The Microwave Interferometer (MWI) will measure the minute variations in distance between the spacecraft.

Diagram of a GRACE Follow-On satellite
Image credit: Airbus-DS, GmbH

GRACE-FO is different from most Earth-observing satellite missions. Except for its atmospheric limb sounder GPS measurements, which will provide data to aid weather forecasting, it does not carry a suite of independent science instruments that point down at Earth’s surface to observe some part of the electromagnetic spectrum. Instead, the twin satellites act in unison as the main instrument, pointing their microwave and laser ranging instruments at each other to obtain data on distance changes between the two spacecraft that are used to generate monthly models of Earth’s gravity field.

While similar in most respects to their predecessor GRACE satellites, GRACE-FO incorporates design upgrades gleaned from 15 years of GRACE operations that will improve satellite performance, reliability and mission operations.


Built by Airbus Defence and Space in Friedrichshafen, Germany, under subcontract to JPL, the twin GRACE-FO satellites -- known respectively as GRACE-FO-A and GRACE-FO-B -- are identical in all respects except for transmit and receive frequencies. Each is 10 feet, 3 inches (3.123 meters) long, 2 feet, 6.7 inches (0.78 meters) high, 6 feet, 4.5 inches (1.943 meters) wide at bottom, 2 feet, 3.3 inches (0.69 meters) wide at top, and weighs 1,323.2 pounds (600.2 kilograms), including onboard propellant.


An electrical power subsystem generates, converts, conditions, regulates, distributes and stores primary electrical power in accordance with instrument and satellite bus user needs. Electrical energy is generated using gallium arsenide solar cell array panels that cover the top and sides of each satellite. Excess energy on each satellite is stored in a lithium-ion battery with a capacity of 78 amp hours. The system provides an average of 355 watts of electrical power on orbit and is manufactured by Airbus.


A thermal control subsystem keeps all spacecraft and science instrument temperatures within allowable limits. It does this using a combination of active and passive control elements. It consists of 64 independent thermistor-controlled heater circuits for in-flight temperature housekeeping, monitoring and heater control, as well as for on-ground verification testing. The thermal control subsystem is manufactured by Airbus.


The Radio Frequency Electronics Assembly allows the satellites to communicate with Earth via radio in the microwave S-band spectrum. It receives telecommands and transmits onboard telemetry and science data from the instruments to the ground. It consists of two redundant transceivers coupled to a transmit/receive helix antenna mounted on the end of a boom that is deployed from the bottom of the satellites after separation. Additional redundant transmit and receive patch antennas are mounted to the top of the satellites. Each satellite uses a separate set of S-band frequencies for transmission and reception.


The onboard data handling system provides the central processor and mass memory software resources for the spacecraft and management of the science and housekeeping data. It provides necessary input and output capabilities for the attitude and orbit control system, and power and thermal systems operations. In addition, it performs spacecraft health functions, including fault detection, isolation and recovery operations.


For the accelerometer to measure only non-gravitational forces, it is important that the spacecraft center of gravity be placed at the center of the proof-mass of the accelerometer. The mass-trim mechanism and associated mass-trim electronics serve this function. The six mass-trim mechanisms each consist of a mass moving on a spindle, with each pair providing center-of-gravity trim along one axis.

The mass trim mechanism will be completely operated from the ground. The mechanism is a rebuild from GRACE with slightly increased mass and trim ranges.


The satellite’s “attitude,” or orientation and orbit control, are controlled by a system consisting of sensors, actuators and software. The Attitude and Orbit Control System provides three-axis stabilized Earth-pointing attitude control during all mission modes and measures spacecraft rates and orbital position. It features numerous improvements from the GRACE design. The system consists of a GPS receiver, Star Tracker Assembly, coarse Earth and sun sensor, fluxgate magnetometer, inertial measurement unit, magnetic torquers and a cold gas propulsion system.

The GPS receivers are used as references to determine the precise location of the two satellites in orbit. The receivers continuously receive location information from the constellation of GPS satellites circling Earth. Each spacecraft has three GPS antennas. One antenna is used to collect navigation data, one collects the mission’s atmospheric occultation data, and the other is used for backup navigation. The GPS receivers were manufactured by JPL.

