|17 July 1991||21 April 1995|
The ERS-1 satellite was launched on 16 July 1991, from Kourou in French Guiana. It had at that time been under development for over a decade, with some of the predecessor studies dating back to the early seventies. ERS-1 is one of the major European satellite developments of the eighties. In June 1996 it was put into hibernation after 5 years of operations, where it is acting as an in-orbit spare for ERS-2..
In the mid-seventies the US was developing Seasat, a satellite based on the much-used Agena bus, which would be dedicated to geophysical measurements of the oceans. Its payload included a Synthetic Aperture Radar (SAR) which would be able to generate high-resolution radar images of land and water surfaces, a radar altimeter accurate to about 10cm over the oceans, to be used for measurements of the marine geoid and ocean circulation, a radar scatterometer which could measure wind speed and direction at sea, and an imaging passive microwave radiometer having many geophysical applications. In Europe meanwhile optical/infra-red observing systems were under development: the Meteosat geostationary weather satellite programme was underway at ESA while CNES were defining the SPOT polar-orbiting optical imaging satellite. The development of microwave technology which would offer the same capabilities as Seasat was seen by many as the natural complement to these optical systems.
Interest in this was spread through many of the member states of ESA and was recognised in the Agency itself as the way forward. Consequently a number of studies were undertaken to determine the feasibility of developing some of the necessary sensors in Europe. The main interest was the SAR, as this was identified as being able to address a potential problem with the exploitation of space imagery in Europe: the combination of often-cloudy skies and relatively small geographical features (agricultural fields etc). It was also evident that the high resolution of SAR (which could be as good as a few metres) and its day/night, all-weather capability would offer many uses of interest to the governments of ESA's member states, and several large European electronics companies were eager to become involved.
At this stage ideas for the overall system were still undefined; the purpose was essentially to understand the technology, to establish the system equations, to learn what would be the critical elements and design drivers and to start thinking about the mission characteristics and constraints. For the next step though, the Phase A, it would be necessary to build up a complete payload for the full system-level feasibility study. It was necessary for ESA to define a set of requirements. Three conceptual missions were defined: the Global Ocean Monitoring Satellite System, GOMSS, the Land Applications Satellite System, LASS and the Coastal Ocean Monitoring Satellite System, COMSS. There would be more than one Phase A study.
By this time it had become evident that CNES would be offering the platform which it was developing for the SPOT program as a multimission platform, capable of hosting other remote sensing missions. This had the potential of saving costs compared to the development of an all-new satellite, though clearly there could be constraints as well. This would be examined in the Phase A studies. Soon the use of the SPOT multimission platform became a system requirement.
Phase-A system studies of LASS and COMSS started in 1978, the year Seasat was launched (and failed: Seasat provided useful measurements for 3 months before a short-circuit in the solar-array slip rings killed the satellite after 99 days in orbit), performed by small teams of industrial contractors. The main tasks which were undertaken during these studies were the following:
The COMSS Phase A study was performed between September 1978 and April 1979, by a consortium of 5 European companies led by BAe. It was necessary to develop initial designs for the various instruments and the payload data telemetry system during this phase so that interface definition could proceed. The exception to this was the SAR, where results from two earlier SAR Phase A studies were available. Nevertheless much of the primary activities were in mission design, overall system definition and satellite configuration.
As the SPOT platform had become a system requirement there were many features introduced which removed choices. It was a sun-synchronous design - the solar array rotation, thermal control and attitude sensor design were all dependent on such an orbit configuration. Furthermore the platform was 3-axis stabilised with the orientation of the various faces of the satellite defined. These factors put many constraints on the system.
Sun-synchronism defines the relationship between orbital altitude and inclination. In the range of orbits under consideration the inclination is about 98 deg - a retrograde, near-polar orbit. This high inclination is well adapted to an earth observation mission and the sun-synchronism is also desirable for optical instruments (as it provides relatively constant solar illumination angles), but is unnecessary for the microwave sensors.
The COMSS Phase A study was followed by a number of riders, mainly driven by the addition of further sensors to the planned payload. Eventually there were more sensors desired than the platform could accommodate. Additionally the need for better definition of the candidate sensors meant that a series of instrument Phase A studies was begun.
