PDS_VERSION_ID = PDS3 OBJECT = INSTRUMENT_HOST INSTRUMENT_HOST_ID = HP OBJECT = INSTRUMENT_HOST_INFORMATION INSTRUMENT_HOST_NAME= "HUYGENS PROBE" INSTRUMENT_HOST_TYPE= "SPACECRAFT" INSTRUMENT_HOST_DESC= " Instrument Host Overview ======================== [From JONES&GIOVAGNOLI1997]: The Huygens Probe is the ESA-provided element of the joint NASA/ESA Cassini/ Huygens mission to Saturn and Titan, the planet's largest moon. The industrial Phase B activities began in January 1991 under the leadership of Aerospatiale, the Huygens prime contractor. The Probe was carried to Titan by the Cassini Saturn Orbiter. Huygens was dormant during the interplanetary journey of 6.7 years, although it was activated about every 6 months to verify and monitor its health. It was released 22 days before the Titan encounter. The Probe's aeroshell decelerated it in less than 3 min from the entry speed of about 6 km/s to 400 m/s (Mach 1.5) by 150-180 km altitude. From that point onwards, a pre-programmed sequence triggered parachute deployment and heatshield ejection. The main part of the scientific mission then started, lasting for the whole descent of 2-2.5 h. The Huygens model philosophy was optimised to achieve the most complete verification possible that the Probe system meets the mission requirements within the cost envelope and the tight schedule constraints imposed by the launch window. Four models were developed at system level: 1. Structural, Thermal & Pyro Model (STPM): to qualify the Probe design (including all mechanisms activated by pyrotechnic devices) for all structural, mechanical and thermal requirements; 2. Electrical Model (EM): to verify the electrical performances of the Probe and of the electrical/functional interfaces with the Orbiter; 3. Special Model (SM2): used for the balloon drop test in May 1995. All the mechanisms and the descent control systems were flight-standard; 4. Flight Model (FM). Overall Configuration --------------------- The Probe System comprises two principal elements: 1. the 318 kg Huygens Probe, which enters Titan's atmosphere after separating from the Saturn Orbiter; 2. the 30 kg Probe Support Equipment (PSE), which remains attached to the Orbiter after Probe separation. The Probe consists of the Entry Assembly (ENA) cocooning the Descent Module (DM). ENA provides Orbiter attachment, umbilical separation and ejection, cruise and entry thermal protection, and entry deceleration control. It is jettisoned after entry, releasing the Descent Module. The DM comprises an aluminium shell and inner structure containing all the experiments and Probe support subsystems, including the parachute descent and spin control devices. The PSE consists of: 1. four electronic boxes aboard the Orbiter: two Probe Support Avionics (PSA), a Receiver Front End (RFE) and a Receiver Ultra Stable Oscillator (RUSO); 2. the Spin Eject Device (SED); 3. the harness (including the umbilical connector) providing power and RF and data links between the PSA, Probe and Orbiter. Front Shield Subsystem (FRSS) ----------------------------- The 79 kg, 2.7 m diameter, 60 degree half-angle coni-spherical Front Shield will decelerate the Probe in Titan's upper atmosphere from about 6 km/s at entry to a velocity equivalent to about Mach 1.5 by around 160 km altitude. Tiles of AQ60 ablative material, a felt of silica fibres reinforced by phenolic resin, provide protection against the 1 MW/m2 thermal flux. The shield is then jettisoned and the Descent Control Subsystem (DCSS) is deployed to control the descent of the DM to the surface. The FRSS supporting structure is a CFRP honeycomb shell which also provides some DM thermal protection during entry. The AQ60 tiles are attached to the CFRP structure by adhesive CAF/730. Prosial, a suspension of hollow silica spheres in silicon elastomer, is sprayed directly on to the aluminium structure of the FRSS rear surfaces, where fluxes are ten times lower. Back Cover Subsystem (BCSS) --------------------------- The Back Cover protects the DM during entry, ensures depressurisation during launch and carries multi-layer insulation (MLI) for the cruise and coast phases. As it does not have stringent aerothermodynamic requirements, it is a stiffened aluminium shell of minimal mass (11.4 kg) protected by Prosial (5 kg). It includes: an access door for late integration and forced-air ground cooling of the Probe; a break-out patch through which the first (drogue) parachute is fired; a labyrinth sealing joint with the Front Shield, providing a non-structural thermal and particulate barrier. Descent Control Subsystem (DCSS) -------------------------------- The DCSS controls the descent rate to satisfy the scientific payload's requirements, and the attitude to meet the requirements of the Probe-Orbiter RF data link and of the descent camera's image-taking. The DCSS is activated nominally at Mach 1.5 and about 160 km altitude. The sequence begins by firing the Parachute Deployment Device (PDD) to eject the pilot 'chute pack through the Back Cover's break-out patch, the attachment pins of which shear under the impact. The 2.59 m diameter Disk Gap Band (DGB) pilot 'chute inflates 27 m behind the DM and pulls the Back Cover away from the rest of the assembly. As it goes, the Back Cover pulls the 8.30 m diameter DGB main parachute from its container. This canopy inflates during the supersonic phase to decelerate and stabilise the Probe through the transonic region. The Front Shield is released at about Mach 0.6. In fact, the main parachute is sized by the requirement to provide sufficient deceleration to guarantee a positive separation of the Front Shield from the Descent Module. The main parachute is too large for a nominal descent time shorter than 2.5 h, a constraint imposed by battery limitations, so it is jettisoned and a 3.03 m diameter DGB stabilising parachute is deployed. All parachutes are made of Kevlar lines and nylon fabric. The main and stabiliser 'chutes are housed in a single canister on the DM's top platform. Compatibility with the Probe's spin is ensured by incorporating a swivel using redundant low-friction bearings in the connecting riser of both the main and stabiliser 'chutes. Separation Subsystem (SEPS) --------------------------- SEPS provides: mechanical and electrical attachment to, and separation from, the Orbiter; the transition between the entry configuration ('cocoon') and the descent configuration (DM under parachute). The three SEPS mechanisms are connected on one side to Huygens' Inner Structure (ISTS) and on the other to the Orbiter's supporting struts. As well as being the Probe-Orbiter structural load path, each SEPS fitting incorporates a pyronut for Probe-Orbiter separation, a rod cutter for Front Shield release and a rod cutter for Back Cover