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.

This book is intended as a concise but broad overview of the engineering, science and flight history of planetary landers and atmospheric probes. Such vehicles are subject to a wide range of design and operational issues that are not experienced by ‘ordinary’ spacecraft such as Earth-orbiting satellites, or even by interplanetary flyby or orbital craft. Such issues deserve special attention, and we have attempted to bring together in one place brief discussions of many of these aspects, providing pointers to more detailed (but dispersed) coverage in the wider published literature. This volume also draws heavily on real examples of landers and probes launched (or, at least, where the launch vehicle's engines were started with that intention!).

More than 45 years have passed since the first vehicles of this type were designed. To a certain extent some past missions, of which there are over one hundred, may now be considered irrelevant from a scientific point of view, outdated from an engineering point of view and perhaps mere footnotes in the broader history of planetary exploration achievements. However, we believe they all have a place in the cultural and technical history of such endeavours, serving to illustrate the evolving technical approaches and requirements as well as lessons learned along the way. They stand as testament to the efforts of those involved in their conception and implementation.

The landers covered in this chapter have the ability to survive an initial landing impact, which may send the vehicle rolling and/or bouncing across the surface, and then commence operations having come to rest in whatever orientation is finally reached. Most achieve this by means of airbags to cushion and dampen the initial impact and subsequent rolling/bouncing motion, followed by the opening out of a system of ‘petals’ to bring the lander itself to its proper orientation for surface operations. The Ranger seismometer capsules are the exception to this; their impact damping was provided by the balsa-wood shell and liquid-bath system surrounding the experimental equipment, and the orientation being achieved by means of the natural position of the equipment within its liquid bath.

Typical payload experiments for such landers include cameras, meteorological, geological, geophysical and environmental sensors for investigation of the landing site. While some can be body-mounted on the probe, others may require deployment by means of masts, arms or a rover. In the case of the Mars Exploration Rovers, the pod landing stage itself plays no further role once the rover has rolled off.

Pod landers are particularly suited to ‘network science’, where simultaneous seismological, meteorological or other geophysical measurements are made at multiple locations. Such a network was the aim of the NetLander mission a network of four Mars landers to be carried on the CNES-led Mars Premier mission. The mission was cancelled in 2003 towards the end of Phase B of the project, however.

The Mars Pathfinder mission began as MESUR (Mars Environmental Survey), a 1991 proposal for a network of as many as 16 Mars landers to perform network science (meteorology and seismology on distributed sites) using nominally inexpensive landers. One prominent approach to reducing the unit cost of the landers was to use a semi-hard landing approach with airbags rather than a retrorocket system. The landing system proposed was sufficiently radical that a technology demonstration/flight validation was designed, originally MESUR Pathfinder, on which work formally began in 1993.

With the loss of Mars Observer and the onset of the Discovery programme in NASA, the Pathfinder concept was ‘adopted’ by the Discovery programme, and became the most widely cited example of the ‘faster, better, cheaper’ (FBC) approach (see McCurdy, 2001). NEAR technically was the first selected Discovery mission, but took rather longer to be built and reach its target. Note also that there were other FBC programmes within NASA, including the Small Explorer Earth orbiters, and the New Millenium technology validation programme. The success of some non-NASA projects like the Clementine moon orbiter, which came out of the Strategic Defense Initiative (the ‘Star Wars’ programme) also set the stage for the FBC era.

As an aside, one viewpoint of the background to the development of Pathfinder is described in Donna Shirley's book Managing Martians (1998). Andrew Mishkin's Sojourner (2004) gives a more detailed but narrower view, of the rover engineering development specifically.

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 carbon– carbon 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.

Missions to small bodies differ from those to larger worlds because the low surface gravity means that an orbiter (or rendezvous) spacecraft can approach close enough to perform a surface mission while hovering (with little or no thrust) and the speed of a landing can be very low. This blurs the distinction between orbiters and landers, and may enable orbiter spacecraft to survive landing, as shown by the landing of NEAR on asteroid Eros. Low gravity also means that a landing vehicle may risk being lost entirely on rebound from the surface, or ejected by outgassing in the case of a comet nucleus. Anchoring systems may thus be required. On the positive side, the low gravity also makes it easy to achieve mobility by jumping, and to perform ‘touch and go’ surface-sampling manoeuvres (e.g. Yano et al., 2003; Sears et al., 2004). Most small bodies are highly irregular, and their gravitational fields can be challenging environments in which to navigate. Dust thrown up from the surface (whether from natural cometary activity or the action of a spacecraft) is another hazard. Many small bodies, particularly comets, are in elliptical orbits and so experience wide variations of temperature and solar power production with time and surface location.

Phobos 1F

The Phobos project involved two large Mars orbiters, Phobos 1 and Phobos 2 (Sagdeev et al., 1988; TsUP, 1988).

