Minimize GPM

GPM (Global Precipitation Measurement) Mission

GPM is a cooperative international US/Japanese Earth science mission with the prime agencies of NASA and JAXA, respectively. GPM is a follow-on and expanded mission to TRMM (Tropical Rainfall Measuring Mission), launched Nov. 27, 1997 and still in operation in late 2012 (in Oct. 2005, NASA decided to continue the TRMM mission until at least 2009 and possibly until 2012). TRMM has been demonstrating the benefits of rain measurement from space. The overall objectives of the GPM mission are to observe global precipitation more frequently and more accurately than TRMM. GPM will build on the work of TRMM and extend the science from understanding tropical rainfall to using that understanding to improve climate, weather, and hydrological forecasts on a global basis. 1) 2) 3) 4) 5) 6) 7) 8) 9) 10) 11) 12) 13) 14) 15) 16) 17)

The top-level science objectives are:

• Climate change: Improve ongoing efforts to predict climate by providing near-global measurement of precipitation, its distribution, and physical processes. Providing this information is a key indicator of the global water cycle and its response to climate change

• Weather prediction: Improve the accuracy of weather and precipitation forecasts through more accurate measurement of rain rates and latent heating. These are key inputs needed by computer models to produce better weather predications

• Flood/fresh water resource prediction (water cycle): Provide more frequent and complete sampling of the Earth's precipitation. This will provide better prediction of flood hazards and management of life-sustaining activities dependent upon fresh water.

The goal is to: a) understand the horizontal and vertical structure of rainfall and its microphysical elements, and b) to achieve global coverage on rain rates with substantially improved sampling of the diurnal cycle.

The delivery of data products in near real-time (3-hour latency max). The bias error should be 1/2 that of TRMM (accuracy threshold) with a 25% precision threshold. The project calls for a precipitation measure of the 4-D structure of rates and drop size distribution at 5 km horizontal and 250 m vertical resolution.

The GPM definition team at NASA has identified four major elements that are necessary for the development of an effective and viable program - in support of the objectives:

1) A Core spacecraft that makes accurate rainfall measurements, collects information on cloud dynamics and rainfall processes, and serves a calibration reference for other instruments used within the GPM Program for taking rainfall measurements

2) A multi-satellite constellation, with each satellite equipped with a passive microwave radiometer that measures rainfall and other forms of precipitation over broad measurement swaths

3) A Ground Validation (GV) program that provides ground truth verification and measurement validation at various locations on the earth that are representative of precipitation events associated with different climates and geo-locations (e.g., tropical, oceanic)

4) A Precipitation Processing System (PPS) that collects and processes the measurement data obtained by the Core and the constellation of spacecraft, and that disseminates precipitation data products to the user community.

The baseline GPM configuration consists of a NASA core satellite to measure the precipitation structure and to provide a calibration standard for the constellation spacecraft, and a constellation of several microsatellites, consisting of a NASA spacecraft and those provided internationally. The objective is to obtain frequent precipitation measurements on a global basis (sufficient to resolve the diurnal cycle). In addition, the mission architecture makes use of ground calibration/validation sites with a broad array of precipitation-measuring instruments. A Global Precipitation Data Center will produce and distribute precipitation maps, application products, and climate quality global products. The success of GPM is very much dependent on international partnerships.


- GPM Implementation Phase Memorandum of Understanding signed on July 30, 2009
- NASA and JAXA follow a dual-gateway approach for GPM partnership building


A CNES/ISRO/NASA trilateral meeting was held in June 2009 to formalize Megha-Tropiques’ participation in GPM – formal agreement being developed


Brazil has a GPM constellation satellite in its 2005-2014 National Space Activities Plan – final draft of NASA/AEB joint study agreement in review


Confirmed MetOp MHS (Microwave Humidity Sounder) data can be redistributed to GPM partners by NASA and expressed interest in a formal GPM partnership


NASA is developing an Inter-Agency Agreement with NOAA

Table 1: GPM partnership development (Ref. 12) 18)

The US GPM project was funded by NASA in early 2007 which includes the Core Observatory with a planned launch in 2014.

The following spacecraft are members (or contributing data members) of the GPM constellation: 19) 20)

• GPM Core spacecraft, provided by NASA and launched by JAXA (orbit of 65º inclination and 400 km altitude) 21)

• GCOM-W1 (Global Change Observation Mission-Water/Wind) of JAXA with a launch on May 17, 2012. Data from the AMSR-2 (Advanced Microwave Scanning Radiometer-2) on GCOM-W1.

• DMSP (Defense Meteorological Satellite Program) of NOAA. The SSMIS (Special Sensor Microwave Imager Sounder) instrument series (F-16 since 2003, F-17 since 2006, F-18 since Oct. 2009, (and F-19 to -20 yet to be launched) will be a key constellation members in the GPM era. 22) 23)

• Megha Tropiques, a joint French (CNES) and Indian (ISRO) low inclination (20º), tropical rainfall measurement satellite (launch Oct. 12, 2011). Data from the MADRAS (Multi-Frequency Microwave Scanning Radiometer). 24)

• MetOp-A of EUMETSAT (launch Oct. 19, 2006). The MetOp-B was launched on Sept. 17, 2012. Both spacecraft operate in a co-planar sun-synchronous orbit, phased 174º.

• NOAA-19 (NOAA-N') with a launch on Feb. 6, 2009 and operational on June 2, 2009. Data from the MHS (Microwave Humidity Sounder). 25)

• Suomi-NPP (NPOESS Preparatory Project) of NASA with a launch on Oct. 28, 2011.

• JPSSS (Joint Polar Satellite System) satellites of NOAA/NASA, each equipped with CMIS (Conical-scanning Microwave Imager/Sounder). A launch of the JPSS-1 spacecraft is planned for 2017. 26)


Figure 1: Overview of the GPM mission architecture (image credit: NASA)


Figure 2: GPM mission architecture (image credit: NASA) 27)



The GPM Core Spacecraft

On July 30, 2009, JAXA and NASA signed a MOU (Memorandum of Understanding) on development and operation activities for the GPM mission. The agreement was signed at the KSC (Kennedy Space Center) in Cape Canaveral, Florida. 28) 29)

The joint mission agreements call for:

• NASA to provide the core (and a companion spacecraft), the companion spacecraft launch, the US ground system, and the Precipitation Data Center. Regarding instrumentation, NASA will contribute a conical-scanning, polarization-sensitive, multi-frequency radiometer, GMI, for both the core and a companion satellite. NASA will also contribute to and participate in algorithm development and data validation activities.

• JAXA to provide the DPR instrument for the core spacecraft, the launch of the core spacecraft, and a data stream from the JAXA GCOM-B1 spacecraft to NASA.

The NASA GPM core spacecraft is of TRMM heritage. The Core spacecraft bus is being developed in-house at NASA/GSFC. The mission CDR (Critical Design Review) took place Dec. 14-17 2009.

The spacecraft bus features an aluminum and composite structure, the bus is modular and has fully redundant avionics consistent with its Class B reliability designation. A steerable high-gain antenna on a dual-hinged boom provides nearly continuous data downlink of science data from the GMI and DPR instruments via the TDRSS (Tracking and Data Relay Satellite System) in MA (Multiple Access) mode at ~230 kbit/s and SA (Single Access) mode at ~2300 kbit/s.

The GPM spacecraft has a mass of about 3850 kg and a power of ~1.95 kW. The design life is 3 years with a goal of at least 5 years of operation. The spacecraft bus uses 12 hydrazine thrusters (4 forward, 8 aft) for the regular orbit maintenance required for its 407 km altitude 65º inclination orbit. The full capacity propulsion tank is providing a fuel margin beyond the 5 year consumable requirement. Also, the increased battery capacity to 200 Ah ensures a mission life beyond the 5 year consumable requirement. 30)


Figure 3: Artist's rendition of the deployed GPM spacecraft in orbit (image credit: NASA)


Figure 4: Overview of the GPM Core spacecraft, bottom view shows instrument locations (image credit: NASA)

ADCS (Attitude Determination and Control Subsystem): A suite of state-of-the-art sensors are used to determine the 3D attitude in space to ensure proper attitude control and Earth-pointing for instrument data acquisition. GPM incorporates star trackers and an IRU (Inertial Reference Unit) as main attitude data sources. Two wide-angle star trackers are used to acquire imagery of the sky that is analyzed by a software algorithm that compares the acquired star pattern with a catalog to precisely determine the spacecraft's orientation in space. The star trackers are connected to the 1553 data bus to relay precise attitude data to the vehicle's control system. 31)

The IRU used on the GPM spacecraft is referred to as SIRU ( Scalable IRU) reference system provided by Northrop Grumman. The system uses gyroscopes to precisely measure changes in rotational attitude on all three axes to provide accurate attitude and rate data to the spacecraft control system. The IRU is internally redundant.