release. Within SEPS, the Spin Eject Device (SED) performs the mechanical separation from the Orbiter: - three stainless steel springs provide the separation force - three guide devices, each with two axial rollers running along a T-profile helical track, ensure controlled ejection and spin, even in degraded cases such as high friction or a weak spring - a carbon fibre ring accommodates the asymmetrical loads from the Orbiter truss and provides the necessary stiffness before and after separation - three pyronuts provide the mechanical link before separation. In addition, the Umbilical Separation Mechanism of three 19-pin connectors, which provide Orbiter-Probe electrical links, is disconnected by the SED. Inner Structure Subsystem (ISTS) -------------------------------- The ISTS provides mounting support for the Probe's payload and subsystems. It is fully sealed except for a vent hole of about 6 cm2 on the top, and comprises: - the 73 mm thick aluminium honeycomb sandwich Experiment Platform; supports the majority of the experiments and subsystems units, together with their associated harness - the 25 mm thick aluminium honeycomb sandwich Top Platform; supports the Descent Control Subsystem and Probe RF antennas, and forms the DM's top external surface - the After Cone and Fore Dome aluminium shells, linked by a central ring - three radial titanium struts; interface with SEPS and ensure thermal decoupling, while three vertical titanium struts link the two platforms and transfer the main parachute deployment loads - 36 spin vanes on the Fore Dome's periphery; provide spin control during descent - the secondary structure; for mounting experiments and equipment. Thermal Subsystem (THSS) ------------------------ While the PSE is thermally controlled by the Orbiter, the Probe's THSS must maintain all experiments and subsystem units within their allowed temperature ranges during all mission phases. In space, the THSS partially insulates the Probe from the Orbiter and ensures only small variations in the Probe's internal temperatures, despite the incident solar flux varying from 3800 W/m2 (near Venus) to 17 W/m2 (approaching Titan after 22 days of the coast phase following Orbiter separation). Probe thermal control is achieved by: - MLI surrounding all external areas, except for the small 'thermal window' of the Front Shield - 35 Radioisotope Heater Units (RHUs) on the Experiment and Top Platforms continuously providing about 1 W each even when the Probe is dormant - a white-painted 0.17 m2 thin aluminium sheet on the Front Shield's forward face acting as a controlled heat leak (about 8 W during cruise) to reduce sensitivity of thermal performances to MLI efficiency. The MLI is burned and torn away during entry, leaving temperature control to the AQ60 high-temperature tiles on the Front Shield's front face, and to Prosial on the Front Shield's aft surface and on the Back Cover. During the descent phase, thermal control is provided by foam insulation and gas-tight seals. Lightweight open-cell Basotect foam covers the internal walls of DM's shells and Top Platform. This prevents convection cooling by Titan's cold atmosphere (70 K at 45 km altitude) and thermally decouples the units mounted on the Experiment Platform from the cold aluminium shells. Gas-tight seals around all elements protruding through the DM's shell minimise gas influx. In fact, the DM is gas tight except for a single 6 cm2 hole in Top Platform that equalises pressure during launch and descent to Titan's surface. Electrical Power Subsystem (EPSS) ================================= Description ----------- The EPSS consists of: - Five batteries: Provide mission power from Orbiter separation until at least 30 min after arrival on Titan's surface. Each battery comprises two modules of 13 LiSO2 (7.6 Ah) cells in series. - Power Conditioning & Distribution Unit (PCDU): Provides the power conditioning and distribution to the Probe's equipment and experiments via a regulated main bus, with protection to ensure uninterrupted operations even in the event of single failure inside or outside the PCDU. During the cruise phase, the Probe is powered by the Orbiter and the PCDU isolates the batteries. The five interface circuits connected to the Orbiter's Solid State Power Switches (SSPSs) provide Probe-Orbiter insulation and voltage adaptation between the SSPS output and the input of the PCDU's Battery Discharge Regulator (BDR) circuits. The BDRs condition the power from either the Orbiter or the batteries and generate the 28 V bus, controlled by a centralised Main Error Amplifier (MEA). The distribution is performed by active current limiters, with the current limitation adapted for each user and with ON/OFF switching capability. The Mission Timer, however, is supplied by three switchable battery voltage lines through series fuses or, when the PCDU is powered by the Orbiter, by dedicated output voltage lines of the Orbiter interface circuits. The PCDU also provides a protected +5 V supply used by the Pyro unit to generate the bi-level status telemetry of the selection relays and for the activation circuit that switches ON the Pyro unit's energy intercept relay. - Pyro Unit (PYRO): Provides two redundant sets of 13 pyro lines, directly connected to the centre taps of two batteries (through protection devices), for activating pyro devices. Safety requirements are met by three independent levels of control relays in series in the Pyro Unit, as well as active switches and current limiters controlling the firing current. The three series relay levels are: energy intercept relay (activated by PCDU at the end of the coast phase); arming relays (activated by the arming timer hardware); selection relays (activated by Command and Data Management Unit, CDMU, software). In addition, safe/arm plugs are provided on the unit itself for ground operations. Operational modes ----------------- - Cruise phase: The EPSS is completely OFF over the whole cruise phase, except for periodic checkout operations. There is no power at the Orbiter interface and direct monitoring by the Orbiter allows verification that all the relays are open. - Cruise phase checkout: The EPSS is powered by the Orbiter for cruise checkout operations. The 28 V bus is regulated by the EPSS BDRs associated with each Orbiter SSPS; a total of 210 W is available from the Orbiter and all the relays are open. - Timer loading: Following the loading (from the Orbiter) of the correct coast time duration into the Mission Timer Unit, battery depassivation is performed to overcome any energy loss due to ageing during cruise. Before Probe separation, the EPSS timer relays are closed to supply the Mission Timer from the batteries and the Orbiter power is switched off. - Coast phase: Only the Mission