The Ulysses Spacecraft

In the following, you will find concise descriptions of the various subsystems that together make up the Ulysses Spacecraft. More detailed technical information, including a large number of illustrations, is available at the Mission Operations section.

[ System Configuration | Structure and Mechanisms | Performance ]

Ulysses at ESTEC


System Configuration

The overall design of the Ulysses spacecraft is dictated by the large distances from the Earth and the Sun at which the spacecraft has to operate (up to 950 million km from Earth, 800 million km from the Sun). Spin-stabilized (at 5 rpm), the spacecraft's main elements are the box-like main body structure on which is mounted the large-diameter (1.65 m), Earth-pointing High-Gain Antenna (HGA) that provides the communication link, and the Radioisotope Thermoelectric Generator (RTG) that supplies the spacecraft's electrical power. The spacecraft mass at launch was 367 kg including 55 kg of payload and 33.5 kg of hydrazine for orbit, attitude and spin rate adjustments.

Schematic of Ulysses
Schematic of the Ulysses spacecraft.

A 5.6 m long radial boom is mounted on the opposite side of the spacecraft to the RTG and carries several experiment sensors. A 72.5 m tip-to-tip dipole wire boom and a 7.5 m axial boom serve as electrical antennas for the Unified Radio and Plasma Wave (URAP) Experiment. Most of the scientific instruments are mounted on the main body, positioned as far as possible from the RTG and in compliance with the field-of-view requirements of the experiment sensors.

Ulysses showing radial boom in flightconfiguration
Ulysses in Hangar-AO at the Cape Canaveral Air Force Station prior to being moved to the Kennedy Space Center. The radial boom is extended in flight configuration.

The radio link between the spacecraft and Earth is achieved via the HGA with 20 W X-band and 5 W S-band transmitters. The X- and S- band downlinks can both be modulated for telemetry and ranging. The spacecraft spin axis in deployed configuration is aligned with the electrical axis of the HGA.

Maximum data coverage throughout the mission is a prime scientific requirement. Since continuous coverage by ground stations is impossible for such a long-duration mission, data are stored on-board and replayed, interleaved with real-time data, during periods of ground contact. The nominal tracking period is 10 h in every 24 h. A variety of downlink bit rates up to 8192 bit/s is selectable, which can provide real-time data rates between 64 and 1024 bit/s and stored data rates between 128 and 512 bit/s. The prime data rates are 1024 bit/s for real-time ("tracking mode") and 512 bit/s for stored data ("storage mode").

A key scientific requirement was to have an electromagnetically and electrostatically clean spacecraft; EMC considerations have therefore driven the mechanical configuration design. The spacecraft is divided into a "quiet" and "noisy" zone. The former comprises an electromagnetically shielded compartment of sensitive experiments, whereas the latter contains the less susceptible - but more emissive - electrical spacecraft subsystems. The preamplifiers for the wire booms and the axial boom are mounted outside the main spacecraft compartment, which acts as a Faraday cage against fields generated inside. A unipoint grounding concept has been implemented in which the main platform constitutes an electrical ground reference with only one area which serves as grounding starpoint for the electrical system. All units that produce significant magnetic fields are removed as far as possible from the magnetometers. Low-impedance ground bus bars are isolated from the platform and connected to it only at the starpoint. Strict control of the magnetic properties of all subsystems and experiments was exercised. For example, the RTG was magnetically compensated.

The electrostatic cleanliness requirement for low-energy particle measurements has been achieved by making all external surfaces of the spacecraft electrically conducting. This measure also prevented differential charging of parts of the spacecraft in the Jovian magnetosphere.

The Jupiter gravity assist necessitated the spacecraft's passage through the Jovian radiation belts. All subsystems and experiments have therefore been designed to survive this environment and radiation-hardened parts (design dose rate 60 krad) have been used throughout the spacecraft.


Structure and Mechanisms

The Ulysses spacecraft has a box-type structure with two overhanging "balconies" and a single aluminium honeycomb equipment platform. All electronics units of the scientific instruments and spacecraft subsystems, most of the sensors and the propellant tank are mounted on this platform. The RTG is mounted on an outrigger structure to minimise its radiation effects and to isolate the main subsystems and the experiments from excess heat.

The two-section radial boom carries the two magnetometer sensors, the solar X-ray/cosmic gamma-ray burst sensors and the magnetic search-coil sensor of the wave experiment. Because of the radiation pattern of the RTG, it was necessary to have the gamma-ray sensor lying as closely as possible along the RTG centre axis. The boom design allows for this, at the same time providing the maximum boom-length consistent with a two-hinge system and also satisfying the spacecraft balance constraints in both the stowed and deployed configurations. The boom is made from carbon-fibre-reinforced plastic (CFRP) tubing 50 mm in diameter and with 1 mm wall thickness.

