore be recorded and downloaded as science data.
DWS (Differential Wavefront Sensing) and Steering Mirror Control: DWS is a well known technique for measuring with high sensitivity the angle between two wavefronts in a laser interferometer. Figure 14 illustrates the basic principle of the DWS technique.
Figure 14: Differential wavefront sensing principle (image credit: LRI consortium)
The photodetector is split into 4 segments which are connected to separate phasemeter channels. The average of the measured phases represents the path length signal similar to what a single-element diode would produce, and the difference between 'left' and 'right' or 'top' and 'bottom' represents the angle between the wavefronts in horizontal and vertical direction, respectively.
The conversion factor from geometrical angle to electrical phase difference can be very large and contributes to the high sensitivity of the method. With some simplifying assumptions it is given by: k ~16 r (3λ)-1 . → for the case of two flat-top beams, where "r "is the beam radius, and "λ" is the optical wavelength.
Numerical simulations yield factors of about 20000 rad/rad for the real situation of one Gaussian and one flat-top beam and the planned LRI parameters.
Further benefits of the DWS sensing method are that the result is to first order independent of small lateral movements of the photodiode (or, equivalently, a common-mode beam walk of both beams), and that many noise sources of the path length measurement such as laser frequency noise cancel. In laboratory experiments with comparable parameters, sensitivities of a few nrad/√Hz have been reached.
In the LRI case, the DWS signal represents the relative angle between the received beam and the local oscillator beam. Since the optical bench is rigidly mounted in the spacecraft, the angle of the received beam w.r.t. the optical bench represents the misalignment of the local spacecraft's optical axis w.r.t. to the the other spacecraft. The angle of the local oscillator beam w.r.t. the optical bench, on the other hand, is directly controlled by and only depends on the steering mirror. Figure 15 illustrates the principle of operation of the steering mirror control loop.
Figure 15: Function of the steering mirror control loop (image credit: LRI consortium)
Legend to Figure 15: From left to right: (1) nominal situation with spacecraft perfectly aligned, (2) spacecraft misalignment produces a non-zero DWS signal on the photodetector, (3) feedback to steering mirror makes DWS signal zero and makes the outgoing beam parallel to the incoming beam. SM - steering mirror, BS - beamsplitter, CP - compensation plate.
Three situations are shown in Figure 15. The first is the nominal situation where the incoming beam and the local beam are aligned. The illustration in the middle shows the situation when the spacecraft alignment changes, leading to a non-zero DWS signal and also to a misalignment between the incoming and outgoing beam directions. Consequently, both the local heterodyne efficiency and the power received at the other spacecraft decrease. Closing the steering mirror control loop (right part of Figure 15) zeros the DWS signal by rotating the steering mirror until the local and incoming beams are parallel on the photodetector. This yields not only optimal interference contrast on the photodetector, but also ensures that local and received beams are parallel at the beamsplitter BS.
Together with the second property of the triple mirror ("the propagation direction of the reflected beam which is always anti-parallel to the incident beam"), this leads to the outgoing beam being parallel to the incoming beam, independent of the local spacecraft orientation. The transmitted beam thus always points back to the other spacecraft, and the local spacecraft can be considered an "active retroreflector". This very useful behavior only occurs when no lenses or other curved optical elements are included in the main round-trip beam path, since such elements would change the beam direction. As a consequence, the same unmodified laser beam needs to serve both as transmit beam the propagation direction of the reflected beam which is always anti-parallel to the incident beam.
Table 3: Parameters chosen as baseline design
Legend to Table 3: The value for the "Transmission efficiency for receive path" parameter includes the reflection from the beamsplitter and loss in the imaging optics, but neither the photoreceiver efficiency nor the heterodyne efficiency. The heterodyne efficiency is, however, included in the effective power computation. The "Transmit efficiency" includes the fiber power splitter, transmission through the components on the optical bench and the TMA.
