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Terra Mission (EOS/AM-1)

Spacecraft    Launch    Mission Status    Sensor Complement    EOS    References

Terra (formerly known as EOS/AM-1) is a joint Earth observing mission within NASA's ESE (Earth Science Enterprise) program between the United States, Japan, and Canada. The US provided the spacecraft, the launch, and three instruments developed by NASA (CERES, MISR, MODIS). Japan provided ASTER and Canada MOPITT. The Terra spacecraft is considered the flagship of NASA's EOS (Earth Observing Satellite) program. In February 1999, the EOS/AM-1 satellite was renamed by NASA to "Terra". 1) 2) 3) 4)

The objective of the mission is to obtain information about the physical and radiative properties of clouds (ASTER, CERES, MISR, MODIS); air-land and air-sea exchanges of energy, carbon, and water (ASTER, MISR, MODIS); measurements of trace gases (MOPITT); and volcanology (ASTER, MISR, MODIS). The science objectives are:

• To provide the first global and seasonal measurements of the Earth system, including such critical functions as biological productivity of the land and oceans, snow and ice, surface temperature, clouds, water vapor, and land cover;

• To improve the ability to detect human impacts on the Earth system and climate, identify the "fingerprint" of human activity on climate, and predict climate change by using the new global observations in climate models;

• To help develop technologies for disaster prediction, characterization, and risk reduction from wildfires, volcanoes, floods, and droughts

• To start long-term monitoring of global climate change and environmental change.

Complemented by aircraft and ground-based measurements, Terra data will enable scientists to distinguish between natural and human-induced changes.

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Figure 1: Illustration of the Terra spacecraft (image credit: NASA)


Spacecraft:

Terra consists of a spacecraft bus built by Lockheed Martin Missiles and Space (LMMS) in Valley Forge, PA. The spacecraft is constructed with a truss-like primary structure built of graphite-epoxy tubular members. This lightweight structure provides the strength and stiffness needed to support the spacecraft throughout its various mission phases. The zenith face of the spacecraft is populated with equipment modules (EMs) housing the various spacecraft bus components. The EMs are sized and partitioned to facilitate pre-launch integration and test of the spacecraft.

EPS (Electrical Power Subsystem): A large single-wing solar array (size of 9 m x 5 m = 45 m2), deployed on the sunlit side of the spacecraft, maximizes both its power generation capability and the cold-space FOV (Field of View) available to instrument and equipment module radiators. The average power of the satellite is 2.53 kW provided by a GaAs/Ge solar array (max of 7.5 kW @ 120 V at BOL). The solar array is based on on a prototype lightweight flexible blanket solar array technology developed by TRW (use of single-junction GaAs/Ge photovoltaics). A coilable mast is used for the deployment of the solar array. The Terra spacecraft represents the first orbiting application of a 120 VDC high voltage spacecraft electrical power system implemented by NASA. A PDU (Power Distribution Unit) has been designed to provide 120 DC (±4%) under any load conditions. This regulated voltage, in turn, is achieved via a sequential shunt unit (SSU) and the 2 BCDUs. A NiH2 (nickel hydrogen) battery is used (54 cells series connected) to provide power during eclipse phases of the orbit. 5) 6) 7)

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Figure 2: Coilable mast deployer for the Terra solar array (image credit: NASA)

GN&C (Guidance Navigation and Control) subsystem: Terra is a three-axis stabilized design with a single rotating solar array. The GN&C subsystem is made up of sensors, actuators, an ACE (Attitude Control Electronics) unit, and software. A three-channel IRU (Inertial Reference Unit) determines body rates in all control modes. Solid-state star trackers provide fine attitude updates, processed by a Kalman filter to maintain precise 3-axis inertial knowledge. A 3-axis magnetometer senses the Earth's geomagnetic field, primarily for magnetic unloading of reaction wheels, but also as a sensor to determine an attitude failure during a deep space calibration maneuver. 8)

The backup sensors include an ESA (Earth Sensor Assembly) for roll and pitch sensing, and coarse sun sensors for pitch and yaw sensing of the sun line relative to the solar array. A fine sun sensor is used in the event that one star tracker fails or during the backup stellar acquisition mode. In addition to these sensors, a gyro-compassing computation is performed for backup yaw attitude determination.

A reaction wheel assembly provides primary attitude control. During normal mode, a wheel speed controller is available to bias the wheel speeds at a range that avoids zero rpm crossings (stagnation point). Magnetic torquer rods regulate the wheel momentum to < 25% capacity in four-wheel mode and < 50% capacity in the three-wheel mode (backup mode). Thrusters are used for attitude control during all velocity change maneuvers and for backup attitude control and wheel momentum unloading.