Figure 5: Photo of the SIRU reference system (image credit: Northrop Grumman)

Two MSS (Medium Sun Sensors) are also part of GPM's guidance suite. These sensors have a smaller field of view than the coarse sun sensors and provide higher accuracy in their measurements. The two units have a FOV (Field of View) of ~17.5º and measure the position of the sun with an accuracy of 2º. The MSS data is used in initial attitude hold mode and during spacecraft safe modes.

Two magnetometer units are installed on GPM to determine the spacecraft attitude relative to Earth's magnetic field. Three magnetic torque rods with redundant coils are used to create angular momentum by running a current through coils in the presence of Earth's magnetic field. The torquers are regulated by computers that control the current that is passing through the coils in order to control the force generated on each axis. The magnetic torquers are used during momentum dumps.

A GPS receiver unit aboard the GPM spacecraft determines the spacecraft position, altitude and velocity for navigation, antenna pointing and science data processing.

Attitude control is primarily provided by a RWA (Reaction Wheel Assembly). The wheels are spun by electric motors at variable speed that is changed when making attitude maneuvers. Each RWA has a mass of 10 kg, the wheels spin as fast as 6,000 rpm. The thrusters are used for periodic angular momentum desaturation - slowing down the reaction wheels and countering the resulting force with the thrusters so that the wheels can then be accelerated during standard attitude operations.

C&DHS (Command and Data Handling Subsystem): The C&DHS is in charge of command reception and execution, payload system operations, housekeeping operations and spacecraft control. The C&DHS uses key-components developed for the LRO (Lunar Reconnaissance Orbiter) mission, based on the VxWorks operating system for realtime operations. GPM uses a Spacewire, 1553 data bus and an analog RS-422 system for data transfer within the spacecraft. Housekeeping data and science data is stored in a solid-state recorder before downlink or is processed and downlinked in realtime.

GPM is equipped with a PowerPC RAD750 microprocessor that features a single-card computer manufactured by BAE Systems of Manassas, VA. The processor can endure radiation doses that are a million times more extreme than what is considered fatal to humans. The RAD750 CPU itself can tolerate 200-1,000 krad. Also, the RAD750 will not suffer more than one event requiring interventions from Earth over a 15-year period.

The RAD750 card is designed to accommodate all those single event effects and survive them. The ultimate goal is one upset is allowed in 15 years. An upset means an intervention from Earth — one 'blue screen of death' in 15 years.

RAD-750 was released in 2001 and made its first launch in 2005 aboard the Deep Impact spacecraft. The CPU has 10.4 million transistors. The RAD750 processors operate at up to 200 MHz, processing at 400 MIPS. The CPU has an L1 cache memory of 2 x 32 kB (instruction + data) - to improve performance, multiple 1MB L2 cache modules can be implemented depending on mission requirements.


Figure 6: Photo of the PowerPC RAD750 microprocessor card (image credit: BAE Systems)

EPS (Electrical Power Subsystem): GPM is equipped with two solar arrays with four panels on each array. The arrays are attached to booms that interface with the second panel in order to individually tilt the arrays to track the sun and optimize power generation. Each panel features 800 to 1,200 individual Gallium-Arsenide solar cells (total solar cell area is 26.5 m2). CSS (Coarse Sun Sensors) are installed on the arrays to provide guidance for array tilting to achieve optimal illumination. Generally, CSS are imagers with a field of view of 85 º to determine the sun direction with an error of 10º.

Power from the arrays is passed to a suite of power controllers facilitated in a single Power System Electronics Box to distribute electrical power to the different subsystems and payloads that use a regulated, redundant power bus at 28 Volts with an operational range of 23 to 35 V. A dedicated controller is used to regulate the state of charge of a single 200 Ah Li-ion battery. Overall, GPM generates 1.95 kW at EOL (End of Life).


Figure 7: Photo of a deployed GPM solar array at NASA/GSFC (image credit: NASA)

TCS (Thermal Control Subsystem): GPM uses a combination of active and passive thermal control. Passive thermal control is accomplished by the use of thermal covers, coatings and multilayer insulation that prevents sunlight from excessively heating the satellite and heat from dissipating into space in darkness. The outer layer of the multilayer insulation is used to minimize heat loss and is also designed to protect the spacecraft against corrosion by atomic oxygen and electrical charging. The outer MLI layer is made of Germanium Black Kapton. GPM uses blanket tents and form-fitting blankets to protect most of its non-radiating surfaces.

Active thermal control uses heat rejection systems as well as heaters and temperature sensors to keep the satellite at an operating temperature. Because all electronics of the satellite generate heat, GPM has to be outfitted with a heat rejection system. Most electronic components of the spacecraft reject heat through their baseplates that are mounted on structural surfaces using a thermal interface material to improve the heat transfer. The heat is then transported via constant conductance heat pipes - the Avionics Module uses one U-shaped and one S-shaped heat pipe while the Power System Electronics Box has two L-shaped heat pipes. The battery assembly has four dedicated heat pipes.

Heat is rejected via radiators installed on the +Y side of the spacecraft that never faces the sun during nominal mission operations. These high-emittance radiators include an avionics radiator, a pocketed battery radiator, a Lower Bus Structure Radiator, dedicated radiators for the two solar array drive assemblies and separate radiators for the RF communications system.

RCS (Reaction Control Subsystem): GPM uses a chemical propulsion system for attitude control, reaction wheels momentum dumps and to maintain its Low Earth Orbit. A total of 12 thrusters are installed on the spacecraft. Eight of those 12 thrusters are installed on the aft section of the satellite while the remaining four thrusters are installed on the forward-facing section. Four thrusters have 90º nozzles, the other eight have straight nozzles. All thrusters are used for attitude control and momentum dumps while orbital maneuvers only use the four forward thrusters that are facing the same direction.

The RCS utilizes high-purity hydrazine fuel that is stored in a single Composite Overwrap Pressure Vessel tank. It uses an outer shell made of graphite composite, a tank skirt consisting of graphite composite with metallic inserts and an Aluminum 6061 alloy liner. The tank is filled with 545 kg of hydrazine at launch pressurized to 27.6 bar. The tank has been designed to operate at pressures of up to 34.5 bar and a burst pressure of 55.2 bar. The minimum flight pressure is 6.8 bar. The tank has been built to maintain a temperature of 2 to 50 ºC with a ten-year minimum storage life of the hydrazine inside without performance degradation. Tank pressurization is accomplished using 6.2 kg of high-pressure nitrogen.

The thrusters generate thrust by the catalytic decomposition of hydrazine propellant using heated platinum/palladium catalyst beds. Operated in blowdown mode, the thrusters provide a maximum thrust of 44.5 N at a feed pressure of 27.6 bar and 13.3 N at 6.8 bar feed pressure. Operation of the thrusters is accomplished in pulse mode for attitude control and steady-state mode for orbital maneuvers. The minimum pulse duration of the thrusters is <50 ms supplying a repeatable impulse bit. Engine thrust is calculated to within 5% for any given feed pressure.

For attitude control during ΔV burns, the thrusters are operated at duty cycles of 33, 67 or 83% followed by steady-state burns of several seconds while momentum unloading requires duty cycles of 17, 33 or 67% for periods of several seconds separated by non-firing periods of up to several minutes. Steady-state burn time of the Orbit Correction Engines is 35 s at the maximum feed pressure at the start of the mission and 70 s at the end of the mission when the minimum feed pressure has been reached.

Overall, the thrusters consume 0.06 kg/s of hydrazine during a 4-thruster burn without attitude control and 0.12 kg/s for a maneuver with attitude control. The expected propellant consumption at the start of the mission is 1.4 kg/maneuver that increases to 1.9 kg at the end of the mission. These drag makeup maneuvers will be performed every 12.4 days (on average).

