European Space Agency


The ERS-2 Spacecraft and its Payload

C.R. Francis, G. Graf, P.G. Edwards, M. McCaig, C. McCarthy, A. Lefebvre, B. Pieper, P.-Y. Pouvreau, R. Wall, F. Weschler, J.Louet, W. Schumann & R. Zobl

ESA Directorate for Observation of the Earth and its Environment, ESTEC, Noordwijk, The Netherlands

The ERS-2 satellite is essentially the same as ERS-1 except that it includes a number of enhancements and it is carrying a new payload instrument to measure the chemical composition of the atmosphere, named the Global Ozone Monitoring Experiment (GOME).

Other major instruments common to ERS-1 and ERS-2 are the Active Microwave Instrument (AMI), the Radar Altimeter (RA), the Along-Track Scanning Radiometer (ATSR), the Microwave Radiometer (MWR) and the Precise Range and Range Rate Experiment (PRARE). The AMI operates in three different modes devoted to radar imagery, and oceanic wind and wave measurements. The RA measures precisely the altitude over ocean ice and land surfaces and also measures oceanic wind and waves. The ATSR measures sea-surface temperatures and has been enhanced for ERS-2 by including visible channels for vegetation monitoring. The MWR and PRARE both support the RA mission by providing information respectively on propagation delays of the radar signal and satellite positioning.

ERS-2
Figure 1. The ERS-2 spacecraft

Note: This article is an updated version of an article describing ERS-1 which appeared in ESA Bulletin 65 (February 1991).

The first European Remote Sensing satellite, ERS-1, was launched on 17 July 1991. The satellite had been developed during the 1980s with the objective of measuring the Earth's atmosphere and surface properties, both with a high degree of accuracy and on a global scale. The primary scientific reason behind acquiring such data is to increase our under-standing of the interaction between the Earth's atmosphere and the oceans, in order to deepen our knowledge of the climate and improve global climate modelling.

Other major benefits have been derived from ERS-1 data, including: improved weather and sea-state forecasting and 'nowcasting'; a greater knowledge of the structure of the sea-floor, which is useful for oil and mineral exploration; detailed measurements of the Earth's movements following seismic events; measurements of ice coverage; and the monitoring of pollution, dynamic coastal pro-cesses, and changes in land use.

In order to ensure the continuity of those measurements, ERS-2, a second flight model of the satellite, was planned. Its development was started in the late 1980s and the satellite was launched on 21 April 1995. Although it is essentially the same as ERS-1, the satellite includes a number of enhancements and, in particular, it is carrying a new payload instru-ment that measures the chemical composition of the atmosphere.

The spacecraft

In common with ERS-1, the major components of the ERS-2 payload are active microwave instruments or radars. Powerful radar pulses are needed to provide sufficient illumination of the Earth's surface to produce detectable echo signals from the satellites' polar orbits, which have a mean altitude of about 780 km. The spacecraft also need large antennas to be able to pick up the returning signals. Consequently, the satellites have to be rather large: they each weigh about 2.3 tonnes. The payload alone weighs about 1000 kg and consumes about 1 kW of electrical power when in full operation. The antennas, after deployment, are up to 10 m long; the main payload support structure hasa2 m x2 mbase and is some 3 m high. To support the payload by providing electrical power, attitude and orbit control, as well as overall satellite operational management, a platform module (derived from the French national SPOT programme) is attached to the payload (Fig. 2). That module is roughly equivalent in size to the payload itself and is equipped with a deployable 12 m x2.4m solar array.

Modules of ERS-2
Figure 2. The major modules of the ERS-2 satellite

When the main ERS-1 development contract started, in 1984, the satellite was far larger and more complex than any that ESA had flown previously. A comparison between the Meteosat satellite and ERS-1, for example, shows that ERS-1 is 7.5 times heavier, transmits 750 times more bits of data per second, and has nine active onboard computers, while Meteosat had none.

The largest of ERS-2's sensors, the Active Microwave Instrument (AMI), is capable, in its imaging mode, of producing highly detailed radar images of a 100 km strip on the Earth's surface. This mode is also known as the Synthetic Aperture Radar or SAR mode. Because that mode consumes a large amount of energy and produces a vast amount of data which cannot be stored on board, it is only used regionally, for periods of approximately 10 min per orbit. The same instrument has alternative global measurement modes, namely the Wind (or Scatterometer) Mode in which the wind speed and direction at the sea-surface can be measured over a 500 km swath, and a Wave Mode which provides small radar images at 200 km intervals. Those images can be used to generate ocean-wave spectra, showing wave energy as a function of wavelength and direction.

A second instrument, the Radar Altimeter, provides very precise measurements of the satellite's height above the ocean, ice and land surfaces. The successful exploitation of those height data - which are to be used to study, among other topics, global ocean circulation and height profiles across the ice caps - is dependent upon precise determination of the satellite's orbit, which is derived from the onboard tracking systems. Those systems are a laser retro-reflector, which is a passive device used by ground-based satellite laser-ranging systems, and the PRARE instrument, which is a two-way microwave ranging system that uses small, dedicated ground stations. The PRARE on ERS-1 failed shortly after launch. For ERS-2, the cause of that failure has been eliminated and, furthermore, a second PRARE has been embarked.

Another payload instrument is the Along-Track Scanning Radiometer (ATSR), which consists of two parts. Detailed images of the sea surface are made by an infra-red scanning radiometer, which allows extremely precise measurements of sea-surface temperature. For ERS-2, additional channels have been incorporated to provide imagery in the visible part of the spectrum as well. The other part is a passive microwave radiometer, which is used to determine the water-vapour content of the vertical column of the Earth's atmosphere passing beneath the satellite.

ERS-2 Satellite
Figure 3. Exploded view of the ERS-2 satellite

ERS-2 is also carrying one completely new instrument compared to ERS-1: the Global Ozone Monitoring Experiment, GOME. That instrument provides spectra of backscattered sunlight in the ultra-violet/visible/near-infrared part of the spectrum, while scanning a swath below the satellite. Processing of those spectra, in combination with direct solar spectra which are also measured by GOME for reference, allows the determination of concentrations and profiles of many trace gases, but particularly ozone, in the atmosphere.

The large amounts of data from those instruments are transmitted to the ground via the Instrument Data Handling and Trans-mission (IDHT) Subsystem. That system includes two high-capacity onboard tape recorders to store the data being gathered while the satellite is outside the visibility of the various ground stations.

