This part of the book is intended to act as a guide to the basic technological principles that are specific to landers, penetrators and atmospheric-entry probes, and to act as a pointer towards more detailed technical works. The chapters of this part aim to give the reader an overview of the problems and solutions associated with each sub-system/flight phase, without going into the minutiae.

Overview and fundamentals

The descent through the atmosphere is often the only part of a planetary probe mission, as for example the Pioneer Venus and Galileo probes; on other missions it is just the last stage of a long journey prior to surface operations. The key parameters are the altitude of deployment – usually the altitude at which the vehicle ends its entry phase, as defined by some Mach number threshold – and the required duration of descent.

The duration of descent for an atmospheric probe is often dictated by an external constraint on the mission duration, such as the visibility window of a flyby spacecraft that is to act as a communications relay. This imposes an upper limit on the descent duration – it may be that (as for the Huygens probe) some part of that mission window is desired to be spent on the surface.

The instantaneous rate of descent (and thus the total duration) is determined at steady state by the balance between weight and drag. The former is simply mass times gravity; the latter depends on ambient air density, the drag area of the vehicle and any drag-enhancement device such as a parachute or ballute. The drag area is usually expressed as a reference area and a drag coefficient. Often these parameters and the mass are lumped together into the so-called ballistic coefficient β.

Often the dynamic pressure of descent is used to force ambient air into sampling instruments such as gas chromatographs.

The transfer of material that is not native to a planet has been happening over the history of the Solar System, with meteorite delivery being a common example of this interchange. With the development of rocket launchers capable of injecting objects into interplanetary trajectories, mankind joined Nature in being able to alter another planet's composition. Generally spacecraft and their associated hardware are designed and assembled so as to minimize the amount of debris that they carry. This chapter examines the problems associated with the unintentional delivery of living or dead organic matter to celestial bodies; so-called ‘forward contamination’. The topic is often referred to by the phrase planetary protection, and its scope includes not only the possible contamination of planetary bodies, but also the potential introduction to the Earth of materal from a non-terrestrial biosphere. Furthermore, the threat that planetary protection seeks to minimize is not restricted to the introduction of non-native organisms to another planetary body. Non-living material, such as DNA fragments and other complex bio-relevant molecules might trigger false-positives from equipment designed to detect extant or extinct life.

A practical definition of a living entity might be that the agent processes matter and energy in such a way that it can reproduce, and in doing so prosper in the face of environmental stresses. If the environment of the organism changes too radically then the organism may be killed or rendered dormant.

While much can be achieved by purely passive observations and measurements of a planetary lander's immediate environment, some key science requires the landed system to interact with the surface mechanically. This may involve the acquisition of samples of material, either to be returned to Earth or delivered to instrumentation internal to the lander. Other instruments, while external, require intimate contact with target rocks – these include alpha-X-ray, X-ray fluorescence or Mössbauer spectrometers, and microscopes. Other interactions may include mechanical-properties investigations using a penetrometer, or current measurements of wheel-drive motors.

Thus a variety of mechanisms have been operated on planetary surfaces, including deployment devices and sampling arms of various types, together with drills, abrasion tools and instrumentation. Soviet/Russian landers have tended to feature simple but robust actuators, usually simple hinged arms, and often actuated by pyro or spring. These include the penetrometers on the Luna and Venera missions. Lunokhods 1 and 2 carried a cone-vane shear penetrometer that was lowered into the lunar regolith and rotated by a motor, to measure bearing strength and shear strength. The rovers made 500 and 740 such measurements, respectively, during their traverses across the lunar surface.

A more sophisticated arm was flown on the Surveyor 3, 4 and 7 lunar landers (Figure 12.1). The Surveyor soil mechanics surface sampler (SMSS) was a tubular aluminium pantograph, five segments long, with a total reach of 1.5 m.

Following the success of the Mars Pathfinder project in 1997, there was a resurgence of interest in the deployment of an untethered rover on the surface of Mars. The concept of a semi-autonomous and freely roving vehicle was mooted as a follow-on to the Viking missions of the late 1970s. Almost twenty years were to pass before a rover was to be operated on Mars. After the Mars Pathfinder mission, NASA had proposed to send a rover equipped with a geology/chemistry payload, dubbed the ‘Athena’ suite, to Mars in 2001. Various constraints led to the redesign of the mission for a 2003 launch, although experiments of the payload were carried on the ill-fated Mars Polar Lander. In 2000 the Mars Exploration Rover mission was selected, with a launch-date flight three years later. This time, the Athena payload was to be duplicated, carried on two identical 174 kg rovers. Designated MER-A and MER-B, the spacecraft carrying the rovers were launched to Mars on separate Delta 2 boosters, making use of the favourable 2003 window for low-energy trajectories. The rovers on each craft were targeted to different regions of Mars. The MER-A craft, carrying the ‘Spirit’ rover, arrived on 4 January 2004 and was directed toward Gusev crater (14.5°S, 175.5°E) in the Aeolis region of Mars. This crater is the terminus of the fluid-cut Ma'adim Vallis, and Gusev was thought to host geological clues to the presence of water on Mars.

