ISO Background articles

The ISO Spacecraft

S. Ximénez de Ferrán

ISO Project, ESA Directorate for Scientific Programmes, ESTEC, Noordwijk, The Netherlands

Note: This article is an update of an article that originally appeared in ESA Bulletin 67 (August 1991).

ESA's Infrared Space Observatory (ISO) consists of two modules: the Payload module, which includes the telescope and the scientific instruments, and the Service Module, which houses the instruments electronics, the hydrazine propellant tank and all other classical spacecraft subsystems. To ensure that the telescope is kept near absolute zero and thus is the least disturbed by the effects of the infrared emissions from other elements of the system, the telescope is enclosed in a helium-cooled cryostat. The cryostat in turn is shaded by a Sun-shield to protect it from the heat of the direct Sun. The shield has a covering of solar cells that provide the electrical power needed for the mission.


The ISO spacecraft, shown in Figure 1, was conceived as two largely independent modules: the Payload Module (PLM) and the Service Module (SVM). The PLM is essentially a large liquid-helium cryostat, which contains the telescope, with four scientific instruments mounted behind the primary mirror and cooled to a temperature near absolute zero. (The various scientific instruments are described in another article, 'The ISO Scientific Instruments', in this issue.) The SVM houses the warm electronics of the scientific instruments, the hydrazine propellant tank, and all the other classical spacecraft subsystems. The Sun shield, with its covering of solar cells, always faces the Sun to provide electrical power whilst at the same time protecting the PLM from direct insolation.

Overall Configuration
Figure 1. Overall configuration of the Infrared Space Observatory (ISO)

The ISO spacecraft will be placed in a 1000 km 70 600 km elliptical orbit (24 h period) by an Ariane-4 launcher in November 1995. This particular orbit will ensure that most observations can be made during the 16 h per orbit when the satellite is travelling outside the Earth's radiation belts. The spacecraft will be tracked from a main ground station, at which the ISO ground segment will be housed, in Villafranca, Spain, and from a secondary ground station in Goldstone, USA.

Mission requirements and system description

Observations of infrared celestial sources.

The primary objective of ISO is to make observations of celestial objects at infrared wavelengths between 2.5 and 200 microns. It is therefore essential that all radiation-gathering equipment, the telescope and scientific detectors, be protected from the disturbing effects of infrared emissions from various elements of the system itself.

All objects emit radiation as a function of their absolute temperature T. The total energy emitted is proportional to T 4 , and the wavelength at which the radiation's spectral density is at a maximum is inversely proportional to T. The first requisite for ISO, therefore, is to provide a telescope, including baffles, that is kept very cold. The focal-plane units (FPUs) of the scientific instruments and the infrared detectors inside those units must also be maintained at temperatures close to absolute zero. The detailed requirements are summarised in Table 1.

Table 1. Temperature requirements Component Temperature, K Temperature stability, deg Detectors interface 1.7<T<1.9 0.05 in 1000 s Optical Support Structure / Focal-Plane Unit interface 2.4<T<3.4 0.10 in 1000 s Primary mirror <3.2 0.10 in 1000 s Secondary mirror <4 Lower baffle <5 Upper baffle <7.5

The solution adopted for ISO is to enclose the telescope in a cryostat. The main element is a toroidal tank containing 2286 litres of super-fluid helium (HeII) at a temperature of 1.8 K. The tank is insulated from external heat inputs by three vapour-cooled radiation shields (VCS) equipped with multi-layer insulation (MLI). The tank, radiation shields and telescope are suspended from the cryo vacuum vessel (CVV) by low-conductivity straps. Boiling helium from the tank provides cooling to the optical support structure (OSS), the focal-plane units and telescope mirrors mounted on the OSS, the optical baffles and, when flowing through the radiation shields to the exhaust nozzles, intercepts incoming heat from the outside environment. Some of the scientific detectors are directly cooled by copper straps connected to the helium tank. A heat shield connected to the OSS encloses all four focal-plane units and provides a light-tight environment.