The orbit

Both ERS-2 and ERS-1 are in a Sun-synchronous polar orbit, highly inclined to the equator, giving the satellites visibility of all areas of the Earth as the planet rotates beneath their orbits. The inclination is such that the precession of the orbit, caused by the non-spherical components of the Earth's gravity field, exactly opposes the annual revolution of the Earth around the Sun. Consequently, the orbital plane will always maintain its position relative to the Sun, crossing the equator with the descending node at about 10:30 am local time. The constant illumination conditions throughout the year which that provides are advantageous for the ATSR and GOME. They also have benefits for the satellite design, in that, for example, the solar array only needs to rotate about one axis, normal to the plane of the orbit, in order to maintain its correct alignment with the Sun.

The orbital inclination required to achieve Sun-synchronism is a weak function of satellite altitude. For a mean altitude of approximately 780 km, it needs to be about 98.5degree, making it a so-called 'retrograde' orbit. The orbital altitude, and consequently the revolution period, may be adjusted by use of the orbit control thrusters provided on both ERS-1 and ERS-2, so that a harmonic relationship exists between the revolution period of the satellite and the rotation period of the Earth. Consequently, after a certain number of orbits, the satellite re-traces its tracks over the Earth's surface. In practice, the orbital altitude of ERS-1 has been changed, by a few kilometres, several times during the four years it has been in orbit. Five orbital patterns have been flown: two had repeat periods of three days but over different ground tracks; for most of the mission, a multi-disciplinary 35-day pattern has been used; and two others had a pattern with a 168-day repeat, offset by half of the track spacing to provide a very dense spatial coverage. Each individual orbit in those patterns lasts approximately 100 min.

ERS-1 and ERS-2 are both flying in the same 35 day orbit, over the same ground tracks. It is foreseen that both ERS-1 and ERS-2 will always remain in that orbit. The phasing of the two satellites around that orbit plane has been adjusted so that they overfly the same track with a one-day separation, ERS-1 being ahead. That provides excellent opportunities to compare the results from the two satellites.

The platform

The spacecraft platform provides the major services required for satellite and payload operation. Those services include attitude and orbit control, power supply, monitoring and control of payload status, telecommunication with ground stations for telecommand reception and telemetry of payload and platform housekeeping data. The platform also houses the two independent PRAREs as passengers.

The platform has been modified with respect to the SPOT programme, in which it was developed as a multi-mission concept, to meet the special needs of the ERS missions. The major modifications have included extension of the solar-array power and battery energy-storage capability, modification of the attitude-control subsystem to provide yaw steering and geodetic pointing, and the development of new software for payload management and control.

The solar array's performance had to be appreciably increased to support ERS-1's power-hungry microwave payload. This has been achieved firstly by increasing the array's effective area (and corresponding power) by about 66%, to approximately 24 m 2 , and secondly by using more efficient solar cells, which produce about 12% more power.

The solar array (Fig. 4) consists of two 5.8 m x 2.4 m wings, manufactured from flexible reinforced Kapton, on which are mounted a total of 22 260 solar cells. The two wings are deployed by means of a pantograph mechanism, and the whole array rotates through 360degree with respect to the satellite during each orbit in order to maintain its Sun pointing.

During the 66-min sunlit phase of each orbit, the array provides electrical power to all of the onboard systems. It also charges the spacecraft's batteries, located in a cylindrical compartment at the solar-array end of the platform, so that they can provide the energy necessary for a similar level of payload operations during the 34-min eclipse periods. The four nickel-cadmium (NiCd) batteries are sized to allow payload operations to be independent of the satellite's orbital position. Connected directly to the spacecraft's unregu-lated 30 V bus, they power it during the 14 eclipses that occur each day, using their combined capacity of 96 Ah. The precise management of the charge and discharge cycles is handled by the onboard computer, with the possibility of ground intervention if required.

Front of Solar Array
Figure 4a. Front side of the solar array (Photo: Aerospatiale)

Rear of Solar Array
Figure 4b. Rear of the solar array, showing the pantograph deployment mechanism

Attitude and orbit control

ERS-2, like ERS-1, is a three-axis-stabilised, Earth-pointing satellite. Its yaw axis is pointed towards the local vertical with respect to a reference ellipsoid, taking the Earth's oblate shape into account. The direction of the pitch axis oscillates slightly during each orbit to keep it oriented normal to the composite ground velocity vector, taking account of the Earth's rotation, to assist the operation of the AMI. The residual attitude errors are no more than 0.06degree on each axis for ERS-1, and ERS-2 is expected to have a similar performance. The attitude control system has the capability to be offset to compensate for any static error that may be observed, but that has not proved to be necessary.

ERS-2 has a range of attitude sensors. The long-term reference in pitch and roll is obtained from one of two continuously operating, redundant infrared horizon sensors. The yaw reference is obtained once each orbit from one of two redundant narrow-field Sun sensors aligned to point at the Sun as the satellite crosses the day/night terminator. The short-term and rate reference are obtained from an inertial core, with a pack of six gyroscopes, of which any three can be in use. Finally, there are two wide-field Sun-acquisition sensors for use in the initial stages of attitude acquisition, and in safe mode, when the satellite is Sun-pointing rather than Earth-pointing.

The primary means of attitude control is provided by a set of momentum wheels (large flywheels), which are nominally at rest. They can be spun in either direction, exchanging angular momentum with the satellite in the process. It is also possible, if there were per-manent torques on the satellite due, for instance, to radiation pressure on the solar array, to bias one or more wheels to a nominal non-zero speed. This has not been necessary with ERS-1. Angular momentum also needs to be dumped from the wheels on a regular basis and a sophisticated system has been devised for this purpose. The onboard computer con-tains a simple model of the Earth's magnetic field, and is also able to control the current in a pair of orthogonal magnetic coils. These coils, called 'magneto-torquers', generate torques by interacting with the Earth's geomagnetic field. Using a servo loop and the built-in field model, the spacecraft's onboard computer continuously adjusts the magneto-torquers to keep the wheel speed close to the nominal values.