Timer is supplied by batteries through specific timer relays during the coast phase. The EPSS is OFF and all other relays are open. - End of coast phase - Probe wake-up: At the end of the coast phase, the Mission Timer wakes the Probe by activating the EPSS. Input relays are closed and the current limiters powering the CDMU are automatically switched ON as soon as the 28 V bus reaches its nominal value (other current limiters are initially OFF at power up). The pyro energy intercept relay is also automatically switched on by a command from the PCDU. - Entry and descent phases: All PCDU relays are closed and the total power (nominal 300 W, maximum 400 W) is available on the 28 V distribution outputs to subsystems and equipment. The Pyro Unit performs the selection and the firing of the squibs, activated by CDMU commands. Command & Data Management Subsystems (CDMS) =========================================== The data handling and processing functions are divided between the Probe Support Equipment (PSE) on the Orbiter and the CDMUs (part of the CDMS) in the Probe. The Probe Data Relay Subsystem (PDRS) provides the RF link function for this purpose, together with the data handling and communication function with the Orbiter's Control and Data Subsystem (CDS) via a Bus Interface Unit (BIU). (During the ground operations and cruise phase checkouts, the Orbiter-Probe RF link is replaced by umbilical connections.) The CDMS has two primary functions: autonomous control of Probe operations after separation; management of data transfer from the equipment, subsystems and experiments to the Probe transmitter for relay to the Orbiter. For these functions, the CDMS uses the Probe On-board Software (POSW), for which it provides the necessary processing, storage and interface capabilities. The driving requirement of the CDMS design is intrinsic single point failure-tolerance. As a result of the highly specific Huygens mission (limited duration and no access by telecommand after separation), a very safe redundancy scheme has been selected. The CDMS comprises: - two identical CDMUs - a triply redundant Mission Timer Unit (MTU) - two mechanical g-switches (backing up MTU) - a triply redundant Central Acceleration Sensor Unit (CASU) - two sets of two mechanical g-switches (backing up CASU) - a Radial Acceleration Sensor Unit (RASU) with two accelerometers - two Radar Altimeter proximity sensors, each comprising separate electronics, transmit antenna and receive antenna The two CDMUs each execute their own POSW simultaneously and are configured with hot redundancy (Chain A and Chain B). Each hardware chain can run the mission independently. They are identical in almost all respects; the following minor differences facilitate simultaneous operations and capitalise on the redundancy: - telemetry is transmitted at two different RF frequencies - chain B telemetry is delayed by about 6 s to avoid loss of data should a temporary loss of the telemetry link occur (e.g. from an antenna misalignment as the Probe oscillates beneath the parachute). Each CDMU chain incorporates a health check (called the Processor Valid status) which is reported to the experiments in the Descent Data Broadcasts (DDBs). A chain declares itself invalid when two bit errors in the same memory word, an ADA exception or an under-voltage on the 5 V line occur within the CDMU. Command and Data Management Unit (CDMU) --------------------------------------- Each CDMU includes a MAS 281 16-bit 1750A micro-processor running at 10 MHz, with 64 kword PROM storing the POSW and 64 kword RAM used for the POSW and other dynamic data when the CDMU is on. A Memory Management Unit is implemented to provide memory flexibility and some growth potential. Direct Memory Access (DMA) is provided to facilitate data transfer between the memory and the input/output registers, thus relieving the microprocessor of repetitive input/output tasks. The RAM-stored program memory is protected against single error occurrence by an Error Detection And Correction (EDAC) device, which detects and corrects single bit errors and reports any double bit errors to the Processor Valid function. TM/TC management is based on an internal On-Board Data Handling (OBDH) bus in order to standardise the internal interfaces, which are based on the classical Central Terminal Unit (CTU) and Remote Terminal Units (RTUs) approach. In addition to conventional CDMS functions, the CDMUs implement the following Huygens-specific functions: - the arming timer function sends pyro and arming commands following a specific hardware-managed timeline, thus offering full decoupling from the POSW operation - the Processor Valid signal is sent to experiments via the Descent Data Broadcast (DDB), indicating the health of the nominal CDMU (unit A) - reprogrammability through the use of 16 kword of Electrically Erasable PROM (EEPROM), thus allowing patching of the POSW if necessary - the EDAC error count reports on internal data transfers - the capability, through specific 16 kword of RAM, to delay one telemetry chain. Mission Timer Unit (MTU) ------------------------ The MTU is used to activate the Probe at the end of the coast phase. To obtain a single point failure-free design, it is based on three independent hot redundant timer circuits followed by two hot-redundant command circuits. Two mechanical g-switches provide backup. MTU power is supplied directly via three 65 V supply lines, one for each Timer Board, from independent batteries. During the pre-separation programming activities, when the Probe is still connected to the Orbiter, all three Timers are programmed with the exact duration of the coast phase via serial memory load interfaces from one of the two CDMUs. Each of the three Timer Boards can be loaded independently from either CDMU. The programmed values can be verified by the serial telemetry channels. When programming is finished, the CDMUs and all other Probe systems except the MTU are turned off and the Probe is separated. During the coast phase of about 22 days, the programmed Timer register is decremented by a very precise clock signal. The MTU consumes about 300 mW during this period as only the necessary circuits (CMOS-based) are powered. When the Command Board majority voting detects either both g-switches active or at least two of the three 'time-out' signals received, five High Level Commands (HLCs) are issued sequentially from each Board to the PCDU in order to switch on both CDMUs. The timer then returns to a standby mode. The two g-switches, which ensure Probe wake-up in the event of atmospheric entry without the time-out signal from any of the Timer boards, are purely mechanical devices closed when deceleration reaches 5.5-6.5 g. Central Acceleration Sensor Unit (CASU) --------------------------------------- The CASU measures axial deceleration at the centre of the Experiment Platform during entry. The signal is processed by the CDMU to calculate the time for parachute deployment (T0). The CASU operates within 0-10 g and uses a scale factor of 0.512 V/g. Its main building blocks are: 1. Power circuit. Two hot-redundant input power lines make it single point failure-tolerant in both nominal and redundant power lines 2. Three accelerometer analogue signal conditioning blocks. A low-pass filter with a 2 Hz cutoff is used and the analogue output from each block is routed to both CDMUs. In addition, the design prevents failure propagation from one conditioning chain to the others, it withstands permanent short circuit conditions without any degradation, and it is single point failure-tolerant toward the input power supply line. Back-up detection of T0 is performed separately for both CDMUs by two pairs of mechanical g-switches in case the prime CASU system is inoperative. The threshold values for each pair of g-switches are 5.5 g and 1.2 g. Radial Acceleration Sensor Unit (RASU) -------------------------------------- The RASU measures radial acceleration at the periphery of the Experiment Platform. The signal is processed by the CDMU to provide the Probe spin rate for insertion into the DDB distributed to experiments. The RASU is designed to measure spin acceleration within 0-120 mg with a 41.67 V/g scale factor. The design is based on CASU's but includes only two accelerometers. Radar Altimeter Unit (RAU) -------------------------- The RAU proximity sensor uses two totally redundant altimeters operating with frequency-modulated carrier waves at 15.4 GHz and 15.8 GHz to measure altitude above Titan's surface, starting from about 25 km. Each of the four antennas (two per altimeter) is a planar slot radiator array providing an antenna gain of 25 dB with a symmetrical full beam width of 7.9 degrees. A continuous signal modulated in frequency with a rising and falling ramp waveform is transmitted; the received signal has a similar form, but delayed by the propagation time. Hence the range to target is proportional (with a linear frequency modulation ramp) to the instantaneous frequency shift between the transmitted and received signals. Received signal data are also provided to the Huygens Atmospheric Structure Instrument (HASI) to establish Titan's surface roughness and topography. Probe Data Relay Subsystem (PDRS) ================================= The PDRS is Huygens' telecommunications subsystem, combining the functions of RF link, data handling and communications with the Orbiter. It transmits science and housekeeping data from the Probe to the Orbiter's PSE, which are then relayed to the Orbiter CDS via a Bus Interface Unit. In addition, the PDRS is responsible for TC distribution from the Orbiter to the Probe by umbilical during the ground and cruise checkouts. It comprises: 1. two hot-redundant S-band transmitters and two circularly polarised Probe Transmitting Antennas (PTAs) on the Probe 2. a Receiver Front End (RFE) unit (enclosing two Low Noise Amplifiers and a diplexer) and two Probe Support Avionics (PSA) units on the Orbiter. The Orbiter's High Gain Antenna (HGA) acts as the PDRS receive antenna. In addition, as part of the Doppler Wind Experiment (DWE), two ultra stable oscillators are available as reference signal sources to allow the accurate measurement of the Doppler shift in the Probe-Orbiter RF link: the Transmitter Ultra Stable Oscillator (TUSO) on Huygens and the Receiver Ultra Stable Oscillator (RUSO) on the Orbiter. The PDRS electrical architecture is fully channelised for redundancy, except that TUSO and RUSO are connected to only one chain. Probe Support Equipment (PSE) ----------------------------- Receiver Front End (RFE) ------------------------ The RFE comprises: - two Low Noise Amplifiers (LNAs) linked to the Orbiter's HGA to amplify the acquired RF signal by 20 dB using two cascaded FET stages - two RF inputs: one linked to the HGA, the other via a coupler and used during checkout to link a dedicated transmitter output (on the Probe) to the RFE via the umbilical - a pre-selection filter (coaxial cavity type with six poles) - an isolator - an output attenuator (fixed value) In addition, owing to the HGA's shared use with the Orbiter, a band pass filter (the TX filter) and a circulator protect the LNA chain B by isolating the Orbiter's S-band transmissions and the Probe's S-band reception, which both use the HGA. These two modes are mutually exclusive. Probe Support Avionics (PSA) ---------------------------- The two RFE outputs are sent to the two PSAs, which perform detection, acquisition (based on a 256-point Fast Fourier Transform algorithm), tracking, signal demodulation and data handling & management. The PSA data handling architecture is divided between analogue and digital sections. The analogue section performs signal down-conversion from S-band to the IF frequency. The IF signal is quantised and the samples processed by the digital section. The digital section performs: - the Digital Signal Processing (DSP) function - the signal acquisition and tracking task based on FFT analysis and frequency acquisition - the Viterbi decoding of the digital signal and delivery of the decoded transfer frame to the data handling section at 8192 bit/s - the data handling task, which consists of: - transforming the received transfer frame into a telemetry packet - generating internal PSA housekeeping data (including the synthesised frequency information) in a packet format - controlling and managing communications with the Orbiter CDS via a Bus Interface Unit - distributing the telecommands from the Orbiter BIU interface. The digital section is composed of the following main modules: - the receiver digital module, comprising the UT1750 microprocessor, 8 kword RAM and 8 kword PROM, and the receiver signal processing ASIC - the interface digital module, using GaAs devices for Numerically Controlled Oscillator (NCO) and Digital to Analogue Converter (DAC) functions - the support interface circuitry module (SIC), which comprises: the 8 kword EEPROM to memorise software patches; the 32 kword PROM containing the Support Avionics Software (SASW) and the testing, telecommand, telemetry and umbilical interfaces; the MAS 281 microprocessor module used by the SASW - the BIU module that controls communications between the PSA and the Orbiter's 1553 bus. Probe Transmitting Terminal (PTT) --------------------------------- The PTT comprises two transmitters and two Probe antennas. Each transmitter includes Temperature Controlled Crystal Oscillator (TCXO) synthesiser and BPSK modulator modules and a 10 W Power Amplifier module using Automatic Level