The pressure inside the HeII tank is 17 mbar, the equilibrium boiling point at a temperature of 1.8 K. This pressure is maintained in orbit by the impedance of the vent line, and on the ground prior to launch by continuous pumping of the tank exhaust. The gaseous-helium exit is located at the highest point of the tank, allowing separation by gravity of the liquid and gas phases during ground operations. Once in orbit, one of the remarkable properties of superfluid helium is exploited, the so-called 'thermodynamic fountain effect', by which a simple porous plug functions as a phase separator, keeping the liquid phase in the tank while allowing the gaseous helium to flow through the vent line.

A flow diagram for the ISO cryostat is shown in Figure 2.

Helium Flow Diagram
Figure 2. Helium flow diagram for ISO when in orbit

To protect the cryostat from external heat inputs and in particular from direct solar illumination, a Sun shield shades the cryo vacuum vessel. This shield is composed of two flat plates (Fig. 1), the outer faces of which carry the solar cells that provide the electrical power needed for the mission.

During ground operations, the vacuum in the cryo vacuum vessel is maintained by a cryo cover, which is also insulated with radiation shields and multi-layer insulation. It is held in place by a clamp band, which will be released in orbit to jettison the cover after satellite outgassing, approximately 15 days after lift-off, at which point the scientific observation programme will commence.

The cryostat is designed for a minimum operational lifetime of 18 months (calculated nominal lifetime of 20 months).

Optical requirements

Two main sets of optical requirements are imposed on the ISO spacecraft. Firstly, there are the light-gathering requirements, which can be summarised as follows:

Entrance-pupil diameter: 	600 mm
Focal length: 9000 mm
Unvignetted field of view: 20 arcmin
Instrument unvignetted
field of view: 3 arcmin
Wavelength range: 2.5 - 200 microns
Image quality: diffraction limit at 5 microns

These requirements are met with a Ritchey-Chrétien Cassegrain telescope configuration, as this is the best solution for an astronomical telescope that must cover a wide spectral range in combination with a limited field of view. This configuration is free from either coma or spherical aberration.

The ISO telescope has a primary mirror with an overall diameter of 640 mm (Fig. 3), a secondary mirror with a diameter of 87.6 mm, and a four-faced pyramidal mirror that distributes the light collected to the four focal-plane units. The pyramidal mirror has a central hole that allows some of the light to impinge on a quadrant star sensor (QSS). This will allow measurement of the alignment offset between the telescope and the spacecraft's attitude-control sensors.

Primary Mirror
Figure 3. The primary mirror of the ISO telescope

The lightweight mirrors are made of fused silica and are gold-coated to give them good reflection characteristics in the infrared. The primary mirror is circumferentially mounted onto the optical support structure via three fixation devices, each consisting of an invar pad fixed to the mirror and crossing blades that provide the required degrees of freedom. The secondary mirror is mounted on a tripod. Both mirrors are cooled by copper straps connecting their rear faces to the helium-cooled optical support structure.

The second set of requirements relates to the stringent control of stray light emanating from bright infrared sources outside the telescope's field of view, for which the following viewing constraints have been defined:

With these viewing constraints, the total stray light falling on the instruments must be less than 10% of the diffuse zodiacal background. This requirement is fulfilled by means of the main baffle with sharp-edged vanes surrounding the telescope, Cassegrain baffles around the secondary mirror and the central hole of the primary, and a gold-coated truncated-cone Sun shade that reflects direct illumination from the Earth back to space (Fig. 4).

Sun Shade
Figure 4. The ISO Sun shade

Moreover, with the temperature distribution described in the previous section, the infrared self-emission of all optical elements gives a noise background that is also less than 10% of the zodiacal background, at wavelengths from 5 to 200 microns.

