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.

The DS-2 mission was the second ‘Deep Space’ mission in NASA's New Millennium technology validation programme (Smrekar et al., 1999). It was to demonstrate miniaturized penetrators to enable subsurface and network science. The spacecraft that flew were radically smaller – by two orders of magnitude – than anything NASA had previously flown to the planets. The project cost a remarkably modest $29.6 million.

The original concept anticipated deployment at low latitude on Mars, and a payload including a microseismometer. As the mission evolved, and the delivery opportunity as a ‘piggyback’ payload on the Mars Polar Lander emerged, the mission concept had to change. In particular, the low-temperature environment at high latitudes on Mars reduced the expected energy capacity of the batteries (and thus the penetrators' lifetime) to the point where it was no longer likely that worthwhile seismic data would be acquired.

The new payload therefore centred on measuring the volatile content of the high-latitude soil. The same thermal environment that eroded the energy capability of the mission also made it likely that water might be trapped as ice in the soil.

Entry performance was driven by the entry conditions (at 6.9 km s−1 with a flight path angle of −13.1°, as for MPL) and the allowed flight parameters (velocity, angle of incidence) at impact (Braun et al. 1999b).

Before journeying through the various specific engineering aspects, it is worth examining two important subjects that have a bearing on many more specific activities later on. First we consider systems engineering as the means to integrate the diverse constraints on a project into a functioning whole. We then look at the choice of landing site for a mission, a decision often based on a combination of scientific and technical criteria, and one that usually has a bearing on the design of several sub-systems including thermal, power and communications.

Systems engineering

Engineering has been frivolously but not inaptly defined as ‘the art of building for one dollar that which any damn fool can build for two’. Most technical problems have solutions, if adequate resources are available. Invariably, they are not, and thus skill and ingenuity are required to meet the goals of a project within the imposed constraints, or to achieve some optimum in performance.

Systems engineering may be defined as

The modern discipline of systems engineering owes itself to the development of large projects, primarily in the USA, in the 1950s and 1960s when projects of growing scale and complexity were undertaken. Many of the tools and approaches derive from operational research, the quantitative analysis of performance developed in the UK during World War II.

The Surveyor spacecraft were a series of seven lunar soft-landing vehicles launched by the USA in the period 1966–1968. They were a second generation of lunar spacecraft, following the Ranger series that ran from 1961 to 1965, and paved the way for the later soft landings required for Apollo. The main aims of the Surveyor project were to accomplish a soft landing on the Moon, provide basic data in support of Apollo, and perform scientific operations on the lunar surface for an extended period. The Ranger 3, 4, 5 soft landing attempts having failed, Surveyor was to achieve the USA's first soft landings on another world. Orbital surveys by the Lunar Orbiter spacecraft complemented the in situ investigations by Surveyor.

Industrial studies for the project that became Surveyor began in mid 1960, with the Hughes Aircraft Company being chosen as prime contractor, under NASA JPL. The first launch was initially planned for late 1963 but a series of technical and programmatic issues forced an accumulated delay of nearly three years, by which time development of the Apollo landers was already well under way, and the Soviet Union had already made the first successful soft landing with Luna 9.

The main challenge for Surveyor was designing one of the first systems for performing a soft landing on another planetary body, with the associated terminal guidance and control problems of braking the spacecraft to land intact, and the then great uncertainty regarding the lunar surface's physical properties.

Atmosphere models

Clearly the design of heatshields and parachute systems requires assumptions on the density structure of the atmosphere to be encountered. Thus atmospheric models must be constructed as a design basis – these models must provide the extreme range of conditions likely to be encountered, since extremes in any direction may drive the design.

Where in situ data from prior missions is available (e.g. at Mars and Venus) this of course adds considerable confidence to the model. More generally, as for the first missions to Mars, Titan, Jupiter, etc., the major source of guidance is an atmospheric refractivity profile derived from radio-occultations by prior flyby or orbiter missions. The refractivity may be converted into a mass–density profile with some assumptions on composition. However, the altitudes probed by radio occultations are generally lower than those at which peak aerodynamic heating and deceleration occur, so some assumptions must be made in propagating those measurements upward. Some of these assumptions are rather robust, such as hydrostatic equilibrium, while others are less so.

There is in model development an inherent tension, just as in the development of a mission as a whole. The engineer designing the heat shield will just want a definitive answer to the question ‘what is the density at 500 km?’ (or whatever), while the scientist developing a model will wish to acknowledge the widest range of uncertainty – there may be intrinsic measurement errors in a refractivity profile, there are uncertainties in the assumed composition or other factors, there may be diurnal and seasonal variations, and variations with solar activity.

This chapter covers the final moments of descent towards, and contact with, a solid (or, as in the case of Titan, possibly liquid) surface. We deal with the issues of surviving impact to deliver a working spacecraft to the surface. This usually requires some sort of prior deceleration achieved during descent. Active guidance, navigation and control can also be performed to avoid hazards and locate a safe landing site. Having arrived, the impact may be damped within the vehicle alone, or by also using the deformability of the surface.

Targeting and hazard avoidance

Thus far, planetary landers have been flown ‘open loop’ in terms of their horizontal targeting with respect to the surface. While feedback control is employed to regulate descent rate to achieve close-to-zero speed at zero altitude, only the horizontal speed tends to be controllable, not the location.

The Mars Exploration Rovers incorporated a camera system (DIMES – Descent Image Motion Estimation System) to sense sideways motion, and a set of rocket motors (TIRS – Transverse Impulse Rocket Subsystem) to null the motion to maximize the probability of successful airbag landing; Surveyor and other lunar landers similarly used multibeam Doppler radar and thrusters to null horizontal motion. However, the latitude and longitude co-ordinates of the landing site were simply those that happened to be under the spacecraft when its height became zero. These were within an expected delivery ellipse specified by entry conditions and uncertainties, etc., but were not controlled.