The decision point arrived at the end of 1980. The situation at that time was the following:
A high level advisory committee of eminent scientists was formed, which, after frequent meetings to discuss the scientific and technical merits, finally, in early 1981, selected a single payload configuration for the new satellite. It was approved by the member states of ESA in March 1981. The selected payload was:
This achievement was an important milestone; it enabled the completion of the mission studies to a Phase A level, and the subsequent system studies to restart in earnest. The committee had also devised a new name for the satellite: it was to be called the first ESA Remote Sensing satellite, ERS-1 (later this became the European Remote Sensing Satellite).
The major loser in this decision was the IMR. By way of compensation an allocation of 50kg and 50W was to be provided for an 'Announcement of Opportunity' (AO) instrument, or package, which could be provided directly by one or more member states. This AO instrument left the payload incompletely defined but was a way of allowing the Phase B preparations to proceed. Actions to select the AO instrument started quickly, with an Announcement in April 1981; evaluation of the offers would be completed by January 1982, for selection in March 1982
ESA started the preparation of the Phase B. The system specification had to be produced and the programme had to be approved and funded by the member states. Industry was able to start preparations for the Phase B by performing advanced work based on what was known of the instruments already. Much of this preparatory work was unfunded as the Phase B was to be competitive. Throughout Europe companies began to court each other with the objective of establishing a competent and winning consortium. At this time in Europe there were three major international consortia which had each worked on various satellite projects. They were called MESH, STAR and COSMOS. For the new satellite MBB and Dornier took the lead in MESH and STAR respectively in trying to expand their consortia to include the new microwave expertise which would be needed. The ultimate failure of this exercise started the decline and eventual collapse of the old consortium system.
But there was to be one more surprise. When ESA finally added up the cost of the proposed system the member states found that it was too expensive. ESA was instructed to reduce the cost significantly. This was finally achieved, in October 1981, by a rather brutal (but common) method: the OCM was removed from the payload and system completely.
The configuration derived by one of the industrial teams at the end of these activities, in late 1981, just before the issue of the Phase B Request for Proposal is shown here. The payload comprises an altimeter, and the AMI. A substantial electronics housing is now evident within the base of the strut assembly, which had been missing in earlier concepts.
At this time it was known that the OCM was gone from the payload, but the replacement AO instrument was not yet known. This illustration shows a number of interesting features: the scatterometer antennas are orientated away from the platform (unlike previous configurations), the struts form a unified structure compared to the individual strut assemblies previously proposed, and all of the payload electronics are now contained in a single large electronics module. These features were all developed during 1981 and are hallmarks of the final flight system.
The AO led, in 1982, to the selection of the Along-Track Scanning Radiometer (ATSR) designed to make high accuracy measurements of sea-surface temperature in a 500km swath, and the Precise Range and Range Rate Equipment (PRARE), a microwave tracking system. The ATSR additionally included a nadir viewing microwave radiometer; both this and the PRARE were selected in support of the altimeter mission.
The aerospace companies of Europe had had great difficulty in establishing competitive consortia. Two camps had formed, led by MBB and Dornier System, but many key companies were in both! This illustrated how important the development of the new technology in ERS-1 was seen. Not only did companies not want to risk being in a losing consortium for commercial reasons but several governments had also joined in, instructing their major electronics companies to bid in both camps.
It was obvious to all that a genuinely competitive Phase B would be impossible in this climate. ESA had therefore decided to hold an initial "beauty contest" between MBB and Dornier by requesting, from the prime contractors only, a condensed technical description of the proposed system and definition of the industrial team, at the end of 1981. Dornier System won this contest and the full Phase B Invitation to Tender was issued in April 1982. Dornier collected a team which included most of the major aerospace companies in Europe.
Although the final payload complement which was now selected had never been the subject of an industrial Phase A study, ESA had been performing its own system studies and the results of these were contained in the ITT; hurriedly modified from an earlier concept which included the OCM.
