Huygens

Among many early concepts for a Titan probe (e.g. Murphy et al., 1981b) it is not surprising that a Galileo-like architecture was envisaged. As initially proposed in 1982, the concept of the Cassini-Huygens mission was to be a joint effort between NASA and ESA, and NASA was to supply the Galileo flight spare probe, and ESA would provide an orbiter delivery vehicle. However, in many respects the Titan probe grew in scope and complexity, in part because of the international nature of the mission.

As the joint study progressed, the roles were reversed, and ESA studied designs for an entry and descent probe (Scoon, 1985). These studies led to some quite novel ideas (e.g. Sainct and Clausen, 1993), which in all probability would not have been explored had the probe development remained in the USA.

The probe changed from an initially spherical shell (the shape adopted by the Galileo probe) to a flatter design. This also opened up novel heat shield architectures, with options such as a beryllium nose cap and a jetisonnable carboncarbon decelerator (although in the end, neither of these concepts was adopted and a more technologically conservative heat-shield design was used - a prudent measure given the novelty of this mission for ESA).

The mass budget (Table 23.1) deserves some brief comment. In broad terms the mass breakdown is typical (e.g. with 15% of the mass devoted to power systems), although the front shield is rather conservative. Note the formidable harness mass, balance mass, and the need for significant mass for separation hardware.

The complexity of the atmospheric photochemistry on Titan required more sophisticated instrumentation (Table 23.2) than Galileo - specifically a gas chro-matograph-mass spectrometer for in situ chemistry measurements during descent, rather than the simpler mass spectrometer used on Galileo. Additionally, the prospect of getting at least close to the surface of Titan, and perhaps surviving contact with it, invited surface science instrumentation, and some kind of surface imager. The Huygens payload allocation defined in the Phase A study was some 40 kg.

Table 23.1. Huygens mass budget (kg)

Item Probe PSE1

Subsystems

Front-shield subsystem 78.75

Back-cover subsystem 16.13

Separation subsystem 11.40 10.29

Descent-control subsystem 12.13

Inner structure subsystem 41.41

Thermal subsystem 20.60 1.50

Electrical-power subsystem 44.73

Probe harness subsystem 12.61

Command and data-management subsystem 23.10

Probe-data relay subsystem 6.04 16.30

Experiments

Doppler wind experiment (DWE) 1.90 1.90

Surface-science package (SSP) 4.87

Gas chromatograph/mass spectrometer (GCMS) 17.2

Huygens atmospheric-structure instrument (HASI) 5.77

Descent imager/spectral radiometer (DISR) 8.07

DISR cover 3.63

Aerosol collector/pyrolyser (ACP) 6.18

Fasteners, etc. 0.95

Balance mass 2.95

Total 318.32 29.99

1 Probe support equipment on the Cassini orbiter.

The dense atmosphere, and the desire to perform scientific measurements from an altitude as high as possible, and certainly above 150 km, meant that a realistic mission would need to last some 2-3 h.

For a 2-3 h mission, primary batteries are the obvious energy source: LiSO2 batteries - indeed using the same cells as flown on Galileo - were selected. The energy budget incorporated a healthy margin, in part because of reasonable conservatism (this being Europe's first planetary probe, and Titan being an almost unknown object) and in part because the energy budget had to be coarsely-quantized - there could only be an integral number of batteries - which was reduced from an initial 6 to 5 in Phase B. Each battery comprises two strings of 13 cells in series.

The command and data handling requirements in some respects are fairly trivial - the CDMU acts as a bent pipe for experiment data, formatting the experiment packets (and some housekeeping information) into transfer frames broadcast by the telecommunications system.

As far as the sequencing of operations is concerned, events are tied to occur an interval after a deceleration threshold is encountered; in a sense, a clockwork

Table 23.2. Huygens payload experiments

Instrument

Allocated mass (kg)

Power (W) (typical/peak)

Energy (Wh)

Typical data rate (bit s_1)

DWE

1.9

10/18

28

101

SSP

3.92

10/11

30

704

GCMS

17.3

28/79

115

960

HASI

6.3

15/85

38

896

DISR

8.13

13/70

42

4800

ACP

6.3

3/85

78

128

1 Housekeeping only (i.e. not collected as packets from the experiment, but voltages, temperatures, etc., as are recorded in other probe subsystems).