The Star Tracker Assembly enables fine attitude and orbit control of the satellites and precise transformation of science data into inertial references. It precisely determines each satellite’s orientation by tracking their relative position in reference to the stars. It consists of three star tracker camera heads and control electronics. The GRACE mission used two star tracker camera heads. The GRACE-FO design increases attitude data availability during Sun/Moon blinding and improves accuracy about all spacecraft axes.

The coarse Earth/Sun sensor provides coarse attitude determination during all mission phases.

The magnetometer provides coarse attitude based on the satellite’s position as determined by onboard GPS position and a model of Earth’s magnetic field.

An inertial measurement unit provides three-axis rate information. The satellites are “three-axis stabilized,” meaning that their orientation is fixed in relation to their momentary flight path, and they do not spin for stability.

Fine corrections of orientation can be adjusted using six 30-Amp-m2 magnetorquers, which help to minimize satellite fuel consumption over the mission lifetime.

The redundant cold gas propulsion subsystem uses small cold gas thrusters to position the twin spacecraft into their operational orbit and establish the satellite constellation. Once inserted into their operational orbit, very little acceleration is required to maintain the constellation. GRACE-FO will use 69 pounds (31.3 kilograms) of gaseous nitrogen as propellant. The subsystem features a set of twelve 10 millinewton thrusters mounted two on each of the six sides of the satellite, and two 40 millinewton orbit-control thrusters mounted on the rear-panel of the satellite.


All science instruments, fuel tanks and batteries and other satellite subsystems are mounted on a carbon-fiber reinforced plastic platform. This material, which has a very low coefficient of thermal expansion, provides the dimensional stability necessary for precise range change measurements between the two spacecraft.


As with GRACE, the key science instrument for GRACE-FO is the microwave tracking system, known on GRACE-FO as the Microwave Instrument (MWI). The MWI provides precise (1 micron, about the diameter of a blood cell or a small fraction of the width of a human hair) measurements of the distance changes between the two satellites -- and, in turn, fluctuations in Earth’s gravity -- by measuring microwave signals sent between the two satellites. Each satellite transmits signals to the other at two frequencies -- 24 gigahertz (K-band) and 32 gigahertz -- (Ka-band), allowing for ionospheric corrections.

The MWI on each satellite consists of a redundant pair of ultra-stable oscillators, a K/Ka-band ranging assembly and an instrument processing unit.


The ultra-stable oscillators serve as the frequency and clock reference for the GRACE-FO satellites. They are built by the Johns Hopkins University Applied Physics Laboratory in Baltimore, Maryland, and are based on the ultra-stable oscillators flown on NASA’s GRACE and Gravity Recovery and Interior Laboratory (GRAIL) missions.


The K/Ka-Band Ranging Assembly is the radio frequency front-end of the GRACE-FO microwave measurement system. It is comprised of a dual-band, dual-linearly polarized horn antenna, waveguide feeds and redundant Microwave Assembly K/Ka-Band transmitter/receivers. The ranging horn transmits and receives K-band (24 gigahertz) and Ka-band (32 gigahertz) carrier signals to and from the other GRACE-FO satellite. The antennas are nearly identical to those flown on GRACE, with a few modifications to the aperture cover and feed components. The MWAs up-convert the ultra-stable oscillator signal to K and Ka-Band for transmission, and down-convert the received K and Ka-Band signals to baseband frequencies of 670 kilohertz and 500 kilohertz. They are based on the original GRACE design, with minor improvements from the design used on GRAIL.


The Instrument Processing Unit (IPU) is the nerve center for the science instruments for the spacecraft. It provides the digital signal processing functions for the K and Ka band signals, as well as for the GPS signals. It also provides various timing references for the satellite. The IPU includes a Trig Navigation Processor, Trig GPS sampler front end, and GRAIL Radio/Frequency Unit. The GPS component provides navigation information, time tagging/correlation of data products and ancillary Earth limb occultation measurements. The IPU subsystem includes the primary, redundant and occultation GPS antennas. The IPU was manufactured by JPL.


The GRACE-FO satellites may speed up or slow down for reasons other than changes in Earth’s gravity field. These other forces acting on the satellites are measured using a science instrument called an accelerometer, mounted at the center of gravity of each satellite. This instrument allows scientists to distinguish between satellite motions due to gravity influences and those caused by other influences such as air drag in the atmosphere or thruster firings. The three-axis electrostatic accelerometers are similar to the SuperSTAR accelerometers flown on GRACE and were developed by ONERA, a French national research laboratory.