Spacecraft bus (GSFC in-house development)

Aluminum and composite structure, modular bus with redundant avionics


Steerable high-gain antenna on dual-hinged boom

Power generation

- Solar arrays track the sun
- 200 Ah Li-ion battery
- 1.95 kW of power at EOL

RCS (Reaction Control Subsystem)

12 thrusters (4 forward, 8 aft), monopropellant thrusters, MR-106L, 22 N, of Aerojet Rocketdyne for attitude control

Size of deployed spacecraft

13 m x 6.5 m x 5 m

Spacecraft mass

3850 kg

Spacecraft design life

3 years with 5 years of consumables (e.g. propellant)

Table 2: Overview of the GPM Core spacecraft parameters

RF communications subsystem: The GPM Core Observatory uses an S-band communications system for data exchange with the ground. The GPM HGA (High Gain Antenna) system features a deployable boom that facilitates the high gain antenna dish of the spacecraft which is installed on a two-axis gimbal mechanism to orient the antenna for communications with NASA's TDRSS (Tracking and Data Relay Satellite System).

Realtime payload and housekeeping data is downlinked via the TDRSS MA (Multi-Access) service that allows a TDRS satellite to relay data from several lower data rate users. Stored science data is downlinked via the SA (Single-Access) service of TDRSS that uses dedicated antennas on the TDRS satellites to achieve high downlink data rates. The data rate is ~230 kbit/s in MA mode and ~2.3 Mbit/s in SA mode.

TDRSS data is downlinked to White Sands, New Mexico from where the GPM data is transferred to the MOC (Mission Operations Center) that then distributes the acquired instrument data for processing and publication. For realtime downlink, instrument data will be available within 15 minutes to achieve a near realtime coverage. Typical data latencies from measurement to user availability are expected to be in the order of 15 minutes. A primary application for the short latency GMI data is for integration into a NRT (Near-Real-Time) global rainfall map created from measurements by all the GPM constellation radiometric sensors and with overall rain map data latency less than 3 hours.

The command uplink is also accomplished via the HGA system in nominal mission modes. GPM is equipped with omnidirectional S-band antennas that are used to communicate with ground stations. These antennas are used for telemetry downlink and command uplink and in case of spacecraft safe modes.


Project status:

• In early March 2012, the GMI (GPM Microwave Imager), built by BATC (Ball Aerospace and Technologies Corp.) of Boulder, CO, arrived at NASA/GSFC. 32) 33)

• In March, 2012, the DPR (Dual-frequency Precipitation Radar) of JAXA was delivered to NASA/GSFC. Following installation of the DPR on the GPM Core Spacecraft, NASA will perform the spacecraft system testing at GSFC. 34)

• Integration of the DPR onto the GPM spacecraft was successfully completed in May 2012. 35) 36)

• In October 2012, the GPM spacecraft went through its first complete CPT (Comprehensive Performance Test), beginning on Oct. 4, 2012 at NASA/GSFC. The testing ran twenty-four hours, seven days a week and lasted ten days as the entire spacecraft was put through its paces. 37)


Figure 8: Photo of the DPR instrument integrated onto the GPM Core Observatory (image credit: NASA) 38)

• The GPM core spacecraft completed the first comprehensive performance test in October 2012 and thermal vacuum test in January 2013.

• In May of 2013, the EMI/EMC tests were completed at NASA/GSFC ( Goddard Space Flight Center). 39) 40)

• The GPM Core satellite successfully completed vibration testing in July 2013, at NASA/GSFC, Greenbelt, MD. The tests ensure that the spacecraft can withstand the vibrations caused by the JAXA H-IIA rocket during satellite’s launch early in 2014. Sitting on a specialized mobile platform, the GPM spacecraft was abruptly moved back and forth in each of its three spatial orientations. 41)


Figure 9: GPM attached to the shaker table for horizontal vibration testing (image credit: NASA)

• On Nov. 23, 2013, a USAF C-5 transport aircraft carrying the GPM (Global Precipitation Measurement) Core Observatory landed at Kitakyushu Airport in Japan. From Kitakyushu Airport, the spacecraft was loaded onto a barge heading to JAXA's Tanegashima Space Center on Tanegashima Island in southern Japan, where it will be prepared for launch in 2014 on an H-IIA rocket. 42)


Figure 10: Photo of the C-5 transport aircraft, carrying the GPM Observatory, landing at Kitakyushu Airport (image credit: JAXA, NASA)

• On Dec. 26, 2013, NASA and JAXA announced the launch date for GPM. They selected Feb. 27, 2014 as the launch date and launch window for a Japanese H-IIA rocket carrying the Global Precipitation Measurement (GPM) Core Observatory satellite from JAXA's Tanegashima Space Center. 43)


Figure 11: The GPM Core Observatory in the clean room at Tanegashima Space Center, Japan (image credit: JAXA, NASA)


Launch: The GPM Core Observatory was launched on February 27, 2014 (at 18:37:00 UTC) on the H-IIA No 23 launch vehicle of JAXA from the Tanegashima Space Center, Japan. JAXA sponsored the launch on the H-IIA vehicle from the Tanegashima Space Center, Japan with MHI (Mitsubishi Heavy Industries, Ltd.) as the service provider. 44) 45) 46) 47) 48) 49) 50)

MHI is Japan's largest aerospace and defence contractor, MHI manufactures the H-II family of rockets for JAXA and has been responsible for conducting H-IIA launches since 2007 – taking over launches of the more powerful H-IIB as well in 2013.

The secondary Japanese payloads manifested by JAXA on the GPM Core mission were: 51)

• ShindaiSat (Shinshu University Satellite), a microsatellite (35 kg) to demonstrate LED light as an optical communications link.

• The STARS-2 (Space Tethered Autonomous Robotic Satellite-2) nanosatellite technology mission of Kagawa University, Takamatsu, Kagawa, Japan

• TeikyoSat-3, a bioscience microsatellite (~20 kg) of Teikyo University

• ITF-1 (Imagine The Future-1), a 1U CubeSat of the University of Tsukuba, Tsukuba, Japan.

• OPUSat (Osaka Prefecture University Satellite), a 1U CubeSat

• INVADER (INteractiVe satellite for Art and Design Experimental Research) of Tama Art University, a 1U CubeSat

• KSat-2 (Kagoshima University Satellite-2), a CubeSat mission with a mass of ~ 1.5 kg. 52)

Orbit of the core satellite: Non-sun-synchronous circular orbit, altitude = 407 km, inclination = 65º. The orbit of the core spacecraft cuts across the orbits of the constellation spacecraft, sample the latitudes where nearly all precipitation occurs, and sample different times of day.


Figure 12: The graphic compares the area covered by three TRMM orbits (yellow) versus three orbits of the GPM Core Observatory (blue), image credit: NASA) 53)

Orbit of constellation satellites: Sun-synchronous (polar) circular orbit, altitude 635 km. The constellation is actually a collection of spacecraft, most with missions independent of GPM (like the DMSP S/C series). Many of the spacecraft are sun-synchronous, but their altitudes and orbital periods are different. The DMSP spacecraft orbit at 833 km, while GCOM-W will orbit at 802 km. The different orbital periods cause the ground tracks to move with respect to each other, oscillating between overlapping coverage and missed coverage.


Figure 13: Schematic view of the observation geometries with the GPM CORE instruments (image credit: NASA)



Mission status:

• March 6, 2014: GPM is performing normally. The initial checkout of the GMI instrument and the spacecraft showed both are performing as expected, and the GMI instrument continues to collect science data on rain and snowfall (Ref. 54).

• March 5, 2014: The GPM Core Observatory is performing normally. The GMI continues in science mode, and GMI data is being sent to the PPS (Precipitation Processing System) at NASA/GSFC in Greenbelt, MD. Using the initial data, the instrument team has verified that GMI is working well on-orbit. The GPM spacecraft will have a 60 day on-orbit check out period (commissioning phase) to ensure the healthy operation of the spacecraft and instruments. Precipitation data will be released from the PPS no later than 6 months post-launch, after the science teams verify their accuracy. 54)

• March 4, 2014: GMI's electronics have been turned on and all seven launch restraints released, deploying the instrument. GMI (GPM Microwave Imager) began spinning today collecting the first science data of the mission. The GMI will complete several additional check-out procedures during the commissioning process. 55)

• March 1, 2014: Following activation and warm up of the GMI electronic systems, the team at NASA’s Goddard Space Flight Center in Greenbelt, Md., deployed the main reflector of the U.S. science instrument for the GPM Core Observatory (Ref. 54).

- A significant step was also achieved today in the activation of the science instrument provided by JAXA with the turning on of the controller for the DPR (Dual-Frequency Precipitation Radar).