ERS-2 has a number of monopropellant-type thrusters, aligned about the spacecraft's three primary axes, in which hydrazine dissociates exothermically as it is passed over a hot-platinum catalyst. They are used in different combinations to maintain and modify the satellite's orbit and to adjust its attitude during non-nominal operations. That is normally done by using pairs of thrusters to provide in-plane thrust when slightly changing the orbital height or speed, or by turning it in yaw to obtain out-of-plane thrust when slightly modifying the orbital inclination.

The payload module

The mechanical structure of the payload has to meet a number of challenging requirements, including rather tight mechanical-stability and thermal-isolation constraints. It was also foreseen that the payload would need to be disassembled many times, and this had to be considered in its basic design. There are two main parts to the payload module (Fig. 5), the Payload Electronics Module (PEM) and the Antenna Support Structure (ASS), for which different design solutions were adopted.

Payload Support Structure
Figure 5. The payload support structure, showing the box-like Payload Electronics Module (PEM) structure and the complex strut assembly of the Antenna Support Structure

The Payload Electronics Module (PEM)
The PEM is an aluminium face-sheet/ honeycomb structure supported by nine internal vertical titanium beams (titanium was selected for its low thermal conductivity and expansion coefficient). The central beam lies at the intersection of two internal cross-walls, so that the PEM is effectively divided into four separate compartments. Each outer panel is dedicated to a particular instrument, to simplify integration logistics, and is fixed to the vertical beams by close-tolerance bushes and titanium screws. This construction minimises settling effects due to vibration and ensures good structural-assembly repeatability.

The payload is separated from the platform by a non-load-bearing electromagnetic (EMC) shield. An aluminium-honeycomb panel closes the opposite end of the structure, stabilising the beams and providing the interface to the ASS at the beam locations. The beams provide a load path from the ASS to the platform.

It was clear that the integration programme would involve many separations of the PEM and the platform and so a system of tapered dowels and shims was developed to ensure repeatability of assembly. To facilitate the con-nection and disconnection of the instrument panels to and from the main harness, there are large connector brackets attached to the lower parts of the panels, with simple covers.

The Antenna Support System (ASS)
The ASS (Fig. 6), requiring structural stiffness while minimising thermal distortion, has been manufactured primarily from high-modulus carbon-fibre-reinforced plastic (CFRP) tubes, with titanium being used for all the highly loaded structural elements such as nodes, strut end-fittings, and interface brackets.

The lower part of the assembly consists of five tripods, three of which provide support points for the SAR antenna and two intermediate support points for the upper assembly. These tripods are also connected to each adjacent node. The CFRP sandwich plate at the top, which carries the Scatterometer antennas, is supported by three further tripods attached to the intermediate points and the SAR central point. The Altimeter's antenna is attached at three node points by a triangulated strut system.

That intricate, highly stable assembly was challenging in terms of design, manufacture and integration. That is amply illustrated by the central titanium node, which interfaces to ten high-tolerance struts without inducing built-in stresses.

The thermal-control system
The thermal-control system is basically a passive design, complemented by an active heater system. The thermal-control approach complements the modular overall design of the satellite, the payload, platform and battery compartment being thermally insulated from one another as far as practicable, allowing separate analysis and testing. The individual modules are also insulated from the external environment by multi-layer insulation blankets, except for the radiators. The latter are covered mainly with materials of low solar absorptance and high infrared emittance, which reject the internally dissipated energy. The radiator areas have been optimised for the extreme hot and cold operating conditions that will be encountered in nominal Earth-pointing attitude, and during the Sun-pointing safe mode in which the payload would be inert. A heat pipe is used to transfer heat from the ATSR to one of the radiators. The GOME, which was added to the payload for ERS-2, obscures one of the original ERS-1 radiators. The necessary thermal dissipation is now provided by a further heat pipe to an enlarged radiator on a nearby panel. The GOME itself has a relatively large built-in radiator panel facing cold space, to which the Peltier coolers, which are used to keep the detector temperatures at 30degreeC, are connected by heat pipes.

ERS-2 Payload
Figure 6. The ERS-2 payload being placed in the Large Space Simulator for environmental testing. The Antenna Support Structure (ASS) can be seen

High heat fluxes in the payload electronics module are spread over larger areas by local skin-thickening of honeycomb side panels or by constant-conductance heat pipes embedded in the panels.

Active heater systems, which are fully redundant during nominal operations and partially redundant in safe mode, provide autonomous thermal control to cope with periods of limited ground contact. The heater systems themselves are controlled pre-dominantly by onboard software in nominal modes and by thermostats in safe mode, or in failure cases where the onboard computer is not available. An anomaly-management system is triggered by failures in the heater systems and/or out-of-range equipment temperatures. It decides on the appropriate corrective action, which can be to switch to redundant heater branches and/or to switch off the payload.

All software parameters used for control or surveillance can be enabled, inhibited or updated by ground command, providing a high degree of flexibility for coping with a variety of unforeseen events or conditions.

On-board command and control

ERS-2 carries a significant number of software packages run by different processors spread throughout the platform and the payload. In the platform, the On-Board Computer (OBC) runs the so-called 'Centralised Flight Software', which is a small software package (44 kwords) incorporating all the basic functions needed to conduct the mission in an optimal fashion. In addition, each payload component (AMI, RA, ATSR and IDHT, described in more detail below) contains its own decentralised Instrument Control Unit (ICU).

These five computers are linked by the On-Board Data-Handling (OBDH) bus, and communicate with each other via a high-level packetised protocol. The PRARE, as a platform passenger, is not a user of the OBDH bus, but also has built-in intelligence, as does the GOME, which is controlled by the control unit of the ATSR.

That set of interdependent computers fulfils a critical requirement. ERS-2 is an extremely complex satellite, with a great many modes, parameters and logical conditions to be set and respected throughout each orbit. It is required to have 24-hour autonomy, and that could only be achieved by providing intelligent payload elements controlled by a capable central computer. A basic concept in that philosophy is the 'macrocommand', a coded instruction expanded and acted upon by the ICU. In that way, the ICU relieves the OBC of many detailed tasks related to internal instrument configuration and operations.

It was a primary requirement that all of the onboard processors be reprogrammable in flight, and many of the operating characteristics are controlled by tables of variable parameters. Commands are provided for manipulation of those tables to enable major changes in the operating characteristics to be easily achieved.