The information about the real spacecraft pointing is still required to compensate residual coupling effects of that pointing into the path length measurement. As the DWS signal has been driven to near zero by the loop, it is not useful for that purpose any more, but the information must be obtained from the feedback signal to the steering mirror instead. This requires either a very linear and predictable mirror actuator, or an internal sensor within the actuator that yields an accurate representation of the actual mirror angle.
Laser frequency stabilization subsystem:
In laser frequency stabilization systems, an independent, stable frequency reference is compared to the frequency of the laser. This measurement is then used by a feedback system to change the laser frequency so that it matches the independent reference. The stability of the frequency reference sets the ideal limit to the frequency stabilization system. 39)
A common means to provide a stable laser frequency is to form an optical cavity by attaching mirrors to the ends of a 'spacer' made of a material of with a very low thermal expansion coefficient. The laser frequency is locked to the length of the cavity by comparing the laser output frequency with light which has resonated in the cavity. Glasses such as zerodur or ULE have thermal expansion coefficients of order 3 x 10-8/K. The optical cavity can be insulated from external sources of heat, so that temperature fluctuations experienced by the cavity are reduced (Ref. 35).
The laser frequency stabilization goal adopted for this development is a frequency noise power spectral density of 30 Hz/√Hz over the frequency range of interest, 10 mHz to 100 mHz, corresponding to length scales of 700 km to 70 km. Performance at lower frequencies (longer length scales) is also of interest but is expected to be limited more by atmospheric drag calibration than by laser frequency stability. Frequency stability is limited by brownian motion noise of the spacer to about one order of magnitude lower than the adopted goal. Performance approaching the brownian-noise limit has been achieved in laboratories. The project adopted a less ambitious goal to ensure that the system can survive launch vibration and on-orbit thermal extremes.
The cavity is based on a design available from Advanced Thin Films and used successfully in laboratory tests. The spacer is made out of ULE. The spacer cross-section tapers from the middle towards each end and is designed to be mounted vertically to reduce distortions due to ground vibration. While mounting in the center is not as important for use in space, the vertical mounting has potentially improved performance when testing prior to launch. The length of the spacer is 77.5 mm. End mirrors were attached using optical contacting. The mirror coating was chosen to achieve a finesse of about 10,000. This is lower than typical for laboratory use, but eases requirements on alignment of the injection optics, which must survive launch and the space thermal environment.
The cavity is mounted from a flange by titanium flexures bonded to the spacer, as shown in Figure 16. The flange is part of a titanium vacuum enclosure which serves as the first stage of a two-stage thermal isolation enclosure. It provides a controlled vacuum environment to reduce fluctuations of refraction within the cavity and eliminates convection between the cavity and the vacuum enclosure. Laser light is injected into the cavity via a single-mode optical fiber. An optical bench made out of zerodur contains optics to match the spatial mode from the output of the fiber into the cavity. The optical bench is also made of zerodur and is mounted to the cavity using titanium flexures. A photodetector is mounted to the vacuum flange to allow the alignment of the injection optics to be checked using light transmitted through the cavity. Light output from the cavity is mixed with the input light from the laser with a beam-splitter and transmitted through a window on the bottom of the vacuum enclosure to a second photodetector.
Figure 16: Laser frequency stabilization cavity and optical bench mounted on vacuum flange (image credit: NASA/JPL)
Figure 17 shows the cavity mounted within the titanium vacuum enclosure which has been coated with gold to reduce radiative heat transfer between the cavity and the enclosure can and also between the vacuum can and the external aluminum thermal shield shown next to the cavity. The current implementation includes a vacuum valve and ion gage on the vacuum flange for testing purposes. For the planned flight unit, the valve and gage will be removed to reduce volume and mass. Heaters and temperature sensors are attached near the top and the bottom of the outer aluminum enclosure to control the temperature of the ends of the enclosure.
Figure 17: Laser frequency stabilization cavity installed in vacuum can, with external thermal shield shown adjacent (image credit: NASA/JPL)
Status of LRI in fall 2014 (Ref. 36):
• LRI has a transponder architecture:
- Cavity stabilization on S/C-1, offset phase lock on S/C-2
- Tripple mirror and steering mirror keep pointing to other S/C, independent of local S/C rotation.