The industrial team had been performing its own preparations for the Phase B. The ensuing Phase B design flowed from this work, not the ESA concept. There were a number of key differences: one particularly noticable one was the location of the scatterometer antennas, and here the decision process was interesting. During the pre-Phase B industrial studies the overall payload configuration was decided during a long brain-storming session held at Dornier System in August 1981. The first element to place was the SAR antenna because of its high mass and complex deployment. It was easy to agree that it should be 'low' on the satellite (when in launch configuration) both for mass properties and vibration reasons. Its movements during deployments then defined a volume which could contain nothing. The OCM (still on the payload at this time) and RA were placed next, just a bit further out, again for mass properties reasons. Then came the Scatterometer antennas. There were two options on the table: deployed back alongside the satellite or deployed in the opposite direction. There were pros and cons for each: 'back' for mass properties, 'forward' for thermal and EMC. There were other opinions too, but the deciding argument was the apparent existence of an ESA paper showing the 'forward' position. This eventually turned out to be misleading as ESA had in fact made the other decision.
The nominal Phase B activities finished in early summer 1983, but difficulties in securing the agreement and subscription of the ESA member states to the full Phase C/D programme led to an extension of Phase B and pre-C/D activities until the C/D start in December 1984. Despite the difficult situation for industry during this period the design of the ERS-1 system, including the detailed space segment design and a detailed concept of the ground segment, were completed during this phase. During this phase the orbital altitude was increased from 650 km to 780 km, primarily to simplify orbit maintenance and determination. Also, by the end of this period, the consortium of more than 50 companies throughout Europe and other countries had been established.
When the C/D phase started, in 1984, ERS-1 was far larger and more complex than any spacecraft previously flown by ESA. In comparison with Meteosat, ERS-1 is 7.5 times heavier, transmits 750 times more data per second, and has nine active onboard computers, while Meteosat had none.
ERS-1 is a large satellite, weighing about 2.3 tonnes - the payload alone weighs about 1000kg and requires about 1 kW of power for full operation.
The multimission SPOT platform was modified to meet the needs of the ERS-1 mission. This 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 24 m2 solar array consists of 22 260 solar cells mounted on two wings, manufactured from flexible reinforced Kapton. These were deployed by a pantograph mechanism, and rotate through 360 during each orbit in order to maintain Sun pointing - this single-axis rotation is a benefit of the sun-synchronous orbit. During each 100 minute orbit the array is in sunlight for 66 minutes, during which time it provides on-board power and charges the four NiCd batteries. Management of the charge and discharge cycles is handled by the onboard computer. The batteries provide sufficient power during the eclipses to enable full payload operations.
ERS-1 is a three-axis-stabilised, Earth-pointing satellite. Its yaw axis points towards the local vertical defined by an ellipsoid which takes the Earth's oblate shape into account. 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 provide doppler compensation. This, nominal, mode of the attitude control system is called Yaw-Steering Mode (YSM), and it is at the top of a tree of modes whose different properties enable the spacecraft to efficiently transfer from any state ( eg separation from the launch vehicle) into operation. The residual attitude errors in YSM are less than 0.06 on each axis.
There are several attitude sensors on board. The long-term pitch and roll reference is from an infrared horizon sensor; the yaw reference is obtained once per orbit, as the satellite crosses the terminator, from a narrow-field Sun sensor. The short-term and rate reference is from an inertial platform, with a pack of three gyroscopes. Each of these sensors has a backup unit. Finally there are two wide-field Sun sensors for use in the initial stages of attitude acquisition, and in safe mode, when the satellite would be Sun- rather than Earth-pointing.
The primary means of attitude control is the set of momentum wheels, nominally at rest. They can be spun in either direction, exchanging angular momentum with the satellite in the process. Angular momentum is continuously dumped from the wheels: the onboard computer contains a simple model of the Earth's magnetic field, and is able to control the current in a pair of orthogonal magneto-torquers (see 10.4.2), which generate torques by interacting with the Earth's geomagnetic field. Using a servo loop and the built-in field model, the computer continuously adjusts the magneto-torquers to keep the wheel speed close to the nominal values. The power of this servo loop was vividly demonstrated shortly after the launch of ERS-1, when the spacecraft first entered the Fine Pointing Mode (FPM) - the first amongst the sequence of attitude control modes to use the wheels. They immediately saturated, an effect which was quickly traced to a polarity error in the magneto-torquers. This was easy to correct by software patch, while the origin of the polarity error was found to be a failure of the relevant Test Procedure to identify the type of "compass" to be used.