- GPM flight controllers at NASA/GSFC began using the satellite’s High Gain Antenna system for high-rate data rate transmissions through NASA’s orbiting fleet of TDRS (Tracking Data Relay Satellites).

• Feb. 28, 2014: The GPM Core Observatory is performing normally. The GPS system has been switched on. This tells the satellite the time and its location with respect to the Earth's surface. The team is readying the spacecraft to use its High Gain Antenna for high data-rate communication through the Tracking and Data Relay Satellite System.

• After the release of the GPM Core Satellite, the second stage performed attitude maneuvers and slightly changed its orbit for the deployment of the seven secondary payloads that include small spacecraft and CubeSats dedicated to scientific missions, technical demonstrations and outreach projects (Ref. 56).

Launch event

Time (minutes:seconds)

Altitude (km)

Inertial speed (km/s)





Solid rocket booster burnout




Solid rocket booster jettison (thrust strut cutoff)




Payload fairing jettison




1st stage engine (main engine) cutoff (MECO)




1st and 2nd stages separation




2nd stage ignition (SEIG)




2nd stage engine cutoff (SECO)




GPM-Core separation




ShindaiSat cubesat separation




STARS-2 CubeSat separation




TeikyoSat-3 microsatellite separation




ITF-1 CubeSat separation




OPUSAT CubeSat separation




INVADER CubeSat separation




KSat-2 CubeSat separation




Table 3: Launch sequence of GPM mission and secondary payloads

• The GPM spacecraft separated from the rocket ~16 minutes after launch, at an altitude of 398 km. Following spacecraft separation, GPM initiated a pre-programmed sequence to establish a stable three-axis orientation in attitude safe mode and acquire communications with ground stations. GPM's signal was received - confirming that the spacecraft was alive and well after its ride into orbit. The solar arrays deployed 10 minutes after spacecraft separation, to power the spacecraft. 56)



Sensor complement of the core mission: (DPR, GMI)

The GPM Core Observatory measurement capabilities are provided by the two main instruments the active microwave DPR and the passive microwave GMI. 57) 58) 59) 60)

The JAXA-supplied DPR, composed of Ka and Ku band radar subsystems, will provide:

• Increased sensitivity (~12 dBZ) for light rain and snow detection relative to TRMM

• Better precipitation measurement accuracy with differential attenuation correction, and

• Detailed precipitation microphysical information of DSD (Drop Size Distribution), mean mass diameter, particle number density) and identification of liquid, ice, and mixed-phase regions.

The multi-frequency (10-183 GHz) GMI conical scan microwave radiometer will provide:

• Higher spatial resolution (IFOV: 6-26 km) than its TRMM Microwave Imager (TMI) predecessor

• Improved light rain & snow detection

• Improved signals of solid precipitation over land (especially over snow- covered surfaces), and

• Four-point calibration to serve as a radiometric reference for constellation radiometers.

The resulting combined radar-radiometer rainfall retrievals utilizing data from the two instruments will together provide greater constraints on possible solutions to improve retrieval accuracy. An observation-based a-priori cloud database will be used for constellation radiometer retrievals.


DPR (Dual-frequency Precipitation Radar):

The DPR instrument, of PR heritage flown on TRMM, is being designed and developed in a collaborative effort between JAXA (Japan Aerospace Exploration Agency) and NICT (National Institute of Information and Communications Technology), Tokyo. (the industrial partner is NEC Toshiba Ltd., Tokyo). The objective is to extend the instrument capability of TRMM in such a way to fully address the key science questions from microphysical to climate time scales. DPR will provide the accurate amount of precipitation including snowfall over both ocean and land. The DPR data will also be used to calibrate the MWRs (Microwave Radiometers) in the GPM constellation. The DPR package will provide a global database of precipitation characteristics, such as storm heights, freezing levels, DSDs (Drop Size Distributions), the mean structure of precipitation profiles, and so on. 61) 62) 63) 64) 65) 66) 67) 68) 69) 70) 71) 72) 73) 74) 75)


Figure 14: Goals and objectives of JAXA's GPM/DPR project (image credit: JAXA) 76)


Figure 15: The GPM Core spacecraft with the DPR on-orbit configuration (image credit: JAXA)

The DPR instrument is comprised of two, essentially independent radars. One radar operates in the Ku-band at 13.6 GHz, it is referred to as PR-U, also known as KuPR (Ku-band Precipitation Radar). The other radar operates in the Ka-band at 35.55 GHz, it is referred to as PR-A (also known as KaPR (Ka-band Precipitation Radar). By measuring the reflectivities of rain at two widely different radar frequencies, it is possible to infer information regarding rain rate, cloud type and its three-dimensional structure, and drop-size distribution. Both radars have almost the same design as the PR instrument on TRMM. The specific objectives of DPR are to:

• To provide the three-dimensional precipitation structure including snowfall over both ocean and land

• To improve the sensitivity and accuracy of precipitation measurement

• To calibrate the estimated precipitation amount by MWRs (Microwave Radiometers) and MWSs (Microwave Sounders) on the constellation satellites.

Each radar has 128 slot array antennas, transmitters (Solid State Power Amplifier: SSPA), receivers (Low Noise Amplifier: LNA), Phase Shifters (PHS), and so on. The FCIF (Frequency Converter Intermediate Frequency) and the SCDP (System Control Data Processing) of both instruments, KuPR and KaPR, have almost the same designs. To make the structures lighter, one SCDP installed on KuPR is used to control both KuPR and KaPR. The other SCDP, which is installed on KaPR, is just for redundancy. There are two major differences from the TRMM/PR: 77)

- One major difference is that the T/R module groups one SSPA, LNA, and PHS together, and one T/R unit consists of 8 T/R modules. In each radar, there are 16 T/R units.

- Another one is the design change of Divider/Combiner (DIV/COMB), Circulator (CIR) and Hybrid (HYB) to eliminate a single failure point in the RF line.

The analysis results using subsystems and components design and parameters reviewed in the critical design of the DPR have achieved the required technical performance of frequency, range resolution, spatial resolution, swath width, minimum detectable rainfall rate, beam matching accuracy, observable range, dynamic range, received power accuracy, and so on.

Figure 16 illustrates the dual-frequency measurement concept of precipitation in the various detectable dynamic ranges. Predominantly, the KaPR will detect snow and light rain, while the KuPR will detect the heavy rain regime. Both instruments have a common effective dynamic range to provide the DSD (Drop Size Distribution) information and more accurate rainfall estimates, implemented by the dual-frequency algorithm. The dual-frequency algorithm employs the difference in rain attenuation from the matched beam data observed by KuPR and KaPR.

The data obtained from DPR will contribute to a global database of precipitation characteristics, to derive such such parameters as precipitation heights, freezing levels, DSDs, the mean structure of precipitation profiles and so on. This database must also serve to improve the MWR and MWS algorithms.


Figure 16: DPR concept of dual-frequency measurement of precipitation (image credit: JAXA, NICT) 78)

Legend to Figure 16: Left: Vertical precipitation structure; Right: Relations between radar reflectivity and height for KuPR and KaPR.


Figure 17: The DPR antenna scanning concept (image credit: JAXA)

Each radar uses a phased array, slotted wave guide antenna. Both radars of the DPR can be electronically steered up to ±17º to either side of the spacecraft nadir, providing a 245 km measurement swath. The KaPR also has a selectable high sensitivity mode which provides an interlacing scan with a swath width of 120 km; this high sensitivity mode will aid in the measurement of light rain and snow. The two phased array antennas will be aligned so that identically sized coincident measurement footprints of 4.5 km diameter can be taken.

The pulse repetition frequency (PRF) of both KuPR and KaPR will vary according to the satellite altitude variation as a function of latitude. This variable PRF technique improves the signal to noise ratio because of the larger sampling numbers it offers. The KuPR has a swath width of about 245 km (comprised of 49 footprints each 5 km in width), which is the same as TRMM PR, while the KaPR observes a swath width of about 120 km. In the overlapping scan area, measurements will be performed synchronously to match the two beams of KuPR and KaPR. While the KuPR observes the outer swath area, the KaPR can measure snow and light rain in the interlacing scan area in a high-sensitivity mode with a double pulse width. Another reason for the narrow swath width of KaPR is that the sidelobe clutter contamination in larger scan angles will hinder measuring shallow snow clouds.