GN&C is a fault-tolerant system that includes an FDIR (Fault Detection, Isolation and Recovery) capability unique to each of the different operational control modes. If an attitude fault is detected, FDIR transfers all control functions to the ACE unit configured to use all redundant hardware. Once in safe mode, FDIR is disabled.

Sensor component

Units

Manufacturer/model

Mission heritage

Solid State Star Tracker (SSST)

2

BATC / CT-601

MSX, XTE

Earth Sensor Assembly) (ESA)

2

Ithaco / conical scanning

UARS

Coarse Sun Sensor (CSS)

2

Adcole / 42060

UARS

Fine Sun Sensor (FSS)

1

Adcole / 42070

TOPEX

Three Axis Magnetometer (TAM)

2

NASA/GSFC

EUVE, UARS

Inertial Reference Unit (IRU)

2

Kearfott / SKIRU-DII

XTE

 

 

 

 

Actuator component

Units

Manufacturer/model

Heritage

Reaction Wheel Assembly (RWA)

4

Honeywell / EOS-AM

Similar to EUVE

Magnetic Torquer Rod (MTR)

3

Ithaco / TR500CFR

EUVE

Attitude Control Thruster

6 (x 2)

Olin Aerospace (Primex)

 

Delta-v thruster

2 (x 2)

Olin Aerospace (Primex)

 

Table 1: Overview of GN&C sensors and actuators

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Figure 3: Artist' view of the Terra spacecraft in orbit (image credit: NASA)

The design life of the Terra spacecraft is six years. The spacecraft bus is of size of 6.8 m (length) x 3.5 m (diameter) and has a total launch mass of 5,190 kg. The total payload mass is 1155 kg.

RF communications: The primary Terra telemetry data transmissions are via TDRS (Tracking & Data Relay Satellite) system. A steerable HGA (High Gain Antenna) and associated electronics are mounted on a deployed boom extending from the zenith side of the spacecraft. This location maximizes the amount of time available for TDRS communications via this antenna without obstruction by other pads of the spacecraft. Emergency communication is done via the nadir or zenith omni antenna. Command and engineering telemetry data are transmitted in S-band. The science data recorded onboard are transmitted via Ku-band at 150 Mbit/s. The nominal mode of operation is to acquire two 12 minute TDRSS contacts per orbit. During each TDRSS contact, both S-band and Ku-band transmission is being used.

The average data rate of the payload is 18.545 Mbit/s (109 Mbit/s peak); onboard recorders for data collection of one orbit. Mission operations are performed at GSFC. 9)

Broadcast of data: Besides Ku-band and S-band communication, Terra is also capable of downlinking science data via X-band. The X-band communication can be operated in three different modes, Direct Broadcast (DB), Direct Downlink (DDL) and Direct Playback (DP). DB and DDL is used to directly transmit real-time MODIS and ASTER science data respectively to users.

The DAS (Direct Access System) provides a backup option for direct transmission in X-band. DAS supports transmission of data to ground stations of qualified EOS users around the world. These users fall into three categories:

- EOS team participants and interdisciplinary scientists

- International meteorological and environmental agencies

- International partners who require data from their EOS instruments

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Figure 4: The Terra spacecraft in the cleanroom of LMMS at Valley Forge (image credit: LMMS)

 

Launch: The launch of the Terra spacecraft took place on Dec. 18, 1999 from VAFB, CA, on an Atlas-Centaur IIAS rocket.

Orbit: Sun-synchronous circular orbit, altitude = 705 km, inclination = 98.5º, period = 99 minutes (16 orbits per day, 233 orbit repeat cycles). The descending nodal crossing is at 10:30 AM.

Orbit determination is performed by TONS (TDRS Onboard Navigation System) which estimates Terra's position and velocity, drag coefficient, and master oscillator frequency bias. TONS is updated by Doppler measurements at the spacecraft's receivers and provides the attitude control software with a desired pointing ephemeris. Ground-based orbital elements are uplinked daily for backup navigation.

As of March 1, 2001, the Landsat-7, EO-1, SAC-C and Terra satellites are flying the so-called "morning constellation" or "morning train" (a loose formation demonstration of a single virtual platform). There is 1 minute separation between Landsat-7 and EO-1, a 15 minute separation between EO-1 and SAC-C, and a 1 minute separation between SAC-C and Terra. The objective is to compare coincident observations (imagery) from various instruments (synergistic effects). 10)