Pointing requirements

ISO will be operated in a similar manner to a ground-based observatory, and therefore the spacecraft has to be able to manoeuvre smoothly from one celestial source to the next, and then maintain accurate pointing on that target. The spacecraft must also be capable of pointing at any region of the sky that satisfies the stray-light constraints described above. The slew speed between sights is set at 7 degrees/min in order to optimise observation time, and the duration of each observation can range from a few seconds to up to 10 h, depending on the type of source.

During a scientific measurement, the following telescope optical-axis pointing accuracy is required:
Absolute pointing error: 11.7 arcsec
Absolute pointing drift: 2.8 arcsec/h
Relative pointing error: 2.8 arcsec

These pointing requirements must be satisfied by the spacecraft's Attitude and Orbit Control Subsystem (AOCS), in combination with careful spacecraft structural design, to avoid thermo-elastic deformation between the telescope's optical axis and the attitude sensors. Three operational pointing modes have been defined:

For the high-accuracy pointing modes, the attitude errors are measured with gyroscopes, a star tracker and fine Sun sensors. In the calibration mode (activated nominally once per orbit), the quadrant star sensor replaces the star tracker. The performances of the various sensors are summarised in Table 2.

           Table 2. Performance figures for the various ISO sensors

Star tracker
Field of view           4 x 3 degrees
Sensitivity:            Visual magnitude
                        from +2 to +8
Bias error:             2arcsec
                        (0.5 arcsec in centre of field
                        of view)
Tracking speed:         5arcsec/s

Random drift:           3arcsec/h
Gyro noise:             0.2 arcsec at 2 Hz
Maximum rate:           1deg/s

Fine Sun sensor
Field of view per slit: 62 x 1 degree
Accuracy:               3 arcmin
                        (1 arcmin in centre of field of view)
Noise equivalent angle: 2 arcsec

A state-reconstructor in the AOCS computer produces minimum-variance estimates for the attitude, angular velocity and disturbance acceleration. This state-reconstructor also serves as a sensor-data smoothing filter.

The control torques for high-performance slews and pointing modes are provided by a reaction-control wheel system, giving a maximum torque of 0.2 Nm, with a total of 126 torque levels, and a maximum angular-momentum storage capability of some 18 Nms. A so-called 'dual control law' is used together with a velocity controller that limits angular velocities to 8 deg/min. The 'dual control law' consists of a nonlinear time-optimal subcontroller and a linear state feedback subcontroller. For large errors during slewing, the time-optimal control prevails, whereas for fine pointing, the linear law predominates.

A schematic of the AOCS is shown in the picture below. The modes of operation include, besides the pointing modes, other functions related to safety, satellite autonomy, and health checking of the subsystem elements, some of which are addressed below.

Schematic of AOCS
Figure 5. Schematic of ISO's Attitude and Orbit Control Subsystem (AOCS)

An important factor in achieving the requisite pointing accuracy for the ISO spacecraft is the limiting of the drift between the optical axis of the telescope and that of the star tracker. Such drift can be induced by transient thermo-elastic deformation of structural elements linking the two optical axes. Consequently, the star-tracker support structure has been mounted on the cryostat's outer wall, rather than on the Service Module. This alone does not prevent local deformation due to temperature gradients in the cryo vacuum vessel from degrading pointing performance. It is also necessary to maintain a stable and uniform temperature distribution in these two structures.

This temperature stability is achieved by covering the cryo vacuum vessel with multi-layer insulation, even at the expense of a penalty in the lifetime of the satellite. In addition, the star-tracker sensors (two for redundancy) are enclosed within a thermal housing, with heaters, which should provide a constant sensor temperature and, even more importantly, maintain a constant temperature gradient between the mounting feet of the operational star tracker (better than 0.1 degC over one orbit, except for 2 h around perigee). The specially stiffened fixing of the housing to the cryo vacuum vessel ensures that the thermal conductance between the two is less than 3 mW/degC.