DPR-KuPR (Ku-band)

DPR-KaPR (Ka-band)

Antenna type

Active Phased Array

Active Phased Array

No of antenna elements


128 (planar array, slotted wave)

Tx peak power

> 1013.5 W

> 146.5 W


13.597 & 13.603 GHz

35.547 & 35.553 GHz


< 18 dBZ (0.5 mm/hr)

< 12 dBZ (0.2 mm/hr)

Average sampling number

> 64

> 64

Z min

< - 110 dBm

< -109 dBm (250 m res.)
< -112 dBm (500 m res.)

Minimum detectable rainfall rate

0.5 mm/hr

0.2 mm/hr (at 500 m resolution)

Pulse width

1.67 μs (2)

1.67 μs (2)

Swath width

245 km (49 footprints at 5 km)

125 km (24 footprints at 5 km)

Scan angle (FOV)



Range resolution

250 m (1.67 µs)

250 m / 500 m (1.67 / 3.34 µs)

Horizontal resolution

5.2 km (at nadir)

5.2 km (at nadir)

Beam width



Beam matching accuracy

< 1000 m

Variable PRF (Pulse Repetition Frequency)

VPRF (2900 ~ 4500 Hz)

VPRF (2900 ~ 4500 Hz)

Data rate

< 108.5 kbit/s

< 81.5 kbit/s

Instrument mass

< 472 kg

< 336 kg

Instrument size (antenna)

2.5 m x 2.4 m x 0.6 m

1.2 m x 1.44 m x 0.7 m

Power consumption

< 446 W

< 344 W

Table 4: Key parameters of DPR (system level requirements)

The two radars are designed to provide temporally matching footprints with the same spatial size and scan pattern. Both radar antennas are carefully aligned to ensure co-alignment of the beams.


Figure 18: Illustration of the EM (Engineering Model) of the KaPR instrument (image credit: JAXA, NICT, Ref. 61)


Figure 19: BBM (Breadboard Model) of KuPR (left) and KaPR (right), 1 T/R unit respectively (image credit: NICT)

DPR status 2010 (Ref. 61) JAXA Dual-frequency Precipitation Radar (DPR) in Phase-C development

- JAXA DPR CDR (Critical Design Review) completed in August, 2009

- DPR engineering model tests for design verification completed

- NASA/JAXA DPR Interface Preliminary Design Review completed in October 2009 and GMI-DPR interference test completed in Dec. 2009 in Japan.

- Currently (2010) manufacturing and testing all of the DPR PFM (Proto-Flight Model) components

- Delivery of DPR simulator to NASA GSFC in the fall of 2010.

- The DPR protoflight test is underway, will be completed in October 2011. 79)


DPR operation modes:

DPR has 7 operation modes (Ref. 78):

1) Observation Mode - is the normal operation mode where the KuPR and the KaPR perform normal rain echo measurements with the ±17º scanning for the KuPR and with the ±8.5º scanning for the KaPR. System noise, surface return, and mirror image data are also collected. In the contingency case that the signal between the KuPR and the KaPR is lost, the SCDP (System Control Data Processing) in the KuPR and the SCDP in the KaPR are operated simultaneously so that both radars perform the observation independently. But beam matching is impossible in this case.

2) Internal Calibration Mode - is used to calibrate FCIF (Frequency Converter Intermediate Frequency) and SCDP. During this mode, RF radiation does not occur.

3) External Calibration Mode - is used for the end to end calibration of the DPR using the ARC (Active Radar Calibrator) on the ground.

4) Analysis Mode - provides the LNAs (Low Noise Amplifiers) and the SSPAs (Solid State Power Amplifiers) status data. During this mode, science observations do not occur.

5) Health-Check Mode - is for checking the ROMs (Random Access Memory) and the ROMs (Read Only Memory) used in the SCDP. During this mode, science observations and RF transmissions do not occur.

6) Standby Mode - is used for re-loading the phase code and the VPRF (Variable Pulse Repetition Frequency) data, changing the timing offset between the KuPR and the KaPR, and re-writing the onboard software in the SCDP. During this mode, science observations and RF transmissions will not occur. This mode is also used when the GPM observatory is in the sun point mode due to the minor spacecraft failure. In this case, only the SCDP continues to be powered on and all other components are powered off.

7) Safety Mode - is the mode that the DPR is off except for the survival heater. This mode is used when the GPM observatory is in the period from launch to early stages of the on-orbit checkout period and when the GPM observatory is in the sun point mode due to the power load-shedding fault.


DPR onboard calibration and validation:

There are two types of calibrations: external and internal calibration. 80)

• External calibration in the initial checkout period: The agreement of the observation volume of the two radars (KuPR and KaPR) must be confirmed by external calibration; that is, the radar beam direction corresponding form the comparison of the antenna patterns. It is necessary to make the assumption of where in the footprint the RC (Radar Calibrator) existed. To proofread the strength of the transmitting and receiving signal, it is assumed by the method of the repetition of less beam direction (from five directions by about ten directions) scanning two or more times. But in the observation, usually 49 direction beams are scanned in one scanning sweep (0.7 seconds).

• External calibration for transmitting on orbit: The RxRC (Receive Radar Calibrator) measures the transmitting power form the DPR, and assumes in which position the RxRC is set up in the footprint when the DPR transmit beam scan forming is changed. The scan beam form change is available.

The TxRC (Transmit Radar Calibrator) transmits to the DPR continuous wave (CW) by f1 and f2. The DPR receives the CW of TxRC; hence, the DPR reception characteristics can be checked.


Figure 20: Illustration of the external and internal calibration scheme (image credit: NICT)

A prototype RC antenna was developed which combines the two frequencies of the Ku- and the Ka-band.



Figure 21: Photo of DPR KuPR (left) and KaPR (right) at GSFC (image credit: NASA, Ref. 27)


GMI (GPM Microwave Imager):

The NASA GMI instrument, of SSM/I, TMI, and SSMIS heritage, is a conical-scanning, polarization-sensitive, multi-frequency passive radiometer for rainfall measurement. GMI will be used to make calibrated, radiometric measurements from space at multiple microwave frequencies and polarizations. In addition, radiometric measurements from GMI and radar measurements from the DPR will be used together to develop a retrieval transfer standard for the purpose of calibrating precipitation retrieval algorithms. This calibration standard will establish a reference against which other retrieval algorithms using only microwave radiometers (and without the benefit of the DPR) on other satellites in the GPM constellation will be compared. 81)

In March 2005, NASA awarded a contract to BATC (Ball Aerospace and Technologies Corporation) to design and built GMI. A successful preliminary design review took place in Nov. 2006. In June 2009, the instrument has completed the Critical Design Review phase of the program. The delivery of the two flight units is planned for 2012 and 2013, respectively.

The conical scan geometry of GMI is shown in Figure 23. The off-nadir-angle defining the cone swept out by the GMI is set at 48.5º which represents an Earth incidence angle of 52.8º (identical to that of TMI on TRMM). The offset parabolic reflector rotates about the vertical axis of the instrument with a rate of 32 rpm; during each revolution the Earth-viewing scan sector is about 140º centered along the S/C velocity vector. The remaining 260º of each scan (revolution) is used for instrument calibration and housekeeping functions. The (140º) GMI swath represents an arc of 885 km on Earth's surface. 82) 83) 84) 85) 86)

The SMA (Spin Mechanism Assembly) is a precision electro-mechanical bearing and power transfer drive assembly mechanism that supports and spins the GMI instrument at a constant rate of 32 rpm continuously for the 3 year plus mission life. The SMA design has to meet a challenging set of requirements and is based on the BATC space mechanisms heritage and lessons learned changes made to the WindSat BAPTA mechanism that is currently (fall 2011) operating on-orbit and has recently surpassed 8 years of successful Flight operation. 87)

The SMA provides a spin accuracy of 0.1% using a pair of angular contact bearings, separated axially on a shaft driven by a 3-phase direct current torque motor with a 2-speed resolver for communication and position feedback. The high-precision electro-optical bearing hosts a power and data transfer drive. The instrument has its own momentum compensation. The control hardware and software that control instrument spinning and momentum compensation reside within the instrument controller assembly. This assembly consists of the controller itself and a momentum wheel for momentum compensation, installed under the structure supporting the GMI sensor.


Figure 22: Schematic view of the SMA (image credit: NASA, BATC)


Figure 23: Scan geometry of GMI (image credit: NASA)

The instrument features 13 microwave channels (similar to those of TMI) in the frequency range of 10-190 GHz as outlined in Table 5. The noise equivalent delta temperature (NEDT) values are valid for the corresponding integration times where the integration times represent scan movement through one antenna beam width. The GMI beam efficiencies for all channels will exceed 90%, where beam efficiency is defined as the percentage of energy collected from an isotropic scene within the solid angle defined by 2.5 times the channel half-power beam widths and approximating the antenna main lobe between first nulls.