Each of the GRACE-FO satellites is equipped with a laser retro-reflector consisting of four corner cubes mounted in a small pyramid, which is located at the underside of the satellites. The laser retro-reflectors were contributed by GFZ and provide a means of tracking the GRACE-FO satellites from the ground for backup and orbit verification purposes. They will be tracked by the Satellite Laser Ranging (SLR) global ground station network of the International Laser Ranging Service (ILRS). Ground controllers can verify the satellite’s orbits by firing lasers upward toward the satellites, where the laser beam bounces back off of the reflector. The data will be valuable in evaluating and strengthening orbit and gravity field solutions.


The experimental Laser Ranging Interferometer (LRI) is a technology demonstration that uses laser interferometry instead of microwaves to measure fluctuations in the separation distance between the two GRACE-FO spacecraft. This is the same measurement made by the MWI, but the LRI offers the potential to improve the precision of range fluctuation measurements by a factor of at least 10, largely due to the laser wavelength being 10,000 times shorter than the microwave wavelength. These improvements will enable the satellites to detect gravitational differences at smaller scales. The LRI will demonstrate precision inter-spacecraft laser interferometry for future GRACE-like geodetic missions. GRACE-FO LRI data are for technology demonstration purposes only and will not be the mission’s data of record for use by the science community.

The LRI was developed jointly by the United States and Germany. JPL managed the development of the laser, laser frequency stabilization reference cavity, and interferometer readout and control electronics, and supported spacecraft integration. Germany provided the optical components of the LRI (optical bench assembly and electronics, triple mirror assembly and baffles) and supported spacecraft integration. The German contribution was managed by the Max Planck Institute for Gravitational Physics Albert Einstein Institute (AEI) in Hannover, with implementation by SpaceTech (STI) in Immenstaad.

The Laser Ranging Interferometer instrument.

The LRI instrument
Image credit: Albert Einstein Institute, Hannover, Germany

The components of the LRI include:

  • An Optical Bench Assembly, which routes, detects and points the laser optical beams. The Optical Bench Assembly was developed by STI. The steering mirror was developed by Airbus. The photoreceivers were developed by the German Aerospace Center (DLR) in Adlershof.
  • Optical Bench Electronics, which provide power to the steering mirror and photoreceiver and signal conditioning between the photoreceiver and laser ranging processor. The Optical Bench Electronics were developed by Apcon Aerospace and Defence in Neubiberg/Munich.
  • A Triple Mirror Assembly, which routes the beam around the MWI. The Triple Mirror Assembly was developed by STI and Hensoldt Optronics in Oberkochen, Germany.
  • Optical Baffles, which prevent obstruction of the laser beams and control scattered light effects in the interferometer. The baffles were developed by STI.
  • A Light Path Closure, which protects the LRI during spacecraft integration and covers the triple mirror assembly mirrors to avoid contamination from the spacecraft. The Light Path Closure was developed by STI.
  • A Laser Ranging Processor, which measures the phase of the laser interferometer signal from the photoreceiver as representative of fluctuations in the separation between the two orbiters, provides control of the laser frequency, and commands the steering mirror angle and implements the search to establish the optical link. The Laser Ranging Processor was developed by JPL.
  • The laser, which provides the light used for laser interferometry and emits approximately 25 milliwatts of light at 1064 nanometers. The laser is based on a commercial spaceflight unit developed by Tesat Corporation for inter-satellite laser telecommunications, and successfully flown on several projects, including the USAF Near Field Infrared Experiment (NFIRE) and German TerraSAR-X projects.
  • An Optical Cavity, which stabilizes the laser light wavelength. The Optical Cavity was built by Ball Aerospace in Boulder, Colorado. The phase-modulator for the cavity was produced by Photline, part of iXblue in Saint-Germain-en-Laye, France.
  • The LRI’s laser frequency stabilization and laser ranging processor are based on prototypes developed under NASA’s Instrument Incubator Program by Ball Aerospace and JPL.
  • Optical fibers for the LRI were produced by Diamond USA Inc. and Diamond SA in Losone, Switzerland.
  • Optical ground support equipment for the LRI was developed by the German Aerospace Center (DLR) Institute of Space Systems in Bremen, the Albert Einstein Institute (AEI) and JPL.