The main functions of the OBC flight software are:

The ICUs run software packages whose functionality depends on the particular instrument that they serve, but there is some commonality. The common ICU tasks are:

The other functions of the ICUs are related to scientific data conditioning/processing, and are therefore more specific to each instrument. Both the AMI and RA ICUs interface with scientific computers, known as the Scatterometer Electronics and Signal-Processor Sub-Assembly (SPSA), respectively. The AMI ICU manages a large memory buffer which accommodates the data originating from the sampling of the radar echo in SAR and Wave modes, while the IDHT ICU manages the tape recorders.

There are two types of time-management functions to be carried out onboard, namely the scheduling of events and the time-stamping of measurements. All timing is referenced to a clock maintained by the OBC, providing time signals with 4 ms resolution and correlated with UTC (Universal Time Coordinated) by the Kiruna (Sweden) ground station. Events are scheduled by associating a time with each macrocommand.

The time-stamping of measurements, known as 'datation', is also performed by the ICUs, which write the appropriate binary time code, transcribed from the ICU clock, into the secondary header of each source (data) packet.

Deployments

During the first few orbits after ERS-2's separation from the launch vehicle, a period known as the Launch and Early-Orbit Phase, or LEOP, several units of the spacecraft were deployed (Fig. 7). They include the solar-array arm and panel on the platform, and the SAR antenna, fore and aft Scatterometer antennas and ATSR antenna on the payload. In designing these deployments and their sequencing, a number of constraints had to be observed:

All of the deployments were controlled by the OBC, some via a pyrotechnic activation sequence triggered by the separation from the launcher, and some via time-tagged macrocommands. The macrocommands were loaded into the OBC before launch and were thus executed at times independent of the actual time of launch during the 5 min window. To maintain synchronism between the two types of deployment, had ERS-2 not been launched at the opening of the launch-window, the macrocommand queue could have been updated very shortly after separation from the launcher, when ERS-2 was visible from the Wallops ground station, on the east coast of the USA. There was also a possibility of updating from Fairbanks in Alaska, or Perth in Australia.

The ATSR microwave antennas were released by pyros 5 s after separation. A spring drive then rotated it into its latched position in just a few seconds. Next, the solar-array arm's deployment began with a pyro release firing, less than 1 min after separation, the further deployment requiring no additional commands and being mechanically sequenced and driven by spring forces (Fig. 8). Deployment of the solar-array panels themselves did not start until about 45 min after separation, when ERS-2 was visible from Perth. The deployment was again passive, with the two panels being pulled out of their container by spring-driven pantographs.

The SAR-antenna deployment (Fig. 9) started 75 min after separation, within the visibility of the Santiago de Chile ground station. The two antenna wings each have spring-driven and motor-driven phases, and the whole sequence was initiated by firing a pyro to release six lever clamps holding the folded antenna in launch configuration.

The Scatterometer antennas were deployed immediately after the SAR antenna. They were stowed at the sides of the PEM for launch, and were also released by pyro firing. Each antenna deployment involves a single motor-driven rotational movement.

Operations Sequence for Launch
Figure 7. Nominal operations sequence for the Launch and Early-Orbit Phase (LEOP)

Deployment of Solar Array Arm
Figure 8a. Deployment of the solar-array arm

Spring Drive
Figure 8b. One of the spring drives for solar-array deployment

Deployment of SAR Antenna
Figure 9. Deployment of the SAR antenna

Finally, after 2,1/4 h and 1,1/4 orbits, when the Sun was directly overhead, the solar array rotation was enabled. All the ERS-2 deployments occurred precisely as planned.

Instrument data-handling and telemetry

ERS-2 has two telemetry systems. The plat-form's needs are served by a classical-type Telemetry, Telecommand and Control (TTC) system operating at S-band. That low-rate (2 kbit/s) system is used to transmit the ICU formats for housekeeping purposes. Because of the high bit rates involved, the science data cannot use this link and the payload therefore includes a so-called 'Instrument Data Handling and Transmission' (IDHT) system. That system allows real-time transmission of AMI Image-Mode data, providing a regional service to local ground stations and global recording and telemetry of the other sensors.

The instruments generate data in the form of 'source packets', which constitute a logical division of telemetry data from the instrument point of view. However, they are not the fundamental unit as far as transmission to the ground is concerned, for which a further division into 'transport frames' is made. The latter are smaller than source packets and, in addition to pieces of source packets, contain synchronisation and transmission error-control information. The source-packet structure is then reassembled from transport frames at the ground stations.

Three data streams are transmitted from the IDHT (Fig. 10). The first contains the high-rate data from the AMI Image Mode, with auxiliary data and a copy of the S-band telemetry data, at a total rate of 105 Mbit/s. This channel has an X-band link dedicated to it. The other sensors have their data combined, again with a copy of the S-band data and satellite ephemeris information, into a (comparatively) low-rate data channel, operating at 1.1 Mbit/s, which is continuously recorded by the onboard tape recorder (Fig. 11). This recorder is replayed at 13.6 times recording speed (in reverse order to save rewind time) when over the ground stations to form a second data channel, at 15 Mbit/s. It shares the second X-band link with the live transmission of the combined low-rate data, which constitutes the third data stream.

IDHT
Figure 10. Block diagram of the X-band science data transmission system (IDHT)

Tape Recorder
Figure 11. One of the two 6.5 Gbit tape recorders, which can hold 3000 ft of ,1/4-inch tape

The tape recorder has been designed to store a full orbit of continuous 1.1 Mbit/s low-rate data on 3000 ft of ,1/4-inch mag-netic tape, leading to a total data-recording capacity of 6.5 Gbit. When performing a data dump to high-latitude ground stations, such as the primary Kiruna station, the spacecraft's solar array might cause a brief occultation of the link, due to the system geometry. On passes when that occurs, the on-board command scheduling includes a stop in playback before the occultation, a slight rewinding of the tape, and a reac-tivation of playback mode after the occul-tation.

The modulation scheme used for the high-rate channel is quadrature phase-shift keying, called QPSK, which allows four distinct states per clock cycle and makes it possible to transport two bits of information per cycle. That reduces the radio-frequency bandwidth required for transmission by a factor of two compared with a simpler modulation scheme. The low-rate link uses unbalanced quadrature phase-shift keying, or UQPSK, to modulate the 15 Mbit/s recorder dump and the convolutionally encoded real-time data onto a single link. If there are no recorder dump data, bi-phase-shift keying (BPSK) is used for the real-time data.