2 Note that this is lower than the value recorded in the system mass budget, suggesting this experiment exceeded at least its initial allocation.

3 This does not include the cover, which was added at a later stage.

1 Housekeeping only (i.e. not collected as packets from the experiment, but voltages, temperatures, etc., as are recorded in other probe subsystems).

2 Note that this is lower than the value recorded in the system mass budget, suggesting this experiment exceeded at least its initial allocation.

3 This does not include the cover, which was added at a later stage.

timer like those on the earliest probes could perform this function. On Huygens the function is implemented by a pair of computers (using MAS 281 silicon-on-sapphire radiation-hard processors). In the latter part of the descent, events are referenced to an altitude determined by two redundant radar altimeters; should they fail, the sequence reverts to a 'time-altitude table' based on a model descent profile. The CDMS also acts as a conduit for reprogramming experiment software and operating cruise checkouts - these functions became of critical importance when the probe mission had to be redesigned following the discovery of poor receiver performance.

The descent control system (Neal and Wellings, 1993; Underwood, 1993) comprises three separate parachutes - a 2.6 m pilot chute, deployed through a breakout patch in the aft cover of the probe by a mortar. This inflates and stabilises the probe, and then pulls off the aft cover and the bag enclosing the main parachute, some 8.3 m in diameter. This slows the probe to a speed of around 50ms~ . The ballistic coefficient of the probe plus main parachute is significantly lower than that of the front shield, which is then allowed to fall away after firing explosive attachment bolts.

The descent would take some 8 h under the main parachute, so this is released after approximately 15 minutes, the bridle holding the main chute being released and a third (3 m) stabiliser chute being inflated. This parachute remains attached until and after surface impact at 4.67 ms~x.

All the parachutes are of the disk-gap-band design, owing to that design's strong space heritage and relatively good damping performance. The riser, bridle, etc. are made of Kevlar, and the canopy itself of nylon.

Consideration was given to metallizing these components to prevent differential charging, but the thermomechanical difficulties in doing so, the possible degradation of communication performance, and the possibility of enhancing ambient electric fields to discharge levels ('probe-induced lightning') argued against doing so.

The thermal design of the probe reflects the several different environments it encounters. First, the 22-day coast in the Saturnian system after release from the orbiter would allow the probe, in a totally dormant state apart from the operation of three redundant clocks, to become unacceptably cold. The probe therefore includes 35 radioisotope heater units. These, and an envelope of multilayer insulation, assure an acceptable radiative equilibrium in free space at 10 AU.

The probe would also get cold during the atmospheric descent, where the thick and cold atmosphere would quickly remove heat from the probe. Conventional multilayer insulation does not insulate well in the presence of an atmosphere, thus a layer (some 10-15 cm thick) of a closed-cell foam (Basotect) retards heat leak from the probe. Note that the probe is dissipating several hundred watts during its descent.

Initially, the foam was applied in discrete panels, each wrapped in a plastic coating, with only millimetre gaps between the panels. It was found with some surprise during testing that substantial convective heat transfer could take place, even in these small gaps, compromising the performance of the insulation. A new packaging technique eliminated the problem, but this episode yet again stresses the importance of testing.

A final, perhaps unexpected, driver on the thermal design arose somewhat late in the project, when the mission design for the interplanetary trajectory to

Saturn was revised to incorporate two Venus flybys, such that Cassini and the

Huygens probe it carried would be exposed to a solar flux of some 3800 W m .

The prime strategy here was to use the Cassini orbiter's 4 m high-gain antenna as a sunshade, to shield both the probe (whose battery performance would be degraded by high temperatures during cruise) and the orbiter's sensitive instrumentation from the solar flux. Even then the equilibrium temperatures would be undesirably high, and thus the MLI coating of the probe incorporates a 'radiative window' - a hole on the antisun side which allows an extra radiative loss of heat to lower the temperature at this point rather more.