Figure 24: Illustration of the GMI instrument (image credit: NASA, BATC)

The GMI instrument design employs a total power type radiometer with through-the-feed hot and cold calibration. featuring an offset parabolic antenna with an aperture size of 1.2 m. The antenna subsystem includes four feedhorns serving the nine channels. Each frequency is allocated an independent feedhorn with the exception of a shared feedhorn for the 18.7 GHz and 23.8 GHz channels. The antenna subsystem and receiver electronics rotate at 32 rpm. A stationary thermal shroud, with an opening to cold space, surrounds the rotating instrument subsystems. GMI features its own momentum compensation. The control circuitry and logic governing instrument spinning and momentum compensation is contained within the instrument controller assembly. The instrument controller assembly and momentum wheel, providing momentum compensation, are mounted beneath the shelf supporting the GMI sensor.

The 1.22 m diameter aperture of GMI provides excellent spatial resolution (IFOVs) for channels 1 through 5, the channels for which the entire aperture is utilized in beam formation. These GMI channels offer fine spatial resolution when compared to other conical-scanning radiometers (they are between 50-60% better than those of TMI on TRMM).


Figure 25: View of the 183 GHz mixer design (image credit: Millitech Inc.)

The GMI uses a set of frequencies that have been optimized over the past two decades to retrieve heavy, moderate, and light precipitation using the polarization difference at each channel as an indicator of the optical thickness and water content. The GMI has the following channel selections (Ref. 57):

• 10 GHz channel for measuring the heaviest precipitation encountered in the tropics

• 9 and 37 GHz channels for measuring moderate to light precipitation over ocean

• 21 GHz channel for correction of the absorption by water vapor in other channels

• 89 GHz channel for detection of the presence of large cloud ice particles, which is used for delineating convective from stratiform precipitation over ocean and for measuring heavy precipitation over land

• 166 GHz channel for measuring light precipitation in frontal structures outside the tropics

• Two 183 GHz water-vapor sounding channels for detecting scattering signals from small ice particles and shielding the surface in regions of high water vapor to estimate light rain and snowfall rates over snowcovered land.

Channel No

Center frequency (GHz)

Ctr. freq. stabilization (±MHz)

Bandwidth (MHz)


Integration time (ms)


Antenna beamwidth @ 3 dB (º)









































































































Table 5: Performance requirements of the GMI instrument


Figure 26: Channel footprint scheme of GMI in successive along-sans (image credit: NASA)

The choice of GMI sampling times is governed by the desire to achieve “Nyquist spatial sampling” in the along-scan direction of the swath. In addition, samples from individual channels must be co-registered on the Earth surface. The sample times are slightly larger than the integration times due to latencies inherent to the digital sampling electronics. To satisfy the Nyquist criterion, all channels are being sampled at a minimum of two times as the GMI scans through a single IFOV. To guarantee co-registration, the sample times for each channel are being made integral multiples of each other.

Instrument mass, power, design life

153 kg, 141 W, 3 years

Data rate, antenna size (offset parabolic reflector)

25 kbit/s, 1.22 m diameter

No of GMI instruments

2 (one on core S/C, one on constellation)

Table 6: Some parameters of GMI


Figure 27: View of the deployed GMI instrument configuration (NASA, BATC)


Instrument calibration: The primary calibration of the GMI instrument is provided through a hot load and cold sky reflector. The hot load design minimizes thermal gradients and provides thermal stability. The hot load has a shroud that limits the exposure of the hot load to solar radiation. In addition 14 thermistors are provided within the hot load to allow spatial and temporal variations to be tracked accurately. The size of the cold sky reflector has been maximized within the mechanical constraints to provide a high beam efficiency to cold space. Calibration is accomplished using a cold sky target and a precisely controlled hot load. The cold sky target is a reflector targeted at space to provide the coldest possible target for calibration purposes. The sky reflector has been sized to provide a high beam efficiency to cold space.

In addition to the hot load and cold sky reflector, GMI has internal noise diodes that provide additional information for tracking the calibration. The noise diodes will be used to track the stability of the non-linearity of the receivers over the life of the instrument. The noise diodes will also be used to verify the short-term stability of the hot and cold sky calibration points and could be used to provide short-term replacement of these loads.


Figure 28: Illustration of the cold load reflector and the hot load device on the GMI platform (image credit: NASA, BATC)

The GMI radiometer will serve as a `transfer standard' in two contexts:

1) as a radiometric transfer standard for the other radiometers of the GPM constellation, and

2) as a constellation. Both transfer standards represent areas of scientific research. 88)

In the context of item 1, the GMI radiometric calibration will serve as a reference for other radiometers. In this method, the brightness temperature calibration of constellation member radiometers will be adjusted to achieve a common basis with that of the GMI. This technique will reduce precipitation retrieval differences between sensors due to biases from inter-sensor calibration.

The context of item 2 refers to a precipitation transfer standard. Specifically, this concerns the measurement synergy created by the GMI and the DPR instruments aboard the Core spacecraft. The mutual overlap of actively sensed, vertically-profiled, radar data at two frequencies in combination with the multi-channel passive data of GMI is a unique capability of the Core observatory.

The GPM project has developed the concept of products that contain inter-calibrated brightness temperatures (Tb) that are mission consistent across constellation radiometers. These products will be the main brightness temperature products distributed both in production and near-realtime.

A process for inter-calibration has been developed that is based on pair-wise comparison of brightness temperatures from constellation radiometers. This process is an outgrowth of the early prototype work using TMI as the surrogate for GMI and further honed by the extensive comparisons carried out by the x-cal (cross-calibration) team. In addition the process led to the discovery of time-dependent biases in TMI brightness temperatures and also led to the development of consistent TMI products. This helps to vindicate that the process developed can lead to consistent, well-calibrated Tb for GPM radiometers thereby leading to greatly improved retrievals. 89)


GMI RDA (Reflector Deployment Assembly):

The GMI RDA is an articulating structure that accurately positions and supports the main reflector of the GMI (Global Microwave Imager) throughout the 3 year mission life. For GMI to fit within the launch vehicle fairing, the main reflector must be stowed toward the front of the spacecraft bus (Figure 29). 90)


Figure 29: Schematic view of the GPM spacecraft with the stowed GMI instrument (image credit: BATC)

Launch restraints secure the main reflector and RDA to the GMI main structure. After launch, the restraints deploy and the RDA must maneuver the reflector from its stowed location and position it into a precise orientation above the instrument for operation. The on-orbit deployed reflector must match the ground alignment orientation to within 0.5 mm in position and 40 arcsec in order to maintain tight off nadir angle pointing requirements. The RDA must maintain stable reflector orientation throughout the 3 year GMI life.

Figure 30 shows the architecture of the GMI instrument. The ISS (Instrument Support Structure) deck consists of a composite panel that contains the interface to the spacecraft, supports the instrument computer and provides the structure to support the stowed main reflector. The IBA (Instrument Bay Assembly) consists of a hexagonal composite structure with a circular top deck that supports the RF subsystem and the RDA (Reflector Deployment Assembly). The calibration targets are supported on a calibration support structure that is despun and connected back through the stator assembly of the SMA (Spin Motor Assembly) which provides the rotational motion and allows for power and signal transfer between the rotating and stationary elements. 91)

The primary connection between the two composite structures is the SMA and the IBS Launch Restraints (IBS LR). The 3 IBS LR’s act as a load bypass mechanism for the bearings contained within the SMA and provide a direct load path for the supported mass into the spacecraft interface. Actuation of the IBS LR’s allows the SMA to spin the rotating elements. The Main Reflector Launch Restraints (MR LR) provide the stowed interface for the Main Reflector providing a load path through the composite structure to the spacecraft interface. The Main Reflector is positioned to its deployed orientation by the RDA after the launch restraints are released. The deployed orientation can be seen in Figure 27.

The calibration targets are connected to the stator elements of the SMA through a bellows that couples the slip ring on the SMA to the Despin Assembly which supports the calibration assembly and holds the calibration assembly stationary. This calibration support structure contains the Calibration Launch Restraint (CAL LR) which provides an alternate load path to support the mass of the calibration targets without over loading the Despin Assembly.