Autonomy safeguards

A further top-level requirement stems from the type of orbit that the ISO spacecraft will be in. The highly elliptical 1000 km 70 600 km orbit makes spacecraft tracking and control impossible during perigee passage.

It is also imperative that the ISO spacecraft should be able to survive a possible failure of the ground stations. The system must therefore ensure that the satellite is safe and that its cryogenic helium is not wasted during any such event. Moreover, even small heat inputs to the optical elements during perigee passages could disturb the telescope's ther-mal equilibrium.

To meet these mission safety requirements, the launch window selected ensures that there is always a region of the sky to which the spacecraft can point without violating the stray-light constraints (the Sun and the Earth; the Moon and Jupiter do not affect the thermal equilibrium). Autonomous functions onboard the satellite would prevent violation of the constraints for at least three orbits in the event that ground control is lost.

Several souces could trigger these autonomous functions:

Two main functions are active during the period of autonomy:

Electromagnetic compatibility

The very weak output signals from the scientific detectors have to be protected from electronic noise, whether conducted or radiated, originating from other elements of the satellite. A stringent spacecraft design requirement, verified by testing at unit, subsystem and satellite levels, ensures full electromagnetic compatibility (EMC) between the onboard instruments and other subsystems.

An important element in this clean EMC design is the spacecraft's power subsystem. It is based on a sequential switching shunt regulator, which provides very good efficiency and reliability as well as a low output impedance and constant bus ripple under all load conditions.

A further requirement is that the spacecraft's external surfaces must not become electrically charged and no electrostatic discharges should take place. Special care has therefore been taken in the design of the thermal-control hardware, which could be susceptible to charging, by making the outside surfaces conductive and grounding all elements to the structure.

Another developmental challenge lay in designing the optical coatings for the baffling system, which have to preclude any electro-static charging close to the focal-plane units and provide a high absorptivity in the far-infrared. A special conductive black paint has been produced for this purpose.

Data collection and transmission

ISO is a real-time mission. The operating principle involves having a single scientific instrument, selected as a function of the type of observation in hand, active at any given moment. An exception to this philosophy is the ISOCAM instrument, which could be active in parallel at low bit rate to provide images of the celestial source under study.

For this reason, four different data formats - one per instrument - can be selected. Each of these formats also includes all of the satellite housekeeping information, the largest part of which is devoted to the AOCS parameters needed for accurate spacecraft attitude reconstitution.

The nominal data rate is 32 768 bits/s, 23 424 bits of which are allocated to the prime instrument. Communications with the ground are via a transponder working at S-band.

Other special features of ISO

Heat balance of the Payload Module

As noted above, heat inputs to the helium tank have been minimised by enclosing it with vapour-cooled shields that intercept ambient heat before it can reach the cryogenics. Another important source of heat is the Service Module, the average temperature of which will be 20 degC when the Payload Module's outer wall is at -150 degC.

To avoid such a temperature gradient causing a net conductive heat transfer to the vessel, the interface between the two Modules consists of 16 tubular glass-fibre struts with very low thermal conductivity. They are filled with Ecofoam resin to prevent radiative heat transfer from taking place inside the tubes. The development programme for these struts has included extensive mechanical and thermal validation testing.

Direct liquid-content measurement

An important factor for the planning of ISO's scientific operations is accurate knowledge of the amount of superfluid helium (HeII) remaining in the tank. The ability to make this measurement under microgravity conditions is a novel development for ISO, which relies on the near-infinite thermal conductivity of the superfluid helium. A calibrated heat pulse is introduced into the tank, which increases the temperature of the helium by an amount directly proportional to the mass remaining.

Launch operations

An important aspect of the Payload Module is the cryogenic operations required prior to launch. The superfluid-helium tank will be topped off when the satellite is already mounted on the launcher, by removing the Ariane fairing (a non-standard operation). The tank will then be closed to minimise helium loss and to avoid having to pump. To maintain the insulation performance after this operation, a second reservoir containing 60 litres of normal liquid helium (HeI) will be used to cool the radiation shields. This HeI tank, which can be accessed through windows in the Ariane fairing, will be completely depleted prior to lift-off.