The telecommunications system was required to return a modest data rate during the descent (initially the data rate was expected to increase from about 1kbps at the start of the descent, when the orbiter would be some 100000 km away, to about 8 kbps at the end of the mission). Two data links were included, to eliminate single-point failures, and the original intent was that data would be sent redundantly on both channels, with one channel delayed by several seconds.

The rationale here is that swinging under the parachute, or some other break in the datastream, would be transient (less than those few seconds), thus the second staggered channel would allow recovery of any data lost during a transient on the first channel.

In 1992, the Cassini mission suffered a heavy descope, as a result of budget pressure from Congress, which essentially pitted the science missions CRAF and Cassini against the International Space Station. In the end, Cassini and the station survived, but CRAF was deleted and Cassini seriously descoped - the most prominent effects being a deferment of software development and the deletion of the scan platforms supporting the science instruments. Also deleted was the probe relay antenna, a dedicated dish that would track the probe. Instead, the whole spacecraft would be slewed, and the probe relay receiver would use the body-fixed 4 m high-gain antenna.

This change required the mission (and in particular the orbiter delay time) to be re-optimised, with the ODT being raised to some 5.2 h. Remarkably, this permitted a data rate (largely due to the size of the HGA) of some 8 kbps per channel, for the whole mission.

The structure of Huygens is moderately simple (see Figure 23.1) - almost all units are bolted onto a large honeycomb disc, the 'experiment platform' (73 mm thick). This is attached via a number of insulating fibreglass brackets to an exterior metal ring, to which the thin metal shell is attached. Also mounted on this ring is the front shield (since clearly the aerodynamic loads from the shield must be transmitted to the bulk of the probe mass, which is attached to the

Figure 23.1. Probe exploded view. Note the small spin vanes ringing the descent module's foredome.

experiment platform). This ring also attaches, via explosive bolts, to the spin-eject device. This is the load-bearing structure transmitting ground support and launch loads to the experiment platform; at probe separation a set of springs and rollers push the probe away at about 30cms_1 (pulling three 19-pin umbilical connectors apart) and set it spinning gently (— 7 rpm) for attitude stability during the coast phase.

Some elements, notably the parachutes and the probe antennae, are attached to a second, smaller honeycomb platform that forms the upper surface of the probe in its descent configuration. This top platform is connected to the experiment platform via a set of titanium struts whose principal function is to carry the parachute inflation loads.

The descent module's shell does not carry significant mechanical loads - it attaches to the experiment sampling ports on the bottom of the probe, and attaches to the foam insulation. In principle it also acts as a shield against electromagnetic interference, notably possible lightning or electrostatic discharge on Titan. It is made from a pressed aluminium alloy about 2 mm thick, with stiffening plates added.

One extra feature on this structure is the presence of 36 spin vanes, small wings protruding radially and a few degrees from vertical. As the probe descends, these vanes exert a slight torque rotating the probe in a horizontal plane. The probe's rotation is decoupled from that of the parachute by means of a swivel in the parachute riser.

In equilibrium, the vanes at an angle T (—tan T) would exert no net torque if the probe were rotating at a rate w = TV/R, where V is the descent rate and R is the radius around which the vanes are mounted. At this rotation rate, the airflow on the vanes would have zero incidence and thus there is no torque. In reality, since the swivel exerts a small retarding torque for non-zero rotation rates, the steady-state spin rate will be slightly lower than this.

However (as for the thermal equilibrium of a satellite in a low orbit) dynamic effects are important. The spin-up time of the probe, defined by the moment of inertia of the probe divided by the derivative of vane torque with spin rate, is not negligibly small compared with the descent time.

Thus from some initial (and unknown, since nondeterministic spin torques may occur due to uneven ablation during the entry phase) spin rate, the spin will slowly tend to a value given by the expression above, but will take some tens of minutes to reach that rate. That equilibrium rate is itself changing, as the probe descent rate drops with time.