Immediately before and after recorder playback, the link is automatically switched between BPSK and UQPSK operation, with minimum impact on the real-time data stream. The ERS-1/ERS-2 ground demodulators have been designed to accommodate that mode-switching automatically.

The fact that the X-band transmission was required to have a minimum power-level fluctuation during the satellite pass led to the design of a shaped-beam antenna able to compensate for losses at low satellite elevation angles, when the distance to the ground station is long, and the attenuation due to the atmosphere's water content is high. To achieve that, the antenna reflector is shaped so that its radiation pattern compensates for the inverse-square-law variation in received power with distance as the satellite passes across the sky at the ground station. The polarisation of the radiated energy is rotated to compensate for Faraday rotation due to the Earth's ionosphere.

The IDHT is physically located on the Earth-facing panel of the PEM, with the tape recorders mounted inside, on one of the cross-walls.

The scientific instruments

The Active Microwave Instrument (AMI)
Two separate radars are incorporated within AMI, a Synthetic-Aperture Radar (SAR) for Image and Wave Mode operation, and a Scatterometer for Wind Mode operation. The operational requirements are such that each mode needs to be able to operate inde-pendently, but the Wind and Wave Modes are also capable of interleaved operation, in so-called 'Wind/Wave Mode'.

Wave-Mode Characteristics
AMI Wave-Mode Characteristics

Image-Mode Characteristics
AMI Image-Mode (SAR) Characteristics

Wind-Mode Characteristics
AMI Wind-Mode Characteristics

In Image Mode, the SAR obtains strips of high-resolution imagery 100 km in width to the right of the satellite track (Fig. 12). The 10 m long antenna, aligned parallel to the flight track, directs a narrow radar beam onto the Earth's surface over the swath. Imagery is built up from the time delay and strength of the return signals, which depend primarily on the roughness and dielectric properties of the surface and its range from the satellite.

The SAR's high resolution in the range direction is achieved by phase coding the transmit pulse with a linear chirp, and compressing the echo by matched filtering. Range resolution is obtained from the travel time. Azimuth resolution is achieved by recording the phase as well as the amplitude of the echoes along the flight path. The set of echoes over a flight path of about 800 m is processed (on the ground) as a single entity, giving an azimuth resolution equivalent to a real aperture 800 m in length. This is the 'synthetic aperture' of the radar.

Operation in Image Mode excludes the other AMI operating modes, and power considerations limit operating time to a maximum of 10 min per orbit. The data rate of 100 Mbit/s is far too high to allow onboard storage, and so images are only acquired within the reception zone of a suitably equipped ground station.

SAR Image Mode
Figure 12. The SAR Image Mode

SAR Wave Mode
Figure 13. The SAR Wave Mode

Wave-Mode operation of the SAR provides 5 km x5 kmimages at intervals of 200 km along track (Fig. 13), which can then be interpreted to provide wave spectra. The relatively low data rate allows onboard data storage, and thus a global sampling of wave spectra is obtained.

The Wind Mode uses three antennas to generate radar beams looking 45degree forward, sideways, and 45degree backward with respect to the satellite's flight direction (Fig. 14). These beams continuously illuminate a 500 km-wide swath as the satellite moves along its orbit, and each provides measurements of radar backscatter from the sea surface on a 25 km grid. The result is three independent backscatter measurements for each grid point, obtained using the three different viewing directions and separated by a short time delay.

SAR Wind Mode
Figure 14. The SAR Wind Mode

As the backscatter depends on the wind speed and direction at the ocean surface, it is possible to calculate the surface wind speed and direction by using those 'triplets' within a mathematical model.

The AMI electronics (see right panel) cover two full 2 m x1 mside panels of the PEM. In addition, the calibration unit is mounted on a cross-wall inside the PEM, the switch matrix and its controller are on the top panel, and the four antennas, one of the most characteristic elements of the ERS-1 and ERS-2 satellites, on the Antenna Support Structure (ASS).

The Radar Altimeter (RA)
The Radar Altimeter is a nadir-pointing pulse radar designed to make precise measure-ments of the echoes from ocean and ice surfaces. It has two measurement modes, optimised for measurements over ocean and ice, respectively. In the so-called 'Ocean Mode', the echo characteristics of interest are:

The radar echoes over ice sheets, particularly the rough surfaces at the continental margins, show much greater variances in shape than oceanic echoes. In order to maximise the data return in those areas, the Ice Mode includes three features designed to improve its 'robust-ness'. The range window width is increased by a factor of four, which also degrades precision by a similar amount. A simplified height-tracking loop greatly improves the ability to keep the echo in the range window, although it cannot distinguish the leading edge of the signal. Finally, the tracker is more agile.

In the Ice Mode, as in the Ocean Mode, the telemetered data stream contains the effective height of the range window, and the digitised echo waveform within this window. They allow ground processing to retrieve topographic information. The returned power level is also telemetered.

The effective pulse width is 3 ns, which is equivalent to about 45 cm in two-way range. The radar is said to be 'pulse-width-limited' because not all of the target is illuminated simultaneously by the short pulse, and the received power is controlled by the illumination.

Over ocean surfaces, the distribution of the heights of reflecting facets is gaussian or near-gaussian, and the echo waveform has a characteristic shape that can be described analytically. It is a function of the standard deviation of the distribution, which is closely related to the ocean wave height.

Different echo waveforms occur over ice surfaces. Over sea ice, there is generally a strong specular component, while the rough topography of continental ice sheets at the margins leads to complex return waveforms. In central ice sheet areas, the height distribution becomes more regular and echoes similar to ocean returns are observed.

Real echoes are composed of the sums of signals from many point-scatterers, each with individual phase and amplitude. To reduce uncertainties in the determination of pulse characteristics, the Radar Altimeter averages pulses together to reduce that statistical effect.

The constraints of available peak transmit power and required pulse width determined that a pulse-compression technique be used to spread the required energy over time, allowing reduced peak power (see panel below).

RA Characteristics
Radar Altimeter (RA) characteristics

The AMI Electronics

The radio-frequency (RF) subsystem units, covering half of a panel in the spacecraft, contain all the electronics needed to generate the transmit pulses and to amplify and filter the received signals.