Payload delivery penetrators are bullet-shaped vehicles designed to penetrate a surface and emplace experiments at some depth. The basic technology for these has existed for several decades based largely on military heritage (e.g. Simmons, 1977; Murphy et al., 1981a; Bogdanov et al., 1988), however only in the mid 1990s did proposals for their use in Solar System exploration begin to be adopted for actual flight. In the US, Mars penetrators were studied for several years (and, indeed, field tested) as part of a possible post-Viking mission, while in the Soviet Union planetary penetrator work seems to have started in the mid 1980s.

Impact speeds range from about 60 to 300 m s−1. The resulting impact load experienced by penetrators as they decelerate in geological materials routinely exceeds 500 g, and terrestrial systems in the military field can be rated at 10 000 g or even 100 000 g, although the choice of components at these levels is severely limited (being more suited to the relatively simple job of triggering a detonator than making planetary science measurements). Additional impact damping may be included in the form of crushable material (e.g. honeycomb or solid rocket motor casing), sacrificial ‘cavitator’ spikes protruding ahead of the penetrator's tip (e.g. Luna-Glob high-speed penetrator concept, with speeds exceeding 1.5 km s−1) and gas-filled cavities (e.g. the Mars 96 penetrators).

Masses have ranged from the tiny DS-2 Mars Microprobes at 2.5 kg each (excluding aeroshell) to 62.5 kg each for the Mars 96 penetrators.

The Galileo mission (e.g. O'Neill, 2002; Bienstock, 2004; Hunten et al., 1986) was conceived early in the 1970s. In 1975 initial work started at NASA Ames for a Jupiter orbiter and probe for launch in 1982 on the Space Shuttle, with Jupiter arrival in 1985 after a Mars flyby en route. The project was transferred to JPL, and was approved by Congress in 1977. Development difficulties with the Space Shuttle led to a slip, and over the following years political pressures from various NASA centres led to several redesigns and different upper stages. Eventually, Galileo was set for a May 1986 launch on the Shuttle with a powerful Centaur upper stage. The Challenger disaster, however, interrupted the Shuttle launch schedule, and a re-examination of safety considerations ruled out the Centaur upper stage with its volatile cryogenic propellants. The revised mission, with a two-stage inertial upper stage (IUS) solid propellant upper stage would launch (after yet more delays) on October 18, 1989.

The low energy of the launcher then required Galileo to make one Venus and two Earth flybys to reach Jupiter. Although this trajectory afforded two asteroid flybys, the thermal design reworking needed to protect the spacecraft in the inner solar system led inadvertently to the failure of the high-gain antenna deployment mechanism, which drastically reduced the downlink performance during the scientific mission.

ESA's Rosetta mission was launched on 2 March 2004, and is destined to reach its target comet, 67P/Churyumov–Gerasimenko, in 2014. The lander of the Rosetta mission, named Philae, is expected to be deployed around November 2014, to make the first ever controlled landing on a comet nucleus. En route, the mission's interplanetary trajectory takes in four gravity assists, three at Earth and one at Mars, and two asteroid flybys. Having matched the comet's orbit, Rosetta will close in to perform a comprehensive remote sensing survey of the nucleus and its environment prior to final selection of the landing site and deployment of the lander.

The finally launched mission had evolved a great deal over several iterations since the initial conception of a ‘mission to the primitive bodies of the Solar System’ around 1985 as a cornerstone of ESA's new Horizon 2000 science programme (this was almost a year before ESA's Giotto spacecraft had encountered comet Halley). The mission plan has always incorporated a surface element, though initially this was to obtain a sample for return to Earth. Known briefly as the Comet Nucleus Sample Return (CNSR) mission, it had by 1987 been renamed Rosetta. By the end of 1985 a joint ESA/NASA Science Definition Team had been formed to define in detail the mission's scientific objectives; NASA being envisaged as a partner for ESA on the mission.