ERS-1 also has a number of hydrazine thrusters, aligned about the spacecraft's three primary axes. They are used in different combinations to maintain and modify the satellite's orbit and to adjust its attitude during certain non-nominal operations. This is done by using pairs of thrusters to provide in-plane thrust when maintaining or changing the orbital height, or by turning the spacecraft in yaw to obtain out-of-plane thrust when maintaining or modifying the orbital inclination.
There are two main parts to the payload module, the Payload Electronics Module (PEM) and the Antenna Support Structure (ASS), for which different design solutions were adopted.
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. The top panel closes the opposite end of the structure, stabilising the beams and providing the interface to the ASS at the beam locations. These 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 assist in the frequent connection 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.
The ASS, requiring structural stiffness whilst minimising thermal distortion, was 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 and strut end-fittings. 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.
ERS-1's thermal-control system is basically a passive design, complemented by heaters. The thermal-control approach complements the modular overall design of the satellite, the payload, platform and battery compartment being thermally insulated from one other 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. These radiator areas have been optimised for the extreme hot and cold operating conditions that will be encountered in nominal attitude (Earth-pointing) and during the safe mode (Sun-pointing) in which the payload would be inert. Active heater systems provide autonomous thermal control, predominantly by onboard software in nominal modes and by thermostats otherwise.
A heat pipe is used to transfer heat from the ATSR to one of the radiators. 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 these panels.
ERS-1 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) performs all the basic functions needed to conduct the mission. In addition, each payload component contains its own decentralised Instrument Control Unit (ICU). These five computers are linked by the On-Board Data-Handling (OBDH) bus, and communicate via a high-level packetised protocol. They are all reprogramable in flight.
This set of interdependent computers fulfils a critical requirement. ERS-1 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 h autonomy, and this could only be achieved by providing intelligent payload elements controlled by a capable central computer. A basic concept in this philosophy is the 'macrocommand', a coded instruction expanded and acted upon by the ICU. In this way the ICU relieves the OBC of many detailed tasks related to internal instrument configuration and operations.
In comparison to contemporary satellites ERS-1 imposes a substantial load on the mission control centre. While a telecommunications satellite may generate and telemeter some hundreds of parameters (520 for Intelsat 5 and 840 for Eutelsat II), and SPOT generates some 500, ERS-1 telemeters no fewer than 5750 housekeeping parameters. This reflects the fact that it is an extremely complex system.
Time-management functions onboard are of two types: the scheduling of events and the time-stamping of measurements. All timing is referenced to a clock maintained by the OBC, which is correlated with UTC at the Kiruna ground station. The ICUs maintain their own clocks, synchronised with the OBC clock. Events are scheduled by associating a time with each macrocommand. The time-stamping of measurements, known as 'datation', is also performed by the ICUs.
Several items were deployed during the first few orbits after ERS-1's separation from the launch vehicle. The design of these deployments had to respect constraints due to dynamics, timing, ground-station visibility, power availability and thermal limits.
The S-band telemetry and ATSR microwave antennas were released by pyros 5 s and 8 s after separation, respectively and then rotated into latched positions by spring drives. Next the solar-array arm's deployment started by pyro release firing, less than 1 m after separation. Further deployment was mechanically sequenced and driven by springs. The two solar-array panels were pulled out of their container by spring-driven pantographs. At this stage the array rotation drive remained locked.
The two SAR antenna wings each had spring-driven and motor-driven phases. The sequence was initiated by pyro firing to release six lever clamps holding the folded antenna in launch configuration. This was in visibility of the Santiago de Chile ground station, 75 m after separation. The timing of this deployment with respect to the solar array was an interesting decision. It had been determined that the shocks induced by the SAR antenna deployment would be unsafe for the unlocked array, while the shock of the array deployment would be unsafe for the deployed SAR antenna. Furthermore it was unsafe for the whole mission to delay the array deployment for the SAR antenna: any difficulties would lead to rapid discharge of the batteries. Hence the two stage array deployment, in which the drive was only unlocked after all other deployments - even in its locked condition it could prevent total loss of battery power.