The intermediate frequency (IF) radar contains a transmit and a receive section. The transmit section, in Image Mode, generates a linearly chirped pulse of 15.8 MHz bandwidth and 37.2 micros length. This pulse is generated by gating the 123 GHz output of the frequency generator into a short pulse and applying it to a dispersive delay line. At the output of the delay line, the pulse is amplified and cut to the correct length of 37.2 ms. In Wind Mode, the transmit pulse is generated by the Scatterometer electronics, and the IF radar acts only as an amplifier.

The up- and down-converters are contained in a single unit. The upconverter converts the output signals of the IF radar to 5.3 GHz and amplifies them to a level of about 250 mW, required for the input of the high-power amplifier (HPA).

The two redundant units of the HPA occupy one complete panel; each consists of a large power conditioning unit (EPC), a travelling-wave-tube amplifier, an output isolator, and an output filter. The latter two elements are located on the outside of the panel. The HPA amplifies the input signals to output levels of about 5000 W.

The output signal from the HPA arrives at the circulator assembly, or switch matrix, on the top panel of the PEM. This matrix of ferrite circulators switches the signal coming from the HPA to any of the four antennas, and on the return path directs the receive signal from the chosen antenna into the receive chain.

The waveguides from the switch matrix to the four antennas are lightweight CFRP units with a rectangular cross-section of 4 cm x 2 cm, internally metallised. They are rigidly connected to the SAR antenna and the mid Scatterometer antenna, while the connection to the deployable fore and aft antennas is by choke flanges, without a fixed connection.

The largest of the AMI antennas is the SAR antenna, with a radiating area of10 m x1 m.Itisaslotted-waveguide array made of metallised CFRP. The antenna itself is subdivided into ten electrical and five mechanical panels. Its planarity across its 10 m length is better than 1.5 mm when in orbit.

The three Scatterometer antennas are made of aluminium alloy. Like the SAR antenna, they are slotted-waveguide arrays, and each is subdivided electrically into two panels. The central unit, measuring 2.3 m x 0.34 m, contains eight waveguides, while the fore and aft arrays, measuring 3.6 m x 0.25 m, each contain six waveguides.

All of the antennas are designed for vertical polarisation.

The receive echo arrives at the receive part of the IF radar, via the circulator assembly, the receiver shutter, which safeguards the sensitive low-noise receiver against transmission-pulse leakage, and the down-converter. In nominal operation, the IF radar works for both SAR and Scatterometer mode as an amplifier and filter stage. In SAR mode, however, onboard range compression can be commanded from the ground, which then switches the signal through an inverse dispersive delay line, compressing the echo pulses by a factor of about 600. Depending on the mode of operation, the output is fed to the SAR processor or the Scatterometer electronics.

The SAR processor filters the signal and down-converts it to baseband. After analogue-to-digital conversion, auxiliary data are added, then the data are buffered and delivered to the IDHT for transmission to the ground. The SAR processor additionally functions as the AMI's ICU. The Scatterometer electronics also has two tasks. It filters and digitises the Wind-Mode echoes and transfers them to the IDHT for transmission to ground. It also controls the AMI during Wind-Mode operation.

The echoes from the fore and aft antennas have rather a high Doppler shift, which varies from approximately 70 to 150 kHz across the swath. This Doppler spread would prevent narrow-band filtering to reduce noise. The Scatterometer electronics therefore, while the echoes are coming in, changes the local oscillator frequency according to the expected instantaneous Doppler shifts. This acts as Doppler compensation. This is also applied to the mid echoes, but here the required compensation is small.

Apart from providing a sample of the transmitted signal into the receiver for calibration purposes, the other task of the calibration unit is to delay a SAR transmit pulse and feed it back to the IF branch of the receiver. This signal is used as a replica of the chirped transmit pulse for on-ground range compression in the ground processor, as an alternative to the onboard range compression mentioned earlier. On-ground range compression is, in fact, the nominal operating mode.

AMI
Figure 15. Functional block diagram of the Active Microwave Instrument (AMI)

Operating Principle
Figure 16. Schematic of the Radar Altimeter's operating principle. The signal at various points is shown as a frequency/time plot

The RA Electronics

The chirp generator, which is based on surface acoustic wave (SAW) devices, is triggered at a fixed rate of almost 1020 Hz. The chirps pass through a 20 s SAW delay line used to separate transmit and receive chirps during calibration. After upconversion to the transmit frequency, they are amplified by the high-power amplifier, a 50 W travelling wave tube (TWT). The pulses pass via the front-end electronics (FEE), which is an arrangement of circulators and the calibration coupler, to the antenna, a front-fed paraboloid.

Returning echoes arrive, via the antenna, FEE, and low-noise amplifier (LNA), at the microwave receiver. When the echo is expected to return, the chirp generator is re-triggered and a second chirp generated. During the upconversion and multiplication process, a slight frequency offset is introduced, and this becomes the first intermediate frequency (IF). This local oscillator chirp is mixed with the received echo in the 'deramping mixer' in the microwave receiver. A series of tones is thus generated, centred on the first IF.

The microwave receiver is a dual-conversion system, and after conversion to baseband the in-phase (I) and quadrature (Q) signals are passed to the signal processor sub-assembly, or SPSA. The next important stage, inside the SPSA, is the spectrum analyser where the spectrum of the tones is found. This spectrum exactly represents the time structure of the echo waveform in 64 points at an equivalent spacing of 3.03 ns. The average power spectrum over 50 successive pulses is formed, and finally this information is used by the parameter estimator. This step is essential in order to provide the estimate of when the next echo is expected to return, for the chirp re-triggering. As an indication of the need for this estimate, the full bandwidth of the spectrum analyser is equivalent to a height window of about 30 m in the ocean mode.

The maximum height rate is about 30 m/s; if the height estimate were not continually updated, the signal could be completely lost in about 1 s.

Sometimes, however, the echo can be lost, for example, as a result of passing over some topographic features such as mountains. In this case, the acquisition mode is automatically entered. This is a sophisticated multi-stage scheme, partially relying on dedicated hardware processing, which virtually guarantees getting any trackable surface into the tracking mode range window in just over 1 s.

The parameter estimator is a microcomputer, within the SPSA. It is used in acquisition and tracking modes. In ocean and ice tracking, it runs software tracking loops which follow the signal characteristics. In the ocean mode, there are three main loops to track echo time-delay (height), leading-edge slope, and echo power. The error signals used as input to these loops are derived from adaptive discriminators.