‘Radiation’ in the spacecraft environment context generally refers to subatomic particles in space. Of course, the Sun and other astrophysical sources yield electromagnetic radiation (hard UV, X-rays and gamma rays) that are somewhat damaging to materials and living things, but these effects are generally small. In this chapter we discuss briefly the sources of energetic particles and their effects on spacecraft systems (Trainor, 1994); effects on living things are discussed in Section 14.3

Note that because the missions of entry probes and landers tend to be short, and the radiation environment at or near a planetary surface is more benign than in orbit, the radiation hazard is generally not as significant a concern as it is for orbiters. Landers on airless bodies (the Moon, Mercury, and especially Europa) may be exceptions, due to secondary radiation from the surface. However, all landers will need a radiation tolerance in that they spend time, perhaps many years, in the space environment.

There are four principal sources of radiation that must be considered. First is any radiation source carried by the spacecraft, such as a radioisotope thermoelectric generator (RTG), radioisotope heaters or sources associated with instruments such as X-ray fluorescence spectrometers. A characteristic of RTGs is their neutron flux.

A second source is galactic cosmic rays (GCRs). These are high-energy particles, usually nuclei of high atomic number (‘heavy-Z’ or ‘high-Z’ particles) from astrophysical sources.

There are two fundamental arrival strategies – from a closed orbit (circular or otherwise) around the target body, and from a hyperbolic or near-linear trajectory directly to the surface.

Landing places some significant requirements on the thrust capability of the landing propulsion. Obviously the thrust-to-weight ratio (in that gravity field) must exceed unity if the vehicle is to be slowed down. The ΔV requirements will depend significantly on the trajectory and thrust level chosen, and can in the case of a hover, be infinite; a lower bound is given by the impulsive approximation analogous to the Hohmann transfer between coplanar orbits – first an impulse is provided to put the vehicle on a trajectory that intersects the surface, on the opposite side in the case of a descent from orbit. A second impulse can then be applied to null the velocity at the impact site.

In practice the trajectory of the vehicle, the performance of the propulsion system and the topography of the target body are inadequately known for such a strategy to be performed open-loop, except in the case of landing on very small bodies where the orbital and impact velocities are low enough that the second, arrival ΔV can be safely provided by impact forces rather than propulsively. Thus some sort of closed-loop control is needed.

Compensation for varying propulsive performance (both due to engine performance variations, especially if feed pressure may vary in blowdown mode, and due to the progressively reducing mass of the vehicle) can be achieved by monitoring the spacecraft acceleration with onboard accelerometers.

The mission of a destructive impact probe ends successfully with a vehicle (or even just a passive projectile) being destroyed on impact with the surface of another world. The first destructive impact probe was Luna 2, which, along with the launcher's upper stage, impacted the Moon in 1959. Luna 2 hit the surface at 3315 m s− 1 (Blagonravov, 1968), demonstrated by the loss of the radio signal. Rangers 6–9 impacted the Moon a few years later, obtaining (in the case of 7, 8, 9) close-up images of the lunar surface prior to impact at 2620–2680 m s− 1 (e.g. Schurmeier et al., 1965; Hall, 1977). The craters made by these impacts were subsequently found in Lunar Orbiter and Apollo images. Discarded Apollo lunar module ascent stages and Saturn IVB rocket stages impacted the Moon and proved useful as artificial, well-characterised seismic sources (Latham et al., 1970, 1978).

Many years later, Lunar Prospector ended its successful mission by impacting the lunar surface at 1700 m s− 1, in an attempt to detect water ice by means of telescopic observations of the ejecta plume from Earth. No plume was seen, however, but the exercise resulted in calculations of possible H2O ejecta cloud propagation that may be applicable to future events (Goldstein et al., 2001). The lunar orbiters Hiten and SMART-1 also ended their missions by impacting the lunar surface. NASA's LCROSS (Lunar CRater Observation and Sensing Satellite) is due to make another attempt to detect ice using the impact technique.

The system design of atmospheric probes is dominated by the atmospheric entry and descent/drift through the atmosphere, even if surface operations are possible (e.g. Venera 7, 8, Pioneer Venus Day Probe, Huygens). Common experiment types for such probes include entry accelerometry, radio science for tracking the probe's motion, sensors for atmospheric temperature, pressure and humidity, mass spectrometry, aerosol analysis, (spectro-) photometry and nephelometry.