The electrical antennas of the URAP wave experiment consist of a pair of radially extending wire booms in the spin plane and an axial boom deployed along the orbital spin axis. The wire booms consist of 5 mm wide and 0.04 mm thick Cu-Be ribbon stowed during launch on two identical drive units. The wires were deployed to a length of 72.5 m tip-to-tip by centrifugal forces acting on tip masses after the second trajectory correction manoeuvre (TCM-2). Each wire boom has a passive tubular root damper which reduces relative motions between the boom and the spacecraft by natural material damping with a time constant of 3.5 h. The axial boom element is formed by a pre-stressed, coilable elastic Cu-Be tube anchored in the axial-boom drive mechanism located on the rear face of the spacecraft. The boom element was deployed to a length of 7.5 m by a traction force through a set of rollers driven by a stepper motor one day after the wire booms.

Thermal Subsystem

Thermal control of the spacecraft, its subsystems and of most of the experiments is achieved by passive means in conjunction with a commandable internal/external power dump and heater system. This involves an optimised layout of subsystems which avoids hot spots on the spacecraft platform, an efficient thermal-blanket design in order to minimise the solar input, the compensation, by the power dump system, of heat fluxes which are caused by the varying solar input and a heater system for individual critical units. The most stringent requirements on the thermal subsystem are to guarantee a temperature above +5 deg C at all times for the hydrazine of the Attitude and Orbit Control Subsystem (AOCS) and a temperature below +35 deg C for all experiment solid-state detectors. All spacecraft walls are covered with thermal multilayer blankets, which are closely fitted around the experiment-sensor apertures. The blankets consist typically of 20 layers of aluminized mylar. The outermost layer is kapton, coated with a transparent conductive coating (Indium Tin Oxide) to provide an electrically conductive outer spacecraft surface. Heat rejection is performed by a thermal radiator, located on the rear of the spacecraft and covered by a 2 mil kapton foil. All units external to the spacecraft (e.g. several experiments) are thermally decoupled from the interior.

RTG and Power Subsystem

Electrical power is provided by the RTG at a level of about 280 W at the beginning of the missing, decreasing to about 220 W at nominal mission end (December 2001). The RTG, which generates 4500 W of thermal energy, has two major components: a heat source and a converter. The General-Purpose Heat Source (GPHS) consists of several elements containing the isotopic fuel 238 Pu, in the form of PuO2. The radioactive decay energy is absorbed at the heat source- converter interface where heat is produced. The Si-GE converter contains thermoelectric elements which convert the heat into electrical energy. Power is delivered to the experiments and subsystems at 28V +/- 2%.

Communication Subsystem

The communication subsystem provides capabilities for telemetry with bit rates up to 8 kbit/s, ranging, telecommand and radio science. It operates in X-band (downlink) and S-band (up- and downlink). The subsystem includes two redundant transponders (each consisting of an X-band exciter, a modulator, an S-band receiver and an S-band power amplifier), two redundant 20 W X-band Travelling-Wave-Tube Amplifiers (TWTA), a TWTA Interface Unit and an S-band Radio-Frequency Distribution Unit. A considerble amount of cross-coupling capability exists within the subsystem.

The parabolic HGA, with both X-band (8.4 GHz) and S-band (2.3 GHz) capabilities, is the prime communications link. Telemetry is provided in X-band, with a 2 deg beamwidth (3 dB); downlink S-band is used for ranging, and radio- science investigations. S- or X-band ranging operations can be performed with or without telemetry transmission. Both transponders can be operated simultaneously, one in X-band and the other in S-band.

A special feature of the HGA is its ability to measure the offset of the spin axis from the direction of the ground station by the CONSCAN (conical scan) system. This is accomplished by a tilt of 1.8 deg between the S-band antenna pattern and the spin axis which results in a spin modulation in the uplink signal strength as the satellite rotates. Processing within the Attitude and Orbit Control Subsystem (AOCS) gives the offset magnitude and direction which is either transmitted for ground analysis or employed in a closed loop control system to minimise the offset. Attitude adjustments are made by operating hydrazine thrusters.

Command and Data-Handling Subsystem

The command and data-handling subsystem provides capabilities for ground commanding, a variety of telemetry formats, on-board data storage, and, in combination with the AOCS, safe automatic manoeuvring.

The telecommand decoder checks commands for validity and distributes them to the experiments and subsystems. There are directly executed commands and memory-load commands. The latter are stored as block commands that are validated prior to execution of critical operations. A command-time-tagging capability over a range of 32 s to 24 d is also available.

The Central Terminal Unit (CTU) processes command messages received from the decoder, provides on-board timing information, and performs formatting and encoding of data to be sent to the ground. It also controls all on-board automatic functions. The CTU contains a provision to auto-check its own functioning and to switch over to the redundant CTU in the event that a failure is detected, assuring that important spacecraft information is maintained.

The CTU contains a master crystal oscillator from which all synchronisation and timing signals for subsystems and payload are derived. 32-bit timing information with a resolution of 2 s is included in every telemetry format, ensuring unambiguous identification of the telemetry data throughout the mission lifetime. A spin reference ("Sun pulse") and spin segment clock (16 384 pulses per spin period) are also supplied by the CTU, based on Sun-sensor data provided by the AOCS subsystem.