• LRI design well chosen to fulfill the requirements
- LRI will improve the distance measurement in comparison to KBR (K-Band Ranging)
- Latest tests show the subsystems are working together as predicted
- Flight model production has started.
• Better thermal stability through a new electronic design
• Improvement of thermal characterization (on ground and through temperature housekeeping)
• On-ground characterization of differential bias and scale factor between both FM accelerometers.
ACC is designed to measure the weak level of acceleration sustained in orbit by the satellites of the GRACE‐Follow‐on mission. The electrostatic space accelerometer of ONERA is very similar to the SuperSTAR instrument operating for the last 10 years on board the twin GRACE-1 satellites. The ACC also takes into account the return of experience learned from the six GRADIO accelerometer models composing the three axis gravity gradiometer of the GOCE mission of ESA, which demonstrates both the robustness and the achieved performances of this family of space instruments. In addition to thermal stability improvement and of better operation reliability, the GRACE Follow-on instrument shall exhibit a more accurate pre launch calibration, a lowered noise level at low frequency bandwidth and a better stability in order to fit with the laser interferometer ranging demonstration.
The GFO accelerometer is similar to the GRACE one. Its operation is based on highly accurate position and attitude control of an inertial Titanium alloy proof mass of 72 gram. Exhibiting a resolution better than 10-10 ms-2 Hz-1/2, the accelerometer measures the satellite's linear acceleration along three axes (denoted by: XACC, YACC and ZACC) and angular acceleration about these same axes (respectively Φ, Θ, Ψ). The three accelerometer axes have different characteristics. The XACC axis is designed to withstand a high voltage to levitate the proof mass and compensate Earth's gravity during ground tests of the YACC and ZACC axes. These axes are defined with respect to the geometry of the electrodes surrounding the proof mass (Ref. 27).
The accelerometer outputs are obtained from the measurements of the analog voltages which provide the electrostatic forces needed for keeping the proof mass at the center of the electrode cage. The accelerometer is thus constituted by two subsystems: the SU (Sensor Unit) and the ICU (Interface Control Unit).
Figure 18: Schematic of the GFO Accelerometer (image credit: ONERA)
The SU consists of the parallelepipedic proof mass, suspended inside an electrode cage made in gold coated ultra low expansion silica. The motion of the proof mass is detected by capacitive sensors and servo-controlled using the same electrodes to apply the quasi DC actuation voltages. This core (mass and electrode cage) is mounted on a sole-plate and enclosed inside a tight housing inside which a good level of vacuum is maintained with a getter material. During on ground storage, a small ion pump also helps to maintain the vacuum and is also used to monitor the level of residual pressure through its current measurement. The SU mechanical (SUM) assembly is surrounded by four boxes containing the FEEU (Front End Electronics Unit). The FEEU boxes contain respectively:
- the position sensing, analog control loops and Analog to Digital Sigma Delta converters for the 2 YACC and ZACC ultra sensitive axes and Φ rotation acceleration outputs
- the position sensing, analog control loops and Analog to Digital Sigma Delta converters for the XACC less sensitive axis and Θ and Ψ rotations acceleration outputs
- the precise and stable references voltages Vp for actuation linearization, Vd for position sensing and Vref for Analog to Digital accurate conversion
- the FPGA which leads the digital operations and insures the interface with nominal and redundant data links to the satellite.
The SU, sensor mechanics associated with the front end electronics units, is thermally controlled in a range of ± 0.1ºC over one orbit.
The ICU (Interface Control Unit), which could be far from the SU, includes only the DC/DC converters for the generation of the secondary regulated power lines (+3.3 V, +5 V, ±15 V, ±48 V) from the satellite nominal and redundant power buses.