The Scatterometer antennas were deployed immediately after the SAR antenna. They were stowed at the sides of the PEM for launch, and also released by pyro firing. Each antenna deployment involved a single motor- driven rotational movement. However the fore Scatterometer antenna stalled shortly after pyro release - it was retracted using the redundant motor windings and then successfully deployed by these redundant windings. This incident vindicated the philosophy of doing everything in visibility of ground stations, as well as the relatively late design change to introduce potentiometers to the antenna hinges.
ERS-1 has two telemetry systems. The platform's needs are served by a classical Telemetry, Telecommand and Control (TTC) system operating at S-band. This low-rate (2 kbit/s) system is used to transmit the ICU data for housekeeping purposes. Because of the high bit rates involved, the science data cannot use this link and the payload therefore includes the Instrument Data Handling and Transmission (IDHT) system. This enables 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': a logical division of telemetry data from the instrument point of view. For transmission to the ground 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 reassembled from transport frames at the ground stations.
Three data streams are transmitted from the IDHT. The first, with a dedicated X-band link, contains the high-rate data from the AMI Image Mode, at a total rate of 105 Mbit/s. The other sensors have their data combined, again with a copy of the S-band data and satellite ephemeris information, into the so-called low bit rate (LBR) data channel, operating at 1.1 Mbit/s, which is continuously recorded by the 6.5 Gbit onboard tape recorder, which provides one full orbit of recording capacity. 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, constituting the third data stream.
When performing a data dump to high-latitude ground stations, such as the primary Kiruna station, the solar array can cause a brief occultation of the link, due to the system geometry. On passes when this occurs, the onboard command scheduling includes a stop in playback before the occultation, a slight rewinding of the tape, and a reactivation of playback mode after the occultation.
The modulation scheme used for the high-rate channel is quadrature phase-shift keying (QPSK, see 13.2.2), which allows four distinct states per clock cycle and makes it possible to transport two bits of information per cycle. This reduces the 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, (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 ground demodulators have been designed to accommodate this 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 this, 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.
Two separate radars are incorporated within the AMI, a Synthetic-Aperture Radar (SAR) for Image and Wave Mode operation, and a Scatterometer for Wind Mode operation. Operation in Image Mode excludes the other AMI operating modes, and its power consumption limits operating time to a maximum of 10 m per orbit. The data rate of 105 Mbit/s is far too high to allow onboard storage, and so images can only be acquired within the reception zone of a suitably equipped ground station.
In Image Mode, the SAR obtains strips of high-resolution imagery 100 km in width to the right of the satellite track. 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. Azimuth resolution is achieved by recording the phase as well as the amplitude of the echoes. The set of echoes over a flight path of about 800 m is processed 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.
Wave-Mode operation of the SAR provides 5 km x 5 km 'imagettes' at intervals of 200 km along track, 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 45 forward, sideways, and 45 backwards with respect to the satellite's flight direction, illuminating a 500 km-wide swath. Three independent backscatter measurements for each grid point are obtained, using the three different viewing directions and separated by a short time delay. As the backscatter depends on the wind speed and direction at the ocean surface, it is then possible to calculate the surface wind speed and direction by using these `triplets' within a mathematical model.
The AMI electronics cover two full 2 m x 1 m side panels of the PEM. In addition, the calibration unit is mounted on one of the cross-walls 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 satellite, on the ASS.
The largest of the AMI antennas is the SAR antenna, with a radiating area of 10 m x 1 m. It is a slotted-waveguide array made of metallised CFRP. Its planarity across its 10 m length is better than 1.5 mm in-orbit. The three Scatterometer antennas, also slotted-waveguide arrays, are made of aluminium alloy. The central antenna, 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.
The Radar Altimeter is a nadir-pointing pulse radar designed to make precise measurements of the return echoes. 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 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. In order to maximise the data return in these areas, the 'Ice Mode' includes features designed to improve its `robustness', at the expense of precision.