Figure 30: Stowed configuration of GMI (image credit: BATC)

The GMI RDA is a kinematically determinate structure consisting of an aft bipod structure and forward and aft side strut assemblies that attach to four locations on the instrument upper deck and three locations on the outer perimeter of the main reflector. The RDA is constructed from composite tubes bonded to titanium fittings that are attached to a variety of joints designed to allow the structure to fold into a stowed configuration. When stowed, the reflector is in a defined location where it is restrained for launch, and the RDA strut tubes are positioned into limited available areas within the stowed envelope.

Deployment force is provided by a torsion spring attached to the aft bipod assembly with speed controlled by a fluid damper. Deployment reliability is enhanced by eliminating the possibility of binding of the strut joints. This is accomplished by using combination of spherical and revolute hinges configured such that the structure is effectively under constrained. To provide control during deployment, a novel auxiliary synchronization linkage directs the motion of the reflector during the majority of the operation until the spring loaded side strut elbow joints lock out completing the deployment and forming a geometrically determinate structure. Figure 31 shows the primary elements of the assembly.


Figure 31: Illustration of the RDA components (image credit: BATC)

The location of the stowed reflector is defined by available volume on the spacecraft and launch vehicle thus driving the geometry of the stowed RDA. The RDA must reliably deploy the 12 kg reflector from this location to a repeatable position, and maintain its orientation when exposed to the on-orbit environment throughout the mission life. Table 7 lists the primary performance requirements of the RDA.


5.7 kg max

Deployed stiffness

11 Hz min

Deployment repeatability

< 0.5 mm, < 40 arcsec

Deployment stability

< 0.25 mm, < 20 arcsec

Deployment duration

< 5 minutes

Table 7: Key RDA performance requirements

The RDA strut tubes are cylindrical graphite epoxy tubers with a titanium fitting bonded to each end. They were manufactured with the aid of precisely aligned bond tooling. Once bonded and cured, they were thermal cycled, proof tested and then assembly began by bolting the appropriate hinge fitting to each end of the strut tube.

Deployment testing: The primary performance tests imposed on the RDA included deployment repeatability, deployment duration, deployed stiffness, torque margin, deployment over operating temperature, off-nominal deployment and kinematic model validation.

The GMI RDA development and validation program has successfully demonstrated the capability to precisely deploy a payload over a large range of motion in a controlled and reliable manner. The RDA avoids the complexity and reliability concerns associated with a metrology/feedback closed-loop motorized deployment scheme in favor of a passively powered, kinematically determinate approach. The RDA manages this using a lightweight strut design that is inherently flexible until fully deployed where it becomes a rigid structure. This configuration can be tailored for a variety of payload sizes and deployment requirements. The performance demonstrated by the RDA is applicable to the requirements for most RF antennas, as well as a wide range of optical payloads. These other applications would benefit by leveraging the increased understanding and capabilities gained by the RDA program (Ref. 90).


Science Discipline Areas of GPM

1) Climate Diagnostics: refining & extending precipitation climatologies including snow climatologies; detecting statistically significant global & regional precipitation trends

2) GWEC (Global Water & Energy Cycle) / Hydrological Predictability: global water & energy cycle analysis & modeling; water transports; water budget closure; hydrometeorological modeling; fresh water resources prediction

3) Climate Change / Climate Predictability: climate-water-radiation states; climate-change analysis & prediction; GWEC response to climate change & feedback

4) Data Assimilation / Weather & Storms Predictability: rainfall data assimilation; global-regional scale NWP techniques

5) MBL (Marine Boundary Layer) Processes: air-sea interface processes & surface flux modeling; ocean mixed layer salinity changes

6) Land Processes: land-atmosphere interface processes & surface flux modeling; integrated surface radiation-energy-water-carbon budget process modeling

7) Coupled Cloud-Radiation Models: diagnosis of cloud dynamics, macrophysical/microphysical processes, & response of 3D radiation field; parameterization of microphysics & radiative transfer in non-hydrostatic mesoscale cloud resolving models

8) Retrieval/Validation/Synthesis: physical retrieval of precipitation & latent heating; algorithm calibration & products normalization; algorithm validation & quantification of uncertainty; synthesis of validation for algorithm improvement

9) Applications/Outreach: weather forecasting; flash flood forecasting; news media products; educational tools.



Ground segment of GPM Core spacecraft:

The GPM ground system architecture builds on the lessons learned from and the experiences of TRMM. Specifically, the ground system supports the generation of radiometer precipitation products from the GMI within one hour of observation and combined radar/radiometer swath products within three hours of observation. The ground system consists of fully integrated elements supporting flight operations, data processing and distribution, and ground validation.

The MOC (Mission Operations Center) is highly automated and staffed 8x5 (8 hours, five days a week) as the GPM instruments operate in survey mode and require very little ground commanding. It interfaces with PPS (Precipitation Processing System) to deliver 5-minute duration science instrument files, 5-minute duration housekeeping data files, metadata associated with data processing and delivery, and ancillary data to support science product generation. The MOC also interfaces with PPS to receive instrument commands and command requests as needed (Ref. 57).


Figure 32: Overview of the GPM Core spacecraft ground segment (image credit: NASA)

The PPS (Precipitation Processing System) is based on an evolution of the TSDIS (TRMM Data Information System) permitting algorithm and other prototyping to aid in the GPM data system development. Its function is to create higher-level science data products, deliver science data products to the user community, provide interface to the instrument science teams, and deliver instrument commands and instrument team command requests to the MOC.


Figure 33: Overview of the GPM Core spacecraft ground segment at JAXA (image credit: JAXA)


Figure 34: Data flow of the GPM Core Observatory and the constellation satellites to the Precipitation Processing System at GSFC (image credit: NASA, Ref. 53)



GPM Ground Validation:

The GPM mission supports a vigorous Ground Validation (GV) program for pre-launch algorithm development and post-launch product evaluation. Based lessons learned from the traditional approach to ground validation is to use ground-based observations to directly assess the quality of satellite products, GPM is establishing joint GV sites with partner agencies and a series of pre- and post-launch field campaigns to carry out one or more of the following three types of validation activities (Ref. 57):

• Direct Validation: These activities facilitate statistical comparisons of GPM satellite precipitation products with ground measurements provided by national networks of radars and rain gages from GPM partners around the world. The purpose of these activities is to identify potential discrepancies between spaceborne and ground-based estimates of precipitation that may require more in-depth studies.

• Precipitation Physics Validation: These activities focus on collecting intensive, targeted, airborne and ground-based measurements of precipitation processes and ancillary observations to provide the basis for developing, testing, and refining satellite retrieval algorithms using both model-simulated and observation-derived microphysical databases. The broad aim of these activities is to gain further insights into the physical relationships between clouds/precipitating particles and simulated microwave radiances at different frequencies to refine the interpretation of GPM satellite measurements and retrieval algorithms.

• Integrated Hydrological Validation: These activities follow the paradigm of an “end to end” assessment of GPM multi-satellite precipitation products using hydrological basins as a time-and-area-integrated measure of data quality in terms of coupled hydrologic and land-surface modeling and prediction.

Accordingly, four field campaigns were conducted or are planned in 2010-2012 to be held in different climatic regimes:

• Pre-CHUVA – This GPM-Brazil & NASA field campaign targeted warm rain retrieval over land and was focused at the Alcântara Launching Center in northeastern Brazil on 3-24 March 2010. 92) 93)

• Light Precipitation Validation Experiment (LPVEx) - This joint CloudSat-GPM collaboration is concentrating on light rain in shallow melting layer situations. It is covering the Helsinki Testbed and the Gulf of Finland during September and October 2010. It involves the FMI (Finnish Meteorological Institute) and Environment Canada in addition to NASA. 94)

• Mid-Latitude Continental Convective Clouds Experiment (MC3E) – This is a NASA-DOE (Department of Energy) field campaign at the DOE-ASR Central Facility in Oklahoma. It is planned for April 15 -June 1, 2011. 95)

• High-Latitude Cold-Season Snowfall Campaign: This GPM-Environment Canada campaign concentrated on snowfall retrieval in Ontario, Canada. The project was conducted from January 15, 2012 until March 3, 2012. However, much of the ground instrumentation was installed during November, 2011.