The time-delay and echo-power loops are also used in the ice mode, although the error signals are derived from different discriminators. Because of the reduced chirp bandwidth, the spectral points are spaced at 12.12 ns intervals, leading to a range window of about 115 m.

Internal open-loop calibration is performed every minute. This procedure is very fast (about 100 ms). The transmitted signal is coupled into the receiver through an attenuator, and analysis of the received signal is performed on the ground to determine the delay around the system. The major item omitted in this scheme is the ultra-stable oscillator (USO), which provides the echo timing. This calibration is obtained by broadcasting the USO frequency via the IDHT to enable measurements to be made on the ground.

The Along-Track Scanning Radiometer and Microwave Sounder (ATSR-M)
The ATSR-M consists of two instruments, an Infrared Radiometer (IRR) and a Microwave Radiometer (MWR). The IRR has been upgraded from the ERS-1 instrument to include three channels in the visible part of the spectrum as well as the possibility of performing some on-board averaging of data, to flexibly allow the extra channels of information to be provided within the same overall data-rate limitations.

ATSR Characteristics
Table 2. Along-track Scanning Radiometer (ATSR) chearacteristics

The primary objective of the IRR is to measure the global Sea-Surface Temperature (SST) for climate-research purposes. Its absolute accuracy is better than 0.5 K when averaged over areas of 50 km x 50 km, assuming that 20% of pixels within the area are cloud-free. For the cloud-free pixels, of 1 km x 1 km, the relative accuracy is about 0.1 K.

To achieve those objectives, the IRR was designed as an imaging radiometer with four co-registered channels with wavelengths of 1.6, 3.7, 11 and 12 µm, defined by beam splitters and multi-layer interference filters. The Instantaneous Field of View (IFOV) at the nadir on the Earth's surface is a 1 km x1 km square, which is imaged onto the detectors via a f/2.3 paraboloidal mirror. These detectors, fixed onto a Focal-Plane Assembly, are cooled to 80 K by a Stirling-cycle cooler in order to reduce their background noise to an acceptable level.

The 1.6 and 3.7 µm channel data are trans-mitted alternately, switched by a day/night logic provided as a service by the platform.

The IFOV is scanned over the Earth's surface by a rotating plane mirror in such a way that it gives two Earth views, namely a 0degree or nadir view and a 47degree or forward view. The rotation period is 150 ms and the scan is subdivided into 2000 pixels of 75 micros each. In order to calibrate the optical and electrical signal chain, two black bodies (one hot and one cold) within the IRR are scanned during the rotation. After onboard data compression, a packet of 960 pixels (555 nadir-view and 371 forward-view pixels, and 16 hot and 16 cold black-body pixels) is transmitted to ground, together with housekeeping and datation. Extensive on-ground data processing then permits retrieval of the IRR final product, namely the Sea-Surface Tempera-ture (SST).

The main objective of the ATSR Microwave Sounder is to measure the atmos-pheric integrated water content (vapour and liquid) in order to compute the most problematic part of the tropospheric path delay in the Radar Alti-meter's signals.

The MWR has two channels, operating at 23.8 and 36.5 GHz, each with a band-width of 400 MHz. The instrument is nadir-viewing, using an offset antenna deployed shortly after the space-craft's separation from the launcher. Onboard calibration is performed by a sky horn pointing to cold space, and internal hot loads. The acquisition cycle is synchronised to the ATSR scan occurrence and the MWR data are merged into the IRR packets described above.

Global Ozone Monitoring Experiment (GOME)
GOME is an optical spectrometer spanning the ultra-violet/visible/ near-infrared wavelength range from 240 to 790 nm, with a spectral resolution of 0.2 to 0.4 nm. Light upwelling from the Earth, decomposed by the instrument into its spectral components and recorded on four silicon array detectors with 1024 pixels each, carries the absorption signatures of ozone and a number of other atmospheric trace gasses. The quantitative concentra-tion of those molecules is derived from the spectrum by fitting the highly banded absorption cross-section of the molecules to the measured spectrum.

LRR Characteristics
Table 3. Laser Retro Reflector(LRR) characteristics

The instantaneous field of view, corresponding to the projection of the spectrometer slit on the Earth, is a narrow rectangle 40 km long, aligned in the along-track direction. In order to observe a large fraction of the atmosphere, a scan mirror sweeps that field of view in the across-track direction. The sweep, which normally sweeps 960 km in 4.5 s with a rapid flyback, together with an integration time on the detectors of 1.5 s (30 s for the UV part), results in ground pixel sizes of 320 x 40 km. With that, global coverage can be achieved within three days. However, smaller pixel sizes can be commanded by reducing the angular extent of the scan, with spatial coverage reduced accordingly.

In order to enable high-accuracy, long-term trend measurements, a calibration unit enables regular views of the Sun. Additionally, a wavelength calibration lamp provides the possibility to regularly check the wavelength stability of the instrument and can also be used to monitor the Sun calibration path for possible degradation. During the times when it is not used, a shutter protects the Sun calibration path.

As the instrument is sensitive to the polarisation of the incoming light, a polarisation detector array monitors one polarisation direction in the broadband channels corresponding essen-tially to the array detector channels 2, 3, and 4.

The Laser Retro-Reflector (LRR)
The Laser Retro-Reflector is a passive device which is used as a target by ground-based laser ranging stations. The operating principle is to measure the time of a round trip of laser pulses reflected from an array of corner cubes mounted on the Earth-facing side of the spacecraft's Payload Electronics Module (PEM) (Fig. 19).

That array consists of a polyhedral housing with a hemispherical arrangement of one nadir-looking corner cube in the centre, surrounded by an angled ring of eight corner cubes. This allows laser ranging for satellite passes in the range of 0degree to 360degree azimuth and 30degree to 90degree elevation at the ground.

The Precise Range and Range-rate Equipment (PRARE)
The PRARE is a satellite tracking system which performs two-way microwave range and range-rate measurements to ground-based transponder stations with high precision. Signal-propagation effects are compen-sated by two-frequency measurements, for ionospheric refraction, and ground-station collection of meteorological data for tropospheric refraction.

PRARE Characteristics
Table 4. PRARE characteristics

Two signals are transmitted to ground, one at S-band (2.2 GHz) and one at X-band (8.5 GHz) frequencies (both signals modulated with the pseudo-random noise code). The ground stations receive the two simultaneously emitted signals with a slight time difference and determine the time delay. This provides a measure of the ionospheric refraction taking place in the atmosphere.