The constraints of available peak transmit power and required pulse width again determined that a pulse-compression technique be used to spread the required energy over time, allowing reduced peak power.
The ATSR-M consists of two instruments, an Infrared Radiometer (IRR) and a Microwave Radiometer (MWR).
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 these objectives, the IRR was designed as an imaging radiometer with four co-registered channels with wavelengths of 1.6, 3.7, 11 and 12 um, defined by beam splitters and multilayer interference filters. The Instantaneous Field of View (IFOV) at the nadir on the Earth's surface is a 1 km x 1 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. This was the first use of such a cooler in space.
The IFOV is scanned over a 500 km swath by a rotating plane mirror in such a way that it gives two Earth views; at nadir, and 57 forward. The rotation period is 150 ms and the scan is subdivided into 2000 pixels of 75 us each. In order to calibrate the optical and electrical signal chain, two black bodies (one hot and one cold) within the IRR are also scanned during the rotation.
The main objective of the ATSR Microwave Sounder is to measure the atmospheric integrated water content (vapour and liquid) in order to compute the most problematic part of the tropospheric path delay in the Radar Altimeter's signals. The MWR has two channels, operating at 23.8 and 36.5 GHz, each with a bandwidth of 400 MHz. The instrument is nadir-viewing, using an offset antenna. Calibration measurements are made by a sky horn pointing to cold space and internal hot loads.
The Laser Retroreflector is a passive device used as a target by ground- based laser ranging stations. They measure the round trip time of laser pulses reflected from this array of corner cubes, mounted on the Earth-facing side of the PEM. The array consists of a polyhedral housing with a hemispherical arrangement of one central, nadir-looking corner cube surrounded by an angled ring of eight corner cubes. This allows laser ranging for satellite passes in the range of 0 - 360 azimuth and 30 - 90 elevation at the ground.
The PRARE is a satellite tracking system which was intended to make two-way microwave range and range-rate measurements to ground-based transponder stations with high precision. Unfortunately it failed shortly after initial switch-on; this was caused by a destructive latch-up of the RAM chips caused by the radiation environment.
The design was very ambitious. Two signals were transmitted to ground, at S-band (2.2 GHz) and X-band (8.5 GHz), both signals modulated with a pseudo-random noise code. The ground station received the two simultaneously emitted signals with a slight time difference, caused by signal-propagation effects, and determined this difference. This provided a measure of the ionospheric refraction, and the ground-station directly measured meteorological data for tropospheric refraction.
The received signals were 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 were carried out, so that the range and range-rate could be determined. Up to four stations could be handled simultaneously by code multiplexing.
Both the space-to-ground and ground-to-space links had additional capacity for data transmission at low bit rates. Control codes and broadcast ephemerides for ground-station operation were transmitted in the downlink, and calibration data, ionospheric-measurement results and meteorological ground data included in the uplink. All measurement data were stored inside the PRARE itself, in 64 kbyte of RAM, and dumped during the next available ground-station pass.
Following the start of the Phase C/D in December 1984, an initial assembly, integration and test programme was devised relying on a classical sequence of three full-size spacecraft models:
The model philosophy finally adopted differed significantly from this as a result of various difficulties encountered during the development programme.
The Structural-Model programme was completed first, in 1987, without major deviations from the baseline. It included the normal vibration and acoustic testing, as well as a static strength test of the payload on the CESTA centrifuge in Bordeaux (F); the only facility in Europe large enough to test the ERS-1 payload, which is at the limit of even its capabilities.
Integration of the Engineering-Model payload began late due to serious development problems with a number of the onboard instruments. The most significant of these was with the AMI SAR Processor, a complex unit handling command and control and scientific data at up to 100 Mbit/s. The project also experienced difficulties with integrating the Electrical Ground-Support Equipment (EGSE) necessary to test the payload. The EGSE occupies a floor area of about 100 m2 and is more complex than the payload itself, so it is perhaps not surprising that problems were encountered.