The entry protection system's most obvious feature is a 2.7 m diameter front shield. This decreases the ballistic coefficient to a level that reduces the peak heating rate to levels that are tolerable by the thermal-protection material on the front surface of the shield. It also ensures that the probe decelerates to a Mach number low enough to permit deployment of the parachute at an altitude consistent with the scientific requirements. (For mass reasons, the shield was reduced from an initial 3 m diameter during phase B; while a smaller front shield gives a higher ballistic coefficient, the incremental area is relatively 'expensive' since rather higher structural rigidity is required - not only the mass of the outboard 0.3 m is saved, but also the additional stiffness needed inboard to support that mass.)

The relatively modest entry heat loads afforded by the large scale height in Titan's atmosphere allow lighter thermal protection to be used than for Galileo. The material used is AQ60, a French resin-doped silica fibre tile.

The peak deceleration during entry (with the relatively steep angle of "64°, and a speed of about 6kms"x at the 1270 km entry interface) was expected to be around 12 g. Stagnation point heat loads peak at around 600 kWm"2, divided approximately equally between radiative and convective fluxes. An attempt was made to observe the radiated emission of entry from the ground (Lorenz et al., 2006) but this was unsuccessful.

In fact, during the early phases of the project, radiative heat loads were re-evaluated in model studies and found to be rather (x2!) larger than originally anticipated. The radiative heat loads are also sensitive to the composition of Titan's atmosphere, which was not exactly known. The radiative emission depends on both the argon and methane abundance. The variation with methane mole fraction is nonmonotonic - initially it increases, as the availability of CN-radiating molecules increases (e.g. by 20% between 2 and 3% CH4); above some amount (3-6%, depending on the argon abundance) the endothermicity of CH4 dissociation takes over and lowers the temperature of the shock layer. Increasing argon abundance increases the electron-number density in the shock layer, which results in a more efficient population of the excited CN states and hence an increase in the radiative flux - for 0-10% argon, the effect is again a 20% increase in flux. These variations underscore that sophisticated entry aero-thermodynamic calculations involving nonequilibrium chemistry and radiative heat transfer, together with as narrow a range of compositions as scientifically justified, are needed to obtain a robust and efficient entry protection design.

The shock layer would radiate onto the back side of the probe (the rear face of the front shield, and the aft cover of the probe itself). These surfaces could be protected with a lighter (and cheaper) thermal protection system. The material used, Prosial, is a resin foam of silica bubbles. Its expense is considerably reduced because it can be sprayed onto the protected surfaces, while AQ60 must be carefully machined into tiles that can be precisely mounted on the front shield.

A significant design flaw in the Italian-built probe relay radio receiver (part of the ESA-supplied support equipment on the orbiter) surfaced some 3 years after launch, shortly after Cassini swung by the Earth. An end-to-end telemetry test was performed, with a DSN antenna performing the role of the probe; although the link behaved nominally, no data was recovered. The problem was eventually traced to inadequate bandwidth on the bit synchronizer in the receiver.

The problem and its solution were subtle and complex; in principle either a lower Doppler shift, or higher signal-to-noise, would improve matters, but the automatic gain-control switches in the receiver would switch in above preset signal levels and in fact degrade performance! These and several other aspects were hard-coded in firmware - straightforward solutions could have been easily implemented by telecommand had these parameters been left flexible.

An option that has not been necessary to implement is to substitute science telemetry packets with dummy 'zero' packets; the bit transition density in these packets would allow the Viterbi lock-state machine in the synchronizer to 'catch up'. Although this clearly results in loss of the effective telemetry bandwidth, this would be better than leaving which packets would be corrupted to chance.

It must be stressed that Huygens is not a lander (the original study called it a Titan Atmosphere Probe), although it has always been recognized that the probe may continue to transmit after surface impact (which occurs at the very modest speed of 5 ms-1). No explicit design features were introduced to permit or enhance surface operations, beyond the mission energy, link and thermal budgets including margin to allow at least 3 minutes of surface operation.

Among environmental hazards that were considered during the development phase were the possibility of lightning discharges in the Titan atmosphere (recall that all four Pioneer Venus probes suffered sensor failures during descent which have been attributed to electrical interactions with the atmosphere). Significant test effort was devoted to demonstrating tolerance of nearby strikes, and the probe incorporates discharge rods to alleviate any triboelectric charge buildup.