The received signals are demodulated and a coherent regenerated copy of the X-band (7.2 GHz) sequence retransmitted to the satellite, where the two-way travel time and the two-way Doppler measurements are carried out, so that the range and range-rate can be determined. Two-way measurements are possible for up to four stations simultaneously via so-called 'code multiplexing'.

Both the space-to-ground and ground-to-space links have additional capacity for data transmission at low bit rates. Control codes and broadcast ephemerides for ground-station operation are transmitted in the downlink, and calibration data, ionospheric-measurement results and meteorological ground data are included in the uplink. All measurement data are stored inside the PRARE itself, in 512 kbytes of RAM, and dumped during the next available ground-station pass.

The PRARE on ERS-1 failed shortly after launch, a failure ascribed to destructive latch-up of a RAM chip caused by the radiation environment. For ERS-2, the parts have been replaced by radiation-resistant devices and various software modifications have been made. Furthermore, a second PRARE is installed on ERS-2 to provide a similar level of redundancy to the majority of the payload.

Integration and testing

The main integration phase of the ERS-2 Programme started in January 1991. Since ERS-2 was, with a few exceptions, an exact rebuild of ERS-1, an assembly, integration and test programme was devised which relied on the satellite qualification having been achieved by ERS-1. Only acceptance testing was required. For ERS-1, the mechanical/ thermal qualification was performed on the SM (structural model) and the EM (engineering model). The only deviations from that concept were GOME, because it was a new instrument, and PRARE, because it had undergone a major redesign after its in-orbit failure on ERS-1.

The manufacturing of the flight units for the payload core instruments and the instrument integration were completed by the end of 1992. The integration and testing of the instruments proceeded very smoothly. Since the work had been carried out by almost the same teams as for ERS-1, there was no 'learning curve' effect and the instruments were delivered well ahead of schedule.

Measurement Principle
Figure 17. Measurement principle of the Along-Track Scanning Radiometer (ATSR)

Ozone Measurement
Figure 18. Computer illustration of the measurement of atmospheric ozone by GOME. Each measure is represented by a rectangle of a different colour, and corresponds to 40 x 80 km of the Earth's surface. Sardinia and Corsica are in the lower left corner, and Denmark is in the upper right corner

LRR
Figure 19. The Laser Retro-Reflector (LRR). Each corner cube is individually made, to compensate for satellite motion in reflecting incident laser energy back exactly along its incoming path

Orbit Determination
Figure 20. Precise orbit determination by PRARE and Laser Retro-Reflector range measurements

The platform units were manufactured at the same time. Some manufacturing problems at unit level, especially in the propulsion subsystem, caused late delivery of items for the platform integration. In parallel to the manufacturing of the recurrent units, the GOME design and development was undertaken. Initially, it was foreseen to have only a GOME breadboard and flight model but during the course of the project, that was found not to be sufficient. Therefore the breadboard model was upgraded to a fully-fledged engineering model which was then exposed to environmental qualification tests. Furthermore, the initial GOME integration tests at payload and satellite level were performed with the engineering model.

After the in-orbit failure of the PRARE onboard ERS-1, the manufacturing of the PRARE for ERS-2 was halted until the results of the failure investigation were known. As a result, a major electrical redesign had to be initiated. The design changes were tested on the ERS-1 PRARE engineering model before being implemented in the flight and flight spare models. The changes affected exclusively the electronic design, and so a protoflight qualification concept was adopted for the flight unit. The flight spare was acceptance tested.

The integration of the flight model payload started at the end of 1992. After the mechanical and electrical integration of all instruments, a successful full-performance test was conducted. The payload assembly integra-tion and test programme was completed well on schedule. It included a three-week thermal vacuum/thermal balance test in the large space simulator at ESTEC in Noordwijk (NL), and the final integration and alignment of the SAR and Scatterometer antennas.

In parallel to the payload integration, the platform was integrated. After a functional performance test and a thermal vacuum test at Intespace, Toulouse (F), the platform was delivered to ESTEC for satellite integration in December 1993, exactly on time despite the late delivery of some units.

The satellite integration started in January 1994. After payload and platform coupling, a series of electrical tests, including hardware/ software compatibility and a system validation test with ESOC, were performed. Prior to the environmental test, the propulsion subsystem was checked for leak-tightness.

Vibration and acoustic testing were then carried out to demonstrate, successfully, that the ERS-2 satellite would not be adversely affected by the vibration and noise induced by the launch vehicle. That was followed by a deployment test, under onboard software and under ESOC control, of the SAR, the Scatterometer and the Microwave Radiometer antennas using special 'zero-gravity' rigs to simulate a realistic deployment.

The satellite assembly, integration and test (AIT) activities concluded with a full functional performance test and a rehearsal of the launch site procedures in August 1994. The AIT programme was completed exactly on the date that had been originally planned at the beginning of the ERS-2 programme in 1990.

Following the flight acceptance review, the flight hardware and the associated ground support equipment were prepared for shipment to the launch range in Kourou, French Guiana. Five sea containers and two dedicated 747 cargo flights were used to transport the equipment to the launch site.

ERS-2 was planned to be launched on Ariane flight V72 in January 1995. The launch campaign started on 14 November 1994. After the set-up of the check-out equipment, the satellite mechanical preparations and the alignment, a post-transport functional performance test was performed. Immediately before the end of that test, on 30 November, however, the launch of Ariane flight V70 failed. The post-transport test of ERS-2 was completed but the satellite then had to be placed in storage until the preparations for flight V72 could be resumed. The satellite was protected by a dedicated tent, which was purged and the ambient conditions inside the tent were permanently monitored.

The launch campaign was interrupted for two months, from 19 December 1994 to 17 February 1995. Only a small 'babysitter' team remained in Kourou.

In February 1995, the campaign resumed with a short functional test, the integration of the solar array and the preparation for fuelling.

After further delays due to hydrogen and oxygen leaks in the third stage of the Ariane V71 launch vehicle, ERS-2 was transferred to the filling and encapsulation hall. Finally, on the night of 20 April 1995, ERS-2 was successfully launched.

Acoustic Testing
Figure 21. The ERS-2 flight model undergoing acoustic testing in the ESTEC LEAF facilities


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Right Left Up Home ESA Bulletin Nr. 83.
Published August 1995.
Developed by ESA-ESRIN ID/D.