Control (ALC) for 40.2 dBm nominal output power (end-of-life, worst-case, including ageing). The reference oscillator for the Phase Locked Loop (PLL) synthesiser is either an (internal) Voltage Controlled Crystal Oscillator (VCXO) with a temperature compensating network or the (external) TUSO signal. The selection between these reference sources is made before separation from the Orbiter. The TUSO has priority unless a failure is detected before separation. The two transmitting antennas linked to the transmitters (dual chains without cross-coupling) are quadrifilar helix designs. The four spirals are fed at the bottom of the helix in phase quadrature. Left Hand Circular Polarisation (LHCP) is used for signal transmission at 2040 MHz and Right Hand Circular Polarisation (RHCP) for transmission at 2098 MHz. The minimum gain for the antennas, mounted on the Top Platform, is 0.9 dB at all Probe-Orbiter aspect angles between +20 degrees and +60 degrees. Probe data relay link budget ---------------------------- During Probe descent, starting from the time of atmospheric entry as predicted from Orbiter trajectory and Probe separation characteristics, the Orbiter HGA is controlled to track a fixed point on Titan's surface - the nominal touchdown point. Orbiter movement along its trajectory significantly reduces the 'space loss' due to link distance during the Probe's 2-2.5 h descent. However, if Huygens does not land at the nominal point, e.g. due to non-nominal entry parameters or zonal winds, the gain from the reduced distance is offset by the HGA's reduced gain from the off-axis angle of the Probe with respect to the HGA's boresight axis. The link budget worst cases occur at the beginning and end of mission. The link design attempts to equalise the BOM/EOM signal level margins. At BOM, the signal level is determined by the range, while the losses owing to off-axis pointing is mainly due to HGA pointing error and Probe delivery error (the additional dispersion arising from the entry phase is relatively minor). At EOM, however, the signal level is critically dependent on the descent duration: the off-axis pointing losses due to the Probe's lateral drift in the assumed Titan wind worsens with descent duration. Software ======== Concept ------- The Huygens software consists of that running in the Probe CDMS, referred to as POSW, and that within the PSA on the Orbiter, referred to as the Support Avionics Software (SASW). The POSW output telemetry is relayed via the SASW and then Cassini's CDS to the ground. Two copies of the data handling hardware (CDMU and PSA) run identical copies of POSW and SASW. The software is based on a top-down hierarchical and modular approach using the Hierarchical Object-Oriented Design (HOOD) method and, except for some specific low level modules, is coded in ADA. The software consists, as much as possible, of a collation for synchronous processes timed by a hardware reference clock (8 Hz repetition rate). In order to avoid unpredictable behaviour, interrupt-driven activities are minimised. Such a design also allows a better observability and reliability of the software. Limited reprogramming accommodates modifications and RAM failure recoveries. The processes are designed to use data tables as much as possible. Mission profile reconfiguration and experiment polling can therefore be changed only by reprogramming these tables. This is possible via an EEPROM. In order to avoid a RAM modification while the software is running (which can lead to unpredictable behaviour and unnecessary complexity), direct RAM patching is forbidden. The POSW communicates with the SASW in different ways depending on mission phase. Before Probe separation, the two software subsystems communicate via an umbilical that provides both command and telemetry interfaces. Huygens cannot be commanded after separation, and telemetry is transmitted to the Orbiter via the PDRS RF link. The overall operational philosophy is that the software runs the nominal mission from power-up without checking its hardware environment or the Probe's connection or disconnection. The specific software actions or inhibitions required for ground or flight check-out must therefore be invoked by special procedures, activated by the delivery of specific telecommands to the software. To achieve this autonomy, POSW's inflight modification is autonomously applied at power-up by using a non-volatile EEPROM. At power-up, the POSW validates the CDMU EEPROM structure and then applies any software patches stored in the EEPROM before running the mission mode. If the EEPROM proves to be invalid at start-up, no patches are applied and the software continues based on the software in the CDMU ROM. A number of other checks are also carried out at start-up (e.g. a DMA check and a main ROM checksum), but the software will continue execution attempts even if the start-up checks fail. POSW functions -------------- The POSW provides the following functions: Probe Mission Management - detecting time T0 as entry begins, based on the Central Accelerometer Sensor Unit signals - forwarding commands at the correct times to the subsystems and experiments according to the pre-defined mission timeline - computation of the spacecraft dynamical state from sensor readings - sending Descent Data Broadcasts to the experiments Telemetry Management - collecting and recording housekeeping data - generation of housekeeping packets from the housekeeping data - collecting experiment packets according to a pre-defined polling scheme - transmitting transfer frames to the PDRS Telecommand Management - reception of TC packets from the PSE (only while attached to the Orbiter) - execution of commands related to these TC packets - forwarding of commands to the experiments POSW operations --------------- Control of the Probe, involving the activation and forwarding of commands to experiments and subsystems, is driven by a pre-defined set of tables, the Mission Timeline Tables (MTTs), that define the actions to be performed as a function of time. The pre-T0 MTT is activated at Probe wake-up, and controls the Probe until the post-T0 MTT is activated by the POSW's detection of T0. The experiments perform most of their activities autonomously based on the mission phase data computed within the POSW and sent to all the experiments every 2 s as a Data Descent Broadcast packet. The DDB contains the time, spin rate (computed by the POSW from the RASU signal or, in the event of failure, from a pre-defined look-up table) and altitude (initially taken from a look-up table based on the time elapsed since T0, but later by processing RAU data). The telemetry management function involves the acquisition and transmission of Probe telemetry as standard packets. Whether they are