Figure 19: Accelerometer principle of operation (image credit: ONERA)
Performance: The performances required for the GFO accelerometer are the same as the GRACE ones. They are mainly driven by the maximum acceleration control range of ±2.5 x 10-4ms-2 requested due to the system excitation by the spacecraft thrusters and by the measurement range, respectively, of ±5 x 10-4ms-2 along the XACC axis and of ±5 x 10-5ms-2 along the YACC and ZACC axes.
Figure 20 presents the exhaustive sources of noise of the acceleration measurement along the satellite to satellite axis (drag axis).
Figure 20: ACC noise performance along S/C to S/C axis (image credit: ONERA)
Development status of GFO accelerometer:
• September 2015: The GFO accelerometer will be delivered to Airbus DS in December 2015. 42)
• The CDR (Critical Design Review) was held in September 2014 and is successful (Ref. 41).
• The accelerometer EM (Engineering Model) has been built in 2013 and the verification of the design and performance is under completion since the beginning of 2014. The CDR (Critical Design Review) was held in September 2014, followed by the realization of the 3 accelerometer Flight Models : one for each satellite and a spare model. The delivery of the accelerometer is foreseen for September 2015, for a launch in August 2017.
Next gravity missions: As GFO prepares, the future NASA mission GRACE-2 (or GRACE-II), which will not start before 2020, ESA is presently performing preliminary studies of the NGGM (Next Generation of Gravity Mission). 43)
SCA (Star Camera Assembly):
SCA is of GRACE-1 and CHAMP heritage. The objective is the precise measurement of satellite attitude. SCA consists actually of two DTU (Technical University of Denmark) star camera assemblies (2 cameras with sensor heads), each with a FOV of 18º x 16º and one DPU (Data Processing Unit). Both assemblies are rigidly attached to the accelerometer, and view the sky at a 45º angle with respect to the zenith, on the port and starboard sides. The SCA is used for both: science as well as AOCS; the two assemblies provide the primary precise attitude determination for each satellite. The baffles are used to avoid the degradation due to solar heating. SCA measures the S/C attitude to an accuracy of < 0.3 mrad (with a goal of 0.1 mrad) by autonomous detection of star constellations using an onboard star catalog.
TriG-RO [Tri-GNSS (GPS+ Galileo+GLONASS) Radio Occultation receiver]
NASA/JPL (Jet Propulsion Laboratory) is developing a next-generation GNSS space science receiver, the TriG receiver. The receiver will upgrade the capabilities offered by the current state of the art BlackJack/IGOR GPS science receivers in order to meet NASA's decadal survey recommendations. This includes the ability to track not only GPS, but additional GNSS signals, including Galileo, CDMA GLONASS and Compass.
The TriG receiver is a NASA funded instrument. The hardware development is at Moog Broad Reach (formerly Broad Reach Engineering), the software development and complete end-to-end testing is at NASA/JPL. While the TriG instrument was in particular developed for the FormoSat-7/COSMIC-2 (Taiwan- US constellation), it is also intended to be used for a number of upcoming missions: DESDynI (Deformation, Ecosystem Structure, and Dynamics of Ice), ICESat-2 (Ice, Cloud and land Elevation Satellite-2), Jason-CS (Jason Continuity of Service), SWOT (Surface Water Ocean Topography), DSAC (Deep Space Atomic Clock), GRACE-FO, and other NASA/NOAA missions. 44) 45)
The next generation TriG science instrument provides the following observation capabilities:
• POD (Precise Orbit Determination)
• GNSS-RO (GNSS Radio Occultation)
• Neutral Atmosphere, Ionosphere, and Scintillation
• GNSS-R (GNSS Reflections)
TriG receives all L-band GNSS signals (GPS, Galileo, GLONASS, Compass) and DORIS.
- TriG has a separate science processor and a navigation processor (dual processor architecture)
- TriG possesses higher SNR compared to its previous generation receivers (particularly for RO and reflection)
- TriG is tolerant to a total ionizing dose of 40 kRad
- TriG offers a greater flexibility to accommodate future requirements.
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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: "Observation of the Earth and Its Environment: Survey of Missions and Sensors" (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates (firstname.lastname@example.org).