The Data Storage Units consist of two redundant tape recorders for storage of data during those periods when the spacecraft is not in communication with the gorund (nontracking periods) for subsequent playback and transmission during the next tracking period. The 45 Mbit capacity of each tape recorder is sufficient to provide continuous storage at 512 bit/s for 22 h or 256 bit/s for 44 h.

Telemetry formats are built up of successive frames. There are three data formats:

  • Scientific format consisting of 32 scientific frames
  • Interleaved forat consisting of a block of 32 frames, interleaving real-time and stored frames with a selectable ratio (1:1, 1:3, 1:7). Formats are played back in reverse time order, but the frames within each format and each individual frame are in forward order
  • Engineering format consisting of two frames of spacecraft housekeeping data and containing no scientific data

Telemetry channels are sampled in a time-ordered fashion and allocated to specific words (8 bit), which are arranged into frames of 128 words. In the scientific and interleaved format one frame consists of 110 digital science words, nine analogue science words, two subcommutated experiment housekeeping words and three synchronisation and identification words. Analogue channels are sampled and converted into 8-bit words with an accuracy of 1% full scale. There are also datation channels which contain accurate time information on an event with a resolution of 0.488 ms (32 s range) or 3.9 ms (256 s range). Datation channels are used by the gamma ray-burst instrument (high resolution) and by the magnetometer and wave experiments (low resolution).

Attitude and Orbit Control Subsystem (AOCS)

The primary operational functions of the AOCS are to maintain the spacecraft spin axis Earth-pointing and control the spin rate. Additional operational functions are dictated by trajectory control requirements, nutation damping, and by the measurement of the attitude for scientific reasons.

The spacecraft Earth-pointing attitude is measured and controlled by the CONSCAN system with spin-rate, spin-phase and solar-aspect-angle information determined from redundant Sun sensors. The sun-sensor outputs are processed in the data-handling subsystem to provide the spin reference pulse and the spin segment clock. The signals and the Sun-sensor data are then used in the AOCS electronics to determine the spacecraft spin rate and solar aspect angle for the purpose of closed loop on-board control, failure detection and recovery. Hydrazine thrusters are actuated either by telecommand or automously. These are fed from a single tank, mounted at the launch centre of gravity, and arranged in two blocks of four thrusters each, providing complete redundancy.

Another AOCS operation is the periodic precession manoeuvring for correction of the apparent Earth drift with respect to the spin axis. These can be performed in closed loop on-board or in open loop via ground or time- tagged command.

A special manoeuvre strategy is required for conjunctions, since the proper spacecraft attitude depends on the operation of the Sun sensors with a safe operational limit of the solar aspect angle greater than 1.25 deg.

The spacecraft carries three nutation dampers containing a fluid whose viscous property dissipates energy. In the operational spin rate range the nominal damping time constant for nutation cone angles from 2.0 deg to 0.02 deg. is less than one hour.

The AOCS also includes failure-mode-detection and protection functions which result in fail-safe operation and a reacquisition capability in both automatic and ground initiated recovery sequences. The AOCS also provides autonomous system capabilities for safe spacecraft reconfiguration. This is required during unexpected and/or predicted periods of nontracking and because of the long signal travel time between ground and spacecraft. The preprogrammable functions include search-mode initiation to reacquire the Earth if no command is received after a preselectabe time of up to 30 d, switch-over to redundant units, and preprogrammed attitude manoeuvres at superior conjunctions.


In-Flight Performance

Following launch and orbit injection of Ulysses on 6 October 1990, the initial flight phase consisted of checking out all spacecraft subsystems (including redundancy), deploying the radial, axial and wire booms and switching on all experiments. The latter were commanded on one-by-one and thoroughly checked over an extended period between 19 October and 16 November 1990. Early in January 1991 the spacecraft was formally declared to be commissioned.


The following is a list of (major) spacecraft anomalies to have occurred since launch:

  • Nutation. Shortly after deployment of the axial boom, build-up of a nutation-like disturbance was observed. This is now believed to be the result of an oscillation induced by non-uniform solar heating of the axial boom coupling into the spacecraft motion, together with under-performance of the passive nutation dampers on board the spacecraft. The onboard CONSCAN system has been successfully employed to control subsequent episodes of nutation, which occurred in 1994 and 1995. Nutation is predicted to return in 2001.
  • Disconnect Non-Essential Loads (DNEL). The DNEL condition is a spacecraft safing mode, and is known to be associated with operation of the latching valve when coinciding with unpredictable peaks in payload current demand. Overcurrent criteria are violated, and the onboard protection logic correctly operates, placing the spacecraft in a minimum current demand mode by disconnecting the scientific payload. To date, 8 DNEL events have occurred.
  • CTU-2 Anomaly. During check-out of the Data Handling Subsystem following Jupiter flyby, telemetry from the redundant Central Terminal Unit (CTU-2) was found to be partially corrupted. CTU-2 is a redundant unit, however, and will only be used in case of CTU-1 failure. If this occurs, extra data processing will be implemented to minimise the impact of the CTU-2 malfunction.