housekeeping or experiment packets, they are all 126 bytes long and forwarded to the SASW in the form of transfer frames comprising header information followed by seven packets and then Reed-Solomon code words, making a total frame size of 1024 bytes. Housekeeping data are acquired from the subsystems (and from the software itself) at different rates according to a pre-determined packet layout, and are loaded into four packets every 16 s. One of the packet types is buffered and issued 6.4 min later as 'History' housekeeping. Experiment data are acquired according to a pre-defined polling strategy and the resulting packets are loaded into the transfer frames. The selection of an appropriate type of telemetry packet to include in each of the frame's seven slots is managed by the polling sequence mechanism on a major acquisition cycle of 16 s (equal to 128 Computer Unit Times) driven by the Polling Sequence Table (PST) and the Experiment Polling Table (EPT). The PST defines if housekeeping or experiment packets are to be included in the transfer frame currently under construction. However, it does not select which experiment is to be included. The EPT defines a prioritised scheme for the collection of experiment data. The table is invoked whenever the PST requests experiment data for the transfer frame and is read in a cyclical manner. It consists of a sequential list of the Huygens experiments, with the number of occurrences of each experiment in the table providing the polling priority. By this method, the CDMS and the POSW are protected against failure modes in the experiments that could affect the data production rates. Each experiment is guaranteed an opportunity to supply data at, as a minimum, its nominal data rate. Furthermore, this polling scheme automatically optimises the data return by reallocating the TM resource in the absence of a 'packet ready' status flag from an experiment when expected. Three EPTs provide different polling priorities during the descent's various stages, switching from one table to the next at a pre-set time. SASW functions -------------- The SASW's main purpose is to provide communications between the Orbiter and Probe. For the SASW, there is no difference between receiving Probe telemetry via the umbilical or via the RF subsystem. All the differences are handled by the PDRS receiver part of the PSA equipment. The SASW provides the following functions: Telecommand Management - reception of TC packets from the BIU interfacing with the Orbiter CDS - execution of commands related to these TC packets - forwarding TC packets to the CDMS (including experiment telecommands) while attached to the Orbiter Telemetry Management - collecting PSE housekeeping data - transmitting PSE housekeeping packets and modified CDMS frames to the Orbiter via the BIU SASW operations --------------- Communication between the SASW and the Orbiter CDS is via a MIL-STD-1553 bus using a BIU. Received telecommands are placed in BIU memory for the SASW to read; the SASW places telemetry packets in BIU memory for transmission by the BIU over the CDS bus. The SASW examines any received telecommands to determine their destination address. Those destined for the Probe (subsystems or experiments) are transmitted over the umbilical TC link. Those for the PSA are handled by the SASW. The SASW handles the reception of Probe telemetry via a Frame Data Interface (FDI). Telemetry from the Probe is transmitted to the SASW either by the umbilical RF link when the probe is connected or by the Probe Relay Link (PRL) after separation. The SASW also generates its own telemetry in the form of housekeeping packets, containing PSA status information, and status data collected from the PDRS subsystem. [From LEBRETONETAL2005]: Launch and Flight to Saturn --------------------------- The Cassini-Huygens spacecraft was launched from Cape Canaveral complex in Florida on 15 October 1997, with the probe mated onto the side of the orbiter. In this configuration, the orbiter provided electrical power to the probe through an umbilical connection. Commands and data were also exchanged by this route. During the seven-year journey to Saturn, the Huygens probe was subjected to 16 in-flight checkouts to monitor the health of its subsystems and scientific instruments. During these in-flight tests, maintenance was performed and calibration data were obtained in preparation for the mission at Titan. The special in-flight tests designed to characterize the communication radio link between the probe and the orbiter were especially important. In the first link test in 2000, a flaw was discovered in the design of the Huygens telemetry receiver on board the orbiter that would have resulted in the loss of a large fraction of the Huygens probe's scientific data during the actual mission at Titan. Originally the Huygens mission was planned to be executed at the end of the first orbit around Saturn. As a remedy to the radio receiver flaw, the first two orbits of the original mission were redesigned into three shorter orbits that enabled the Huygens mission to be carried out on the third orbit. The re-designed orbiter trajectory provided a Doppler shift on the probe–orbiter radio link that was compatible with the well-characterized receiver performance and it also smoothly reconnected with the already-designed post-Huygens orbiter four-year trajectory. As a bonus, the new trajectory allowed early orbiter observations of Titan's upper atmosphere in order to validate the so-called Titan atmosphere engineering model well before the Huygens probe release. It led to improvements in our knowledge of the structure and the composition of the upper atmosphere; in particular, it provided better constraints on the argon concentration and indicated that methane was not present in sufficient quantity to affect the probe entry adversely (that is, via excessive radiative heating). Indeed, the new mission scenario led to the Huygens mission being completely successful. This achievement was the culmination of more than 20 years of work and shows that the in-flight rework of the mission was necessary and was successfully implemented. Probe release ------------- In preparation for releasing the probe, the Cassini-Huygens spacecraft had been set on a Titan-impact trajectory. Following its release, the Huygens probe had no manoeuvring capability and had to function autonomously. The Huygens release trajectory was achieved via a 'probe targeting manoeuvre' with a speed adjustment of 12 m/s on 17 December 2004, followed by a 'probe targeting clean-up manoeuvre' on 23 December 2004. After the separation of the Huygens probe on 25 December at 02:00 UTC, Cassini performed an 'orbiter deflection manoeuvre', so that it would not crash into Titan, and a 