GCPEx (GPM Cold Season Precipitation Experiment) was carried out in the winter of 2011/012 in Ontario, Canada. Its goal was to provide information on the precipitation microphysics and processes associated with cold season precipitation to support GPM snowfall retrieval algorithms that make use of a dual-frequency precipitation radar and a passive microwave imager onboard the GPM core satellite, and radiometers on constellation member satellites. 96)

• Japan's GV: Japan is developing the DPR based on the excellent heritage of the TRMM PR development. Japan's GV focuses on the ground experiment relevant to DPR. An airborne experiment is planned (Ref. 61). 97)


Figure 35: Illustration of field campaign locations in the time frame 2010-2012 (image credit: NASA)


GV (Ground Validation) site in Okinava: NICT (National Institute of Information and Communications Technology) has three facilities on the main island of Okinawa. The main office is the Okinawa Electromagnetic Technology Center (NICT Okinawa), and two radar sites are the Ogimi Wind profiler Facility (NICT Ogimi) and the Nago Precipitation Radar Facility (NICT Nago) as shown in Figure 36. 98)

• COBRA (C-band polarimetric Radar) is installed at the NICT Nago site. The COBRA system is a ground-based, monostatic pulse Doppler radar using a single wave (5340 MHz) in the C-band. The maximum observation range is a radius of approximately 300 km, although this depends on the repetition frequency and the transmitted pulse. The spatial resolution is 37.5–600 m, depending on the pulse width and the over-sampling rate.

• 400-MHz Wind Profiler: The 400-MHz Wind Profiler (400-MHz WPR) is installed at the NICT Ogimi site. The 400-MHz WPR is able to observe simultaneously the atmospheric turbulence echo and the echo from precipitation. Hence, by analyzing the echo power spectrum of the received signal, the rain drop size distribution can be estimated for which the effects of wind speed, the intensity of atmospheric turbulence, and background winds have been removed.

Using these vertical and ground-based measurements of raindrop size distributions, the extinction cross-section and the back scattering cross section can be processed by the Mie scattering theory. Then, the specific attenuation (k) and the radar reflectivity (Z) for Ku-band are estimated. The vertical variations and characteristics (depending on rain type) of rain attenuation for Ku-band can be analyzed. The GPM/DPR for the Ku- and Ka-band algorithm can be also evaluated using this ground validation observation network in Okinawa.


Figure 36: Site locations of the ground validation site in Okinawa (image credit: NICT)

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75) Kinji Furukawa, Masahiro Kojima, Takeshi Miura, Yasutoshi Hyakusoku, Toshio Iguchi, Hiroshi Hanado, Katsuhiro Nakagawa, Minoru Okumura, “Proto-flight test of the Dual-frequency Precipitation Radar for the global precipitation measurement,” Proceedings of IGARSS (International Geoscience and Remote Sensing Symposium), Vancouver, Canada, July 24-29, 2011

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77) T. Miura, M. Kojima, K. Furukawa, Y. Hyakusoku, T. Ishikiri, H. Kai, T. Iguchi, H. Hanado, K. Nakagawa, “Status of proto-flight model of the dual-frequency precipitation radar for the global precipitation measurement,” Proceedings of SPIE Remote Sensing 2012, 'Sensors, Systems, and Next-Generation Satellites,' Edinburgh, Scotland, UK, Vols. 8531-8539, Sept. 24-27, 2012

78) Takeshi Miura, Masahiro Kojima, Kinji Furukawa, Yasutoshi Hyakusoku, Takayuki Ishikiri, Hiroki Kai, Toshio Iguchi, Hiroshi Hanado, Katsuhiro Nakagawa, “Status of Proto-flight Model of the Dual-frequency Precipitation Radar for the Global Precipitation Measurement,” Proceedings of the 29th ISTS (International Symposium on Space Technology and Science), Nagoya-Aichi, Japan, June 2-8, 2013, paper: 2013-n-53

79) Masahiro Kojima, Kinji Furukawa, Takeshi Miura, Takayuki Ishikiri, Yasutoshi Hyakusoku, Toshio Iguchi, Hiroshi Hanado, Katsuhiko Nakagawa, “Development Status of the Dual-frequency Precipitation Radar,” Proceedings of the 28th ISTS (International Symposium on Space Technology and Science), Okinawa, Japan, June 5-12, 2011, paper: 2011-n-34

80) Shigeo Sugitani, Hiroshi Hnado, Seiji Kawamura, Katsuhiro Nakagawa, “Development of Radar Calibrators for the Dual-frequency Precipitation (DPR) installed on the Global Precipitation Measurement (GPM) primary satellite,” Proceedings of the 28th ISTS (International Symposium on Space Technology and Science), Okinawa, Japan, June 5-12, 2011, paper: 2011-n-43

81) David Newell, Sergey Krimchansky, “GPM Microwave Imager Design, Predicted Performance and Status,” URL:

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84) David A. Newell, Gary Rait, Thach Ta, Barry Berdanier, David Draper, Michael Kubitschek, “GPM microwave imager design, predicted performance and status,” Proceedings of IGARSS (IEEE International Geoscience and Remote Sensing Symposium) 2010, Honolulu, HI, USA, July 25-30, 2010

85) “GPM Microwave Imager (GMI),” NASA, Feb. 7, 2011, URL:

86) “GMI Media Kit,” BATC, URL:

87) Michael Kubitschek, Scott Woolaway, Larry Guy, Chris Dayton, Barry Berdanier, David Newell, Joseph Pellicciotti, “Global Microwave Imager (GMI) Mechanism Assembly Design, Development, and Performance Test Results,” Proceedings of the 14th European Space Mechanisms & Tribology Symposium – ESMATS 2011, Constance, Germany, Sept. 28–30 2011 (ESA SP-698)

88) J, B. Sechler, “GPM microwave imager selected calibration features and predicted performance,” Proceedings of IGARSS 2007 (International Geoscience and Remote Sensing Symposium), Barcelona, Spain, July 23-27, 2007

89) Erich Franz Stocker, John Stout, Joyce Chou, “GPM Plans for Radiometer Intercalibration,” Proceedings of the 28th ISTS (International Symposium on Space Technology and Science), Okinawa, Japan, June 5-12, 2011, paper: 2011-n-37, URL:

90) Larry Guy, Mike Foster, Mike McEachen, Joseph Pellicciotti, Michael Kubitschek, “Design, Development and Testing of the GMI Reflector Deployment Assembly,” Proceedings of the 14th European Space Mechanisms & Tribology Symposium – ESMATS 2011, Constance, Germany, Sept. 28–30 2011 (ESA SP-698), URL:

91) Adam Sexton, Chris Dayton, Ron Wendland, Joseph Pellicciotti, “Design, Development and Testing of the GMI Launch Locks,” Proceedings of the 14th European Space Mechanisms & Tribology Symposium – ESMATS 2011, Constance, Germany, Sept. 28–30 2011 (ESA SP-698), URL:

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94) J. Koskinen, J. Koistinen, J. Pulliainen, J. Lemmetyinen, J. Leinonen, T. Lauri, L. Nevvonen, H. Pohjola , D. Moiseev, T. Nousiainen, J. Tyynelä , M. Hallikainen, A. K. von Lerberl, A.Sihvola, B. Vehviläinen M. Huttunen, V. Podsechin, “Finnish GPM program status,” 4th GPM International GV (Ground Validation) Workshop, Helsinki, Finland, June 21-23, 2010, URL:

95) Walt Petersen, NASA MSFC Matt Schwaller, Arthur Hou, “GPM MC3E, Cold Season, and HydroMet Field Campaigns (2010-2013),” 4th GPM International GV (Ground Validation) Workshop, Helsinki, Finland, June 21-23, 2010, URL:

96) David Hudak, Walt Petersen, Gail Skofronick-Jackson, Mengistu Wolde, Mathew Schwaller, Paul Joe, Chris Derksen, Kevin Strawbridge, Pavlos Kollias, Ronald Stewart, “GPM Cold Season Precipitation Experiment (GCPEx),” Proceedings of the 2012 EUMETSAT Meteorological Satellite Conference, Sopot, Poland, Sept. 3-7, 2012, URL:

97) Shuji Shimizu, Katsuhiro Nakagawa, Kenji Nakamura, “Overview of GV activities in Japan,” 4th GPM International GV (Ground Validation) Workshop, Helsinki, Finland, June 21-23, 2010, URL:

98) Seiji Kawamura, Hiroshi Hanado, Shigeo Sugitani, Katsuhiro Nakagawa, Toshio Iguchi, “GPM/DPR Ground Validation Super Site in Okinawa, Japan,” Proceedings of the 28th ISTS (International Symposium on Space Technology and Science), Okinawa, Japan, June 5-12, 2011, paper: 2011-n-42

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.