Each of the missions or spacecraft in this part has been selected for description in greater detail because they have faced and overcome an unusual challenge in their design and/or mission. Collectively, the seven case studies cover: atmospheric probes and surface/sub-surface missions; worlds with and without atmospheres; low and high gravity environments, and both static and mobile elements.

This part of the book provides a basic description, key data and a drawing for all planetary atmospheric or surface vehicles launched, or attempted, from the earliest examples to 2007. Key references concerning the design, payload and results of each craft or mission are given in each case so that the reader may find more detailed information elsewhere. For the payload experiments, the names in parentheses indicate the Principal Investigators (PIs) or otherwise-titled responsible experimenters. Details of the particular experiments and the results obtained (if any) can in most cases be found by searching publications authored (or co-authored) by those named.

The many vehicles are divided into six categories, reflecting the way in which they encounter an atmosphere or surface.

• Destructive impact probes (where the mission is intended to end with the vehicle being destroyed on impact with the surface). These probes are discussed only very briefly, since they are not landers yet do play a role in planetary surface exploration.

• Atmospheric entry probes (where the vehicle's design is driven by its mission in the atmosphere).

• Pod landers (where the vehicle is designed to land initially in any orientation).

• Legged landers (where the vehicle is provided with landing gear).

• Payload delivery penetrators (where the vehicle decelerates in the sub-surface to emplace a payload).

• Small-body surface missions (where the vehicle operates in a low surface gravity environment). These can include many operations that are possible in low gravity, and various types of surface element.

The diagrams in this part of the book were drawn using information gleaned from a variety of sources.

These landers use a system of legs to cushion the landing and provide a stable platform for surface operations. With the exception of the Venera landers and the forthcoming Mars Science Laboratory (MSL) rover, all legged landers have three or four legs with footpads, and retro-thrusters perform final braking before landing. This was not required for the Veneras, whose terminal velocity at the surface was low enough (~8 m s− 1) such that the landing gear alone was able to provide sufficient damping. The landing gear was toroidal and we thus consider it as effectively a single ‘leg’. Mars Science Laboratory is due to make use of the rover's wheels as landing gear. A key feature of legged landers is that they must be the right way up for landing – beyond some tolerable limit such landers would topple over and fail. This attitude control must be performed during descent, usually by thrusters. Only for sufficiently thick atmospheres, such as that of Venus, can aerodynamic stabilisation be used.

Beyond those described here, future possible legged landers include robotic and crewed lunar landers from the US, robotic lunar landers from China and Japan, and a Mars sample-return mission.

Surveyor landers

The Surveyor landers performed soft landings on the Moon, largely as reconnaissance of the surface for the later Apollo landings. For more details see the Case Study, Chapter 21 (Figure 18.1).

A sobering thought experiment is to contemplate a world without electricity. Not only is electricity exploited as a convenient means of delivering mechanical or thermal energy to remote locations, but electricity is vital in information transmission and in sensing and control. Although the first planetary missions contemplated involved launching to the Moon a vehicle containing flash powder with which it would optically announce its arrival, and some early spacecraft used clockwork timers to sequence operations, every mission actually flown has been electrically powered.

In this chapter we first consider the overall requirements on the probe's power system, and how these requirements favour the various means adopted to meet them. The power supply and storage possibilities are then discussed, with particular reference to planetary probes. A general reference for power considerations is the book by Angrist (1982).

It is instructive to consider the electrical power requirements of various household devices to place spacecraft requirements in context. A modern PC may consume perhaps 200 W; a laptop perhaps an order of magnitude less. The Viking lander ran on 90 W. The Huygens probe's batteries supplied around 300 W for about 5 hours. The Sojourner rover had a solar array that delivered a mere 15 W.

System requirements

The total energy requirement of a mission (i.e. its integrated power requirement) is the most fundamental parameter for designing the power system.

In the late 1960s the Mariner 6 and 7 spacecraft sent back images of Mars that seemed to indicate that it was a dead, cratered planet, much like the Moon. The volcanoes gave us the first indications that this was a false impression. When Mariner 9 approached Mars late in 1971, the planet was engulfed in a global dust storm. The surface in the equatorial regions could barely be seen. There were, however, four dark spots, and in the center of each spot was a complex crater. The craters resembled terrestrial summit calderas, volcanic depressions that form by collapse as a result of withdrawal of magma from below. Most calderas look very different from impact craters because they form by collapse rather than by excavation, and because they are commonly the result of multiple events rather than a single event. The craters were clearly calderas and the dark spots were clearly volcanoes poking up above the surrounding dust storm.

We now know that Mars has had a long and complex history of volcanism, and that the planet is probably still active today. The largest volcanoes and the most extensive lava plains are in and around the region of Tharsis. Volcanism probably started very early in the planet's history. By the end of the Noachian, around 3.8 billion years ago, the Tharsis bulge, a volcanic pile 10 km high and 5000 km across, had been largely built.