'clean-up manoeuvre' for final adjustment of its trajectory. These were on 28 December 2004 and 03 January 2005 respectively and placed Cassini on the correct trajectory for receiving data from the Huygens probe during the descent. The responsibilities for meeting the probe's trajectory requirements were shared between NASA/JPL and ESA. The targeting of the probe, the NASA/JPL responsibility, was specified at an altitude of 1,270 km, very close to the atmosphere's upper layer, above which no significant drag was expected. From this point onward ESA was responsible for the probe's trajectory. The spring-loaded Huygens separation mechanism, called the Spin Eject Device, had three points of attachment to the probe. It provided a speed increment relative to the orbiter of 33 cm/s. The Spin Eject Device also imparted to the probe an anti-clockwise spin of 7.5 r.p.m. (when viewed from the orbiter). This provided inertial stability during the ballistic trajectory and atmospheric entry. Coast and probe 'wake up' ------------------------- The Huygens probe was set on a ballistic trajectory that took a little over 20 days. During this time, the probe was dormant, with only three redundant timers counting down to a specific time programmed to end 4 h and 23 min before the predicted entry. At this time, battery power was turned on and the on-board computers, their sensors (accelerometers, and later in the descent the radar altimeters), and the scientific instruments were energized according to the pre-programmed sequence. The probe 'woke up' as planned, at 04:41:33 UTC on 14 January 2005. The Huygens probe's receivers on board the Cassini orbiter were powered on from 06:50:45 to 13:37:32 UTC. The Huygens probe arrived at the 1,270 km interface altitude on the predicted trajectory on 14 January 2005 at 09:05:53 UTC, just a few seconds before the expected time. Entry, descent and landing -------------------------- The Huygens scientific mission proper took place during the entry, descent, landing and post-landing phases. The descent of the probe through Titan's atmosphere was controlled by parachutes. The aerodynamic conditions under which the main parachute had to be deployed were critical. The correct instant for parachute deployment (mission time event, T0) was determined by the probe on-board computers that processed the measurements from the accelerometers that monitored the probe's deceleration. Pyrotechnic devices fired a mortar that pulled out a pilot chute, which in turn removed the probe's back cover and pulled out the main parachute. Then, 30 s later, the front shield was released. It was expected that, by this time, the probe would have stabilized under the main parachute. During the entry phase, telemetry could not be transmitted by the probe until its back cover was removed. Thus, a limited set of engineering housekeeping data and the HASI science accelerometer data acquired during entry was stored onboard the probe for transmission to the orbiter after the radio link was established. Post-flight data analysis showed that only one of the receivers(channel B) was phase-locked and functioned properly. Channel A had an anomaly that was later identified as being due to the unfortunate omission of the telecommand to apply power to the ultra-stable oscillator driving the channel A receiver. Subsequent on-board events were determined by the on-board software that initiated a set of commands at times all related to the moment the pilot chute was released. These commands included switching on other instruments and the replacement of the main parachute by a smaller 'stabiliser chute' after 15 min, to ensure that the probe would reach the surface of Titan within the designed duration of the mission (150 min maximum for the descent under parachute). The actual duration of the descent following the T0 event was 2 h 27 min 50 s. During the first part of the descent, the probe followed the nominal time-based sequence with the instrument operations defined by commands in the on-board mission timeline. The later part of the descent sequence was optimized by taking into account the altitude measurements provided by two redundant radar altimeters. The altimeters were switched on 32 min after T0 which corresponded to an altitude of around 60 km. They provided altitude measurements to the on-board computers, which filtered and compared the measurements to the predicted altitude, in order to exclude erratic measurements at high altitude and to provide reliable measured altitude information to the payload instruments. This allowed for optimization of the measurements during the last part of the descent. The DISR measurements sequence was adjusted to measured altitude below 10 km and its lamp was switched on at 700 m above the surface. The HASI and SSP instruments were set to their proximity and surface modes at low altitude above the surface. The probe landed safely with a vertical speed of about 5 m/s and continued thereafter to transmit data for at least another 3 h 14 min, as determined by the detection and monitoring of the probe's 2.040-GHz carrier signal by the Earth-based radio telescopes. Throughout this time, Cassini was oriented to receive the two incoming radio signals from the probe by continuously pointing its high gain antenna to the predicted Huygens landing point. After listening for the longest possible duration of the Huygens probe's visibility, the orbiter was commanded to re-point its high gain antenna to Earth for transmission of the stored Huygens telemetry data. At that time, Cassini was at a distance of 1,207 million kilometres (8.07 AU) from the Earth (the one-way light-time was 67 min 6 s). The data were received by the ground stations of the NASA Deep Space Network (DSN) and eventually delivered to the Huygens Probe Operations Centre (HPOC) in ESA’s European Space Operation Centre (ESOC, Darmstadt, Germany) for science and engineering analysis. A 1-h margin was built into the orbiter sequence to cope with uncertainties as to when the orbiter would disappear below the horizon. As seen from the probe landing site, the orbiter actually set below the horizon at 12:50:24 UTC. The probe's channel A carrier signal was still being received on Earth by radio telescopes at the time of the planned completion of the observations, at 16:00 UTC (Earth received time), meaning that the probe was still operating at 14:53 UTC (Titan time). Post-flight analysis of the probe telemetry data indicates that the batteries probably became fully discharged at about 15:10 UTC, a mere 17 min after the Huygens radio signal was last verified on Earth. It is thought that the probe continued to function until the batteries were exhausted. Trajectory reconstruction ------------------------- The probe arrived