First Soviet Venera and Mars entry probes

This section covers early (1961–65) Soviet entry probes to Venus and Mars designed and built by Korolev's OKB-1 design bureau (now RKK Energia), all of which failed during launch or cruise. In 1965 further development of the deep space and lunar probes was handed over to NPO Lavochkin's Babakin Space Centre (then called OKB-301). Very few published details exist concerning the entry probes.

1VA entry probes

The first launches of atmospheric entry probes were those of Venera 1, lost en route, and its ‘twin’ that failed to leave Earth orbit. The Venera 1 entry capsule was not designed to transmit a signal from the Venusian atmosphere. One could thus argue that these 1VA entry probes should be classed as ‘destructive entry probes’ rather than an atmospheric entry probe in the modern sense of the phrase. The carrier spacecraft part of the 1VA probes were somewhat similar to those of the two Mars 1M craft, which were lost in launch failures in October 1960 (Figure 16.1).

The challenges involved in designing optimal trajectories for planetary landers or atmospheric probes are shared by many other types of spacecraft projects. Spacecraft, at least for the foreseeable future, have to be launched from the Earth's surface and then placed on a path that intersects the orbit of the target body. How this is achieved depends on the mission of the spacecraft and its associated cost and design details.

The launch environment

Spacecraft have been delivered to space on a wide variety of launchers, all of which subject their payloads to different acoustic, dynamic and thermal regimes. These parameters vary with the size and nature of the launcher, yet the complex launch vehicle industry often makes it difficult to isolate a preferred launcher type for a given mission. In Table 2.1 pertinent features of current launch vehicles are shown with data taken from their user manuals; the launcher market currently has over a dozen vehicles capable of lifting interplanetary payloads. Costs are not listed as many of the vehicles offer dual manifest capability, or other partial-occupancy accommodation (such as Ariane's ASAP) which can make heavy launchers and their capability available to even modestly funded missions.

Of particular interest are the mass values shown for the parameter C3. This quantity is the square of the hyperbolic escape speed; the speed an object would have upon leaving the influence of a gravitating body. Paths with a C3 greater than 0 km2 s−2 are trajectories which never return to their origin.

Telecommunication is one of the most important functions of entry probes: it transmits to Earth all the science and engineering data that are the main goal of the mission. Tracking of the probes is another function that can help to analyse the probe dynamics during the entry and descent, providing independent science data on the atmosphere of a planet.

During entry, if communications are to be attempted at all, only status tones or very low data rates are possible. This is because the attitude during entry and descent may be very dynamic, preventing pointing of high-gain antennas. Depending on the wavelength of the communication link and the aerothermochemistry of the plasma sheath, transmissions may be completely blocked for a short period (the entry ‘blackout’).

During the highly dynamic entry phase data rates in direct-to-Earth (DTE) links are usually very small due to the great distance to the Earth and the use of low-gain antennas on probes. A relay link (Figure 10.1) uses a much shorter distance to the relay orbiter to boost the received signal strength though using a less efficient receiving antenna than on Earth. The probe data received on the orbiter is re-transmitted to the Earth using the high-gain antenna of the orbiter.

Motion of the probe affects the frequency, amplitude and phase of the signal at the receiving station. The entry process includes phases that are significantly different from the point of view of the communications link.

Planetary probes present a very diverse range of structural problems and solutions. This is in contrast to free-flying spacecraft (i.e. satellites and deep-space probes) which generally have a simple box or drum structure because there is only a single dominant loading (launch). On the other hand, landers and probes can range from resembling spiders to cannonballs, with the range generally being driven by thermal as well as structural requirements. Landers may be spidery open frames with equipment boxes bolted on, like the Surveyor landers; in contrast, entry probes for hot, deep atmospheres are constructed as pressure vessels and have thus been spherical in shape.

On most satellites the largest accelerations and thus structural loads are encountered during launch (typically 5–10 g): however, entry probes to Venus or Jupiter may encounter decelerations of 100–500 g. In such situations, load paths must be kept as short as possible to minimize the structural mass. The Pioneer Venus and Galileo probes (which had thermal constraints) used thick-walled pressure vessels supporting solid deck plates to which equipment was bolted. Spherical geometries are also appropriate where landing attitude is not initially controlled (e.g. Luna 9, 13; though note that the interiors of these spacecraft were pressurized, which also tended to favour a spherical design).

The Huygens probe did not need to exclude the atmosphere and therefore had an unsealed, thin-walled shell to preserve an aerodynamic shape and support light foam insulation.