The Engineering-Model payload programme included a three-week thermal-balance test in the Large Space Simulator at ESTEC, in Noordwijk (NL). Here the ATSR was found to be running some 15 K too hot; a problem traced to the thermal modelling of the complex obstructions of its thermal view to deep space. The solution was to fit a long heat pipe to a radiator surface, in the Flight-Model payload.
The next important test phases were conducted with the so-called `Flight-Engineering Model' (FEM) satellite, namely a test to prove the compatibility of the payload with the flight software loaded into the platform's onboard computer, and the Radio-Frequency Compatibility (RFC) test, which was a first in Europe. In the early stages of the ERS-1 Programme, this RFC test was the subject of intense debate. Clearly, with active onboard radars producing field strengths of up to 100 V/m in the vicinity of electronics boxes, and with a sensitive microwave radiometer located less than 2 m from the SAR antenna, the question `Would it all work together?' needed to be addressed very seriously.
Calculations had shown that the coupling between receivers and transmitters onboard would be fatally compromised by even the smallest reflections from any building in which the RFC test was conducted. At first, therefore, it was proposed to conduct the test with the deployed satellite lying on its back outdoors, using a giant roll-on/roll-off cover much like those used to protect cricket pitches during rain.
As the date for testing grew closer, the unpleasantness of exposing even Engineering-Model hardware to the elements provoked a re-examination of these plans. A set of reflection measurements were made to decide between testing inside an inflatable plastic tent (like a portable tennis hall) or in a normal EMC anechoic chamber. It was subsequently decided to enlarge the EMC Chamber at Intespace in Toulouse (F) to accommodate the deployed satellite. The test itself was conducted in late 1989; there was no detectable disturbance of the satellite by the transmitters.
The Flight-Model payload was shipped to Matra in March 1990, with the Flight-Model SAR Processor following some six weeks later. The test programme included a further compatibility test between the payload and the onboard software, which had by now completed its validation programme. The various antennas were then installed and aligned. Vibration and acoustic testing were then carried out to demonstrate, successfully, that the ERS-1 satellite would not be adversely affected by the vibration and noise induced by the launch vehicle. This was followed by a deployment test, under onboard software control, of the SAR and Scatterometer antennas, using special `zero-gravity' rigs to simulate a realistic deployment.
In the late summer of 1990 the Flight-Model satellite was de-coupled and the payload subjected to a further three-week thermal- balance/thermal-vacuum test. The primary purpose of this was to validate the in-orbit thermal predictions and to provide a realistic simulation of the in-orbit environment under which the payload will have to operate. This also confirmed the performance of the heat-pipe modification to the ATSR.
This sequence of satellite mechanical and payload thermal tests allowed an intensive post-vibration-test check-out of the platform without affecting the satellite critical path. Finally, in October 1990, the flight platform was delivered to ESTEC for final integrated check-out, and an extended three-week operations test during which the mission control centre at ESOC operated the satellite by modem.
Arianespace had already announced a launch delay until 3 May 1991, so while the EGSE took the slow (but cheaper) route to Kourou by sea, ERS-1 waited in ESTEC. The satellite itself went by air in February 1991. The launch campaign had been extended to 8 weeks in view of the complexity of ERS-1; the normal campaign in Kourou, typically for telecommunications satellites, is 6 weeks. However in the event it turned out to be significantly longer.
With just a few days to go before launch, with the spacecraft fully fuelled with hydrazine and waiting on top of the launch vehicle, Arianespace announced a further, indefinite delay, while anomalies previously observed with the Ariane 4 third stage were further investigated. The spacecraft was ultimately removed from the launch vehicle and put into storage, fully fuelled - it had never been foreseen to prepare a procedure for removing it. The launch was rescheduled for 16 July 1991 and a health test was performed before reintegration with the launcher. This test, which involved measuring the RF field in front of the antennas, was uncomfortably hazardous: the 300 kg of hydrazine on-board was not only highly poisonous but also, pressurised to 30 bar, explosive.
All went well, and ERS-1 was launched at the opening of the 10 minute launch window. The Ariane 44L launcher delivered the satellite to the orbital insertion point with extremely small residual errors, so that only about 10 kg of hydrazine was required for orbit acquisition, a fact which has left ERS-1 with a healthy fuel margin to support a long lifetime.