at the 1,270 km interface altitude with the spin imparted at separation in the anticlockwise direction. No significant spin modification was observed during the entry. The spin decreased more than expected under the main parachute and unexpectedly changed direction after 10 min. The probe continued spinning in the unexpected direction (clockwise) for the rest of the descent. No explanation was found for this behaviour, which is still under investigation. The determination of the landing site coordinates is a complex and iterative task and requires several assumptions. At present, the best estimate, based on the combined Descent Trajectory Working Group (DTWG), DISR and DWE reconstruction, is a latitude of 10.3 degrees (+-0.4 degrees) south and a longitude of 167.7 degrees (+-0.5 degrees) east. Summary and discussion ---------------------- The probe and its scientific payload performed close to and sometimes beyond expectations. The in-flight modifications of the Huygens part of the mission, to cope with the receiver design flaw detected in 2000, was highly successful. The loss of data on channel A, due to a telecommand omission, was largely compensated for by the flawless transmission on channel B, with not a single bit missing until the radio link signal-to-noise decreased below the design limit of 3.3 dB, in the last 10 min of surface transmission, and the fact that the DWE scientific objectives were largely recovered by using data from the Earth-based radio telescope observations. Deceleration and load levels measured during the hypersonic entry were well within the expected limits and all prime systems worked well, with no need to have recourse to the two back-up systems (g-switches) that had also been activated. The parachute performance was within the expected envelope, although the descent time, at slightly less than 2 h 28 min, was only just within the predicted envelope of 2 h 15 min +- 15 min. The descent was rather smooth under the main parachute but rougher than anticipated during the first hour under the last parachute. A detailed profile of the atmosphere is being worked out from the scientific measurements to allow the parachute performance to be studied in detail. An exciting scientific data set was returned by the Huygens probe, offering a new view of Titan, which appears to have an extraordinarily Earth-like meteorology, geology and fluvial activity (in which methane would play the role of water on Earth). While many of Earth's familiar geophysical processes appear to occur on Titan, the chemistry involved is quite different. Instead of liquid water, Titan has liquid methane. Instead of silicate rocks, Titan has frozen water ice. Instead of dirt, Titan has hydrocarbon particles settling out of the atmosphere. Titan is an extraordinary world having Earth-like geophysical processes operating on exotic materials under very alien conditions. The Huygens data set provides the ground-truth reference for the interpretation of the remote observations of the Huygens landing site by orbiter instruments, and more generally the global observations of Titan. Future observations of the Huygens landing site by Cassini should allow us to place the local Huygens maps into their global context and are expected to tell us whether changes can be seen. Probe–orbiter synergistic studies are a key aspect for achieving the very ambitious Cassini-Huygens objectives at Titan. Channel A anomaly ----------------- The mission had two probe–orbiter radio link channels, which are referred to as channels A and B. Both transmitters (onboard the probe) and both receivers (onboard Cassini) were equipped with a temperature-controlled crystal oscillator (TCXO) which provided sufficient frequency stability (~10^-6) for telemetry. One of the channels (channel A) was additionally equipped with ultra-stable oscillators (USOs) that were needed for the Doppler Wind Experiment (DWE), which required a stable carrier frequency signal. As part of finalising the Huygens probe's configuration for its mission, it had been decided to use the channel A USOs instead of the TCXOs because the performance of the USOs had been very satisfactory during the seven-year cruise. The command to power on the USO on the receiver side was unfortunately omitted. As a result, the Channel A receiver onboard Cassini did not have a reference oscillator and was unable to lock on to the Huygens signal. Consequently, the frequency measurements for the Doppler Wind Experiment (DWE), together with the non-redundant telemetry data on Channel A, were lost. The loss of the DWE data was, fortunately, largely mitigated by the radio astronomy segment of the mission consisting of a network of ground-based radio telescopes. The Channel A carrier signal, driven by the probe's USO, was received by 15 radio telescopes and tracked for post-flight data analysis. Real-time Doppler tracking information was obtained through the two largest telescopes of the network: the NRAO R. C. Byrd Green Bank Telescope (West Virginia, USA) and the CSIRO Parkes Radio Telescope (New South Wales, Australia). Both telescopes were equipped with NASA Deep Space Network's Radio Science Receivers (RSR) operated by the Radio Science Group of the Jet Propulsion Laboratory. In addition, the other 13 radio telescopes recorded the Channel A carrier signal for non-real-time Doppler and VLBI analysis. Huygens mission timeline on 14 January 2005 =========================================== Activity Time (UTC) Mission time, t - t0 ------------------------------------------------------------------------ Probe power-on 04:41:18 -4:29:03 Probe support avionics power-on 06:50:45 -2:19:56 Arrival at interface altitude (1,270 km) 09:05:53 -0:04:28 t0 (start of the descent sequence) 09:10:21 0:00:00 Main parachute deployment 09:10:23 0:00:02 Heat shield separation 09:10:53 0:00:32 Transmitter ON 09:11:06 0:00:45 GCMS inlet cap jettison 09:11:11 0:00:50 GCMS outlet cap jettison 09:11:19 0:00:58 HASI boom deployment (latest) 09:11:23 0:01:02 DISR cover jettison 09:11:27 0:01:06 ACP inlet cap jettison 09:12:51 0:01:30 Stabilizer parachute deployment 09:25:21 0:15:00 Radar altimeter power-on 09:42:17 0:31:56 DISR surface lamp on 11:36:06 2:25:45 Surface impact 11:38:11 2:27:50 End of Cassini–probe link 12:50:24 3:40:03 Probe support avionics power-off 13:37:32 4:27:11 Last channel A carrier signal reception ~14:53 5:42:39 by Earth-based radio telescopes 16:00 (ERT) ------------------------------------------------------------------------- The second column gives the time in UTC (for the probe), while the third column gives the time relative to t0, where t0 is the official start of the descent associated with the pilot chute deployment event. 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