During the launcher's flight, commands issued by Ariane's electronics will operate a set of cryogenic valves that will open the helium vent line to space, and also the main helium tank and its porous-plug phase separator. Initially, the vented helium mass flow rate will be about 20 mg/s, rising to a peak of about 27 mg/s and then falling until, after about 20 days in orbit, it will be about 5 mg/s, the in-orbit equilibrium point.

To cope with this range of flow rates, the system is equipped with two sets of nozzles. Initially, both will be open to accommodate the high mass flow rate: as the rate falls and the temperatures decrease, the larger nozzles will be valved off, leaving only the smaller set open.

Development plan

The development plan for ISO was based on two independent development and integration activities for the PLM and the SVM. A system integration phase followed, during which the two modules were mated and the system tests (mechanical, acoustic, electromagnetic compatibility, functional and thermal tests) were carried out.

The development of the PLM was a challenging task, involving more laboratory experimental effort than a classical qualification procedure, and requiring inventive solutions to unforeseen problems. The operation of the cryostat was demonstrated on a qualification model, first at room temperature and later in a vacuum chamber that simulated the in-orbit environment. The tests revealed that some elements of the external vent line had too great a pressure drop. Also, the filling port, together with the helium subsystem, indicated that there was a possibility of thermo-acoustic oscillations due to unwanted heat transport. The mechanical tests of the model showed that the main helium tank was not rigid enough. (Later, a non-negligible interaction between the compressible liquid helium and the structure was also found.)

The telescope underwent its own qualification programme, and a special facility was developed at the Centre Spatial de Liège in Belgium to test it optically at cryogenic temperatures. The tests demonstrated the correct behaviour of the optics in terms of both image quality and alignment of the mirrors with respect to the scientific instrument focal plane units. The same facility was used to test the flight model of the telescope, using tests similar to those for the qualification model.

After rectification of the problems encountered during the qualification programme, and a re-development of the cryogenic helium valves to improve the leak characteristics, the PLM flight model was built. After completing extensive testing, it was delivered in June 1995 for system integration.

The SVM was first qualified thermally and mechanically on a structure/thermal model. In parallel, the spacecraft's electrical subsystem was developed and qualified at unit/subsystem level. The more challenging unit was the star tracker, with its stringent requirements. However, it represents today the state of the art for that type of sensor. Also, the gyroscopic units used by the AOCS were given special attention, and problems encountered on other satellite programmes were resolved. The final integration of the SVM flight model was finished at the end of 1993.

The system-level integration and test phase started upon the arrival of the PLM at ESTEC (Noordwijk, The Netherlands). It was mated with the SVM. The system was functional-tested successfully and the electromagnetic compatibility was established. Afterward, the satellite underwent a series of mechanical/dynamic tests at proto-flight level to verify its behaviour and compatibility with the launcher, the Ariane 44P.

The final environmental tests were the thermal balance/thermal vacuum tests at ESTEC's Large Space Simulator (LSS). These tests confirmed the correct functioning of the cryostat, in particular the transient phase after launch, and the lifetime of ISO, and the adequacy of the SVM thermal control. The final settings of the active thermal control of the star tracker were established based on the data gathered during these tests.

The final activities prior to shipment were dedicated to proving the compatibility with the control centre in Villafranca and to developing the operational procedures. The satellite was then shipped to the launch site (CSG) in Kourou, French Guyana, in June 1995.

The ISO satellite was functionally tested upon arrival at CSG and, at the end of July, was declared ready for flight. Final preparations - refilling the Helium tank and filling the propulsion tank with hydrazine-will now follow as the countdown for the launch in November 1995 nears.

Payload Module Undergoing Tests
Figure 6. ISO Payload Module undergoing testing at DASA, Ottobrun