Parachute performance is always an area of concern on probe projects. Some scientific effort was expended in order to try to understand the likely constraints on wind gust amplitudes (it is impossible to guarantee parachute dynamic performance - getting the probe back to vertical within a few seconds so that the staggered radio links are not broken longer than the overlap period).

The mission took place on 14 January 2005, and can be judged to have been a great success. After parachute deployment, the presence of a transmission from Huygens was detected by radio telescopes on the ground (although a transmitter in which the carrier is suppressed would be more energy efficient in terms of data transmission, the existence of the unsuppressed carrier made this sort of radio science much more feasible). Analysis of the directly detected radio signals includes Doppler wind measurement, and very long baseline interferometry (VLBI), measuring the position of the spacecraft in the sky.

The parachute descent (Figure 23.2) took 2.5 hours, right at the maximum end of the predicted range of descent times (presumably due to parachute drag performance and/or deployment altitude - the atmospheric models seem to have predicted the actual pressure-density profile rather accurately). After impact, the probe continued working normally, with some 72 minutes of data received by Cassini before it passed below the horizon. The surface data included the impact (with a peak deceleration of — 15 g, indicating a soft solid surface) and images showing a cobble-littered plain, suggestive of past fluvial activity.

Some performances of the probe systems deserve comment. Two radio channels A and B, corresponding to entirely independent data handling systems, were carried, in part for simple redundancy, and in particular to guard against data loss from swinging under the parachute by having one stream delayed by several seconds with respect to the other. Channel A was equipped with an

2000

4000

6000

8000

Figure 23.2. Huygens descent profiles. Altitude (divided by 2 to use common scale) is shown by the solid line. After initial deceleration and front-shield release in the first —100s of descent, the probe was at a terminal velocity (dashed line) that steadily decreased with time as the probe descended into denser air. At 900 s (—115 km altitude) the main chute was released and the probe accelerated under gravity to a new terminal descent speed, which declines steadily until surface impact at 4.6 ms-1 at 8969 s.

2000

4000

6000

8000

Figure 23.2. Huygens descent profiles. Altitude (divided by 2 to use common scale) is shown by the solid line. After initial deceleration and front-shield release in the first —100s of descent, the probe was at a terminal velocity (dashed line) that steadily decreased with time as the probe descended into denser air. At 900 s (—115 km altitude) the main chute was released and the probe accelerated under gravity to a new terminal descent speed, which declines steadily until surface impact at 4.6 ms-1 at 8969 s.

ultrastable oscillator (which allowed the direct radio science from the ground), and the corresponding receiver on Cassini was also equipped with such an oscillator. However, although both transmitting chains operated nominally (with essentially no data loss from chain B), and both receivers were powered on, the oscillator for Cassini's chain A receiver was not, and so data modulated on that channel, and the on-board measurement of Doppler frequency, was lost.

The two radar altimeters initially generated false altitudes, roughly half the true value, until the altitude became low enough for the echo to suppress the false lock. There was minimal impact to the mission, but this was a salutary lesson that comprehensive testing is needed of such systems.

Another unexpected behaviour was strong short-period 1 Hz) motion of the probe - buffeting. Relatively little parachute swinging was noted, but the more rapid motions dominated accelerometer and tilt sensor data, making it difficult to retrieve information on wind gusts. Again, analysis of drop-tests on Earth would have helped anticipate (or alleviate) such effects. Additionally, the spin history of the probe (Figure 23.3) appears not to have been as anticipated.

Figure 23.3. The expected spin-rate profile of the probe during descent (solid line) and that reconstructed from various datasets (Lebreton et al., 2005, dots). The reason for the reversal in spin direction is not yet understood.

Figure 23.3. The expected spin-rate profile of the probe during descent (solid line) and that reconstructed from various datasets (Lebreton et al., 2005, dots). The reason for the reversal in spin direction is not yet understood.

The location of the landing site was determined by combining the knowledge of the entry state with Doppler measurements and on-board altitude data. This iterative process yielded a location within 5 km of the location determined by correlating surface features seen by the descent imager with a map made by an imaging radar on Cassini about 10 months later. Landing co-ordinates were estimated as 10.2° S, 192.4° W.

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